JOINT AVIATION AUTHORITIES
AIRLINE TRANSPORT PILOT'S LICENCE
Theoretical Knowledge Manual
'- J-031
080 PRINCIPLES OF FLIGHT
Second Edition, First Impression.
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This 'earning materia' has been approved as
JAA com pliant by the United Kingdom
Civil Aviation Authority.
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© Oxford Aviation Services Limited 2001
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FOREWORD
Joint Aviation Authorities (JAA) pilot licences were first introduced in 1999. By the end of2002, all 33
JAA member states will have adopted the new, pan-European licensing system. Many other countries
world-wide have already expressed interest in aligning their training with the syllabi for the various JAA
licences. These syllabi and the regulations governing the award and the renewal of licences are defined
by the JAA's licensing agency, known as "Joint Aviation Requirements-Flight Crew Licensing", or JARFCL.
The introduction of JAA licences is, naturally, accompanied by associated JAR-FCL practical skill tests
(tests of flying ability) and theoretical knowledge examinations corresponding to each level of licence:
Private Pilot Licence (PPL), Commercial Pilot Licence (CPL), CPL with Instrument Rating and Air
Transport Pilot Licence (ATPL). The JAR-FCL skill tests and the ground examinations, though similar
in content and scope to those conducted by many national authorities, are inevitably different in detail
from the tests and examinations set by any individual JAA member state under its own national scheme.
Consequently, students who wish to train for JAA licences need access to study material which has been
specifically designed to meet the requirements of the new licensing system.
As far as the JAA ATPL ground examinations are concerned, the subject matter to be tested is set out in
the A TPL training syllabus contained in the JAA publication, ' JAR -FCL 1 (Aeroplanes)'. Inevitably, this
syllabus represents a compromise between the differing academic contents of the national ATPL training
syllabi it replaces. Thus, it follows that the advent of the new examinations has created a need for
completely new reference texts to cover the requirements of the new syllabus. This series of manuals,
prepared by Oxford Aviation Training and published by Jeppesen, aims to cover those requirements and
to help student pilots prepare for the JAA ATPL theoretical knowledge examinations.
Oxford Aviation Training (OAT) is one of the world's leading professional pilot schools. It has been in
operation for over thirty years and has trained more than 12,000 professional pilots for over 80 airlines,
world-wide. OAT was the first pilot school in the United Kingdom to be granted approval to train for the
JAA ATPL. As one of the most active members of the European Association of Airline Pilot Schools,
OAT has been a leading player in the pan-European project to define, in objective terms, the depth and
scope of the· academic content of JAA A TPL ground training as outlined in ' JAR -FCL 1 (Aeroplanes)'.
OAT led and coordinated this joint-European effort to produce the JAA A TPL Learning Objectives which
are now published by the JAA itself as a guide to the theoretical knowledge requirements of ATPL
training.
In less than two years since beginning JAA ATPL training, and despite the inevitable teething problems
that national aviation authorities have experienced in intr.oducing the new examination system, OAT has
achieved an unsurpassed success rate in terms of the passes its students have gained in the JAA ATPL
examinations. This achievement is the result of OAT's whole-hearted commitment to the introduction
of the new JAA licensing system and of its willingness to invest heavily in the research and development
required to make the new system work for its students. OAT has not only been at the forefront of the
effort made to document JAA ATPL theoretical knowledge requirements, but it has also produced
associated academic notes of the highest quality and created computer-generated and web-based ATPL
lessons which ensure that its students are as well-prepared as possible to succeed in the ground
examinations. OAT's experience and expertise in the production ofJAA ATPL training material make
this series of manuals the best learning material available to students who aspire to hold a JAA ATPL.
continued ....
Jeppesen, established in 1934, is acknowledged as the world's leading supplier of flight information
services, and provides a full range of print and electronic flight information services, including navigation
data, computerised flight planning, aviation software products, aviation weather services, maintenance
information, and pilot training systems and supplies. Jeppesen counts among its customer base all US
airlines and the majority of international airlines world-wide. It also serves the large general and business
aviation markets.
The combination of Jeppesen and OAT expertise embodied in these manuals means that students aiming
to gain a JAA ATPL now have access to top-quality, up-to-date study material at an affordable cost.
Manuals are not, of course, the complete answer to becoming an airline pilot. For instance, they cannot
teach you to fly . Neither may you enter for the new JAA ATPL theoretical knowledge examinations as
a "self-improver" student. The new regulations specify that all those who wish to obtain a JAA ATPL
must be enrolled with a flying training organisation (FTO) which has been granted approval by a JAAauthorised national aviation authority to deliver JAA ATPL training. The formal responsibility to prepare
you for both the flying tests (now known as "skill tests") and the ground examinations lies with your
FTO. However, these OAT/J eppesen manuals represent a solid foundation on which your formal training
can rest.
For those aspirant airline pilots who are not yet able to begin formal training with an FTO, but intend to
do so in the future, this series of manuals will provide high-quality study material to help them prepare
themselves thoroughly for their formal training. The manuals also make excellent reading for general
aviation pilots or for aviation enthusiasts who wish to further their knowledge of aeronautical subjects
to the standard required of airline pilots.
At present, the JAA ATPL theoretical knowledge examinations are in their infancy. The examinations
will inevitably evolve over the coming years. The manuals are supported by a free on-line amendment
service which aims to correct any errors and/or omissions, and to provide guidance to readers on any
changes to the published JAA ATPL Learning Objectives. The amendment service is accessible at
http://www.oxfordaviation.net/shop/notes.htm
OAT's knowledge of and involvement in JAR-FeL developments are second to none. You will benefit
from OAT's expertise both in your initial purchase of this text book series and from the free amendment
service. OAT and Jeppesen have published what they believe to be the highest quality JAA ATPL
theoretical knowledge manuals currently available. The content of these manuals enables you to draw
on the vast experience of two world-class organisations, each of which is an acknowledged expert in its
field ofthe provision of pilot training and the publication of pilot training material, respectively.
We trust that your study of these manuals will not only be enjoyable but, for those of you undergoing
training as airline pilots, will also lead to success in the JAA ATPL ground examinations.
Whatever your aviation ambitions, we wish you every success and, above all, happy landings.
Oxford, England. January 2002
PREFACE TO EDITION TWO, FIRST IMPRESSION
Edition Two of this work has been recompiled to give a higher quality of print and diagram. The
opportunity has also been taken to update the contents in line with Oxford Aviation Training's experience
of the developing JAA ATPL Theoretical Knowledge Examinations.
Oxford, England. September 2002
Textbook Series
Book
Title
1
010 Air Law
2
020 Aircraft General Knowledge 1
3
4
5
020 Aircraft General Knowledge 2
020 Aircraft General Knowledge 3
020 Aircraft General Knowledge 4
JAR Ref. No.
Subiect
021 01
Airframes & Systems
021
021
021
021
021
021
021
021
Fuselage, Wings & Stabilising Surfaces
Hydraulics
Landing Gear
Flight Controls
Air Systems & Air Condition ing
Anti-icing & De-icing
Emergency Equipment
Fuel Systems
01 01/04
01 07
01 05
01 06
01 08/09
01 09/10
0400
01 11
02102
Electrics - Electronics
021 0201
021 0202
021 0205
Direct Current
Alternating Current
Basic Radio Propagation.
021 00
Powerplant
021 0301
021 0302
Piston Engines
Gas Turbines
22
Instrumentation
02201
02203
02202
02204
Flight Instruments
Warning & Recording
Automatic Flight Control
Power Plant & System Monitoring Instruments
6
030 Flight Performance & Planning 1
031
032
Mass & Balance
Performance
7
030 Flight Performance & Planning 2
033
Flight Planning & Monitoring
8
040 Human Performance &
Limitations
9
050 Meteorology
10
060 Navigation 1
061
General Navigation
11
060 Navigation 2
062
Radio Navigation
12
070 Operational Procedures
13
080 Principles of Flight
14
090 Communications
15
Reference Material
PRINCIPLES OF FLIGHT
TABLE OF CONTENTS
Chapter 1
Overview and Definitions
Chapter 2
The Atmosphere
Chapter 3
Basic Aerodynamic Theory
Chapter 4
Subsonic Airflow
Chapter 5
Lift
Chapter 6
Drag
Chapter 7
Stalling
Chapter 8
High Lift Devices
Chapter 9
Airframe Contamination
Chapter 10
Stability and Control
Chapter 11
Controls
Chapter 12
Flight Mechanics
Chapter 13
High Speed Flight
Chapter 14
Limitations
Chapter 15
Windshear
Chapter 16
Propellers
CHAPTER 1 - OVERVIEW AND DEFINITIONS
Contents
Page
OVERVIEW ............................................................ 1 - 1
GENERAL DEFINITIONS
MASS ........................................................... 1 - 5
FORCE
WEIGHT
CENTRE OF GRAVITY
WORK .......................................................... 1- 6
POWER
ENERGY
KINETIC ENERGY
NEWTON'S FIRST LAW OF MOTION ............................... 1 - 7
INERTIA
NEWTON'S SECOND LAW OF MOTION
ACCELERATION
VELOCITY
MOMENTUM .................................................... 1 - 8
NEWTON'S THIRD LAW OF MOTION
GLOSSARy ............................................................ 1 - 9
LIST OF SYMBOLS .................................................... 1 - 14
SELF ASSESSMENT QUESTIONS ........................................ 1 - 15
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--~.=E3
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OVERVIEW AND DEFINITIONS
PRINCIPLES OF FLIGHT
1.1
OVERVIEW
The primary requirements of an aircraft are as follows:
(a)
A wing to generate a lift force.
(b)
A fuselage to house the payload.
(c)
Tail surfaces to add stability.
(d)
Control surfaces to change the direction of flight and,
(e)
Engines to make it go forward.
The process of lift generation is fairly straightforward and easy to understand. Over the years
aircraft designers, aerodynamicists and structural engineers have refined the basics and by subtle
changes of shape and configuration have made maximum use of the current understanding of the
physical properties of air to produce aircraft best suited to a particular role.
Aircraft come in different shapes and sizes, usually, each designed for a specific task. All aircraft
share certain features, but to obtain the performance required by the operator the designer will
configure each type of aeroplane in a specific way.
As can be seen from the illustrations on the facing page, the position of the features shared by all
types of aircraft - i.e. wings, fuselage, tail surfaces and engines varies from type to type.
Why are wing plan shapes different?
Why are wings mounted sometimes on top of the fuselage instead of the bottom?
Why are wings mounted in that position and at that angle?
Why is the horizontal stabiliser mounted sometimes high on top of the fin rather than on
either side of the rear fuselage?
Every feature has a purpose and is never includ~d merely for reasons of style.
1- 1
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OVERVIEW AND DEFINITIONS
PRINCIPLES OF FLIGHT
An aeroplane, like all bodies, has mass. With the aircraft stationary on the ground it has only the
force due to the acceleration of gravity acting upon it. This force, its WEIGHT, acts vertically
downward at all times.
w
Figure 1.1 The Force of Weight
Before an aeroplane can leave the ground and fly the force of weight must be balanced by a force
which acts upwards. This force is called LIFT. The lift force must be increased until it is the same
as the aeroplane's weight.
L
w
Figure 1.2 The Forces of Weight and Lift
1-2
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PRINCIPLES OF FLIGHT
OVERVIEW AND DEFINITIONS
To generate a lift force the aeroplane must be propelled forward through the air by a force called
THRUST, provided by the engine(s).
L
w
Figure 1.3 The Forces of Weight, Lift and
Thrust
From the very moment the aeroplane begins to move, air resists its forward motion with a force
called DRAG.
L
w
Figure 1.4 The Forces of Weight, Lift,
Thrust and Drag
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OVERVIEW AND DEFINITIONS
PRINCIPLES OF FLIGHT
When an aeroplane is moving there are four main forces acting upon it:WEIGHT, LIFT, THRUST and DRAG.
These are all closely interrelated. i.e.:The greater the weight - the greater the lift requirement.
The greater the lift - the greater the drag.
The greater the drag - the greater the thrust required, and so on ...
Air has properties which change with altitude. Knowledge of these variables, together with their
effect on an aeroplane, is a prerequisite for a full understanding of the principles of flight.
The structural and aerodynamic design of an aeroplane is a masterpiece of compromise. An
improvement in one area frequently leads to a loss of efficiency in another.
An aeroplane does not 'grip' the air as a car does the road. An aeroplane is often not pointing
in the same direction in which it is moving.
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OVERVIEW AND DEFINITIONS
PRINCIPLES OF FLIGHT
1.2
GENERAL DEFINITIONS
Mass
Unit - Kilogram (kg) - 'The quantity of matter in a body.' The mass of a body is a
measure of how difficult it is to start or stop. ("a body", in this context, means a
substance. Any substance; a gas, a liquid or a solid).
e.g.
(a)
The larger the mass, the greater the FORCE required to start or stop it
in the same distance.
(b)
Mass has a big influence on the time and/or distance required to change
the direction of a body.
Force
Unit - Newton (N) - 'A push or a pull'. That which causes or tends to cause a change in
motion of a body.
There are four forces acting on an aircraft in flight - pushing or pulling in different
directions.
Weight
Unit - Newton (N) - 'The force due to gravity'. (F = m x g)
Where (m) is the mass of the object and (g) is the acceleration due to the gravity
constant, which has the value of9·81 mls 2 • (A 1 kg mass "weighs" 9·81 newtons)
e.g.
If the mass of a B737 is 60,000 kg
and F = m x g
it is necessary to generate: [60,000 kg x 9.81 mls2]
588,600 N of lift force.
Centre of Gravity (CG)
The point through which the weight of an aircraft acts.
(a)
An aircraft in flight is said to rotate around its CG.
(b)
The CG of an aircraft must remain within certain forward and aft limits, for
reasons of both stability and control.
1-5
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
OVERVIEW AND DEFINITIONS
Work
Unit - Joule (J) - A force is said to do work on a body when it moves the body in the
direction in which the force is acting. The amount of work done on a body is the product
of the force applied to the body and the distance moved by that force in the direction in
which it is acting. If a force is exerted and no movement takes place, no work has been
done.
e.g.
(a)
Work = Force x Distance (through which the force is applied)
(b)
If a force of ION ewton's moves a body 2 metres along its line of action
it does 20 Newton metres (Nm) of work. [10 N x 2 m = 20 Nm]
(c)
A Newton metre, the unit of work, is called ajoule (J).
Power
Unit - Watt (W) - Power is simply the rate of doing work. (the time taken to do work)
e.g.
(a)
Power (W) =
Force (N) x Distance (m)
-----'--"'-----~....:....
Time (s)
(b)
If a force of ION moves a mass 2 metres in 5 seconds, then the power
is 4 Joules per second. A Joule per second (J/s) is called a Watt (W),
the unit of power. So the power used in this example is 4 Warts.
Energy
Unit - Joule (J) - Mass has energy ifit has the ability to do work. The amount of energy
a body possesses is measured by the amount of work it can do. The unit of energy will
therefore be the same as those of work, joules.
Kinetic Energy
Unit - Joule (J) - 'The energy possesseq by mass because of its motion'. 'A mass that
is moving can do work in coming to rest'.
KE = ~ m V 2 joules
The kinetic energy ofa 1 kg mass of air moving at 52 mls (100 knots) is 1352 joules; it
possesses 1352 joules of kinetic energy. [0.5 x 1 x 52 x 52 = 1352 J ]
From the above example it can be seen that doubling the velocity will have a greater
impact on the kinetic energy than doubling the mass. (velocity is squared).
1-6
© Oxford Aviation Services Limited
OVERVIEW AND DEFINITIONS
PRINCIPLES OF FLIGHT
Newton's first law of motion
'A body will remain at rest or in uniform motion in a straight line unless acted on by an
external force' .
To move a stationary object or to make a moving object change its direction a force must
be applied.
Inertia
'The opposition which a body offers to a change in motion'. A property of all bodies.
Inertia is a quality, but measured in terms of mass, which is a quantity.
e.g.
(a)
The larger the mass, the greater the force required for the same result.
(b)
A large mass has a lot of inertia.
(c)
Inertia refers to both stationary and moving masses.
Newton's second law of motion
'The acceleration of a body from a state of rest, or uniform motion in a straight line, is
proportional to the applied force and inversely proportional to the mass' .
Velocity
Unit - Metres per second (m/s). - 'Rate of change of displacement'
Acceleration
Unit - Metres per second per second (m/s2) - 'Rate of change of velocity' .
A force of 1 newton acting on a mass of 1 kg will produce an acceleration of 1 m/s2
Acceleration
Force
=- Mass
e.g.
(a)
For the same mass; the bigger the force, the greater the acceleration ..
(b)
For the same force; the larger the mass, the slower the acceleration.
1-7
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
OVERVIEW AND DEFINITIONS
Momentum
Unit - Mass x Velocity (kg-mls) - 'The quantity of motion possessed by a body'. The
tendency of a body to continue in motion after being placed in motion.
e.g.
(a)
A body of 10 kg mass moving at 2 mls has 20 kg-mls of momentum.
(b)
At the same velocity, a large mass has more momentum than a small
mass.
Newton's third law
'Every action has an equal and opposite reaction'
e.g.
(a)
If a force accelerates a mass in one direction, the body supplying the
force will be subject to the same force in the opposite direction.
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© Oxford Aviation Services Limited
OVERVIEW AND DEFINITIONS
PRINCIPLES OF FLIGHT
1.3
GLOSSARY
Aerofoil - A body so shaped as to produce aerodynamic reaction normal to the direction of its
motion through the air without excessive drag.
Aft - To the rear, back or tail of the aircraft.
Air brake - Any device primarily used to increase drag of an aircraft at will.
Ambient - Surrounding or pertaining to the immediate environment.
Amplitude - Largeness; abundance; width; range; extent of repetitive movement (from
extreme to extreme).
Attitude - The nose-up or nose-down orientation of an aircraft relative to the horizon.
Boundary Layer - The thin layer of air adjacent to a surface, in which the viscous forces are
dominant.
Buffeting - An irregular oscillation of any part of an aircraft produced and maintained directly
by an eddying flow.
Cantilever (wing) - A wing whose only attachment to the fuselage is by fittings at the wing root,
has no external struts or bracing. The attachments are faired-in to preserve the streamline shape
Control lock (Gust lock) - A mechanical device designed to safeguard, by positive lock, the
control surfaces and flying control system against damage in high winds or gusts when the
aircraft is parked.
Control Reversal - At high speed: the displacement of a control surface producing a moment
on the aircraft in a reverse sense because of excessive structural distortion. At low speed: the
displacement of an aileron increasing the angle of attack of one wing to or beyond the critical
angle, causing a roll in the direction opposite to that required.
Convergent - Tend towards or meet in one point or value.
Critical Mach number (M CRIT) - The free stream Mach number at which the peak velocity on
the surface of a body first becomes equal to the local speed of sound.
Damping - To slow down the rate; to diminish the amplitude of vibrations or cycles.
Geometric Dihedral - The angle between the horizontal datum of an aeroplane and the plane of
a wing or horizontal stabiliser semi -span.
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PRINCIPLES OF FLIGHT
OVERVIEW AND DEFINITIONS
Divergent - To incline or tum apart. Divergence - A disturbance which increases continually
with time.
Eddy - An element of air having intense vorticity.
Effective Angle of Attack (ue) - The angle between the chord line and the mean direction of a
non-uniform disturbed airstream.
Equilibrium - A condition that exists when the sum of all moments acting on a body is zero
AND the sum of all forces acting on a body is zero.
Fairing - A secondary structure added to any part to reduce its drag.
Feel - The sensations of force and displacement experienced by the pilot from the aerodynamic
forces on the control surfaces.
Fence - A projection from the surface of the wing and extending chordwise to modify the wing
surface pressure distribution.
Fillet - A fairing at the junction of two surfaces to improve the airflow.
Flightpath - The path of the Centre ofOravity (CO) of an aircraft.
Fluid - A substance, either gaseous or liquid, that will conform to the shape of the container that
holds it.
Free stream velocity - The velocity of the undisturbed air relative to the aircraft.
Gradient (Pressure) - Rate of change in pressure with distance.
Gust - A rapid variation, with time or distance, in the speed or direction of air.
Gust lock - See control lock.
Instability - The quality whereby any disturbance from steady motion tends to increase.
Laminar Flow - Flow in which there is no mixing between adjacent layers.
Load Factor - The ratio of the weight of an aircraft to the load imposed by lift. The correct
symbol for load factor is (n), but is colloquially known as (g).
Load Factor
1 - 10
=
Lift
Weight
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
OVERVIEW AND DEFINITIONS
Mach Number (M) - The ratio of the True Air Speed to the speed of sound under prevailing
atmospheric conditions.
M
=
TAS
Local Speed of Sound (a)
Magnitude - Largeness; size; importance.
Moment (N-m) - The moment of a force about a point is the product of the force and the distance
through which it acts. The distance in the moment is merely a leverage and no movement is
involved, so moments cannot be measured in joules.
Nacelle - A streamlined structure on a wing for housing engines (usually).
Normal- Perpendicular; at 90°.
Oscillation - Swinging to and fro like a pendulum; a vibration; variation between certain limits;
fluctuation.
Parallel - Lines which run in the same direction and which will never meet or cross.
Pitot tube - A tube, with an open end facing up-stream, wherein at speeds less than about four
tenths the speed of sound the pressure is equal to the total pressure. For practical purposes, total
pressure may be regarded as equal to pitot pressure at this stage.
Pod - A nacelle supported externally from a fuselage or wing.
Propagate - To pass on; to transmit; to spread from one to another.
Relative Airflow, (Relative Wind), (Free Stream Flow) - The direction of airflow produced by
the aircraft moving through the air. The relative airflow flows in a direction parallel and opposite
to the direction offlight. Therefore, the actual flight path of the aircraft determines the direction
of the relative airflow. Also, air in a region where pressure, temperature and relative velocity are
unaffected by the passage of the aircraft through it.
Scale - If a 1I10th scale model is considered, all the linear dimensions are 1I10th of the real
aircraft, but the areas are III OOth; and if the model is constructed of the same materials, the mass
is 1I1000th of the real aircraft. So the model is.to scale in some respects, but not others.
Schematic - A diagrammatic outline or synopsis; an image of the thing; representing something
by a diagram.
Separation - Detachment of the airflow from a surface with which it has been in contact.
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OVERVIEW AND DEFINITIONS
PRINCIPLES OF FLIGHT
Shockwave - A narrow region, crossing the streamlines, through which there occur abrupt
increases in pressure, density, and temperature, and an abrupt decrease in velocity. The normal
component of velocity relative to the shock wave is supersonic upstream and subsonic
downstream.
Side-slip - Motion of an aircraft, relative to the relative airflow, which has a component of
velocity along the lateral axis.
Slat - An auxiliary, cambered aerofoil positioned forward of the main aerofoil so as to form a
slot.
Spar - A principal spanwise structural member of a wing, tailplane, fin or control surface.
Speed - Metres per second (m1s) is used in most formulae, but nautical miles per hour or knots
(kt) are commonly used to measure the speed of an aircraft. There are 6080 ft in 1 nautical mile
and 3.28 ft in 1 metre.
Speed of Sound (a) - Sound is pressure waves which propagate spherically through the
atmosphere from their source. The speed of propagation varies ONLY with the temperature of
the air. The lower the temperature, the lower the speed of propagation. On a 'standard' day at
sea level the speed of sound is approximately 340 m1s (660 kt TAS).
Stability - The quality whereby any disturbance of steady motion tends to decrease.
Stagnation point - A point where streamlines are divided by a body and where the fluid speed
is zero, relative to the surface.
Static vent - A small aperture in a plate fixed to form part of the fuselage and located
appropriately for measuring the ambient static pressure.
Throat - A section of minimum area in a duct.
True Air Speed (T AS) or (V) - The speed at which the aircraft is travelling through the air.
Turbulent Flow - Flow in which irregular fluctuations with time are superimposed on a mean
flow.
Velocity - The same as speed, but with direction specified as well.
Viscosity - The resistance of fluid particles to flow over each other. All fluids have the property
of viscosity. A fluid with high viscosity would not flow very easily. The viscosity of air is low
in comparison to something like syrup, but the viscosity that air does have is a very important
consideration when studying aerodynamics.
Vortex - A region of fluid in circulatory motion, having a core of intense vorticity, the strength
of the vortex being given by its circulation.
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OVERVIEW AND DEFINITIONS
PRINCIPLES OF FLIGHT
Vortex generator - A device, often a small vane attached to a surface, to produce one or more
discrete vortices which trail downstream adjacent to the surface, promote mixing in the boundary
layer and delay boundary layer separation. (Increases the kinetic energy of the boundary layer).
Vorticity - Generally, rotational motion in a fluid, defined, at any point in the fluid, as twice the
mean angular velocity of a small element of fluid surrounding the point.
Wake - The region of air behind an aircraft in which the total pressure has been changed by the
presence of the aircraft.
Wash-out - Decrease in angle of incidence towards the tip of a wing or other aerofoil.
Wing Loading - Ratio of aircraft weight to wing area.
Wing Loading =
Aircraft Weight
Wing Area
Zoom - Using kinetic energy to gain height.
1 - 13
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OVERVIEW AND DEFINITIONS
PRINCIPLES OF FLIGHT
1.4
LIST OF SYMBOLS
The following symbols are used throughout these notes. However, no universal defining standard
for their use exists. Other books on the subject may use some of these symbols with different
definitions. Every effort has been made to employ symbols that are widely accepted and that
conform to the JAA Learning Objectives.
a
AC
AR
b
C
c
CD
CG
CP
CL
Cm
D
Di
F
g
K
L
LID
M
m
n
p
Qorq
S
T
speed of sound
aerodynamic centre
aspect ratio
span
Centigrade
chord length
drag coefficient
centre of gravity
centre of pressure
lift coefficient
pitching moment coefficient
drag
induced drag
force
acceleration due to gravity
also used for load factor
Kelvin
lift
lift to drag ratio
Mach number
tIc
V
Vs
W
mass
load factor
pressure
dynamic pressure
area; wing area
temperature
thickness-chord ratio
free stream speed (T AS)
stall speed
weight
GREEK SYMBOLS
a.
~
y
11
Jl
p
0'
<P
(alpha) angle of attack
(beta) sideslip angle
(gamma) angle of climb or
descent
(delta) increment in
(mu) Mach angle
(rho) density
(sigma) relative density
(phi) angle of bank
OTHERS
.
ex
..!.
NB.
proportional to
is approximately equal to
The Greek symbol y (gamma) has been used in these notes to denote angle of climb and descent.
The JAA Learning Objectives use e (theta). Evidence exists that a question in the exam uses
y (gamma) for angle of climb and descent. The notes have been amended to use y, but consider
either y or e to indicate angle of climb and descent. (15th Sept 2000).
1 - 14
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
OVERVIEW AND DEFINITIONS
SELF ASSESSMENT QUESTIONS
Aircraft (2)
Mass: 2,000 kilograms (kg)
Engine thrust: 8,000 Newtons (N)
VI speed: 130 knots (kt)
Take-off run to reach VI: 1,500 metres (m)
Time taken to reach VI: 40 seconds (s)
Aircraft (1)
Mass: 2,000 kilograms (kg)
Engine thrust: 4,000 Newtons (N)
V I speed: 65 knots (kt)
Take-off run to reach VI: 750 metres (m)
Time taken to reach VI: 30 seconds (s)
where, 1 nautical mile
=
6080 ft and 1 metre = 3.28 ft
At VI both aircraft experience an engine failure and take-off is abandoned.
a)
How much work was done to aircraft (1) getting to V I
b)
How much power was used to get aircraft (1) to V I
c)
How much work was done to aircraft (2) getting to VI
d)
How much power was used to get aircraft (2) to VI
e)
How much momentum does aircraft (1) possess at VI
t)
How much momentum does aircraft (2) possess at VI
g)
How many times greater is the momentum of aircraft (2)
h)
How much kinetic energy does aircraft (1) possess at V I
i)
How much kinetic energy does aircraft (2) possess at VI
j)
How many times greater is the kinetic
k)
State the mass and velocity relationship of both aircraft and compare to their momentum
and kinetic energy
1)
Which has the greater effect on kinetic energy, mass or velocity
m)
What must be done with the kinetic energy so the aircraft can be brought to a stop
1 - 15
~nergy
of aircraft (2)
© Oxford Aviation Services Limited
OVERVIEW AND DEFINITIONS
PRINCIPLES OF FLIGHT
1.
An aircraft's mass is a result of:
a)
b)
c)
d)
2.
The unit of mass is the:
a)
b)
c)
d)
3.
Mass-kilogram
Newton-metre
louIe
Newton
The unit of weight is the:
a)
b)
c)
d)
6.
That which causes a reaction to take place
Thrust and drag only
A push or a pull
The result of an applied input
The unit of force is the:
a)
b)
c)
d)
5.
louIe
Watt
Newton
Kilogram
The definition of a force is:
a)
b)
c)
d)
4.
Its weight
How big it is
How much matter it contains
Its volume
Kilogram
Newton
Watt
Kilowatt
Weight is the result of:
a)
b)
c)
d)
The force on mass due to gravity
The action of a falling mass
How much matter the object contains
The rate of mass per unit area
1 - 16
© Oxford Aviation Services Limited
OVERVIEW AND DEFINITIONS
PRINCIPLES OF FLIGHT
7.
About which point does an aircraft rotate:
a)
b)
c)
d)
8.
If a force is applied to a mass and the mass does not move:
a)
b)
c)
d)
9.
Pascal
Joule
Watt
Kilogram
The unit of power is called the:
a)
b)
c)
d)
11.
Work is done even though there is no movement of the mass
Work is done only if the mass moves a long way
Power is exerted, but no work is done
No work is done
The unit of work is called the:
a)
b)
c)
d)
10.
The wings
The main undercarriage
The centre of gravity
The rudder
Joule
Newton-metre
Watt
Metre per second
Ifa force of20 Newton's moves a mass 5 metres:
1 - the work done
2 - the work done
3 - the work done
4 - the work done
is 100 Nm
is 100 Joules
is 4 Joules
is 0.25 Joules
The correct statements are:
a)
b)
c)
d)
1 only
1 and 3
1 and 2
2 only
1 - 17
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
12.
Ifa force of 50 Newton's is applied to a 10 kg mass and the mass moves 10 metres and a force
of 50 Newton's is applied to a 100 kg mass which moves 10 metres:
a)
b)
c)
d)
13.
4167 Watts
250 Kilowatts
1 Megawatt
4 Watts
Kinetic energy is:
a)
b)
c)
d)
16.
The rate of force applied
The rate of movement per second
The rate of doing work
The rate of applied force
If a force of 500 Newton's moves a mass 1000 metres in 2 mins, the power used is:
a)
b)
c)
d)
15.
The work done is the same in both cases
Less work is done to the 10 kg mass
More work is done to the 10 kg mass
More work is done to the 100 kg mass
The definition of power is:
a)
b)
c)
d)
14.
OVERVIEW AND DEFINITIONS
The energy a mass possesses due to its position in space
The energy a mass possesses when a force has been applied
The energy a mass possesses due to the force of gravity
The energy a mass possesses because of its motion
The unit of kinetic energy is the:
a)
b)
c)
d)
Joule
Metre per second
Watt
Newton-metre per second
1 - 18
© Oxford Aviation Services Limited
OVERVIEW AND DEFINITIONS
PRINCIPLES OF FLIGHT
17.
When considering kinetic energy:
1 - a moving mass can apply a force by being brought to rest
2 - kinetic energy is the energy possessed by a body because of its motion
3 - if a body's kinetic energy is increased, a force must have been applied
4 - kinetic energy = Y2 m V2 joules
The combination of correct statements is:
a)
b)
c)
d)
18.
The property of inertia is said to be:
a)
b)
c)
d)
19.
1 and 2
1, 2, 3 and 4
4 only
2 and 4
The energy possessed by a body because of its motion
The opposition which a body offers to a change in motion
That every action has an equal and opposite reaction
The quantity of motion possessed by a body
Considering Newton's first law of motion:
1 - a body is said to have energy if it has the ability to do work
2 - the amount of energy a body possesses is measured by the amount of work it can do
3 - a body will tend to remain at rest, or in uniform motion in a straight line, unless acted upon
by an external force
4 - to move a stationary object or to make a moving object change its direction, a force must be
applied
The combination with the correct statements is:
a)
b)
c)
d)
3 and 4
3 only
1 and 2
1, 2, 3 and 4
1 - 19
© Oxford Aviation Services Limited
OVERVIEW AND DEFINITIONS
PRINCIPLES OF FLIGHT
20.
Considering Newton's second law of motion:
1 - every action has an equal and opposite reaction
2 - if the same force is applied, the larger the mass the slower the acceleration
3 - if two forces are applied to the same mass, the bigger the force the greater the acceleration
4 - the acceleration of a body from a state of rest, or uniform motion in a straight line, is
proportional to the applied force and inversely proportional to the mass
The combination of true statements is:
a)
b)
c)
d)
21.
Newton's third law of motion states:
a)
b)
c)
d)
22.
The energy possessed by a mass is inversely proportional to its velocity
Every force has an equal and opposite inertia
For every force there is an action
Every action has an equal and opposite reaction
The definition of velocity is the:
a)
b)
c)
d)
23.
1 only
1, 2, 3 and 4
2,3,and4
3 and 4
Rate of change of acceleration
Rate of change of displacement
The quantity of motion possessed by a body
The acceleration of a body in direct proportion to its mass
When considering acceleration:
1 - acceleration is the rate of change of velocity
2 - the units of acceleration are metres per second
3 - the units of acceleration are kilogram-metres per second
4 - the units of acceleration are seconds per metre per metre
The combination of correct statements is:
a)
b)
c)
d)
4 only
1 and 4
1 only
1 and 2
1 - 20
© Oxford Aviation Services Limited
OVERVIEW AND DEFINITIONS
PRINCIPLES OF FLIGHT
24.
The definition of momentum is:
a)
b)
c)
d)
25.
The quantity of mass possessed by a body
The quantity of inertia possessed by a body
The quantity of motion possessed by a body
The opposition which a body offers to a change in velocity
A force of24 Newton's moves a 10 kg mass 60 metres in 1 minute, the power used is:
1 - 24 Watts
2 - 240 Watts
3 - Force times distance moved in one second
4 - Force times the distance the mass is moved in one second
Which of the preceding statements are correct
a)
b)
c)
d)
26.
1 and 3
1,3 and 4
2 and 4
4 only
When considering momentum:
1 - Momentum is the quantity of motion possessed by a body
2 - Momentum is the tendency of a body to continue in motion after being placed in motion
3 - A mass of2000 kg moving at 55 mls has 110,000 kg-mls of momentum
4 - A large mass moving at 50 mls will have less momentum than a small mass moving at 50 mls
The correct combination of statements is:
a)
b)
c)
d)
1 and 3
1, 2, 3 and 4
1,2 and 3
2,3 and 4
1 - 21
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
OVERVIEW AND DEFINITIONS
ANSWERS
Aircraft number (1) VI speed of65 knots = 33.5 mls
Aircraft number (2) VI speed of 130 knots = 67 mls
a)
b)
c)
d)
e)
f)
g)
h)
i)
j)
k)
1)
m)
3,000,000 joules
100,000 watts
12,000,000 joules
300,000 watts
67,000 kg-mls
134,000 kg-mls
Twice
1,122,250 joules
4,489,000 joules
Four times greater
Same mass, speed doubled, momentum doubled, but kinetic energy four times greater.
Velocity has a greater effect on kinetic energy than mass.
It must be dissipated by the braking systems.
I No I A I B I c I DII
1
III
D
3
16
D
B
5
D
8
B
18
B
22
A
C
D
B
10
C
23
C
11
C
24
C
12
13
A
25
26
C
1 - 23
I
A
21
B
9
REF
D
20
C
I
A
17
19
A
7
IA IB Ic ID
15
C
4
No
14
C
2
6
REF
B
C
© Oxford Aviation Services Limited
CHAPTER 2 - THE ATMOSPHERE
Contents
Page
INTRODUCTION ....................................................... 2 - 1
THE PHYSICAL PROPERTIES OF AIR
STATIC PRESSURE
TEMPERATURE ................................................. 2 - 2
AIR DENSITY
INTERNATIONAL STANDARD ATMOSPHERE ....................... 2 - 3
DYNAMIC PRESSURE ................................................... 2 - 4
MEASURING DYNAMIC PRESSURE ................................ 2 - 6
RELATIONSHIPS BETWEEN AIR SPEEDS .................................. 2-7
INDICATED AIR SPEED
CALIBRATED AIR SPEED
EQUIVALENT AIR SPEED
TRUE AIR SPEED
SPEED OF SOUND ............................................... 2 - 8
MACH NUMBER
CRITICAL MACH NUMBER
AIRSPEED
ERRORS AND CORRECTIONS ..................................... 2 - 9
INSTRUMENT ERROR
POSITION ERROR (PRESSURE ERROR)
COMPRESSIBILITY ERROR
'V'SPEEDS
SUMMARY ........................................................... 2 - 10
SELF ASSESSMENT QUESTIONS ........................................ 2 - 11
ANSWERS ..................................................... 2 -17
PRINCIPLES OF FLIGHT
2.1
THE ATMOSPHERE
INTRODUCTION
The atmosphere is the medium in which an aircraft operates. It is the properties of the
atmosphere, changed by the shape of the wing, that generate the required Lift force.
The most important property is air density (the "thickness" of air)
KEY FACT: If air density decreases, the mass of air flowing over the aircraft in a
given time will decrease. Not usually considered during the study of Principles of
Flight, keeping the idea of Mass flow (Kg/s) in the 'back of your mind' can aid general
understanding.
A given mass flow will generate the required Lift force, but a decrease in air density will
reduce the mass flow.
To maintain the required Lift force if density is decreased, the speed of the aircraft
through the air must be increased. The increased speed of airflow over the wing will
restore the mass flow and Lift force to its previous value.
2.2
THE PHYSICAL PROPERTIES OF AIR
Air has substance! Air has mass; not very much if compared to other matter, but nevertheless
a significant amount. A mass of moving air has considerable kinetic energy, e.g. when moving
at 100 knots the kinetic energy of air can inflict severe damage to man-made structures.
Air is a compressible fluid and is able to flow or change its shape when subjected to even minute
pressure differences. (Air will flow in the direction of the lower pressure). The viscosity of air
is so low that very small forces are able to move the molecules in relation to each other.
When considering the portion of atmosphere in which most aircraft operate (up to 40,000ft);
with increasing altitude the characteristics of air undergo a gradual transition from those at sea
level. Since air is compressible, the lower layers contain much the greater part of the whole
mass of the atmosphere. Pressure falls steadily with increasing altitude, but temperature falls
steadily only to about 36,000 ft, where it then remains constant through the stratosphere.
2.2.1
STATIC PRESSURE
The unit for static pressure is N/m 2, the symbol is lower case 'p'.
a)
Static pressure is the result of the weight of the atmosphere pressing down on the air
beneath.
b)
Static pressure will exert the same force per square metre on all surfaces of an
aeroplane. The lower the altitude the greater the force per square metre.
2-1
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
THE ATMOSPHERE
c)
It is called static pressure because of the air's stationary or static presence.
d)
An aircraft always has Static pressure acting upon it.
Newtons per square metre is the SI unit for pressure. 1 N/m2 is called a Pascal and is quite a
small unit. In aviation the hectoPascal (hPa) is used. (,hecto' means 100 and 1 hectoPascal is
the same as 1 millibar).
Static pressure at a particular altitude will vary from day to day, and is about 1000 hPa at sea
level. In those countries that measure static pressure in inches of mercury (ins Hg), sea level
static pressure is about 30 ins Hg.
2.2.2
TEMPERATURE
The unit for temperature is °C, or K. Degrees Celsius (or centigrade) when measured relative
to the freezing point of water, or Kelvin when measured relative to absolute zero. (O°C is
equivalent to 273 K).
Temperature decreases with increasing altitude, up to about 36,000ft and then remains constant.
2.2.3
AIR DENSITY
The unit for density is kg/m3 and the symbol is the Greek letter p [rho].
a)
Density is 'Mass per unit volume' (The "number" of air particles in a given space).
b)
Density varies with static pressure, temperature and humidity.
(i)
Density decreases if static pressure decreases.
(ii)
Density decreases if temperature increases.
(iii)
Density decreases if humidity increases. (This will be discussed later).
Air Density is proportional to pressure and inversely proportional to temperature. This is shown
in the ideal gas law formula below.
P
T
p
=
where, p
constant,
=
more usefully it can be said that
pressure;
T
=
temperature and p
p
=
oc
P
T
density
Density decreases with increasing altitude because of decreasing static pressure. However, with
increasing altitude temperature also decreases, which would tend to increase density, but the
effect of decreasing static pressure is dominant.
2-2
© Oxford Aviation Services Limited
THE ATMOSPHERE
PRINCIPLES OF FLIGHT
2.2.4
INTERNATIONAL STANDARD ATMOSPHERE (ISA)
The values of temperature, pressure and density are never constant in any given layer of the
atmosphere. To enable accurate comparison of aircraft performance and the calibration of
pressure instruments, a 'standard' atmosphere has been adopted. The standard atmosphere
represents the mean or average properties of the atmosphere.
Europe uses the standard atmosphere defined by the International Civil Aviation Organisation
(lCAO).
The ICAO standard atmosphere assumes the following mean sea level values:
15 °C
1013 ·25 hPa
1·225 kg/m3
Temperature
Pressure
Density
The temperature lapse rate is assumed to be uniform at a rate of 2 °C per 1,000 ft (1 '98 0c) from
mean sea level up to a height of36,090 ft (11,000 m) above which the lapse rate becomes zero
and the temperature remains constant at - 56·5 °C.
ICAO Standard Atmosphere
CC)
Pressure (hPa)
(p)
Density (kg/m3)
(p)
Relative Density
(0)
0
15
1013·25
1·225
1·0
5,000
5·1
843·1
1·056
0·86
10,000
- 4·8
696·8
0·905
0·74
15,000
- 14·7
571·8
0·771
0·63
20,000
- 24·6
465·6
0·653
0·53
25,000
- 34·5
376·0
0·549
0-45
30,000
- 44-4
300·9
0-458
0·37
35,000
- 54·3
238-4
0·386
0·31
40,000
-56·5
187·6
0·302
0·25
45 ,000
- 56,5
147·5
0·237
0·19
50,000
- 56,5
116·0
0·186
0·15
Altitude
(ft)
Temperature
NOTE: The air density at 40,000 ft is only 114 of the sea-level value.
2-3
© Oxford Aviation Services Limited
THE ATMOSPHERE
PRINCIPLES OF FLIGHT
2.3
DYNAMIC PRESSURE
The unit for dynamic pressure is N/m2 and the symbol is lower case 'q' or upper case 'Q'.
a)
Because air has mass, air in motion must possess kinetic energy, and will exert a force
per square metre on any object in its path. (KE = 112 m y2)
b)
It is called DYNAMIC pressure because the air is moving in relation to the object being
considered, in this case an aircraft.
c)
Dynamic pressure is proportional to the density of the air and the square of the speed of
the air flowing over the aircraft.
An aircraft immersed in moving airflow will therefore experience both Static AND Dynamic
pressure. (Remember, static pressure is always present).
The kinetic energy of one cubic metre of air moving at a stated speed is given by the formula:
Kinetic Energy = 'is P y2
joules
where p is the local air density in kglm3 and Y is the speed in m1s
If this cubic metre of moving air is completely trapped and brought to rest by means of an openended tube the total energy will remain constant, but by being brought completely to rest the
kinetic energy will become pressure energy which, for all practical purposes, is equal to:
Dynamic Pressure
=
'is p y2
N/m2
Consider air flowing at 52 m1s (100 kt) with a density of 1·225 kg/m3
(100 kt x 6080ft = 608000ft/hour -:- 3·28 = 185366metreslhour -:- 60 -:- 60
Dynamic pressure
=
0·5 x 1·225 x 52 x 52
1656
52m1s)
N/m2 (16·56 hPa)
If speed is doubled, dynamic pressure will be four times greater
0·5 x 1·225 x 104 x 104 = 6625 N/m2
2-4
(66·25 hPa)
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
THE ATMOSPHERE
If the area of the tube is 1m2 a force of lh p y2
(F orce = Pressure x Area)
Newtons will be generated.
Dynamic pressure ( lh p y2) is common to ALL aerodynamic forces and determines the air
loads imposed on an aeroplane moving through the air.
The symbol for dynamic pressure ( lh p y2) is q or Q
(Q
= 1/2p V
2
J
KEY FACTS
1.
A pilot needs to know how much dynamic pressure is available, but dynamic pressure
cannot be measured on its own because static pressure will always be present. The sum
of Static and Dynamic pressure, in this context, is known as "Total" Pressure.
(Dynamic + Static pressure can also be referred to as Stagnation or Pitot Pressure).
Total Pressure = Static Pressure + Dynamic Pressure
This can be re-arranged to show that:
Total Pressure - Static Pressure
2.
=
Dynamic Pressure
The significance of dynamic pressure to the understanding of Principles of Flight cannot
be over-emphasised.
Because dynamic pressure is dependent upon air density and the speed of the aircraft
through the air, it is necessary for students to fully appreciate the factors which affect
air density.
a)
Temperature - increasing temperature decreases air density. Changes in air
density due to air temperature are significant during all phases of flight.
b)
Static pressure - decreasing static pressure decreases air density. Changes in
air density due to static pressure are significant during all phases of flight.
c)
Humidity - increasing humidity decreases air density. (The reason increasing
humidity decreases air density is that the density of water vapour is about 5/8
that of dry air). Humidity is most significant during take-off and landing.
Increasing altitude will decrease air density because the effect of decreasing static
pressure is more dominant than decreasing temperature.
2-5
© Oxford Aviation Services Limited
THE ATMOSPHERE
PRINCIPLES OF FLIGHT
2.3.1
MEASURING DYNAMIC PRESSURE
All aerodynamic forces acting on an aircraft are determined by dynamic pressure, so it is
essential to have some means of measuring dynamic pressure and presenting that information
to the pilot.
A sealed tube, open at the forward end, is located where it will collect air when the aircraft is
moving. The pressure in the tube (Pitot tube) is Dynamic + Static and, in this context, is called
"Pi tot" Pressure. (Because the air is inside the Pitot tube).
Some way of 'removing' the static pressure from the pitotpressure must be found. A hole (vent)
in a surface parallel to the airflow will sense static pressure. Referring to Fig 2.1, if the pressure
from the pitot tube is fed to one side of a diaphragm mounted in a sealed case, and static pressure
is fed to the other side, the two static pressures will cancel each other and the diaphragm
movement will be influenced only by changes in dynamic pressure.
Movement of the diaphragm moves a pointer over a scale so that changes in dynamic pressure
can be observed by the flight crew. But the instrument is calibrated at ISA sea level density,
so the instrument will only give a 'true' indication of the speed of the aircraft through the air
when the air density is 1·225 kg/m3.
This is not a problem because the pilot needs an indication of dynamic pressure, and this is what
the instrument provides. The instrument is made in such a way that it indicates the square root
of the dynamic pressure in nautical miles per hour (knots) or statute miles per hour (MPH). So,
if this "Indicated Air Speed" is doubled, the speed of the aircraft through the air will also be
doubled.
The Air Speed Indicator
is a pressure gauge
••• Airflow
PITOT TUBE -
1
PITOT PRESSURE
(Static + Dynamic)
/
Needle indicates
changes in
DYNAMIC PRESSURE
/-~-
••• Airflow
STATIC
~::~
' - I_
__
Figure 2.1
__
P_R_S~_~_~_G_CR_E_ _ '
Schematic of Air Speed Indicator (ASI)
2-6
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
2.4
THE ATMOSPHERE
RELATIONSHIPS BETWEEN AIR SPEEDS
Indicated Air Speed: (lAS). The speed registered on the Air Speed Indicator.
Calibrated Air Speed: (CAS). An accurate measure of dynamic pressure when the aircraft is
flying slowly. The position of the pitot tube(s) and static vent(s), together with the aircraft's
configuration (Flaps, landing gear etc.) and attitude to the airflow (Angle of attack and sideslip)
will affect the pressures sensed; particularly the pressures sensed at the static vent(s).
Under the influence of the above conditions a false dynamic pressure (lAS) will be displayed.
When lAS is corrected for this 'position' or 'pressure' error, as it's called, the resultant is
Calibrated Air Speed. (The airspeed corrections to be applied may be displayed on a placard on
the flight-deck, or in the Flight Manual, and will include any instrument error).
Equivalent Air Speed: (EAS). An accurate measure of dynamic pressure when the aircraft is
flying fast. Air entering the pitot tube( s) is compressed, which gives a false dynamic pressure
(lAS) reading, but only becomes significant at higher speeds.
At a given air density, the amount of compression depends on the speed of the aircraft through
the air. When the lAS is corrected for 'position' AND 'compressibility' error, the resultant is
Equivalent Air Speed.
True Air Speed: (TAS) or (V). The speed of the aircraft through the air. THE ONLY SPEED
THERE IS - All the other, so called, speeds are pressures.
TAS
=
EAS
Jcr
Where, (J is Relative Density
The Air Speed Indicator is calibrated for 'standard' sea level density, so it will only read TAS
if the density of the air through which the aircraft is flying is 1·225 kg/m3. Thus at 40,000 ft
where the 'standard' density is one quarter of the sea-level value, to maintain the same EAS
the aircraft will have to move through the air twice as fast!
2-7
© Oxford Aviation Services Limited
THE ATMOSPHERE
PRINCIPLES OF FLIGHT
The Speed of Sound: (a). Sound is 'weak' pressure waves which propagate spherically through
the atmosphere from their source. The speed at which pressure waves propagate is
proportional to the square root of the absolute temperature of the air. The lower the
temperature, the lower the speed of propagation. On a 'standard' day at sea level the speed of
sound is approximately 340 m/s (660 kt TAS).
At higher aircraft True Air Speeds (TAS) and/or higher altitudes, it is essential to know the
speed of the aircraft in relation to the local speed of sound. This speed relationship is known as
the Mach Number (M).
M
=
TAS
a
where (a) is the local speed of sound
i.e. If the True Air Speed of the aircraft is four tenths the speed at which pressure waves
propagate through the air mass surrounding the aircraft, the Mach meter will register 0·4 M
Critical Mach Number: (M CRIT) The critical Mach number is the Mach number of the aircraft
when the speed of the airflow over some part of the aircraft (usually the point of maximum
thickness on the aerofoil) first reaches the speed of sound.
2.5
AIRSPEED
This information is to reinforce that contained in the preceding paragraphs.
The airspeed indicator is really a pressure gauge, the 'needle' of which responds to changes in
dynamic pressure (1/2 p y2).
I
!he Air Speed Indicator
IS a pressure gauge
~
I
~
Calibration of the airspeed indicator is based on standard sea level density (1·225 kg/m3). The
"airspeed" recorded will be different from the actual speed of the aircraft through the air unless
operating under standard sea-level conditions (unlikely). The actual speed ofthe aircraft relative
to the free stream is called true airspeed (TAS), and denoted by (Y). The "speed" recorded by
the airspeed indicator calibrated as above, if there are no other errors, is called equivalent
airspeed (EAS).
It may seem to be a drawback that the instrument records equivalent rather than true airspeed.
But the true airspeed may always be determined from it. Also, many of the handling
characteristics of an aircraft depend mainly on the dynamic pressure, i.e. on the equivalent
airspeed, so it is often more useful to have a direct reading ofEAS than TAS.
2-8
© Oxford Aviation Services Limited
THE ATMOSPHERE
PRINCIPLES OF FLIGHT
2.5.1
ERRORS AND CORRECTIONS
An airspeed indicator is, however, also subject to errors other than that due to the difference
between the density of the air through which it is flying and standard sea level density.
a)
Instrument Error: This error may arise from the imperfections in the design and
manufacture of the instrument, and varies from one instrument to another. Nowadays
this type of error is usually very small and for all practical purposes can be disregarded.
Where any instrument error does exist, it is incorporated in the calibrated airspeed
correction chart for the particular aeroplane.
b)
Position Error (Pressure Error): This error is of two kinds, one relating to the static
pressure measurement, the other to the pitot (total) pressure measurement. The pitot
tube(s) and static port(s) may be mounted in a position on the aircraft where the flow is
affected by the presence of the aircraft, changes in configuration (flaps and maybe gear)
and proximity to the ground (ground effect). If so, the static pressure recorded will be
the local and not the free stream value. The pitot pressure may be under-recorded
because of incorrect alignment - the tube( s) may be inclined to the airstream instead of
facing directly into it (changes in angle of attack, particularly at low speeds). The
magnitude of the consequent errors will generally depend on the angle of attack, and
hence the speed of the aircraft.
c)
Compressibility Error: At high speeds, the dynamic pressure is not simply 1/2 p y2,
but exceeds it by a factor determined by Mach number. Thus the airspeed indicator will
over-read.
Because of the errors listed, the "speed" recorded on the airspeed indicator is generally not the
equivalent airspeed. It is called instead the indicated airspeed. Corrections to rectify the
instrument and position errors are determined experimentally. In flight, using special
instruments, measurements are taken over the whole range of speeds and configurations, from
which a calibration curve is obtained which gives the corrections appropriate to each indicated
airspeed. The compressibility error correction may be obtained by calculation.
The indicated airspeed, after correction for instrument, position (pressure) and compressibility
errors, gives the equivalent airspeed 1/2 p y2.
2.5.2
'V' SPEEDS
These include: Y s, YI' Y R , Y2, YMD , YMC , YYSE and many others - these are all Calibrated Air
Speeds because they relate to aircraft operations at low speed. However, the appropriate
corrections are made and these speeds are supplied to the pilot in the Flight Manual as lAS.
Y MO - The maximum operating lAS is however an EAS because it is a high speed, but again is
supplied to the pilot in the Flight Manual as an lAS.
2-9
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
2.6
THE ATMOSPHERE
SUMMARY
1.
Dynamic pressure (Q) is affected by changes in air density.
a)
air density decreases if atmospheric pressure decreases,
b)
air density decreases if air temperature increases,
c)
air density decreases if humidity increases.
(i)
With the aircraft on the ground:
Taking-off from an airfield with low atmospheric pressure and/or high
air temperature and/or high humidity, will require a higher TAS to
achieve the same dynamic pressure (lAS).
F or the purpose of general understanding:
A constant lAS will give constant dynamic pressure.
d)
increasing altitude decreases air density because of decreasing static pressure.
(i)
With the aircraft airborne:
As altitude increases, a higher TAS is required to maintain a constant
dynamic pressure. Maintaining a constant lAS will compensate for
changes in air density.
2.
There is only one speed, the speed of the aircraft through the air, the TAS. All the other,
so called, speeds are pressures.
The Air Speed Indicator is a pressure gauge.
3.
Aircraft 'V' speeds are CAS, except V MO which is an EAS, but all are presented to the
pilot in the Flight Manual as lAS.
2 - 10
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
THE ATMOSPHERE
SELF ASSESSMENT QUESTIONS
1.
When considering air:
1 - Air has mass
2 - Air is not compressible
3 - Air is able to flow or change its shape when subject to even small pressures
4 - The viscosity of air is very high
5 - Moving air has kinetic energy
The correct combination of all true statements is:
a)
b)
c)
d)
2.
Why do the lower layers contain the greater proportion of the whole mass of the atmosphere:
a)
b)
c)
d)
3.
1,2,3 and 5
2,3 and 4
1 and 4
1,3,and5
Because air is very viscous
Because air is compressible
Because of greater levels of humidity at low altitude
Because air has very little mass
With increasing altitude, up to about 40,000 ft, the characteristics of air change:
1 - Temperature decreases continuously with altitude
2 - Pressure falls steadily to an altitude of about 36,000 ft, where it then remains constant
3 - Density decreases steadily with increasing altitude
4 - Pressure falls steadily with increasing altitude
The combination of true statements is:
a)
b)
c)
d)
3 and 4
1,2 and 3
2 and 4
1 and 4
2 - 11
© Oxford Aviation Services Limited
THE ATMOSPHERE
PRINCIPLES OF FLIGHT
4.
When considering static pressure:
1 - In aviation, static pressure can be measured in hectopascal' s
2 - The SI units for static pressure is N/m 2
3 - Static pressure is the product of the mass of air pressing down on the air beneath
4 - Referred to as static pressure because of the air's stationary or static presence
5 - The lower the altitude, the greater the static pressure
The correct statements are:
a)
b)
c)
d)
5.
2,4 and 5
1,2,3,4and5
1,3 and 5
1 and 5
When considering air density:
I - Density is measured in millibar's
2 - Density increases with increasing altitude
3 - If temperature increases the density will increase
4 - As altitude increases, density will decrease
5 - Temperature decreases with increasing altitude, this will cause air density to increase
The combination of correct statements is:
a)
b)
c)
d)
6.
4 only
4 and 5
5 only
2,3 and 5
Air density is:
a)
b)
c)
d)
Mass per unit volume
Proportional to temperature and inversely proportional to pressure
Independent of both temperature and pressure
Dependent only on decreasing pressure with increasing altitude
2 - 12
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
7.
THE ATMOSPHERE
When considering the ICAO International Standard Atmosphere and comparing it with the actual
atmosphere, which of the following statements is correct:
1 - Temperature, pressure and density are constantly changing in any given layer of the actual
atmosphere
2 - A requirement exists for a hypothetical 'standard' atmosphere
3 - The values given in the International Standard Atmosphere exist at a the same altitudes in the
actual atmosphere
4 - The International Standard Atmosphere was designed for the calibration of pressure
instruments and the comparison of aircraft performance calculations
a)
b)
c)
d)
8.
1,2 and 3
2,3 and 4
1, 2, 3 and 4
1,2 and 4
When considering the ICAO International Standard Atmosphere, which of the following
statements is correct:
1 - The temperature lapse rate is assumed to be uniform at 2°C per 1,000 ft (1.98°C) up to a
height of 11,000 ft
2 - Sea level temperature is assumed to be 15°C
3 - Sea level static pressure is assumed to be 1.225 kg/m3
4 - Sea level density is assumed to be 1013.25 hPa
a)
b)
c)
d)
9.
1,2,3 and 4
No statements are correct
1,3 and 4
2 only
A moving mass of air possesses kinetic energy. An obj ect placed in the path of such a moving
mass of air will be subject to which of the following:
a)
b)
c)
d)
Dynamic pressure
Static Pressure
Static pressure and dynamic pressure
Dynamic pressure minus static pressure
2 - 13
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
10.
Dynamic pressure is:
a)
b)
c)
d)
11.
Static pressure
Dynamic pressure
Static pressure plus dynamic pressure
The difference between total pressure and static pressure
A static pressure vent must be positioned:
a)
b)
c)
d)
14.
Density times speed squared
Half the density times the indicated airspeed squared
Half the true airspeed times the density squared
Half the density times the true airspeed squared
A tube facing into an airflow will experience a pressure in the tube equal to:
a)
b)
c)
d)
13.
The total pressure at a point where a moving airflow is brought completely to rest
The amount by which the pressure rises at a point where a moving airflow is brought
completely to rest
The pressure due to the mass of air pressing down on the air beneath
The pressure change caused by heating when a moving airflow is brought completely
to rest
Dynamic pressure is equal to:
a)
b)
c)
d)
12.
THE ATMOSPHERE
On a part of the aircraft structure where the airflow is undisturbed, in a surface at right
angles to the airflow direction
On a part of the structure where the airflow is undisturbed, in a surface parallel to the
airflow direction
At the stagnation point
At the point on the surface where the airflow reaches the highest speed
The inputs to an Air Speed Indicator are from:
a)
b)
c)
d)
A static source
Pitot pressure
A pitot and a static source
Pitot, static and density
2 - 14
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
15.
The deflection of the pointer of the Air Speed Indicator is proportional to:
a)
b)
c)
d)
16.
Dynamic pressure
Static pressure
The difference between static and dynamic pressure
Static pressure plus dynamic pressure
Calibration of the Air Speed Indicator is based upon the density:
a)
b)
c)
d)
17.
THE ATMOSPHERE
At the altitude at which the aircraft is flying
At sea level ICAO International Standard Atmosphere temperature
At sea level
At sea level ICAO International Standard Atmosphere + 15°C temperature
When considering the relationship between different types of air speed:
1 - True Air Speed (TAS) is read directly from the Air Speed Indicator
2 - Equivalent Air Speed is Indicated Air Speed corrected for position error
3 - Indicated Air Speed is not a speed at all, it's a pressure
4 - True Air Speed is the speed of the aircraft through the air
Which of the above statements are true:
a)
b)
c)
d)
18.
1 only
2 and 3
3 and 4
1 and 4
When considering the relationship between different types of air speed:
1 - Calibrated Air Speed is Indicated Air Speed corrected for position error
2 - Equivalent Air Speed is Indicated Air Speed corrected for position error and compressibility
3 - Position error, which causes false Indicated Air Speed readings, is due to variations in the
pressures sensed at the pitot and static ports
4 - The Air Speed Indicator is calibrated to read True Air Speed when the ambient density is that
of the ICAO International Standard Atmosphere at sea level
The combination of correct statements is:
a)
b)
c)
d)
Non of the statements are correct
1,2 and 4
2 and 3
1, 2, 3 and 4
2 - 15
© Oxford Aviation Services Limited
THE ATMOSPHERE
PRINCIPLES OF FLIGHT
19.
The speed of sound:
a)
b)
c)
d)
20.
Mach number is:
a)
b)
c)
d)
21.
Is dependent upon the True Air Speed and the Mach number of the aircraft
Is inversely proportional to the absolute temperature
Is proportional to the square root of the absolute temperature of the air
Is directly proportional to the True Air Speed of the aircraft
The aircraft True Air Speed divided by the local speed of sound
The speed of sound in the ambient conditions in which the aircraft is flying
The True Air Speed of the aircraft at which the relative airflow somewhere on the
aircraft first reaches the local speed of sound
The Indicated Air Speed divided by the local speed of sound sea level
An aircraft's critical Mach number is;
(a)
(b)
(c)
(d)
The speed of the airflow when the aircraft first becomes supersonic
The speed of the aircraft when the airflow somewhere reaches the speed of sound
The Indicated Airspeed when the aircraft first becomes supersonic
The aircraft's Mach number when airflow over it first reaches the local speed of sound
2 - 16
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
THE ATMOSPHERE
ANSWERS
I No I A I B I c I D II
1
B
A
4
B
5
A
6
A
7
D
8
D
9
C
10
B
11
D
12
C
13
B
14
15
C
A
16
B
17
C
18
D
19
20
21
I
D
2
3
REF
C
A
D
2 - 17
© Oxford Aviation Services Limited
CHAPTER 3 - BASIC AERODYNAMIC THEORY
Contents
Page
THE PRINCIPLE OF CONTINUITY ........................................
BERNOULLI'S THEOREM ...............................................
STREAMLINES AND THE STREAMTUBE ..................................
SUMMARY ............................................................
SELF ASSESSMENT QUESTIONS .........................................
3-1
3-2
3-3
3-4
3-5
BASIC AERODYNAMIC THEORY
PRINCIPLES OF FLIGHT
3.1
THE PRINCIPLE OF CONTINUITY
One ofthe fundamental laws of the universe is "ENERGY and MASS can neither be created
nor destroyed", only changed from one form to another. To demonstrate the effect this basic
'Principle of Continuity' has on aerodynamic theory, it is instructive to consider a streamline
flow of air through a tube which has a reduced cross sectional area in the middle.
The air mass flow , or mass per unit time, through the tube will be the product of the crosssectional-area (A), the airflow velocity (V) and the air density (p). Mass flow will remain a
constant value at all points along the tube. The Equation of Continuity is:
A x V x p
=
Constant
Because air is a compressible fluid, any pressure change in the flow will affect the air density.
However, at low subsonic speeds « 0 . 4 M) density changes will be insignificant and can be
disregarded. The equation of continuity can now be simplified to: A x V = constant, or:
constant
Velocity (V) = - - Area (A)
I AirflON
Cross
Sectional
Area (A)
Velocity (V)
Mass FION
(Constant)
:>
~~~
I
1 m2
y, m 2
1 m2
52 m's (100 kt)
104 m's (200 kt)
52 m's (100 kt)
52 kg/s
Figure 3.1
52 kg/s
52 kg/s
The Principle of Continuity
Because the mass flow must remain constant, it can be seen from the equation of continuity that
the reduction in the tube's cross-sectional area results in an increase in velocity and, vice versa.
The equation of continuity enables the velocity changes of airflow around a given shape to be
predicted mathematically, « 0·4 M).
3-1
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
3.2
BASIC AERODYNAMIC THEORY
BERNOULLI'S THEOREM
"In the steady flow of an ideal fluid the sum of the pressure and kinetic energy per unit volume
remains constant".
(NB: An ideal fluid is both incompressible and has no viscosity).
This statement can be expressed as: Pressure + Kinetic energy
p
+hp
V
2
=
Constant or:
constant
Consider a mass of air: Static Pressure 101325 N/m2 , Density 1.225 kg/m3 and Velocity 52 mis,
its dynamic pressure will be: 1656 N/m2 • [Q = 'li x 1.225 x 52 x 52]
Pressure (101325 N/m2 ) + Kinetic energy (1656 N/m2 )
=
Constant (102981 N/m2 )
(200 kt)
(100 kt)
52 m/s
104 m/s
(100 kt)
52 m/s
1656 N/m2
6624 N/m2
1656 N/m 2
Static
Pressure
101325 N/m2
96357 N/m 2
101325 N/m 2
TOTAL
PRESSURE
102981 N/m 2
102981 N/m 2
102981 N/m 2
Dynamic
Pressure
Figure 3.2 Bernoulli's Theorem
Because the velocity of air at the throat has doubled, its dynamic pressure has risen by a value
of four, and the static pressure has decreased. The significant point is that:
Static Pressure + Dynamic Pressure is a constant. This constant can be referred to either
as:
TOTAL PRESSURE, STAGNATION PRESSURE or PITOT PRESSURE.
It can be seen that flow velocity is dependent on the shape of the object over which it flows.
And, from Bernoulli's theorem it is evident that an increase in velocity will cause a decrease in
static pressure, and vice versa.
3-2
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
BASIC AERODYNAMIC THEORY
The tube illustrated in Figures 3.1 and 3.2 is used only to demonstrate the Principle of Continuity
and Bernoulli's Theorem and is of no practical use in making an aeroplane fly.
But, an aerodynamic force to oppose the weight of an aircraft can be generated by using a
specially shaped body called an aerofoil.
Figure 3.3
"Typical" Aerofoil Section
The airflow velocity over the top surface of a lifting aerofoil will be greater than that beneath,
so the pressure differential that results will produce a force per unit area acting upwards. The
larger the surface area, the bigger the force that can be generated.
In the next section we see that the flow over the top of the aero foil looks very like the tube in
Fig. 3.2 and the principle of continuity and Bernoulli's theorem still apply.
3.3
STREAMLINES AND THE STREAMTUBE
A streamline is the path traced by a particle of air in a steady airflow, and streamlines cannot
cross. When streamlines are shown close together it illustrates increased velocity and vice versa.
Diverging streamlines illustrate a decelerating airflow and resultant increasing pressure and
converging streamlines illustrate an accelerating airflow, with resultant decreasing pressure.
STREAMTUBE
Figure 3.4 Streamlines and a Streamtube
A streamtube is an imaginary tube made of streamlines. There is no flow into or out of the
streamtube through the "walls", only a flow along the tube. With this concept it is possible to
visualise the airflow around an aerofoil being within a tube made-up of streamlines.
3-3
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
3.4
BASIC AERODYNAMIC THEORY
SUMMARY
(I)
At flow speeds of less than about 0·4 M, pressure changes will not affect air density.
(2)
Continuity:
(3)
(4)
Bernoulli:
(a)
Air accelerates when the cross-sectional-area of a streamline
flow is reduced
(b)
Air decelerates when the cross-sectional-area increases again.
(a)
If a streamline flow of air accelerates, its kinetic energy will
increase and its static pressure will decrease.
(b)
When air decelerates, the kinetic energy will decrease and the
static pressure will increase again.
By harnessing the Principle of Continuity and Bernoulli's Theorem an aerodynamic
force can be generated.
3-4
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
BASIC AERODYNAMIC THEORY
SELF ASSESSMENT QUESTIONS
1.
If the cross sectional area of an airflow is mechanically reduced:
a)
b)
c)
d)
2.
The statement, "Pressure plus Kinetic energy is constant", refers to:
a)
b)
c)
d)
3.
Bernoulli's theorem
The principle of continuity
Newton's second law of motion
The Magnus effect
If the velocity of an air mass is increased:
a)
b)
c)
d)
4.
The velocity of the airflow remains constant and the kinetic energy increases
The velocity of the airflow remains constant and the mass flow increases
The mass flow remains constant and the static pressure increases
The mass flow remains constant and the velocity of the airflow increases
The dynamic pressure will decrease and the static pressure will increase
The static pressure will remain constant and the kinetic energy will increase
The kinetic energy will increase, the dynamic pressure will increase and the static
pressure will decrease
The mass flow will stay constant, the dynamic pressure will decrease and the static
pressure will increase
When considering a streamlined airflow, which of the following statements is correct:
1 - A resultant decrease in static pressure is indicated by streamlines shown close together
2 - An increase in velocity is indicated by streamlines shown close together
3 - Accelerating airflow with a resultant decreasing static pressure is indicated by converging
streamlines
4 - Diverging streamlines indicate decelerating airflow with a resultant increasing static pressure
a)
b)
c)
d)
2 and 4
1,3 and 4
2,3 and 4
1, 2, 3 and 4
3-5
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
5.
If the pressure on one side of a surface is lower than on the other side:
a)
b)
c)
d)
6.
BASIC AERODYNAMIC THEORY
A force per unit area will exist, acting in the direction of the lower pressure
No force will be generated, other than drag
A force will be generated, acting in the direction of the higher pressure
The pressure will leak around the sides of the surface, cancelling-out any pressure
differential
When considering a streamtube, which of the following statements is correct:
1 - Different sizes of streamtube are manufactured to match the wing span of the aircraft to
which they will be fitted
2 - A streamtube is a concept to aid understanding of aerodynamic force generation
3 - There is no flow into or out of the streamtube through the "walls", only flow along the tube
4 - A streamtube is an imaginary tube made-up of streamlines
a)
b)
c)
d)
7.
At flow speeds less than four tenths the speed of sound, the following will be insignificant:
a)
b)
c)
d)
8.
1 only
1 and 3
2,3 and 4
1,2 and 3
Changes in static pressure due to temperature
Changes in density due to static pressure
Changes in density due to dynamic pressure
Changes in static pressure due to kinetic energy
In accordance with the principle of continuity:
1 - Air accelerates when the cross-sectional area of a streamline flow is reduced
2 - When air accelerates the density of air in a streamline flow is increased
3 - Air decelerates when the cross-sectional area of a streamline flow is increased
4 - Changes in cross-sectional area of a streamline flow will affect the air velocity
Which of the preceding statements are true:
(a)
(b)
(c)
(d)
1,2,3and4
1 and 4
3 and 4
1,3 and 4
3-6
© Oxford Aviation Services Limited
BASIC AERODYNAMIC THEORY
PRINCIPLES OF FLIGHT
9.
In accordance with Bernoulli's theorem:
1 - If a streamline flow of air decelerates, its kinetic energy will decrease and the static pressure
will increase
2 - If a streamline flow of air accelerates, its kinetic energy will increase and the static pressure
will decrease
3 - If a streamline flow of air is accelerated, the dynamic pressure will increase and the static
pressure will increase
4 - If a streamline flow of air is decelerated, its dynamic pressure will decrease and the static
pressure will increase
the combination of correct statements is:
a)
b)
c)
d)
10.
1,2,3 and 4
3 only
1,2 and 4
3 and 4
The statement, "Energy and mass can neither be created nor destroyed, only changed from one
form to another", refers to:
a)
b)
c)
d)
Bernoulli's theorem
The equation of kinetic energy
The principle of continuity
Bernoulli's principle of continuity
3-7
© Oxford Aviation Services Limited
BASIC AERODYNAMIC THEORY
PRINCIPLES OF FLIGHT
I No I A I B I c I DII
1
2
I
D
A
3
C
D
4
5
REF
A
6
C
7
C
D
8
9
C
28
C
3-9
© Oxford Aviation Services. Limited
CHAPTER 4 - SUBSONIC AIRFLOW
Contents
Page
AEROFOIL TERMINOLOGY .............................................. 4 - 2
BASICS ABOUT AIRFLOW ............................................... 4 - 4
TWO DIMENSIONAL AIRFLOW .......................................... 4 - 4
INFLUENCE OF DYNAMIC PRESSURE .............................. 4 - 5
INFLUENCE OF ANGLE OF ATTACK ............................... 4 - 6
CENTRE OF PRESSURE ........................................... 4 - 8
MOVEMENT OF THE CENTRE OF PRESSURE
AERODYNAMIC FORCE COEFFICIENT
DEVELOPMENT OF AERODYNAMIC PITCHING MOMENTS .......... 4- 10
AERODYNAMIC CENTRE
PITCHING MOMENTS OF A SYMMETRICAL AEROFOIL ............. 4 - 11
SUMMARY ........................................................... 4 - 12
SELF ASSESSMENT QUESTIONS ........................................ 4 - 13
'lJ
"SUCTION" PEAK DUE TO ACCELERATED
FLOW AROUND LEADING EDGE PROFILE
(INCREASING KINETIC ENERGY
DECREASING STATIC PRESSURE) I
"
~
LIFT FORCE
z
o
ANr(-.. . '
II
'lJ
r
m
en
o."
" ,
I
I
I
'
,,
,,
,,
."
r
(5
"
::I:
-f
"'\"
,
,
\
\
\
\
\
\
\
\
\
FLOW DECELERATING
(DECREASING KINETIC ENERGY)
',DUE TO ADVERSE PRESSURE GRADIENT
,
(-)
"""'"
\
\
\
(PRESSURE INCREASING FROM MINIMUM
STATIC PRESSURE BACK TO FREE STREAM)
\
"",
\
UPWASH IN FRONT OF AEROFOIL
BECAUSE OF LOWER PRESSURE
ON TOP SURFACE
~
----
---------..
T
L
-
(HIGHER THAN STATIC)
en
c
[D
en
oz
o
~
::0
."
r
o
~
SUBSONIC AIRFLOW
PRINCIPLES OF FLIGHT
MAXIMUM
THICKNESS
LOCATION OF
MAX. THICKNESS
MAXIMUM
CAMBER
LEADING
EDGE ,
RADIUS
_L __
MEAN CAMBER LINE
_ _ _C
=-:H
-'-O
=-:R
--"D
=---_ __ _ .___._______~
LOCATION OF
MAX. CAMBER
TOTAL
REACTION
~
LIFT
/
/1
I
/
/
/
/
ANGLE OF
ATTACK
RELATIVE AIRFLOW
/
..-.....~.I DRAG
.-------------------- AIRCRAFT FLiGHTPATH
Figure 4.1
4.1
AEROFOIL TERMINOLOGY
Aerofoil: A shape capable of producing lift with relatively high efficiency.
Chord Line: A straight line joining the centres of curvature of the leading and trailing edges
of an aerofoil.
Chord: The distance between the leading and trailing edges measured along the chord line.
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PRINCIPLES OF FLIGHT
SUBSONIC AIRFLOW
Angle of Incidence: The angle between the chord line and the horizontal datum of the aircraft.
(This angle is fixed for the wing, but may be variable for the tailplane).
Mean Line or Camber Line: A line joining the leading and trailing edges of an aerofoil,
equidistant from the upper and lower surfaces.
Maximum Camber: The maximum distance of the mean line from the chord line. Maximum
camber is expressed as a percentage of the chord, with its location as a percentages of the chord
aft of the leading edge. When the camber line lies above the chord line the aerofoil is said to
have positive camber, and if the camber line is below the chord line it is said to have negative
camber. A symmetrical aerofoil has no camber because the chord line and camber line are coincidental.
Thickness/Chord ratio: The maximum thickness or depth of an aerofoil section expressed as
a percentage of the chord, with its location as a percentages of the chord aft of the leading edge.
The thickness and thickness distribution of the aero foil section have a great influence on its
airflow characteristics.
Leading edge radius: The radius of curvature of the leading edge. The size of the leading edge
radius can significantly effect the initial airflow characteristics of the aerofoil section.
Relative Air Flow (Relative Wind or Free Stream Flow): Relative Air Flow has three qualities.
(1) DIRECTION - air parallel to, and in the opposite direction to the flight path of the aircraft,
in fact the path of the CG; the direction in which the aircraft is pointing is irrelevant.
(2) CONDITION - air close to, but unaffected by the presence of the aircraft; its pressure,
temperature and velocity are not affected by the passage of the aircraft through it.
(3) MAGNITUDE - The magnitude of the Relative Air Flow is the TAS.
If air flow does not possess all three of these qualities, it is referred to as EFFECTIVE
AIRFLOW.
Total Reaction: The resultant of all the aerodynamic forces acting on the aerofoil section.
Centre of Pressure (CP): The point on the chord line, through which Lift is considered to act.
Lift: The aerodynamic force which acts at 90 ,to the Relative Air Flow.
0
Drag: The aerodynamic force which acts parallel to and in the same direction as the Relative
Air Flow (or opposite to the aircraft flight path).
Angle of Attack (a or alpha) (can also be referred to as Aerodynamic Incidence) The angle
between the chord line and the Relative Air Flow.
The angle between the chord line and the effective airflow is referred to as the
EFFECTIVE ANGLE OF ATTACK.
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PRINCIPLES OF FLIGHT
4.2
SUBSONIC AIRFLOW
BASICS ABOUT AIRFLOW
When considering airflow velocity, it makes no difference to the pressure pattern if the aircraft
is moving through the air or the air is flowing over the aircraft: it is the relative velocity which
is the important factor. To promote a full understanding, references will be made to both windtunnel experiments, where air is flowing over a stationary aircraft, and aircraft in flight moving
through 'stationary' air.
Three dimensional airflow: Three dimensional flow is the true airflow over an aircraft and
consists of a hypothetical two dimensional flow modified by various pressure differentials.
Three dimensional airflow will be examined later.
Two dimensional airflow: Assumes a wing with the same aerofoil section along the entire span
with no spanwise pressure differential or flow.
4.3
TWO DIMENSIONAL AIRFLOW
This CONCEPT, figures 4.2 and 4.3, is used to illustrate the basic principles of aerodynamic
force generation.
As air flows towards an aerofoil it will be turned towards the lower pressure at the upper surface;
this is termed upwash. After passing over the aerofoil the airflow returns to its original position
and state; this is termed downwash.
Figure 4.2
DOWNWASH
UPWASH
Figure 4.3
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PRINCIPLES OF FLIGHT
SUBSONIC AIRFLOW
Influence of Dynamic Pressure: Fig. 4.4 shows an aerofoil section at a representative angle
of attack subject to a given dynamic pressure (lAS). "If the static pressure on one side of a body
is reduced more than on the other side, a pressure differential will exist".
Fig. 4.5 shows the same aero foil section at the same angle of attack, but subject to a higher
dynamic pressure (lAS). "If the dynamic pressure (lAS) is increased, the pressure differential
will increase".
REPRESENTATIVE ANGLE OF ATTACK
AND A GIVEN DYNAMIC PRESSURE
Figure 4.4
SAME ANGLE OF ATTACK
INCREASED DYNAMIC PRESSURE
Figure 4.5
The pressure differential acting on the surface area will produce an upward acting force. "If the
dynamic pressure (lAS) is increased, the upward force will increase".
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PRINCIPLES OF FLIGHT
SUBSONIC AIRFLOW
Influence of Angle of Attack: At a constant dynamic pressure (lAS), increasing the angle of
attack (up to about 16°) will likewise increase the pressure differential, but will also change the
pattern of pressure distribution.
The aerofoil profile presented to the airflow will determine the distribution of velocity and hence
the distribution of pressure on the surface. This profile is determined by the aerofoil geometry,
i.e. thickness and distribution (fixed), camber and distribution (assumed to be fixed for now)
and by the angle of attack (variable).
The greatest positive pressure occurs at the stagnation point where the relative flow velocity is
zero. This stagnation point is located somewhere near the leading edge. As the angle of attack
increases from - 4 0 the leading edge stagnation point moves from the upper surface around the
leading edge to the lower surface. It is at the front stagnation point where the flow divides to
pass over and under the section. The pressure at the stagnation point is Static + Dynamic.
The flow over the top of the section accelerates rapidly around the nose and over the leading
portion of the surface - inducing a substantial decrease in static pressure in those areas. The rate
of acceleration increases with increase in angle of attack, up to about 16 0 • (Anything which
changes the accurately manufactured profile of the leading portion of the surface can seriously
disrupt airflow acceleration in this critical area. e.g. ice, snow, frost, dirt or dents). The pressure
reduces continuously from the stagnation value through the free stream value to a position on
the top surface where a peak negative value is reached. From there onwards the flow
continuously slows down again and the pressure increases back to the free stream value in the
region of the trailing edge.
0
At angles of attack less than 8 the flow under the section is accelerated much less, reducing the
pressure to a small negative value, also with subsequent deceleration and increase in pressure
back to the free stream value in the region of the trailing edge.
The pressure differential between the leading edge stagnation point and the lower pressure at the
trailing edge creates a force acting backward which is called 'form (pressure) drag' . (This will
be discussed in more detail later).
Angle of Attack ( -4 0) - The decrease in pressure above and below the section are equal and no
differential exists. There will, thus, be no lift force. (Fig. 4.6). This can be called the "zero lift
angle of attack".
Figure 4.6
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PRINCIPLES OF FLIGHT
SUBSONIC AIRFLOW
Angles of Attack (0° to 8°) - Compared to
free stream static pressure, there is a pressure
decrease over the upper surface and a lesser
decrease over most of the lower surface. For
a cambered aerofoil there will be a small
amount oflift even at small negative angles (4 ° to 0 °).
~
-)
~~.._~O_-Ic-I.
~
(-)
Angles of attack (0° to 16°) - Increasing the
angle of attack increases the lift force because
the acceleration of the airflow over the top
surface is increased by the reduction in
effective cross-sectional area of the local
streamtube.
The reduced pressure 'peak' moves
forward as the angle of attack increases.
The greatest contribution to overall lift
comes from the upper surface.
Pressure Gradient: Is a change 10 air
pressure over distance. The greater the
difference in pressure between two points, the
steeper the gradient. A favourable gradient is
when air pressure is falling in the direction of
airflow. An adverse pressure gradient is when
air pressure is rising in the direction of
airflow, such as between the point of
minimum pressure on the top surface and the
trailing edge. The higher the angle of attack,
the steeper the pressure gradient. At angles of
attack higher than approximately 16°, the
extremely steep adverse pressure gradient
prevents air that is flowing over the top
, surface from following the aerofoil contour
and the previously smooth streamline flow
will separate from the surface, causing the low
pressure area on the top of the section to
suddenly collapse. Any pressure differential
remaining is due to the pressure increase on
the lower surface only. This condition is
known as the stall and will be described in
detail in Chapter 7.
Figure 4.7
4-7
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PRINCIPLES OF FLIGHT
SUBSONIC AIRFLOW
Centre of Pressure (CP): The whole surface of the aerofoil contributes to lift, but the point
along the chord where the distributed lift is effectively concentrated is termed the Centre of
Pressure (Fig. 4.8). The location of the CP is a function of camber and section lift coefficient.
i.e. angle of attack.
AERODYNAMIC FORCE
LIFT
Figure 4.8
Movement of the Centre of Pressure: As the angle of attack increases from 0 0 to 16 0 the
upper ' suction' peak moves forward (Fig. 4.7) so the point at which the lift is effectively
concentrated, the CP, will move forward . The CP moves forward and the magnitude ofthe lift
force increases with increase in angle of attack until the stall is reached when the lift force
decreases abruptly and the CP generally moves back along the chord (Fig. 4.9). Note that the
CP is at its most forward location just before the stall (C L ~
Aerodynamic Force Coefficient: A coefficient is a dimensionless number expressing degree
of magnitude. An aerodynamic force coefficient is a common denominator for all N C of
whatever weight, size and speed. An aerodynamic force coefficient is a dimensionless ratio
between the average aerodynamic pressure and the airstream dynamic pressure.
By this definition a lift coefficient (C L) is the ratio between lift divided by the wing planform
area and dynamic pressure and a drag coefficient (C D) is the ratio between drag divided by the
wing planform area and dynamic pressure.
The use of the coefficient of an aerodynamic force is necessary since the force coefficient is:
a)
An index of the aerodynamic force independent of area, density and velocity. It is
derived from the relative pressure and velocity distribution.
b)
Influenced only by the shape of the surface and angle of attack since these factors
determine the pressure distribution.
4-8
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PRINCIPLES OF FLIGHT
SUBSONIC AIRFLOW
(5
~
~
«
u..
o
UJ
...J
Cl
z
«
o
10%
20%
30 %
40%
50%
60%
70%
80%
90%
100%
CP POSITION
(Percentage chord , aft of leading edge)
LEADING
EDGE
TRAILING
EDGE
Figure 4.9 CP Movement with Angle of Attack
4-9
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PRINCIPLES OF FLIGHT
SUBSONIC AIRFLOW
Development of Aerodynamic Pitching Moments: The distribution of pressure over a surface
is the source of aerodynamic moments as well as forces. There are two ways to consider the
effects of changing angle of attack on the pitching moment of an aerofoil.
a)
Changes in the magnitude oflift acting through a moving CP, or more simply:
b)
Changes in the magnitude of lift always acting through an Aerodynamic Centre, which
is fixed.
Aerodynamic Centre (AC): The AC is a 'fixed' point on the chord line and is defined as: 'The
point where all changes in the magnitude of the lift force effectively take place', AND: 'The
point about which the pitching moment will remain constant at 'normal' angles of attack' . A
nose-down pitching moment exists about the AC which is the product of a force (lift at the
CP) and an arm (distance from the CP to the AC). Since an increase in angle of attack will
increase the lift force, but also move the CP towards the AC (shortening the lever arm), the
moment about the AC remains the same at any angle of attack within the "normal" range.
Figure 4.10 ,
MOMENT (M) REMAINS THE SAME AT 'NORMAL' ANGLES OF ATTACK BECAUSE
When considering subsonic airflows ofless than MO-4, the AC is located at the 25 % chord point
for any aerofoil regardless of camber, thickness and angle of attack.
The aerodynamic centre (AC) is an aerodynamic reference point. The most direct application
being to the longitudinal stability of an aircraft, which will be discussed in Chapter 10.
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PRINCIPLES OF FLIGHT
SUBSONIC AIRFLOW
Pitching Moment for a Symmetrical Aerofoil: Note the change in pressure distribution with
angle of attack for the symmetrical aerofoil in Fig 4.11. When at zero angle of attack, the upper
and lower surface forces are equal and located at the same point. With an increase in angle of
attack, the upper surface force increases while the lower surface force decreases. A change in
the magnitude of lift has taken place with no change in the CP position - a characteristic of
symmetrical aerofoils. Thus, the pitching moment about the AC for a symmetrical aerofoil will
be zero at ' normal' angles of attack - one of the big advantages of symmetrical aero foils.
SYMMETRICAL AERO FOIL
AT ZERO LIFT
SYMMETRICAL AEROFOIL
AT POSITIVE LIFT
Figure 4.11
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SUBSONIC AIRFLOW
PRINCIPLES OF FLIGHT
4.4
SUMMARY
Airflow pattern, and ultimately lift and drag, will depend upon:
a)
Angle of attack - airflow cross-sectional area change
b)
Aerofoil shape (thickness & camber). - airflow cross-sectional area change
c)
Air density - mass flow of air (decreases with increased altitude)
d)
Velocity - mass flow of air (changes with aircraft T AS)
The Lift force is the result of the pressure differential between the top and bottom surfaces of
an aerofoil; the greatest contribution to overall lift comes from the top surface.
Anything (Ice in particular, but also frost, snow, dirt, dents and even water droplets) which
changes the accurately manufactured profile of the leading portion of the upper surface can
seriously disrupt airflow acceleration in that area, and hence the magnitude of the lift force.
An increase in dynamic pressure (lAS) will increase the lift force, and vice versa.
An increase in angle of attack will increase the lift force, and vice versa, (0° to 16°)
The centre of pressure (CP) of a cambered aerofoil moves forward as the angle of attack
increases. The (CP) of a symmetrical aerofoil does not move under the influence of angle of
attack. (within the confines of 'normal range').
Throughout the normal range of angles of attack the aerofoil nose down pitching moment about
the aerodynamic centre (AC) will remain constant. The AC is located at the quarter chord
position for subsonic flow of less than MO·4.
The coefficient of lift (C L) is the ratio between lift per unit wing area and dynamic pressure.
As the angle of attack increases from -4 ° the leading edge stagnation point moves from the
upper surface around the leading edge to the lower surface.
The greatest positive pressure occurs at the leading edge stagnation point, where the relative
flow velocity is zero.
Form (pressure) drag is the result of the pressure differential between the leading edge and
trailing edge of the aero foil.
An increase in dynamic pressure (lAS) will increase form drag, and vice versa.
The coefficient of drag (CD) is the ratio between drag per unit wing area and dynamic pressure.
4 - 12
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SUBSONIC AIRFLOW
PRINCIPLES OF FLIGHT
SELF ASSESSMENT QUESTIONS
1.
With reference to aerofoil section terminology, which of the following statements is true:
1 - The chord line is a line joining the centre of curvature of the leading edge to the centre of the
trailing edge, equidistant from the top and bottom surface of the aerofoil.
2 - The angle of incidence is the angle between the chord line and the horizontal datum of the
aircraft.
3 - The angle between the chord line and the relative airflow is called the aerodynamic incidence
or angle of attack.
4 - The thickness/chord ratio is the maximum thickness of the aerofoil as a percentage of the
chord; the location of maximum thickness is measured as a percentage of the chord aft of the
leading edge.
a)
b)
c)
d)
2.
The definition of lift is:
a)
b)
c)
d)
3.
the aerodynamic force which acts perpendicular to the chord line of the aerofoil
the aerodynamic force that results from the pressure differentials about an aerofoil
the aerodynamic force which acts perpendicular to the upper surface of the aerofoil
the aerodynamic force which acts at 90 to the relative airflow
0
An aerofoil section is designed to produce lift resulting from a difference in the:
a)
b)
c)
d)
4.
1,2,3and4
1,2 and 4
2,3 and 4
2 and 4
negative air pressure below and a vacuum above the surface.
vacuum below the surface and greater air pressure above the surface.
higher air pressure below the surface and lower air pressure above the surface.
higher air pressure at the leading edge than at the trailing edge.
On an aerofoil section, the force of lift acts perpendicular to, and the force of drag acts parallel
to the:
a)
b)
c)
d)
flightpath.
longitudinal axis.
chord line.
aerofoil section upper surface.
4 - 13
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PRINCIPLES OF FLIGHT
5.
When the angle of attack of a symmetrical aerofoil is increased, the centre of pressure will:
a)
b)
c)
d)
6.
b)
c)
d)
midpoint of the chord.
centre of gravity.
centre of pressure.
aerodynamic centre.
The angle between the chord line of the aerofoil section and the longitudinal axis of the aircraft
is known as:
a)
b)
c)
d)
9.
The increased impact of the relative wind on an aerofoil's lower surface creates a greater
amount of air being deflected downward.
The increased speed of the air passing over an aerofoil's upper surface decreases the
static pressure, thus creating a greater pressure differential between the upper and lower
surface.
The increased velocity of the relative wind overcomes the increased drag.
Increasing speed decreases drag.
The point on an aerofoil section through which lift acts is the:
a)
b)
c)
d)
8.
have very limited movement.
move aft along the aerofoil surface.
remain unaffected.
move forward to the leading edge.
Why does increasing speed also increase lift?
a)
7.
SUBSONIC AIRFLOW
the angle of attack.
the angle of incidence.
dihedral.
sweep back.
The angle between the chord line of an aerofoil section and the relative wind is known as the
angle of:
a)
b)
c)
d)
incidence.
lift.
attack.
sweepback
4 - 14
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SUBSONIC AIRFLOW
PRINCIPLES OF FLIGHT
10.
A line drawn from the leading edge to the trailing edge of an aerofoil section and equidistant at
all points from the upper and lower contours is called the:
a)
b)
c)
d)
11.
At zero angle of attack, the pressure along the upper surface of a symmetrical aerofoil section
would be:
a)
b)
c)
d)
12.
amount of airflow above and below the section.
angle of incidence of the section.
distribution of positive and negative pressure acting on the section.
the angle relative to the horizontal datum
When the angle of attack of a positively cambered aerofoil is increased, the centre of pressure
will:
a)
b)
c)
d)
14.
greater than atmospheric pressure.
equal to atmospheric pressure.
less than atmospheric pressure.
non existent.
The angle of attack of an aerofoil section directly controls:
a)
b)
c)
d)
13.
chord line.
camber.
mean camber line.
longitudinal axis.
have very little movement.
move forward along the chord line.
remain unaffected.
move back along the chord
The term "angle of attack" is defined as the angle:
a)
b)
c)
d)
formed by the longitudinal axis of the aeroplane and the chord line of the section.
between the section chord line and the relative wind.
between the aeroplane's climb angle and the horizon.
formed by the leading edge of the section and the relative airflow.
4 - 15
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PRINCIPLES OF FLIGHT
15.
SUBSONIC AIRFLOW
Which of the following statements is true:
1 - Relative airflow, free stream flow, relative wind and aircraft flightpath are parallel.
2 - Aircraft flightpath, relative airflow, relative wind and free stream flow are parallel, but the
aircraft flightpath is opposite in direction.
3 - The pressure, temperature and relative velocity of the free stream flow are unaffected by the
presence of the aircraft.
4 - The relative wind is produced by the aircraft moving through the air.
5 - The direction of flight is parallel with and opposite to the relative airflow.
a)
b)
c)
d)
16.
5 only
3,4 and 5
1 and 2
1,2,3,4 and 5
Which of the following statements is correct:
1 - Maximum camber is the maximum distance between the top and bottom surface of an aerofoil
section.
2 - The thickness/chord ratio is expressed as a percentage of the chord.
3 - It is easier for air to flow over a well rounded leading edge radius than a sharp leading edge.
4 - Two dimensional airflow assumes a wing with the same aerofoil section along its entire span,
with no spanwise pressure differential.
5 - Air flowing towards the lower pressure of the upper surface is called upwash.
a)
b)
c)
d)
17.
1,2,3,4and5
2,3 and 4
2, 3, 4 and 5
1 and 5
When considering an aerofoil section at a constant angle of attack, which of the following
statements is true:
a)
b)
c)
d)
If the static pressure on one side is reduced more than on the other side, a pressure
differential will exist.
If dynamic pressure is increased, the pr€ssure differential will decrease.
The pressure differential will increase if the dynamic pressure is decreased
Dynamic pressure and pressure differential are not related.
4 - 16
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SUBSONIC AIRFLOW
PRINCIPLES OF FLIGHT
18.
Considering an aerofoil section subject to a constant dynamic pressure, which of the following
statements is correct:
a)
b)
c)
d)
19.
0
0
If the angle of attack is increased from 4 to 14 the pressure differential will not
change but lift will be greater due to increased dynamic pressure acting on the lower
surface.
Up to about 16 increasing the angle of attack will increase the pressure differential
between the top and bottom surface of the aerofoil.
Changing the angle of attack does not affect the pressure differential, only changes in
dynamic pressure affect the pressure differential.
Up to about 16 increasing the angle of attack decreases the pressure differential
between the top and bottom surface of the aerofoil section.
0
,
0
,
When considering the effect of changing angle of attack on the pitching moment of an aerofoil,
which of the following statements is correct:
1 - At 'normal' angles of attack the pitching moment is nose up.
2 - The pitching moment about the aerodynamic centre (AC) is constant at 'normal' angles of
attack.
3 - The aerodynamic centre (AC) is located approximately at the 25% chord point.
4 - The moment about the aerodynamic centre (AC) is a product of the distance between the
aerodynamic centre (AC) and the centre of pressure (CP) and the magnitude of the lift force.
a)
b)
c)
d)
20.
1,2,3 and 4
4 only
3 and 4
2,3 and 4
Ice contamination of the leading portion of the aerofoil has which of the following
consequences:
1 - The profile of the leading portion of the surface can be changed, preventing normal
acceleration of the airflow and substantially reducing the magnitude of the lift force.
2 - Form (pressure) drag will be increased because of the increased frontal area of the aerofoil
section.
3 - Loss of lift will have a greater effect than all increase in form (pressure) drag.
4 - At 'normal' angles of attack lift can be lost entirely if enough ice accumulates.
a)
b)
c)
d)
1, 2, 3 and 4
1,3 and 4
1,2 and 3
3 and 4
4 - 17
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SUBSONIC AIRFLOW
PRINCIPLES OF FLIGHT
I No I A I B I c
1
C
2
D
3
4
C
A
5
6
C
REF
III
No
IA I B I c ID
11
C
12
C
13
B
14
B
15
C
17
9
C
19
10
C
20
4 - 19
REF
I
C
A
18
B
I
D
16
B
7
8
I DII
B
D
A
© Oxford Aviation Services Limited
CHAPTER 5 - LIFT
Contents
Page
AERODYNAMIC FORCE COEFFICIENT .................................... 5 - 1
THE BASIC LIFT EQUATION ............................................. 5 - 2
INTERPRETATION OF THE BASIC LIFT EQUATION
THE LIFT CURVE ....................................................... 5 - 6
INTERPRETATION OF THE LIFT CURVE
VELOCITY - DYNAMIC PRESSURE RELATIONSHIP ........................ 5 - 8
DENSITY ALTITUDE ............................................. 5 - 9
AEROFOIL SECTION LIFT CHARACTERISTICS
MINIMUM FLIGHT SPEEDS
INTRODUCTION TO DRAG CHARACTERISTICS ........................... 5 - 10
LIFT/DRAG RATIO
EFFECT OF WEIGHT ON MINIMUM FLIGHT SPEED ....................... 5 - 12
CONDITION OF THE SURFACE
FLIGHT AT HIGH LIFT CONDITIONS
EFFECTS OF HIGH LIFT DEVICES
THREE DIMENSIONAL AIRFLOW ....................................... 5 - 15
WING TERMINOLOGY
WING AREA
WINGSPAN
AVERAGE CHORD
ASPECT RATIO
ROOT CHORD
TIP CHORD
TAPER RATIO
SWEEP ANGLE
MEAN AERODYNAMIC CHORD
WING TIP VORTICES .................................................. 5 - 16
INDUCED DOWNWASH
WAKE TURBULENCE .................. ' ...... , ................... , ..... 5 - 18
WAKE VORTEX CHARACTERISTICS
DISTRIBUTION OF TRAILING VORTICES .......................... 5 - 19
VORTEX MOVEMENT NEAR THE GROUND ....................... 5 - 20
THE DECAY PROCESS OF TRAILING VORTICES ................... 5 - 21
PROBABILITY OF WAKE TURBULENCE ENCOUNTER
WAKE TURBULENCE AVOIDANCE
GROUND EFFECT .....................................................
THE IMPACT OF GROUND EFFECT
HIGH AND LOW TAIL CHARACTERISTICS .........................
INFLUENCE OF TAILPLANE CAMBER ON PITCHING MOMENT ......
TAILPLANE ANGLE OF ATTACK .................................
ENTERING GROUND EFFECT ....................................
LEAVING GROUND EFFECT .....................................
SUMMARY ...........................................................
ANSWERS FROM PAGE 5 - 7 ............................................
ANSWERS FROM PAGE 5 - 8 ............................................
SELF ASSESSMENT QUESTIONS ........................................
5 - 22
5 - 23
5 - 24
5 - 25
5 - 26
5 - 27
5 - 28
5 - 29
5 - 30
5 - 31
LIFT
PRINCIPLES OF FLIGHT
5.1
AERODYNAMIC FORCE COEFFICIENT
The aerodynamic forces of both lift and drag depend on the combined effect of many variables.
The important factors being:
a)
b)
c)
d)
e)
f)
g)
Airstream velocity (Y) }
Dynamic Pressure (Yl p y2)
Air density (p)
Shape or profile of the surface }
Pressure Distribution (C Lor CD)
Angle of attack
Surface area (S)
Condition of the surface
Compressibility effects (to be considered in later chapters)
Dynamic Pressure: The dynamic pressure CI2 p y2) of the airflow is a common denominator
of aerodynamic forces and is a major factor since the magnitude of a pressure distribution
depends on the energy given to the airflow (KE = 1/2 m y2).
Pressure Distribution: Another major factor is the relative pressure distribution existing on
the surface. The distribution of velocities, with resulting pressure distribution, is determined by
the shape or profile of the surface and the angle of attack (C Lor CD).
Surface Area: Since aerodynamic forces are the result of various pressures distributed on a
surface, the surface area (S) is the remaining major factor - the larger the surface area for a given
pressure differential, the greater the force generated.
Thus, any aerodynamic force can be represented as the product of three major factors:
1.
2.
3.
The dynamic pressure of the airflow CI2 p y2)
The coefficient of force determined by the relative pressure distribution (C Lor C D)' and
The surface area of the object (S)
The relationship of these three factors is expressed by the following equation:
F
where, F
=
=Q
CF S
aerodynamic force (Lift or Drag)
Q = dynamic pressure (Yl p y2 )
C F = coefficient of aerodynamic force (C Lor CD)
S = surface area (S)
5-1
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LIFT
PRINCIPLES OF FLIGHT
5.2
THE BASIC LIFT EQUATION
Lift is defined as the net force generated normal (at 90 0 ) to the relative airflow or flight path of
the aircraft.
The aerodynamic force of lift results from the pressure differential between the top and bottom
surfaces of the wing. This lift force can be defined by the following equation:
Correct interpretation of the Lift formula is a key element in the complete understanding of
Principles of Flight.
ANGLE OF ATTACK
/
TAS
TO
CL
BALANCE--- L =
WEIGHT
S-
FIXED WING AREA
~ DYNAMIC PRESSURE (lAS)
AIR DENSITY
NB.
For the sake of clarity; during this initial examination of the lift formula it is stated that C L is
determined by angle of attack. This is true, but CL is also influenced by the shape or profile of
the surface and other factors which will be amplified in later sections.
•
An aircraft spends most of its time in straight and level flight.
How much lift is required?
The same as the weight.
•
Consider that at any moment in time weight is constant, so lift must be constant.
•
While generating the required lift force the less drag the better, because drag has to be
balanced by thrust and thrust costs money.
The value of lift divided by drag is a measure of aerodynamic efficiency. This has a
maximum value at one particular angle of attack. For a modem wing this is about 4 0 •
If this "optimum" angle of attack is maintained, maximum aerodynamic efficiency will
be achieved. Note : Maximum CL and minimum CD are not obtained at best LID.
5-2
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PRINCIPLES OF FLIGHT
LIFT
Lift is generated by a pressure differential between the top and bottom surface of the
wing. Pressure is reduced by the air accelerating over the top surface of the wing. The
wing area must be big enough to generate the required lift force.
Air gets thinner as altitude increases. If the speed of the aircraft through the air is kept
constant as altitude is increased, the amount of air flowing over the wing in a given time
would decrease - and lift would decrease.
•
For a constant Lift force as altitude is increased, a constant mass flow must be
maintained. As air density decreases with altitude the speed of the wing through the air
must be increased; the true airspeed (TAS).
If you refer to the ICAO Standard Atmosphere chart on page 2 - 2, the air density at 40,000 ft
is only one quarter of the sea level value. We can use this as an example to illustrate the
relationship between the changes in T AS that are required as air density changes with altitude.
TO KEEP LIFT CONSTANT AT 40.000 ft.
TAS MUST BE DOUBLED
x4
x2
/
CONSTANT ~
L
=
CL
KEEP CONSTANT TO
MAINTAIN UD max
S-
FIXED AREA
CONSTANT
DYNAMIC PRESSURE (lAS)
For this example we will assume the optimum angle of attack of 4 0
aerodynamic efficiency and that the wing area is constant.
IS
maintained for
At 40,000 ft the air density is '/4 of the sea level value, so the speed of the aircraft through the
air must be doubled to maintain dynamic pressure (hence lift) constant. TAS is squared because
essentially we are considering the kinetic energy of the airflow (KE = '/2 m V2).
5-3
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LIFT
PRINCIPLES OF FLIGHT
The lift formula can also be used to consider the relationship between speed and angle of attack
at a constant altitude (air density).
IF SPEED IS DOUBLED,
TO
%
CL
MUST BE REDUCED
OF ITS PREVIOUS VALUE
x4
X
x2
CONSTANT~
I
L
=
CL
S - FIXED AREA
DYNAMIC PRESSURE
FOUR TIMES GREATER
(lAS) DOUBLED
CONSTANT
ALTITUDE
As speed is changed, angle of attack must be adjusted to keep lift constant.
As an example: ifIAS is doubled, TAS will double, and the square function would increase
dynamic pressure (hence lift) by a factor of four. As the aircraft is accelerated, the angle of
attack must be decreased so that the C L reduces to one quarter of its previous value to maintain
a constant lift force.
It is stated on page 2 - 4 that IAS will vary approximately as the square root of the dynamic
pressure. The proportionality between IAS and dynamic pressure is :
lAS
0<
fo
For the sake of simplicity and to promote a general understanding of this basic principle (though
no longer true when considering speeds above 0.4 M), it can be said that TAS will change in
proportion to IAS, at constant altitude, (double one, double the other etc).
The lift formula can be transposed to calculate many variables which are of interest to a
professional pilot. For example: if speed is increased in level flight by 30% from the minimum
level flight speed, we can calculate the new C L as a percentage of C L M AX :
5-4
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PRINCIPLES OF FLIGHT
LIFT
transposed becomes:
as density, lift and wing area are constant, this can be written:
CL
ex
1
V2
30% above minimum level flight speed can be written as 1.3
1
V2
now becomes
1
(1.3)2
=
1
1.69
= 0.59
100
x -1-
While maintaining level flight at a speed 30% above minimum level flight speed,
the C L would be 590/0 of C L MAX
REVIEW: Lift must balance weight in straight and level flight so at any moment in time,
weight and the lift required is constant.
a)
to maintain constant lift if density varies because of altitude change, the TAS must be
changed.
(i)
if altitude is increased, density decreases, so TAS must be increased.
(ii)
if altitude is decreased, density increases, so TAS must be decreased.
Maintaining a constant lAS will compensate for density changes.
b)
c)
to maintain constant lift if speed is changed at a constant altitude (density), the angle of
attack must be adjusted.
(i)
if speed is increased, angle of attack must be decreased, (if speed is doubled,
angle of attack must be decreased to make CL one quarter of its previous value).
(ii)
if speed is decreased, angle of attack must be increased, (if speed is halved,
angle of attack must be increased to make CL four times its previous value).
generally, a cruise speed is chosen so the aircraft operates at its optimum angle of attack
(LID max - approximately 4°).
5-5
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PRINCIPLES OF FLIGHT
5.3
LIFT
THE LIFT CURVE
Fig. 5.1 shows the lift curve of an aerofoil section, with lift coefficient (C L ) plotted against angle
of attack. It is evident that the section is symmetrical because no lift is produced at zero angle
of attack.
The lift curve is a convenient way to illustrate the properties of various configurations and will
be used extensively throughout these notes.
Lift coefficient increases with angle of attack up to a maximum (C L MAX)' which corresponds to
the 'Critical' angle of attack. Continuing to increase the angle of attack beyond this point makes
it impossible for the airflow to maintain its previous smooth flow over the contour of the upper
surface, and lift will reduce. This phenomena, stall, will be discussed in detail later.
5.3.1
INTERPRETATION OF THE LIFT CURVE
a)
To generate a constant lift force, any adjustment in dynamic pressure must be
accompanied by a change in angle of attack. (At C L less than CL MAX).
b)
F or a constant lift force, each dynamic pressure requires a specific angle of attack.
c)
Minimum dynamic pressure is determined by the maximum lift coefficient (C L MAX),
which occurs at a specific angle of attack (approximately 16 0 ) .
d)
The angle of attack for CL MAX is constant. (This is true for a given configuration).
e)
If more lift is required due to greater operating weight, a greater dynamic pressure is
required to maintain a given angle of attack.
f)
The greater the operating weight, the higher the minimum dynamic pressure.
To use the lift formula with specific values, it is necessary to convert each item to SI units.
The mass of the aircraft is 60,000 kg, to convert to a weight the mass must be multiplied by the
acceleration of gravity (9·81 mls 2). The wing area is 105 m2 • Density is the ICAO Standard
Atmosphere sea level value of 1·225 kg/m3 •
The speed resulting from the calculation will be in mls. There are 6080 ft in one nautical mile
and 3·28 ft in one metre.
The lift formula:
L
=
1P
v2
CL S
when transposed to calculate speed becomes:
5-6
V=~~P~LS
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
LIFT
l
1.532
---- -- ~=-r--i
max I
+-
0.863 -- -
-
Knots
I
-
0.552- 0.384- -
ANGLE OF ATTACK ( DEGREES)
Figure 5.1 Typical Lift Curve
Please answer the following questions: ("Answers" are provided on Page 5 - 27)
a)
How many Newtons of lift are required for straight and level flight?
b)
Calculate the airspeed in knots for each highlighted coefficient of lift.
c)
What is the lowest speed at which the aircraft can be flown in level flight?
d)
What coefficient of lift must be used to fly as slowly as possible in level flight?
e)
Does each angle of attack require a particular speed?
f)
As speed is increased what must be done to the angle of attack to maintain level flight?
g)
At higher altitude air density will be lower, what must be done to maintain the required
lift force if the angle of attack is kept constant?
h)
At a constant altitude, if speed is halved, what must be done to the angle of attack to
maintain level flight?
5-7
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PRINCIPLES OF FLIGHT
LIFT
CAMBERED
WITH 12% THICKNESS
(
CAMBER GIVES
)
\INCREASE IN C L max
SYMMETRICAL
WITH 12% THICKNESS
.~TH 6~I _
r
, (GREATER THICKNESS)
GIVES 70% INCREASE
IN C Lmax
SYMMETRICAL
j'
THICKNESS '
o~~----~~--------------------------------------------------------
o
SECTION ANGLE OF ArrACK (DEGREES)
Figure 5.2
Using the above graph, please answer the following questions:("Answers" on Page 5 - 28)
5.4
a)
Why does the cambered aerofoil section have a significantly higher C L MAX?
b)
F or the same angle of attack, why do the symmetrical aerofoil sections generate less lift
than the cambered aerofoil section?
c)
Why does the cambered aerofoil section of 12% thickness generate a small amount of
lift at slightly negative angles of attack?
d)
For a given angle of attack, the symmetrical aerofoil section of 6% thickness generates
the smallest amount of lift. In what way can this be a favourable characteristic?
e)
What are the disadvantages of the symmetrical aerofoil section of 6% thickness?
VELOCITY - DYNAMIC PRESSURE RELATIONSHIP
It is very important to understand the relationship between the velocity used in the force
equations and dynamic pressure. The velocity in the force equation is the speed of the aircraft
relative to the air through which it is moving - the True Air Speed (T AS).
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PRINCIPLES OF FLIGHT
LIFT
At a given angle of attack: "F or a constant lift force a constant dynamic pressure must be
maintained". When an aircraft is flying at an altitude where the air density is other than sea level
ISA, the TAS must be varied in proportion to the air density change. With increasing altitude;
the TAS must be increased to maintain the same dynamic pressure (Q = Yi P V2).
5.4.1
DENSITY ALTITUDE
Air density at the time of take-off and landing can significantly affect aircraft performance. If
air density is low, a longer take-off run will be needed. Air density is a product of pressure,
temperature and humidity. Humidity reduces air density because the density of water vapour is
about 5/8 that of dry air.
On an airfield at sea level with standard pressure, 1013 hPa set in the window will cause the
altimeter to read zero. This is the 'Pressure Altitude', which can be very misleading because
dynamic pressure depends on the TAS and air density, not just air pressure. If the temperature
is above standard, the density of the air will be less, perhaps a lot less, with no direct indication
of this fact visible to the pilot. If the temperature is 25 0 C it would be 10 0 C above standard
(25 - 15 = 10). The air density would be that which would exist at a higher altitude and is given
the name, 'high density altitude'.
In practical terms, this means that the aircraft will need a higher TAS for a given dynamic
pressure, hence a longer take-off run to achieve the required lAS.
To remember what 'high density altitude' means, think of it as 'HIGH density ALTITUDE'.
5.4.2
AEROFOIL SECTION LIFT CHARACTERISTICS
Fig. 5.2 shows aerofoil sections with different thickness and camber combinations producing
specific CL against a plots.
a)
An increase in the thickness ofa symmetrical aerofoil gives a higher CLMAX •
b)
The introduction of camber also has a beneficial effect on CL MAX.
The importance of maximum lift coefficient is obvious: The greater the CL MAX , the lower the
minimum flight speed (stall speed). However, thickness and camber necessary for a high CLMAX
will produce increased form drag and large twisting moments at high speed. So a high CL MAX
is just one of the requirements for an aerofoil section. The major point is that a high C LMAX wit"l
give a low minimum flight speed (lAS).
If an aero foil section of greater camber is used to give a lower minimum flight speed, the
efficient cruise speed will be lower due the generation of excessive drag. It is better to use an
aerofoil section that is efficient at high cruise speed, with the ability to temporarily increase the
camber of the wing when it is necessary to fly slowly. This can be achieved by the use of
adjustable hinged sections of the wing leading and trailing edges (Flaps).
5-9
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PRINCIPLES OF FLIGHT
5.5
LIFT
INTRODUCTION TO DRAG CHARACTERISTICS
Drag is the aerodynamic force parallel to the relative airflow and opposite in direction to the
flight path. (Drag, as a complete subject, will be discussed in detail later). As with other
aerodynamic forces, drag forces may be expressed in the form of a coefficient which is
independent of dynamic pressure and surface area.
D = Q
Co S
Drag is the product of dynamic pressure, drag coefficient and surface area. Co is the ratio of drag
per unit wing area to dynamic pressure. If the CD of a representative wing were plotted against
angle of attack, the result typically would be a graph similar to that shown in Fig. 5.3. At low
angles of attack CD is low and small changes in angle of attack create only small changes in CD.
But at higher angles of attack, the rate of change in CD per degree of angle of attack increases;
CD change with angle of attack is exponential. Beyond the stalling angle of attack (C L MAX ), a
further large increase in CD takes place.
5.5.1
LIFTIDRAG RATIO
An appreciation of the efficiency of lift production is gained from studying the ratio between lift
and drag; a high LID ratio being more efficient.
The proportions of CL and CD can be calculated for each angle of attack. Fig. 5.4 shows that the
LID ratio increases with angle of attack up to a maximum at about 4 0 ; this is called the
'optimum' angle of attack. The LID ratio then decreases with increasing angle of attack until
C L MAX is reached.
NB:
The plot of lift, the plot of drag and the plot of LID ratio shown in Fig. 5.4 are all at different
scales and no conclusions should be drawn from the intersection of plots.
The maximum liftldrag ratio (LID MAX ) of a given aerofoil section will occur at one specific
angle of attack. If the aircraft is operated in steady level flight at the optimum angle of attack,
drag will be least while generating the required lift force. Any angle of attack lower or higher
than that for LID MAX reduces the LID ratio and consequently increases drag for the required lift.
Assume the LID MAX of Fig. 5.4. is 12·5. In steady flight at a weight of588,600N and lAS to give
the required lift at 4 0 angle of attack, the drag would be 47,088N (588,600 -7- 12·5). Any higher
or lower speed would require a different angle of attack to generate the required lift force. Any
angle of attack other than 4 0 will generate more drag than 47,088 N. Of course, this same
'aircraft' could be operated at a different weight and the same LID MAX of 12·5 could be obtained
at the same angle of attack. But a change in weight requires a change in lAS to support the new
weight at the same angle of attack. The lower the weight, the lower lAS required to stay at the
LID MAX angle of attack, and vice versa.
For a given configuration (Flaps, gear, spoilers and airframe contamination) and at speeds less
than MO·4, changes in weight will not change LID MAX.
5 - 10
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LIFT
PRINCIPLES OF FLIGHT
L -_ _ _ __ __ __ __ __ __ _ _ _ _ _ _ _ _ _ _ __ _
ANGLE OF ATTACK (DEGREES)
Figure 5.3
-oL
CD
I
~
: 4·
\
: 16·
OPTIMUM
OF
ATTACK
~ ANGLE
ANGLE OF ATTACK (DEGREES)
Figure 5.4
5 - 11
© Oxford Aviation Services Limited
LIFT
PRINCIPLES OF FLIGHT
The design of an aircraft has a great effect on the LID ratio. Typical values are listed below for
various types.
Aircraft Type
LID MAX
High performance sailplane
Modern jet transport
Propeller powered trainer
5.6
from 25 to 60
from 12 to 20
from 10 to 15
EFFECT OF AIRCRAFT WEIGHT ON MINIMUM FLIGHT SPEED
A given aero foil section will always stall at the same angle of attack, but aircraft weight will
influence the lAS at which this occurs. Modern large jet transport aircraft may have just over
half their maximum gross take-off weight made-up offuel. So stall speed can vary considerably
throughout the flight.
5.7
CONDITION OF THE SURFACE
Surface irregularities, especially near the leading edge, have a considerable effect on the
characteristics of aerofoil sections. CLMAX in particular, is sensitive to leading edge roughness.
Fig. 5.5 illustrates the effect of a rough leading edge compared to a smooth surface. In general,
C LMAX decreases progressively with increasing roughness ofthe leading edge. Roughness further
downstream than about 20 percent of the chord from the leading edge has little effect on C L MAX
or the lift-curve slope. Frost, snow and even rainwater can significantly increase surface
roughness. Dirt or slush picked-up from contaminated parking areas, taxiways and runways can
also have a serious affect. In-flight icing usually accumulates at the leading edge of aerofoils
and will severely increase surface roughness causing a significant decrease in C L MAX'
5.8
FLIGHT AT HIGH LIFT CONDITIONS
The aerodynamic lift characteristics of an aircraft are shown by the curve of lift coefficient
versus angle of attack in Fig. 5.6, for a specific aircraft in the clean and flap down
configurations. A given aerodynamic configuration experiences increases in lift coefficient with
increases in angle of attack until the maximum lift coefficient is obtained. A further increase in
angle of attack produces stall and the lift coefficient then decreases.
Effect of High Lift Devices: The primary purpose of high lift devices (flaps, slots, slats, etc)
is to reduce take-off and landing distance by increasing the C L MAX of the aerofoil section and so
reduce the minimum speed. The effect of a "typical" high lift device is shown by the lift curves
of Fig. 5.6. The principal effect of the extension of flaps is to increase C LMAX and reduce the
angle of attack for any given lift coefficient. The increase in CL MAX afforded by flap deflection
reduces the stall speed in a certain proportion. (High lift devices will be fully covered later).
5 - 12
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PRINCIPLES OF FLIGHT
LIFT
C L max -------l~..__ - -Take-off
CL_
_ _
,
,-.,- .........
\
IZ
Basic Smooth Wing
w
' ...
1-(3
....
_----- ...
-
Ll.._Ll..
....JLl..
-
-
I
Wing with Frost,
Dirt, Water or Slush
W
ou
Wing with Ice
ANGLE OF ATTACK
Figure 5.5
IZ
W
(3
u::
Ll..
W
o
U
t;::
::J
•
~CLEAN
CONFIGURATION
ANGLE OF ATTACK
Figure 5.6
5 - 13
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
-
1~
1·
c
LIFT
b
·1
S = WING AREA, sq. m (b x c)
(~____________________~)
b = SPAN, m
T
c = AVERAGE CHORD, m
b
::::>
~I
AR = ASPECT RATIO
AR = b/c
AR = b 2/S
CR
- -- - - - b - - - - --1
= ROOT
CHORD, m
Cl = TIP CHORD, m
CliCR
=
I----~-------- SWEEP
TAPER RATIO
ANGLE, degrees
Cl
MAC
---------.,
= MEAN
AERODYNAMIC CHORD , m
I
I
MAC
I
I
I
I
Figure 5.7 Wing Terminology
5 - 14
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LIFT
PRINCIPLES OF FLIGHT
5.9
THREE DIMENSIONAL AIRFLOW
So far we have considered only two dimensional airflow. This has been a foundation for an
appreciation of the actual pattern of airflow over an aircraft. Even minute pressure differences
will modify airflow direction by inducing air to flow towards any region of lower pressure.
Three dimensional airflow modifies the effective angle of attack, increases drag, alters stalling
characteristics and can influence the control and stability of the aircraft. From now on, instead
of just an aerofoil section, the entire wing will be considered.
5.10
WING TERMINOLOGY
Wing Area (S): The plan surface area of the wing. Although a portion of the area may be
covered by fuselage or engine nacelles, the pressure carryover on these surfaces allows
legitimate consideration of the entire plan area.
Wing Span (b): The distance from tip to tip.
Average Chord (c): The geometric average. The product of the span and the average chord
is the wing area (b x c = S).
Aspect Ratio (AR): The proportion of the span and the average chord (AR = b/c). If the
planform has curvature and the average chord is not easily determined, an alternative expression
is (b 2/S). The aspect ratio of the wing determines the aerodynamic characteristics and structural
weight. Typical aspect ratios vary from 35 for a high performance sailplane to 3 for a j et fighter.
The aspect ratio of a modem high speed jet transport is about 12.
Root Chord (C.J: The chord length at the wing centreline.
Tip Chord (C T): The chord length at the wing tip
Taper Ratio (C T/C.J: The ratio of the tip chord to the root chord. The taper ratio affects the
lift distribution and the structural weight of the wing. A rectangular wing has a taper ratio of 1.0
while the pointed tip delta wing has a taper ratio of 0.0
Sweep Angle: Usually measured as the angle between the line of 25% chords and a
perpendicular to the root chord. The sweep of a wing causes definite changes in compressibility,
maximum lift, and stall characteristics.
Mean Aerodynamic Chord (MAC): The chord drawn through the geographic centre of the
plan area. A rectangular wing of this chord and the same span would have broadly similar
pitching moment characteristics. The MAC is located on the reference axis of the aircraft and
is a primary reference for longitudinal stability considerations.
5 - 15
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PRINCIPLES OF FLIGHT
LIFT
5.11
WING TIP VORTICES
Air flowing over the top surface of a wing is
at a lower pressure than that beneath. The
trailing edge and the wing tips are where the
airflows interact. The pressure differential
modifies the directions of flow, inducing a
span-wise vector towards the root on the
upper surface and generally, towards the tip
on the lower surface, Fig. 5 .8.
"Conventionally", an aircraft is viewed from
the rear. An anti-clockwise vortex will be
induced at the right wing-tip and a clock-wise
vortex at the left wing-tip, Figs. 5.9, 5.10 &
5.11.
Figure 5.8
UPPER SURFACE
(LCMler Pressure)
$Q~":' @'~
At higher angles of attack (Lower lAS) the
decreased chordwise vector will increase the
effect of the resultant spanwise flow, making
the vortices stronger.
Figure 5.9
Induced Downwash: (Fig. 5.12) Trailing
vortices create certain vertical velocity
components in the airflow in the vicinity of
the wing, both in front of and behind it. These
vertical velocities cause a downwash over the
wing resulting in a reduction in the effective
angle of attack. The stronger the vortices, the
greater the reduction in effective angle of
attack. Because of this local reduction in
effective angle of attack, the overall lift
generated by a wing will be below the value
that would be generated if there were no
spanwise pressure differential. It is the
production of lift itself which reduces the
, magnitude of the lift force being generated.
To replace the lift lost by the increased
downwash, the aircraft must be flown at a
higher angle of attack. This increases drag.
This extra drag is called induced drag. The
stronger the vortices, the greater the induced
drag.
Figure 5.10
Figure 5.11
5 - 16
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
LIFT
1'.;
UpNash Increased
_ilill ,.
-~--
Vertical Velocities
in the vicinity of
the wing are a function of
tip vortex strength
DoNwash Increased
Angular deflection of effective airflcw
is a function of both vortex strength
v
Relative Airflow
Lift Inclined Rearwards because of
I~-- Decreased Effective Angle of Attack
'-...
Effective
Airflcw
____
(
'-...
~'
I
<J(.e
<J(.
I
\ d;
.~--~
i
I
0 _ '_
Relative Airflow
<J(. e
<J(.
Figure 5.12
i
= effective angle of attac~
= induced angle of attack
J
Wing tip vortices, in particular their influence on upwash and downwash, have a significant
effect on several important areas of aircraft aerodynamics, stability and control. Some ofthese
effects will be examined now and throughout the remaining chapters.
5 - 17
© Oxford Aviation Services Limited
LIFT
PRINCIPLES OF FLIGHT
5.12
WAKE TURBULENCE: (Ref: AIC 17/1999)
Trailing wingtip vortices extend behind aircraft for a considerable distance and can present an
extreme hazard to any aircraft unfortunate enough to encounter them. Maximum tangential
airspeed in the vortex system may be as high as 90 m/s (300 ft/sec) immediately behind a large
aircraft. Wake turbulence cannot be detected, so it is important for pilots to be aware of the
potential distribution and duration of trailing vortices, plus modifications made to the "classic"
vortex system by surface wind speed and direction.
Aircraft Wake Vortex Characteristics: Wake vortex generation begins when the nosewheel
lifts off the runway on take-off and continues until the nosewheel touches down on landing.
Wake vortices exist behind every aircraft, including helicopters, when in flight, but are most
severe when generated by heavy aircraft. They present the greatest danger during the take-off,
initial climb, final approach and landing phases of flight - in other words, at low altitude where
large numbers of aircraft congregate. A wake turbulence encounter is a hazard due to potential
loss of control and possible structural damage, and if the experience takes place near the ground
there may be insufficient time and/or altitude to recover from an upset.
Touchdown
(Wake ends)
I
~
.
~J
~"'" ..
Figure 5.13
The characteristics oftrailing vortices are determined by the "generating" aircraft's:
(a)
Gross weight - the higher the weight, the stronger the vortices.
(b)
Wingspan - has an influence upon the proximity of the two trailing vortices.
(c)
Airspeed - the lower the speed, the stronger the vortices.
(d)
Configuration - vortex strength is greatest with aircraft in a "clean" configuration (for
a given speed and weight).
(e)
Attitude - the higher the angle of attack, the stronger the vortices.
5 - 18
© Oxford Aviation Services Limited
LIFT
PRINCIPLES OF FLIGHT
As a general rule, the larger the "generating" aircraft relative to the aircraft encountering the
wake turbulence, the greater the hazard. There is also evidence that for a given weight and speed
a helicopter produces a stronger vortex than a fixed-wing aircraft.
Distribution of Trailing Vortices: Typically the two trailing vortices remain separated by
about three quarters of the aircraft's wingspan and in still air they tend to drift slowly
downwards and level off, usually between 500 and 1000 ft below the flight path of the aircraft.
Behind a large aircraft the trailing vortices can extend as much as nine nautical miles.
Figure 5.14
Figure 5.15
~~ ____ ___
500 to 1000 ft
__________ J__
.
~"J'."J"J'ZJ<J\J\j\J\j \j"J":.f"..~
1---
1......
- --
-
---
Approx. 9 nautical miles
behind a large aircraft
~
-r--
--------~
Figure 5.16
5 - 19
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
LIFT
Vortex Movement near the Ground: Fig. 5.17 shows that if the generating aircraft is within
1000 ft of the ground, the two vortices will "touch-down" and move outwards at about 5 kts from
the track of the generating aircraft at a height approximately equal to Y2 the aircraft's wingspan.
A
I
I
I
1000 ft
880 5k~
~ Drift
I
I
I
y
STILL AIR - (viewed from the rear)
Figure 5.17
In a crosswind, if the surface wind is light and steady, the wake vortex system "in contact" with
the ground will drift with the wind. Fig. 5.18 shows the possible effect of a crosswind on the
motion of a vortex close to the ground. With parallel runways, wake turbulence from an aircraft
operating on one runway can be a potential hazard to aircraft operating from the other.
10 kts Drift
e-
~~~-------
(5 kts + 5 kts)
~---~--~
5 kt CROSSWIND - (Viewed from the rear)
Figure 5.18
5 - 20
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
LIFT
The Decay Process of Trailing vortices: Atmospheric turbulence has the greatest influence
on the decay of wake vortices; the stronger the wind, the quicker the decay.
Probability of Wake Turbulence Encounter: Certain separation minima are applied by Air
Traffic Control (ATC), but this does not guarantee avoidance. ATC applied separation merely
reduces the probability of an encounter to a lower level, and may minimise the magnitude of the
upset if an encounter does occur. Particular care should be exercised when following any
substantially heavier aircraft, especially in conditions of light wind. The majority of serious
incidents, close to the ground, occur when winds are light.
Wake Turbulence Avoidance: If the location of wake vortices behind a preceding or crossing
aircraft are visualised, appropriate flight path control will minimise the probability of a wake
turbulence encounter. Staying above and/or upwind of a preceding or crossing aircraft will
usually keep your aircraft out of the generating aircraft's wake vortex. Unfortunately, deviating
from published approach and departure requirements in order to stay above/upwind of the flight
path of a preceding aircraft may not be advisable. Maintaining proper separation remains the
best advice for avoiding a wake turbulence encounter.
5 - 21
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
5.13
LIFT
GROUND EFFECT
When landing and taking off, the closeness of the wing to the ground prevents full development
of the trailing vortices, Fig. 5.19, making them much weaker. Upwash and downwash are
reduced, causing the effective angle of attack of the wing to increase, (ref: Fig. 5.12). Therefore,
when an aircraft is "in ground effect" lift will generally be increased and induced drag (C Oi )
will be decreased. In addition, the reduced downwash will effect both longitudinal stability
because of CP movement, and the pitching moment because of changes to the effective angle of
attack of the tailplane, (Ref: Fig. 5.21).
Upwash
Aeroplane out of
Ground Effect
Aeroplane in
Ground Effect
height
h
Reduced
Upwash
~
~
Tip Vortices reduced by contact with Ground
Figure 5.19
The Impact of Ground Effect: The influence of ground effect depends on the distance of the
wing above the ground. A large reduction in C Oi will take place only when the wing is very
close to the ground, (within halfthe wingspan). '
For a representative aircraft with a 40m span, (Ref. Fig. 5.20):
(a)
At a height of 40m, the reduction in C Oi is only 1.4%.
(b)
At a height of 10m, the reduction in C Oi is 23.5%, but
(c)
At a height of 4m, the reduction in C Oi is 47.6%
5 - 22
© Oxford Aviation Services Limited
LIFT
PRINCIPLES OF FLIGHT
60
Percent
Reduction
in
Induced
Drag
Coefficient
50
CL
40
Constant
30
20
10 +--+--+--+--+-~~+
O +-~~~-t~-J--~C=~~~~~
o
0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0 1.1
Ratio of wing height to span (hlb)
Figure 5.20
The height of the wing above the ground when the aircraft is in the landing attitude is influenced
by its mounting position on the fuselage. From the graph in Fig. 5.20 it can be seen that the last
few metres makes a big difference to the reduction of C Oi ' In general, it can be said that a low
wing aircraft will experience a greater degree of ground effect than an aircraft with a high
mounted wing.
High and Low Tail Characteristics: While ground effect may possibly change the
aerodynamic characteristics ofthe tailplane in its own right, a low mounted tailplane will have
its effective angle of attack modified by the changing downwash angle behind the wing. A high
mounted tailplane may be outside the influence of the changing downwash angle and not suffer
the same disadvantages.
Downwash Decreased
by Ground Effect
----------~
----------------------=
-~
- -~
•
•
Decreased
Dcwmload
on Tailplane
Figure 5.21
5 - 23
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
LIFT
REDUCED
DOW NWASH
ANGLE
" NORMAL"
DOW NWASH
INCREASED
UP - FORCE
POSITIVE
CAMBER
Tailplane (Every illustration)
NEGATIVE
CAMBER
-
DECREASED
DOWN - FORCE
SYMMETRICAL
DECREASED
DOWN - FORCE
Figure 5.22 Influence of Camber on Effect of downwash on pitching moment
Influence of Tailplane Camber on Pitching Moment: It can be seen from Fig. 5.22 that the
type of tailplane camber does not influence the pitching moment generated when downwash
from the wing changes. Decreased downwash will always result in an aircraft nose down
pitching moment. The opposite will be true of increased downwash.
Downwash will change not only because of ground effect, but also when flaps are operated and
when a shockwave forms on the wing at speeds higher than M eRIT> so appreciation of this
phenomena is a key element towards a full understanding of Principles of Flight.
5 - 24
© Oxford Aviation Services Limited
LIFT
PRINCIPLES OF FLIGHT
D
DECREASING
ANGLE OF ATTACK
B
A
TAILPLANE CHORD
-.~~----
LINE
E
G
INCREASING
ANGLE OF ATTACK
Figure 5.23 Influence of Downwash on Tailplane Angle of Attack
Tailplane Angle of Attack: Angle of Attack is the angle between the chord line and the
Relative Airflow. The Relative Airflow has three characteristics:
1.
Magnitude - the speed ofthe aircraft through the air; the True Air Speed (TAS)
2.
Direction - parallel to and in the opposite direction to the aircraft flight path, and
3.
Condition - unaffected by the presence of the aircraft.
Air flowing offthe wing trailing edge (downwash) cannot be defined as relative airflow because
it does not conform to the definitions. Neither is it possible to think strictly of a tailplane angle
of attack. Airflow which has been influenced by the presence of the aircraft (direction of flow
and dynamic pressure) must be thought of as Effective Airflow. And the angle between the
chord line and the effective airflow must be thought of as Effective Angle of Attack.
Consider Fig. 5.23 . Airflow from direction (A) gives the tailplane zero (effective) angle of
attack. Airflow from direction (E, F or G) would be an increase in (effective) angle of attack.
If airflow from direction (G) is now considered, flow from (F, E, A, B, C or D) would be a
decrease in (effective) angle of attack. The term "negative angle of attack" is not used .
CONCLUSION
Increasing downwash (G to D) gives a decrease in tailplane (effective) angle of attack and
decreasing downwash (D to G) gives an increase in tailplane (effective) angle of attack.
It is necessary to understand the effect of changing down wash on tailplane angle of attack, but
it is vital to understand the influence of downwash on aircraft pitching moment.
5 - 25
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
LIFT
Entering Ground Effect: Consider an aircraft entering ground effect, assuming that a constant
CL and lAS is maintained. As the aircraft descends into ground effect the following changes will
take place:
(a)
The decreased downwash will give an increase in the effective angle of attack, requiring
a smaller wing angle of attack to produce the same lift coefficient. If a constant pitch
attitude is maintained as ground effect is encountered, a "floating" sensation may be
experienced due to the increase in CL and the decrease in CD i (thrust requirement),
Fig's. 5.12 & 5.24. The decrease of induced drag will cause a reduction in deceleration,
and any excess speed may lead to a considerable "float" distance. The reduction in
thrust required might also give the aircraft a tendency to climb above the desired glide
path, "balloon", if a reduced throttle setting is not used.
(b)
If airspeed is allowed to decay significantly during short finals and the resulting sinkrate arrested by increasing the angle of attack, upon entering ground effect the wing
could stall, resulting in a heavy landing.
(c)
The pilot may need to increase pitch input (more elevator back-pressure) to maintain the
desired landing attitude. This is due to the decreased downwash increasing the effective
angle of attack of the tailplane, Fig. 5.21. The download on the tail is reduced,
producing a nose down pitching moment.
(d)
Due to the changes in the flowfield around the aircraft there will be a change in position
error which may cause the ASI to misread. In the majority of cases, local pressure at the
static port will increase and cause the ASI and altimeter to under read.
/
Aircraft in
Ground Effect
\
~
/// /~ \
/
/
/
1ii
2
.s:::
I-
Aircraft out of
Ground Effect
Ai rc raft out of
Ground Effect
/
/
\/
/
~ Aircraft
/
in
Ground Effect
/
V
lAS
Angle of Attack
Figure 5.24
5 - 26
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
LIFT
Leaving Ground Effect: The effects of climbing out of ground effect will generally be the
opposite to those of entering. Consider an aircraft climbing out of ground effect while
maintaining a constant C L and lAS. As the aircraft climbs out of ground effect the following
changes will take place:
(a)
The CL will reduce and the CD i (thrust requirement) will increase. The aircraft will
require an increase in angle of attack to maintain the same CL •
(b)
The increase in downwash will generally produce a nose up pitching moment. The pitch
input from the pilot may need to be reduced (less elevator back-pressure).
(c)
Position error changes may cause the ASI to misread. In the majority of cases, local
pressure at the static port will decrease and cause the ASI and altimeter to over read.
(d)
It is possible to become airborne in ground effect at an airspeed and angle of attack
which would, after leaving ground effect, cause the aircraft to settle-back on to the
runway It is therefore vitally important that correct speeds are used for take-off.
(e)
The nose up pitching moment may induce an inadvertent over rotation and tail strike.
5 - 27
© Oxford Aviation Services Limited
LIFT
PRINCIPLES OF FLIGHT
5.14
SUMMARY
Three major factors influence production of the required lift force:
(a)
Dynamic Pressure (lAS)
(b)
Pressure Distribution (Section profile & Angle of attack)
(c)
Wing Area (S)
To provide a constant lift force, each lAS corresponds to a particular angle of attack.
The angle of attack at CL MAX is constant.
A higher aircraft weight requires a lift force to balance it; an increased lAS is needed to provide
the greater lift at the same angle of attack.
As altitude increases a constant lAS will supply the same lift force at a given angle of attack.
A thinner wing will generate less lift at a given angle of attack, and have a higher minimum
speed.
A thinner wing can fly faster before shock wave formation increases drag.
A thinner wing requires high lift devices to have an acceptably low minimum speed.
The Lift/Drag ratio is a measure of aerodynamic efficiency.
Contamination of the wing surface, particularly the front 20% of the chord, will seriously
decrease aerodynamic performance.
Wing tip vortices:
(a)
Decrease overall lift production.
(b)
Increase drag.
(c)
Modify the downwash which changes the effective angle of attack of the tailplane.
(d)
Generate trailing vortices which pose a serious hazard to aircraft that encounter them.
(e)
Affect the stall characteristics of the wing
(f)
Change the lift distribution.
The sudden full effects of vortices or their absence must be anticipated during take-off and
landing.
5 - 28
© Oxford Aviation Services Limited
LIFT
PRINCIPLES OF FLIGHT
ANSWERS FROM PAGE 5 - 7
CL
1.532
- - - ~J-
~
-.[!TA ~
0.863- -
-
-
-
-
0.552- -
ANGLE OF ATTACK ( DEGREES)
a)
How many Newtons of lift are required for straight and level flight? 588,600 N.
b)
Calculate the airspeed in knots for each highlighted coefficient of lift. As above.
c)
What is the lowest speed at which the aircraft can be flown in level flight? 150 kt.
d)
What coefficient of lift must be used to fly as slowly as possible in level flight? C LMAX
e)
Does each angle of attack require a particular speed? Yes.
f)
As speed is increased what must be done to the angle of attack to maintain level flight?
Angle of attack must be decreased .
g)
At higher altitude air density will be lower, what must be done to maintain the required
lift force? Increase the True Airspeed (TAS) .
h)
At a constant altitude, if speed is halved, what must be done to the angle of attack to
maintain level flight? Increased so that C L is four times greater.
5 - 29
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
LIFT
ANSWERS FROM PAGE 5 - 8
CAMBERED
WITH 12% THICKNESS
CAMBER GIVES )
( INCREASE INC Lmax
GREATER THICKNESS~
GIVES 70% INCREASE
(
IN C Lmax
o~~----~~--------------------------------------------------------
o
SECTION ANGLE OF ATTACK (DEGREES)
(a)
Why does the cambered aerofoil section have a significantly higher CL MAX?
When compared to a symmetrical section ofthe same thickness: at approximately
the same stall angle, the cross sectional area of the "streamtube" over the top
surface is smaller with a more gradual section change. This allows greater
acceleration of the air over the top surface, and a bigger pressure differential.
(b)
For the same angle of attack, why do the symmetrical aerofoil sections generate less lift
than the cambered aerofoil section? Angle of attack is the angle between the chord
line and the relative airflow. At the same angle of attack, the cross sectional area
of the symmetrical section upper surface "streamtube" is larger.
(c)
Why does the cambered aerofoil section of 12% thickness generate a small amount of
lift at slightly negative angles of attl}ck? At small negative angles of attack, a
cambered aerofoil is still providing a reduced cross sectional area "streamtube"
over the top surface, generating a small pressure differential. (Ref. Page 4 - 6 & 7).
(d)
For a given angle of attack, the symmetrical aerofoil section of6% thickness generates
the smallest atnount oflift. In what way can this be a favourable characteristic?
At the high speeds at which modern high speed jet transport aircraft operate, a
thin wing can generate the required lift force with minimum drag caused by the
formation of shock waves. (This will be fully explained in later chapters).
(e)
What are the disadvantages of the symmetrical aerofoil section of 6% thickness?
It will give a high minimum speed, requiring complex high lift devices to enable the
aircraft to use existing runways.
5 - 30
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PRINCIPLES OF FLIGHT
LIFT
SELF ASSESSMENT QUESTIONS
1.
To maintain altitude, what must be done as Indicated Air Speed (lAS) is reduced:
a)
b)
c)
d)
2.
If more lift force is required because of greater operating weight, what must be done to fly at the
angle of attack which corresponds to CL MAX:
a)
b)
c)
d)
3.
Decrease angle of attack to reduce the drag.
Increase angle of attack to maintain the correct lift force.
Deploy the speed brakes to increase drag.
Reduce thrust.
Increase the angle of attack.
Nothing, the angle of attack for CL MAX is constant.
It is impossible to fly at the angle of attack that corresponds to CL MAX.
Increase the Indicated Air Speed (lAS).
Which of the following statements is correct:
1 - To generate a constant lift force, any adjustment in lAS must be accompanied by a change
in angle of attack.
2 - For a constant lift force, each lAS requires a specific angle of attack.
3 - Minimum lAS is determined by CL MAX.
4 - The greater the operating weight, the higher the minimum lAS.
a)
b)
c)
d)
4.
What effect does landing at high altitude airports have on ground speed with comparable
conditions relative to temperature, wind, and aeroplane weight:
a)
b)
c)
d)
5.
1,2 and 4
4 only
2,3 and 4
1,2,3 and 4
Higher than at low altitude.
The same as at low altitude.
Lower than at low altitude.
Dynamic pressure will be the same at any altitude.
What flight condition should be expected when an aircraft leaves ground effect:
a)
b)
c)
d)
A decrease in parasite drag permitting a lower angle of attack.
An increase in induced drag and a requirement for a higher angle of attack.
An increase in dynamic stability.
A decrease in induced drag requiring a smaller angle of attack.
5 - 31
© Oxford Aviation Services Limited
LIFT
PRINCIPLES OF FLIGHT
6.
What will be the ratio between airspeed and lift if the angle of attack and other factors remain
constant and airspeed is doubled. Lift will be:
a)
b)
c)
d)
7.
What true airspeed and angle of attack should be used to generate the same amount of lift as
altitude is increased:
a)
b)
c)
d)
8.
Lift and airspeed, but not drag.
Lift, gross weight, and drag.
Lift, airspeed, and drag.
Lift and drag, but not airspeed.
Which flight conditions of a large jet aeroplane create the most severe flight hazard by
generating wingtip vortices of the greatest strength:
a)
b)
c)
d)
11.
A lower angle of attack.
A higher angle of attack.
The same angle of attack.
The same angle of attack, but a lower lAS.
By changing the angle of attack of a wing, the pilot can control the aeroplane's:
a)
b)
c)
d)
10.
A higher true airspeed for any given angle of attack.
The same true airspeed and angle of attack.
A lower true airspeed and higher angle of attack.
A constant angle of attack and true air speed.
How can an aeroplane produce the same lift in ground effect as when out of ground effect:
a)
b)
c)
d)
9.
Two times greater.
Four times greater.
The same.
One quarter.
Heavy, slow, gear and flaps up.
Heavy, fast, gear and flaps down.
Heavy, slow, gear and flaps down.
Weight, gear and flaps make no difference.
Hazardous vortex turbulence that might be encountered behind large aircraft is created only
when that aircraft is:
a)
b)
c)
d)
U sing high power settings.
Operating at high airspeeds.
Developing lift.
Operating at high altitude.
5 - 32
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
12.
Wingtip vortices created by large aircraft tend to:
a)
b)
c)
d)
13.
b)
c)
d)
Wake turbulence behind a propeller-driven aircraft is negligible because jet engine
thrust is a necessary factor in the formation of vortices.
Vortices can be avoided by flying 300 feet below and behind the flightpath of the
generating aircraft.
The vortex characteristics of any given aircraft may be altered by extending the flaps or
changing the speed.
Vortices can be avoided by flying downwind of, and below the flight path of the
generating aircraft.
What effect would a light crosswind have on the wingtip vortices generated by a large aeroplane
that has just taken off:
a)
b)
c)
d)
16.
Inward, upward, and around the wingtip.
Counterclockwise.
Outward, upward, and around the wingtip.
Outward, downward and around the wingtip.
Which statement is true concerning the wake turbulence produced by a large transport aircraft:
a)
15.
Rise from the surface to traffic pattern altitude.
Sink below the aircraft generating the turbulence.
Accumulate and remain for a period of time at the point where the takeoff roll began.
Dissipate very slowly when the surface wind is strong.
How does the wake turbulence vortex circulate around each wingtip, when viewed from the rear:
a)
b)
c)
d)
14.
LIFT
The downwind vortex will tend to remain on the runway longer than the upwind vortex.
A crosswind will rapidly dissipate the strength of both vortices.
A crosswind will move both vortices clear of the runway.
The upwind vortex will tend to remain on the runway longer than the downwind vortex.
To avoid the wingtip vortices ofa departing jet aeroplane during takeoff, the pilot should:
a)
b)
c)
d)
Remain below the flightpath of the jet aeroplane.
Climb above and stay upwind of the jet aeroplane's flightpath.
Lift off at a point well past the jet aeroplane's flightpath.
Remain below and downwind of the jet aeroplane's flightpath.
5 - 33
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LIFT
PRINCIPLES OF FLIGHT
17.
What wind condition prolongs the hazards of wake turbulence on a landing runway for the
longest period of time:
a)
b)
c)
d)
18.
If you take off behind a heavy jet that has just landed, you should plan to lift off:
a)
b)
c)
d)
19.
d)
Differential pressure acting perpendicular to the chord of the wing.
Force acting perpendicular to the relative wind.
Reduced pressure resulting from a laminar flow over the upper camber of an aerofoil,
which acts perpendicular to the mean camber.
Force acting parallel with the relative wind and in the opposite direction.
Which statement is true relative to changing angle of attack:
a)
b)
c)
d)
22.
Increased thrust.
A decreased stall speed.
An increased stall speed.
An aircraft will always stall at the same indicated airspeed.
Lift on a wing is most properly defined as the:
a)
b)
c)
21.
Prior to the point where the jet touched down.
At the point where the jet touched down and on the upwind edge of the runway.
Before the point where the jet touched down and on the downwind edge of the runway.
Beyond the point where the jet touched down.
The adverse effects of ice, snow, or frost on aircraft performance and flight characteristics
include decreased lift and:
a)
b)
c)
d)
20.
Light quartering headwind.
Light quartering tailwind.
Direct tailwind.
Strong, direct crosswind.
A decrease in angle of attack will increase pressure below the wing, and decrease drag.
An increase in angle of attack will decrease pressure below the wing, and increase drag.
An increase in angle of attack will increase drag.
An increase in angle of attack will decrease the lift coefficient.
The angle of attack of a wing directly controls the:
a)
b)
c)
d)
Angle of incidence of the wing.
Distribution of pressures acting on the wing.
Amount of airflow above and below the wing.
Dynamic pressure acting in the airflow.
5 - 34
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
23.
In theory, if the angle of attack and other factors remain constant and the airspeed is doubled,
the lift produced at the higher speed will be:
a)
b)
c)
d)
24.
d)
Thrust is greater than drag and weight and lift are equal.
These forces are equal.
Thrust is greater than drag and lift is greater than weight.
Thrust is slightly greater than Lift, but the drag and weight are equal.
At higher elevation airports the pilot should know that indicated airspeed:
a)
b)
c)
d)
28.
Camber line.
Longitudinal axis.
Chord line.
Flightpath.
Which statement is true, regarding the opposing forces acting on an aeroplane in steady-state
level flight:
a)
b)
c)
d)
27.
Negative air pressure below and a vacuum above the wing's surface.
Vacuum below the wing's surface and greater air pressure above the wing's surface.
Higher air pressure below the wing's surface and lower air pressure above the wing's
surface.
Higher pressure at the leading edge than at the trailing edge.
On a wing, the force of lift acts perpendicular to, and the force of drag acts parallel to the:
a)
b)
c)
d)
26.
The same as at the lower speed.
Two times greater than at the lower speed.
Four times greater than at the lower speed.
One quarter as much.
An aircraft wing is designed to produce lift resulting from a difference in the:
a)
b)
c)
25.
LIFT
Will be unchanged, but ground speed will be faster.
Will be higher, but ground speed will be unchanged.
Should be increased to compensate for the thinner air.
Should be higher to obtain a higher landing speed.
An aeroplane leaving ground effect will:
a)
b)
c)
d)
Experience a reduction in ground friction and require a slight power reduction.
Require a lower angle of attack to maintain the same lift coefficient.
Experience a reduction in induced drag and require a smaller angle of attack
Experience an increase in induced drag and require more thrust.
5 - 35
© Oxford Aviation Services Limited
LIFT
PRINCIPLES OF FLIGHT
29.
If the same angle of attack is maintained in ground effect as when out of ground effect, lift will:
a)
b)
c)
d)
30.
Which is true regarding the force of lift in steady, unaccelerated flight:
a)
b)
c)
d)
31.
There is a corresponding indicated airspeed required for every angle of attack to
generate sufficient lift to maintain altitude.
An aero foil will always stall at the same indicated airspeed; therefore, an increase in
weight will require an increase in speed to generate sufficient lift to maintain altitude.
At lower airspeeds the angle of attack must be less to generate sufficient lift to maintain
altitude.
The lift force must be exactly equal to the drag force.
At a given Indicated Air Speed, what effect will an increase in air density have on lift and drag:
a)
b)
c)
d)
32.
Increase, and induced drag will increase.
Increase, and induced drag will decrease.
Decrease, and induced drag will increase.
Decrease and induced drag will decrease.
Lift will increase but drag will decrease.
Lift and drag will increase.
Lift and drag will decrease.
Lift and drag will remain the same.
If the angle of attack is increased beyond the critical angle of attack, the wing will no longer
produce sufficient lift to support the weight of the aircraft:
a)
b)
c)
d)
Unless the airspeed is greater than the normal stall speed.
Regardless of airspeed or pitch attitude.
Unless the pitch attitude is on or below the natural horizon.
In which case, the control column should be pulled-back immediately.
5 - 36
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
33.
LIFT
GIVEN THAT:
Aircraft A.
Wingspan:
Average wing chord:
51 m
4m
Aircraft B.
Wingspan:
Average wing chord:
48 III
3.5 m
Determine the correct aspect ratio and wing area.
a)
b)
c)
d)
34.
Aspect ratio of the wing is defined as the ratio of the:
a)
b)
c)
d)
35.
Wingspan to the wing root.
Square of the chord to the wing span.
Wing span to the average chord.
Square of the wing area to the span.
What changes to aircraft control must be made to maintain altitude while the airspeed is being
decreased:
a)
b)
c)
d)
36.
Aircraft A has an aspect ratio of 13.7, and has a larger wing area than aircraft B.
Aircraft B has an aspect ratio of 13.7, and has a smaller wing area than aircraft A.
Aircraft B has an aspect ratio of 12.75, and has a smaller wing area than aircraft A.
Aircraft A has an aspect ratio of 12.75, and has a smaller wing area than aircraft B.
Increase the angle of attack to compensate for the decreasing dynamic pressure.
Maintain a constant angle of attack until the desired airspeed is reached, then increase
the angle of attack.
Increase angle of attack to produce more lift than weight.
Decrease the angle of attack to compensate for the decrease in drag.
Take-off from an airfield with a low density altitude will result in:
a)
b)
c)
d)
a longer take-off run.
a higher than standard lAS before lift off.
a higher TAS for the same lift offIAS.
a shorter take off run because of the lower TAS required for the same lAS.
5 - 37
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
I No I A I B I c
1
LIFT
I DII
REF
III
No
19
B
2
D
20
3
D
21
4
IA IB I c ID
C
B
B
23
C
6
B
24
C
A
25
8
A
26
9
10
D
B
A
28
A
11
12
27
C
D
29
C
30
B
B
A
13
C
31
14
C
32
B
33
B
15
D
16
B
34
17
B
35
18
D
36
5 - 39
I
B
5
7
REF
C
22
A
I
D
C
A
D
© Oxford Aviation Services Limited
CHAPTER 6 - DRAG
Contents
Page
INTRODUCTION ....................................................... 6 - 1
PARASITE DRAG. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .. 6 - 2
SKIN FRICTION DRAG
SURFACE CONDITION
SPEED OF FLOW
ADVERSE PRESSURE GRADIENT ........................... 6 - 3
FORM (PRESSURE) DRAG
LAMINAR AND TURBULENT SEPARATION .................. 6 - 4
STREAMLINING
PROFILE DRAG
INTERFERENCE DRAG ........................................... 6 - 5
FACTORS AFFECTING PARASITE DRAG
INDUCED DRAG ....................................................... 6 - 6
WING TIP VORTICES
INDUCED DOWNWASH
FACTORS THAT AFFECT INDUCED DRAG .......................... 6 - 8
SIZE OF THE LIFT FORCE
SPEED OF THE AIRCRAFT
ASPECT RATIO OF THE WING
THE INDUCED DRAG COEFFICIENT .............................. 6 - 11
METHODS OF REDUCING INDUCED DRAG ........................ 6 - 12
WING END-PLATE
TIP TANKS
WINGLETS
WINGTIP SHAPE
EFFECT OF LIFT ON PARASITE DRAG ................................... 6 - 13
EFFECT OF CONFIGURATION
EFFECT OF ALTITUDE
EFFECT OF SPEED
AEROPLANE TOTAL DRAG ............' ................................ 6 - 14
THE EFFECT OF AIRCRAFT GROSS WEIGHT ....................... 6 - 16
THE EFFECT OF ALTITUDE ...................................... 6 - 17
THE EFFECT OF CONFIGURATION
SPEED STABILITY ..................................................... 6 - 18
POWER REQUIRED .................................................... 6 - 20
VMPNMD RELATIONSHIP ............................................. 6 - 21
SUMMARY ........................................................... 6 - 22
SELF ASSESSMENT QUESTIONS ........................................ 6 - 24
ANSWERS ..................................................... 6 - 33
TOTAL DRAG
I
I
INDUCED DRAG
PARASITE DRAG
I
SKIN FRICTION
FORM
INTERFERENCE
DRAG
DRAG
DRAG
I
I
PROFILE
DRAG
Figure 6.0 Total Drag
DRAG
PRINCIPLES OF FLIGHT
6.1
INTRODUCTION
Drag is the force which resists the forward motion of the aircraft. Drag acts parallel to and in
the same direction as the relative airflow (in the opposite direction to the flight path). Please
remember that when considering airflow velocity it does not make any difference to the airflow
pattern whether the aircraft is moving through the air or the air is flowing past the aircraft: it is
the relative velocity which is the important factor.
TOTAL
REACTION
LIFT
I
I
I
I
I
I
(~
...
:
RELATIVE AIRFLOW ~~~~~~~_ _ _---'-_ _ _ _ .ll!c~~~.
DRAG
.------------------ AIRCRAFT FLiGHTPATH
Figure 6.1
Every part of an aeroplane exposed to the airflow produces different types of resistance to
forward motion which contribute to the Total Drag. Total Drag is sub-divided into two main
types:
(1)
PARASITE DRAG - independent of lift generation, and
(2)
INDUCED DRAG - the result oflift geperation.
Parasite drag is further sub-divided into:
Note:
a)
Skin Friction Drag
b)
Form (Pressure) Drag, and
c)
Interference Drag
Skin Friction and Form Drag are together known as PROFILE DRAG.
Induced drag will be considered later. We will first consider the elements of parasite drag.
6-1
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DRAG
PRINCIPLES OF FLIGHT
6.2
PARASITE DRAG
If an aircraft were flying at zero lift angle of attack the only drag present would be parasite drag.
Parasite drag is made-up of 'Skin Friction','Form' and 'Interference' Drag.
6.2.1
SKIN FRICTION DRAG: Particles of air in direct contact with the surface are
accelerated to the speed of the aircraft and are carried along with it. Adjacent particles
will be accelerated by contact with the lower particles, but their velocity will be slightly
less than the aircraft because the viscosity of air is low. As distance from the surface
increases less and less acceleration of the layers of air takes place. Therefore, over the
entire surface there will exist a layer of air whose relative velocity ranges from zero at
the surface to a maximum at the boundary of the air affected by the presence of the
aircraft. The layer of air extending from the surface to the point where no viscous effect
is detectable is known as the boundary layer. In flight, the nature of the boundary layer
will determine the maximum lift coefficient, the stalling characteristics, the value of
form drag, and to some extent the high speed characteristics of an aircraft.
TRANSITION
POINT
I
Figure 6.2
Consider the flow of air across a flat surface, as in Fig 6.2. The boundary layer will
exist in two forms, either laminar or turbulent. In general, the flow at the front will be
laminar and become turbulent some distance back, known as the transition point. The
increased rate of change in velocity at the surface in the turbulent flow will give
more skin friction than the laminar flow. A turbulent boundary layer also has a
higher level of kinetic energy than a laminar layer.
Forward movement of the transition point will increase skin friction because there will
be a greater area of turbulent flow. The position of the transition point is dependent
upon:
a)
Surface condition - The thin laminar layer is extremely sensitive to surface
irregularities. Any roughness on the skin of a leading portion of an aircraft will cause
transition to turbulence at that point and the thickening, turbulent boundary layer will
spread out fanwise down-stream causing a marked increase in skin friction drag.
6-2
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DRAG
PRINCIPLES OF FLIGHT
b)
Adverse pressure gradient (Fig. 6.3) - A laminar layer cannot exist when pressure is
rising in the direction of flow. On a curved surface, such as an aerofoil, the transition
point is usually at, or near to the point of maximum thickness. Because of the adverse
pressure gradient existing on a curved surface the transition point will be further forward
than if the surface was flat.
TRANSITION
VELOCITY DECREASING
~
ADVERSE
PRESSURE INCREASING -...j~~ PRESSURE
(in the direction of flON)
GRADIENT
SEPARATION
Figure 6.3
NB:
The vertical scale of the boundary layer in the above sketch is greatly exaggerated.
Typically, boundary layer thickness is from 2 millimetres at the leading edge, increasing
to about 20 millimetres at the trailing edge.
6.2.2
FORM (PRESSURE) DRAG: Results from the pressure at the leading edge of a body
being greater than the pressure at the trailing edge. Overall, skin friction causes a
continual reduction of boundary layer kinetic energy as flow continues back along the
surface. The adverse pressure gradient behind the transition point will cause an
additional reduction in kinetic energy of the boundary layer. If the boundary layer does
not have sufficient kinetic energy in the presence of the adverse pressure gradient, the
lower levels of the boundary layer stop moving (stagnate). The upper levels of the
boundary layer will overrun at this point (separation point) and the boundary layer will
separate from the surface at the separation point. See Fig. 6.3 . Also, surface flow aft
of the separation point will be forward, toward the separation point - a flow reversal.
Because of separation there will be a lower pressure at the trailing edge than the leading
edge. An aerodynamic force will act in the direction of the lower pressure - form drag.
Separation will occur when the boundary layer does not have sufficient kinetic
energy in the presence of a given adverse pressure gradient.
6-3
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DRAG
PRINCIPLES OF FLIGHT
Loss of kinetic energy in the boundary layer can be caused by various factors.
a)
As angle of attack increases, the transition point moves closer to the leading edge and
the adverse pressure gradient becomes stronger. This causes the separation point to
move forward. Eventually, boundary layer separation will occur so close to the leading
edge that there will be insufficient wing area to provide the required lift force, C L will
decrease and stall occurs.
b)
When a shock wave forms on the upper surface, the increase of static pressure through
the shock wave will create an extreme adverse pressure gradient. If the shock wave is
sufficiently strong, separation will occur immediately behind the shock wave. This will
be explained fully in Chapter 13 - High Speed Flight.
Laminar and Turbulent Separation: Separation has been shown to be caused by the airflow
meeting an adverse pressure gradient, but it is found that a turbulent boundary layer is more
resistant to separation than a laminar one when meeting the same pressure gradient. In this
respect the turbulent boundary layer is preferable to the laminar one, but from the point of view
of drag the laminar flow is preferable.
Streamlining: Each part of an aircraft will be subj ect to form (pressure) drag. To reduce form
drag it is necessary to delay separation to a point as close to the trailing edge as possible.
Streamlining increases the ratio between the length and depth ofa body, reducing the curvature
of the surfaces and thus the adverse pressure gradient. Fineness ratio is the measure of
streamlining. It has been found that the ideal fineness ratio is 3:1, as illustrated in Fig. 6.4.
NB:
The addition of fairings and fillets (see Glossary, Page 1-10) at the junction of
components exposed to the airflow is also referred to as "Streamlining".
Fineness Ratio
=
Length
Depth
Depth
~--
Length
--~I
Figure 6.4
Profile Drag: The combination of skin friction and form drag is known as profile drag. It can
be considered that these drags result from the "profile" (or cross-sectional area) of the aircraft
presented to the relative airflow.
6-4
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PRINCIPLES OF FLIGHT
DRAG
6.2.3
INTERFERENCE DRAG: When considering a complete aircraft, parasite drag will
be greater than the sum of the parts. Additional drag results from boundary layer
'interference' at wing/fuselage, wing/engine nacelle and other such junctions. Filleting
is necessary to minimise interference drag.
6.2.4
FACTORS AFFECTING PARASITE DRAG
1.
Indicated Air Speed
Parasite Drag varies directly with the square of the Indicated Air Speed (lAS).
If lAS is doubled the Parasite Drag will be four times greater - if lAS is halved
the Parasite Drag will be one quarter of its previous value.
2.
Configuration
Parasite Drag varies directly in proportion to the frontal area presented to the
airflow; this is known as 'Parasite Area'. If flaps are deployed, the
undercarriage lowered, speed brakes selected or roll control spoilers operated,
'Parasite Area' is increased and Parasite drag will increase.
3.
Airframe Contamination
Contamination by ice, frost, snow, mud or slush will increase the Parasite Drag
Coefficient, and in the case of severe airframe icing, the area.
THE PARASITE DRAG FORMULA
V2
C Dp
s
where,
Parasite Drag
Dynamic Pi"essure (Q)
Parasite Drag Coefficient
Area (Parasite Area)
6-5
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DRAG
PRINCIPLES OF FLIGHT
6.3
INDUCED DRAG
Induced drag is an undesirable by-product of lift. Wingtip vortices modify upwash and
downwash in the vicinity of the wing which produces a rearward component to the lift vector
known as induced drag.
The lower the lAS, the higher the angle of attack - the stronger the vortices.
The stronger the vortices - the greater the induced drag.
Wing Tip Vortices: Airflow over the top surface of a wing is at a lower pressure than that
beneath. The trailing edge and the wing tips are where the airflows interact, Fig. 6.5. The
pressure differential modifies the directions of flow, inducing a span-wise vector towards the
root on the upper surface and towards the tip on the lower surface. "Conventionally", an aircraft
is viewed from the rear. An anti-clockwise vortex will be induced at the right wing-tip and a
clock-wise vortex at the left wing-tip, Fig. 6.6. At higher angles of attack (Lower IAS) the
decreased chordwise vector will increase the resultant spanwise flow , making the vortices
stronger.
UPPER SURFACE
(LONer Pressure)
$8@ : @ ~
Figure 6.6
Figure 6.5
Induced Downwash: Wingtip vortices create certain vertical velocity components in the
airflow in the vicinity of the wing, both in front of and behind it, Fig. 6.8. These vertical
velocities strengthen upwash and downwash which reduces the effective angle of attack. The
stronger the vortices, the greater the reduction in effective angle of attack.
Due to the localised reduction in effective angle of attack, the overall lift generated by a wing
will be below the value that would be generated ifthere were no spanwise pressure differential.
It is the production of lift itself which reduces the magnitude of the lift force being generated.
To replace the lift lost by the increased upwash and downwash the wing must be flown at a
higher angle of attack, than would otherwise be necessary. This increases drag. This extra drag
is called Induced drag, Fig. 6.9.
6-6
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DRAG
PRINCIPLES OF FLIGHT
EFFECTIVE AIRFLOW
RELATIVE AIRFLOW
!
Tip vortices increase upwash
Oller outer portions of span
Tip vortices increase dONf"MIash
over outer portions of span
INCREASED DOWNWASH AND UPNASH
REDUCES EFFECTIVE ANGLE OF ATTACK
OVER OUTER PORTIONS OF SPAN
Figure 6.7
Upwash Increased
,AI...
t
,AI...
.~
.
IIIHn
Vertical Velocities
in the vicinity of
the wing are a function of
tip vortex strength
Downwash Increased
EFFECTIVE AIRFLOW
......
v
Relative Airflow
~
--~)
~-"V-~~
Angular deflection of effective airflow
is a function of both vortex strength
and True Air Speed (TAS).
-
-,
lod"ood
Downwash
Figure 6.8
6-7
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DRAG
PRINCIPLES OF FLIGHT
oc. e = effective angle of attack
oc. i = induced angle of attack
Induced Drag (Oi)
Normal
1+-_
_ Lift Inclined Rearwards because of
Decreased Effective Angle of Attack
'--Effective
AirflON
(
'---
_ _ _ _ oc.
..;r-
!
e
\
----.~-~---~
Relative AirflCNV
Figure 6.9
Factors that Affect Induced Drag:
(a)
The size of the lift force - Because induced drag is a component of the lift force, the
greater the lift, the greater will be the induced drag. Lift must be equal to weight in
level flight so induced drag will depend on the weight of the aircraft. Induced drag will
be greater at higher aircraft weights. Certain manoeuvres require the lift force to be
greater than the aircraft weight. The relationship oflift to weight is known as the "Load
Factor" (or 'g'). For example, lift is greater than weight during a steady tum so induced
drag will be higher during a steady tum than in straight and level flight. Therefore,
induced drag also increases as the Load Factor increases. Induced drag will increase
in proportion to the square of the lift force.
Load Factor
=
Lift
Weight
(b)
The speed of the aircraft - Induced drag decreases with increasing speed (for a
constant lift force). This is because as speed increases the downwash caused by the tip
vortices becomes less significant, the ' rearward inclination of the lift is less, and
therefore induced drag is less. Induced drag varies inversely as the square of the
speed. (Refer to page 6 - 11 for a detailed explanation)
(c)
The aspect ratio of the wing - The tip vortices of a high aspect ratio wing affect a
smaller proportion of the span so the overall change in downwash will be less, giving
a smaller rearward tilt to the lift force. Induced drag therefore decreases as aspect ratio
increases (for a given lift force). The induced drag coefficient is inversely
proportional to the aspect ratio.
6-8
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
DRAG
From the previous three factors it is possible to develop the following equation:
It can be seen that the relationship for the induced drag coefficient, (C D i), emphasises the need
of a high aspect ratio wing for aeroplane configurations designed to operate at the higher lift
coefficients during the major portion of their flight, i.e. conventional high speed jet transport
aircraft.
The effect of aspect ratio on lift and drag characteristics is shown in Figs. 6.10 and 6.11 . The
basic aerofoil section properties are shown on these plots and these properties would be typical
only of a wing planform of extremely high (infinite) aspect ratio. When a wing of some finite
aspect ratio is constructed of this basic section, the principal differences will be in the lift and
drag characteristics - the moment characteristics remain essentially the same.
The effect of increasing aspect ratio on the lift curve, Fig. 6.10, is to decrease the wing angle of
attack necessary to produce a given lift coefficient. Higher aspect ratio wings are more sensitive
to changes in angle of attack, but require a smaller angles of attack for maximum lift.
1.4
AR= 5
WING
CL
12
AR=2
BASIC SECTION
AR =INFINITE
1.0
0 .8
(NO SWEEPBACK) - - 1 - - -
0 .6
0.4
02
0
5
10
15
20
25
WING ANGLE OF ATTACK
Figure 6.10
6-9
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
DRAG
From Fig. 6.11 it can be seen that at any lift coefficient, a higher aspect ratio gives a lower wing
drag coefficient since the induced drag coefficient varies inversely with aspect ratio. When the
aspect ratio is high the induced drag varies only slightly with lift. At high lift coefficients (low
lAS), the induced drag is very high and increases very rapidly with lift coefficient.
\
1.4
\(
WING
C L
BASIC SECTION
AR = INFINITE
12
I
I
AR= 18
I
f..--
0.8
V
0 .6
1(/
0.4
If
02
./
~
/-
.--
I
---
-
AR=2
\
~
- I
J.
./
I
AR=5
AR= 12
t
VV Y
/I V
I 1/
IIV
1.0
-
~
~
,..~-,--
-
-
.
( LOW MACH NUMBER)
/
I
I
I
f
I
0 I
0.05
0 .10
0 .15
020
025
WING DRAG COEFFICIENT CD
Figure 6.11
The lift and drag curves for a high aspect ratio wing, Figs. 6.10 and 6.11, show continued strong
increase in C L with a up to stall and large changes in Co only at the point of stall.
Continuing to increase aspect ratio is restricted by the following considerations.
Very high aspect ratio wings will experience the following:a)
b)
c)
Excessive wing bending moments: which can be reduced by carrying fuel in
the wings and mounting the engines in pods beneath the wing.
Reduced rate of roll (particularly at low airspeed): This is caused by the
down-going wing (only while it is actually moving down) experiencing an
increased effective angle of attack. The increased effective angle of attack is
due to the resultant oftbe forward T AS of the wing and the angular T AS of the
tip. The higher the aspect ratio, the greater the vertical TAS of the tip for a
given roll rate, leading to a greater increase in effective angle of attack. The
higher the effective angle of attack at the tip, the greater the resistance to roll.
This phenomena is called aerodynamic damping and will be covered in more
detail in later chapters.
Reduced ground clearance in roll during take-off and landing.
6 - 10
© Oxford Aviation Services Limited
DRAG
PRINCIPLES OF FLIGHT
The Induced Drag Coefficient (C D i)
s
This equation would seem to imply that induced drag (D j ) increases with speed, but the induced
drag coefficient (CD i) is proportional to C L2 and inversely proportional to wing aspect ratio. As
speed increases, to maintain a constant lift force CL must be reduced. Thus, with an increase
in speed CD j decreases:
=
2
CL
AR
The following example illustrates the change in CDj with speed, which leads to the change in D j •
If an aircraft's speed is increased from 80 kt (41 m/s) to 160 kt (82 m/s) the dynamic pressure
will be four times greater. (Sea level ISA density is used in the example, but any constant
density will give the same result).
Q
=
0.5 X 1.225 X 41 X 41
= 1029.6
Q
=
0.5 X 1.225 X 82 X 82
= 4118.4
Referring to the lift formula: L
=
Q
CL
S
If dynamic pressure is four times greater because speed is doubled, C L must be reduced to Y4 of
its previous value to maintain a constant lift force.
Applying 114 of the previous CL to the CDj formula: CDi
because AR is constant C Oi
=
2
CL
AR
= (~)2
If 1116 of the previous CDj is applied to the Induced drag formula:
-
1/
/4
Conclusion: If speed is doubled in level flight: dynamic pressure will be four times greater, CL
must be decreased to 114 of its previous value, C Di will be 1116 of its previous value and D j will
be reduced to 114 of its previous value.
If speed is halved in level flight: dynamic pressure will be 114 of its previous value, CL will need
to be four times greater, C Di will be 16 times greater, giving four times more D j
6 - 11
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DRAG
PRINCIPLES OF FLIGHT
6.4
METHODS OF REDUCING INDUCED DRAG
Induced drag is low at high speeds, but at low speeds it comprises over half the total drag.
Induced drag depends on the strength of the trailing vortices, and it has been shown that a high
aspect ratio wing reduces the strength ofthe vortices for a given lift force. However, very high
aspect ratios increase the wing root bending moment, reduce the rate of roll and give reduced
ground clearance in roll during take-off and landing, therefore aspect ratio has to be kept within
practical limits. The following list itemises other methods used to minimise induced drag by
weakening the wing tip vortices.
a)
Wing End-plates: A flat plate placed at the wing tip will restrict the tip vortices and
have a similar effect to an increased aspect ratio, but without the extra bending loads.
However, the plate itself will cause parasite drag, and at higher speeds there may be no
overall saving in drag.
b)
Tip Tanks: Fuel tanks placed at the wing tips will have a similar beneficial effect to
an end plate, will reduce the induced drag, and will also reduce the wing root bending
moment.
c)
Winglets: Small vertical aerofoils which form part of the wing tip (Fig. 6.12). Shaped
and angled to the induced airflow, they generate a small forward force (i.e. "negative
drag", or thrust). Winglets partly block the air flowing from the bottom to the top
surface of the wing, reducing the strength of the tip vortex. In addition, the small vortex
generated by the winglet interacts with and further reduces the strength of the main
wingtip vortex.
d)
Wing tip shape: The shape of the wing tip can affect the strength of the tip vortices,
and designs such as turned down or turned up wing tips have been used to reduce
induced drag.
Winglet
Figure 6.12
6 - 12
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DRAG
PRINCIPLES OF FLIGHT
6.5
EFFECT OF LIFT ON PARASITE DRAG
The sum of drag due to form, friction and interference is termed "parasite" drag because it is not
directly associated with the development of lift. While parasite drag is not directly associated
with the production of lift, in reality it does vary with lift. The variation of parasite drag
coefficient, CD p' with lift coefficient, C L , is shown for a typical aeroplane in Fig. 6.13.
I
1.4 I - - - - COp
11
1.2
CL
/
1.0
0.2
o
/v -il
V
CO i
y
0.6
0.4
k'
//
0 .8
2
c
= ARL
CL
1.0
_CO Pmin
-
I
0 .8
0 .6
0.4
0 .2 I
V- COpmin
I
o
~~-
1.2 -
/
L
I
1.4
0 .05
I
0 .10
0 .15
o
o
I
I
A\
-J- =1=--./
~
1/
I
Co
= COp min.
+
co·I
I
~~
0 .05
0 .15
0 .10
Co
Figure 6.13
Figure 6.14
However, the part of parasite drag above the minimum at zero lift is included with the
induced drag coefficient. Fig. 6.14.
Effect of Configuration: Parasite drag, D p' is unaffected by lift, but is variable with dynamic
pressure and area. If all other factors are held constant, parasite drag varies significantly with
frontal area. As an example, lowering the landing gear and flaps might increase the parasite
areaby as much as 80%. At any given lAS this aeroplane would experience an 80% increase
in parasite drag.
Effect of Altitude: In most phases of flight the aircraft will be flown at a constant lAS, the
dynamic pressure and, thus parasite drag will not vary. The T AS would be higher at altitude to
provide the same lAS.
Effect of Speed: The effect of speed alone on parasite drag is the most important. If all other
factors are held constant, doubling the speed will give four times the dynamic pressure and
hence, four times the parasite drag, (or one quarter as much parasite drag at half the original
speed). This variation of parasite drag with speed points out that parasite drag will be of greatest
importance at high lAS and of much lower significance at low dynamic pressures. To illustrate
this fact, an aeroplane in flight just above the stall speed could have a parasite drag which is only
25% of the total drag. However, this same aeroplane at maximum level flight speed would have
a parasite drag which is very nearly 100% of the total drag. The predominance of parasite drag
at high flight speeds emphasises the necessity for great aerodynamic cleanliness (streamlining)
to obtain high speed performance.
6 - 13
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DRAG
PRINCIPLES OF FLIGHT
6.6
AEROPLANE TOTAL DRAG
The total drag of an aeroplane in flight is the sum of induced and parasite drag. Fig. 6.15
illustrates the variation of total drag with IAS for a given aeroplane in level flight at a particular
weight and configuration.
TOTAL DRAG
DRAG
L/Omax
Parasite Drag
Induced Drag
Vrrd
lAS
Figure 6.15
Fig. 6.15 shows the predominance of induced drag at low speed and parasite drag at high speed.
Because of the particular manner in which parasite and induced drags vary with speed the speed
at which total drag is a minimum (V md) occurs when the induced and parasite drag are
equal. The speed for minimum drag is an important reference for many items of aeroplane
performance. Range, endurance, climb, glide, manoeuvre, landing and take-off performance are
all based on some relationship involving the aeroplane total drag curve. Since flying at V Old
incurs the least total drag for lift-equal-weight flight, the aeroplane will also be at LID max angle
of attack (approximately 4 °).
6 - 14
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
DRAG
It is important to remember that LID max is obtained at a specific angle of attack and also that the
maximum Lift/Drag ratio is a measure of aerodynamic efficiency.
NB:
If an aircraft is operated at the LID max angle of attack, drag will be a minimum while generating
the required lift force. Any angle of attack lower or higher than that for LID max increases the
drag for a given lift force; greater drag requires more thrust, which would be inefficient, and
expensive. It must also be noted that if lAS is varied, LID will vary.
Fig. 5.4 illustrated LID ratio plotted against angle of attack. An alternative presentation of LID
is a polar diagram in which C L is plotted against CD' as illustrated in Fig. 6.16.
I
I
·-~-------+----+---" '-- --t-- ---1
I
I
-K-I
-----~t-+- I
i--+---+l----+------+----+---+----+- -
I
L
t --
I
J
--'--------'_u-t-tJ
I
I
Figure 6.16
The C L / CD , whole aeroplane polar diagram in Fig. 6.16 shows C L increasing initially much
more rapidly than CD' but that ultimately CD increases more rapidly. The condition for maximum
Lift/Drag ratio may be found from the drag polar by drawing the tangent to the curve from the
ongIn.
NB:
This is a very common method of displaying LID ratio, so the display in Fig. 6.16 should become
well known.
6 - 15
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PRINCIPLES OF FLIGHT
6.7
DRAG
THE EFFECT OF AIRCRAFT GROSS WEIGHT ON TOTAL DRAG
The effect of variation in aircraft gross weight on total drag can be seen from Fig. 6.17. As fuel
is consumed gross weight will decrease. As the aircraft weight decreases less lift is required
(lower C L) which will reduce induced drag. Total drag will be less and V md will occur at a lower
lAS.
If an aircraft is operated at a higher gross weight, more lift will be required. If more lift is
generated, induced drag will be higher. Total drag will be greater and V md will occur at a higher
lAS. If an aircraft is manoeuvred so that the load factor is increased, the result will be similar
to that caused by an increase in gross weight. i.e. induced drag will increase.
DRAG
Decreased
TOTAL DRAG
at ICMler weight
Parasite Drag
Less
Induced Drag
at lower weight
Decreased
V rnd
lAS
because of ICMler weight
Figure 6.17
6 - 16
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PRINCIPLES OF FLIGHT
6.8
DRAG
THE EFFECT OF ALTITUDE ON TOTAL DRAG
Aircraft usually operate within limits ofIndicated Air Speed (lAS), so it is relevant to consider
the variation of drag with lAS. If an aircraft is flown at a constant lAS, dynamic pressure will
be constant. As density decreases with increasing altitude, TAS must be increased to maintain
the constant lAS (Q = 'is p V 2). If the aircraft is flown at a constant lAS, drag will not vary with
altitude.
6.9
THE EFFECT OF CONFIGURATION ON TOTAL DRAG
Extension of the landing gear, airbrakes, or flaps will increase parasite drag, but will not
substantially affect induced drag. The effect of increasing parasite drag is to increase total drag
at any lAS but to decrease the speed V m d compared to the clean aircraft, (Fig. 6.18).
Increased
TOTAL DRAG
DRAG
/
/
/
..............
/
~ -- -- -- --
/
/
/
/
/
",
I
\
I
Increased
'\ Parasite Drag
(e .g . flaps , undercarraige or
speed-brakes)
"I",
/'
Induced Drag
Decreased
Vmd
because of increased paraSite drag
lAS
Figure 6.18
6 - 17
© Oxford Aviation Services Limited
DRAG
PRINCIPLES OF FLIGHT
6.10
SPEED STABILITY
For an aircraft to be in steady flight the aircraft must be in equilibrium - there can be no out of
balance forces or moments. When an aircraft is trimmed to fly at a steady speed, thrust and drag
are equal. Therefore, when an aircraft is in steady flight it can be said that the term DRAG and
the term 'THRUST REQUIRED' have the same meaning.
Consequently, an alternative to considering DRAG against lAS as in the graph of Fig. 6.15, the
term 'THRUST REQUIRED' can be substituted for drag.
For an aircraft in steady flight, ifthere is a variation in speed with no change in throttle setting,
(which is called ' THRUST AVAILABLE'), depending on the trim speed, there will be either
an excess or a deficiency ofthrust available. This phenomena is illustrated in Fig. 6.19.
DRAG
Thrust
Excess
or
Thrust
Required
I
I
Thrust
Available
Thrust
Deficiency
Thrust
Excess
Thrust
I
I ~~
~------'~
Non
Stable
lAS
Region I
Stable lAS region
Neutral
lAS
Region
Figure 6.19
6 - 18
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
DRAG
If an aircraft is established in steady flight at point 'A' in Fig. 6.19, lift is equal to weight and
the thrust available is set to match the thrust required. If the aircraft is disturbed to some
airspeed slightly greater than point 'A', a thrust deficiency will exist and, if the aircraft is
disturbed to some airspeed slightly lower than point 'A', a thrust excess will exist. This
relationship provides a tendency for the aircraft to return to the equilibrium of point 'A' and
resume the original trim speed. Steady flight at speeds greater than V md is characterised by a
relatively strong tendency of the aircraft to maintain the trim speed quite naturally; the aircraft
is speed stable.
Speed stability is an important consideration, particularly at speeds at and below V md' most often
encountered during the approach to landing phase of flight.
If an aircraft is established in steady flight at point 'B' in Fig. 6.19, lift is equal to weight and
the thrust available is set to match the thrust required. If the aircraft is disturbed and goes faster
than the trim speed there will be a decrease in drag giving an excess of thrust which will cause
the aircraft to accelerate. If a disturbance slows the aircraft below the trim speed there will be
an increase in drag which will give a thrust deficiency causing the aircraft to slow further. This
relationship is basically unstable because the variation of excess thrust to either side of point' B'
tends to magnify any original disturbance. Steady flight at speeds less than V md is characterised
by a tendency for the aircraft to drift away from the trim speed and the aircraft is speed
unstable. If a disturbance reduces speed it will naturally continue to reduce. If a disturbance
increases speed it will continue to drift up to V md. For this reason, the pilot must closely monitor
lAS during the approach phase of flight. Any tendency for the aircraft to slow down must be
countered immediately by a 'generous' application of thrust to quickly return to the desired trim
speed.
Consider Fig. 6.19. If an aircraft maintains a constant lAS in the speed unstable region, the
addition of parasite drag by selecting undercarriage down or by deploying flaps has the benefit
of reducing V md which can improve speed stability by moving the speed stable region to the left.
At speeds very close to V md an aircraft usually exhibits no tendency towards either speed
stability or speed instability - the neutral lAS region.
6 - 19
© Oxford Aviation Services Limited
DRAG
PRINCIPLES OF FLIGHT
6.11
POWER REQUIRED (Introduction)
We will now consider the relationship between Thrust, Drag and Power. These sound like
engine considerations which might be better studied in Book 4, but it has already been shown
that Drag can also be referred to as 'Thrust Required' and you will now see that a similar
relationship exists with 'Power Required' - they are both important airframe considerations.
1.
Thrust is a FORCE (a push or a pull), used to oppose Drag,
but Power is the RATE of doing WORK, or
and
WORK
=
so POWER must be
POWER
=
WORK
TIME
FORCE x DISTANCE
FORCE x DISTANCE
TIME
For Power Required:
a)
Which Force?
Drag.
b)
Distance divided by time is speed.
Which speed?
The only speed there is - the speed of the aircraft through the
air, True Air Speed (TAS).
Therefore,
2.
POWER REQUIRED
=
DRAG x TAS
If an aircraft climbs at a constant lAS, Drag will remain constant, but T AS must be
increased - so Power Required will increase.
It is necessary to consider Power Required when studying Principles of Flight because Work
must be done on the aircraft to "raise" it to a higher altitude when climbing. Logically,
maximum work can be done on the aircraft in the minimum time when the power available from
the engine(s) is greatest and the power required,by the airframe is least.
For easy reference, associate the word POWER with the word RATE. e.g. minimum rate of
descent is achieved in a steady glide when the aircraft is flown at the minimum power required
speed (VMP ).
These and other considerations will be examined more fully during the study of Aircraft
Performance in Book 6 and Flight Mechanics in Chapter 12 of this Book.
6 - 20
© Oxford Aviation Services Limited
DRAG
PRINCIPLES OF FLIGHT
POWER
REQUIRED (kW )
(DRAG x TAS)
THRUST REQUIRED
or
\jD~
,
.-
- - - -
DRAG (kN)
- - - - -
TAS (kts)
Figure 6.20
Figure 6.20 is drawn for sea level conditions where TAS = lAS and is valid for one particular
aircraft, for one weight, only in level flight, and shows how a graph of TAS against 'Power
Required' has been constructed from a TAS /Drag curve by multiplying each value of drag by
the appropriate TAS and converting it to kilowatts.
The speed for minimum power required is known as V MP and is an Indicated Air Speed (lAS).
Note that the speed corresponding to minimum Power Required (V MP)' is slower than the
speed for minimum drag (VMD)'
Effect of Altitude: An aircraft flying at VMD will experience constant drag at any altitude
because VMD is an lAS. At altitude the TAS for a given lAS is higher, but the power required
also increases by the amount (Power Required = Drag x TAS). So the ratio of TAS to Power
Required is unaffected and V MP will remain slower than V MD.
This information primarily concerns aircraft performance, but the relationship of speed for
minimum power required (V MP) and speed for minimum drag (V MO) is important for the study
of rate and angle of descent in a steady glide, outlined in Chapter 12.
6 - 21
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
6.11
DRAG
SUMMARY
(1)
Parasite drag is made up of:
(a)
Skin friction drag.
(b)
Form (Pressure) drag.
Skin friction plus Form drag is known as Profile Drag.
(c )
Interference drag.
Parasite Drag varies directly as the square of the Indicated Air Speed (lAS) - Double the
speed, four times the parasite drag. Halve the speed, one quarter the parasite drag.
The designer can minimise Parasite Drag by:(i)
(ii)
(iii)
Streamlining.
Filleting and
The use of laminar flow wing sections.
Flight crews must ensure the airframe, and the wing in particular, is not contaminated by ice,
snow, mud or slush.
(2)
Induced drag:(a)
Spanwise airflow generates wingtip vortices.
(b)
The higher the C L (the lower the lAS) the stronger the wingtip vortices.
(c)
Wingtip vortices strengthen downwash.
(d)
Strengthened down wash inclines wing lift rearwards.
(e)
The greater the rearward inclination of wing lift the greater the Induced Drag.
Induced Drag varies inversely as the square of the Indicated Air Speed (lAS) - Halve the
speed, 16 times the induced drag coefficient (CDD and four times the induced drag (Di)' Double
the speed, one sixteenth the CDi and one quarter the Di.
The designer can minimise Induced Drag by:(i)
(ii)
(iii)
Using a high aspect ratio wing planform.
A tapered wing planform with wing twist and/or spanwise camber variation, or
Incorporation of wing end-plates, tip tanks, winglets or various wing tip shapes.
6 - 22
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
(3)
(4)
(5)
DRAG
Total drag.
(a)
Total drag is the sum of Parasite and Induced drag.
(b)
Total drag is a minimum when Parasite and Induced drag are equal.
(c)
At low lAS Induced drag is dominant.
(d)
At high lAS Parasite drag dominates.
(e)
The lAS at which Parasite and Induced drag are equal is called minimum drag
speed (V md).
(f)
As gross weight decreases in flight Induced drag decreases, Total drag
decreases and V md decreases.
(g)
At a constant lAS, altitude has no affect on Total drag, but TAS will increase
as density decreases with increasing altitude.
(h)
Configuration changes which increase the "Parasite Area", such as
undercarriage, flaps or speed brakes, increases Parasite drag, increases Total
drag and decreases V md.
Speed stability.
(a)
An aircraft flying at a steady lAS higher than V md with a fixed throttle setting
will have speed stability.
(b)
An aircraft flying at a steady lAS at V md or slower with a fixed throttle setting
will usually NOT have speed stability.
(c)
If an aircraft flying at a steady lAS and a fixed throttle setting within the nonstable lAS region encounters a disturbance which slow the aircraft, the aircraft
will tend to slow further; lAS will tend to continue to decrease and Total drag
increase.
Power Required
(a)
V MP, the Indicated Air Speed for minimum' Power Required' is slower than the
minimum drag speed (V MD).
(b)
Maximum TAS/Power ratio (1·32 V MP) occurs at V MD.
6 - 23
© Oxford Aviation Services Limited
DRAG
PRINCIPLES OF FLIGHT
SELF ASSESSMENT QUESTIONS
1.
What is the effect on total drag of an aircraft if the airspeed decreases in level flight below that
speed for maximum L/D?
a)
b)
c)
d)
2.
By changing the angle of attack of a wing, the pilot can control the airplane's:
a)
b)
c)
d)
3.
Parasite drag increases more than induced drag.
Induced drag increases more than parasite drag.
Both parasite and induced drag are equally increased.
Both parasite and induced drag are equally decreased.
In theory, if the airspeed of an airplane is doubled while in level flight, parasite drag will
become:
a)
b)
c)
d)
5.
lift and airspeed, but not drag.
lift, gross weight, and drag.
lift, airspeed, and drag.
lift and drag, but not airspeed.
What is the relationship between induced and parasite drag when the gross weight is increased?
a)
b)
c)
d)
4.
Drag increases because of increased induced drag.
Drag decreases because of lower induced drag.
Drag increases because of increased parasite drag.
Drag decreases because of lower parasite drag.
twice as great.
half as great.
four times greater.
one quarter as much.
As airspeed decreases in level flight below that speed for maximum lift/drag ratio, total drag of
an aeroplane:
a)
b)
c)
d)
decreases because of lower parasite drag.
increases because of increased parasite drag.
increases because of increased induced drag.
decreases because of lower induced drag.
6 - 24
© Oxford Aviation Services Limited
DRAG
PRINCIPLES OF FLIGHT
6.
(Refer to annex 'A') At the airspeed represented by point B, in steady flight, the airplane will
a)
b)
c)
d)
7.
Which statement is true relative to changing angle of attack?
a)
b)
c)
d)
8.
induced drag, and is greatly affected by changes in airspeed.
induced drag, and is not affected by changes in airspeed.
parasite drag, and is greatly affected by changes in airspeed.
parasite drag, which is inversely proportional to the square of the airspeed
The best LID ratio of an aircraft occurs when parasite drag is:
a)
b)
c)
d)
11.
flightpath.
longitudinal axis.
chord line.
longitudinal datum
That portion of the aircraft's total drag created by the production of lift is called:
a)
b)
c)
d)
10.
A decrease in angle of attack will increase pressure below the wing, and decrease drag.
An increase in angle of attack will decrease pressure below the wing, and increase drag.
An increase in angle of attack will increase drag.
A decrease in angle of attack will decrease pressure below the wing and increase drag.
On a wing, the force of lift acts perpendicular to, and the force of drag acts parallel to the:
a)
b)
c)
d)
9.
have its maximum LID ratio.
have its minimum LID ratio.
be developing its maximum coefficient of lift.
be developing its minimum coefficient of drag
a minimum.
less than induced drag.
greater than induced drag.
equal to induced drag.
An aircraft has a LID ratio of 15:1 at 50 kts in calm air. What would the LID ratio be with a
direct headwind of 25 kts?
a)
b)
c)
d)
30 : 1
15 : 1
25 : 1
7.5 : 1
6 - 25
© Oxford Aviation Services Limited
DRAG
PRINCIPLES OF FLIGHT
12.
Which is true regarding aerodynamic drag?
a)
b)
c)
d)
13.
At a given True Air Speed, what effect will increased air density have on the lift and drag of an
aircraft?
a)
b)
c)
d)
14.
two times greater.
four times greater.
half as much.
one quarter as much.
The best LID ratio of an aircraft in a given configuration is a value that:
a)
b)
c)
d)
17.
four times greater.
six times greater.
two times greater.
one quarter as much.
If the Indicated Air Speed of an aircraft is decreased from 100 kts to 50 kts, induced drag will
be:
a)
b)
c)
d)
16.
Lift will increase but drag will decrease.
Lift and drag will increase.
Lift and drag will decrease.
Lift and drag will remain the same.
If the Indicated Air Speed of an aircraft is increased from 50 kts to 100 kts, parasite drag will be:
a)
b)
c)
d)
15.
Induced drag is a by-product of lift and is greatly affected by changes in airspeed.
All aerodynamic drag is created entirely by the production of lift.
Induced drag is created entirely by air resistance.
Parasite drag is a by-product of lift.
varies with Indicated Air Speed.
varies depending upon the weight being carried.
varies with air density.
remains constant regardless of Indicated Air Speed changes.
The tendency of an aircraft to develop forces which restore it to its original condition, when
disturbed from a condition of steady flight, is known as:
a)
b)
c)
d)
manoeuverability.
controllability.
stability.
instability.
6 - 26
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
18.
As Indicated Air Speed increases in level flight, the total drag of an aircraft becomes greater than
the total drag produced at the maximum lift/drag speed because of the:
a)
b)
c)
d)
19.
decrease, and parasite drag will decrease.
increase, and induced drag will decrease.
decrease, and parasite drag will increase.
increase and induced drag will increase.
Which statement is true regarding aeroplane flight at L/Dmax?
a)
b)
c)
d)
23.
Parasite drag varies inversely as the square of the airspeed.
Induced drag increases as the square of the airspeed.
Parasite drag increases as the square of the lift coefficient divided by the aspect ratio.
Induced drag varies inversely as the square of the airspeed.
If the same angle of attack is maintained in ground effect as when out of ground effect, lift will:
a)
b)
c)
d)
22.
induced drag.
form drag.
parasite drag.
interference drag.
Which relationship is correct when comparing drag and airspeed?
a)
b)
c)
d)
21.
decrease in induced drag only.
increase in induced drag.
increase in parasite drag.
decrease in parasite drag only.
The resistance, or skin friction, due to the viscosity of the air as it passes along the surface of a
wing is a type of:
a)
b)
c)
d)
20.
DRAG
Any angle of attack other than that for LlDmax increases parasite drag.
Any angle of attack other than that for L/Dmax increases the lift/drag ratio.
Any angle of attack other than that for L/Dmax increases total drag for a given
aeroplane's lift.
Any angle of attack other than that for L/Dmax increases the lift and reduces the drag~
Aspect ratio of a wing is defined as the ratio of the:
a)
b)
c)
d)
square of the chord to the wingspan.
wingspan to the wing root.
area squared to the chord.
wingspan to the mean chord.
6 - 27
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
24.
A wing with a very high aspect ratio (in comparison with a low aspect ratio wing) will have:
a)
b)
c)
d)
25.
4.
2.
3.
1.
(Refer to annex 'B') Consider only aspect ratio (other factors remain constant). Which aircraft
will generate greatest lift?
a)
b)
c)
d)
29.
3.
4.
2.
1.
(Refer to annex 'B') Which aircraft has the lowest aspect ratio?
a)
b)
c)
d)
28.
increased drag, especially at a low angle of attack.
decreased drag, especially at a high angle of attack.
increased drag, especially at a high angle of attack.
decreased drag, especially at low angles of attack.
(Refer to annex 'B') Which aircraft has the highest aspect ratio?
a)
b)
c)
d)
27.
poor control qualities at low airspeeds.
increased drag at high angles of attack.
a low stall speed.
reduced bending moment on its attachment points.
At a constant velocity in airflow, a high aspect ratio wing will have (in comparison with a low
aspect ratio wing):
a)
b)
c)
d)
26.
DRAG
1.
2.
3.
4.
(Refer to annex 'B') Consider only aspect ratio (other factors remain constant). Which aircraft
will generate greatest drag?
a)
b)
c)
d)
1.
4.
3.
2.
6 - 28
© Oxford Aviation Services Limited
DRAG
PRINCIPLES OF FLIGHT
30.
What happens to total drag when accelerating from C Lmax to maximum speed?
a)
b)
c)
d)
31.
Increases.
Increases then decreases.
Decreases.
Decreases then increases.
(Refer to annex 'C'), the whole aircraft CL against CD polar. Point 'B' represents:
1 - Best LiftlDrag ratio.
2 - The critical angle of attack.
3 - Recommended approach speed.
4 - Never exceed speed (V NE).
a)
b)
c)
d)
32.
1 and 2
1 only
2 only
4 only
If the Indicated Air Speed of an aircraft in level flight is increased from 100 kt to 200 kt, what
change will occur in (i) TAS (ii) COi (iii) Di ?
a)
b)
c)
d)
(i)
(ii)
(iii)
2
0
114
4
1/16
1116
1116
16
114
114
4
2
6 - 29
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
DRAG
ANNEX 'A'
DRAG
lAS
ANNEX 'B'
Aircraft 1.
Span 22.5 metres
Chord 4 metres
Aircraft 2.
Wing Area 90 square metres
Span 45 metres
Aircraft 3.
Span 30 metres
Chord 3 metres
Aircraft 4.
Wing Area 90 square metres
Span 40 metres
6 - 31
© Oxford Aviation Services Limited
DRAG
PRINCIPLES OF FLIGHT
ANNEX
'e'
!
--
/
/~
/c
v
/
v
-~~-.------------.
B
L
I
6 - 32
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
DRAG
ANSWERS
I No I A I B I c I DII
1
A
2
C
3
B
REF
III
No
17
C
18
C
19
C
4
C
20
5
C
21
6
7
8
A
24
9
A
25
10
D
12
13
14
15
16
D
A
B
26
C
D
28
A
29
B
B
A
30
A
31
B
D
32
6 - 33
I
C
27
B
REF
B
23
C
I
D
22
A
11
IA I B I c ID
D
B
D
© Oxford Aviation Services Limited
CHAPTER 7 - STALLING
Contents
Page
INTRODUCTION ....................................................... 7 - 1
CAUSE OF THE STALL
THE LIFT CURVE ....................................................... 7 - 2
STALL RECOVERY
AIRCRAFT BEHAVIOUR CLOSE TO THE STALL ........................... 7 - 3
USE OF FLIGHT CONTROLS CLOSE TO THE STALL
STALL RECOGNITION .................................................. 7 - 4
STALL SPEED
STALL WARNING ...................................................... 7 - 6
ARTIFICIAL STALL WARNING DEVICES
STICK SHAKER .................................................. 7 - 7
STALL WARNING VANE .......................................... 7 - 8
ANGLE OF ATTACK VANE ........................................ 7 - 9
ANGLE OF ATTACK PROBE
BASIC STALL REQUIREMENTS ......................................... 7 - 10
WING DESIGN CHARACTERISTICS
THE EFFECT OF AEROFOIL SECTION
THE EFFECT OF WING PLAN FORM ..................................... 7 - 12
RECTANGULAR WING
TAPERED WING ................................................ 7 - 13
WASHOUT
THICKNESS AND CAMBER
LEADING EDGE SLOTS
STALL STRIPS ........................................... 7 - 14
VORTEX GENERATORS ................................... 7 - 15
SWEPT BACK WING ............................................ 7 - 16
TIP STALL AND PITCH-UP
WING FENCES ........................................... 7 - 17
VORTILONS
SA W TOOTH LEADING ED9E
KEY FACTS 1 - SELF STUDy ........................................... 7 - 18
SUPER STALL (DEEP STALL) ........................................... 7 - 22
SUPER STALL PREVENTION - STICK PUSHER ...................... 7 - 23
FACTORS THAT EFFECT STALL SPEED .................................. 7 - 24
1G STALL SPEED
WEIGHT CHANGE .............................................. 7 - 25
COMPOSITION AND RESOLUTION OF FORCES .............. 7 - 26
PARALLELOGRAM OF FORCES
BASIC TRIGONOMETRY
LIFT INCREASE IN A LEVEL TURN ......................... 7 - 27
FACTORS THAT EFFECT STALL SPEED - CONTINUED.
LOAD FACTOR .................................................
HIGH LIFT DEVICES ............................................
CG POSITION ...................................................
LANDING GEAR POSITION .......................................
ENGINE POWER ................................................
PROPELLER
JET
MACH NUMBER ................................................
WING CONTAMINATION ........................................
ICE
FROST
SNOW
WARNING TO PILOT OF ICING-INDUCED STALLS ..................
STABILISER STALL DUE TO ICE ..................................
EFFECTS OF HEAVY RAIN ON STALL SPEED
STALL AND RECOVERY CHARACTERISTICS OF CANARDS
SPINNING ............................................................
PRIMARY CAUSES OF A SPIN
PHASES OF A SPIN ..............................................
EFFECT OF MASS & BALANCE ON SPINS ..........................
SPIN RECOVERY
SPECIAL PHENOMENA OF STALL .......................................
CROSSED-CONTROL STALL
ACCELERATED STALL
SECONDARY STALL
LARGE AIRCRAFT
SMALL AIRCRAFT ..............................................
POWER ON AND POWER OFF
CLIMBING AND DESCENDING TURNS ............................
HIGH SPEED BUFFET
ANSWERS (V s IN 25 0 AND 30 0 BANK) ...................................
KEY FACTS 2 - SELF STUDY ............................................
SELF ASSESSMENT QUESTIONS ........................................
KEY FACTS 1 AND 2 (COMPLETE) ......................................
NB:
7 - 28
7 - 29
7 - 30
7 - 31
7 - 32
7 - 34
7 - 36
7 - 38
7 - 39
7 - 40
7 - 41
7 - 42
7 - 44
7 - 45
7 - 46
7 - 48
7 - 49
7 - 53
7 - 63
Throughout this chapter reference will be made to JAR stall requirements etc, but it must
be emphasised that these references are for training purposes only and are not subject to
amendment action.
PRINCIPLES OF FLIGHT
7.1
STALLING
INTRODUCTION
Stalling is a potentially hazardous manoeuvre involving loss of height and loss of control. A
pilot must be able to clearly and unmistakably identify an impending stall, so that it can be
prevented. Different types of aircraft exhibit various stall characteristics; some less desirable
than others. Airworthiness authorities specify minimum stall qualities that an aircraft must
possess.
7.2
CAUSE OF THE STALL
The C L of an aero foil increases with angle of attack up to a maximum (C L MAX). Any further
increase above this stalling, or critical angle of attack, will make it impossible for the airflow
to smoothly follow the upper wing contour, the flow will separate from the surface, causing C L
to decrease and drag to increase rapidly. Since the CL MAX of an aerofoil corresponds to the
minimum steady flight speed (the Ig stall speed), it is an important point of reference.
A stall is caused by airflow separation. Separation can occur when either the boundary layer has
insufficient kinetic energy or the adverse pressure gradient becomes too great.
Fig. 7.1 shows that at low angles of attack
virtually no flow separation occurs before the
trailing edge, the flow being attached over the
rear part of the surface in the form of a
turbulent boundary layer.
As angle of attack increases, the adverse
pressure gradient increases, reducing the
kinetic energy, and the boundary layer will
begin to separate from the surface at the
trailing edge.
Further increase in angle of attack makes the
separation point move forward and the wing
area that generates a pressure differential
, becomes smaller. At angles of attack higher
than approximately 16 the extremely steep
adverse pressure gradient will have caused so
much separation that insufficient lift IS
generated to balance the aircraft weight.
0
,
Figure 7.1
It is important to remember that the angle of attack is the
angle between the chord line and the relative airflow.
Therefore, if the angle of attack is increased up to or
beyond the critical angle, an aeroplane can be stalled at
any airspeed or flight attitude.
7-1
An aeroplane can be stalled
at any airspeed or attitude
© Oxford Aviation Services Limited
STALLING
PRINCIPLES OF FLIGHT
7.3
THE LIFT CURVE
I St~1I
I
o
I
4
I
8
I
12
Angle of Attack in Degrees
I
16
(00)
Figure 7.2
Fig. 7.2 shows that as the angle of attack increases from the zero lift value, the curve is linear
over a considerable range. As the effects of separation begin to be felt, the slope of the curve
begins to fall off. Eventually, lift reaches a maximum and begins to decrease. The angle at
which it does so is called the stalling angle or critical angle of attack, and the corresponding
value of lift coefficient is C L MAX . A typical stalling angle is about 16 0 •
7.4
STALL RECOVERY
To recover from a stall or prevent a full stall, the angle of attack must be decreased to
reduce the adverse pressure gradient. This may consist of merely releasing back pressure, or
it may be necessary to smoothly move the pitch control forward, depending on the aircraft design
and severity of the stall. (Excessive forward movement of the pitch control however, may
impose a negative load on the wing and delay recovery). For most modem jet transport aircraft
it is usually sufficient to lower the nose to the horizon or just below, while applying maximum
authorised power to minimise height loss.
On straight wing aircraft the rudder should bet used to prevent wing drop during stall and
recovery. On swept wing aircraft it is recommended that the ailerons be used to prevent wing
drop, with a small amount of smoothly applied co-ordinated rudder. (The rudder on modern high
speed jet transport aircraft is very powerful and careless use can give too much roll, leading to
pilot induced oscillation - PIO).
Allow airspeed to increase and recover lost altitude with moderate back pressure on the pitch
control. Pulling too hard could trigger a secondary stall, or worse, could exceed the limit load
factor and damage the aircraft structure. As angle of attack reduces below the critical angle, the
adverse pressure gradient will decrease, airflow will re-attach, and lift and drag will return to
their normal values.
7-2
© Oxford Aviation Services Limited
STALLING
PRINCIPLES OF FLIGHT
7.5
AIRCRAFT BEHAVIOUR CLOSE TO THE STALL
Stall characteristics vary with different types of aircraft. However, for modem aircraft during
most normal manoeuvres, the onset of stall is gradual. The first indications of a stall may be
provided by any or all of the following:a)
unresponsive flight controls,
b)
a stall warning or stall prevention device, or
c)
aerodynamic buffet.
The detailed behaviour of various aircraft types will be discussed later.
7.6
USE OF FLIGHT CONTROLS CLOSE TO THE STALL
At low speeds normally associated with stalling, dynamic pressure is at a very low value and
greater control deflection will be required to achieve the same response; also the flying controls
will feel unresponsive or "mushy". If an accidental stall does occur it is vitally important that
the stall and recovery should occur without too much wing drop. Moving a control surface
modifies the chord line and hence the angle of attack. An aircraft being flown close to the stall
angle may have one wing that produces slightly less lift than the other; that wing will tend to
drop. Trying to lift a dropping wing with aileron will increase its angle of attack, Fig. 7.3, and
may cause the wing to stall completely, resulting in that wing dropping at an increased rate. At
speeds close to the stall ailerons must be used with caution. On straight wing aircraft the
rudder should be used to yaw the aircraft just enough to increase the speed of a dropping wing
to maintain a wings level attitude. Swept wing aircraft basic stall requirements are designed to
enable the ailerons to be used successfully up to ' stall recognition' (Page 7-4 and Para. 7.11),
but small amounts of rudder can be used if smoothly applied and co-ordinated with the ailerons.
Figure 7.3
7-3
© Oxford Aviation Services Limited
STALLING
PRINCIPLES OF FLIGHT
7.7
STALL RECOGNITION
The aeroplane is considered stalled when the behaviour of the aeroplane gives the pilot a clear
and distinctive indication of an acceptable nature that the aeroplane is stalled.
Acceptable indications of a stall, occurring either individually or in combination, are:-
7.8
(1)
A nose-down pitch that cannot be readily arrested;
(2)
Buffeting, of a magnitude and severity that is a strong and effective deterrent to further
speed reduction; or
(3)
The pitch control reaches the aft stop and no further increase in pitch attitude occurs
when the control is held full aft for a short time before recovery is initiated.
STALL SPEED
It is necessary to fly at slow speeds (high angles of attack) during take-off and landing in order
to keep the required runway lengths to a reasonable minimum. There must be an adequate
safety margin between the minimum speed allowed for normal operations and the stall
speed.
Prototype aircraft are stalled and stall speeds established for inclusion in the Flight Manual
during the flight testing that takes place before type certification.
'Small' aircraft (JAR-23) use Vso and V S1 on which to base the stall speed.
For 'Large' aircraft (JAR-25) a reference stall speed, V SR , is used.
a)
The reference stall speed (V SR ) is a calibrated airspeed defined by the aircraft
manufacturer. V SR may not be less than a I-g stall speed. V SR is expressed as:-
Where:-
Calibrated airspeed obtained when the load factor corrected lift coefficient is
first a maximum during the manoeuvre prescribed in sub-paragraph (c) of this
paragraph.
In addition, when the manoeuvre is limited by a device that abruptly pushes the
nose down at a selected angle of attack (e.g. a stick pusher), V CLMAX may not be
less than the speed existing at the instant the device operates.
Load factor normal to the flight path at
7-4
VCLMAX
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
NB
STALLING
On aircraft without a stick pusher, VSR can be considered to be the same as the Ig stall
speed (VSlg). But it is impossible to fly at speeds less than that at which the stick pusher
activates, so for aircraft fitted with a stick pusher, VSR will be 2 knots or 2% greater than
the speed at which the stick pusher activates. (See Figures 7.3a and 7.3b for an illustration
of the designations of stall speed and stall warning).
From the 'sample' aeroplane on Page 5-6, the speed at CLMAX was 150 kts. This can be
considered as that aeroplane's VCLMAX. At Ig, VSR would therefore be 150 kts.
(b)
VCLMAX is determined with:(1)
Zero thrust at the stall speed
(2)
Propeller pitch controls (if applicable) in the take-off position
(3)
The aeroplane in other respects (such as flaps and landing gear) in the condition
existing in the test or performance standard in which V SR is being used
(4)
The weight used when V SR is being used as a factor to determine compliance
with a required performance standard
(5)
The centre of gravity position that results in the highest value of reference stall
speed; and
(6)
The aeroplane trimmed for straight flight at a speed selected by the
manufacturer, but not less than 1·13 V SR and not greater than 1·3 V SR.
(c)
Starting from the stabilised trim condition [see (b)( 6) above], apply the longitudinal
control to decelerate the aeroplane so that the speed reduction does not exceed one knot
per second.
(d)
In addition to the requirements of sub-paragraph (a) of this paragraph, when a device
that abruptly pushes the nose down at a selected angle of attack (e.g. a stick pusher) is
installed, the reference stall speed V SR, may not be less than 2 knots or 2%, whichever
is the greater, above the speed at which the device operates.
V SR will vary with each of the above conditions. Additional factors which affect V SR are load
factor, thrust in excess of zero and wing contamination. All these effects will be detailed later.
Density altitude does not
effect indicated stall speed
7-5
© Oxford Aviation Services Limited
STALLING
PRINCIPLES OF FLIGHT
7.9
STALL WARNING
Having established a stall speed for each configuration there must be clear and distinctive
warning, sufficiently in advance of the stall, for the stall itself to be avoided.
(a)
Stall warning with sufficient margin to prevent inadvertent stalling with the flaps and
landing gear in any normal position must be clear and distinctive to the pilot in straight
and turning flight.
(b)
The warning may be furnished either through the inherent aerodynamic qualities of the
aeroplane or by a device that will give clearly distinguishable indications under expected
conditions of flight. However, a visual stall warning device that requires the
attention of the crew within the cockpit is not acceptable by itself. If a warning
device is used, it must provide a warning in each of the aeroplane configurations
prescribed in sub-paragraph ( a) of this paragraph at the speed prescribed in subparagraph (c) and (d) of this paragraph.
(c )
When the speed is reduced at rates not exceeding 1 knot per second, stall warning must
begin, in each normal configuration, at a speed, V sw, exceeding the speed at which the
stall is identified in accordance with Paragraph 7.7 by not less than 5 knots or 5%
CAS, whichever is the greater. Once initiated, stall warning must continue until the
angle of attack is reduced to approximately that at which stall warning began.
(d)
In addition to the requirements of sub-paragraph (c) of this paragraph, when the speed
is reduced at rates not exceeding one knot per second, in straight flight with engines
idling and CO position specified in Paragraph 7.8 (b)(5), V sw , in each normal
configuration must exceed V SR by not less than 3 knots or 3% CAS, whichever is
greater.
(e)
The stall warning margin must be sufficient to allow the pilot to prevent stalling (as
defined in Paragraph 7.7) when recovery is initiated not less than one second after the
onset of stall warning in slow-down turns with at least 1·5g load factor normal to the
flight path and airspeed deceleration rates of at least 2 knots per second, with the flaps
and landing gear in any normal position, with the aeroplane trimmed for straight flight
at a speed of 1·3 V SR, and with the power or thrust necessary to maintain level flight at
1·3 V SR .
(f)
Stall warning must also be provided in each abnormal configuration of the high lift
devices that is likely to be used in flight following system failures (including all
configurations covered by Flight Manual procedures).
7-6
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
V
STALLING
C Ltv1AX
I
V SR
I
VS
VSw
I
1g
5kt or 5%--'
~.--------~.--------------------------~~ CAS
I
Para. 7.7
(1,2&3)
Figure 7.3a Aircraft without stick pusher
STICK
PUSH
V SR
2kt
or--.
2%
Vsw
3kt
or --.
3%
~----~------~------------------~CAS
Figure 7.3b Aircraft fitted with Stick Pusher
7.10
ARTIFICIAL STALL WARNING DEVICES
Adequate stall warning may be provided by the airflow separating comparatively early and
giving aerodynamic buffet by shaking the wing and by buffeting the tailplane, perhaps
transmitted up the elevator control run and shaking the control column, but this is not usually
sufficient, so a device which simulates natural buffet is usually fitted to all aircraft.
Artificial stall warning on small aircraft is usually given by a buzzer or hom. The artificial stall
warning device used on modern large aircraft is a stick shaker, in conjunction with lights and a
noise-maker.
Stick shaker: A stick shaker represents what it is replacing; it shakes the stick and is a tactile
warning. Ifthe stick shaker activates when the pilot's hands are not on the controls: when the
aircraft is on autopilot, for example, a very quiet stick shaker could not function as a stall
warning so a noise maker is added in parallel.
The stick shaker is a pair of simple electric motors, one clamped to each pilot's control column,
rotating an out of balance weight. When the motor runs it shakes the stick.
7-7
© Oxford Aviation Services Limited
STALLING
PRINCIPLES OF FLIGHT
An artificial stall warning device can receive its signal from a number of different types of
detector switch, all activated by changes in angle of attack,
FLAPPER SW ITCH
( activated by movement of stagnation point)
STAGNATION POINT -------'
(has moved downwards and backwards around leading edge)
Figure 7.4
Flapper Switch
Flapper switch (leading edge stall warning vane): Fig. 7.4. As angle of attack increases, the
stagnation point moves downwards and backwards around the leading edge. The flapper switch
is so located, that at the appropriate angle of attack, the stagnation point moves to its underside,
and the increased pressure lifts and closes the switch.
7-8
© Oxford Aviation Services Limited
STALLING
PRINCIPLES OF FLIGHT
AS ANGLE OF ATIACK INCREASES , VANE ROTATES RELATIVE TO FUSELAGE
FUSELAGE
SKIN
Figure 7.5 Angle of Attack Vane
Angle of attack vane: Fig. 7.5. Mounted on the side of the fuselage, the vane streamlines with
the relative airflow and the fuselage rotates around it. The stick shaker is activated at the
appropriate angle of attack.
Angle of attack probe: Also mounted on the side of the fuselage; consists of slots in a probe,
which are sensitive to changes in angle of relative airflow.
All of these sense angle of attack and, therefore, automatically take care of changes in aircraft
mass; the majority also compute the rate of change of angle of attack and give earlier warning
in the case of faster rates of approach to the stall. The detectors are usually datum compensated
for configuration changes and are always heated or anti-iced. There are usually sensors on both
sides to counteract any sideslip effect.
7-9
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STALLING
PRINCIPLES OF FLIGHT
7.11
BASIC STALL REQUIREMENTS (JAR and FAR)
(a)
It must be possible to produce and to correct roll and yaw by unreversed use of
aileron and rudder controls, up to the time the aeroplane is stalled. No abnormal noseup pitching may occur. The longitudinal control force must be positive up to and
throughout the stall. In addition, it must be possible to promptly prevent stalling and
to recover from a stall by normal use of the controls.
7.12
(b)
For level wing stalls, the roll occurring between the stall and the completion of the
recovery may not exceed approximately 20 0 •
(c)
For turning flight stalls, the action of the aeroplane after the stall may not be so violent
or extreme as to make it difficult, with normal piloting skill, to effect a prompt recovery
and to regain control of the aeroplane. The maximum bank angle that occurs during the
recovery may not exceed:
(1)
Approximately 60 degrees in the original direction of the tum, or 30 degrees in
the opposite direction, for deceleration rates up to 1 knot per second; and
(2)
Approximately 90 degrees in the original direction of the tum, or 60 degrees in
the opposite direction, for deceleration rates in excess of 1 knot per second.
WING DESIGN CHARACTERISTICS
It has been shown that stalling is due to airflow separation, characterised by a loss of lift, and
an increase in drag, that will cause the aircraft to lose height. This is generally true, but there
are aspects of aircraft behaviour and handling at or near the stall which depend on the design of
the wing aerofoil section and planform.
7.13
THE EFFECT OF AEROFOIL SECTION
Shape of the aerofoil section will influence the manner in which it stalls. With some sections,
stall occurs very suddenly and the drop in lift is very marked. With others, the approach to stall
is more gradual, and the decrease in lift is less disastrous.
In general, an aeroplane should not stall too suddenly, and the pilot should have adequate
warning, in terms of handling qualities, of the approach of a stall. This warning generally takes
the form of buffeting and general lack of response to the controls. If a particular wing design
stalls too suddenly, it will be necessary to provide some sort of artificial pre-stall warning device
or even a stall prevention device.
~
,
~
A given aerofoil section will always
stall at the same angle of attack
7 - 10
~
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PRINCIPLES OF FLIGHT
STALLING
Features of aerofoil section design which affect behaviour near the stall are:
a)
leading edge radius,
b)
thickness-chord ratio,
c)
camber, and particularly the amount of camber near the leading edge, and
d)
chordwise location of the points of maximum thickness and maximum camber.
Generally, the sharper the nose (small leading edge radius), the thinner the aerofoil section, or
the further aft the position of maximum thickness and camber, the more sudden will be the stall.
i.e. an aerofoil section designed for efficient operation at higher speeds, Fig. 7.6.
The stall characteristics of the above listed aerofoil sections can be used to either encourage a
stall to occur, or delay stalling, at a particular location on the wing span.
1. ROUNDED LEADING EDGE
2 . HIGHER THICKNESS-CHORD RATIO
3. MAX . THICKNESS AND CAMBER MORE FWD .
1. SHARP LEADING EDGE
2. LOW THICKNESS-CHORD RATIO
3. AFT MAXIMUM THICKNESS AND CAMBER
Figure 7.6
7 - 11
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PRINCIPLES OF FLIGHT
7.14
STALLING
THE EFFECT OF WING PLAN FORM
On basic wing plan forms, airflow separation will not occur simultaneously at all spanwise
locations.
STRONG TIP VORTICES
DECREASE EFFECTIVE
ANGLE OF ATTACK
AT WING TIP, THUS
DELAYING TIP STALL.
CP MOVES REARWARDS ,
AIRCRAFT NOSE DROPS .
Figure 7.7
Rectangular Wing
The Rectangular Wing: Fig. 7.7. On a rectangular wing, separation tends to begin at the root,
and spreads out towards the tip. Reduction in lift initially occurs inboard near the aircraft eG,
and if it occurs on one wing before the other, there is little tendency for the aircraft to roll.
The aircraft loses height, but in doing so remains more or less wings level. Loss of lift is felt
ahead of the centre of gravity of the aircraft and the CP moves rearwards, so the nose
drops and angle of attack is reduced. Thus, there is a natural tendency for the aircraft to move
away from the high angle of attack which gave rise to the stall. The separated airflow from the
root immerses the rear fuselage and tail area, and aerodynamic buffet can provide a warning of
the approaching stall. Being located outside of the area of separated airflow, the ailerons tend
to remain effective when the stalling process starts. All of these factors give the most desirable
kind of response to a stall (Ref. para. 7.7, 7.9(b), and 7.11):a)
Aileron effectiveness,
b)
nose drop,
c)
aerodynamic buffet, and
d)
absence of violent wing drop.
Unfortunately, a rectangular wing has unacceptable wing bending characteristics and is not very
aerodynamically efficient, so most modem aircraft have a tapered and/or swept planform.
7 - 12
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STALLING
PRINCIPLES OF FLIGHT
WING TIP IS UNABLE TO
SUPPORT TIP VORTICES ,
CAUSING THEM TO FORM
CLOSER TO THE ROOT.
THIS GIVES A DECREASED
EFFECTIVE ANGLE OF ATTACK
AT THE WING ROOT, THUS
DELAYING THE ROOT STALL.
Figure 7.8
Tapered Wing
The Tapered Wing: Fig. 7.8. Separation tends to occur first in the region of the wing tips,
reducing lift in those areas. If an actual wing were allowed to stall in this way, stalling would
give aileron buffet, and perhaps violent wing drop. (Wing drop at the stall gives an increased
tendency for an aircraft to enter a spin). There would be no buffet on the tail , no strong nose
down pitching moment, and very little, if any, aileron effectiveness. To give favourable stall
characteristics, a tapered wing must be modified using one or more of the following:a)
Geometric twist (washout), a decrease in incidence from root to tip. This decreases the
angle of attack at the tip, and the root will tend to stall first.
b)
The aerofoil section may be varied throughout the span such that sections with greater
thickness and camber are located near the tip. The higher CL M AX of such sections
delays stall so that the root will tend to stall first.
AILERON!
A
SECTION
A-A
Figure 7.9
c)
Leading Edge Slot
Leading edge slots, Fig. 7.9, towards the tip re-energise (increase the kinetic energy of)
the boundary layer. They increase local CL M AX and are useful, both for delaying
separation at the tip, and retaining aileron effectiveness. The function of slats and slots
will be fully described in Chapter 8.
7 - 13
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PRINCIPLES OF FLIGHT
STALLING
STALL STRIP
Figure 7.10
d)
Stall Strip
Another method for improving the stall pattern is by forcing a stall to occur from the
root. An aerofoil section with a smaller leading edge radius at the root would promote
airflow separation at a lower angle of attack but decrease overall wing efficiency. The
same result can be accomplished by attaching stall strips (small triangular strips),
Fig. 7.10, to the wing leading edge.
At higher angles of attack stall strips promote separation, but will not effect the
efficiency of the wing in the cruise.
7 - 14
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
STALLING
Figure 7.11
e)
Vortex Generators
Vortex generators, Fig. 7.11, are rows of small, thin aerofoil shaped blades which
project vertically (about 2'5cm) into the airstream. They each generate a small vortex
which causes the free stream flow of high energy air to mix with and add kinetic energy
to the boundary layer. This re-energises the boundary layer and tends to delay
separation.
7 - 15
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
STALLING
~LATERAL
AXIS
OUTBOARD SUCTION
PRESSURES TEND TO
DRAW BOUNDARY LAYER
TOWARDS TIP.
CP MOVES FORWARD AND
CREATES AN UNSTABLE
NOSE UP PITCHING MOMENT
Figure 7.12
Sweepback: Fig. 7.12. A swept wing is fitted to allow a higher maximum speed, but it has an
increased tendency to stall first near the tips. Loss of lift at the tips moves the CP forward,
giving a nose up pitching moment.
Effective lift production is concentrated inboard and the maximum downwash now impacts the
tailplane, Fig. 7.13, adding to the nose up pitching moment.
Pitch-up: As soon as a swept wing begins to stall, both forward CP movement and increased
downwash at the tailplane cause the aircraft nose to rise rapidly, further increasing the angle of
attack. This is a very undesirable and unacceptable response at the stall and can result in
complete loss of control in pitch from which it may be very difficult, or even impossible, to
recover. This phenomenon is known as pitch-up, and is a very dangerous characteristic of
many high speed, swept wing aircraft.
TIP STALL
UNSTALLED
--------CP
MAX IMUM
DOWNWASH
Figure 7.13
7 - 16
Pitch - Up
© Oxford Aviation Services Limited
STALLING
PRINCIPLES OF FLIGHT
The tendency of a swept back wing to tip stall is due to the induced spanwise flow of the
boundary layer from root to tip. The following design features can be incorporated to
minimise this effect and give a swept wing aircraft more acceptable stall characteristics:-
WING
FENCE
Figure 7.14 Wing Fence
Wing fences (boundary layer fences), Fig. 7.14, are thin metal fences which generally extend
from the leading edge to the trailing edge on the top surface and are intended to prevent outward
drift of the boundary layer.
/
- SAW TOOTH
Figure 7.16 Saw Tooth
Figure 7.15 Vortilon
Vortilons, Fig. 7.15 , are also thin metal fences, but are smaller than a full chordwise fence.
They are situated on the underside of the wing leading edge. The support pylons of pod mounted
engines on the wing also act in the same way. At high angles of attack a small but intense vortex
is shed over the wing top surface which acts as an aerodynamic wing fence.
Saw tooth leading edge, Fig. 7.16, will also generate a strong vortex over the wing upper surface
at high angles of attack, minimising spanwise flow of the boundary layer. (Rarely used on
modem high speed jet transport aircraft).
7 - 17
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PRINCIPLES OF FLIGHT
STALLING
The following four pages contain a revision aid to encourage students to become familiar with
any new terminology, together with the key elements of 'stalling'.
KEY FACTS 1 - Self Study
Insert the missing words in these statements, using the foregoing paragraphs for reference.
Stalling involves loss of ____ and loss of _ _ _ _ . (Para. 7.1)
A pilot must be able to clearly and unmistakably _ _ _ _ _ a stall. (Para. 7.1)
A stall is caused by airflow _ _ _ _ _ _ . (Para. 7.2)
Separation can occur when either the boundary layer has insufficient
_ _ _ _ _ gradient becomes too great. (Para. 7.2)
Adverse pressure gradient increases with increase in angle of
energy or the
. (Para. 7.2)
Alternative names for the angle of attack at which stall occurs are the
_ _ _ _ angle of attack. (Para. 7.3)
angle and the
The coefficient of lift at which a stall occurs is ____ . (Para. 7.3)
A stall can occur at any _ _ _ _ _ or flight _ _ _ _ _ . (Para. 7.2)
A typical stalling angle is approximately
o.
(Para. 7.3)
To recover from a stall the angle of ____ must be _ _ _ _ _ . (Para. 7.4)
Maximum power is applied during stall recovery to minimise _ _ _ _ loss. (Para. 7.4)
7 - 18
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PRINCIPLES OF FLIGHT
STALLING
On small aircraft, the _ _ _ _ should be used to prevent wing _ _ _ at the stall. (Para. 7.4)
On swept wing aircraft the _ _ _ should be used to prevent wing _ _ at the stall.(Para. 7.4)
Recover height lost during stall recovery with moderate _ _ _ pressure on the _ _ __
control. (Para. 7.4)
The first indications of a stall may be
or aerodynamic
. (Para. 7.5)
At speeds close to the stall,
(Para. 7.6)
flight controls, stall
must be used with caution to
device
a dropping wing.
Acceptable indications of a stall are: (Para. 7.7)
(1)
a nose
pitch that can not be readily arrested.
(2)
severe - - - - (3)
pitch control reaching _ _ stop and no further increase in - - - attitude
occurs.
Reference stall speed (V SR ) is a CAS defined by the _ _ _ _ _ _ _ _ _ .(Para. 7.8(a))
V SR may not be _ _ than a __ stall speed. (Para.7.8(a))
is installed,
When a device that abruptly pushes the _ _ _ _ at a selected angle of
V SR may not be _ _ than _ knots or _ %, whichever is
, above the speed at which
the
operates. (Para.7.8(d))
Stall warning with sufficient
to prevent inadvertent stalling must be
_ _ _ _ _ _ to the pilot in straight and turning flight. (Para. 7.9(a))
and
qualities of the aeroplane
Acceptable stall warning may consist of the inherent
that will give clearly distinguishable indications under expected conditions
or by a
offlight. (Para. 7.9(b))
7 - 19
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PRINCIPLES OF FLIGHT
STALLING
Stall warning must begin at a speed exceeding the stall speed by not less than _ knots or _ %
CAS, whichever is the greater. (Para. 7.9(c))
Artificial stall warning on a small aircraft is usually given by a _ _ _ or ____ . (Para. 7.10)
Artificial stall warning on a large aircraft is usually given by a
with
and a noisemaker. (Para. 7.10)
An artificial stall warning device can be activated by a
vane or an angle of attack
. (Para. 7.10)
shaker, in conjunction
switch, an angle of _ _ __
Most angle of attack sensors compute the
of change of angle of attack to give _ _ __
warning in the case of accelerated rates of stall approach. (Para. 7.10)
JAR required stall characteristics, up to the time the aeroplane is stalled, are:- (Para. 7.11)
a)
It must be possible to produce and correct _ _ by unreversed use of the ____ and
b)
c)
d)
No abnormal nose up
may occur
Longitudinal control force must be _ _ __
It must be possible to promptly prevent ____ and recover from a stall by normal use
of the - - - There should be no excessive _ _ between the stall and completion of recovery
For turning flight stalls, the action of the aeroplane after the stall may not be so _ __
or
as to make it difficult, with normal piloting _ _ , to effect prompt
_ _ _ _ and to regain
of the aeroplane
e)
f)
An aerofoil section with a small leading edge ___ will stall at a _ _ _ angle of attack and
the stall will be more
. (Para. 7.13)
An aerofoil section with a large thickness-chord ratio will stall at a ___ angle of attack and
will stall more
. (Para. 7.13)
An aero foil section with camber near the - - - attack. (Para. 7.13)
7 - 20
___ will stall at a higher angle of
© Oxford Aviation Services Limited
STALLING
PRINCIPLES OF FLIGHT
A rectangular wing plan form will tend to stall at the _ _ first. (Para. 7.14)
A rectangular wing planform usually has ideal stall characteristics, these are:- (Para. 7.14)
a)
b)
c)
d)
Aileron - - - - - - at the stall
at the stall
Nose
Aerodynamic
at the stall
Absence of violent wing _ _ at the stall
To give a wing with a tapered planform the desired stall characteristics, the following devices
can be included in the design:- (Para. 7.14)
a)
b)
c)
d)
e)
(decreasing incidence from root to tip)
An aero foil section with
thickness and camber at the tip
Leading edge
at the tip
Stall
fitted to the wing inboard leading edge
generators which re-energise the
layer at the tip
A swept back wing has an increased tendency to tip stall due to the spanwise flow of boundary
layer from root to tip on the wing top surface. Methods of delaying tip stall on a swept wing
planform are:- (Para. 7.14)
a)
b)
c)
d)
e)
Wing
, thin metal fences which generally extend from the leading edge
to the trailing edge on the wing top surface
, also thin metal fences, but smaller and are situated on the underside
of the wing leading edge
Saw _ _ leading edge, generates vortices over wing top surface at high angles
of attack
Engine
of pod mounted wing engines also act as vortilons
generators are also used to delay tip stall on a swept wing
Tip stall on a swept wing planform gives a tendency for the aircraft to _ _- _ at the stall.
This is due to the _ moving forwards when the wing tips stall
. (Para. 7.14)
KEY FACTS 1, WITH WORD INSERTS CAN BE FOUND ON PAGE 7 - 63
7 - 21
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STALLING
PRINCIPLES OF FLIGHT
7.15
SUPER STALL (DEEP STALL)
A swept-back wing tends to stall first near the tips. Since the tips are situated well aft of the eG,
the loss of lift at the tips causes the pitch attitude to increase rapidly and further increase the
angle of attack. Fig. 7.17.
Figure 7.17
Pitch-up
This "automatic" increase in angle of attack, caused by pitch-up, stalls more of the wing. Drag
will increase rapidly, lift will reduce, and the aeroplane will start to sink at a constant, nose
high, pitch attitude. This results in a rapid additional increase in angle of attack, Fig. 7.18.
DOWNWARD INCLINED FLIGHT PATH
TAILPLANE IMMERSED
IN SEPARATED AIRFLOW
FROM STALLED WING
Figure 7.18 Super Stall
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STALLING
PRINCIPLES OF FLIGHT
Separated airflow from the stalled wing will immerse a high-set tailplane in low energy turbulent
air, Fig. 7.18. Elevator effectiveness is greatly reduced making it impossible for the pilot to
decrease the angle of attack. The aeroplane will become stabilized in what is known as the
"super-stall" or "deep-stall" condition.
Clearly, the combination of a swept-back wing and a high mounted tailplane ('T' -Tail) are the
factors involved in the "super or deep-stall". Of the two:THE SWEPT-BACK WING IS THE MAJOR CONTRIBUTORY FACTOR.
It has been shown that the tendency for a swept-back wing to pitch-up can be reduced by design
modifications (wing fences, vortilons and saw tooth leading edge) which minimise the root-to-tip
spanwise flow of the boundary layer. These devices delay tip stall. Vortex generators are also
frequently used on a swept wing to delay tip stall and improve the stall characteristics.
The wing root can also be encouraged to stall first. This can be done by modifying the aerofoil
section at the root, fitting stall strips and by fitting less efficient leading edge flaps (Kruger flaps)
to the inboard section of the wing.
Aircraft such as the DC-9, MD-80, Boeing 727, Fokker 28 and others, have swept-back wings
and high mounted tailplanes ('T'-Tail). They also have rear, fuselage mounted engines. The
only contribution rear mounted engines make is that they are the reason the designer placed the
tailplane on top of the fin in the first place. In-and-of-itself, mounting the engines on the rear
fuselage does not contribute to super stall.
7.16
SUPER STALL PREVENTION - STICK PUSHER
An aircraft design which exhibits super-stall characteristics must be fitted with a device to
prevent it from ever stalling. This device is a stick pusher. Once such an aircraft begins to
stall it is too late; the progression to super stall is too fast for a human to respond, and the aircraft
cannot then be un-stalled.
A stick pusher is a device attached to the elevator control system, which physically pushes the
control column forward, reducing the angle of attack before super-stall can occur.
The force of the push is typically about 80 lbs.' This is regarded as being high enough to be
effective, but not too high to hold in a runaway situation. Provision is made to "dump" the stick
pusher system in the event of a malfunction. Once dumped, the pusher cannot normally be reset
in flight.
Once actuated, the stick pusher will automatically disengage once the angle of attack reduces
below a suitable value.
7 - 23
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STALLING
PRINCIPLES OF FLIGHT
7.17
FACTORS THAT AFFECT STALL SPEED
Paragraph 7.8 details the CAS at which an aircraft stalls (V SR)' We know that stalling is caused
by exceeding the critical angle of attack. Stalling has nothing to do with the speed of the
aircraft; the critical angle of attack can be exceeded at any aircraft speed. However, it has
been shown that if an aircraft is flown in straight and level flight and speed reduced at a rate not
exceeding 1 knot per second, the CAS at which it stalls can be identified. It is upon this
reference stall speed (V SR) that the recommended take-off, manouevre, approach and landing
speeds are based, to give an adequate margin from the stall during normal operations (1'05 V SR,
1·1 V SR ' 1·2 V SR , 1·3 V SR etc).
Factors which can affect V SR are:
a)
b)
c)
d)
e)
f)
g)
7.18
Changes in weight.
Manoeuvring the aircraft (increasing the load factor).
Configuration changes (changes in CL MAX and pitching moment).
Engine thrust and propeller slipstream
Mach number
Wing contamination
Heavy rain
19 STALL SPEED
In straight and level flight the weight of the aircraft is balanced by the lift.
Load Factor (n) or' g'
=
Lift
Weight
While (n) is the correct symbol for load factor, the relationship between lift and weight has for
years been popularly known as 'g'. (1 g corresponds to the force acting on us in every day life).
If more lift is generated than weight the load factor or 'g' will be greater than one; the force
acting on the aircraft and everything in it, including the pilot will be greater.
If Lift = Weight, the load factor will be one and from the lift formula:
it can be seen that lift will change whenever any of the other factors in the formula change. We
consider density (p) and wing area (S) constant for this example. If the engine is throttled
back, drag will reduce speed (V) and, from the formula, it can be seen that lift would decrease.
To keep lift constant and maintain 1g flight at a reduced speed, CL must be increased by
increasing the angle of attack.
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PRINCIPLES OF FLIGHT
STALLING
Any further reduction in speed would need a further increase in angle of attack; each succeeding
lower CAS corresponding to a greater angle of attack. Eventually, at a certain CAS, the wing
reaches its stalling angle (C L MAX)' beyond which any further increase in angle of attack, in an
attempt to maintain lift, will precipitate a stall. We can transpose the lift formula to show this
relationship:-
=
7.19
r
L
S
I
~
Density altitude does not
effect indicated stall speed
~
~
EFFECT OF WEIGHT CHANGE ON STALL SPEED
At CL MAX for 1g flight, a change in weight requires a change in lift and it can be seen from the
V Slg formula that, for instance, an increase in weight (lift) will increase V Slg
The relationship between basic stalling speeds at two different weights can be obtained from the
following formula:-
V
S19 new
=
V
new weight
S19 old
old weight
The angle of attack at which stall occurs will NOT be affected by the weight. (Provided that the
appropriate value ofC L MAX is not affected by speed - as it will be at speeds greater than 0·4M,
ref. para. 7.29). To maintain a given angle of attack in level flight, it is necessary to change
the dynamic pressure (CAS) if the weight is changed.
As an example: at a weight of 588,600 N an aircraft stalls at 150 Kt CAS, what is the V Slg stall
speed at a weight of 470,880 N?
vS19 new
470880
=
150
=
134 knots CAS
588600
Weight does not
effect stall angle
It should be noted that a 20% reduction in weight has resulted in an approximate 10% reduction
in stall speed. (As a "rule of thumb", this relationship can be used to save calculator batteries,
and time in the exam!). The change in stall speed due to an increase in weight can be calculated
in the same way.
7 - 25
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PRINCIPLES OF FLIGHT
7.20
STALLING
COMPOSITION AND RESOLUTION OF FORCES
A force is a vector quantity. It has magnitude and direction, and can be represented by a straight
line, passing through the point at which it is applied, its length representing the magnitude of the
force, and its direction corresponding to that in which the force is acting.
FORCE
VECTOR
FORCE
VECTOR
FORCE
FORCE
VECTOR
FORCE
FORCE
VECTOR
Figure 7.19
The Resolution of A Force Into Two Vectors
As vector quantities, forces can be added or subtracted to form a resultant force, or they can be
resolved - split into two or more component parts by the simple process of drawing the
vectors to represent them. Fig. 7.19.
7.21
THE PARALLELOGRAM OF FORCES
If three forces which act at a point are in equilibrium, they can be represented by the sides of a
triangle taken in order. This is called the principle of the triangle of forces, and the so called
parallelogram of forces is really the same thing, two sides and the diagonal of the parallelogram
corresponding to the triangle.
7.22
USING TRIGONOMETRY TO SOLVE A PARALLELOGRAM OF FORCES
If one of the angles and the length of one of the sides of a right angled triangle is known, it is
possible to calculate the length of the other sides. In a parallelogram of forces the sides of the
triangle represent the magnitude of the force veotors, so it is possible to calculate the magnitude
of the forces.
Opposite
LS~
~
Adjacent
TANQ>
=
Opp
Adj
SINQ>
7 - 26
=
Opp
Hyp
COSQ>
Adj
=-Hyp
© Oxford Aviation Services Limited
STALLING
PRINCIPLES OF FLIGHT
7.23
LIFT INCREASE IN A LEVEL TURN
~
LIFT INCREASE
REQUIRED
ADJACENT
w
Figure 7.20
Fig. 7.20 shows an aircraft in a level, 45 0 bank turn. Weight always acts vertically downwards.
For the aircraft to maintain altitude, the UP force must be the same as the DOWN force. Lift is
inclined from the horizontal by the bank angle of 45 0 and can be resolved into two components,
or vectors; one vertical and one horizontal. It can be SEEN from the illustration that in a level
turn, lift must be increased in order to produce an upwards force vector equal to weight. We
know the vertical force must be equal to the weight, so the vertical force can represented by (I ).
The relationship between the vertical force and lift can be found using trigonometry, where <p
(phi) is the bank angle:-
cos <f> =
L =
ADJ (1)
HYP (L)
0.707
=
transposing this formula gives, L
=
1
cos <f>
1.41
This shows that :-
In a 45° bank, LIFT must be greater than weight by a factor of 1.41
Another way of saying the same thing: in a level 45 0 bank turn, lift must be increased by 41 %
7 - 27
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PRINCIPLES OF FLIGHT
7.24
STALLING
EFFECT OF LOAD FACTOR ON STALL SPEED
It has been demonstrated that to bank an aircraft and maintain altitude, lift has to be greater than
weight. And that, additional lift in a tum is obtained by increasing the angle of attack. To
consider the relationship between lift and weight we use Load Factor.
LOAD FACTOR (n) or 'g'
=
LIFT
WEIGHT
(a)
Increasing lift in a tum, increases the load factor.
(b)
As bank angle increases, load factor increases.
In straight and level flight at CL MAX it would be impossible to tum AND maintain altitude.
Trying to increase lift would stall the aircraft. If a tum was started at an IAS above the stall
speed, at some bank angle CL would reach its maximum and the aircraft would stall at a speed
higher than the 1g stall speed.
The increase of lift in a level turn is a function of the bank angle only. Using the following
formula, it is possible to calculate stall speed as a function of bank angle or load factor.
= VS
where:
VSt
-~
cos <P
Load factor does not "'
effect stall angle
is the stall speed in a turn
Using our example aeroplane: the Ig stall speed is 150 knots CAS, what will be the stall speed
in a 45 ° bank?
VSt
=
150
~
0.;07
=
178 knots CAS
=
212 knots CAS
In a 60° bank the stall speed will be:
=
150
~ 0~5
Stall speed in a 45° bank is 19% greater than V Sig and in a 60° bank the stall speed is 41%
greater than V S ig' and since these are ratio's, this will be true for any aircraft.
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STALLING
PRINCIPLES OF FLIGHT
As bank angle is increased, stall speed will increase at an increasing rate. While operating at
high Cu during take-off and landing in particular, only moderate bank angles should be used to
manoeuvre the aircraft. For a modem high speed jet transport aircraft, the absolute maximum
bank angle which should be used in service is 30 ° (excluding emergency manoeuvres). The
normal maximum would be 25 ° , at higher altitude the normal maximum is 10° to 15 ° .
If the Ig stall speed is 150 kt, calculate the stall speed in a 25° and a 30° bank tum. (Answers
on page 7 - 48).
If the stall speed in a 15° bank tum is 153 kt CAS and it is necessary to calculate the stall speed
in a 45 ° bank tum, you would need to calculate the Ig stall speed first, as follows:-
Vs
t
=
Vs
19
~
cos 150
153
1.02
7.25
=
transposition gives
VS19
=
150 kt CAS
EFFECT OF HIGH LIFT DEVICES ON STALL SPEED
Modem high speed jet transport aircraft have swept wings with relatively low thickness/chord
ratio's (e.g. 12% for an A310). The overall value of C L MAX for these wings is fairly low and the
clean stalling speed correspondingly high. In order to reduce the landing and take off speeds,
various devices are used to increase the usable value ofC LMAX • In addition to decreasing the stall
speed, these high-lift devices will usually alter the stalling characteristics. The devices include:a)
leading-edge flaps and slats
b)
trailing edge flaps
L
From the 1g stall formula:
S
it can be seen that an increase in CL MAX will reduce the stall speed. It is possible, with the most
modem high lift devices, to increase CLMAX by as much as 100%. High lift devices will be fully
described in chapter 8. High lift devices decrease stall speed, hence minimum flight speed, so
provide a shorter take-off and landing run - this is their sole purpose.
7 - 29
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PRINCIPLES OF FLIGHT
7.26
STALLING
EFFECT OF CG POSITION ON STALL SPEED
JAR 25 .103(b) states that V CLMAX is determined with the CG position that results in the highest
value of reference stall speed.
L
t
TAIL
DOWNLOAD
w
Figure 7.21
If the CG is in front ofthe CP, Fig. 7.21, giving a nose down pitching moment and there is no
thrust/drag moment to oppose it, the tailplane must provide a down load to maintain equilibrium.
Lift must be increased to maintain an upwards force equal to the increased downwards force.
From the Ig stall formula it can be seen that CLMAX will be divisible into the increased lift force
more times.
L
=
s
Forward movement of the CG increases stall speed.
7 - 30
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PRINCIPLES OF FLIGHT
7.27
STALLING
EFFECT OF LANDING GEAR ON THE STALL SPEED
L
•
r------,rr----~-------- ,
TA IL
DOWNLOAD
PROFILE DRAG
FROM GEAR
w
Figure 7.22
From Fig. 7.22 it can be seen that with the undercarriage down, profile drag below the CG is
increased. This will give a nose down pitching moment which must be balanced by increasing
the tail down load. Lift must be increased to balance the increased downwards force.
(CG movement due to the direction in which the undercarriage extends will have an insignificant
influence on stall speed). By far the greater influence is the increased profile drag of the gear
when it is extended.
Extending the undercarriage increases stall speed.
7 - 31
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PRINCIPLES OF FLIGHT
7.28
STALLING
EFFECT OF ENGINE POWER ON STALL SPEED
JAR 25 .103(b) states that V CLMAX is determined with zero thrust at the stall speed.
When establishing V CLMAX the engines must be at zero thrust and it is assumed that the weight
of the aircraft is entirely supported by lift. If thrust is applied close to the stall the nose high
attitude of the aircraft produces a vertical component of thrust, Fig. 7.24, which assists in
supporting the weight and less lift is required. Aircraft with propellers will have an additional
effect caused by the propeller slipstream.
The most important factors affecting this relationship are engine type (propeller or jet), thrust
to weight ratio and inclination of the thrust vector at C L MAX'
/
-/-
INDUCED FLOW
FROM PROPELLER
SLIPSTREAM
Figure 7.23
Propeller: Fig. 7.23. The slipstream velocity behind the propeller is greater than the free
stream flow, depending on the thrust developed. Thus, when the propeller aeroplane is at low
airspeeds and high power, the dynamic pressure within the propeller slipstream is much greater
than that outside and this generates much more lift than at zero thrust. The lift of the aeroplane
at a given angle of attack and airspeed will be greatly affected. If the aircraft is in the landing
flare, reducing power suddenly will cause a significant reduction in lift and a heavy landing
could result. On the other hand, a potentially heavy landing can be avoided by ajudicious 'blast'
from the engines.
Jet: The typical jet aircraft does not experience the induced flow velocities encountered in
propeller driven aeroplanes, thus the only significant factor is the vertical component of thrust,
Fig. 7.24. Since this vertical component contributes to supporting the weight of the aircraft, less
aerodynamic lift is required to hold the aeroplane in flight. If the thrust is large and is given a
large inclination at maximum lift angle, the effect on stall speed can be very large. Since there
is very little induced flow from the jet, the angle of attack at stall is essentially the same poweron as power-off.
7 - 32
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STALLING
PRINCIPLES OF FLIGHT
Figure 7.24
Power-on stall speed is less than power-off. This will be shown to be significant during the
study of winds hear in chapter. IS.
7 - 33
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STALLING
PRINCIPLES OF FLIGHT
7.29
EFFECT OF MACH NUMBER (COMPRESSIBILITY) ON STALL SPEED
As an aircraft flies faster, the streamline pattern around the wing changes. Faster than about four
tenths the speed of sound (OAM) these changes start to become significant. This phenomena is
known as compressibility. This will be discussed fully in Chapter 13.
Pressure waves, generated by the passage of a wing through the air, propagate ahead of the wing
at the speed of sound. These pressure waves upwash air ahead of the wing towards the lower
pressure on the top surface.
. HIGH SPEED
LOW SPEED
Figure 7.25
Fig. 7.25 shows that at low speed, the streamline pattern is affected far ahead of the wing and
the air has a certain distance in which to upwash. As speed increases the wing gets closer to its
leading pressure wave and the streamline pattern is affected a shorter distance ahead, so must
approach the wing at a steeper angle.
This change in the streamline pattern accentuates the adverse pressure gradient near the leading
edge and flow separation occurs at a reduced angle of attack. Above OAM C L MAX decreases as
shown in Fig. 7.26.
_._--
- - - - - - - ' - - - - - -- .
0'4
1· 0
M
Figure 7.26
7 - 34
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STALLING
PRINCIPLES OF FLIGHT
Referring to the 1g stall speed formula:
If C LMAX decreases, V Sig will increase.
To maintain a constant EAS as altitude mcreases, T AS is increased. Also, outside air
temperature decreases with increasing altitude, causing the local speed of sound to decrease.
Mach number is proportional to TAS and inversely proportional to the local speed of sound (a):
M
=
TAS
a
Therefore, at a constant EAS, Mach number will increase as altitude increases.
~ 19 Stall Speed
~--~~----------.
EAS
Figure 7.27
Fig. 7.27 shows the variation of stalling speed with altitude at constant load factor (n). Such a
curve is called the stalling boundary for the given load factor, in which altitude is plotted against
equivalent airspeed. At this load factor (lg), the aircraft cannot fly at speeds to the left of this
boundary. It is clear that over the lower range of altitude, stall speed does not vary with altitude.
This is because at these low altitudes, the Mach number at V s is less than 004 M, too low for
compressibility effects to be present. Eventually (approximately 30,000 ft) , Mach number at V s
has increased with altitude to such an extent that these effects are important, and the rise in
stalling speed with altitude is apparent.
Using the example aeroplane from earlier, the VSi g of 150 kt is equal to MO'4 at approximately
29 ,000 ft using ISA values.
As altitude increases, stall speed is initially constant then increases, due to compressibility.
7 - 35
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PRINCIPLES OF FLIGHT
7.30
STALLING
EFFECT OF WING CONTAMINATION ON STALL SPEED
Refer to:AIC 104/1998 "Frost Ice and Snow on Aircraft", and
AIC 98/1999 "Turbo-Prop and other Propeller Driven Aeroplanes: Icing Induced Stalls".
Any contamination on the wing, but particularly ice, frost or snow, will drastically alter the
aerodynamic contour and affect the nature of the boundary layer.
ICE
The formation of ice on the leading edge of the wing will produce:a)
Large changes in the local contour, leading to severe local adverse
pressure gradients.
b)
High surface friction and a considerable reduction of boundary layer
kinetic energy.
These cause a large decrease in CL MAX and can increase stall speed by
approximately 30% with no change in angle of attack.
The added weight of the ice will also increase the stall speed, but the major
factor is the reduction in CL MAX.
FROST
The effect of frost is more subtle. The accumulation of a hard coat of frost on
the wing upper surface will produce a surface texture of considerable
roughness.
Tests have shown that ice, snow or frost, with the thickness and surface
roughness similar to medium or coarse sandpaper on the leading edge and upper
surface ofa wing can reduce lift by as much as 30% (10% to 15% increase in
stall speed) and increases drag by 40%.
While the basic shape and aerodynamic contour is unchanged, the increase in
surface roughness increases skin friction and reduces the kinetic energy of the
boundary layer. Separation will occur at an angle of attack and lift coefficients
lower than for the clean smooth wing.
SNOW
The effect of snow can be similar to frost in that it will increase surface
roughness. If there is a coating of snow on the aircraft it must be removed
before flight. Not only will the snow itself increase skin friction drag, but may
obscure airframe icing. Snow will NOT blow-off during taxi or take-off.
The pilot in command is legally required to ensure the aeroplane is aerodynamically clean at the
time of take-off. It is very important that the holdover time of any de-icing or anti-icing fluid
applied to the airframe is known. If this time will be exceeded before take-off, the aircraft must
be treated again.
7 - 36
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PRINCIPLES OF FLIGHT
STALLING
While the reduction in C L MAX due to frost fonnation is not usually as great as that due to ice
formation, it is usually unexpected because it may be thought that large changes in the
aerodynamic shape (such as due to ice) are necessary to reduce C L MAX • However, kinetic energy
of the boundary layer is an important factor influencing separation of the airflow and this energy
is reduced by an increase in surface roughness. The general effects of ice and frost fonnation
on C L MAX is typified by the illustrations in Fig. 7.28.
Ice , frost and snow change
the aerofoil section,
decrease the stall angle
and increase the stall speed
LEADING EDGE ICE FORMATION
UPPER SURFACE FROST
I
BASIC SMOOTH WING
FROST
ICE
ANGLE OF ATIACK
Figure 7.28
The increase in stall speed due to ice fonnation is not easy to quantify, as the accumulation and
shape of the ice fonnation is impossible to predict. Even a little ice is too much. Ice or frost
must never be allowed to remain on any aerodynamic surfaces in flight, nor must ice, frost, snow
or other contamination be allowed to remain on the aircraft immediately before flight.
7 - 37
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PRINCIPLES OF FLIGHT
7.31
STALLING
WARNING TO THE PILOT OF ICING-INDUCED STALLS
There have been recent cases involving loss of control in icing conditions due to undetected
stalling at speeds significantly above the normal stalling speed, accompanied by violent roll
oscillations.
Control of an aeroplane can be lost as a result of an icing-induced stall, the onset of which can
be so insidious * as to be difficult to detect.
The following advice is offered on the recognition of, and the recovery from, insidious icinginduced wing-stalls:a)
Loss of performance in icing conditions may indicate a serious build-up of airframe
icing (even if this cannot be seen) which causes a gradual loss of lift and a significant
increase in drag;
b)
this build-up of ice can cause the aeroplane to stall at approximately 30% above the
normal stall speed;
c)
The longitudinal characteristics of an icing-induced wing-stall can be so gentle that the
pilot may not be aware that it has occurred;
d)
the stall warning system installed on the aeroplane may not alert the pilot to the
insidious icing-induced wing-stall (angle of attack will be below that required to trigger
the switch), so should not be relied upon to give a warning of this condition. Airframe
buffet, however, may assist in identifying the onset of wing-stall;
e)
the first clue may be a roll control problem. This can appear as a gradually increasing
roll oscillation or a violent wing drop;
f)
a combination of rolling oscillation and onset of high drag can cause the aeroplane to
enter a high rate of descent unless prompt recovery action is taken;
g)
if a roll control problem develops in icing conditions, the pilot should suspect that the
aeroplane has entered an icing-induced wing-stall and should take immediate stall
recovery action (decrease the angle of attack). The de-icing system should also be
activated. If the aeroplane is fitted with an anti-icing system this should have been
activated prior to entry into icing conditions in accordance with the Flight
Manual/Operations Manual procedures and recommendations. If the anti-icing system
has not been in use then it should be immediately activated. Consideration should also
be given to leaving icing conditions by adjusting track and/or altitude if possible.
*Insidious - advancing imperceptibly: without warning
7 - 38
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PRINCIPLES OF FLIGHT
7.32
STALLING
STABILISER STALL DUE TO ICE
The tailplane is an aerofoil, and because it is thinner than the wing, it is likely to experience
icing before the wing does. The effect will be the same as for the wing, the stall will occur at
a lower angle of attack. The tailplane is normally operating at a negative angle of attack,
producing a download, so if the tailplane stalls and the download is lost, the nose of the aircraft
will drop and longitudinal control will be lost.
Stalling of an ice contaminated tailplane could be precipitated by extension of the wing flaps.
Lowering the flaps increases the downwash, and this increases the negative angle of attack of
the tailplane. If the tailplane has ice contamination this could be sufficient to cause it to stall.
Recovery procedure in this situation would be to retract the flaps again, thus reducing the
downwash.
7.33
EFFECT OF HEAVY RAIN ON STALL SPEED
WEIGHT: Heavy rain will form a film of water on an aircraft and increase its weight slightly,
maybe as much as 1 - 2%, this in itself will increase stall speed.
AERODYNAMIC EFFECT: The film of water will distort the aerofoil, roughen the surface,
and alter the airflow pattern on the whole aircraft. CL MAX will decrease causing stall speed to
Increase.
DRAG: The film of water will increase interference drag, profile drag and form drag. In light
rain, drag may increase by 5%, moderate 20% and heavy rain up to 30%. This obviously
increases thrust required.
IMPACT: An additional consideration, while not affecting stall speed, is the effect of the
impact of heavy rain on the aircraft. Momentum will be lost and airspeed will decrease,
requiring increased thrust. At the same time, heavy rain will also be driving the aircraft
downwards. The volume of rain in any given situation will vary, but an aircraft on final
approach which suddenly enters a torrential downpour of heavy rain will be subject to a loss of
momentum and a decrease in altitude, similar to the effect of microburst windshear. (Chapt. 16).
7.34
STALL AND RECOVERY CHARACTERISTICS OF CANARDS
With the conventional rear tailplane configuration the wing stalls before the tailplane, and
longitudinal control and stability are maintained at the stall. On a canard layout if the wing stalls
first, stability is lost, but if the foreplane stalls first then control is lost and the maximum value
of CL is reduced.
7 - 39
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PRINCIPLES OF FLIGHT
7.35
STALLING
SPINNING
When an aircraft is accidentally or deliberately stalled, the motion of the aircraft may in some
cases develop into a spin. The important characteristics of a spin are:a)
the aircraft is descending along a steep helical path about a vertical spin axis,
b)
the angle of attack of both wings is well above the stall angle,
c)
the aircraft has a high rate of rotation about the vertical spin axis,
d)
viewed from above, the aircraft executes a circular path about the spin axis, and the
radius of the helix is usually less than the semi-span of the wing,
e)
the aircraft may be in the "erect" or "inverted" position in the spin.
The spin is one of the most complex of all flight manoeuvres. A spin may be defined as an
aggravated stall resulting in autorotation, which means the rotation is stable and will continue
due to aerodynamic forces if nothing intervenes. During the spin the wings remain unequally
stalled.
7.36
PRIMARY CAUSES OF A SPIN
A stall must occur before a spin can take place. A spin occurs when one wing stalls more than
the other, Fig. 7.29. The wing that is more stalled will drop and the nose of the aircraft will yaw
in the direction of the lower wing.
The cause of an accidental spin is exceeding the critical angle of attack while performing a
manoeuvre with either too much or not enough rudder input for the amount of aileron being used
(crossed-controls). If the correct stall recovery is not initiated promptly, the stall could develop
into a spin.
Co-ordinated use of the flight controls is important, especially during flight at low airspeed and
high angle of attack. Although most pilots are able to maintain co-ordinated flight during routine
manoeuvres, this ability often deteriorates when distractions occur and their attention is divided
between important tasks. Distractions that have caused problems include preoccupation with
situations inside or outside the flight deck, manoeuvring to avoid other aircraft, and manoeuvring
to clear obstacles during take-off, climb, approach or landing.
A spin may also develop if forces on the aircraft are unbalanced in other ways, for example, from
yaw forces due to an engine failure on a multi-engine aircraft, or if the CG is laterally displaced
by an unbalanced fuel load.
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STALLING
PRINCIPLES OF FLIGHT
STALL
o
UPGOING
SEMI-SPAN
/
U
o
z
«
l
DONNGOING
~ SEMI-SPAN
ANGLE OF ATTACK
Figure 7.29
7.37
PHASES OF A SPIN
There are three phases of a spin.
1.
The incipient spin is the first phase, and exists from the time the aeroplane stalls and
rotation starts until the spin is fully developed.
2.
A fully developed spin exists from the time the angular rotation rates, airspeed, and
vertical descent rate are stabilized from one turn to the next.
3.
The third phase, spin recovery, begins when the anti-spin forces overcome the pro-spin
forces.
If an aircraft is near the critical angle of attack, and more lift is lost from one wing than the other)
that wing will drop. Its relative airflow will be inclined upwards, increasing its effective angle
of attack. As the aeroplane rolls around its CG, the rising wing has a reduced effective angle of
attack and remains less stalled than the other. This situation of unbalanced lift tends to increase
as the aeroplane yaws towards the low wing, accelerating the high, outside wing and slowing the
inner, lower wing. As with any stall, the nose drops, and as inertia forces begin to take effect,
the spin usually stabilizes at a steady rate of rotation and descent.
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PRINCIPLES OF FLIGHT
STALLING
It is vitally important that recovery from an unintentional spin is begun as soon as possible, since
many aeroplanes will not easily recover from a fully developed spin, and others continue for
several turns before recovery inputs become effective. Recovery from an incipient spin normally
requires less altitude and time than the recovery from a fully developed spin. Every aeroplane
spins differently, and an individual aeroplane's spin characteristics vary depending on
configuration, loading, and other factors.
7.38
THE EFFECT OF MASS AND BALANCE ON SPINS
Both the total mass of the aircraft and its distribution influence the spin characteristics of the
aeroplane. Higher masses generally mean slower initial spin rates, but as the spin progresses,
spin rates may tend to increase. The higher angular momentum extends the time and altitude
necessary for recovery from a spin in a heavily loaded aeroplane.
CG location is even more significant, affecting the aeroplane's resistance to spin as well as all
phases of the spin itself.
a)
CG towards the forward limit makes an aircraft more stable, and control forces will be
higher which makes it less likely that large, abrupt control movements will be made.
When trimmed, the aeroplane will tend to return to level flight if the controls are
released, but the stall speed will be higher.
b)
CG towards the aft limit decreases longitudinal static stability and reduces pitch control
forces, which tends to make the aeroplane easier to stall. Once a spin is entered, the
further aft the CG, the flatter the spin attitude.
c)
If the CG is outside the aft limit, or if power is not reduced promptly, the spin is more
likely to go flat. A flat spin is characterised by a near level pitch and roll attitude with
the spin axis near the CG. Although the altitude lost in each tum of a flat spin may be
less than in a normal spin, the extreme yaw rate (often exceeding 400 per second)
results in a high descent rate. The relative airflow in a flat spin is nearly straight up,
keeping the wings at high angles of attack. More importantly, the upward flow over the
tail may render the elevator and rudder ineffective, making recovery impossible.
0
7.39
SPIN RECOVERY
Recovery from a simple stall is achieved by reducing the angle of attack which restores the
airflow over the wing; spin recovery additionally involves stopping the rotation. The extremely
complex aerodynamics of a spin may dictate vastly different recovery procedures for
different aeroplanes, so no universal spin recovery procedure can exist for all aeroplanes.
The recommended recovery procedure for some aeroplanes is simply to reduce power to idle and
release pressure on the controls. At the other extreme, the design of some aircraft is such that
recovery from a developed spin requires definite control movements, precisely timed to coincide
with certain points in the rotation, for several turns.
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STALLING
PRINCIPLES OF FLIGHT
The following is a general recovery procedure for erect spins. Always refer to the Flight Manual
for the particular aircraft being flown and follow the manufacturers recommendations.
1.
Move the throttle or throttles to idle. This minimises altitude loss and reduces the
possibility of a flat spin developing. It also eliminates possible asymmetric thrust in
multi-engine aeroplanes. Engine torque and gyroscopic propeller effect can increase the
angle of attack or the rate of rotation in single-engine aeroplanes, aggravating the spin.
2.
Neutralise the ailerons. Aileron position is often a contributory factor to flat spins, or
to higher rotation rates in normal spins.
3.
Apply full rudder against the spin. Spin direction is most reliably determined from the
tum co-ordinator. Do not use the ball in the slip indicator, its indications are not reliable
and may be affected by its location within the flight deck.
4.
Move the elevator control briskly to approximately the neutral position. Some aircraft
merely require a relaxation of back pressure, while others require full forward pitch
control travel.
The above four items can be accomplished simultaneously.
5.
Hold the recommended control positions until rotation stops.
6.
As rotation stops, neutralise the rudder. If rudder deflection is maintained after rotation
stops, the aircraft may enter a spin in the other direction!
7.
Recover from the resulting dive with gradual back pressure on the pitch control.
a)
Pulling too hard could trigger a secondary stall, or exceed the limit load factor
and damage the aircraft structure.
b)
Recovering too slowly from the dive could allow the aeroplane to exceed its
airspeed limits, particularly in aerodynamically clean aeroplanes.
Avoiding excessive speed build-up during recovery is another reason for closing the
throttles during spin recovery
c)
Add power as you resume normal flight, being careful to observe power and
rpm limitations.
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STALLING
PRINCIPLES OF FLIGHT
7.40
SPECIAL PHENOMENA OF STALL
CROSSED-CONTROL STALL
A crossed-control stall can occur when flying at high angles of attack while applying rudder in
the opposite direction to aileron, or too much rudder in the same direction as aileron. This will
be displayed by the ball in the slip indicator being displaced from neutral.
Crossed-control stalls can occur with little or no warning; one wing will stall a long time before
the other and a quite violent wing drop can occur. The "instinctive" reaction to stop the wing
drop with aileron must be resisted (ref. para 7.6). The rudder should be used to keep the aircraft
in balanced, co-ordinated flight at all times (ball in the middle), especially at low airspeedslhigh
angles of attack.
ACCELERATED STALL
An accelerated stall is caused by abrupt or excessive control movement. An accelerated stall can
occur during a sudden change in the flight path, during manoeuvres such as steep turns or a
rapid recovery from a dive. It is called an 'accelerated stall' because it occurs at a load factor
greater than 1g. An accelerated stall is usually more violent than a 1g stall, and is often
unexpected because of the relatively high airspeed.
SECONDARY STALL
A secondary stall may be triggered while attempting to recover from a stall. This usually
happens as a result of trying to hasten the stall recovery; either by not decreasing the angle of
attack enough at stall warning or by not allowing sufficient time for the aircraft to begin flying
again before attempting to regain lost altitude. With full power still applied, relax the back
pressure and allow the aeroplane to fly before reapplying moderate back pressure to regain lost
height.
LARGE AIRCRAFT
During airline 'type' conversion training on large aircraft, full stalls are not practised. To
familiarise pilots with the characteristics of their aircraft, only the approach to stall (stick shaker
activation) is carried out.
(a)
Jet Aircraft (swept wing): there are no special considerations during the approach to the
stall.
(i)
Power-Off stall: at stick shaker, smoothly lower the nose to the horizon, or just
below, to un-stall the wing; simultaneously increase power to the maximum
recommended to minimise height loss, prevent wing drop with roll control, raise
the gear and select take-off flaps.
(ii)
Power-On stall: as with power-off.
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STALLING
PRINCIPLES OF FLIGHT
(b)
Multi-engine propeller.
(i)
Power-Off stall: at stick shaker, smoothly lower the nose to the horizon, or just
below, to un-stall the wing; simultaneously increase power to the maximum
recommended to minimise height loss, prevent wing drop with rudder and
aileron control, raise the gear and select take-off flaps.
(ii)
Power-On stall: as with power-off.
The primary difference between j et and propeller aircraft is the rapidly changing propeller torque
and slipstream that will be evident during power application. It is essential for the pilot to
maintain co-ordination between rudder and aileron while applying the control inputs required
to counter the changing rolling and yawing moments generated by the propeller when the engine
is at high power settings or during rapid applications of power. Yaw must be prevented during
a stall and recovery.
SMALL AIRCRAFT
(c)
Single engine propeller
(i)
Power-Off stall: at stall warning, smoothly lower the nose to the horizon, or
just below, to un-stall the wing; simultaneously increase power to the maximum
recommended to minimise height loss, prevent wing drop with rudder and raise
the gear if applicable.
(ii)
Power-on stall and recovery in a single engine propeller aircraft has additional
complications. At the high nose attitude and low airspeed associated with a
power-on stall, there will be considerable "turning effects" from the propeller.
(These are fully detailed in Chapt. 16).
To maintain co-ordinated flight during the approach to, and recovery from, a power-on
stall, the pilot of a single engine propeller aircraft must compensate for the turning
effects of the propeller with the correct combination of rudder and aileron. It is essential
to maintain co-ordinated flight (ball in the middle) when close to the stall AND during
recovery. Any yawing tendency could easily develop into a spin. When the aircraft
nose drops at the stall, gyroscopic effect will also be apparent, increasing the nose left
yawing moment - with a clockwise rotating propeller.
An accidental power-on stall, during take-off or go-around, when a pilot's attention is
diverted, could easily tum into a spin. It is essential that correct stall recovery action is
taken at the first indication of a stall. (Forward movement of the pitch control; neutralise
the roll control; and prevent wing drop with the rudder).
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STALLING
PRINCIPLES OF FLIGHT
STALL AND RECOVERY IN A CLIMBING AND DESCENDING TURN
When an aircraft is in a level co-ordinated tum at a constant bank angle, the inside wing is
moving through the air more slowly than the outside wing and consequently generates less lift.
If the ailerons are held neutral, the aircraft has a tendency to continue to roll in the direction of
bank (over-banking tendency). Rather than return the ailerons to neutral when the required
degree of bank angle is reached, the pilot must hold aileron opposite to the direction of bank; the
lower the airspeed, the greater the aileron input required.
The inner (lower) wing may have a greater effective angle of attack due to the lowered aileron
and may reach the critical angle of attack first. The rudder must be used at all times to maintain
co-ordinated flight (ball in the middle).
In a climbing tum, airspeed will be lower and in a single engine propeller aircraft, the rolling
and yawing forces generated by the propeller and its slipstream will add their own requirements
for unusual rudder and aileron inputs. e.g. for an aircraft with a clockwise rotating propeller in
a climbing tum to the left at low speed it may be necessary for the pilot to be holding a lot of
right roll aileron and right rudder. If an aircraft in this situation were to stall, the gross control
deflections can make the aircraft yaw or roll violently. Correct co-ordination of the controls is
essential, in all phases of flight, to prevent the possibility of an accidental spin.
CONCLUSIONS
In whatever configuration, attitude, or power setting a stall warning occurs, the correct pilot
action is to decrease the angle of attack below the stall angle to un-stall the wing, apply
maximum allowable power to minimise altitude loss and prevent any yaw from developing to
minimise the possibility of spinning (pretty much, in that order). "Keep the ball in the middle".
7.41
HIGH SPEED BUFFET (SHOCK STALL)
When explaining the basic Principles of Flight, we consider air to be incompressible at speeds
less than four tenths the speed of sound (OAM). That is, pressure is considered to have no effect
on air density. At speeds higher than OAM it is no longer practical to make that assumption
because density changes in the airflow around the aircraft begin to make differences to the
behaviour of the aircraft.
SHOCKWAVE
SEPARATED AIRFLOW
Figure 7.30
Shock Induced Stall
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PRINCIPLES OF FLIGHT
STALLING
At high altitude, a large high speed jet transport aircraft will be cruising at a speed marginally
above its critical Mach number, and will have a small shock wave on the wing. If such an
aircraft over speeds the shock wave will rapidly grow larger, causing the static pressure to
increase sharply in the immediate vicinity of the shock wave. The locally increased adverse
pressure gradient will cause the boundary layer to separate immediately behind the shock wave,
Fig 7.30. This is called a 'shock stall'. The separated airflow will engulf the tail area in a very
active turbulent wake and cause severe airframe buffeting - a very undesirable phenomenon.
High speed buffet (shock stall) can seriously damage the aircraft structure, so an artificial
warning device is installed that will alert the pilot if the aircraft exceeds its maximum
operational speed limit (YMO IMMO)* by even a small margin. The high speed warning is aural
("clacker", hom or siren) and is easily distinguishable from the "low speed" high angle of attack
"stick shaker" warning.
We have seen that approaching the critical angle of attack can cause airframe buffeting ("low
speed" buffet) and we have now shown that flying too fast will also cause airframe buffeting
("high speed" buffet). ANY airframe buffeting is undesirable and can quickly lead to structural
damage, besides upsetting the passengers.
It will be shown that at high cruising altitudes (36,000 to 42,000 ft) the margin between the high
angle of attack stall warning and the high speed warning, may be as little as 15 kt.
*YMO is the maximum operating Indicated Air Speed,
MMO is the maximum operating Mach number. (These will be fully discussed in Chapter 14).
NB:
It is operationally necessary to fly as fast as economically possible and designers are constantly
trying to increase the maximum speed at which aircraft can fly, without experiencing any
undesirable characteristics. During certification flight testing the projected maximum speeds
are investigated and maximum operating speeds are established. The maximum operating
speed limit (YMO IM MO ) gives a speed margin into which the aircraft can momentarily overspeed
and be recovered by the pilot before any undesirable characteristics occur. (Tuck, loss of control
effectiveness, and several stability problems - these will all be detailed in later chapters).
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STALLING
PRINCIPLES OF FLIGHT
ANSWERS:
Stall speed in a 25° and 30° bank ifV s1g = 150 kt CAS. (with % comparisons)
25°
158 kt CAS (5% increase in stall speed above V S1g) [lift 10% greater]
30°
161 kt CAS (7% increase in stall speed above V S1g) [lift 15% greater]
45 °
178 kt CAS (190/0 increase in stall speed above V Slg) [lift 41 % greater]
60 ° = 212 kt CAS (41 % increase in stall speed above V S1g) [lift 100% greater]
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PRINCIPLES OF FLIGHT
STALLING
KEY FACTS 2 - Self Study
Insert the missing words in these statements, using the foregoing paragraphs for reference.
The swept-back wing is the major contributory factor to _ _ _ stall. (Para. 7.15).
An aircraft design with super stall tendencies must be fitted with a stick _ _ _ . (Para. 7.16).
Factors which can affect V SR are: (Para. 7.17).
a)
b)
c)
d)
e)
t)
g)
Changes in _ __
Manoeuvring the aircraft (increasing the _______ )
Configuration changes (changes in _ _ _ and _ _ _ _ moment)
Engine
and propeller _ _ _ _ __
_ _ _ number
Wing _ _ _ _ _ __
Heavy _ _ __
In straight and level flight the load factor is _ _ . (Para. 7.18).
At a higher weight, the stall speed of an aircraft will be _ _ _ _ (Para. 7.19)
If the weight is decreased by 50% the stall speed will _ _ _ _ _ by approximately _ _%.
(Para. 7.19).
Load factor varies with - - - - - -. (Para. 7.24)
The increase in stall speed in a turn is proportional to the square root of the _ _ _ _ _ __
(Para. 7.24).
High lift devices will _ _ _ _ the stall speed because CL MAX is
. (Para. 7.25)
Forward CG movement will _ _ _ _ stall speed due to the increased tail
(Para. 7.26).
load.
Lowering the landing gear will increase stall speed due to the increased tail - - - - load.
(Para. 7.27).
Increased engine power will decrease stall speed due to propeller _ _ _ _ and/or the
_ _ _ _ inclination of thrust. (Para. 7.28).
The effect of increasing Mach number on stall speed begin at
The effects of compressibility increases stall speed by decreasing
7 - 49
M. (Para. 7.29).
. (Para. 7.29)
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STALLING
PRINCIPLES OF FLIGHT
The formation of ice on the leading edge of the wing can _ _ _ _ _ stall speed by __ %.
(Para. 7.30).
Frost formation on the wing can _____ stall speed by _ _ %. (Para. 7.30).
An aircraft must be free of all _ _, _ _ and _ _ immediately before ___ . (Para. 7.30)
Airframe contamination
stall speed by reducing
, increasing the adverse
_ _ _ _ _ _ _ _ and/or reducing the
energy of the boundary layer. (Para. 7.30).
oscillations or
Indications of an icing-induced stall can be loss of aircraft _ _ _ _ _
_ _ drop and high rate of
. Artificial stall warning will be _ _ _, but aerodynamic
_ _ _ may assist in identifying the onset of wing stall. (Para. 7.31).
Very heavy _ _ can
the stall speed due to the film of water altering the _ _ _ __
contour of the wing. (Para. 7.33).
A
must occur before a spin can take place. (Para. 7.36).
In a steady spin, _ _ wings are stalled, one more than the other. (Para. 7.36).
A spin may also develop if forces on the aircraft are unbalanced in other ways, for example, from
yaw forces due to an
failure on a multi-engine aircraft, or if the _ is laterally displaced
by an unbalanced _ _ load. (Para. 7.36).
The following is a general recovery procedure for erect spins:- (Para. 7.39).
1.
2.
3.
4.
5.
6.
7.
Move the throttle or throttles to
the ailerons.
Apply full
against the spin.
Move the
control briskly to approximately the neutral position.
the recommended control positions until rotation stops.
As rotation stops, neutralise the _ __
Recover from the resulting dive with
back pressure on the
A crossed-control stall can be avoided by maintaining the _
(Para. 7.40).
control.
of the slip indicator in the _ __
A stall can occur at any ___ or flight _ _ _ _ if the ____ angle of attack is exceeded.
(Para. 7.2 and 7.40).
A secondary stall can be triggered either by not
the angle of
enough at stall
warning or by not allowing sufficient _ for the aircraft to begin _ _ again before attempting
to
lost altitude. (Para. 7.40).
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PRINCIPLES OF FLIGHT
STALLING
An added complication during an accidental stall and recovery of a single engine propeller
aircraft is due to the
and
forces generated by the
. It is essential to
maintain balanced, co-ordinated flight, particularly at _ _ airspeed, high angles of _ __
(Para. 7.40).
In whatever configuration, attitude, or power setting a stall warning occurs, the correct pilot
action is to
the angle of attack below the _ _ angle to un-stall the _ _ , apply
maximum allowable
to minimise altitude loss and prevent any _ _ from developing to
minimise the possibility of
. "Keep the _ _ in the middle". (Para. 7.40).
If a large shockwave forms on the wing, due to an inadvertent overspeed. The locally increased
_ _ _ pressure gradient will cause the
_ _ to separate immediately
the
shock wave. This is called '
stall'. (Para. 7.41).
KEY FACTS 2, WITH WORD INSERTS CAN BE FOUND ON PAGE 7 - 66
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STALLING
PRINCIPLES OF FLIGHT
SELF ASSESSMENT QUESTIONS
1.
An aeroplane will stall at the same:
a)
b)
c)
d)
2.
A typical stalling angle of attack for a wing without sweepback is:
a)
b)
c)
d)
3.
lift and drag will both decrease.
lift will decrease and drag will increase.
lift will increase and drag will decrease.
lift and drag will both increase.
The angle of attack at which an aeroplane stalls:
a)
b)
c)
d)
6.
remain the same.
decrease.
increase.
the position of the CG does not affect the stall speed.
If the angle of attack is increased above the stalling angle:
a)
b)
c)
d)
5.
4°
16°
30°
45°
If the aircraft weight is increased without change of C of G position, the stalling angle of attack
will:
a)
b)
c)
d)
4.
angle of attack and attitude with relation to the horizon
airspeed regardless of the attitude with relation to the horizon
angle of attack regardless of the attitude with relation to the horizon
indicated airspeed regardless of altitude, bank angle and load factor
will occur at smaller angles of attack flying downwind than when flying upwind
is dependent upon the speed of the airflow over the wing
is a function of speed and density altitude
will remain constant regardless of gross weight
An aircraft whose weight is 237402 N stalls at 132 kt. At a weight of356103 N it would stall
at:
a)
b)
c)
d)
88 kt
162 kt
108 kt
172 kt
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PRINCIPLES OF FLIGHT
7.
For an aircraft with a Ig stalling speed of60 kt lAS, the stalling speed in a steady 60° tum would
be:
a)
b)
c)
d)
8.
wing loading
lift/drag ratio
aspect ratio
load factor
The stalling speed of an aeroplane is most affected by:
a)
b)
c)
d)
12.
not affect the stalling speed.
cause the aircraft to stall at a higher speed and a higher angle of attack.
cause the aircraft to stall at a higher speed and a lower angle of attack.
cause the aircraft to stall at a lower speed.
Dividing lift by weight gives:
a)
b)
c)
d)
11.
the same as in level flight
at a lower speed than in level flight
at a higher speed than in level flight, and a lower angle of attack.
at a higher speed than in level flight and at the same angle of attack.
Formation of ice on the wing leading edge will:
a)
b)
c)
d)
10.
43 kt
60 kt
84 kt
120 kt
For an aircraft in a steady tum the stalling speed would be:
a)
b)
c)
d)
9.
STALLING
changes in air density
variations in aeroplane loading
variations in flight altitude
changes in pitch attitude
Stalling may be delayed to a higher angle of attack by:
a)
b)
c)
d)
increasing the adverse pressure gradient
increasing the surface roughness of the wing top surface
distortion of the leading edge by ice build-up
increasing the kinetic energy of the boundary layer
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STALLING
PRINCIPLES OF FLIGHT
13.
A stall inducer strip will:
a)
b)
c)
d)
14.
On a highly tapered wing without wing twist the stall will commence:
a)
b)
c)
d)
15.
to re-energise the boundary layer and prevent separation.
to control spanwise flow and delay tip stall.
to generate a vortex over the upper surface of the wing.
to maintain a laminar boundary layer.
A wing with washout would have:
a)
b)
c)
d)
18.
reduce induced drag at low speed.
increase the tendency to tip stall.
reduce the tendency to tip stall.
cause the stall to occur at a lower angle of attack.
The purpose of a boundary layer fence on a swept wing is:
a)
b)
c)
d)
17.
simultaneously across the whole span.
at the centre of the span.
at the root.
at the tip.
Sweepback on a wing will:
a)
b)
c)
d)
16.
cause the wing to stall first at the root
cause the wing to stall at the tip first
delay wing root stall
re-energise the boundary layer at the wing root
the tip chord less than the root chord.
the tip incidence less than the root incidence.
the tip incidence greater than the root incidence.
the tip camber less than the root camber.
On an untapered wing without twist the downwash:
a)
b)
c)
d)
increases from root to tip.
increases from tip to root.
is constant across the span.
is greatest at centre span, less at root and tip.
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PRINCIPLES OF FLIGHT
19.
A wing of constant thickness which is not swept back:
a)
b)
c)
d)
20.
to reduce induced drag
to reduce boundary layer separation
to induce a root stall
to counteract the effect of the wing-tip vortices.
A stick shaker is:
a)
b)
c)
d)
24.
wing root providing adequate stall warning
wingtip providing inadequate stall warning
wingtip providing adequate stall warning
leading edge, where the wing root joins the fuselage
V ortex generators are used:
a)
b)
c)
d)
23.
Increasing leading edge camber.
delaying separation.
Reducing the effective angle of attack.
Reducing span-wise flow.
A rectangular wing, when compared to other wing planforms, has a tendency to stall first at the:
a)
b)
c)
d)
22.
will stall at the tip first due to the increase in spanwise flow.
could drop a wing at the stall due to the lack of any particular stall inducing
characteristics.
will pitch nose down approaching the stall due to the forward movement of the centre
of pressure.
will stall evenly across the span.
Slots increase the stalling angle of attack by:
a)
b)
c)
d)
21.
STALLING
an overspeed warning device that operates at high Mach numbers.
an artificial stability device.
a device to vibrate the control column to give a stall warning.
a device to prevent a stall by giving a pitch down.
A stall warning device must be set to operate:
a)
b)
c)
d)
at the stalling speed.
at a speed just below the stalling speed.
at a speed about 5% to 10% above the stalling speed.
at a speed about 20% above the stalling speed.
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STALLING
PRINCIPLES OF FLIGHT
25.
Just before the stall the wing leading edge stagnation point is positioned:
a)
b)
c)
d)
26.
A wing mounted stall warning detector vane would be situated:
a)
b)
c)
d)
27.
a device to prevent an aircraft from stalling.
a type of trim system.
a device to assist the pilot to move the controls at high speed.
a device which automatically compensates for pitch changes at high speed.
In a developed spin:
a)
b)
c)
d)
30.
angle of attack only.
angle of attack, and in some systems rate of change of angle of attack.
airspeed only.
airspeed and sometimes rate of change of airspeed.
A stick pusher is:
a)
b)
c)
d)
29.
on the upper surface at about mid chord.
on the lower surface at about mid chord.
at the leading edge on the lower surface.
at the leading edge on the upper surface.
The input data to a stall warning device (e.g. stick shaker) system is:
a)
b)
c)
d)
28.
above the stall warning vane
below the stall warning vane
on top of the stall warning vane
on top of the leading edge because of the extremely high angle of attack
the angle of attack of both wings will be positive
the angle of attack of both wings will be negative
the angle of attack of one wing will be positive and the other will be negative
the down going wing will be stalled and the up going wing will not be stalled
To recover from a spin, the elevators should be:'
a)
b)
c)
d)
moved up to increase the angle of attack
moved down to reduce the angle of attack
set to neutral
allowed to float
7 - 57
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
31.
High speed buffet (shock stall) is caused by:
a)
b)
c)
d)
32.
Water increases the viscosity of air
Heavy rain can block the pitot tube, giving false airspeed indications
The extra weight and distortion of the aerodynamic surfaces by the film of water
The impact of heavy rain will slow the aircraft
If the tailplane is supplying a download and stalls due to contamination by ice:
a)
b)
c)
d)
35.
7%
30%
1.07%
15%
Heavy rain can increase the stall speed of an aircraft for which of the following reasons?
a)
b)
c)
d)
34.
the boundary layer separating in front of a shockwave at high angles of attack
the boundary layer separating immediately behind the shock wave
the shock wave striking the tail of the aircraft
the shock wave striking the fuselage
In a 30° bank level tum, the stall speed will be increased by:
a)
b)
c)
d)
33.
STALLING
the wing will stall and the aircraft will pitch-up due to the weight of the ice behind the
aircraft CO
the increased weight on the tailplane due to the ice formation will pitch the aircraft nose
up, which will stall the wing
because it was supplying a download the aircraft will pitch nose up
the aircraft will pitch nose down
Indications of an icing-induced stall can be:
1.
2.
3.
4.
5.
6.
An artificial stall warning device
Airspeed close to the normal stall speed
Violent roll oscillations
Airframe buffet
Violent wing drop
Extremely high rate of descent while in a 'normal' flight attitude
a)
b)
c)
d)
1,2,4and5
1,3 and 5
1,4 and 6
3,4, 5 and 6
7 - 58
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
36.
If a light single engine propeller aircraft is stalled, power-on, in a climbing tum to the left, which
of the following is the preferred recovery action?
a)
b)
c)
d)
37.
STALLING
elevator stick forward, ailerons stick neutral, rudder to prevent wing drop.
elevator stick neutral, rudder neutral, ailerons to prevent wing drop, power to idle.
elevator stick forward, ailerons and rudder to prevent wing drop.
elevator stick neutral, rudder neutral, ailerons stick neutral, power to idle.
If the stick shaker activates on a swept wing jet transport aircraft immediately after take-off
while turning, which of the following statements contains the preferred course of action?
a)
b)
c)
d)
Decrease the angle of attack
Increase thrust
Monitor the instruments to ensure it is not a spurious warning
Decrease the bank angle
7 - 59
© Oxford Aviation Services Limited
STALLING
PRINCIPLES OF FLIGHT
I No IAIBI c I D II
1
C
B
2
3
A
B
4
D
5
B
6
7
C
D
8
9
C
D
10
B
11
12
13
D
A
D
14
REF
II~.-'
A
I B I c I D II
REF
Page 7 - 12(c)
B
Para. 7.2
20
Para. 7.3
21
Para 7.19
22
Para. 7.2
23
C
Para. 7.10
Para. 7.19
24
C
Para. 7 .9(c)
Para. 7.19
25
Para. 7.24
26
Para. 7.24
27
Para. 7.30
28
A
Para. 7.16
Para. 7.24
29
A
Para. 7.35
Para. 7.24
30
B
Para. 7.39
Page 7 - 12(c)
31
B
Para. 7.41
Page 7 - 13(d)
32
Fig. 7.8
33
Para. 7.14
A
Page 7 - 14(e)
B
Fig. 7.4
B
Fig. 7.4
C
B
Para. 7.10
Para. 7.24
A
Para. 7.33
C
15
B
Page 7 - 15
34
D
Para. 7.32
16
B
Page 7 - 16
35
D
Para. 7.31
17
B
Page 7 - 12(a)
36
A
Para. 7.40
Fig. 7.7
37
A
Para. 7.40
18
19
A
B
I
Para. 7.14
7 - 61
© Oxford Aviation Services Limited
STALLING
PRINCIPLES OF FLIGHT
KEY FACTS 1 - Correct Statements
Stalling involves loss of height and loss of control
A pilot must be able to clearly and unmistakably identify a stall
A stall is caused by airflow separation
Separation can occur when either the boundary layer has insufficient kinetic energy or the
adverse pressure gradient becomes too great.
Adverse pressure gradient increases with increase in angle of attack
Alternative names for the angle of attack at which stall occurs are the stall angle and the critical
angle of attack
The coefficient of lift at which a stall occurs is C L MAX
A stall can occur at any airspeed or flight attitude
A typical stalling angle is approximately 16°
To recover from a stall the angle of attack must be decreased
Maximum power is applied during stall recovery to minimise height loss
On small aircraft, the rudder should be used to prevent wing drop at the stall
On swept wing aircraft the ailerons should be used to prevent wing drop at the stall
Recover height lost during stall recovery with moderate back pressure on the elevator control
The first indications of a stall may be unresponsive flight controls, stall warning device or
aerodynamic buffet
At speeds close to the stall, ailerons must be used with caution to lift a dropping wing
Acceptable indications of a stall are:
(1)
a nose down pitch that can not be readily arrested
(2)
severe buffeting
(3)
pitch control reaching aft stop and no further increase in pitch attitude occurs
7 - 63
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
STALLING
Reference stall speed (V SR) is a CAS defined by the aircraft manufacturer.
V SR may not be less than a Ig stall speed.
When a device that abruptly pushes the nose down at a selected angle of attack is installed, V SR
may not be less than 2 knots or 2 %, whichever is greater, above the speed at which the device
operates.
Stall warning with sufficient margin to prevent inadvertent stalling must be clear and
distinctive to the pilot in straight and turning flight
Acceptable stall warning may consist of the inherent aerodynamic qualities of the aeroplane or
by a device that will give clearly distinguishable indications under expected conditions of flight
Stall warning must begin at a speed exceeding the stall speed by not less than 5 knots or 5 %
CAS, whichever is the greater.
Artificial stall warning on a small aircraft is usually given by a horn or buzzer
Artificial stall warning on a large aircraft is usually given by a stick shaker, in conjunction with
lights and a noisemaker
An artificial stall warning device can be activated by a flapper switch, an angle of attack vane
or an angle of attack probe
Most angle of attack sensors compute the rate of change of angle of attack to give earlier
warning in the case of accelerated rates of stall approach
JAR required stall characteristics, up to the time the aeroplane is stalled, are:a)
b)
c)
d)
It must be possible to produce and correct yaw by unreversed use of the ailerons and
rudder
No abnormal nose up pitching may occur
Longitudinal control force must be positive
It must be possible to promptly prevent stalling and recover from a stall by normal use
of the controls
There should be no excessive roll between the stall and completion of recovery
For turning flight stalls, the action of the aeroplane after the stall may not be so violent
or extreme as to make it difficult, with normal piloting skill, to effect prompt recovery
and to regain control of the aeroplane
#
e)
f)
7 - 64
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
STALLING
An aero foil section with a small leading edge radius will stall at a smaller angle of attack and
the stall will be more sudden
An aerofoil section with a large thickness-chord ratio will stall at a higher angle of attack and
will stall more gently
An aerofoil section with camber near the leading edge will stall at a higher angle of attack
A rectangular wing plan form will tend to stall at the root first
A rectangular wing planform usually has ideal stall characteristics, these are:a)
b)
c)
d)
Aileron effectiveness at the stall
Nose drop at the stall
Aerodynamic buffet at the stall
Absence of violent wing drop at the stall
To give a wing with a tapered planform the desired stall characteristics, the following devices
can be included in the design:a)
b)
c)
d)
e)
Washout (decreasing incidence from root to tip)
An aero foil section with greater thickness and camber at the tip
Leading edge slots at the tip
Stall strips fitted to the wing inboard leading edge
Vortex generators which re-energise the boundary layer at the tip
A swept back wing has an increased tendency to tip stall due to the spanwise flow of boundary
layer from root to tip on the wing top surface. Methods of delaying tip stall on a swept wing
planform are:a)
b)
c)
Wing fences, thin metal fences which generally extend from the leading edge
to the trailing edge on the wing top surface
Vortilons, also thin metal fences, but smaller and are situated on the underside
of the wing leading edge
Saw tooth leading edge, generates vortices over wing top surface at high angles
of attack
Engine pylons of pod mounted wing engines also act as vortilons
Vortex generators are also used to delay tip stall on a swept wing
#
d)
e)
Tip stall on a swept wing planform gives a tendency for the aircraft to pitch-up at the stall. This
is due to the CP moving forwards when the wing tips stall first.
7 - 65
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
STALLING
KEY FACTS 2 - Correct Statements
The swept-back wing is the major contributory factor to super stall.
An aircraft design with super stall tendencies must be fitted with a stick pusher.
Factors which can affect V SR are:
a)
b)
c)
d)
e)
f)
g)
Changes in weight.
Manoeuvring the aircraft (increasing the load factor).
Configuration changes (changes in C L MAX and pitching moment).
Engine thrust and propeller slipstream
Mach number
Wing contamination
Heavy rain
In straight and level flight the load factor is one.
At a higher weight, the stall speed of an aircraft will be higher.
If the weight is decreased by 50% the stall speed will decrease by approximately 25%.
Load factor varies with bank angle.
The increase in stall speed in a tum is proportional to the square root of the load factor.
High lift devices will decrease the stall speed because C L MAX is increased.
Forward CG movement will increase stall speed due to the increased tail down load.
Lowering the landing gear will increase stall speed due to the increased tail down load.
Increased engine power will decrease stall speed due to propeller slipstream and!or the
upwards inclination of thrust.
The effect of increasing Mach number on stall speed begin at O·4M.
The effects of compressibility increases stall speed by decreasing C L MAX'
7 - 66
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
STALLING
The formation of ice on the leading edge of the wing can increase stall speed by 30%.
Frost formation on the wing can increase stall speed by 150/0.
An aircraft must be free of all snow, frost and ice immediately before flight.
Airframe contamination increases stall speed by reducing C L MAX' increasing the adverse
pressure gradient and/or reducing the kinetic energy of the boundary layer.
Indications of an icing-induced stall can be loss of aircraft performance, roll oscillations or
wing drop and high rate of descent. Artificial stall warning will be absent, but aerodynamic
buffet may assist in identifying the onset of wing stall.
Very heavy rain can increase the stall speed due to the film of water altering the aerodynamic
contour of the wing.
A stall must occur before a spin can take place.
In a steady spin, both wings are stalled, one more than the other.
A spin may also develop if forces on the aircraft are unbalanced in other ways, for example, from
yaw forces due to an engine failure on a multi-engine aircraft, or if the CG is laterally displaced
by an unbalanced fuel load.
The following is a general recovery procedure for erect spins:1.
2.
3.
4.
5.
6.
7.
Move the throttle or throttles to idle.
Neutralise the ailerons.
Apply full rudder against the spin.
Move the elevator control briskly to approximately the neutral position.
Hold the recommended control positions until rotation stops.
As rotation stops, neutralise the rudder.
Recover from the resulting dive with gradual back pressure on the pitch control.
A crossed-control stall can be avoided by maintaining the ball of the slip indicator in the middle.
A stall can occur at any speed or flight attitude if the critical angle of attack is exceeded.
A secondary stall can be triggered either by not decreasing the angle of attack enough at stall
warning or by not allowing sufficient time for the aircraft to begin flying again before
attempting to regain lost altitude.
7 - 67
© Oxford Aviation Services Limited
STALLING
PRINCIPLES OF FLIGHT
An added complication during an accidental stall and recovery of a single engine propeller
aircraft is due to the rolling and yawing forces generated by the propeller. It is essential to
maintain balanced, co-ordinated flight, particularly at low airspeed, high angles of attack.
In whatever configuration, attitude, or power setting a stall warning occurs, the correct pilot
action is to decrease the angle of attack below the stall angle to un-stall the wing, apply
maximum allowable power to minimise altitude loss and prevent any yaw from developing to
minimise the possibility of spinning. "Keep the ball in the middle".
If a large shockwave forms on the wing, due to an inadvertent overspeed. The locally increased
adverse pressure gradient will cause the boundary layer to separate immediately behind the
shock wave. This is called 'shock stall' .
7 - 68
© Oxford Aviation Services Limited
CHAPTER 8 - HIGH LIFT DEVICES
Contents
Page
PURPOSE OF HIGH LIFT DEVICES ........................................ 8 - 1
TAKE-OFF AND LANDING SPEEDS
C LMAX AUGMENTATION
FLAPS
TRAILING EDGE FLAPS
PLAIN FLAP
SPLIT FLAP ..................................................... 8 - 2
SLOTTED AND MULTIPLE SLOTTED FLAPS
FOWLER FLAP .................................................. 8 - 3
COMPARISON OF TRAILING EDGE FLAPS
C L MAX AND STALLING ANGLE ..................................... 8 - 4
DRAG .......................................................... 8 - 5
LIFT / DRAG RATIO .............................................. 8 - 6
PITCHING MOMENT
CENTRE OF PRESSURE MOVEMENT
CHANGE OF DOWNWASH
OVERALL PITCH CHANGE ........................................ 8 - 7
AIRCRAFT ATTITUDE WITH FLAPS LOWERED
LEADING EDGE HIGH LIFT DEVICES ..................................... 8 - 8
LEADING EDGE FLAPS
KRUEGER FLAP
V ARIABLE CAMBER LEADING EDGE FLAP ......................... 8 - 9
EFFECT OF LEADING EDGE FLAPS ON LIFT
LEADING EDGE SLOTS .......................................... 8 - 10
LEADING EDGE SLAT
AUTOMATIC SLOTS ............................................ 8 - 12
DISADVANTAGES OF THE SLOT
DRAG AND PITCHING MOMENT OF LEADING EDGE DEVICES
TRAILING EDGE PLUS LEADING EDGE DEVICES
SEQUENCE OF OPERATION ...................................... 8 - 13
ASYMMETRY OF HIGH LIFT DEVICES ................................... 8 - 14
FLAP LOAD RELIEF SYSTEM
CHOICE OF FLAP SETTING FOR TAKE-OFF CLIMB AND LANDING
MANAGEMENT OF HIGH LIFT DEVICES .' ................................ 8 - 16
FLAP RETRACTION AFTER TAKE OFF
FLAP EXTENSION PRIOR TO LANDING ........................... 8 - 18
SELF ASSESSMENT QUESTIONS ........................................ 8 - 19
ANSWERS ..................................................... 8 - 25
HIGH LIFT DEVICES
PRINCIPLES OF FLIGHT
8.1
PURPOSE OF HIGH LIFT DEVICES
Aircraft are fitted with high lift devices to reduce the take-off and landing distances. This
permits operation at greater weights from given runway lengths and enables greater payloads to
be carried.
8.2
TAKE-OFF AND LANDING SPEEDS
The take-off and landing distances depend on the speeds required at the screen, and these are laid
down in the performance regulations. For both take-off and landing, one of the requirements is
for a safe margin above the stalling speed (1.2 V S1 for take-off and 1.3 Vso for landing). The
stalling speed is determined by the C LMAX of the wing, and so to obtain the lowest possible
distances, the C LMAX , must be as high as possible.
8.3
C LMAX AUGMENTATION
One of the main factors which determine the C LMAX of an aerofoil section is the camber. It has
been shown (Pages 5-8 and 5-9) that increasing the camber of an aerofoil section increases the
C L at a given angle of attack and increases C LMAX . For take-off and landing a cambered section
is desirable, but this would give high drag at cruising speeds and require a very nose down
attitude. It is usual to select a less cambered aerofoil section to optimise cruise and modify the
section for take-off and landing by the use of flaps.
8.4
FLAPS
A flap is a hinged portion of the trailing or leading edge which can be deflected downwards and
so produce an increase of camber. For low speed aerofoils the flaps will be on the trailing edge
only, but on high speed aerofoils where the leading edge may be symmetrical or have a negative
camber, there will usually be flaps on both the leading edge and the trailing edge.
8.5
TRAILING EDGE FLAPS
The basic principle of the flap has been adapted in many ways. The more commonly used types
of trailing edge flap are considered below.
8.6
PLAIN FLAP
The plain flap , illustrated in Fig. 8.1, has a simple construction and gives a good increase in
C LMAX ' although with fairly high drag. It is used, mainly on low speed aircraft and where very
short take-off and landing is not required.
Figure 8.1
8-1
Plain Flap
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
8.7
HIGH LIFT DEVICES
SPLIT FLAP
The flap forms part of the lower surface of the wing trailing edge, the upper surface contour
being unaffected when the flap is lowered.
Figure 8.2
Split Flap
The split flap gives about the same increase in lift as the plain flap at low angles of attack, but
gives slightly more at higher angles as the upper surface camber is not increased and so
separation is delayed. The drag however is higher than for the plain flap due to the increased
depth of the wake.
8.8
SLOTTED AND MULTIPLE SLOTTED FLAPS
When the slotted flap is lowered a slot or gap is opened between the flap and the wing.
Figure 8.3
Slotted Flap
The purpose of the slot is to direct higher pressure air from the lower surface over the flap and
re-energise the boundary layer. This delays the separation of the airflow on the upper surface
of the flap. The slotted flap gives a bigger increase in C LMAX than the plain or split flap and much
less drag, but has a more complex construction.
8-2
© Oxford Aviation Services Limited
HIGH LIFT DEVICES
PRINCIPLES OF FLIGHT
8.9
THE FOWLER FLAP
The Fowler flap, Fig. 8.4, moves rearwards and then down, initially giving an increase in wing
area and then an increase in camber. The Fowler flap may be slotted.
Fowler Flap
Triple Slotted Fowler Flap
Figure 8.4
Because of the combined effects of increased area and camber, the Fowler flap gives the greatest
increase in lift of the flaps considered, and also gives the least drag because of the slot and the
reduction of thickness : chord ratio. However the change of pitching moment is greater because
of the rearward extension of the chord.
8.10
COMPARISON OF TRAILING EDGE FLAPS
SLOTTED FLAP
SPLIT FLAP
Figure 8.5
shows a
comparison ofthe lift curves
for the flaps considered
above, for the same angle of
flap deflection. It should be
noted however that the
different types of flap do not
all give their greatest
increase in lift at the same
flap angle.
~~--~~---------------.
U
Figure 8.5
8-3
© Oxford Aviation Services Limited
HIGH LIFT DEVICES
PRINCIPLES OF FLIGHT
Figure 8.6 shows the variation of the lift increment with increasing flap angle. It can be seen that
the increments are reducing as the flap angle increases, and that beyond some optimum flap
angle the increments decrease.
10"
FLAP ANGLE
Figure 8.6
8.11
C LMAX AND STALLING ANGLE
It can be seen from Figure 8.S that with the flap lowered C L MAX is increased, but the stalling
angle is reduced. This is because lowering the flap increases the effective angle of attack.
REDUCED ANGLE OF ATTACK
OF BASIC SECTION
EFFECTIVE
ANGLE OF ATTACK
~
•••
Figure 8.7
It is conventional to plot the CL
~
a curve using the angle of attack for the basic section.
Consequently, as shown in Fig. 8.7, at the stalling angle of attack for the section with flap
lowered, the basic wing section is at a reduced angle.
8-4
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
8.12
HIGH LIFT DEVICES
DRAG
Fig. 8.8 shows a comparison of the drag polar curves for the various types offlap. It can be seen
that for a given flap deflection the drag produced by the different types of flap varies
considerably, the split flap giving the highest drag and the Fowler flap the least.
FOWLER ~
,
SLOTTED -
-------~ BASIC SECTION
Figure 8.8
During take-off, drag reduces the acceleration, and so the flap should give as little drag as
possible. For landing however, drag adds to the braking force and so the flap drag is beneficial.
The addition of drag during approach also improves speed stability, see Para. 6.11.
As in the case of the lift increments, the drag increments with increasing flap angle are not
constant, the increments in drag get larger as the flap angle increases.
8-5
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
8.13
HIGH LIFT DEVICES
LIFT / DRAG RATIO
Lowering flap increases both the lift and the drag, but not in the same proportion. Although the
lift is the larger force , the proportional increase in the drag is greater, and so the maximum
obtainable lift / drag ratio decreases. The maximum lift / drag ratio occurs where the tangent
from the origin of the drag polar touches the curve, and the gradient of the tangent line is a
measure ofthe maximum lift / drag ratio (Figure 8.9).
RATIO
Figure 8.9
L / D Ratio
The lift / drag ratio is a measure of aerodynamic efficiency and affects the aircraft's performance
in areas such as range, climb angle and glide angle. With flaps lowered, range will be decreased,
climb angle reduced and glide angle increased.
8.14
PITCHING MOMENT
Flap movement, up or down, will usually cause a change of pitching moment. This is due to
Centre of Pressure (CP) movement and downwash at the tailplane.
8.15
CENTRE OF PRESSURE MOVEMENT
Moving a trailing edge flap will modify the prtlssure distribution over the whole chord of the
aerofoil, but the greatest changes will occur in the region of the flap . When flap is lowered, the
Centre of Pressure will move rearwards giving a nose down pitching moment, Fig. 8.1 Oa. In the
case of a Fowler flap, rearward movement of the flap will also cause the CP to move aft,
resulting in an even greater increase in the nose-down pitching moment.
8.16
CHANGE OF DOWNWASH
Tailplane effective angle of attack is determined by the downwash from the wing. If the flaps
are lowered the downwash will increase and the tailplane angle of attack will decrease, causing
a nose-up pitching moment, Fig. 8.10b.
8-6
© Oxford Aviation Services Limited
HIGH LIFT DEVICES
PRINCIPLES OF FLIGHT
WING
TAILPLANE
1
1
1
1
1
1
1
1
1
1
1
1
1
1
1
INCREASED
~ASH
1
1
1r~
1
1
1
1
1
NOSE DOWN
PITCHING MOMENT
__
_
__
_
_
NOSE UP
PITCHING MOMENT
1
__
_
I _ _ __ _ _ --- - - - J
1
Figure 8.10b
Figure 8.10a
8.17
OVERALL PITCH CHANGE
The resultant aircraft pitching moment will depend upon which of the two effects is dominant.
The pitching moment will be influenced by the type offlap, the position of the wing and relative
position of the tailplane, and may be nose-up, nose-down, or almost zero. For example, on flap
extension, a tailplane mounted on top of the fin will be less influenced by the change of
downwash, resulting in an increased aircraft nose down pitching moment.
8.18
AIRCRAFT ATTITUDE WITH FLAPS LOWERED
When the aircraft is in steady flight the lift must be equal to the weight. If the flaps are lowered
but the speed kept constant, lift will increase, and to maintain it at its original value the angle of
attack must be decreased. The aircraft will therefore fly in a more nose-down attitude if the flaps
are down. On the approach to landing this is an advantage as it gives better visibility of the
landing area.
8-7
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
8.19
HIGH LIFT DEVICES
LEADING EDGE HIGH LIFT DEVICES
There are two forms ofleading edge high lift device commonly in use, the leading edge flap and
the leading edge slot or slat.
8.20
LEADING EDGE FLAPS
On high speed aerofoil sections the leading edge may have very little camber and have a small
radius. This can give flow separation just aft of the leading edge at quite low angles of attack.
This can be remedied by utilising a leading edge flap which increases the leading edge camber.
Figure 8.11
Krueger flap
8.20.1 KRUEGER FLAP
The Krueger flap is part of the lower surface of the leading edge, which can be rotated
about its forward edge as shown in Fig. 8.11. To promote root stall on a swept wing,
Krueger flaps are used on the inboard section because they are less efficient than the
variable camber shown opposite.
8-8
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
HIGH LIFT DEVICES
RETRACTED
EXTENDED
Figure 8.12
Variable camber leading edge flap
8.20.2 VARIABLE CAMBER LEADING EDGE FLAP
To improve efficiency by giving a better leading edge profile, the camber of a leading
edge flap may be increased as it is deployed. Unlike trailing edge flaps which can be
selected to intermediate positions, leading edge devices are either fully extended
(deployed) or retracted (stowed).
8.21
EFFECT OF LEADING EDGE FLAPS ON LIFT
The main effect of the leading edge flap is to delay separation, so increase the stalling angle
and the corresponding CL MAX ' However there will be some increase of lift at lower angles of
attack due to the increased camber of the aerofoil section. Figure 8.13 shows the effect of these
flaps on the lift curve.
wrrH LEADING EDGE FLAP
Figure 8.13
8-9
© Oxford Aviation Services Limited
HIGH LIFT DEVICES
PRINCIPLES OF FLIGHT
8.22
LEADING EDGE SLOTS
A leading edge slot is a gap from the lower surface to the upper surface of the leading edge, and
may be fixed, or created by moving part of the leading edge (the slat) forwards.
C Lmax
WING
(Given Adverse Pressure Gradient)
SLAT
SLAT OPEN - Boundary Layer Re - Energised
(Same Adverse Pressure Gradient)
Figure 8.14
8.23
Leading edge slat
LEADING EDGE SLAT
A slat is a small auxiliary aerofoil attached to the leading edge of the wing, Fig. 8.14. When
deployed, the slat forms a slot which allows passage of air from the high pressure region below
the wing to the low pressure region above it. Additional Kinetic Energy is added to the airflow
through the slot by the slat forming a convergent duct.
When slats are deployed the boundary layer is re-energised
If Kinetic Energy is added to the boundary layer, boundary layer separation will be delayed to
a much higher angle of attack. At approximately 25 0 , the increased adverse pressure gradient
will once again overwhelm the Kinetic Energy of the boundary layer and separation will occur.
If the slot is permanently open, i.e. a fixed slot, the extra drag at high speed is an unnecessary
disadvantage, so most slats in commercial use are opened and closed by a control mechanism.
The slot can be closed for high speed flight and opened for low speeds, usually in conjunction
with the trailing edge flaps and actuated by the same selector on the flight deck.
8 - 10
© Oxford Aviation Services Limited
HIGH LIFT DEVICES
PRINCIPLES OF FLIGHT
The graph at Fig.8.1S shows the comparative figures for a slatted and un-slatted wing of the
same basic dimensions.
WING PLUS
/"-r ",\ SLATS
/
1.5
C
1.0
7:
/
m
L
I
/
I
I
I
WING
I
0.5
I
I
5
10
15
20 25
30
Angle of Attack
Figure 8.15
The effect of the slat is to prolong the lift curve by delaying boundary layer separation until a
higher angle of attack. When operating at high angles of attack the slat itself is generating a high
lift coefficient because of its marked camber. The action of the slat is to flatten the marked peak
of the low-pressure envelope at high angles of attack and to change it to one with a more gradual
pressure gradient. The flattening of the lift distribution envelope means that the boundary layer
does not undergo the sudden thickening that occurred through having to negotiate the very steep
adverse pressure gradient that existed immediately behind the former suction peak, and so it
retains much of its Kinetic Energy, thus enabling it to penetrate almost the full chord of the wing
before separating. Fig. 8.16 shows the alleviating effect of the slat on the low pressure peak and
that, although flatter, the area of the low pressure region, which is proportional to its strength,
is unchanged or even increased. The 'suction' peak does not move forward, so the effect of
the slot on pitching moment is insignificant.
With Slat
No Slat
Figure 8.16
8 - 11
© Oxford Aviation Services Limited
HIGH LIFT DEVICES
PRINCIPLES OF FLIGHT
8.24
AUTOMATIC SLOTS
On some aircraft the slots are not controlled by the pilot, but operate automatically. Their
movement is caused by the changes of pressure which occur around the leading edge as the angle
of attack increases. At low angles of attack the high pressures around the stagnation point keep
the slat in the closed position. At high angles of attack the stagnation point has moved
underneath the leading edge and 'suction' pressures occur on the upper surface of the slat. These
pressures cause the slat to move forward and create the slot.
This system is used mainly on small aircraft as a stall protection system. On larger aircraft the
position of the slats is selected when required by the pilot, their movement being controlled
electrically or hydraulically.
8.25
DISADVANTAGES OF THE SLOT
The slot can give increases in CLMAX of the same magnitude as the trailing edge flap, but whereas
the trailing edge flap gives its C LMAX at slightly less than the normal stalling angle, the slot
requires a much increased angle of attack to give its C LMAX • In flight this means that the aircraft
will have a very nose-up attitude at low speeds and on the approach to land, visibility of the
landing area could be restricted.
8.26
DRAG AND PITCHING MOMENT OF LEADING EDGE DEVICES
Compared to trailing edge flaps the changes of drag and pitching moment resulting from the
operation of leading edge devices are small.
8.27
TRAILING EDGE PLUS LEADING EDGE DEVICES
Most large transport aircraft employ both trailing edge and leading edge devices. Fig. 8.17
shows the effect on the lift curve of both types of device.
8 - 12
© Oxford Aviation Services Limited
HIGH LIFT DEVICES
PRINCIPLES OF FLIGHT
t
TRAILING EDGE FLAP
~SIC
SECTION
Figure 8.17
8.28
SEQUENCE OF OPERATION
For some aerofoils the sequence of flap operation is critical. Lowering a trailing edge flap
increases both the downwash and the upwash. For a high speed aerofoil, an increase of up wash
at the leading edge when the angle of attack is already fairly high, could cause the wing to stall.
The leading edge device must therefore be deployed before the trailing edge flap is lowered.
When the flaps are retracted the trailing edge flap must be retracted before the leading edge
device is raised.
8 - 13
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
8.29
HIGH LIFT DEVICES
ASYMMETRY OF HIGH LIFT DEVICES
Deployment of high lift devices can produce large changes of lift, drag and pitching moment.
If the movement of the devices is not symmetrical on the two wings, the unbalanced forces could
cause severe roll control problems. On many flap control systems the deflection on the two sides
is compared while the flaps are moving, and if one side should fail, movement on the other side
is automatically stopped. However on less sophisticated systems, a failure of the operating
mechanism could lead to an asymmetric situation. The difference in lift will cause a rolling
moment which must be opposed by the ailerons, and the difference in drag will cause a yawing
moment which must be opposed by the rudder. Whether the controls will be adequate to
maintain straight and level flight will depend on the degree of asymmetry and the control power
available.
8.30
FLAP LOAD RELIEF SYSTEM
On large high speedjet transport aircraft, a device is fitted in the flap operating system to prevent
the flaps deploying if the aircraft speed is too high. The pilot can select the flaps, but they will
not extend until the airspeed is below the flap extend speed (VFE)' If a selection is made, and the
flaps do not run because the speed is too high, they will extend as soon as the airspeed decreases
to an appropriate value.
8.31
CHOICE OF FLAP SETTING FOR TAKE-OFF, CLIMB AND LANDING
1. TAKEOFF
Takeoff distance depends upon unstick speed and rate of acceleration to that speed.
a)
Lowest unstick speed will be possible at the highest CLMAX and this will be achieved at
a large flap angle, Fig. 8.18.
b)
But large flap angles also give high drag, Fig. 8.19, which will reduce acceleration and
increase the distance required to accelerate to unstick speed.
c)
A lower flap angle will give a higher un stick speed, but better acceleration, and so give
a shorter distance to unstick.
Thus there will be some optimum setting which will give the shortest possible take-off distance.
Ifleading edge devices are fitted they will be used for take-off as they increase the CLMAX for any
trailing edge flap setting.
2. CLIMB
After take-off, a minimum climb gradient is required in the take-off configuration. Climb
gradient is reduced by flap, so if climb gradient is limiting, a lesser flap angle may be selected
even though it gives a longer take-off distance.
3. LANDING
Landing distance will depend on touchdown speed and deceleration. The lowest touchdown
speed will be given by the highest CL MAX, obtained at a large flap angle, Fig. 8.18. Large flap
angle will also give high drag, Fig. 8.19, and so good deceleration. For landing, a large flap
angle will be used. Leading edge devices will also be used to obtain the highest possible C LMAX '
8 - 14
© Oxford Aviation Services Limited
HIGH LIFT DEVICES
PRINCIPLES OF FLIGHT
FLAPS UP
ANGLE OF ATTACK
Figure 8.18
FLAPS UP
LANDING
TAKEOFF
ANGLE OF ATTACK
Figure 8.19
8 - 15
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
8.32
HIGH LIFT DEVICES
MANAGEMENT OF HIGH LIFT DEVICES
To take full advantage of the capabilities of flaps the flight crew must properly manage their
retraction and extension.
FLAP RETRACTION AFTER TAKE OFF
With reference to Fig. 8.20, assume the aircraft has just taken offwith flaps extended and is at
point 'A' on the lift curve. If the flaps are retracted, with no change made to either angle of attack
or lAS, the coefficient of lift will reduce to point 'C' and the aircraft will sink.
1.
From point 'A' on the lift curve the aircraft should be accelerated to point 'B.
2.
From point' B' , as the flaps are retracted the angle of attack should be increased to point
'C' to maintain the coefficient of lift constant.
The pilot should not retract the flaps until the aircraft has sufficient lAS. Of course, this same
factor must be considered for any intermediate flap position between extended and retracted.
(Refer to Page 5-6 for a review of the Interpretation of the Lift Curve if necessary.)
As the configuration is altered from the flaps down to the flaps up or "clean" configuration, three
important changes take place:
a)
Changes of pressure distribution on the wing generates a nose up pitching moment. But
reduced wing downwash increasing the tailplane effective angle of attack generates a
nose down pitching moment. The resultant, actual, pitching moment experienced by the
aircraft will depend upon which of these two pitching moments is dominant.
(Ref. Para. 8.15)
b)
With reference to Fig. 8.21, the retraction of flaps ('B' to 'C') causes a reduction of drag
coefficient. This drag reduction improves the acceleration of the aircraft.
c)
Flap retraction usually takes place in stages and movement of the flaps between stages
will take a finite period of time. It has been stated that as flaps are retracted, an increase
in angle of attack is required to maintain the same lift coefficient.
If aircraft acceleration is low throughout the flap retraction speed range, the angle of
attack must be increased an appreciable amount to prevent the aircraft from sinking.
This situation is typical after take offw.hen gross weight and density altitude are high.
However, most modem jet transport aircraft have enough acceleration throughout the
flap retraction speed range that the resultant rapid gain in airspeed requires a much less
noticeable increase in angle of attack.
8 - 16
© Oxford Aviation Services Limited
HIGH LIFT DEVICES
PRINCIPLES OF FLIGHT
FLAPS EXTENDED
FLAPS RETRACTED
ANGLE OF ATTACK
Figure 8.20
Figure 8.21
8 - 17
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
HIGH LIFT DEVICES
FLAP EXTENSION PRIOR TO LANDING
With reference to Fig. 8.22, assume the aircraft is in level flight in the terminal area prior to
landing and is at point 'A' on the lift curve. If the flaps are extended, with no change made to
angle of attack, the coefficient oflift will increase to point 'C' and the aircraft will gain altitude
(balloon).
1.
From point 'A', as the flaps are extended the angle of attack should be decreased to
point 'B' to maintain the coefficient of lift constant.
2.
From point 'B' on the lift curve the aircraft should be decelerated to point ' C' .
(Refer to Page 5-6 for a review of the Interpretation of the Lift Curve if necessary.)
FLAPS EXTENDED
FLAPS RETRACTED
ANGLE OF ATTACK
Figure 8.22 Deployment of flaps for landing
8 - 18
IC Oxford Aviation Services Limited
HIGH LIFT DEVICES
PRINCIPLES OF FLIGHT
SELF ASSESSMENT QUESTIONS
1.
With the flaps lowered, the stalling speed will:
a)
b)
c)
d)
2.
When flaps are lowered the stalling angle of attack of the wing:
a)
b)
c)
d)
3.
Increase
decrease
remain the same but will occur at a higher angle of attack.
remain the same but will occur at a lower angle of attack.
The purpose of a leading edge droop is:
a)
b)
c)
d)
6.
increases and the stalling angle increases
decreases and the stalling speed decreases
remains the same and the stalling angle remains the same
remains the same and the stalling angle decreases
When a leading edge slot is opened, the stalling speed will:
a)
b)
c)
d)
5.
remains the same, but C L max increases.
increases and C L max increases.
decreases, but C L max increases.
decreases, but C L max remains the same.
With full flap, the maximum Lift/drag ratio:
a)
b)
c)
d)
4.
increase.
decrease.
increase, but occur at a higher angle of attack.
remain the same.
to give a more cambered section for high speed flight.
to increase the wing area for take-off and landing.
to increase wing camber, and delay separation of the airflow when trailing edge flaps
are lowered.
to decrease the lift during the landing run.
Lowering flaps sometimes produces a pitch moment change due to:
a)
b)
c)
d)
decrease of the angle of incidence.
movement of the centre of pressure.
movement of the centre of gravity.
increased angle of attack of the tailplane.
8 - 19
© Oxford Aviation Services Limited
HIGH LIFT DEVICES
PRINCIPLES OF FLIGHT
7.
Which type of flap would give the greatest change in pitching moment?
a)
b)
c)
d)
8.
A split flap is:
a)
b)
c)
d)
9.
the
the
the
the
lift would not change until the aircraft is airborne.
lift would increase when the flaps are lowered.
lift would decrease.
acceleration would increase.
When flaps are lowered the spanwise flow on the upper surface of the wing:
a)
b)
c)
d)
12.
must be reduced.
must be increased.
must be kept constant but power must be increased.
must be kept constant and power required will be constant.
If flaps are lowered during the take-off run:
a)
b)
c)
d)
11.
a flap divided into sections which open to form slots through the flap.
a flap manufactured in several sections to allow for wing flexing.
a flap which can move up or down from the neutral position.
a flap where the upper surface contour of the wing trailing edge is fixed and only the
lower surface contour is altered when the flaps are lowered
If the flaps are lowered in flight, with the airspeed kept constant, to maintain level flight the
angle of attack:
a)
b)
c)
d)
10.
Split
Plain
Fowler
Plain slotted
does not change.
increase towards the tip.
increases towards the root.
increases in speed but has no change of direction.
If a landing is to be made without flaps the landing speed must be:
a)
b)
c)
d)
reduced.
increased.
the same as for a landing with flaps.
the same as for a landing with flaps but with a steeper approach.
8 - 20
© Oxford Aviation Services Limited
HIGH LIFT DEVICES
PRINCIPLES OF FLIGHT
13.
Lowering the flaps during a landing approach:
a)
b)
c)
d)
14.
With reference to Annex A, the type of flap illustrated is a:
a)
b)
c)
d)
15.
increases the angle of descent without increasing the airspeed
decreases the angle of descent without increasing power
eliminates floating
permits approaches at a higher indicated airspeed
Slotted Krueger flap
Slotted Variable camber flap
Slotted Slat
Slotted Fowler flap
With reference to Annex F , the type of flap illustrated is a:
a)
b)
c)
d)
Slat
Fowler flap
Krueger flap
Variable camber flap
8 - 21
© Oxford Aviation Services Limited
HIGH LIFT DEVICES
PRINCIPLES OF FLIGHT
Annex A
Annex F
8 - 23
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
HIGH LIFT DEVICES
ANSWERS
I No I A I B I c I DII
1
C
3
B
4
B
5
C
B
6
7
C
D
8
A
10
B
11
C
12
13
I
B
2
9
REF
B
A
14
D
15
D
8 - 25
© Oxford Aviation Services Limited
CHAPTER 9 - AIRFRAME CONTAMINATION
Contents
Page
INTRODUCTION .......................................................
TYPES OF CONTAMINATION
EFFECT OF FROST AND ICE ON THE AIRCRAFT
EFFECT ON INSTRUMENTS .............................................
EFFECT ON CONTROLS
WATER CONTAMINATION
AIRFRAME AGING
SELF ASSESSMENT QUESTIONS .........................................
ANSWERS ......................................................
9-1
9-2
9-3
9-5
AIRFRAME CONTAMINATION
PRINCIPLES OF FLIGHT
9.1
INTRODUCTION
The airframe may become contaminated by ice, frost or water either whilst it is flight, or when
standing on the ground.
The meteorological conditions that cause ice and frost to form are dealt with elsewhere, but the
effect is an accumulation of ice or frost on the surface of the aircraft which will affect its
performance and handling.
9.2
TYPES OF CONTAMINATION
a)
Frost. Frost can form on the surface of the aircraft either when it is standing on the
ground when the temperature falls below O°C, or in flight, if the aircraft, after flying in
a region where the temperature is below O°C, moves into a warmer layer of air. It
consists of a fairly thin coating of crystalline ice.
b)
Ice. The main forms of icing are clear ice, rime ice and rain ice. Clear ice (glaze ice)
is a translucent layer of ice with a smooth surface, caused by large super cooled water
droplets, striking the leading edges of the airframe. As there is some delay in freezing,
there is some flow back along the surface behind the leading edge.
Rime ice forms when small supercooled water droplets strike the leading edges and
freeze almost immediately so that there is no flow back. It is a white opaque formation.
Rain ice is caused by rain which becomes supercooled by falling from an inversion into
air which is below O°C. It does not freeze immediately and forms considerable flow
back, and builds up very quickly.
9.3
EFFECT OF FROST AND ICE ON THE AIRCRAFT
The formation of ice and frost on the airframe will:
a)
modify the profile of the aerofoil
b)
increase the roughness of the aircraft surface
c)
increase the weight of the aircraft
The main effect of frost will be to increase the surface roughness and this will increase the
energy loss in the boundary layer. The skin friction drag will increase and the boundary layer
will have an earlier separation, giving a reduced C LMAX • Take-offwith frost on the wings could
result in a stall after lift off if the normal take-off speed is used.
Tests have shown that frost, ice or snow with the thickness and surface roughness of
medium or coarse sandpaper, reduces lift by as much as 300/0 and increases drag by 40%
Ice will normally form on and behind the leading edges of wings and tailplane and can result in
severe distortion of the leading edge profile. This will give a large increase in drag and a
substantial decrease in C L MAX.
9-1
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
AIRFRAME CONTAMINATION
The reduced CL MAX of the wing will give a higher stalling speed and the decreased CL MAX of the
tailplane could cause it to stall when the aircraft is flying at low speed, particularly if the wing
downwash is increased as a result of flap extension.
Tailplane stall will result in loss of longitudinal control. Clear ice and rain ice especially can
add considerable weight to the airframe, and this will in tum give a higher stalling speed, as well
as increased induced drag. The margin of thrust to drag will be decreased, reducing the ability
to climb. Increased power will be required to maintain height, resulting in increased fuel
consumption.
Ice formation on propeller blades can upset the balance of the propeller and cause severe
vibration, particularly if pieces of ice break off from one blade. Pieces of ice shed from
propellers can also cause damage to the fuselage.
9.4
EFFECT ON INSTRUMENTS
Formation of ice on static vents and pitot heads could cause errors in the readings of pressure
instruments, and eventually, failure to show any reading.
9.5
EFFECT ON CONTROLS
Any moveable surface could become jammed by ice forming in the gaps around the control, or
by pieces of ice breaking off and becoming jammed in the control gaps. The controls could
become difficult to operate or immovable.
9.6
WATER CONTAMINATION
If the wings are contaminated with water due to heavy rain, the boundary layer may become
turbulent further forward on the wing, particularly if the section is ofthe laminar flow type. This
will cause increased drag and may disrupt the boundary layer resulting in a significantly higher
stall speed.
Adjustments to operational speed should be made in accordance with the
recommendations of the aircraft manufacturer or aircraft operator, when taking-off and
landing in heavy rain.
9.7
AIRFRAME AGING
Over a period of years the condition of the airframe will deteriorate due to small scratches, minor
damage, repairs, and general accumulation of dirt and grease.
The overall effect of this will be to increase the drag of the aircraft (mainly skin friction drag)
with a consequent increase in fuel consumption. The cost of operating the aircraft will therefore
increase with the age of the airframe. The normal deterioration of the airframe is allowed for
in the performance charts of the aeroplane.
9-2
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
AIRFRAME CONTAMINATION
SELF ASSESSMENT QUESTIONS
1.
After an aircraft has been exposed to severe weather:
a)
b)
c)
d)
2.
Icing conditions may be encountered in the atmosphere when:
a)
b)
c)
d)
3.
Increased angle of attack for stalls.
Increased stall speed.
Increased pitch down tendencies.
Decreased speed for stalling
Frost covering the upper surface of an aircraft wing will usually cause:
a)
b)
c)
d)
5.
relative humidity is low and temperature rises.
pressure is high and humidity falls.
relative humidity is high and temperature is low.
relative pressure is high and temperature is high.
Which is an effect of ice, snow, or frost formation on an aeroplane?
a)
b)
c)
d)
4.
snow should be removed but smooth ice may be left.
all snow and ice should be removed.
loose snow may be left but ice must be removed.
providing the contamination is not too thick, it may be left in place.
the aircraft to stall at an angle of attack that is lower than normal
no problems to pilots
drag factors so large that sufficient speed cannot be obtained for take-off
the aircraft to stall at an angle of attack that is higher than normal
If it is suspected that ice may have formed on the tailplane and longitudinal control difficulties
are experienced following flap selection, the prudent action to take would be:
a)
b)
c)
d)
immediately decrease the flap setting
allow the speed to increase
select a greater flap deflection because this will increase CL max
reduce the angle of attack
9-3
© Oxford Aviation Services Limited
AIRFRAME CONTAMINATION
PRINCIPLES OF FLIGHT
6.
When considering in-flight airframe contamination with frost or ice, which of the following
statements is correct?
a)
b)
c)
d)
7.
Build-up can be identified by the ice detection equipment fitted to the aircraft.
The pilot can visually identify build-up on the wings, tailplane or flight controls by
looking through the flight deck windows; at night by using the ice detection lights.
Visual evidence of the accumulation of airframe icing may not exist.
Due to the high speed of modem aircraft, significant airframe contamination with frost,
ice or snow will not occur.
In the event of an icing-induced wing stall, which of the following indications will reliably be
available to the flight crew?
1 - Activation of the stall warning device (hom or stick shaker).
2 - The aircraft pitching nose down.
3 - Loss of elevator effectiveness.
4 - Airframe buffet.
5 - A roll control problem (increasing roll oscillation or violent wing drop).
6 - A high rate of descent.
a)
b)
c)
d)
1,2,3,4, 5 and 6
1,3 and 4
1,4 and 6
4,5 and 6
9-4
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
AIRFRAME CONTAMINATION
ANSWERS
I No I A I B I c I DII
2
C
B
3
4
A
5
A
7
I
B
1
6
REF
C
D
9-5
© Oxford Aviation Services Limited
CHAPTER 10 - STABILITY AND CONTROL
Contents
Page
INTRODUCTION ...................................................... 10 - 1
STATIC STABILITY
AEROPLANE REFERENCE AXES .................................. 10 - 4
STATIC LONGITUDINAL STABILITy .............................. 10 - 5
NEUTRAL POINT ............................................... 10 - 9
STATIC MARGIN .............................................. 10 - 10
TRIM AND CONTROLLABILITY ................................. 10 - 11
KEY FACTS 1 (SELF STUDY) .................................... 10 - 14
GRAPHIC PRESENTATION ...................................... 10 - 17
CONTRIBUTION OF THE COMPONENT SURFACES ................ 10 - 20
WING
FUSELAGE AND NACELLES .............................. 10 - 22
HORIZONTAL TAIL
LONGITUDINAL DIHEDRAL ....................... 10 - 23
DOWNWASH ..................................... I0-24
POWER-OFF STABILITY ........................................ 10 - 25
EFFECT OF CG POSITION ................................ 10 - 26
POWER EFFECTS ........................................ 10 - 27
HIGH LIFT DEVICES ..................................... 10 - 29
CONTROL FORCE STABILITY ................................... 10 - 30
MANOEUVRE STABILITY ...................................... 10 - 35
STICK FORCE PER 'G' ................................... 10 - 36
TAILORING CONTROL FORCES ................................. 10 - 38
STICK CENTRING SPRING
DOWN SPRING
BOBWEIGHT ............................................ 10 - 39
LONGITUDINAL CONTROL ..................................... 10 - 40
MANOEUVRING CONTROL REQUIREMENTS
TAKE OFF CONTROL REQUIREMENTS ........................... 10 - 41
LANDING CONTROL REQUIREMENTS ........................... 10 - 42
DYNAMIC STABILITY ................................................ 10 - 43
LONGITUDINAL DYNAMIC STABILITY .......................... 10 - 47
LONG PERIOD OSCILLATION (PHUGOID) .................. 10 - 48
SHORT PERIOD OSCILLATION ............................ 10 - 49
DIRECTIONAL STABILITY AND CONTROL ..............................
SIDESLIP ANGLE ..............................................
STATIC DIRECTIONAL STABILITY ..............................
CONTRIBUTION OF THE AEROPLANE COMPONENTS .............
FUSELAGE
DORSAL AND VENTRAL FINS ............................
FIN ....................................................
WING AND NACELLES ...................................
POWER EFFECTS ........................................
CRITICAL CONDITION
CG POSITION
HIGH ANGLE OF ATTACK ................................
VENTRAL FIN
LATERAL STABILITY AND CONTROL ..................................
STATIC LATERAL STABILITY ...................................
CONTRIBUTION OF THE AEROPLANE COMPONENTS .............
WING
WING POSITION .........................................
SWEEPBACK ...........................................
FIN ....................................................
PARTIAL SPAN FLAPS ...................................
LATERAL DYNAMIC EFFECTS .........................................
SPIRAL DIVERGENCE
DUTCH ROLL
PILOT INDUCED OSCILLATIONS (PIO) ..................................
HIGH MACH NUMBERS ...............................................
MACH TRIM
KEY FACTS 2 (SELF STUDY) ..........................................
SUMMARY ..........................................................
SELF ASSESSMENT QUESTIONS .......................................
ANSWERS ....................................................
KEY FACTS 1 (ANSWERS) ............ ; ...............................
KEY FACTS 2 (ANSWERS) ............................................
10 - 51
10 - 52
10 - 53
10 - 54
10 - 55
10-57
10 - 58
10 - 60
10 - 61
10 - 62
10 - 63
10 - 65
10 - 66
10 - 67
10 - 68
10 - 69
10 - 70
10 -71
10 - 72
10 - 73
10 -77
10 - 81
10 - 87
10 - 89
10 - 92
STABILITY AND CONTROL
PRINCIPLES OF FLIGHT
10.1
INTRODUCTION
Stability is the tendency of an aircraft to return to a steady state of flight without any help from
the pilot, after being disturbed by an external force.
An aircraft must have the following qualities:
a)
Adequate stability to maintain a uniform flight condition.
b)
The ability to recover from various disturbing influences.
c)
Sufficient stability to minimise the workload of the pilot and
d)
Proper response to the controls so that it may achieve its design performance
with adequate manoeuvrability.
There are two broad categories of stability, static and dynamic. Dynamic stability will be
considered later.
10.2
STATIC STABILITY
An aircraft is in a state of equilibrium (trim) when the sum of all forces is zero and the sum of
all moments is zero; there are no accelerations and the aircraft will continue in steady flight. If
equilibrium is disturbed by a gust, or deflection of the controls, the aircraft will experience
accelerations due to an unbalance of moments or forces.
The type of static stability an aircraft possesses is defined by its initial tendency, following the
removal of some disturbing force.
1 - Positive static stability (or static stability) exists if an aircraft is disturbed from
equilibrium and has the tendency to return to equilibrium.
2 - Neutral static stability exists if an aircraft is subject to a disturbance and has neither
the tendency to return nor the tendency to continue in the displacement direction.
3 - Negative static stability (or static instability) exists if an aircraft has a tendency to
continue in the direction of disturbance.
Examples of the three types of static stability are shown in Fig's. 10.1, 10.2 and 10.3
10- 1
© Oxford Aviation Services. Limited
STABILITY AND CONTROL
PRINCIPLES OF FLIGHT
Fig. 10.1, illustrates the condition of positive static stability (or static stability). The ball is
displaced from equilibrium at the bottom of the trough. When the disturbing force is removed,
the initial tendency of the ball is to return towards the equilibrium condition. The ball may roll
back and forth through the point of equilibrium but displacement to either side creates the initial
tendency to return.
POSITIVE STATIC STABILITY
Tendency to Return
---- to Equilibrium
Figure 10.1
Fig. 10.2, illustrates the condition of neutral static stability. The ball encounters a new
equilibrium at any point of displacement and has no tendency to return to its original
equilibrium.
Displacement
,
\,
NEUTRAL STATIC STABILITY
Figure 10.2
10 - 2
© Oxford Aviation Services Limited
STABILITY AND CONTROL
PRINCIPLES OF FLIGHT
Fig. 10.3, illustrates the condition of negative static stability (or static instability).
Displacement from equilibrium at the hilltop gives a tendency for greater displacement.
Tendency to Continue
Direction
\
NEGATIVE STATIC STABILITY
Figure 10.3
The term "static" is applied to this form of stabi lity since any resulting motion is not considered.
Only the initial tendency to return to equilibrium is considered in static stability.
The static longitudinal stability of an aircraft is assessed by it being displaced from some
trimmed angle of attack.
Ifthe aerodynamic pitching moments created by this displacement tend to return the aircraft to
the equilibrium angle of attack the aircraft has positive static longitudinal stability.
10 - 3
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
10.3
STABILITY AND CONTROL
AEROPLANE REFERENCE AXES
In order to visualise the forces and moments on the aircraft, it is necessary to establish a set of
reference axes passing through the centre of gravity. Figure 10.4 illustrates a conventional
right hand axis system.
The longitudinal axis passes through the CG from nose to tail. A moment about this axis is a
rolling moment, L, a roll to the right is a positive rolling moment.
The normal axis passes "vertically" through the CG at 90 ° to the longitudinal axis. A moment
about the normal axis is a yawing moment, N, and a positive yawing moment would yaw the
aircraft to the right.
The lateral axis is a line passing through the CG, parallel to a line passing through the wing tips.
A moment about the lateral axis is a pitching moment, M, and a positive pitching moment is nose
up.
Lateral Axis
o. sitive
Pitching
Moment,
~ I'
~
'~,~
J
Centre Of
Gravity
I
2i
~,
~
~i
rt.~
\
Longitudinal
Axis
------7
~
~------~~
~ Positive
~\)
~
~.
I
~-
'. J
D~II;:ositive
Yawin
Moment,
N
Rolling Moment,
L
Normal Axis
Figure 10.4
10 - 4
© Oxford Aviation Services Limited
STABILITY AND CONTROL
PRINCIPLES OF FLIGHT
10.4
STATIC LONGITUDINAL STABILITY
Longitudinal stability is motion about the lateral axis. To avoid confusion, consider the axis
about which the particular type of stability takes place. Thus, lateral stability is about the
longitudinal axis (rolling), directional stability is about the normal axis (yawing) and
longitudinal stability is about the lateral axis (pitching). Static longitudinal stability is
considered first because it can be studied in isolation; in general it does not interact with motions
about the other two axes. Lateral and directional stability tend to interact (coupled motion), and
these will be studied later.
a)
An aircraft will exhibit static longitudinal stability ifit tends to return towards the trim
angle of attack when displaced by a gust OR a control input.
(i)
It is essential that an aircraft has positive static longitudinal stability. If it is
stable, an aeroplane is safe and easy to fly since it seeks and tends to maintain
a trimmed condition of flight. It also follows that control deflections and
control "feel" (stick force) must be logical, both in direction and magnitude.
b)
If the aircraft is neutrally stable, it tends to remain at any displacement to which it is
disturbed.
(i)
c)
Neutral static longitudinal stability usually defines the lower limit of aeroplane
stability since it is the boundary between stability and instability. The
aeroplane with neutral static stability may be excessively responsive to controls
and the aircraft has no tendency to return to trim following a disturbance generally this would not be acceptable.
The aircraft which is unstable will continue to pitch in the disturbed direction until the
displacement is resisted by opposing control forces.
(i)
The aeroplane with negative static longitudinal stability is inherently divergent
from any intended trim condition. If it is at all possible to fly the aircraft, it
cannot be trimmed and illogical control forces and deflections are required to
provide equilibrium with a change of attitude and airspeed - clearly, this would
be totally unacceptable.
10 - 5
© Oxford Aviation Services Limited
STABILITY AND CONTROL
PRINCIPLES OF FLIGHT
For the study of stability it is convenient to consider the changes in magnitude of lift force due
to changes in angle of attack, acting through a stationary point; the aerodynamic centre (AC).
It will be remembered that the location of the AC is at the quarter chord (or 25% aft of the
leading edge). It should be noted that the pitching moment about the AC is negative (nose down)
and that this negative (nose down) pitching moment about the AC does not change with changes
in angle of attack. Fig. 10.5.
a 1
MOMENT (M) REMAINS THE SAME AT "NORMAL" ANGLES OF ATTACK BECAUSE
Figure 10.5
Aerodynamic Centre (AC)
The pitching moment about the AC remains constant as the angle of attack is increased because
the magnitude of the lift force increases but acts through a smaller arm due to the CP moving
forward. It is only at the AC (25% chord) that this will occur. If a point in front of, or to the rear
of the AC were considered, the pitching moment would change with angle of attack.
For the study of stability we will consider the lift to act at the AC. The AC is a stationary point
located at the 25% chord, only when the airflow is subsonic.
10 - 6
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
STABILITY AND CONTROL
L L
...
AC:/)
Flight
Path ...
X
I
I
~I
I
I
I
wing
-,----
Momentary Relative Airflow
due to Gust
Figure 10.6
A wing alone is unstable
A wing considered alone is statically unstable, because the AC is in front of the CG, Fig. 10.6.
A vertical gust will momentarily increases the angle of attack and increase lift (~L), which, when
multiplied by arm 'x' , will generate a positive (nose up) pitching moment about the CG. This
will tend to increase the angle of attack further, an unstable pitching moment. The wing on its
own would rotate nose up about the CG, Fig.10.7.
AN AIRCRAFT ROTATES)
( AROUND ITS CG
AC :x)- ---+
--~CG
UNSTABLE (NOSE UP)
PITCHING MOMENT
ABOUT THE CG
Figure 10.7
10 - 7
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
STABILITY AND CONTROL
Now consider a wing together with a tailplane. The tailplane is positioned to generate a
stabilising pitching moment about the aircraft CG. The same vertical gust will increases the
angle of attack of the tailplane and increase tailplane lift (~Lt), which, when multiplied by arm
'y', will generate a negative (nose down) pitching moment about the aircraft CG.
If the tail moment is greater than the wing moment the sum of the moments will not be zero and
the resultant nose down moment will give an angular acceleration about the CG. The nose down
angular acceleration about the CG will return the aircraft towards its original position of
equilibrium. The greater the tail moment relative to the wing moment, the greater the rate of
acceleration towards the original equilibrium position. (Too much angular acceleration is not
good).
I
I
,
Flight
Path
~
Y - - -- -
---j~.
+6Lt
,
I
~_-l_~_~ /~ AIRCRAFT
~~
C __ ____~ AC _~
I
••Itt
CG
--~~
---~=--------
09 AC talr------~
-,---~
~:C---------~-
--------~-
~--------Momentary Relative Airflow
Momentary Relative Airflow
due to Gust
due to Gust
AC
L
LL
Change in angle of attack due to gust
Tailplane lift
Aerodynamic Centre
Change in tail plane lift
Wing lift
x
Arm from wing AC to aircraft CG
Change in wing lift
y
Arm from tailplane AC to aircraft CG
Figure 10.8
There are two moments to consider; the wing moment and the tail moment. The wing moment
is a function of the change in wing lift multiplied by arm 'x'. The tail moment is a function of
the change in tailplane lift multiplied by arm 'y', Fig. 10.8. The length of both arms is dependent
upon CG position. If the CG is considered in a more forward position, the tail arm is larger and
the wing arm is smaller. A more forward CG position increases static longitudinal stability.
If the nose down (negative) tail moment is greater than the nose up (positive) wing moment, the
aircraft will have static longitudinal stability.
10 - 8
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
STABILITY AND CONTROL
-+
INCREASED
I" L
/
/
.-- X
/
I
----I.~ I .~----
DECREASED
Y
I
NEUTRAL POINT
POSITION OF CG WHEN
TAIL MOMENT AND
WING MOMENT
ARE EQUAL
Figure 10.9
10.5
Neutral Point
NEUTRAL POINT
If you consider the CG moving rearwards from a position of static longitudinal stability:a)
the tail arm 'y' will decrease and the wing arm 'x' will increase; consequently,
b)
the (negative) tail moment will decrease and the (positive) wing moment will increase,
Fig. 10.9.
Eventually the CG will reach a position at which the tail moment is the same as the wing
moment. If a vertical gust were to displace the aircraft nose up, the sum of the moments will be
zero and there will be no angular acceleration about the CG to return the aircraft towards its
original position of equilibrium.
Because there is no resultant moment, either nose up or nose down, the aircraft will remain in
its new position of equilibrium; the aircraft will have neutral static longitudinal stability.
Para. lOA (b).
The position of the CG when the sum of the changes in the tail moment and wing moment caused
by the gust is zero, is known as the neutral point, Fig. 10.9.
10 - 9
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
10.6
STABILITY AND CONTROL
STATIC MARGIN
We have established that: with the CG on the neutral point the aircraft will have neutral static
longitudinal stability. i.e. the sum changes in the wing moment and the tail moment caused by
a disturbance is zero.
If the CG is positioned just forward of the neutral point, the tail moment will be slightly greater
than the wing moment (arm ' y' increased and arm 'x' decreased). A vertical gust which
increases the angle of attack will generate a small nose down angular acceleration about the CG,
which will gently return the aircraft towards its original position of trim (equilibrium).
The further forward the CG, the greater the nose down angular acceleration about the CG - the
greater the degree of static longitudinal stability.
STATIC MARGIN
/
L
L ....
/
I
X
/
/
IIi'T~
k~
Y
1_",>
I
AC
••
-AFTCG~
LIMIT
NEUTRAL POINT
Figure 10.10
Static Margin & Aft CG Limit
The neutral point is an important point of reference in the study of static longitudinal stability.
In practice, the CG wi Il never be allowed to move so far aft that it reached the neutral point. The
aircraft would be much too sensitive to the controls. Para. 10.4 (b).
It has been stated that the further forward the CG is from the neutral point, the greater the static
longitudinal stability. The distance the CG is forward of the neutral point will give a measure
of the static longitudinal stability; this distance is called the static margin, Fig. 10.10. The
greater the static margin, the greater the static longitudinal stability.
A certain amount of static longitudinal stability is always required, so the aft CG limit will be
positioned some distance forward of the neutral point. The distance between the neutral point
and the aft CG limit gives the required minimum static stability margin.
10 - 10
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
10.7
STABILITY AND CONTROL
TRIM AND CONTROLLABILITY
An aircraft is said to be trimmed (in trim) if all moments in pitch, roll, and yaw are equal to zero.
The establishment of trim (equilibrium) at various conditions offlight may be accomplished by:a)
pilot effort,
b)
trim tabs,
c)
variable incidence trimming tailplane,
d)
moving fuel between the wing tanks and an aft located trim tank, or
e)
bias of a surface actuator (powered flying controls).
STATIC STABILITY RESISTS
PILOT EFFORT AND GUSTS
The term controllability refers to the ability of the aircraft to respond to control surface
displacement and achieve the desired condition of flight. Adequate controllability must be
available to perform takeoff and landing and accomplish the various manoeuvres in flight.
A contradiction exists between stability and controllability. A high degree of stability gives
reduced controllability. The relationship between static stability and controllability is
demonstrated by the following four illustrations.
POSITIVE STATIC STABILITY
Figure 10.11
Degrees of static stability are illustrated by a ball placed on various surfaces. Positive static
stability is shown by the ball in a trough, Fig. 10.11; ifthe ball is displaced from equilibrium at
the bottom of the trough, there is an initial tendency to return to equilibrium. If it is desired to
"control" the ball and maintain it in the displaced position, a force must be supplied in the
direction of displacement to balance the inherent tendency to return to equilibrium.
This same stable tendency in an aircraft resists displacement from trim equally, whether by
pilot effort on the controls (stick force) or atmospheric disturbance.
10- 11
© Oxford Aviation Services Limited
STABILITY AND CONTROL
PRINCIPLES OF FLIGHT
INCREASED POSITIVE
STATIC STABILITY
INCREASED STATIC STABILITY
INCREASES STICK FORCE
Figure 10.12
The effect of increased static stability (forward CG movement) on controllability is illustrated
by the ball in a steeper trough, Fig. 10.12. A greater force is required to "control" the ball to the
same position of displacement when the static stability is increased. In this manner, a large
degree of static stability tends to make the aircraft less controllable. It is necessary to achieve
the proper proportion between static stability and controllability during the design of an aircraft
because too much static stability (Forward CG position) reduces controllability. The forward
CG limit is set to ensure minimum controllability, Fig. 10.13 .
('A5;}_1
FWD
LIMIT
~:/
HIGH
STICK
FORCE
AFT CG
LIMIT
LOW
STICK
FORCE
Figure 10.13
Fwd & Aft CG Limits
10 - 12
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
STABILITY AND CONTROL
........ -
,
I
,
\
,
.......,
\
,
I
.... 7 -"
I
o
NEUTRAL STATIC STABILITY
I
Figure 10.14
DECREASED STATIC STABILITY )
REDUCES STICK FORCE
The effect of reduced static stability on controllability is shown by the ball on a flat surface,
Fig. 10.14. If neutral static stability existed (CG on the neutral point), the ball may be displaced
from equilibrium and there is no tendency to return. A new point of equilibrium is obtained and
no force is required to maintain the displacement. As static stability approaches zero,
controllability increases to infinity and the only resistance to displacement is a resistance to the
motion of displacement, aerodynamic damping. For this reason, decreased static stability (Aft
CG movement) increases controllability. If the stability of the aircraft is too low, control
deflections may create exaggerated displacements of the aircraft.
NEGATIVE STATIC STABILITY
Figure 10.15
The effect of static instability on controllability (CG aft of the neutral point) is shown in
Fig. 10.15 by the ball on a hill. If the ball is displaced from equilibrium at the top of the hill, the
initial tendency is for the ball to continue in the displaced direction. In order to "control" the
ball at this position of displacement, a force must be applied opposite to the direction of
displacement.
This effect would be apparent during flight by an unstable "feel" to the aircraft. If the controls
were deflected to increase the angle of attack, the aircraft would need to be ' held' at the higher
angle of attack by a push force to keep the aircraft from continuing in the nose up direction. The
pilot would be supplying the stability by his attempt to maintain the equilibrium, this is totally
unacceptab Ie!
10 - 13
© Oxford Aviation Services Limited
STABILITY AND CONTROL
PRINCIPLES OF FLIGHT
KEY FACTS 1 - Self Study (Insert the missing words, with reference to the preceding paragraphs).
Stability is the ____ of an aircraft to return to a _ _ state of flight, after being disturbed
by an external
, without any help from the _ _ . (Para. 10.1).
There are two broad categories of stability; ____ and ____ . (Para. 10.1).
An aircraft is in a state of _ _ _ _ (trim) when the sum of all forces is _ _ and the sum of
all
is zero. (Para. 10.2).
The type of static stability an aircraft possesses is defined by its ___ tendency, following the
removal of some disturbing force. (Para. 10.2).
The three different types of static stability are: (Para. 10.2).
a)
static stability exists if an aircraft is disturbed from equilibrium and has the
tendency to return to equilibrium.
b)
static stability exists if an aircraft is subject to a disturbance and has neither the
tendency to return nor the tendency to continue in the displacement direction.
c)
static stability exists if an aircraft has a tendency to continue in the direction
of disturbance.
The longitudinal axis passes through the __ from _ _ to _ _ . (Para. 10.3)
The normal axis passes "vertically" through the_at_ 0 tothe _ _ _ _ _ axis. (Para. 10.3)
The lateral axis is a line passing through the _ , parallel to a line passing through the __ tips.
(Para. 10.3).
The three reference axes all pass through the _ __
_ _ _ . (Para. 10.3)
Lateral stability involves motion about the _ _ _ _ axis (_ _~). (Para. 10.4).
Longitudinal stability involves motion about the ___ axis ('-_ _-/). (Para. 10.4).
Directional stability involves motion about the _ _ _ axis ('-_ _-'). (Para. 10.4).
10 - 14
© Oxford Aviation Services Limited
STABILITY AND CONTROL
PRINCIPLES OF FLIGHT
We consider the changes in _____ oflift force due to changes in angle of ____, acting
through a
point; the
. (Page 10 - 6)
The aerodynamic centre (AC) is located at the _% chord position. (Page 10 - 6)
The
(Page 10 - 6).
pitching moment about the AC remains ____ at normal angles of attack.
A wing on its own is statically ____ because the _
is in front of the _ . (Page 10 - 7).
An upward vertical gust will momentarily
the angle of attack of the wing. The
____ lift force magnitude acting through the _ will increase the
pitching moment
pitching moment. (Page 10 - 7).
about the
. This is an
The
is positioned to generate a _ _ _ _ pitching moment about the aircraft _ .
(Page 10 - 8).
If the tail moment is greater than the wing moment the sum of the moments will not be __ and
about the __ . (Page 10 - 8).
the resultant nose _ _ moment will give an angular
The
the tail moment relative to the wing moment, the _ _ _ the rate of return _ __
the original
position. (Page 10 - 8).
The tail moment is increased by moving the aircraft _
and decreases the _ _ arm. (Page 10 - 8).
forwards, which _ _ _ _ the tail arm
If the nose down (_ _- » tail moment is greater than the nose up (_ _- » wing moment, the
aircraft will have
stability. (Page 10 - 8).
The position of the CG when changes in the sum of the tail moment and wing moment due to a
_ _ . (Para. 10.5).
disturbance is zero, is known as the
The further forward the _ , the ___ the nose down angular _____ about the
___ the degree of__
stability. (Para. 10.6).
-the
The
the
is forward of the
point will give a measure of the _ _
longitudinal stability; this distance is called the static
. (Para. 10.6).
The greater the static margin, the ___ the _ _ _ _ _ _ _ _ stability. (Para. 10.6).
The __ CG limit will be positioned some distance _ _ _ of the _ _ _ _ . (Para. 10.6).
The distance between the _ _ limit and the neutral point gives the required _ _ _ _ static
stability
. (Para. 10.6).
10 - 15
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
STABILITY AND CONTROL
An aircraft is said to be _ _ _ ifall ____ in pitch, roll, and yaw are equal to _ _ .
(Para. 10.7).
Trim ( _ _ _ _ ) is the function ofthe _ _ _ andmaybe accomplished by:- (Para. 10.7).
a)
effort
b)
trim _ _,
tank, or
c)
moving _ _ between the wing ___ and an aft located
d)
bias of a surface
(
flying controls).
The term
refers to the ability of the aircraft to respond to control surface
of flight. (Para. 10.7).
displacement and achieve the desired
A high degree of stability tends to reduce the _ _ _ _ _ of the aircraft. (Para. 10.7).
The stable tendency of an aircraft resists displacement from _
on the controls (
force) or _ _ . (Para. 10.7).
equally, whether by _ _ effort
If the CG moves forward, static longitudinal stability ____ and controllability _ _ __
). (Para. 10.7).
(stick force
If the CG moves aft, static longitudinal stability _ _ _ _ and controllability ____ (stick
). (Para. 10.7).
force
With the CG on the forward limit, static longitudinal stability is _ _ _, controllability is _ _
and stick force is _ _ . (Para. 10.7).
With the CG on the aft limit, static longitudinal stability is _ _, controllability is
stick force is __ . (Para. 10.7).
and
The aft CG limit is set to ensure a _ _ _ _ degree of static longitudinal stability. (Para. 10.7).
The fwd CG limit is set to ensure a _ _ _ _ degree of controllability under the worst
circumstance. (Para. 10.7).
KEY FACTS 1, WITH THE MISSING WORDS INSERTED CAN BE FOUND ON PAGE 10 - 89.
10 - 16
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
10.8
STABILITY AND CONTROL
GRAPHIC PRESENTATION OF STATIC LONGITUDINAL STABILITY
Static longitudinal stability depends upon the relationship of angle of attack and pitching
moment. It is necessary to study the pitching moment contribution of each component of the
aircraft. In a manner similar to all other aerodynamic forces, the pitching moment about the
lateral axis is studied in the coefficient form.
M
=
CM
Q
S
(MAC)
or
M
Q
S
(MAC)
where,
M
=
pitching moment about the CG
(positive if in a nose - up direction)
Q
=
dynamic pressure
S
=
wing area
MAC
=
mean aerodynamic chord
CM
=
pitching moment coefficient
The pitching moment coefficients contributed by all the various components of the aircraft are
summed up and plotted versus lift coefficient (angle of attack).
Study of the plots of CM versus CL is a convenient way to relate the static longitudinal stability
of an aeroplane.
10 - 17
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
STABILITY AND CONTROL
+
®
!zwO~
~
-
01-
~z
W
(90
LIFT COEFFICIENT
z_LL
ILL
OW
1-0
a.. 0
Graph A illustrates the variation of pitching moment coefficient (C M ) with lift coefficient (C L )
for an aeroplane with positive static longitudinal stability. Evidence of static stability is shown
by a tendency to return to equilibrium, or "trim", upon displacement. The aeroplane described
by graph A is in trim or equilibrium when CM = 0 and, if the aeroplane is disturbed to some
different CL the pitching moment change tends to return the aircraft to the point of trim. If the
aeroplane were disturbed to some higher CL (point y), a negative or nose-down pitching moment
is developed which tends to decrease angle of attack back to the trim point. If the aeroplane
were disturbed to some lower CL (point x), a positive or nose up pitching moment is developed
which tends to increase the angle of attack back to the trim point. Thus, positive static
longitudinal stability is indicated by a negative slope of CM versus CL . The degree of static
longitudinal stability is indicated by the slope of the curve (red line).
+
@
STABLE
NEUTRAL
\
"-~ UNSTABLE
Graph B provides comparison of a stable and an unstable condition. Positive static stability is
indicated by the red curve with negative slope. Neutral static stability would be the result if the
curve had zero slope. If neutral stability exists, the aeroplane could be disturbed to some higher
or lower lift coefficient without change in pitching moment coefficient.
10 - 18
© Oxford Aviation Services Limited
STABILITY AND CONTROL
PRINCIPLES OF FLIGHT
Such a condition would indicate that the aeroplane would have no tendency to return to some
original equilibrium and would not hold trim. An aeroplane which demonstrates a positive slope
of the CM versus CL curve (blue line) would be unstable. If the unstable aeroplane were subject
to any disturbance from equilibrium at the trim point, the changes in pitching moment would
only magnify the disturbance. When the unstable aeroplane is disturbed to some higher CL a
positive change in CM occurs which would illustrate a tendency for continued, greater
displacement. When the unstable aeroplane is disturbed to some lower CL a negative change in
CM takes place which tends to create continued displacement.
+
©
STABLE
LESS STABLE
NEUTRAL
Ordinarily, the static longitudinal stability of a conventional aeroplane configuration does not
vary with lift coefficient. In other words, the slope of CM versus CL does not change with CL .
However, if:
a)
the aeroplane has sweep back,
b)
there is a large contribution of 'power effect' on stability, or
c)
there are significant changes in down wash at the horizontal tail,
noticeable changes in static stability can occur at high lift coefficients (low speed) . This
condition is illustrated by graph C. The curve ofC M versus CL of this illustration shows a good
stable slope at low values of CL (high speed). Increasing CL gives a slight decrease in the
negative slope hence a decrease in stability occurs. With continued increase in C L the slope
becomes zero and neutral stability exists. Eventually, the slope becomes positive and the
aeroplane becomes unstable or "pitch-up" results.
Remember, at any lift coefficient, the static stability of the aeroplane is depicted by the slope of
the curve of CM versus CL .
10 - 19
© Oxford Aviation Services limited
PRINCIPLES OF FLIGHT
10.9
STABILITY AND CONTROL
CONTRIBUTION OF THE COMPONENT SURFACES
The net pitching moment about the lateral axis is due to the contribution of each of the
component surfaces acting in their appropriate flow fields.
By studying the contribution of each component, their effect on static stability may be
appreciated. It is necessary to recall that the pitching moment coefficient is defined as:
M
Q
S
(MAC)
Thus, any pitching moment coefficient (C M ) - regardless of source - has the common
denominator of dynamic pressure (Q), wing area (S), and wing mean aerodynamic chord (MAC).
This common denominator is applied to the pitching moments contributed by the:
a)
fuselage and nacelles,
b)
horizontal tail, and
c)
power effects as well as pitching moments contributed by the wing.
WING: The contribution of the wing to stability depends primarily on the location of the
aerodynamic centre (AC) with respect to the aeroplane centre of gravity. Generally, the
aerodynamic centre is defined as the point on the wing Mean Aerodynamic Chord (MAC)
where the wing pitching moment coefficient does not vary with lift coefficient., All changes
in lift coefficient effectively take place at the wing aerodynamic centre. Thus, if the wing
experiences some change in lift coefficient, the pitching moment created will be a direct function
of the relative location of the AC and CG.
*The degree of positive camber of the wing has no effect on longitudinal stability. The
pitching moment about the AC is always negative regardless of angle of attack.
Stability is given by the development of restoring moments. As the wing AC is forward of the
CG, the wing contributes an unstable pitching moment to the aircraft, as shown in Fig. 10.16.
10 - 20
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
STABILITY AND CONTROL
+CHANGE IN LIFT
I
I
I
AERODYNAMIC CENTRE
CG AFT
OFAC~
" - UNSTABLE SLOPE
Figure 10.16
Unstable Wing Contribution
Since the wing is the predominating aerodynamic surface of an aeroplane, any change in the
wing contribution may produce a significant change in the aeroplane stability.
10 - 21
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
STABILITY AND CONTROL
SYMMETRICAL BODY (fuselage or nacelle)
Figure 10.17
FUSELAGE AND NACELLES: In most cases, the contribution of the fuselage and nacelles
is destabilising. A symmetrical body in an airflow develops an unstable pitching moment when
given an angle of attack. In fact, an increase in angle of attack produces an increase in the
unstable pitching moment without the development of lift. Fig. 10.17 illustrates the pressure
distribution which creates this unstable moment on the body. An increase in angle of attack
causes an increase in the unstable pitching moment but a negligible increase in lift.
HORIZONTAL TAIL: The horizontal tail usually provides the greatest stabilising influence
of all the components of the aeroplane.
I
I
L
I
~ X
------. 1. . .1--- - - - - - y -------I~~
T
1
' Lt
1
Lt
Flight
Path .-
Momentary Relative Airflow
due to Gust
Momentary Relative Airflow
due to Gust
Figure 10.18
10 - 22
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
STABILITY AND CONTROL
To appreciate the contribution of the horizontal tail to stability, inspect Fig. 10.18. If the
aeroplane is given an increase in angle of attack (by a gust OR control displacement), an increase
in tail lift will occur at the aerodynamic centre of the tail. An increase in lift at the horizontal
tail produces a negative (stabilising) moment about the aircraft CG.
For a given vertical gust velocity and aircraft TAS, the wing moment is essentially determined
by the CG position. BUT, the tail moment is determined by the CG position AND the
effectiveness ofthe tailplane. For a given moment arm (CG position), the effectiveness of the
tailplane is dependent upon:
a)
Downwash from the wing
b)
Dynamic pressure at the tailplane
c)
Longitudinal dihedral
Downwash from the wing and dynamic pressure at the tailplane will be discussed in due course,
but the effect oflongitudinal dihedral is shown below.
LONGITUDINAL DIHEDRAL: The difference between tailplane and wing incidence. For
longitudinal static stability the tailplane incidence is smaller. As illustrated in Fig. 10.19, this
will generate a greater percentage increase in tailplane lift than wing lift for a given vertical gust.
This guarantees that the positive contribution of the tailplane to static longitudinal stability will
be sufficient to overcome the sum of the de-stabilising moments from the other components of
the aeroplane.
D
L
= 100%
Lt =200%
4° INCIDENCE
2° INCIDENCE
r
1 ~~0:-A-C
-
--
IN ANGLE OF ATTACK
DUE TO VERTICAL GUST
Figure 10.19 Longitudinal Dihedral
10 - 23
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
STABILITY AND CONTROL
DOWNWASH AT
HORIZONTAL TAIL
Figure 10.20
DOWNWASH DECREASES
LONGrrUDINAL STATIC STABILITY
DOWNWASH: It should be appreciated that the flow at the horizontal tail does not have the
same flow direction or dynamic pressure as the free stream. Due to the wing wake, fuselage
boundary layer, and power effects, the dynamic pressure at the horizontal tail may be greatly
different from the dynamic pressure of the free stream. In most instances, the dynamic pressure
at the tail is usually less and this reduces the efficiency of the tail.
When the aeroplane is given a change in angle of attack, the horizontal tail does not experience
the same change in angle of attack as the wing, Fig. 10.20.
Because of the increase in downwash behind the wing, the horizontal tail will experience a
smaller change in angle of attack, e.g., if a 10° change in wing angle of attack causes a 4 °
increase in downwash at the horizontal tail, the horizontal tail experiences only a 6 ° change in
angle of attack. In this manner, the downwash at the horizontal tail reduces the contribution to
stability.
Any factor which alters the rate of change of downwash at the horizontal tail (e.g. flaps or
propeller slipstream) will directly affect the tail contribution and aeroplane stability. Downwash
decreases static longitudinal stability.
10 - 24
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
10.10
STABILITY AND CONTROL
POWER-OFF STABILITY
When the aerodynamic stability of a configuration is of interest, power effects are neglected and
the stability is considered by a buildup of the contributing components.
Fig. 10.21 illustrates a typical buildup of the components of a conventional aeroplane
configuration. If the CG is arbitrarily set at 30 percent MAC, the contribution of the wing alone
is destabilising, as indicated by the positive slope ofC M versus C L • The combination of the wing
and fuselage increases the instability. The contribution of the tail alone is highly stabilizing from
the large negative slope of the curve. The contribution of the tail must be sufficiently stabilising
so that the complete configuration will exhibit positive static stability at the anticipated CG
locations.
TYPICAL BUILD-UP OF COMPONENTS
+
...
I
WING + FUSELAGE ..
WING ONLY
----~----~~~ --~~------------- ----------------.
'" '"
CL
'" '"
CG @ 30% MAC
'"
Lp~::"
ONLY
Figure 10.21
10 - 25
© Oxford Aviation Services Limited
STABILITY AND CONTROL
PRINCIPLES OF FLIGHT
EFFECT OF CG POSITION
+
Figure 10.22
10.11
EFFECT OF CG POSITION
A variation of CG position can cause large changes in the static longitudinal stability. In the
conventional aeroplane configuration, the large changes in stability with CG variation are
primarily due to the large changes in the wing contribution. If the incidence of all surfaces
remains fixed , the effect of CG position on static longitudinal stability is typified by the chart
in Fig. 10.22. As the CG is gradually moved aft, the aeroplane static stability decreases, then
becomes neutral then unstable. The CG position which produces zero slope and neutral static
stability is referred to as the "neutral point." The neutral point may be imagined as the effective
aerodynamic centre of the entire aeroplane configuration, i.e., with the CG at the neutral point,
all changes in net lift effectively occur at that point and no change in pitching moment
results. The neutral point defines the most aft CG position without static instability.
10 - 26
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
10.12
STABILITY AND CONTROL
POWER EFFECTS
The effects of power may cause significant changes in trim lift coefficient and static longitudinal
stability. Since the contribution to stability is evaluated by the change in moment coefficients,
power effects will be most significant when the aeroplane operates at high power and low
airspeeds such as during approach and while taking-off.
DESTABILISING
Figure 10.23
The effects of power are considered in two main categories. First, there are the direct effects
resulting from the forces created by the propulsion unit. Next, there are the indirect effects of
the slipstream and other associated flow which alter the forces and moments of the aerodynamic
surfaces. The direct effects of power are illustrated in Fig. 10.23. The vertical location of the
thrust line defines one of the direct contributions to stability. If the thrust line is below the CG,
as illustrated, a thrust increase will produce a positive or nose up moment and the effect is
destabilising.
...
~
~
,
NORMAL FORCE DUE TO
MOMENT CHANGE
Figure 10.24
A propeller located ahead of the CG contributes a destabilising effect. As shown in Fig. 10.24,
a rotating propeller inclined to the relative airflow causes a deflection of the airflow. The
momentum change of the slipstream creates a normal force at the plane of the propeller. As this
normal force will increase with an increase in aeroplane angle of attack, the effect will be
destabilising when the propeller is ahead of the CG. The magnitude of the unstable
contribution depends on the distance from the CG to the propeller and is largest at high
power and low dynamic pressure.
10 - 27
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
STABILITY AND CONTROL
W lNG, NACELLE AND FUSELAGE
MOMENTS AFFECTED BY
SLIPSTREAM
-
-
DYNAMIC PRESSURE
AT TAIL AFFECTED
BY SLIPSTREAM
WING LIFT AFFECTED
BY SLIPSTREAM
Figure 10.25
The indirect effects of power are of greatest concern in the propeller powered aeroplane rather
than the jet powered aeroplane. As shown in Fig. 10.25, the propeller powered aeroplane creates
slipstream velocities on the various surfaces which are different from the flow field typical of
power-off flight. Since the various wing, nacelle, and fuselage surfaces are partly or wholly
immersed in this slipstream, the contribution of these components to stability can be quite
different from the power-off flight condition. Ordinarily, the change of fuselage and nacelle
contribution with power is relatively small. The added lift on the portion of the wing immersed
in the slipstream requires that the aeroplane operate at a lower angle of attack to produce the
same effective lift coefficient. Generally, this reduction in angle of attack to effect the same C L
reduces the tail contribution to stabi.1ity. However, the increase in dynamic pressure at the tail
tends to increase the effectiveness of the tail and may be a stabilising effect. The magnitude of
this contribution due to the slipstream velocity on the tail will depend on the CG position and
trim lift coefficient.
~-----.;;:.",-~~____
..
------------
---------
rf
DOWNWASH AT TAIL
AFFECTED BY
SLIPSTREAM DIRECTION
------ -....------------
~-~--- ~
----------
Figure 10.26
The deflection of the slipstream shown in Fig. 10.26 by the normal force at the propeller tends
to increase the downwash at the horizontal tail and reduce the contribution to stability.
10 - 28
© Oxford Aviation Services Limited
STABILITY AND CONTROL
PRINCIPLES OF FLIGHT
FLOW INDUCED BY
_
7 / -JEr/FAN
~/j~
----------C5===~=-~.~~t ~
-- --
EXHAUST
~
Figure 10.27
Essentially the same destabilising effect is produced by the flow induced at the exhaust of turbojet/fan engines, Fig. 10.27. Ordinarily, the induced flow at the horizontal tail ofajet aeroplane
is slight and is destabilising when the jet passes underneath the horizontal tail. The magnitude
of the indirect power effects on stability tends to be greatest at high C L , high power, and low
flight speeds.
CONCLUSIONS TO THE EFFECTS OF POWER
The combined direct and indirect power effects contribute to a general reduction of static
stability at high power, high C L and low dynamic pressure. It is generally true that any
aeroplane will experience the lowest level of static longitudinal stability under these
conditions. Because ofthe greater magnitude of both direct and indirect power effects, the
propeller powered aeroplane usually experiences a greater effect than the jet powered
aeroplane.
10.13
HIGH LIFT DEVICES
An additional effect on stability can be from the extension of high lift devices. High lift devices
tend to increase down wash at the tail and reduce-the dynamic pressure at the tail, both of which
are destabilising. However, high lift devices may prevent an unstable contribution of the wing
at high CL • While the effect of high lift devices depends on the aeroplane configuration, the
usual effect is destabilising. Hence, the aeroplane may experience the most critical forward
neutral point during the power approach or overshoot/missed approach. During this
condition of flight the static stability is usually the weakest and particular attention must
be given to precise control of the aeroplane.
The power - on neutral point may set the most aft limit of CG position.
10 - 29
© Oxford Aviation Services Limited
STABILITY AND CONTROL
PRINCIPLES OF FLIGHT
10.14
CONTROL FORCE STABILITY
The static longitudinal stability of an aeroplane is defined by the tendency to return to
equilibrium upon displacement. In other words, a stable aeroplane will resist displacement from
trim or equilibrium. The control forces of the aeroplane should reflect the stability of the
aeroplane and provide suitable reference to the pilot for precise control of the aeroplane.
EFFECT OF ELEVATOR DEFLECTION
c__--- -'.-\~~_
ELEVATOR
DEFLECTION
,
?
t
+
30° Up
TRIM FOR
10° UP
20° Up
TRIM FOR Oo ~
I---
10° Up
..
CG @ 20% MAC
Figure 10.28
The effect of elevator deflection on pitching moments is illustrated by the graph of Fig. 10.28.
If the elevators of the aeroplane are held at zero deflection, the resulting line ofC M versus CL for
0 0 depicts the static stability and trim lift coefficient. If the elevators are held at a deflection of
10 up (aircraft trimmed at a lower speed), the aeroplane static stability is unchanged but the
trim lift coefficient is increased.
0
A CHANGE IN ELEVATOR POSITION DOES NOT ALTER
THE TAIL CONTRIBUTION TO STABILITY
10 - 30
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
STABILITY AND CONTROL
As the elevator is held in various positions, equilibrium (trim) will occur at various lift
coefficients, and the trim C L can be correlated with elevator deflection as shown in Fig. 10.29.
TRIM
C L VERSUS
ELEVATOR DEFLECTION
CG LOCATION -_____
--..
z
()
W
UP
W
o
0:::
~
/
~
20% MAC
o
i=
....J
U.
/
10% MAC
~
~
~
.----------------
30%MAC ---
--+-----~~-=-----==-~
~=------------------------------ ~
W
CL
40% MAC
....J
W
(NEUTRAL POINT)
DONN
Figure 10.29
When the CG position of the aeroplane is fixed, each elevator position corresponds to a
particular trim lift coefficient. As the CG is moved aft the slope of this line decreases and the
decrease in stability is evident by a given control displacement causing a greater change in trim
lift coefficient. This is evidence that decreasing stability causes increased controllability
and, of course, increasing stability decreases controllability.
If the CG is moved aft until the line of trim C L versus elevator deflection has zero slope, neutral
static stability is obtained.
10 - 31
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
STABILITY AND CONTROL
TRIM AIRSPEED VERSUS ELEVATOR DEFLECTION
z
o
i=
()
W
...J
u..
W
UNSTABLE
UP
o
a:::
o
~
EQUIVALENT
~
-+-------~---~~~---"-------:AC""CI=RS
=-:P
=-:E
=-=E=-=D:--~
...J
W
DOWN
Figure 10.30
Since each value of lift coefficient corresponds to a particular value of dynamic pressure
required to support an aeroplane in level flight, trim airspeed can be correlated with elevator
deflection as in the graph of Fig. 10.30.
If the CG location is ahead of the neutral point and control position is directly related to surface
deflection, the aeroplane will give evidence of stick position stability. In other words, the
aeroplane will require the stick to be moved aft to increase the angle of attack and trim at a lower
airspeed and to be moved forward to decrease the angle of attack and trim at a higher airspeed.
It is highly desirable to have an aeroplane demonstrate this feature. If the aeroplane were to have
stick position instability, the aeroplane would require the stick to be moved aft to trim at a higher
airspeed or to be moved forward to trim at a lower airspeed.
10 - 32
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
STABILITY AND CONTROL
There is an increment offorce dependent on the trim tab setting which varies with the dynamic
pressure or the square of equivalent airspeed. Fig.10.31 indicates the variation of stick force
with airspeed and illustrates the effect of tab setting on stick force.
EFFECT OF TRIM TAB SETTING
PULL
w
o
IX
ou..
o
~
o
CG @ 20% MAC
i=
en PUSH
Figure 10.31
In order to trim the aeroplane at point (1) a certain amount of up elevator is required and zero
stick force is obtained with the use of the trim tab. To trim the aeroplane for higher speeds
corresponding to points (2) and (3), less and less aircraft nose-up tab is required.
Note that when the aeroplane is properly trimmed, a push force is required to increase airspeed
and a pull force is required to decrease airspeed. In this manner, the aeroplane would have
positive stick force stability with a stable "feel" for airspeed.
10 - 33
© Oxford Aviation Services Limited
STABILITY AND CONTROL
PRINCIPLES OF FLIGHT
i
EFFECT OF CG POSITION
CG POSITION
r--_~1 Q!~ rv1AC
PULL
I
UJ
o
0::::
o
LL
:::.r:::
EAS
40% rv1AC
O l---~~~---------------~~~~------------~
o
I-
(j)
PUSH
Figure 10.32
If the CG of the aeroplane were varied while maintaining trim at a constant airspeed, the effect
ofCG position on stick force stability could be appreciated. As illustrated in Fig. 10.32, moving
the CG aft decreases the slope of the line of stick force through the trim speed. Thus, on
decreasing stick-force stability it is evident that smaller stick forces are necessary to displace the
aeroplane from the trim speed. When the stick force gradient (or slope) becomes zero, the CG
is at the neutral point and neutral stability exists. If the CG is aft of the neutral point, stick force
instability will exist, e.g. the aeroplane will require a push force at a lower speed or a pull force
at a higher speed. It should be noted that the stick force gradient is low at low airspeeds and
when the aeroplane is at low speeds, high power, and a CG position near the aft limit, the "feel"
for airspeed will be weak.
EFFECT OF CONTROL SYSTEM FRICTION
PULL
w
0
0::::
0
LL
~
EAS~
0
0
I(j)
PUSH
FRICTION FORCE
BAND
Figure 10.33
Control system friction can create very undesirable effects on control forces. Fig. 10.33
illustrates that the control force versus airspeed is a band rather than a line. A wide friction force
band can completely mask the stick force stability when the stick force stability is low. Modern
flight control systems require precise maintenance to minimize the friction force band and
preserve proper feel to the aeroplane.
10 - 34
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
10.15
STABILITY AND CONTROL
MANOEUVRE STABILITY
When the pilot pitches the aircraft, it rotates about the CG and the tailplane is subject to a
pitching velocity, in this example, downwards. Due to the pitching velocity in manoeuvring
flight, the longitudinal stability of the aeroplane is slightly greater than in steady flight
conditions.
CHANGE IN TAIL LIFT
~
TAS
,~
~
-
r!
-
-
- ~--
I
\
-*----~\
/
(
~.
RELATIVE AIRFLOW
FROM ANGULAR ROTATION
PITCHING VELOCITY
INCREASE IN TAIL ANGLE OF
ATTACK DUE TO PITCHING
VELOCITY
Figure 10.34
Aerodynamic Damping
Fig. 10.34 shows that the tailplane experiences an upwards component of airflow due to its
downwards pitching velocity. The vector addition of this vertical component to the T AS
provides an increase in effective angle of attack of the tail, which creates an increase in tail lift,
opposing the nose up pitch displacement.
Since the negative pitching moment opposes the nose up pitch displacement but is due to the
nose up pitching motion, the effect is a damping in pitch (aerodynamic damping).
It can be seen that an increase in TAS, for a given pitching velocity, decreases the angle of attack
due to pitching velocity
INCREASING ALTITUDE AT A CONSTANT lAS
DECREASES AERODYNAMIC DAMPING
10 - 35
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
STABILITY AND CONTROL
MANOEUVRE MARGIN
/
I
/
I
y-
I
~I
MANOEUVRE POINT
·~T
I".;t·,
Figure 10.35
Manoeuvre Point
The pitching moment from aerodynamic damping will give greater stability in manoeuvres than
is apparent in steady flight. The CG position when the tail moment would be the same as the
wing moment during manoeuvring is known as the manoeuvre point and this "neutral point"
will be further aft than for Ig flight, as shown in Fig.l0.35.
In most cases the manoeuvre point will not be a critical item; if the aeroplane demonstrates static
stability in I g flight, it will definitely have stability in manoeuvring flight.
10.16
STICK FORCE PER 'g'
The most direct appreciation of the manoeuvring stability of an aeroplane is obtained from a plot
of stick force versus load factor such as shown in Fig. 10.36. The aeroplane with positive
manoeuvring stability should demonstrate a steady increase in stick force with increase in load
factor or "g". The manoeuvring stick force gradient - or stick force per "g" - must be positive
but should be of the proper magnitude. The stick force gradient must not be excessively high
or the aeroplane will be difficult and tiring to manoeuver. Also, the stick force gradient must
not be too low or the aeroplane may be overstressed inadvertently when light control forces
exist.
10 - 36
© Oxford Aviation Services Limited
STABILITY AND CONTROL
PRINCIPLES OF FLIGHT
(J)
.0
~'NGSTK::K
w
o
0:::
oLL
20
I
FORCE GRADIENT
---~
~
o
I-
en
10
~~---------------------------- -----------.--.
2
1
3
4
5
6
7
8
LOAD FACTOR, (n)
or (g)
CG POSITION
% MAC
w
o
0:::
o
u..
w
o
10
0:::
oLL
20
~
o
~
o
I-
en
LOW
ALTITUDE
\
I-
30
\
en
HIGH
ALTITUDE
./
--------------40
:/
--j'--------------.-.-~
LOAD FACTOR
LOAD FACTOR
Figure 10.36
When the aeroplane has high static stability, the manoeuvring stability will be high and a high
stick force gradient will result, Fig. 10.36. A possibility exists that the forward CG limit could
be set to prevent an excessively high manoeuv~ing stick force gradient. As the CG moves aft,
the stick force gradient decreases with decreasing manoeuvring stability and the lower limit of
stick force gradient may be reached.
When asked to calculate ' stick force per g', remember that the aircraft is at 1g to start with. So
I g must be subtracted from the 'g' limit before dividing by the pull force.
The pitch damping of the aeroplane is related to air density. At high altitudes, the high T AS
reduces the change in tail angle of attack for a given pitching velocity and reduces the pitch
damping. Thus, a decrease in manoeuvring stick force stability can be expected with increased
altitude.
10 - 37
© Oxford Aviation Services Limited
STABILITY AND CONTROL
PRINCIPLES OF FLIGHT
10.17
TAILORING CONTROL FORCES
Control forces should reflect the stability of the aeroplane but, at the same time, should be of a
tolerable magnitude. A manual flying control system may employ an infinite variety of
techniques to provide satisfactory control forces throughout the speed, CG and altitude range of
the aircraft.
EFFECT OF STICK CENTRING SPRING
STICK CENTRING
\
PRINcGF = = = = = ==O
~/II\/IV\/ 0 \/\l\/\/\/\/\~
Figure 10.37
10.17.1
STICK CENTRING SPRING
If a spring is added to the control system as shown in Fig 10.37, it will tend to centre the stick
and provide a force increment depending on stick displacement.
When the control system has a fixed gearing between stick position and surface deflection, the
centring spring will provide a contribution to stick force stability according to stick position.
The contribution to stick force stability will be largest at low flight speeds where relatively large
control deflections are required. The contribution will be smallest at high airspeed because of
the smaller control deflections required. Thus, the stick centring spring will increase the
airspeed and manoeuvring stick force stability but the contribution decreases at high airspeeds.
A variation of this device would be a spring stiffness which would be controlled to vary with
dynamic pressure (Q - Feel). In that case, the contribution of the spring to stick force stability
would not diminish with speed.
10.17.2
DOWN SPRING
A down spring added to a control system is a means of increasing airspeed stick force stability
without a change in aeroplane static stability.
As shown in Fig. 10.38, a down spring consists of a long pre-loaded spring attached to the
control system which tends to rotate the elevators down (aircraft nose down). The effect of the
down spring is to contribute an increment of pull force independent of control deflection or
airspeed.
10 - 38
© Oxford Aviation Services Limited
STABILITY AND CONTROL
PRINCIPLES OF FLIGHT
EFFECT OF DOW NSPRING
PRELOADED SPRING
j
~ VW/VV'L ....- - -
Figure 10.38
When the down spring is added to the control system of an aeroplane and the aeroplane is retrimmed for the original speed, the airspeed stick force gradient is increased and there is a
stronger feel for airspeed. The down spring would provide a "synthetic" improvement to an
aeroplane deficient in airspeed stick force stability. Since the force increment from the down
spring is unaffected by stick position or normal acceleration, the manoeuvring stick force
stability would be unchanged.
10.17.3
BOBWEIGHT
The bobweight is an effective device for improving stick force stability. As shown in Fig. 10.39,
the bobweight consists of an eccentric mass attached to the control system which, in
unaccelerated flight, contributes an increment of pull force identical to the down spring. In fact,
a bobweight added to the control system of an aeroplane produces an effect identical to the down
spring. The bobweight will increase the airspeed stick force gradient and increase the feel for
airspeed.
The bobweight also has an effect on the manoeuvring stick force gradient since the bobweight
mass is subjected to the same acceleration as the aeroplane. Thus, the bobweight will provide
an increment of stick force in direct proportion to the manoeuvring acceleration ofthe aeroplane
(load factor applied). This will prevent the pilot applying too much 'g' during manoeuvres; the
more you pull back, the more resistance the bobweight adds to the control system.
EFFECT OF BOBW EIGHT
BOBWEIGHT
~,
Figure 10.39
10 - 39
© Oxford Aviation Services Limited
STABILITY AND CONTROL
PRINCIPLES OF FLIGHT
10.18
LONGITUDINAL CONTROL
To be satisfactory, an aeroplane must have adequate controllability as well as adequate stability.
An aeroplane with high static longitudinal stability will exhibit great resistance to displacement
from equilibrium. Hence, the most critical conditions of controllability will occur when the
aeroplane has high static stability, i.e., the lower limits of controllability will set the upper
limits of static stability. (Fwd. CG limit).
There are three principal conditions of flight which provide the critical requirements of
longitudinal control power (manoeuvring, take-off and landing). Anyone or combination of
these conditions can determine the overall longitudinal control power and set a limit to the
forward CG position.
10% MAC
t-
z
o
I-
o
UP
i
MAXIMUM _
_
18% MAC
~/ r ~
MOST FORWARD
CG FOR MANOEUVERING
DEFLECTION
W
CONTROLLABILITY
--l
20% MAC
LL
W
o
~ 30% MAC
0:::
o
I-
:;
W
--l
W
DOWN
I
CG
POSITION
C L max
Figure 10.40
10.19
MANOEUVRING CONTROL REQUIREMENT
The aeroplane should have sufficient longitudinal control power to attain the maximum usable
lift coefficient or the limit load factor during manoeuvres. As shown in Fig. 10.40, forward
movement of the CG increases the longitudinal stability of an aeroplane and requires larger
control deflections to produce changes in trim -lift coefficient. For the example shown, the
maximum effective deflection of the elevator is not capable of trimming the aeroplane at C Lmax
for CG positions ahead of 18 percent MAC.
This particular control requirement can be most critical for an aeroplane in supersonic flight.
Supersonic flight is usually accompanied by large increases in static longitudinal stability (due
to aft CP movement) and a reduction in the effectiveness of control surfaces. In order to cope
with these trends, powerful all-moving surfaces must be used to attain limit load factor or
maximum usable CL in supersonic flight. This requirement is so important that once satisfied,
the supersonic configuration usually has sufficient longitudinal control power for all other
conditions of flight.
10 - 40
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
10.20
STABILITY AND CONTROL
TAKEOFF CONTROL REQUIREMENT
At take-off, the aeroplane must have sufficient elevator control power to assume the takeoff
attitude prior to reaching takeoff speed.
LIFT
V
IL
LOAD
---------------------
...
~.-?
------- ROLLING FRICTION
WEIGHT
Figure 10.41
Fig. 10.41 illustrates the principal forces acting on an aeroplane during takeoff roll. When the
aeroplane is in the three point attitude at some speed less than the stall speed, the wing lift will
be less than the weight of the aeroplane. As the elevators must be capable of rotating to the
takeoff attitude, the critical condition will be with zero load on the nose wheel and the net of lift
and weight supported on the main gear.
Rolling friction resulting from the normal force on the main gear creates an adverse nose down
moment.
Also, the CG ahead of the main gear contributes a nose down moment and this consideration
could decide the most aft location of the main landing gear during design.
To balance these two nose down moments, the -horizontal tail must be capable of producing a
nose up moment big enough to attain the takeoff attitude at the specified speed.
The propeller aeroplane at take-off power may induce considerable slipstream velocity at the
horizontal tail which can provide an increase in the efficiency of the surface. The jet aeroplane
does not experience a similar magnitude ofthis effect since the induced velocities from the jet
are relatively small compared to the slipstream velocities from a propeller.
10 - 41
© Oxford Aviation Services Limited
STABILITY AND CONTROL
PRINCIPLES OF FLIGHT
10.21
LANDING CONTROL REQUIREMENT
At landing, the aeroplane must have sufficient control power to ensure adequate control at
specified landing speeds. The most critical requirement will exist when the CG is in the most
forward position, flaps are fully extended, and power is set at idle. This configuration will
provide the most stable condition which is most demanding of controllability.
The landing control requirement has one particular difference from the manoeuvring control
requirement of free flight. As the aeroplane approaches the surface, there will be a change in
the three-dimensional flow over the aeroplane due to ground effect. A wing in proximity to the
ground plane will experience a decrease in tip vortices and downwash at a given lift coefficient.
The decrease in downwash at the tail tends to increase the static stability and produce a nose
down moment from the reduction in download on the tail. Thus, the aeroplane just off the
runway surface, Fig. 10.42, will require additional control deflection to trim at a given lift
coefficient and the landing control requirement may be critical in the design of longitudinal
control power.
REDUCED DOWNWASH
~ DUE TO GROUND EFFECT
Figure 10.42
As an example of ground effect, a typical propeller powered aeroplane may require as much as
15 more up elevator to trim at C L MAX in ground effect than in free flight.
0
In some cases the effectiveness of the elevator is adversely affected by the use of trim tabs. If
trim is used to excess in trimming stick forces, the effectiveness of the elevator may be
reduced to hinder landing or takeoff control.
Each of the three principal conditions requiring adequate longitudinal control are critical for high
static stability. If the forward CG limit is exceeded, the aeroplane may encounter a deficiency
of controllability in any of these conditions.
The forward CG limit is set by the minimum permissible controllability
The aft CG limit is set by the minimum permissible stability.
10 - 42
© Oxford Aviation Services Limited
STABILITY AND CONTROL
PRINCIPLES OF FLIGHT
10.22
DYNAMIC STABILITY
While static stability is concerned with the initial tendency of an aircraft to return to
equilibrium, dynamic stability is defined by the resulting motion with time. If an aircraft is
disturbed from equilibrium, the time history of the resulting motion indicates its dynamic
stability. In general, an aircraft will demonstrate positive dynamic stability if the amplitude of
motion decreases with time. The various conditions of possible dynamic behaviour are
illustrated in the following six history diagrams. The nonoscillatory modes shown in diagrams
A, Band C depict the time histories possible without cyclic motion.
Initial
Disturbance
®
SUBSIDENCE
(or Dead Beat Return)
I-
Z
W
~
W
o
«-.J
TIME
a..
CJ)
o
(Positive StatiC)
(Positive Dynamic)
Chart A illustrates a system which is given an initial disturbance and the motion simply subsides
without oscillation, the mode is termed "subsidence" or "deadbeat return." Such a motion
indicates positive static stability by the initial tendency to return to equilibrium and positive
dynamic stability since the amplitude decreases with time.
I-
Z
W
@ DIVERGENCE
~
W
o
«-.J
TIME
a..
CJ)
o
(Negative Static)
(Negative Dynamic)
Chart B illustrates the mode of "divergence" by a non-cyclic increase of amplitude with time.
The initial tendency to continue in the displacement direction is evidence of static instability
and the increasing amplitude is proof of dynamic instability.
10 - 43
© Oxford Aviation Services Limited
STABILITY AND CONTROL
PRINCIPLES OF FLIGHT
©
NEUTRAL STATIC STABILITY
(Neutral Static)
(Neutral Dynamic)
TIME
Chart C illustrates the mode of pure neutral stability. If the original disturbance creates a
displacement which then remains constant, the lack of tendency for motion and the constant
amplitude indicate neutral static and neutral dynamic stability.
The oscillatory modes shown in diagrams D, E and F depict the time histories possible with
cyclic motion. One feature common to each of these modes is that positive static stability is
demonstrated by the initial tendency to return to equilibrium conditions. However, the resulting
dynamic behaviour may be stable, neutral , or unstable.
t
IZ
@
DAMPED OSCILLATION
w
~
w
u
~~-----------~
«
--l
TIME
D....
en
o
(Positive Static)
(Positive Dynamic)
Chart D illustrates the mode of a damped oscillation where the amplitude decreases with time.
The reduction of amplitude with time indicates there is resistance to motion and that energy
is being dissipated. Dissipation of energy or damping is necessary to provide positive dynamic
stability.
10 - 44
© Oxford Aviation Services Limited
STABILITY AND CONTROL
PRINCIPLES OF FLIGHT
®
fZ
(Positive Static)
(Neutral Dynamic)
w
:2
w
o
<{
TIME
-l
0....
if)
o
If there is no damping in the system, the mode of chart E is the result, an undamped oscillation.
Without damping, the oscillation continues with no reduction of amplitude with time. While
such an oscillation indicates positive static stability, neutral dynamic stability exists. Positive
damping is necessary to eliminate the continued oscillation. As an example, a car with worn
shock absorbers (or "dampers") lacks sufficient dynamic stability and the continued oscillatory
motion is both unpleasant and potentially dangerous. In the same sense, an aircraft must have
sufficient damping to rapidly dissipate any oscillatory motion which would affect the safe
operation of the aircraft. When natural aerodynamic damping cannot be obtained, artificial
damping must be provided to give the necessary positive dynamic stability.
®
(Positive Static)
(Negative Dynamic)
fZ
w
:2
w
o
«
--.J
0...
if)
o
Chart F illustrates the mode of a divergent oscillation. This motion is statically stable since it
tends to return to the equilibrium position. However, each subsequent return to equilibrium is
with increasing velocity such that amplitude continues to increase with time. Thus, dynamic
instability exists.
10 - 45
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
STABILITY AND CONTROL
Divergent oscillation results when energy is supplied to the motion rather than dissipated by
positive damping. An example of divergent oscillation occurs if a pilot unknowingly makes
control inputs which are near the natural frequency of the aeroplane in pitch; energy is added to
the system, negative damping exists, and Pilot Induced Oscillation ( P.LO.) results.
The existence of static stability does not guarantee the existence of dynamic stability. However,
the existence of dynamic stability implies the existence of static stability.
I,
IF AN AIRCRAFT IS STATICALLY UNSTABLE,"]
rr CANNOT BE DYNAMICALLY STABLE
~
Any aircraft must demonstrate the required degrees of static and dynamic stability. Ifthe aircraft
were allowed to have static instability with a rapid rate of divergence, it would be very difficult,
if not impossible to fly. In addition, positive dynamic stability is mandatory in certain areas to
prevent objectionable continued oscillations of the aircraft.
10 - 46
© Oxford Aviation
Servic~s
Limited
STABILITY AND CONTROL
PRINCIPLES OF FLIGHT
10.23
LONGITUDINAL DYNAMIC STABILITY.
The considerations of longitudinal dynamic stability are concerned with the time history
response ofthe aeroplane to disturbances, i.e., the variation of displacement amplitude with time
following a disturbance.
From previous definition:
a)
dynamic stability will exist when the amplitude of motion decreases with time and
b)
dynamic instability will exist if the amplitude increases with time.
An aeroplane must demonstrate positive dynamic stability for the major longitudinal motions.
In addition, the aeroplane must demonstrate a certain degree oflongitudinal stability by reducing
the amplitude of motion at a certain rate. The required degree of dynamic stability is usually
specified by the time necessary for the amplitude to reduce to one-half the original value: the
time to damp to half-amplitude.
The aeroplane in free flight has six degrees of freedom: rotation in roll, pitch, and yaw and
translation in the horizontal, vertical, and lateral directions. In the case of longitudinal dynamic
stability, the degrees of freedom can be limited to pitch rotation, plus vertical and horizontal
translation.
Since the aeroplane is usually symmetrical from left to right, there will be no need to
consider coupling between longitudinal and lateral/directional motions.
Thus, the principal variables in the longitudinal motion of an aeroplane will be:
1.
The pitch attitude of the aeroplane.
2.
The angle of attack (which will differ from the pitch attitude by the inclination of the
flight path).
3.
True airspeed (TAS)
The longitudinal dynamic stability of an aeroptane generally consists of two basic modes of
oscillation: a)
long period oscillation (phugoid)
b)
short period motion
While the longitudinal motion of the aeroplane may consist of a combination of these modes,
the characteristics of each mode are sufficiently distinct that each oscillatory tendency may be
studied separately.
10 - 47
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
10.24
STABILITY AND CONTROL
LONG PERIOD OSCILLATION (PHUGOID)
The first mode of dynamic longitudinal stability consists of a long period oscillation referred
to as the phugoid.
The phugoid or long period oscillation involves noticeable variations in :i)
pitch attitude,
ii)
altitude and
iii)
airspeed, but
iv)
nearly constant angle of attack. (not much change in load factor)
The phugoid is a gradual interchange of potential and kinetic energy about some equilibrium
airspeed and altitude. Fig. 10.43 illustrates the characteristic motion of the phugoid.
ANGLE OF ATTACK AT EACH
INSTANT ALONG THE FLIGHT
PATH IS ESSENTIALLY
CONSTANT
PERIOD
I
U
I-
o..w
o
TIME
Z=:l
- I-- --
•
wf=
(91--
Z«
«
I
U
Figure 10.43
Long Period Oscillation (Phugoid)
The period of oscillation in the phugoid is between 1 and 2 minutes. Since the pitch rate is quite
low and only negligible changes in angle of attack take place, damping of the phugoid is weak.
However, such weak damping does not necessarily have any great consequence. Since the
period of oscillation is so great, long period oscillation is easily controlled by the pilot. Due
to the nature of the phugoid, it is not necessary to make any specific aerodynamic provisions to
counteract it.
10 - 48
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
10.25
STABILITY AND CONTROL
SHORT PERIOD OSCILLATION
The second mode of dynamic longitudinal stability is the short period oscillation.
Short period oscillation involves significant changes in angle of attack (load factor), with
approximately constant speed, height and pitch attitude; it consists of rapid pitch oscillations
during which the aeroplane is constantly being restored towards equilibrium by its static stability
and the amplitude of the short period oscillations being decreased by pitch damping.
MOTION OCCURS AT ESSENTIALLY CONSTANT SPEED
_I TIME
TO DAMP TO
HALF AMPLITUDE
y:
U
«
w«
c.9LL
Zo
«
IW
ZI-I-
U--.J
c.9
Z
«
--.
TIME
--
-
- -
- --
I
-I
SHORT PERIOD
Figure 10.44
Short Period Oscillation
Short period oscillation at high dynamic pressures with large changes in angle of attack could
produce severe 'g' loads. (Large changes in load factor) .
Shown in Fig. 10.44, the second mode has relatively short periods that correspond closely with
the normal pilot response lag time, e.g., 1 or 2 seconds or less. There is the possibility that an
attempt by the pilot to forcibly damp an oscillation may actually reinforce the oscillation (PIO)
and produce instability.
Short period oscillation is not easily controlled by the pilot.
If short period oscillation occurs, release the controls; the aeroplane is designed to demonstrate
the necessary damping. Even an attempt by the pilot to hold the controls stationary when the
aeroplane is oscillating may result in a small unstable input into the control system which can
reinforce the oscillation to produce failing flight loads.
Modern large high speedjet transport aircraft are fitted with pitch dampers, which automatically
compensate for any dynamic longitudinal instability.
10 - 49
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
STABILITY AND CONTROL
Of the two modes of dynamic longitudinal stability, the short period oscillation is of greatest
importance. The short period oscillation can generate damaging flight loads due to the rapid
changes in 'g' loading, and it is adversely affected by pilot response lag (PIO).
It has been stated that the amplitude of the oscillations are decreased by pitch damping, so the
problems of dynamic stability can become acute under the conditions of flight where reduced
aerodynamic damping occurs.
High altitude, and consequently low density (high TAS), reduces aerodynamic damping, as
detailed in paragraph 10.15.
r
I
~
DYNAMIC STABILITY IS REDUCED AT HIGH ALTITUDE
DUE TO REDUCED AERODYNAMIC DAMPING
~
~
10 - 50
© Oxford Aviation Services Limited
STABILITY AND CONTROL
PRINCIPLES OF FLIGHT
10.26
DIRECTIONAL STABILITY AND CONTROL
The directional stability of an aeroplane is essentially the "weathercock" stability and involves
moments about the Normal axis and their relationship with yaw or sideslip angle. An aeroplane
which has static directional stability will tend to return to equilibrium when subjected to some
disturbance. Evidence of static directional stability would be the development of yawing
moments which tend to restore the aeroplane to equilibrium.
DEFINITIONS
The axis system of an aeroplane defines a positive yawing moment, N, as a moment about the
normal axis which tends to rotate the nose to the right. As in other aerodynamic considerations,
it is convenient to consider yawing moments in the coefficient form so that static stability can
be evaluated independent of weight, altitude, speed, etc. The yawing moment, N, is defined in
the coefficient form by the following equation:
N = Cn
Q
S
b
or
N
Q S
b
where,
N
Q
S
b
Cn
= yawing moment (positive to the right)
= dynamic pressure
= wing area
= wing span
= yawing moment coefficient (positive to the right)
The yawing moment coefficient, en is based on the wing dimensions Sand b as the wing is
the characteristic surface of the aeroplane.
10 - 51
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
10.27
STABILITY AND CONTROL
SIDESLIP ANGLE
The sideslip angle relates the displacement of the aeroplane centerline from the relative airflow.
Sideslip angle is provided the symbol ~ (beta) and is positive when the relative wind is displaced
to the right of the aeroplane centerline. Fig. 10.45 illustrates the definitions of sideslip angle.
RELATIVE
AIRFLOW
~~---I
SIDESLIP
ANGLE
~ // '~
I
+ N,
Figure 10.45
A YAW TO THE LEFT GIVES
A SIDESLIP TO THE RIGHT
YAWING MOMENT
Sideslip Angle
(~)
The sideslip angle, ~ , is essentially the "directional angle of attack" of the aeroplane and is the
primary reference in directional stability as well as lateral stability considerations. Static
directional stability of the aeroplane is appreciated by response to sideslip.
10 - 52
© Oxford Aviation Services Limited
STABILITY AND CONTROL
PRINCIPLES OF FLIGHT
i
YAWING MOMENT
COEFFICIENT, C n .
STABLE- :
l
r
- -- - - - -
NEUTRAL
,
-
........ ........
---~-
SIDESLIP ANGLE,
........
. . "\. . . .
UNSTABLE \
------~
+ !3
!3
I
"",-
II'
i~
(
;
L.
C::c
Figure 10.46
10.28
STATIC DIRECTIONAL STABILITY
Static directional stability can be illustrated by a graph of yawing moment coefficient, Cn versus
sideslip angle, ~ , such as shown in Fig 10.46. When the aeroplane is subject to a positive
sideslip angle, static directional stability will beevident if a positive yawing moment coefficient
results. Thus, when the relative airflow comes from the right (+ ~ ) a yawing moment to the right
(+Cn) should be created which tends to "weathercock" the aeroplane and return the nose into
the wind. Static directional stability will exist when the curve ofCn versus ~ has a positive slope
and the degree of stability will be a function of the slope of this curve. If the curve has zero
slope, there is no tendency to return to equilibrium and neutral static directional stability exists.
When the curve ofCn versus ~ has a negative slope, the yawing moments developed by sideslip
tend to diverge rather than restore and static directional instability exists.
10 - 53
© Oxford Aviation Services Limited
STABILITY AND CONTROL
PRINCIPLES OF FLIGHT
t
en
+
+
Figure 10.47
Fig. 10.47 illustrates the fact that the instantaneous slope of the curve of en versus
describe the static directional stability of the aeroplane.
~
will
a)
At small angles of sideslip a strong positive slope depicts strong directional stability.
b)
Large angles of sideslip produce zero slope and neutral stability.
c)
At very high sideslip the negative slope of the curve indicates directional instability.
This decay of directional stability with increased sideslip is not an unusual condition. However,
directional instability should not occur at the angles of sideslip of ordinary flight conditions.
Static directional stability must be in evidence for all the critical conditions of flight. Generally,
good directional stability is a fundamental quality directly affecting the pilots' impression of an
aeroplane.
10.29
CONTRIBUTION OF THE AEROPLANE COMPONENTS.
Because the contribution of each component depends upon and is related to the others, it is
necessary to study each separately.
FUSELAGE: The fuselage is destabilizing, Fig. 10.48. It is an aerodynamic body and a
condition of sideslip can be likened to an "angle of attack", so that an aerodynamic side force
is created. This side force acts through the fuselage aerodynamic centre (AC), which is close
to the quarter-length point. If this aerodynamic centre is ahead of aircraft centre of gravity, as
is usually the case, the effect is de-stabilizing.
10 - 54
© Oxford Aviation Services Limited
STABILITY AND CONTROL
PRINCIPLES OF FLIGHT
FUSELAGE (Plan View)
FORCE
FLIGHT PATH
UNSTABLE
YAWING
MOMENT
Figure 10.48
DORSAL AND VENTRAL FINS: To overcome the instability in the fuselage it is possible
to incorporate into the overall design, dorsal or ventral fins. A dorsal fin is a small aerofoil, of
very low aspect ratio, mounted on top of the fuselage near the rear. A ventral fin is mounted
below. Such fins are shown in Fig. 10.49.
DORSAL FIN
I
>
\
VENTRAL FIN
Figure 10.49
If the aircraft is yawed to the right the dorsal and ventral fins will create a side force to the right.
The line of action of this force is well aft of the aircraft eG, giving a yawing moment to the left
(a stabi lising effect). However, at small angles of yaw they are ineffective.
10 - 55
© Oxford Aviation Services Limited
STABILITY AND CONTROL
PRINCIPLES OF FLIGHT
The side force created by dorsal and ventral fins at small sideslip angles will be very small
because:a)
the dorsal and ventral fins are at a low angle of attack,
b)
they have a small surface area, and
c)
their aspect ratio is very low, resulting in small lift-curve slope. Fig. 10.50.
HIGH ASPECT
RATIO
LOW ASPECT
RATIO
(or sweepback)
Figure 10.50
When fitted with dorsal and ventral fins a fuselage which is unstable in yaw, will remain
unstable at low sideslip angles. Dorsal and ventral fins become more effective at relatively high
sideslip angles. Due to their low aspect ratio they do not tend to stall at any sideslip angle which
an aircraft is likely to experience in service.
The effectiveness of dorsal and ventral fins increases with increasing sideslip angle, so that the
combination of a fuselage with dorsal or ventral fin is stable at large sideslip angles.
While dorsal and ventral fins contribute in exactly the same way to directional static stability,
a dorsal fin contributes positively to lateral static stability, while a ventral fin is destabilising in
this mode, as will be demonstrated later. For this reason, the dorsal fin is much more common.
Figure 10.51
10 - 56
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
STABILITY AND CONTROL
RELATIVE
AIRFLOW
Figure 10.52
FIN: The fin (vertical stabiliser) is the major source of directional stability for the aeroplane.
As shown in Fig. lO.52, in a sideslip the fin will experience a change in angle of attack. The
change in lift (side force) on the fin creates a yawing moment about the centre of gravity which
tends to yaw the aeroplane into the relative airflow. The magnitude of the fin contribution to
static directional stability depends on both the change in fin lift and the fin moment arm.
Clearly, the fin moment arm is a powerful factor.
The contribution of the fin to directional stability depends on its ability to produce changes in
lift, or side force , for a given change in sideslip angle. The contribution of the fin is a direct
function of its area. The required directional stability may be obtained by increasing the fin area.
However, increased surface area has the obvious disadvantage of increased parasite drag.
The lift curve slope of the fin determines how sensitive the surface is to change in angle of
attack. While it is desirable to have a high lift curve slope for the fin, a high aspect ratio surface
is not necessarily practical or desirable - bending, lower stalling angle (Fig. 10.51), hangar roof
clearance, etc. The stall angle of the surface must be sufficiently great to prevent stall and
subsequent loss of effectiveness at expected sideslip angles. (sweepback or low aspect ratio
increases the stalling angle of attack of the fin). '
The flow field in which the fin operates is affected by other components of the aeroplane as well
as power effects. The dynamic pressure at the fin could depend on the slipstream of a propeller
or the boundary layer of the fuselage. Also, the local flow direction at the fin is influenced by
the wing wake, fuselage crossflow, induced flow of the horizontal tail, or the direction of
slipstream from a propeller. Each ofthese factors must be considered as possibly affecting the
contribution of the fin to directional stability.
A high mounted tailplane CT' - tail) makes the fin more effective by acting as an "end plate".
10 - 57
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
STABILITY AND CONTROL
The side force on the fin may still be relatively small compared to that on the fuselage, which
is destabilising, but because its line of action is far aft of the CG, the yawing moment it creates
is relatively large, and gives overall stability to the fuselage-fin combination. The principle
behind the effect of the fin as a stabiliser is just the same as in the case of the dorsal or ventral
fin. However, because it is much larger, and in particular, has a much higher aspect ratio, it is
effective at low angles of sideslip. It remains effective until the angle of sideslip is such that the
fin angle of attack approaches its stalling angle, but above this value the side force on the fin
decreases with increasing sideslip angle, and the fin ceases to be effective as a stabiliser. It is
at this point that the dorsal or ventral fin becomes important. Because it stalls at a very much
higher angle of attack, it takes over the stabilising role of the fin at large angles of sideslip.
WING and NACELLES: The contribution of the wing to static directional stability is usually
small.
a)
The contribution of a straight wing alone is usually negligible.
b)
Sweep back produces a stabilizing effect, which increases with increase in CL (i.e. at
lower lAS).
c)
Engine nacelles on the wings produce a contribution that will depend on such factors as
their size and position and the shape of the wing planform. On a straight wing, they
usually produce a de stabilising effect.
A swept wing provides a stable contribution depending on the amount of sweepback but the
contribution is relatively weak when compared with other components. Consider a sideslipping
swept wing, as illustrated in Fig. 10.53.
NORMAL
NORMAL
Figure 10.53
10 - 58
© Oxford Aviation Services Limited
STABILITY AND CONTROL
PRINCIPLES OF FLIGHT
The inclination of the forward, right, wing to the relative airflow is greater than that of the
rearward wing, so there is more lift, and hence more induced drag on the right side, (the
influence of increased lift on the forward wing will be explained when lateral static stability is
considered). The result of this discrepancy in drag on the two sides of the wing is a yawing
moment to the right, which tends to eliminate the sideslip . This is a stabilising effect, and may
be important if the sweepback angle is quite large.
Fig. 10.54 illustrates a typical buildup of the directional stability of an aeroplane by separating
the contribution of the fuselage and fin. As shown by the graph of en versus ~ , the contribution
of the fuselage is destabilising but the instability decreases at large sideslip angles. The
contribution ofthe fin alone is highly stabilising up to the point where the surface begins to stall.
The contribution of the fin must be large enough so that the complete aeroplane (wing-fuselagefin combination) exhibits the required degree of stability.
AEROPLANE WITH
DORSAL FIN
r
ADDED
STALL
en
+
---
FIN
ALONE
'":: COMPLETE
AEROPLANE
+/9
Figure 10.54
The dorsal fin has a powerful effect on preserving the directional stability at large angles of
sideslip which would produce stall of the fin.
The addition of a dorsal fin to the aeroplane will reduce the decay of directional stability at high
sideslip in two ways.
a)
The least obvious but most important effect is a large increase in the fuselage stability
at large sideslip angles.
b)
In addition, the effective aspect ratio of the fin is reduced which increases the stall angle
for the surface.
By this twofold effect, the addition of the dorsal fin is a very useful device. The decreased lift
curve slope of a sweptback fin will also decrease the tendency for the fin to stall at high sideslip
angles.
10 - 59
© Oxford Aviation Services Limited
STABILITY AND CONTROL
PRINCIPLES OF FLIGHT
POWER EFFECT : The effects of power on static directional stability are similar to the power
effects on static longitudinal stability. The direct effect is confined to the normal force at the
propeller plane and, of course, is destabilising when the propeller is located ahead of the CG.
In addition, the air in the slipstream behind a propeller spirals around the fuselage , and this
results in a sidewash at the fin (from the left with a clockwise rotating propeller). The indirect
effects of power induced velocities and flow direction changes at the fin (spiral slipstream effect)
are quite significant for the propeller driven aeroplane and can produce large directional trim
changes. As in the longitudinal case, the indirect effects are negligible for the jet powered
aeroplane.
The contribution of the direct and indirect power effects to static directional stability is greatest
for the propeller powered aeroplane and usually slight for the jet powered aeroplane. In either
case, the general effect of power is destabilising and the greatest contribution will occur at
high power and low dynamic pressure.
CRITICAL CONDITIONS : The most critical conditions of static directional stability are usually
the combination of several separate effects. The combination which produces the most critical
condition is much dependent upon the type of aeroplane. In addition, there exists a coupling
oflateral and directional effects such that the required degree of static directional stability
may be determined by some of these coupled conditions.
CENTRE OF GRAVITY POSITION: Centre of gravity position has a relatively negligible
effect on static directional stability. The usual range ofCG position on any aeroplane is set by
the limits oflongitudinal stability and control. Within this limiting range of CG position, no
significant changes take place in the contribution of the vertical tail, fuselage, nacelles, etc.
Hence, static directional stability is essentially unaffected by the variation ofCG position within
the longitudinal limits.
l- e
Z 0
W
:::2: I-
t
LOW ANGLE
OF ATTACK
0 z
:::2: w
C)
0
Z
LL
LL
S
«
>-
w
0
0
..................
~-
----------- .... "
HIGH' ANGLE
OF ATTACK
"......
SIDESLIP ANGLE,
f3
Figure 10.55
10 - 60
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
STABILITY AND CONTROL
HIGH ANGLE OF ATTACK: When the aeroplane is at a high angle of attack a decrease in
static directional stability can be anticipated. As shown by Fig. 10.55, a high angle of attack
reduces the stable slope of the curve ofCn versus p. The decrease in static directional stability
is due in great part to the reduction in the contribution of the fin. At high angles of attack, the
effectiveness of the fin is reduced because of increase in the fuselage boundary layer at the fin
location. The decay of directional stability with angle of attack is most significant for an
aeroplane with sweepback since this configuration requires a high angle of attack to achieve high
lift coefficients.
VENTRAL
FIN
Figure 10.56
Ventral Fin
VENTRAL FIN: Ventral fins may be added as an additional contribution to directional stability,
Fig. 10.56. Landing clearance requirements may limit their size, require them to be retractable,
or require two smaller ventral fins to be fitted instead of one large one.
The most critical demands of static directional stability will occur from some combination ofthe
following effects :
I)
high angle of sideslip
2)
high power at low airspeed
3)
high angle of attack
4)
high Mach number
The propeller powered aeroplane may have such considerable power effects that the critical
conditions may occur at low speed while the effect of high Mach numbers may produce the
critical conditions for the typical transonic, jet powered aeroplane. In addition, the coupling of
lateral and directional effects may require prescribed degrees of directional stability.
10 - 61
© Oxford Aviation Services Limited
STABILITY AND CONTROL
PRINCIPLES OF FLIGHT
10.30
LATERAL STABILITY AND CONTROL
The static lateral stability of an aeroplane involves consideration of rolling moments due to
sideslip. If an aeroplane has favourable rolling moment due to sideslip, a lateral displacement
from wing level flight produces sideslip and the sideslip creates rolling moments tending to
return the aeroplane to wing level flight. By this action, static lateral stability will be evident.
Of course, a sideslip will produce yawing moments depending on the nature of the static
directional stability but the consideration of static lateral stability will involve only the
relationship of rolling moments and sideslip.
DEFINITIONS
The axis system of an aeroplane defines a positive rolling, L, as a moment about the longitudinal
axis which tends to rotate the right wing down. As in other aerodynamic considerations, it is
convenient to consider rolling moments in the coefficient form so that lateral stability can be
evaluated independent of weight, altitude, speeds, etc. The rolling moment, L, is defined in the
coefficient form by the following equation:
L
=
C1
Q
S
b
or
L
Q S
b
where,
L
Q
S
b
C1
= rolling moment (positive to the right)
= dynamic pressure
= wing area
= wing span
= rolling moment coefficient (positive to the right)
The angle of sideslip, ~ has been defined previously as the angle between the aeroplane
centerline and the relative wind and is positive when the relative wind is to the right of the
centerline.
10 - 62
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
10.31
STABILITY AND CONTROL
STATIC LATERAL STABILITY
Static lateral stability can be illustrated by a graph of rolling moment coefficient, C I, versus
sideslip angle, ~ , such as shown in Fig. 10.57. When the aeroplane is subject to a positive
sideslip angle, lateral stability will be evident if a negative rolling moment coefficient results.
Thus, when the relative airflow comes from the right (+~), a rolling moment to the left (-C I)
should be created which tends to roll the aeroplane to the left. Lateral stability will exist when
the curve of C1 versus ~ has a negative slope and the degree of stability will be a function of the
slope of this curve. Ifthe slope of the curve is zero, neutral lateral stability exists; ifthe slope
is positive lateral instability is present.
ROLLING MOMENT COEFFICIENT
+
UNSTABLE
~~
~~
~~
~~
~~
~~
,
~~
~.
~~~~~~~ NEUTRAL7
~~
~~
~~
~~
~~
~~
~~
+
~~
SIDESLIP ANGLE,
j3
Figure 10.57
10 - 63
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
STABILITY AND CONTROL
It is desirable to have static lateral stability (favourable roll due to sideslip), Fig. 10.58.
However, the required magnitude oflateral stability is determined by many factors. Excessive
roll due to sideslip complicates crosswind take-off and landing and may lead to undesirable
oscillatory coupling with the directional motion of the aeroplane. In addition, high lateral
stability may combine with adverse yaw to hinder rolling performance. Generally, good
handling qualities are obtained with a relatively 1ight, or weak positive, lateral stability.
STABLE ROLL DUE
TO SIDESLIP , .. - ......
I
: h/
-.:0:.--
~
~ .'
"':!... .::--./
NEUTRAL
UNSTABLE ROLL
DUE TO SIDESLIP
Figure 10.58
Static Lateral Stability
10 - 64
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STABILITY AND CONTROL
PRINCIPLES OF FLIGHT
10.32
CONTRIBUTION OF THE AEROPLANE COMPONENTS.
In order to appreciate the development of lateral stability in an aeroplane, each of the
components which contribute must be inspected. There will be interference between the
components, which will alter the contribution to stability of each component on the
aeroplane.
EFFECTIVE INCREASE IN
LIFT DUE TO SIDESLIP
EFFECTIVE DECREASE IN
LIFT DUE TO SIDESLIP
Figure 10.59
Geometric Dihedral
WING: The principal surface contributing to the lateral stability of an aeroplane is the wing.
The effect of *geometric dihedral is a powerful contribution to lateral stability.
As shown in Fig. 10.59, a wing with geometric dihedral will develop stable rolling moments with
sideslip. If the relative wind comes from the side, the wing into the wind is subject to an
increase in angle of attack and develops an increase in lift. The wing away from the wind is
subj ect to a decrease in angle of attack and develops a decrease in lift. The changes in lift gives
a rolling moment tending to raise the into-wind wing, hence geometric dihedral contributes
a stable roll due to sideslip.
Since geometric dihedral is so powerful in producing lateral stability it is taken as a common
denominator of the lateral stability contribution of all other components. Generally, the
contribution of wing position, flaps , power, etc., is expressed as "DIHEDRAL EFFECT".
*Geometric Dihedral:
The angle between the plane of each wing and the horizontal, when the aircraft
is unbanked and level; positive when the wing lies above the horizontal, as in
Fig. 10.59. Negative geometric dihedral is used on some aircraft, and is known
as anhedral.
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PRINCIPLES OF FLIGHT
STABILITY AND CONTROL
WING POSITION: The contribution of the fuselage alone is usually quite small; depending
on the location of the resultant aerodynamic side force on the fuselage .
However, the effect of the wing - fuselage - tail combination is significant since the vertical
placement of the wing on the fuselage can greatly affect the combination. A wing located at the
mid wing position will generally exhibit a "dihedral effect" no different from that of the wing
alone.
Fig. 10.60 illustrates the effect of wing position on static lateral stability.
a)
A low wing position gives an unstable contribution. The direction of relative airflow
decreases the effective angle of attack of the wing into wind and increases the effective
angle of attack of the wing out of wind - tending to increase the rolling moment.
b)
A high wing location gives a stable contribution. The direction of relative airflow
increases the effective angle of attack of the wing into wind and decreases the effective
angle of attack of the wing out of wind, tending to decrease the rolling moment.
SIDESLIP
\--
HIGH WING PosrnON
\
Figure 10.60
Wing - Fuselage Interference Effect
The magnitude of "dihedral effect" contributed by the vertical position of the wing is large and
may require a noticeable dihedral angle for the low wing configuration. A high wing position,
on the other hand, usually requires no geometric dihedral at all.
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STABILITY AND CONTROL
PRINCIPLES OF FLIGHT
SWEEPBACK: The contribution of sweep back to "dihedral effect" is important because of the
nature of the contribution. As shown in Figs. 10.61 and 10.62, if the wing is at a positive lift
coefficient, the wing into the wind has less sweep and an increase in lift and the wing out of the
wind has more sweep and a decrease in lift; a negative rolling moment will be generated, tending
to roll the wings towards level. In this manner the swept back wing contributes a positive
"dihedral effect". (A swept forward wing would give a negative dihedral effect).
+/3
1-
INCREASED ,
EFFECTIVE
SWEEP
V~
v---DECREASED
EFFECTIVE
SWEEP
Figure 10.61
The Effect of Sweepback
The contribution of sweepback to "dihedral effect" is proportional to the wing lift coefficient as
well as the angle of sweepback. Because high speed flight requires a large amount of
sweepback, an excessively high "dihedral effect" will be present at low speeds (high C L ). An
aircraft with a swept back wing requires less geometric dihedral than a straight wing.
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STABILITY AND CONTROL
PRINCIPLES OF FLIGHT
/
LEADING WING
IN SIDESLIP
/
DIFFERENCE IN
LIFT ON THE
TWO WINGS
/
/
ZERO
SIDESLIP
/
/
/
/'
/'
TRAILING WING
IN SIDESLIP
/
/
/
/'
/'
'
----~----------------------------------,~
o
HIGH
SPEED
LOW
SPEED
Figure 10.62 Effect of speed on 'dihedral effect' of swept wing
The fin can provide a small "dihedral effect" contribution, Fig. 10.63. If the fin is large, the side
force produced by sideslip may produce a rolling moment as well as the important yawing
moment contribution. The fin contribution to purely lateral static stability, is usually very
small.
SIvlALL STABILISING
ROLLING MOMENT IN
SIDESLIP
~
RELATIVE AIRFLOW
Figure 10.63 Fin Contribution
The ventral fin, being below the aircraft CG, has a negative influence on lateral static stability,
as illustrated in Fig. 10.63a.
SIvlALL DE - STABILISING
ROLLING MOMENT
IN SIDESLIP
\
VENTRAL
FIN
~~
"
RELATIVE AIRFLOW
Figure 10.63a Ventral Fin Contribution
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STABILITY AND CONTROL
PRINCIPLES OF FLIGHT
Generally, the "dihedral effect" should not be too great since high roll due to sideslip can create
certain problems.
Excessive "dihedral effect" can lead to "Dutch roll," difficult rudder coordination in rolling
manoeuvres, or place extreme demands for lateral control power during crosswind takeoff and
landing. If the airplane demonstrates satisfactory "dihedral effect" during cruise, certain
exceptions can be tolerated when the airplane is in the takeoff and landing configuration. Since
the effects offlaps and power are destabilizing and reduce the "dihedral effect", a certain amount
of negative "dihedral effect" may be possible due to these sources.
REDUCED ARM
Figure 10.64
Partial Span Flaps Reduce Lateral Stability
The deflection of flaps causes the inboard sections of the wing to become relatively more
effective and these sections have a small spanwise moment arm, Fig. 10.64. Therefore, the
changes in wing lift due to sideslip occur closer inboard and the dihedral effect is reduced.
The effect of power on "dihedral effect" is negligible for the jet aeroplane but considerable for
the propeller driven aeroplane. The propeller slipstream at high power and low airspeed makes
the inboard wing sections much more effective and reduces the dihedral effect.
The reduction in "dihedral effect" is most critical when the flap and power effects are combined,
e.g., the propeller driven aeroplane in a power-on approach.
With certain exceptions during the conditions of landing and takeoff, the "dihedral effect" or
lateral stability should be positive but light. The problems created by excessive "dihedral effect"
are considerable and difficult to contend with. Lateral stability will be evident to a pilot by stick
forces and displacements required to maintain sideslip . Positive stick force stability will be
evident by stick forces required in the direction of the controlled sideslip.
CONCLUSION: the designer is faced with a dilemma. An aircraft is given sweepback to
increase the speed at which it can operate, but a by-product of sweep back is static lateral
stability. A sweptback wing requires much less geometric dihedral than a straight wing. If a
requirement also exists for the wing to be mounted on top of the fuselage, an additional "dihedral
effect" is present. A high mounted and sweptback wing would give excessive "dihedral effect",
so anhedral is used to reduce "dihedral effect" to the required level.
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STABILITY AND CONTROL
PRINCIPLES OF FLIGHT
10.33
LATERAL DYNAMIC EFFECTS
Previous discussion has separated the lateral and directional response of the aeroplane to
sideslip, in order to give each the required detailed study.
However, when an aeroplane is placed in a sideslip, the lateral and directional response will be
coupled, i.e. sideslip will simultaneously produce a rolling and a yawing moment.
The principal effects which determine the lateral dynamic characteristics of an aeroplane are:
10.34
1)
Rolling moment due to sideslip or "dihedral effect" (static lateral stability).
2)
Yawing moment due to sideslip or static directional stability.
SPIRAL DIVERGENCE
Spiral divergence will exist when static directional stability is very large when compared
to the "dihedral effect".
The character of spiral divergence is not violent. The aeroplane, when disturbed from the
equilibrium oflevel flight, begins a slow spiral which gradually increases to a spiral dive. When
a small sideslip is introduced, the strong directional stability tends to restore the nose into the
wind while the relatively weak "dihedral effect" lags in restoring the aeroplane laterally. The
rate of divergence in the spiral motion is usually so gradual that the pilot can control the
tendency without difficulty.
10.35
DUTCH ROLL
Dutch roll will occur when the "dihedral effect" is large when compared to static
directional stability.
Dutch roll is a coupled lateral and directional oscillation which is objectionable because of the
oscillatory nature.
When a yaw is introduced, the strong "dihedral effect" will roll the aircraft due to the lift
increase on the wing into wind. The increased. induced drag on the rising wing will yaw the
aircraft in the opposite direction, reversing the coupled oscillations.
Aircraft with a tendency to Dutch Roll are fitted with a Yaw Damper. This automatically
displaces the rudder proportional to the rate of yaw to damp-out the oscillations.
If the Yaw Damper fails in flight, it is recommended that the ailerons be used by the pilot to
damp-out Dutch Roll. Because of the response lag, if the pilot uses the rudder, pilot induced
oscillation (PIO) will result and the Dutch Roll may very quickly become divergent, leading to
loss of control.
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PRINCIPLES OF FLIGHT
STABILITY AND CONTROL
Dutch roll is objectionable, and spiral divergence is tolerable if the rate of divergence is low.
For this reason the "dihedral effect" should be no more than that required for satisfactory lateral
stability.
If the static directional stability is made adequate to prevent objectionable Dutch roll, this will
automatically be sufficient to prevent directional divergence. Since the more important handling
qualities are a result of high static directional stability and minimum necessary "dihedral effect",
most aeroplanes demonstrate a mild spiral tendency. As previously mentioned, a weak spiral
tendency is of little concern to the pilot and certainly preferable to Dutch roll.
The contribution of sweepback to the lateral dynamics of an aeroplane is significant. Since the
"dihedral effect" from sweepback is a function of 1ift coefficient, the dynamic characteristics
may vary throughout the flight speed range.
When the swept wing aeroplane is at low CL the "dihedral effect" is small and the spiral
tendency may be apparent. When the swept wing aeroplane is at high CL the "dihedral effect"
is increased and the Dutch Roll oscillatory tendency is increased.
10.36
PILOT INDUCED OSCILLATIONS (PIO)
Certain undesirable motions may occur due to inadvertent action on the controls. This can occur
about any of the axes, but the most important condition exists with the short period longitudinal
motion of the aeroplane where pilot control system response lag can produce an unstable
oscillation. The coupling possible in the pilot/control system/aeroplane combination is capable
of producing damaging flight loads and loss of control of the aeroplane.
When the normal human response lag and control system lag are coupled with the aeroplane
motion, inadvertent control reactions by the pilot may furnish negative damping to the oscillatory
motion and dynamic instability will exist.
Since short period motion is of relatively high frequency, the amplitude of the pitching
oscillation can reach dangerous proportions in an unbelievably short time.
When pilot induced oscillation is encountered, the most effective solution is an immediate
release of the controls. Any attempt to forcibly damp the oscillation simply continues the
excitation and amplifies the oscillation. Freeing the controls removes the unstable (but
inadvertent) excitation and allows the aeroplane to recover by virtue of its inherent dynamic
stability.
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PRINCIPLES OF FLIGHT
10.37
STABILITY AND CONTROL
HIGH MACH NUMBERS
Generally, flight at high Mach numbers will take place at high altitude, hence the effect of high
altitude must be separated for study. Aerodynamic damping is due to moments created by
pitching, rolling, or yawing of the aircraft. These moments are derived from the changes in
angles of attack of the tail, wing and fin surfaces with angular rotation (see Fig. 10.35).
Higher TAS common to high altitude flight reduces the angle of attack changes and reduces
aerodynamic damping. In fact, aerodynamic damping is proportional to the square root of the
relative density, similar to the proportion of True Air Speed to Equivalent Air Speed. Thus, at
an ISA altitude of 40,000 ft., aerodynamic damping would be reduced to one-half the ISA sea
level value.
10.38
MACH TRIM
As speed increases beyond the Critical Mach number (MeRIT)' shock wave formation at the root
of a swept back wing will:
a)
reduce lift forward of the CG, and
b)
reduce downwash at the tailplane.
Together, these factors will generate a nose down pitching moment. At high Mach numbers, an
aircraft will become unstable with respect to speed; instead of an increasing push force being
required as speed increases, a pull force becomes necessary to prevent the aircraft accelerating
further. This is potentially dangerous. A small increase in Mach number will give a nose down
pitch which will further increase the Mach number. This in turn leads to a further increase in
the nose down pitching moment. This unfavourable high speed characteristic, known as "Mach
Tuck", "High Speed Tuck" or "Tuck Under" would restrict the maximum operating speed of
a modem high speed jet transport aircraft.
To maintain the required stick force gradient at high Mach numbers, a Mach trim system
must be fitted. This device, sensitive to Mach number, may:
a)
deflect the elevator up
b)
decrease the incidence of the variable incidence trimming tailplane or
c)
move the CG rearwards by transferring fuel from the wings to a rear trim tank
by an amount greater than that required merely to compensate for the trim change. This ensures
the required stick force gradient is maintained in the cruise at high Mach numbers.
Whichever method of trim is used by a particular manufacturer, a Mach trim system will adjust
longitudinal trim and operates only at high Mach numbers.
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PRINCIPLES OF FLIGHT
STABILITY AND CONTROL
KEY FACTS 2 - Self Study (Insert the missing words, with reference to the preceding paragraphs).
Positive static longitudinal stability is indicated by a _ _ _ slope ofCM versus C L • The degree
of
longitudinal stability is indicated by the
of the curve. (Para. 10.8).
The net pitching moment about the
axis is due to the contribution of each of the
component
acting in their appropriate _ _ fields. (Para. 10.9).
In most cases, the contribution of the fuselage and nacelles is _ _ _ _ _ . (Para. 10.9).
(Para. 10.9) Noticeable changes in static stability can occur at high C L (low speed) if:
a)
the aeroplane has _ _ _ __
b)
there is a large contribution of ' - - - effect', or
c)
there are significant changes in
at the horizontal tail,
The horizontal tail usually provides the ____ stabilising influence of all the components of
the aeroplane. (Para. 10.9).
_____ decreases static longitudinal stability. (Para. 10.9).
If the thrust line is below the CG, a thrust increase will produce a _ _ _ or nose
and the effect is
. (Para. 10.12).
moment
High lift devices tend to _ _ _ _ downwash at the tail and _ _ the dynamic pressure at the
. (Para. 10.13).
tail, both of which are
An increase in TAS, for a given pitching velocity, _ _ _ _ aerodynamic damping.
(Para. 10.15).
The aeroplane with positive manoeuvring stability should demonstrate a steady - - - in stick
force with
in load factor or "_". (Para. 10.16).
The stick force gradient must not be excessively __ or the aeroplane will be difficult and tiring
to manoeuver. Also, the stick force gradient must not be too _ _ or the aeroplane may be
overstressed inadvertently when light control forces exist. (Para. 10.16).
When the aeroplane has high static stability, the manoeuvring stability will be _ _ and a __
stick force gradient will result. The
CG limit could be set to prevent an excessively
high manoeuvring stick force gradient. As the CG moves aft, the stick force gradient - - - with
manoeuvring stability and the
limit of stick force gradient may be
reached. (Para. 10.16).
the change in tail angle of attack for a given pitching
At high altitudes, the high TAS
velocity and
the pitch damping. Thus, a decrease in manoeuvring stick force stability
can be expected with
altitude. (Para. 10.16).
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STABILITY AND CONTROL
PRINCIPLES OF FLIGHT
A flying control system may employ
springs, _ _ springs or _ _ weights to provide
satisfactory control forces throughout the speed, CG and altitude range of an aircraft.
(Para. 10.17).
While static stability is concerned with the initial tendency of an aircraft to return to equilibrium,
dynamic stability is defined by the resulting
with
(Para. 10.22).
An aircraft will demonstrate positive dynamic stability if the _ _ _ _ of motion _ _ __
with time. (Para. 10.22).
When natural aerodynamic damping cannot be obtained, _ _ _ _ damping must be provided
to give the necessary positive dynamic stability. (Para. 10.22).
(Para. 10.23) The longitudinal dynamic stability of an aeroplane generally consists of two basic
modes of oscillation:a)
_ _ period (phugoid)
b)
period
The phugoid oscillation occurs with nearly constant ___ of _ __
(Para. 10.24).
The period of oscillation is so great, the pilot is easily able to counteract _____ oscillation.
(Para. 10.24).
Short period oscillation involves significant changes in ___ of _ _ _ . (Para. 10.25).
Short period oscillation is ______ controlled by the pilot. (Para. 10.25).
The problems of dynamic stability can become acute at _ _ altitude because of _ _ __
aerodynamic
. (Para. 10.25).
To overcome the directional instability in the fuselage it is possible to incorporate into the
overall design,
or
fins. (Para. 10.29).
The _ _ is the major source of directional stability for the aeroplane. (Para. 10.29).
A _
- tail makes the fin more effective by acting as an " __ plate". (Para. 10.29).
Because the
fin stalls at a very much higher angle of attack, it takes over the stabilising
role of the fin at large angles of sideslip. (Para. 10.29).
_ _ _ _ _ produces a directional stabilising effect, which increases with increase in CL •
(Para. 10.29).
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STABILITY AND CONTROL
PRINCIPLES OF FLIGHT
_ _ _ _ fins increase directional stability at _ _ angles of attack. Landing clearance
requirements may limit their size, require them to be retractable, or require two smaller ventral
fins to be fitted instead of one large one. (Para. 10.29).
Generally, good handling qualities are obtained with a relatively _ _, or _ _ positive, lateral
stability. (Para. 10.31).
The principal surface contributing to the lateral stability of an aeroplane is the _ _ . The effect
of geometric
is a powerful contribution to lateral stability. (Para. 10.31).
A low wing position gives an ____ contribution to static lateral stability. (Para. 10.32).
A _ _ wing location gives a stable contribution to static lateral stability. (Para. 10.32).
The magnitude of "dihedral effect" contributed by the vertical position of the wing is _ _ and
may require a noticeable dihedral angle for the ___ wing configuration. A high wing position,
on the other hand, usually requires _ geometric
at all. (Para. 10.32).
The
back wing contributes a positive "dihedral effect". (Para. 10.32).
An aircraft with a swept back wing requires _ _ geometric dihedral than a straight wing.
(Para. 10.32).
The fin contribution to purely lateral static stability, is usually very ___ . (Para. 10.32).
Excessive "dihedral effect" can lead to"- - - - roll," difficult rudder coordination in - - - manoeuvres, or place extreme demands for
control power during crosswind takeoff and
landing. (Para. 10.32).
Deploying partial span flaps gives a _ _ _ _ dihedral effect. (Para. 10.32).
A sweptback wing requires much less geometric dihedral than a straight wing. If a requirement
also exists for the wing to be mounted on top of the fuselage, an additional "dihedral effect" is
present. A high mounted and sweptback wing would give excessive "dihedral effect", so
_ _ _ _ is used to reduce "dihedral effect" to the required level. (Para. 10.32).
When an aeroplane is placed in a sideslip, the lateral and directional response will be _ _ __
i.e. sideslip will simultaneously produce a
and a
moment. (Para. 10.33).
Spiral divergence will exist when static directional stability is very _ _ when compared to the
"dihedral effect". (Para. 10.34).
The rate of divergence in the spiral motion is usually so ____ that the pilot can control the
tendency without
. (Para. 10.34).
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STABILITY AND CONTROL
PRINCIPLES OF FLIGHT
Dutch roll will occur when the "dihedral effect" is ___ when compared to static directional
stability. (Para. 10.35).
Aircraft which Dutch Roll are fitted with a _ _ Damper. This automatically displaces the
rudder proportional to the _ _ of yaw to damp-out the oscillations. (Para. 10.35).
If the Yaw Damper fails in flight, it is recommended that the ____ be used by the pilot to
damp-out Dutch Roll. (Para. 10.35).
If the pilot uses the ____ , pilot induced oscillation (PIO) will result and the Dutch Roll may
. (Para. 10.35).
very quickly become
, leading to loss of
When the swept wing aeroplane is at low CL the "dihedral effect" is small and the _ __
tendency may be apparent. When the swept wing aeroplane is at high CL the "dihedral effect"
is increased and the
_ _ oscillatory tendency is increased. (Para. 10.35).
When pilot induced oscillation is encountered, the most effective solution is an immediate
_ _ _ of the controls. Any attempt to forcibly damp the oscillation simply
the
excitation and
the oscillation. (Para. 10.36).
Higher TAS common to high altitude flight _ _ _ the
aerodynamic
. (Para. 10.37).
Mach Tuck is caused by _
the formation of a
of ___ changes and reduces
of lift in front of the _ and
downwash at the tail due to
on a swept back wing at _ _ Mach numbers. (Para. 10.38).
The Mach trim system will adjust
_ _ to maintain the required _ _ __
gradient and operates only at _ _ Mach numbers. (Para. 10.38).
KEY FACTS 2, WITH THE MISSING WORDS INSERTED CAN BE FOUND ON PAGE 10 - 92.
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STABILITY AND CONTROL
PRINCIPLES OF FLIGHT
10.39
SUMMARY - Self Study
Stability is the inherent quality of an aircraft to correct for conditions that may disturb its
equilibrium and to return to, or continue on its origin flight path. An aircraft can have two basic
types of stability: static and dynamic, and three condition of each type: positive, neutral, and
negative.
Static stability describes the initial reaction of an aircraft after it has been disturbed from
equilibrium about one or more of its three axes.
Positive static stability is the condition of stability in which restorative forces are setup that will tend to return an aircraft to its original condition anytime it's disturbed from
a condition of equilibrium. If an aircraft has an initial tendency to return to its original
attitude of equilibrium, it has positive static stability. (statically stable).
An aircraft with neutral static stability produces neither forces that tend to return it to
its original condition, nor cause it to depart further from this condition. If an aircraft
tends to remain in its new, disturbed state, it has neutral static stability. (statically
neutral).
If an aircraft has negative static stability, anytime it is disturbed from a condition of
equilibrium, forces are set up that will tend to cause it to depart further from its original
condition. Negative static stability is a highly undesirable characteristic as it can cause
loss of control. When an aircraft continues to diverge, it exhibits negative static
stability. (statically unstable).
Most aeroplanes have positive static stability in pitch and yaw, and are close to neutrally
statically stable in roll.
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PRINCIPLES OF FLIGHT
STABILITY AND CONTROL
When an aircraft exhibits positive static stability about any of its three axes, the term "dynamic
stability" describes the long term tendency of the aircraft.
When an aircraft is disturbed from equilibrium and then tries to return, it will invariably
overshoot the original ATTITUDE (due to its momentum) and then start to return again. This
results in a series of oscillations.
Positive dynamic stability is a condition in which the forces of static stability decrease
with time. Positive dynamic stability is desirable. If oscillations become smaller with
time, an aircraft has positive dynamic stability. (dynamically stable).
Neutral dynamic stability causes an aircraft to hunt back and forth around a condition
of equilibrium, with the corrections getting neither larger or smaller. (dynamically
neutral). Neutral dynamic stability is undesirable.
If an aircraft diverges further away from its original attitude with each oscillation, it has
negative dynamic stability. Negative dynamic stability causes the forces of static
stability to increase with time. (dynamically unstable). Negative dynamic stability is
extremely undesirable.
The overall design of an aircraft contributes to its stability (or lack of it) about each of its three
axes of motion.
The vertical stabiliser (fin) is the primary source of directional stability (yaw).
The horizontal stabiliser (tailplane) is the primary source of longitudinal stability (pitch).
The wing is the primary source of lateral stability (roll).
CG location also affects stability.
If the CG is close to its aft limit, an aircraft will be less stable in both pitch and yaw.
As the CG is moves forward, stability increases.
Even though an aeroplane will be less stable with an aft CG, it will have some desirable
aerodynamic characteristics due to reduced aerodynamic loading of the horizontal tail surface.
This type of an aeroplane will have a slightly lower stall speed and will cruise faster for a given
power setting.
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STABILITY AND CONTROL
PRINCIPLES OF FLIGHT
Manoeuverability is the quality of an aircraft that permits it to be manoeuvred easily and to
withstand the stresses imposed by those manoeuvres.
Controllability is the capability of an aircraft to respond to the pilot's control, especially with
regard to flight path and attitude.
An aircraft is longitudinally stable if it returns to a condition of level flight after a disturbance
in pitch, caused by either a gust or displacement of the elevator by the pilot. The location of the
CG and the effectiveness of the tailplane determines the longitudinal stability, and thus the
controllability of an aircraft.
Increasing stability about any axis:a)
Decreases manoeuverability and controllability, and
b)
Increases stick (or pedal) forces.
Phugoid oscillation is a long-period oscillation in which the pitch attitude, airspeed, and altitude
vary, but the angle of attack remains relatively constant. It is a gradual interchange of potential
and kinetic energy about some equilibrium airspeed and altitude. An aircraft experiencing
longitudinal phugoid oscillation is demonstrating positive static stability, and it is easily
controlled by the pilot.
An aircraft will return towards wing level after a wing drop if it has static lateral stability.
The wing of most aircraft has a positive geometric dihedral angle (dihedral). This is the angle
produced by the wing tips being higher than the wing root. If the left wing drops in flight, an
aircraft will momentarily begin to slip to the left, and the effective angle of attack of the left
wing will increase and the effective angle of attack of the right wing will decrease. The change
in angle of attack of both wings will cause the wing to return back towards a level attitude.
Sweepback also has a "dihedral effect". This is a by-product. A wing is swept back to give an
aircraft a higher MeRIT. An aircraft with a swept-back wing will not require as much geometrical
dihedral as a straight wing.
Some aircraft have the wing mounted on top of the fuselage for various reasons. Also as a byproduct, a high mounted wing will give a "dihedral effect" due to the direction of airflow around
the fuselage and wing during a sideslip. An aircraft with a high mounted wing does not require
as much geometric dihedral.
An aircraft which has a high mounted, swept-back wing will have so much lateral stability that
the wing is usually given anhedral (negative dihedral).
Too much static lateral stability could result in dynamic instability - Dutch Roll.
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STABILITY AND CONTROL
PRINCIPLES OF FLIGHT
Static Directional stability is the tendency of the nose of an aircraft to yaw towards the relative
airflow. It is achieved by the keel surface behind the CG being larger than that in front of the
CG.
A sweptback wing also provides a measure of static directional stability.
Too much static directional stability could result in dynamic instability - Spiral Instability.
Interaction between static lateral stability and static directional stability. If a wing drops
and the aircraft begins to slip to the side, directional stability will cause the nose to yaw into the
relative airflow.
"Dihedral effect" tends to roll an aircraft when a wing drops, and directional stability causes the
nose to yaw into the direction of the low wing.
These two forces interact (coupled motion):1.
An aircraft with strong static directional stability and weak "dihedral effect" will have
a tendency towards spiral instability.
When a wing drops, the nose will yaw toward the low wing and the airplane will begin
to tum. The increased speed of the wing on the outside of the tum will increase the
angle of bank, and the reduction in the vertical component of lift will force the nose to
a low pitch angle. This will cause the aircraft to enter a descending spiral.
2.
An aircraft with strong "dihedral effect" and weak directional stability will have a
tendency towards dutch roll instability.
A Mach Trim system maintains the required stick force gradient at high Mach numbers
by adjusting the longitudinal trim. The Mach Trim system only operates at high Mach
numbers.
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STABILITY AND CONTROL
PRINCIPLES OF FLIGHT
SELF ASSESSMENT QUESTIONS
1.
An aeroplane which is inherently stable will:
a)
b)
c)
d)
2.
After a disturbance in pitch an aircraft oscillates in pitch with increasing amplitude. It is:
a)
b)
c)
d)
3.
lateral stability about the longitudinal axis.
longitudinal stability about the lateral axis.
lateral stability about the normal axis.
directional stability about the normal axis.
Lateral stability is reduced by increasing:
a)
b)
c)
d)
6.
the fin.
the wing dihedral.
the horizontal tailplane.
the ailerons.
An aircraft is constructed with dihedral to provide:
a)
b)
c)
d)
5.
statically and dynamically unstable.
statically stable but dynamically unstable.
statically unstable but dynamically stable.
statically and dynamically stable.
Longitudinal stability is given by:
a)
b)
c)
d)
4.
require less effort to control.
be difficult to stall.
not spin.
have a built-in tendency to return to its original state following the removal of any
disturbing force.
Anhedral.
Dihedral.
Sweepback.
Fuselage and fin area.
If the wing AC is forward of the CG:
a)
b)
c)
d)
changes in lift produce a wing pitching moment which acts to reduce the change of lift.
changes in lift produce a wing pitching moment which acts to increase the change oflift.
changes in lift give no change in wing pitching moment.
when the aircraft sideslips the CG causes the nose to tum into the sideslip thus applying
a restoring moment.
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PRINCIPLES OF FLIGHT
7.
The longitudinal static stability of an aircraft:
a)
b)
c)
d)
8.
the CO and the AC.
the AC and the neutral point.
the CO and the neutral point.
the CO and the CO datum point.
If a disturbing force causes the aircraft to roll:
a)
b)
c)
d)
12.
very small forces are required on the control column to produce pitch.
longitudinal stability is reduced.
very high stick forces are required to pitch because the aircraft is very stable.
stick forces are the same as for an aft CO.
The static margin is equal to the distance between:
a)
b)
c)
d)
11.
must always coincide with the AC.
must be forward of the Neutral Point.
must be aft of the Neutral Point.
must not be forward of the aft CO limit.
When the CO is close to the forward limit:
a)
b)
c)
d)
10.
is reduced by the effects of wing downwash.
is increased by the effects of wing downwash.
is not affected by wing downwash.
is reduced for nose up displacements, but increased for nose down displacements by the
effects of wing downwash.
To ensure some degree of longitudinal stability in flight, the position of the CO:
a)
b)
c)
d)
9.
STABILITY AND CONTROL
wing dihedral will cause a rolling moment which reduces the sideslip.
the fin will cause a rolling moment which reduces the sideslip.
dihedral will cause a yawing moment which reduces the sideslip.
dihedral will cause a nose up pitching moment.
With flaps lowered, lateral stability:
a)
b)
c)
d)
will be increased because of the effective increase of dihedral.
will be increased because of increased lift.
will be reduced because the centre of lift of each semi-span is closer to the wing root.
will not be affected.
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PRINCIPLES OF FLIGHT
13.
Dihedral gives a stabilising rolling moment by causing an increase in lift:
a)
b)
c)
d)
14.
c)
d)
remains the same.
increases because the TAS increases.
decreases because the ailerons are less effective.
decreases because the density decreases.
Sweepback of the wings will:
a)
b)
c)
d)
17.
greater longitudinal stability.
the same degree of longitudinal stability as any other configuration because dihedral
gives longitudinal stability.
less lateral stability than a low wing configuration.
greater lateral stability due to the airflow pattern around the fuselage when the aircraft
is sideslipping increasing the effective angle of attack of the lower wing.
At a constant lAS, what affect will increasing altitude have on damping in roll:
a)
b)
c)
d)
16.
on the up going wing when the aircraft rolls.
on the down going wing when the aircraft rolls.
on the lower wing if the aircraft is sideslipping.
on the lower wing whenever the aircraft is in a banked attitude.
A high wing configuration with no dihedral, compared to a low wing configuration with no
dihedral, will provide:
a)
b)
15.
STABILITY AND CONTROL
not affect lateral stability.
decrease lateral stability.
increases lateral stability at high speeds only.
increases lateral stability at all speeds.
At low forward speed:
a)
b)
c)
d)
increased downwash from the wing will cause the elevators to be more responsive.
due to the increased angle of attack of the wing the air will flow faster over the wing
giving improved aileron control.
a large sideslip angle could cause the fin to stall.
a swept back wing will give an increased degree of longitudinal stability.
10 - 83
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STABILITY AND CONTROL
PRINCIPLES OF FLIGHT
18.
Following a lateral disturbance, an aircraft with Dutch roll instability will:
a)
b)
c)
d)
19.
To correct dutch roll on an aircraft with no automatic protection system:
a)
b)
c)
d)
20.
wash out.
taper.
sweep.
anhedral.
The lateral axis of an aircraft is a line which:
a)
b)
c)
d)
23.
increases rudder effectiveness.
must be disengaged before making a tum.
augments stability.
increases the rate of yaw.
A wing which is inclined downwards from root to tip is said to have:
a)
b)
c)
d)
22.
use roll inputs
use yaw inputs
move the CG
reduce speed below M MO
A yaw damper:
a)
b)
c)
d)
21.
go into a spiral dive.
develop simultaneous oscillations in roll and yaw.
develop oscillations in pitch.
develop an unchecked roll.
passes through the wing tips.
passes through the centre of pressure, at right angles to the direction of the airflow.
passes through the quarter chord point of the wing root, at right angles to the
longitudinal axis.
passes through the centre of gravity, parallel to a line through the wing tips.
Loading an aircraft so that the CG exceeds the aft limits could result in:
a)
b)
c)
d)
loss of longitudinal stability, and the nose to pitch up at slow speeds
excessive upward force on the tail, and the nose to pitch down
excessive load factor in turns
high stick forces
10 - 84
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PRINCIPLES OF FLIGHT
24
The tendency of an aircraft to suffer from dutch roll instability can be reduced:
a)
b)
c)
d)
25.
the angle between the main plane and the longitudinal axis
the angle measured between the main plane and the normal axis
the angle between the quarter chord line and the horizontal datum
the upward and outward inclination of the main planes to the horizontal datum
Stability around the normal axis:
a)
b)
c)
d)
28.
The relationship of thrust and lift to weight and drag
The effectiveness of the horizontal stabilizer, rudder, and rudder trim tab
The location of the CG with respect to the AC
the size of the pitching moment which can be generated by the elevator
Dihedral angle is:
a)
b)
c)
d)
27.
by sweeping the wings
by giving the wings anhedral
by reducing the size of the fin
by longitudinal dihedral
What determines the longitudinal static stability of an aeroplane?
a)
b)
c)
d)
26.
STABILITY AND CONTROL
is increased if the keel surface behind the CG is increased
is given by the lateral dihedral
depends on the longitudinal dihedral
is greater if the wing has no sweepback
The Centre of Gravity of an aircraft is found to be within limits for take-off::
a)
b)
c)
d)
the C of G will be within limits for landing
the C of G for landing must be checked, allowing for fuel consumed
the C of G will not change during the flight
the flight crew can adjust the CG during flight to keep it within acceptable limits for
landing
10 - 85
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PRINCIPLES OF FLIGHT
29.
The ailerons are deployed and returned to neutral when the aircraft has attained a small angle
of bank. If the aircraft then returns to a wings-level attitude without further control movement
it is:
a)
b)
c)
d)
30.
neutrally stable
statically and dynamically stable
statically stable, dynamically neutral
statically stable
The property which tends to decreases rate of displacement about any axis, but only while
displacement is taking place, is known as:
a)
b)
c)
d)
31.
STABILITY AND CONTROL
stability
controllability
aerodynamic damping
manoeuverability
If an aircraft is loaded such that the stick force required to change the speed is zero
a)
b)
c)
d)
the CO is on the neutral point
the CO is behind the neutral point
the CO is on the manouevre point
the CO is on the forward CO limit
10 - 86
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PRINCIPLES OF FLIGHT
STABILITY AND CONTROL
ANSWERS
I No I A I B I c I DII
1
REF
III
No
17
D
2
19
C
20
5
A
21
D
22
D
7
23
A
8
9
C
25
10
C
26
11
C
A
24
B
27
A
B
C
D
A
12
C
28
B
13
C
29
B
14
D
30
15
D
31
16
D
/
10 - 87
I
A
A
B
REF
B
4
6
II
C
18
B
3
IA I B I c ID
C
A
V
~ ~~ ~
© Oxford Aviation Services Limited
STABILITY AND CONTROL
PRINCIPLES OF FLIGHT
KEY FACTS 1 - Self Study
Stability is the tendency of an aircraft to return to a steady state of flight, after being disturbed
by an external force, without any help from the pilot. (Para. 10.1).
There are two broad categories of stability; static and dynamic. (Para. 10.1).
An aircraft is in a state of equilibrium (trim) when the sum of all forces is zero and the sum of
all moments is zero. (Para. 10.2).
The type of static stability an aircraft possesses is defined by its initial tendency, following the
removal of some disturbing force. (Para. 10.2).
The three different types of static stability are: (Para. 10.2).
a)
Positive static stability exists if an aircraft is disturbed from equilibrium and has the
tendency to return to equilibrium.
Neutral static stability exists if an aircraft is subj ect to a disturbance and has neither the
b)
tendency to return nor the tendency to continue in the displacement direction.
c)
Negative static stability exists if an aircraft has a tendency to continue in the direction
of disturbance.
The longitudinal axis passes through the CG from nose to tail. (Para. 10.3)
The normal axis passes "vertically" through the CG at 90° to the longitudinal axis. (Para. 10.3)
The lateral axis is a line passing through the CG, parallel to a line passing through the wing tips.
(Para. 10.3).
The three reference axes all pass through the centre of gravity. (Para. 10.3)
Lateral stability involves motion about the longitudinal axis (rolling). (Para. 10.4).
Longitudinal stability involves motion about the lateral axis (pitching). (Para. 10.4).
Directional stability involves motion about the 'normal axis (yawing). (Para. 10.4).
10 - 89
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STABILITY AND CONTROL
PRINCIPLES OF FLIGHT
We consider the changes in magnitude of lift force due to changes in angle of attack, acting
through a stationary point; the aerodynamic centre. (Page 10 - 6)
The aerodynamic centre (AC) is located at the 25% chord position. (Page 10 - 6)
The negative pitching moment about the AC remains constant at normal angles of attack.
(Page 10 - 6).
A wing on its own is statically unstable because the AC is in front of the CG. (Page 10 - 7).
An upward vertical gust will momentarily increase the angle of attack of the wing. The
increased lift force magnitude acting through the AC will increase the positive pitching moment
about the CG. This is an unstable pitching moment. (Page 10 - 7).
The tailplane is positioned to generate a stabilising pitching moment about the aircraft CG.
(Page 10 - 8).
If the tail moment is greater than the wing moment the sum of the moments will not be zero and
the resultant nose down moment will give an angular acceleration about the CG. (Page 10- 8).
The greater the tail moment relative to the wing moment, the greater the rate of return towards
the original equilibrium position. (Page 10 - 8).
The tail moment is increased by moving the aircraft CG forwards, which increases the tail arm
and decreases the wing arm. (Page 10 - 8).
If the nose down (negative) tail moment is greater than the nose up (positive) wing moment, the
aircraft will have static longitudinal stability. (Page 10 - 8).
The position of the CG when changes in the sum of the tail moment and wing moment due to a
disturbance is zero, is known as the neutral point. (Para. 10.5).
The further forward the CG, the greater the nose down angular acceleration about the CG - the
greater the degree of static longitudinal stability. (Para. 10.6).
The distance the CG is forward of the neutral point will give a measure of the static
longitudinal stability; this distance is called the static margin. (Para. 10.6).
The greater the static margin, the greater the static longitudinal stability. (Para. 10.6).
The aft CG limit will be positioned some distance forward of the neutral point. (Para. 10.6).
The distance between the aft CG limit and the neutral point gives the required minimum static
stability margin. (Para. 10.6).
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PRINCIPLES OF FLIGHT
STABILITY AND CONTROL
An aircraft is said to be trimmed if all moments in pitch, roll, and yaw are equal to zero.
(Para. 10.7).
Trim (equilibrium) is the function of the controls and may be accomplished by:- (Para. 10.7).
a)
pilot effort
b)
trim tabs,
c)
moving fuel between the wing tanks and an aft located trim tank, or
d)
bias of a surface actuator (powered flying controls).
The term controllability refers to the ability of the aircraft to respond to control surface
displacement and achieve the desired condition of flight. (Para. 10.7).
A high degree of stability tends to reduce the controllability of the aircraft. (Para. 10.7).
The stable tendency of an aircraft resists displacement from trim equally, whether by pilot effort
on the controls (stick force) or gusts. (Para. 10.7).
If the CG moves forward, static longitudinal stability increases and controllability decreases
(stick force increases). (Para. 10.7).
If the CG moves aft, static longitudinal stability decreases and controllability increases (stick
force decreases). (Para. 10.7).
With the CG on the forward limit, static longitudinal stability is greatest, controllability is least
and stick force is high. (Para. 10.7).
With the CG on the aft limit, static longitudinal stability is least, controllability is greatest and
stick force is low. (Para. 10.7).
The aft CG limit is set to ensure a minimum degree of static longitudinal stability. (Para. 10.7).
The fwd CG limit is set to ensure a minimum degree of controllability under the worst
circumstance. (Para. 10.7).
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PRINCIPLES OF FLIGHT
STABILITY AND CONTROL
KEY FACTS 2 - Self Study
Positive static longitudinal stability is indicated by a negative slope ofCM versus CL • The degree
of static longitudinal stability is indicated by the slope of the curve. (Para. 10.8).
The net pitching moment about the lateral axis is due to the contribution of each of the
component surfaces acting in their appropriate flow fields. (Para. 10.9).
In most cases, the contribution of the fuselage and nacelles is destabilising. (Para. 10.9).
(Para.
a)
b)
c)
10.9) Noticeable changes in static stability can occur at high CL (low speed) if:
the aeroplane has sweepback,
there is a large contribution of 'power effect', or
there are significant changes in downwash at the horizontal tail,
The horizontal tail usually provides the greatest stabilising influence of all the components of
the aeroplane. (Para. 10.9).
Downwash decreases static longitudinal stability. (Para. 10.9).
If the thrust line is below the CG, a thrust increase will produce a positive or nose up moment
and the effect is destabilizing. (Para. 10.12).
High lift devices tend to increase downwash at the tail and reduce the dynamic pressure at the
tail, both of which are destabilizing. (Para. 10.13).
An increase in TAS, for a given pitching velocity, decreases aerodynamic damping.
(Para. 10.15).
The aeroplane with positive manoeuvring stability should demonstrate a steady increase in stick
force with increase in load factor or "g". (Para. 10.16).
The stick force gradient must not be excessively high or the aeroplane will be difficult and tiring
to manoeuver. Also, the stick force gradient must not be too low or the aeroplane may be
overstressed inadvertently when light control forces exist. (Para. 10.16).
When the aeroplane has high static stability, the manoeuvring stability will be high and a high
stick force gradient will result. The forward CG limit could be set to prevent an excessively
high manoeuvring stick force gradient. As the CG moves aft, the stick force gradient decreases
with decreasing manoeuvring stability and the lower limit of stick force gradient may be
reached. (Para. 10.16).
At high altitudes, the high TAS reduces the change in tail angle of attack for a given pitching
velocity and reduces the pitch damping. Thus, a decrease in manoeuvring stick force stability
can be expected with increased altitude. (Para. 10.16).
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PRINCIPLES OF FLIGHT
STABILITY AND CONTROL
A flying control system may employ centring springs, down springs or bob weights to provide
satisfactory control forces throughout the speed, CG and altitude range of an aircraft.
(Para. 10.17).
While static stability is concerned with the initial tendency of an aircraft to return to equilibrium,
dynamic stability is defined by the resulting motion with time. (Para. 10.22).
An aircraft will demonstrate positive dynamic stability if the amplitude of motion decreases
with time. (Para. 10.22).
When natural aerodynamic damping cannot be obtained, artificial damping must be provided
to give the necessary positive dynamic stability. (Para. 10.22).
(Para. 10.23) The longitudinal dynamic stability of an aeroplane generally consists of two basic
modes of oscillation:long period (phugoid)
a)
short period
b)
The phugoid oscillation occurs with nearly constant angle of attack. (Para. 10.24).
The period of oscillation is so great, the pilot is easily able to counteract long period oscillation.
(Para. 10.24).
Short period oscillation involves significant changes in angle of attack. (Para. 10.25).
Short period oscillation is not easily controlled by the pilot. (Para. 10.25).
The problems of dynamic stability can become acute at high altitude because of reduced
aerodynamic damping. (Para. 10.25).
To overcome the directional instability in the fuselage it is possible to incorporate into the
overall design, dorsal or ventral fins. (Para. 10.29).
The fin is the major source of directional stability for the aeroplane. (Para. 10.29).
T - tail makes the fin more effective by acting as an "end plate". (Para. 10.29).
Because the dorsal fin stalls at a very much higher angle of attack, it takes over the stabilizing
role of the fin at large angles of sideslip. (Para. 10.29).
Sweepback produces a directional stabilising effect, which increases with increase in C L •
(Para. 10.29).
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STABILITY AND CONTROL
PRINCIPLES OF FLIGHT
Ventral fins increase directional stability at high angles of attack. Landing clearance
requirements may limit their size, require them to be retractable, or require two smaller ventral
fins to be fitted instead of one large one. (Para. 10.29).
Generally, good handling qualities are obtained with a relatively light, or weak positive, lateral
stability. (Para. 10.31).
The principal surface contributing to the lateral stability of an aeroplane is the wing. The effect
of geometric dihedral is a powerful contribution to lateral stability. (Para. 10.31).
A low wing position gives an unstable contribution to static lateral stability. (Para. 10.32).
A high wing location gives a stable contribution to static lateral stability. (Para. 10.32).
The magnitude of "dihedral effect" contributed by the vertical position of the wing is large and
may require a noticeable dihedral angle for the low wing configuration. A high wing position,
on the other hand, usually requires no geometric dihedral at all. (Para. 10.32).
The swept back wing contributes a positive "dihedral effect". (Para. 10.32).
An aircraft with a swept back wing requires less geometric dihedral than a straight wing.
(Para. 10.32).
The fin contribution to purely lateral static stability, is usually very small. (Para. 10.32).
Excessive "dihedral effect" can lead to "Dutch roll," difficult rudder coordination in rolling
manoeuvres, or place extreme demands for lateral control power during crosswind takeoff and
landing. (Para. 10.32).
Deploying partial span flaps gives a reduced dihedral effect. (Para. 10.32).
A sweptback wing requires much less geometric dihedral than a straight wing. If a requirement
also exists for the wing to be mounted on top of the fuselage, an additional "dihedral effect" is
present. A high mounted and sweptback wing would give excessive "dihedral effect", so
anhedral is used to reduce "dihedral effect" to the required level. (Para. 10.32).
When an aeroplane is placed in a sideslip, the lateral and directional response will be coupled,
i.e. sideslip will simultaneously produce a rolling and a yawing moment. (Para. 10.33).
Spiral divergence will exist when static directional stability is very large when compared to the
"dihedral effect". (Para. 10.34).
The rate of divergence in the spiral motion is usually so gradual that the pilot can control the
tendency without difficulty. (Para. 10.34).
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PRINCIPLES OF FLIGHT
STABILITY AND CONTROL
Dutch roll will occur when the "dihedral effect" is large when compared to static directional
stability. (Para. 10.35).
Aircraft which Dutch Roll are fitted with a Yaw Damper. This automatically displaces the
rudder proportional to the rate of yaw to damp-out the oscillations. (Para. 10.35).
If the Yaw Damper fails in flight, it is recommended that the ailerons be used by the pilot to
damp-out Dutch Roll. (Para. 10.35).
If the pilot uses the rudder, pilot induced oscillation (PIO) will result and the Dutch Roll may
very quickly become divergent, leading to loss of control. (Para. 10.35).
When the swept wing aeroplane is at low CL the "dihedral effect" is small and the spiral
tendency may be apparent. When the swept wing aeroplane is at high CL the "dihedral effect"
is increased and the Dutch Roll oscillatory tendency is increased. (Para. 10.35).
When pilot induced oscillation is encountered, the most effective solution is an immediate
release of the controls. Any attempt to forcibly damp the oscillation simply continues the
excitation and amplifies the oscillation. (Para. 10.36).
Higher TAS common to high altitude flight reduces the angle of attack changes and reduces
aerodynamic damping. (Para. 10.37).
Mach Tuck is caused by loss of lift in front of the CG and reduced downwash at the tail due to
the formation ofa shockwave on a swept back wing at high Mach numbers. (Para. 10.38).
The Mach trim system will adjust longitudinal trim to maintain the required stick force
gradient and operates only at high Mach numbers. (Para. 10.38).
10 - 95
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CHAPTER 11 - CONTROLS
Contents
Page
INTRODUCTION ...................................................... 11 - 1
HINGE MOMENTS ..................................................... 11 - 2
CONTROL BALANCING ................................................ 11 - 3
AERODYNAMIC BALANCE
INSET HINGE
HORN BALANCE
INTERNAL BALANCE ........................................... 11 - 4
BALANCE TAB
ANTI-BALANCE TAB ............................................ 11 - 5
SERVO TAB
SPRING TAB
POWERED FLYING CONTROLS ......................................... 11 - 6
POWER ASSISTED
FULLY POWERED .............................................. 11 - 7
ARTIFICIAL FEEL ('Q' FEEL)
MASS BALANCE ...................................................... 11 - 8
LONGITUDINAL CONTROL
EFFECT OF ELEVATOR DEFLECTION
DIRECTION OF THE TAILPLANE LOAD .................................. 11 - 9
ELEVATOR ANGLE WITH 'G'
EFFECT OF ICE ON THE TAILPLANE
LATERAL CONTROL
EFFECT OF AILERON DEFLECTION .................................... 11 - 10
EFFECT OF WINGSPAN ON RATE OF ROLL
ADVERSE AILERON YAW
DIFFERENTIAL AILERONS ...................................... 11 - 11
FRISE AILERONS
AILERON - RUDDER COUPLING
ROLL CONTROL SPOILERS
INBOARD AILERONS ................................................. 11 - 12
FLAPERONS
SPOILERS ........................................................... 11 - 13
COMBINED AILERON AND SPOILER CONTROLS ........................ 11 - 14
SPEED BRAKES
TYPES OF SPEED BRAKE
EFFECT OF SPEED BRAKES ON THE DRAG CURVE ...................... 11 - 15
GROUND SPOILERS ( LIFT DUMPERS)
DIRECTIONAL CONTROL ............................................. 11 - 16
EFFECT OF RUDDER DEFLECTION
FIN STALL
ASYMMETRIC THRUST ...............................................
RUDDER RATIO CHANGER
SECONDARY EFFECTS OF CONTROLS .................................
YA WING MOMENT DUE TO ROLL
ROLLING MOMENT DUE TO YAW
TRIMMING ..........................................................
METHODS OF TRIMMING
TRIM TAB ....................................................
FIXED TABS
V ARIABLE INCIDENCE TAILPLANE
SPRING BIAS ..................................................
CG ADJUSTMENT
ARTIFICIAL FEEL TRIM
SELF ASSESSMENT QUESTIONS .......................................
ANSWERS ....................................................
11 - 17
11 - 18
11 - 19
11 - 20
11 - 22
11 - 23
11 - 31
PRINCIPLES OF FLIGHT
11.1
CONTROLS
INTRODUCTION
All aircraft are fitted with a control system to enable the pilot to manoeuvre and trim the aircraft
in flight about each of its three axes. The aerodynamic moments required to rotate the aircraft
about the axes are usually supplied by means of 'flap ' type control surfaces positioned at the
extremities of the aircraft so that they have the longest possible moment arm about the CG.
There are usually three separate control systems and three sets of control surfaces:
a)
Rudder for control in yaw about the normal axis (directional control).
b)
Elevator for control in pitch about the lateral axis (longitudinal control).
c)
Ailerons for control in roll about the longitudinal axis (lateral control). Spoilers may
also be used to assist or replace the ailerons for roll control.
The effect of two of these controls may be combined in a single set of control surfaces:
d)
Elevons: combine the effects of elevator and aileron.
e)
Ruddervator: ('V' or butterfly tail), combine the effects of rudder and elevator.
f)
Tailerons: slab horizontal tail surfaces that move either together, as pitch control, or
independently for control in roll.
The moment around an axis is produced by changing the aerodynamic force on the appropriate
aerofoil. The magnitude of the force is a product of the dynamic pressure (IAS2) and the angular
displacement of the control surface. Aerodynamic force can be changed by:g)
h)
i)
adjusting the camber of the aerofoil.
changing the incidence of the aerofoil.
decreasing lift and increasing drag by "spoiling" the airflow.
Changing the camber of any
aerofoil (wing, tailplane or
fin) will change its lift.
Deflecting a control surface
effectively changes its
camber. Fig. 11 .1 shows the
effect on CL of movement of
a control surface.
ANGLE OF ATIACK
Figure 11.1
Control surface changes camber and lift
11 - 1
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
CONTROLS
Changing the incidence of an aerofoi I
will change its lift.
The usual
application of this system is for pitch
control - the all moving (slab)
tailplane. There is no elevator, when
the pilot makes a pitch input the
incidence of the whole tailplane
changes.
(
Figure 11.2 All moving (slab) tailplane
~ / ..
Spoilers are a device for reducing the
lift of an aero foil, by disturbing the
airflow over the upper surface. They
assist lateral control by moving up on
the side with the up-going aileron, as
illustrated in Fig. 11.3.
SPOILER SURFACES
/~~~.--~~l --
AILERONS
~_y.
~~~~
,"
Figure 11.3
11.2
.
~
Spoilers
HINGE MOMENTS
If an aerodynamic force acts on a control surface, it will try to rotate the control around its hinge
in the direction of the force. The moment is a product of the force (F) times the distance (d)
from the hinge line to the control surface CPo This is called the hinge moment. The force is due
to the surface area, the angular displacement of the control surface and the dynamic pressure.
HINGE MOMENT
=F
x d
F2
...,
j
Figure 11.4
Hinge moment (Feel)
To move the control surface to the required angular displacement and maintain it in that position
the pilot has to overcome, then balance, the hinge moment by applying a force (stick force) to
the cockpit control. The stick force will therefore depend on the size of the hinge moment.
11 - 2
© Oxford Aviation Services Limited
CONTROLS
PRINCIPLES OF FLIGHT
11.3
CONTROL BALANCING
The aerodynamic force on the controls will depend on the area of the control surface, its angular
displacement and the lAS. For large and fast aircraft the resulting aerodynamic force can give
hinge moments / stick forces which are too high for easy operation of the controls.
The pilot will require assistance to move the controls in these conditions, and this can be done
either by using (hydraulic) powered flying controls, or by using some form of aerodynamic
balance.
11.3.1
AERODYNAMIC BALANCE
Aerodynamic balance involves using the aerodynamic forces on the control surface to reduce the
hinge mom~nt / stick force and may be done in several ways:
a)
HINGE SET - BACK
INTO SURFACE
F2
)
Figure 11.5
If the aerodynamic force (F2) were to
move forward of the hinge, a
condition known as "overbalance"
would exist. As the force moved
forward, a reduction then a reversal
of the stick force would occur. This
would be very dangerous and the
designer must ensure the aerodynamic
force can never move forward of the
hinge.
Inset hinge
, b)
HINGE
/' / ' LiNE
AEROFOIL
~~ /
/
HORN
Figure 11.6
Inset hinge: If distance (d) is reduced
the hinge moment will be reduced.
The smaller the hinge moment, the
smaller the stick force and the easier
it will be for the pilot to move the
controls. Setting the hinge back does
not reduce the effectiveness of the
control, only the hinge moment.
-~ ~
/
CONTROL
SURFACE
Horn balance
11 - 3
Horn Balance: The principle of the
hom balance is similar to that of the
inset hinge, in that part of the surface
is forward of the hinge line, and
forces on this part of the surface give
hinge moments which are in the
opposite direction to the moments on
the main part of the surface. The
overall moment is therefore reduced,
but not the control effectiveness.
© Oxford Aviation Services Limited
CONTROLS
PRINCIPLES OF FLIGHT
c)
Internal Balance: This balance works on the same principle as the inset hinge, but the
aerodynamic balance area is inside the wing.
Figure 11.7
Internal balance
Movement of the control causes pressure changes on the aerofoil, and these pressure
changes are felt on the balance area. For example, if the control surface is moved down,
pressure above the aerofoil is reduced and pressure below it is increased. The reduced
pressure is felt on the upper surface of the balance 'panel', and the increased pressure
on the lower surface. The pressure difference on the balance therefore gives a hinge
moment which is the opposite to the hinge moment on the main control surface, and the
overall hinge moment is reduced.
See Page 11-22 for a Tab Quick Reference Guide.
d)
Balance Tab: The preceding types of aerodynamic balance work by causing some of
the dynamic pressure on the control surface to act forward of the hinge line. The
balance tab provides a force acting on the control surface trailing edge opposite to the
force on the main control surface. The balance tab moves in the opposite direction
to the control surface. The pilot moves the surface, the surface moves the tab.
TAB
FORCE
PILOT INPUT . . . "------~
----? ~~
--
CONTROL
FORCE
Figure 11.8
Balance Tab
Unlike the previous types of balance, the balance tab will give some reduction in control
effectiveness, as the tab force is opposite to the control force.
11 - 4
© Oxford Aviation Services Limited
CONTROLS
PRINCIPLES OF FLIGHT
PILOT INPUT"
c==
_-------115~=:~
i7=~QJ~~
TAB
FORCE
CONTROL
FORCE
Figure 11.9 Anti - Balance tab
e)
Anti-balance Tab: The anti-balance tab moves in the same direction as the control
surface and increases control effectiveness, but will increase the hinge moment and give
heavier stick forces . The pilot moves the surface, the surface moves the tab.
PILOT INPUT . .
c ·_-
CONTROL "HORN" FREE TO
PIVOT ON HINGE AXIS
(
Figure 11 .10
f)
Servo Tab
Servo Tab: Pilot control input deflects the servo tab only, the aerodynamic force on
the tab then moves the control surface until an equilibrium position is reached. If
external control locks are fitted to the control surface on the ground, the cockpit
control will still be free to move. Older types of high speed jet transport aircraft
(B707) successfully used servo tab controls, the disadvantage of the servo tab is reduced
control effectiveness at low lAS.
HORN FREE TO
PIVOT ON HINGE AXIS
PILOT INPUT . . ~(_ _--,-'
Figure 11.11
g)
Spring Tab
Spring Tab: The spring tab is a modification of the servo tab, such that tab movement
is proportional to the applied stick force. Maximum tab assistance is obtained at high
speed when the stick forces are greatest. High dynamic pressure will prevent the surface
from moving, so the spring is compressed by the pilot input and the tab moves the
surface. The spring is not compressed at low lAS, so the pilot input deflects the control
surface and the tab, increasing the surface area and control effectiveness at low speed.
11 - 5
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CONTROLS
PRINCIPLES OF FLIGHT
11.3.2 (HYDRAULIC) POWERED FLYING CONTROLS
If the required assistance for the pilot to move the controls cannot be provided by the preceding
types of aerodynamic balance, then power assisted or fully powered controls have to be used.
POWER FLYING
CONTROL UNIT (PFCU)
..
SERVO~
VALVE
Figure 11.12
a)
Power assisted flying control
Power Assisted Controls
With a power assisted flying control, Fig. 11 .12, only a certain proportion of the force required
to oppose the hinge moment is provided by the pilot, the hydraulic system provides most of the
force. Although the pilot does not have to provide all the force required, the natural 'feel' of the
controls is retained and the stick force increases as the square of the lAS, just as in a completely
manual control.
POWER FLYING
CONTROL UNrT (PFCU) \
+-
..
Figure 11.13
Fully powered flying control
11 - 6
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PRINCIPLES OF FLIGHT
b)
CONTROLS
Fully Powered Controls
For bigger and/or faster aircraft, hinge moments are so large that fully powered controls must
be used. In a fully powered control system, none of the force to move the control surface is
supplied by the pilot. The only force the pilot supplies is that required to overcome system
friction and to move the servo valve, all the necessary power to move the control surface is
supplied by the aircraft's hydraulic system.
Fig. 11 .13 shows that movement of the servo valve to the left allows hydraulic fluid to enter the
left chamber of the PFCU. The body of the unit will move to the left, its movement being
transferred to the control surface. As soon as the PFCU body reaches the position into which
the pilot placed the servo valve, the PFCU body, and hence the control surface, stops moving.
The unit is now locked in its new position by "incompressible" liquid trapped on both sides of
the piston and will remain in that position until the servo valve is again moved by the pilot.
Aerodynamic loads on the control surface are unable to move the cockpit controls, so powered
flying controls are known as "irreversible" controls.
c)
Artificial Feel ('Q' Feel)
POWER FLYING
CONTROL UNIT (PFCU) \
r'-,.,.-----,
STATIC -
\
--
PITOT-
ARTIFICLAL FEEL UNIT
('Q' FEEL)
SERVO
VALVE
J
Figure 11.14
Artificial feel ('Q' feel)
With a fully powered flying control the pilot is unaware of the aerodynamic force on the
controls, so it is necessary to incorporate "artificial" feel to prevent the aircraft from being
overstressed. As shown schematically in Fig. 11.14, a device sensitive to dynamic pressure
(Yz p V2) or 'Q' is used. Pitot pressure is fed to one side of a chamber and static pressure to the
other, which moves a diaphragm under the influence of changing dynamic pressure with airspeed
and causes "regulated" hydraulic pressure to provide a resistance or "feel" on the pilot's input
controls proportional to IAS 2,just as in a manual control. In addition, stick force should increase
as stick displacement increases.
11 - 7
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PRINCIPLES OF FLIGHT
11.4
CONTROLS
MASS BALANCE
Mass balance is a WEIGHT attached to the control surface forward of the hinge. Most
control surfaces are mass balanced. The purpose of this is to prevent control surface flutter.
Flutter is an oscillation of the control surface which can occur due to the bending and twisting
of the structure under load. If the control surface CG is behind the hinge line, inertia will cause
the surface to oscillate about its hinge line. The oscillations can be divergent, and cause
structural failure. A detailed explanation of flutter is given in Chapter 14.
£[;0 Figure 11 .15
Mass balance weights
Flutter may be prevented by adding weight to the control surface in front of the hinge line.
This brings the centre of gravity of the control forward to a position on, or slightly in front of
the hinge, but always to the point required by the designers . This reduces the inertia moments
about the hinge and prevents flutter developing. Fig. 11.15 illustrates some common methods
of mass balancing.
11.5
LONGITUDINAL CONTROL
Control in pitch is usually obtained by elevators or by an all moving (slab) tailplane, and the
controls must be adequate to balance the aircraft throughout its speed range at all permitted CG
positions and configurations and to give an adequate rate of pitch for manoeuvres.
11.5.1
EFFECT OF ELEVATOR DEFLECTION
Suppose that the aircraft is flying in balance at a steady speed with the elevator neutral. If the
elevator is deflected upwards, the tail will develop a download which will begin to pitch the
aircraft nose upwards. As the angle of attack increases, the tailplane download decreases and
the aircraft will reach an equilibrium pitch position. It will then remain in that pitch position
with the elevator kept at the selected angle. If the elevator is returned to neutral the tail will
develop an upload which will begin to pitch the aeroplane down again. At a given C G position
there will be a given pitch attitude for each elevator position.
11 - 8
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CONTROLS
PRINCIPLES OF FLIGHT
11.5.2
DIRECTION OF THE TAILPLANE LOAD
The elevator angle required to give balance depends on lAS and the CG position. At normal
cruising speeds and CG positions, the elevator should ideally be approximately neutral. The
tailplane will be giving a download, and consequently a nose-up pitching moment. This will
balance the nose-down moment created by the wing with its centre of pressure fairly well aft.
At higher than normal speeds the CP will move further rearwards giving a stronger nose-down
pitch, and needing a larger down-load from the tailplane. However at higher speed the aircraft's
angle of attack will be decreased, requiring some down elevator to provide the correct tail-load.
At lower than normal speeds the CP will move forward and the wing and fuselage may cause a
nose-up pitching moment. The tailplane will be required to give an up-load for balance. At low
speed the aircraft will be at a high angle of attack, and to reach this attitude the elevator will have
been moved up.
11.5.3
ELEVATOR ANGLE WITH 'g'
When the aircraft is performing a pitching manoeuvre the tailplane angle of attack is increased
by the effect of the rotational velocity and the aerodynamic damping is increased. This means
that a larger elevator angle will be required than for the same conditions in 1g flight. The
additional elevator angle required will be proportional to the' g' being experienced. The elevator
movement available should be sufficient to allow the design limit' g' to be reached. The most
demanding requirement for elevator up authority will be when the aircraft is being flared
for landing, in ground effect with most forward CG (see Paragraph 10.21).
11.5.4
EFFECT OF ICE ON THE TAILPLANE
The tailplane is an aerofoil, usually symmetrical as it is required to produce both up and down
loads. It is set at an angle of incidence which is less than that of the wing. This ensures that it
will not stall before the wing, and so control can be maintained up to the stall. It is usually
affected by the downwash from the wing and this reduces its effective angle of attack.
Typically the tail will be at a negative angle of attack, producing a download for balance. If ice
forms on the tailplane leading edge, its aerofoil shape will be distorted, and its stalling angle
reduced. This could cause the tailplane to stall, particularly if the downwash is increased as
result of lowering flaps. With the tailplane stalled its download would be reduced and the
aircraft would pitch down and could not be recovered.
11.6
LATERAL CONTROL
Control in roll is usually obtained by ailerons or by spoilers, or by a combination of the two. The
main requirement for lateral control is to achieve an adequate rate of roll.
On the ground with the control wheel in the neutral position both ailerons should be slightly
below alignment with the wing trailing edge, "drooped". When airborne, the lower pressure on
the top surface will "suck" both ailerons up into a position where they are perfectly aligned with
the wing trailing edge, thus reducing drag.
11 - 9
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CONTROLS
PRINCIPLES OF FLIGHT
11.6.1
EFFECT OF AILERON DEFLECTION (Aerodynamic Damping)
In steady level flight with the ailerons neutral, the lift on the two wings will be equal. If the
control wheel is turned to the left, the left aileron will move up and the right aileron down. The
up aileron will decrease the lift of the left wing which will begin to 'drop'. The downward
movement of the wing creates a relative airflow upwards, which increases its effective angle of
attack. The opposite effects will occur on the right (up going) wing.
HIGH TAS
~--~~
/
LOW TAS
~~---~- ~
RELATIVE AIRFLOW
FROM ANGULAR ROTATION
EQUAL WING TIP VELOCITY
INCREASE IN EFFECTIVE ANGLE OF ATTACK
DUE TO WING TIP DOWNWARDS VELOCITY
Figure 11.16
Aerodynamic damping in roll
The increased effective angle of attack of the down going wing increases its lift, which opposes
the roll. This is called aerodynamic damping. The greater the rate of roll, the greater the
damping. It can also be seen from Fig. 11.16 that the greater the TAS, the smaller the increase
in effective angle of attack for a given roll rate.
The change in wing lift for a given aileron deflection depends on the lAS, but the change of
effective angle of attack due to roll velocity depends on TAS . At high TAS (constant lAS,
higher altitude) the change in effective angle of attack will be reduced and a higher rate of roll
will be possible. Rate of roll therefore increases (aerodynamic damping decreases) with higher
TAS for a given aileron deflection. The aileron is known as a rate control since a given
aileron angle of deflection determines a rate of roll, not a roll displacement.
11.6.2
EFFECT OF WINGSPAN ON RATE OF ROLL
For a given rate of roll, the wing tip rolling velocity will increase with the wing span.
Aerodynamic damping will therefore be greater if the span is greater. Under the same
conditions, a short span wing will have a greater rate of roll than a large span wing.
11.6.3
ADVERSE AILERON YAW
The ailerons produce a rolling moment by increasing the lift on one wing and decreasing it on
the other. The increased lift on the up-going wing gives an increase in the induced drag, whereas
the reduced lift on the downgoing wing gives a decease in induced drag. The difference in drag
on the two wings produces a yawing moment which is opposite to the rolling moment, that is,
a roll to the left produces a yawing moment to the right. This is known as adverse aileron yaw.
11 - 10
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PRINCIPLES OF FLIGHT
11.6.4
CONTROLS
REDUCING ADVERSE AILERON YAW
LARGE
UPWARD
MOVEMENT
~ ,(, -
Figure 11 .17
a)
SMALL
DOWNWARD
MOVEMENT
Differential ailerons
Differential ailerons: The aileron linkage causes the up-going aileron the move
through a larger angle than the down-going aileron, Fig. 11.17. This increases the drag
on the up aileron, and reduces it on the down aileron, and so reduces the difference in
drag between the two wings.
Figure 11 .18
Frise ailerons
b)
Frise ailerons: have an asymmetric leading edge, as illustrated in Fig. 11 .18. The
leading edge of the up-going aileron protrudes below the lower surface of the wing,
causing high drag. The leading edge of the down-going aileron remains shrouded and
causes less drag.
c)
Aileron-rudder coupling: In this system the aileron and rudder controls are
interconnected, so that when the ailerons are deflected the rudder automatically moves
to counter the adverse yaw.
d)
Roll control spoilers: If roll spoilers are used to augment the roll rate obtained from
the ailerons, they will reduce the adverse yaw, as the down-going wing will have an
increase in drag due to the raised spoiler.
11 - 11
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PRINCIPLES OF FLIGHT
11.6.5
CONTROLS
INBOARD AILERONS
Ailerons are normally situated at the wing tips to give the greatest rolling moment for the force
produced. However, this means they are also able to generate the maximum twisting loads on
the wing. For instance, a down going aileron will twist the wing tip and decrease wing tip
incidence. The wing is not a rigid structure and any twist will cause a decrease of aileron
effectiveness. As lAS increases, a down going aileron will give more wing twist (decreased
wing tip incidence). Eventually an lAS will be reached at which the decrease in tip incidence
will give a larger down force than the up force produced by the aileron. This is called high speed
"aileron reversal"; the wing will go down, rather than up as the pilot intended. To reduce this
effect the ailerons could be mounted further inboard. Unfortunately, this would reduce aileron
effectiveness at low speed.
SPOILER SURFACES
(LIFT DUMP POSITION)
OUTBOARD AILERONS
(LOW SPEED ONLY)
Figure 11.19
Inboard and outboard ailerons & Spoiler surfaces
Alternatively, two sets of ailerons may be fitted, as illustrated in Fig. 11 .19. One set at the wing
tips for use only at low speeds when the forces involved are low, and one set inboard for use at
high speeds when the forces are greater and could cause greater structural distortion. Outboard
(low speed) ailerons are "locked-out" as the flaps retract. At low speed both sets of ailerons
work, but at high speed only the inboard ailerons respond to control input.
11.6.6
FLAPERONS
The flaps and the ailerons both occupy part of the wing trailing edge. For good take-off and
landing performance the flaps need to be as large as possible, and for a good rate of roll the
ailerons need to be as large as possible. However, the space available is limited, and one
solution is to droop the ailerons symmetrically to augment the flap area. They then move
differentially from the drooped position to give lateral control. Another system is to use the
trailing edge moveable surfaces to perform the operation of both flaps and ailerons.
11 - 12
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PRINCIPLES OF FLIGHT
11.6.7
CONTROLS
ROLL CONTROL SPOILERS
Spoilers may be used to give lateral control, in addition to, or instead of ailerons. Spoilers
consist of movable panels on the upper wing surface, hinged at their forward edge, which can
be raised hydraulically, as illustrated in Fig. 11.20. A raised spoiler will disturb the airflow over
the wing and reduce lift.
OUTBOARD AILERONS LOCKED - OUT AT
HIGH SPEED
-
SPOILER SURFACES ASSISTING
WITH ROLL CONTROL
-•• \
/
INBOARD AILERONS OPERATE
AT ALL SPEEDS
Figure 11.20
Roll control spoilers
To function as a lateral control, the spoilers rise on the wing with the up going aileron (down
going wing), proportional to aileron input. On the wing with the down going aileron, they
remain flush . Unlike ailerons, spoilers cannot give an increase of lift, so a roll manoeuvre
controlled by spoilers will always give a net loss of lift. However the spoiler has several
advantages compared to the aileron:
a)
There is no adverse yaw: The raised spoiler increases drag on that wing, so the yaw
is in the same direction as the roll.
b)
Wing twisting is reduced: The aerodynamic force on the spoilers acts further forward
than is the case with ailerons, reducing the moment which tends to twist the wing.
c)
At transonic speed its effectiveness is not reduced by shock induced separation.
d)
It cannot develop flutter.
e)
Spoilers do not occupy the trailing edge, which can then be utilised for flaps.
11 - 13
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PRINCIPLES OF FLIGHT
11.6.8
CONTROLS
COMBINED AILERON AND SPOILER CONTROLS
On a few aircraft, lateral control is entirely by spoilers, but in the majority of applications the
spoilers work in conjunction with the ailerons. Ailerons alone may be inadequate to achieve the
required rate of roll at low speeds when the dynamic pressure is low, and at high speeds they
may cause excessive wing twist and begin to lose effectiveness if there is shock induced
separation. Spoilers can be used to augment the rate of roll, but may not be required to operate
over the whole speed range. On some aircraft the spoilers are only required at low speed, and
this can be achieved by making them inoperative when the flaps are retracted.
Movement of the cockpit control for lateral control is transmitted to a mixer unit which causes
the spoiler to move up when the aileron moves up, but to remain retracted when the aileron
moves down.
11.7
SPEED BRAKES
Speed brakes are devices to increase the drag of an aircraft when it is required to decelerate
quickly or to descend rapidly. Rapid deceleration is required if turbulence is encountered at high
speed, to slow down to the Rough Air Speed as quickly as possible. A high rate of descent may
be required to conform to Air Traffic Control requirements, and particularly if an emergency
descent is required.
11.7.1
TYPES OF SPEED BRAKE
Ideally the speed brake should produce an increase in drag with no loss of lift or change in
pitching moment. The fuselage mounted speed brake is best suited to meet these requirements,
Fig. 11.21.
WING MOUNTED
SPEED BRAKES
~1
/0//
Figure 11.21
P
/
FUSELAGE MOUNTED
l:--~ I ~ SPEED BRAKE
/
Wing mounted and fuselage mounted speed brakes
11 - 14
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
CONTROLS
However, as the wing mounted spoiler gives an increase in drag, it is convenient to use the
spoiler surfaces as speed brakes in addition to their lateral control function. To operate as speed
brakes they are controlled by a separate lever in the cockpit and activate symmetrically.
Speed brakes are normally cleared for operation up to V MO / M MO but may "blow back" from
the fully extended position at high speeds. Spoilers will still function as a roll control whilst
being used as speed brakes, by moving asymmetrically from the selected speed brake position.
An example is illustrated in Fig. 11.22. Speed brakes have been selected and then a turn to the
left is initiated. The spoiler surfaces on the wing with the up going aileron will stay deployed,
or modulate upwards, depending on the speed brake selection and the roll input. The spoiler
surfaces on the wing with the down going aileron will modulate towards the stowed position.
The spoiler surfaces on the wing with the down going aileron may partially or fully stow, again
depending on the speed brake selection and the roll input.
o
SPEED BRAKES
SPEED BRAKE AND ROLL INPUT
Figure 11.22
11.7.2
Mixed speed brake and roll input
EFFECT OF SPEED BRAKES ON THE DRAG CURVE
The drag resulting from the operation of speed brakes is profile drag, so will not only increase
the total drag but will also decrease V md' This is an advantage at low speeds as the speed
stability will be better than with the aircraft in the clean configuration.
11.7.3
GROUND SPOILERS (LIFT DUMPERS)
During the landing run the decelerating force is given by aerodynamic drag, reverse thrust and
the wheel brakes. Wheel brake efficiency depends on the weight on the wheels, but this will be
reduced by any lift that the wing is producing. Lift can be reduced by operating the speed brake
lever to the lift dump position, Fig. 11.19. Both the wheel brake drag and the aerodynamic drag
are increased, and the landing run reduced. On many aircraft types, additional spoiler surfaces
are activated in the lift dumping selection than when airborne. These ground spoilers are made
inoperative in flight by a switch on the undercarriage leg which is operated by the extension of
the leg after take-off.
11 - 15
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PRINCIPLES OF FLIGHT
11.8
CONTROLS
DIRECTIONAL CONTROL
Control in yaw is obtained by the rudder. The rudder is required to:a)
maintain directional control with asymmetric power
b)
correct for crosswinds on take off and landing
c)
correct for adverse yaw
d)
recover from a spin
e)
correct for changes in propeller torque on single engined aircraft
11.8.1
EFFECT OF RUDDER DEFLECTION
If the rudder is deflected to the left the aircraft will begin to yaw to the left. This will create a
sideslip to the right. The sideslip airflow from the right acting on the fixed part of the fin will
cause a side load to the left, opposing the effect of the rudder. As the yaw increases this
damping force will increase until it balances the rudder force. The aircraft wi II then stop yawing,
and will maintain that angle of yaw, with the rudder deflected to its original position. If the
rudder is returned to the neutral position, both the fin and the rudder will give a force to the left
which will return the aircraft to its original position with zero yaw. A given rudder angle will
therefore correspond to a given yaw displacement.
11.8.2
FIN STALL
The sideslip angle is effectively the angle of attack of the fin, and as for any aerofoil, there will
be a critical angle at which it will stall. If the rudder is deflected in the direction to correct the
sideslip, the stalling angle will be reduced.
DORSALFIN ~
~
= ~
The stalling angle of an aerofoil is
affected by its aspect ratio, and so the
stalling angle of the fin could be
increased by decreasing its aspect
ratio. This can be done by fitting a
dorsal fin, Fig. 11.23.
Figure 11.23
11 - 16
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PRINCIPLES OF FLIGHT
11.8.3
CONTROLS
ASYMMETRIC THRUST
For a twin engined aircraft, if engine failure occurs, the thrust from the operating engine will
cause a yawing moment. This must be counteracted by the rudder. The rudder force will vary
with speed squared, and so there will be a minimum speed at which the force will be sufficient
to balance the engine yawing moment. This is the minimum control speed (V Md.
11.8.4
RUDDER RATIO CHANGER
450
400
350
(/)
-<
300
(/)
~
0
z
250
::.::
0
200
w
W
0..
(/)
150
100
50
0
5
10
15
20
25
30
RUDDER ANGLE - DEGREES
Figure 11.24
With a simple control system, full
rudder pedal movement will provide
full rudder deflection. With high
speed aircraft, while it is necessary to
have large rudder deflections
available at low speed, when flying at
high speed, full rudder deflection
would cause excessive loads on the
structure. To prevent this occurring a
gear change system can be
incorporated into the rudder control
system. This may be a single gear
change which gives a smaller rudder
deflection for full pedal movement
above a certain speed, or a
progressive gear change which gives
a decreasing rudder deflection with
full pedal movement as speed
increases, Fig. 11.24.
Rudder ratio
11 - 17
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PRINCIPLES OF FLIGHT
11.9
CONTROLS
SECONDARY EFFECTS OF CONTROLS
The controls are designed to give a moment around a particular axis, but may additionally give
a moment around a second axis. This coupling occurs particularly with the rolling and yawing
moments.
11.9.1
YAWING MOMENT DUE TO ROLL
a)
A rolling moment is normally produced by deflecting the ailerons, and it has been seen
that they can also produce an adverse yawing moment due to the difference in drag on
the two ailerons. Induced drag is increased on the wing with the down going aileron,
making the aircraft, for instance, roll left and at the same time, yaw right.
b)
If the aircraft is rolling, the down-going wing experiences an increased angle of attack
and the up-going wing a decreased angle of attack, increasing the adverse yawing
moment..
11.9.2
ROLLING MOMENT DUE TO YAW
a)
If the aircraft is yawing to the left, the right wing has a higher velocity than the left wing
and so will give more lift. The difference in lift will give a rolling moment to the left.
b)
If the rudder is deflected to the left (to give yaw to the left) the force on the fin is to the
right. This will give a small rolling moment to the right because the fin CP is above the
aircraft CG. This effect is usually very small, but a high fin may give an adverse roll.
One way to counteract this effect is to interconnect the ailerons and rudder so that when
the rudder is moved the ailerons move automatically to correct the adverse roll.
11 - 18
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PRINCIPLES OF FLIGHT
11.10
CONTROLS
TRIMMING
An aeroplane is trimmed when it will maintain its attitude and speed without the pilot having to
apply any load to the cockpit controls. If it is necessary for a control surface to be deflected to
maintain balance of the aircraft, the pilot will need to apply a force to the cockpit control to hold
the surface in its deflected position. This force may be reduced to zero by operation of the
trim controls.
The aircraft may need to be trimmed in pitch as a result of :
a)
changes of speed
b)
changes of power
c)
varying CO positions
Trimming in yaw will be needed:
d)
on a multi-engined aircraft if there is asymmetric power
e)
as a result of changes in propeller torque
Trimming in roll is less likely to be needed, but could be required if the configuration is
asymmetric, or if there is a lateral displacement of the CO.
11.10.1
METHODS OF TRIMMING
Various methods of trimming are in use, the main ones are:
a)
the trimming tab
b)
variable incidence (trimming) tailplane
c)
spring bias
d)
CO adjustment
e)
adjustment of the artificial feel unit
11 - 19
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
11.10.2
CONTROLS
TRIM TAB
A trim tab is a small adjustable surface set into the trailing edge of a main control surface. It' s
deflection is controlled by a trim wheel or electrical switch in the cockpit, usually arranged to
operate in an instinctive sense. To maintain the primary control surface in its required position,
the tab is moved in the opposite direction to the control surface, until the tab moment
balances the control surface hinge moment.
Fig. 11.25 shows (f x D) from tab
opposes (F x d) from control surface.
If the two moments are equal the
control will be trimmed, i.e. the stick
force will be zero. Operation of the
trim tab will slightly reduce the force
being produced by the main control
surface.
Figure 11.25
11.10.3
FIXED TABS
Some trim tabs are not adjustable in flight, but can be adjusted on the ground, to correct a
permanent out of trim condition. They are usually found on ailerons and rudder. They operate
in the same manner as the adjustable trim tab.
11.10.4
VARIABLE INCIDENCE (Trimming) TAIL PLANE
This system of trimming may be used on manually operated and power operated controls. To
trim, the tailplane incidence is adjusted by the trim wheel until the tailplane load is equal to the
previous elevator balancing load required, Fig. 1.26. Stick force is now zero.
The main advantages of a variable incidence (trimming) tailplane are:
a)
the drag is less in the trimmed state, as the aerofoil is more streamlined
b)
trimming does not reduce the effective range of pitch control, as the elevator remains
approximately neutral when the aircraft is trimmed.
c)
it is very powerful and gives an increased ability to trim for larger CG and speed range.
The disadvantage of a variable incidence (trimming) tailplane is that it is more complex and
heavy than a conventional trim tab system.
11 - 20
© Oxford Aviation Services Limited
CONTROLS
PRINCIPLES OF FLIGHT
ELEVATOR POSITIONED TO TRIM
~-
A I C STRUCTURE
SCREW JACK
-------AFTER TRIM INPUT
Figure 11.26
Variable incidence (trimming) tailplane
The amount of trim required will depend on the CG position, and recommended stabiliser takeoff settings will be given in the aircraft Flight Manual. It is important that these are correctly
set before take-off as incorrect settings could give either an excessive rate of pitch when the
aircraft is rotated, leading to possible tail strikes, or very heavy stick forces on rotation, leading
to increased take-off distances required.
5
c~
Figure 1.27
Reduced aircraft nose up pitch authority
The disadvantage of a "conventional" elevator and trim tab, Fig. 11.27, is that the aircraft nose
up pitch authority reduces with forward CG movement. Forward CG positions will require the
elevator to be trimmed more aircraft nose up. The illustration shows up elevator authority
reduced from 10 0 to 5 o .
11 - 21
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
11.10.5
CONTROLS
SPRING BIAS
In the spring bias trim system, an adjustable spring force is used to decrease the stick force. No
tab is required for this system.
11.10.6
CGADJUSTMENT
If the flying controls are used for trimming, this results in an increase of drag due to the deflected
surfaces. The out of balance pitching moment can be reduced by moving the CG, thus reducing
the balancing load required and therefore the drag associated with it. This will give an increase
of cruise range. CG movement is usually achieved by transferring fuel between tanks at the nose
and tail of the aircraft.
11.10.7
ARTIFICIAL FEEL TRIM
If the flying controls are power operated, there is no feedback of the load on the control surface
to the cockpit control. The feel on the controls has to be created artificially. When a control
surface is moved the artificial feel unit provides a force to resist the movement of the cockpit
control. To remove this force (i.e. to trim) the datum of the feel unit can be adjusted so that it
no longer gives any load on the flight deck controls.
TABS - Quick Reference Guide
Irype
of Tab ope~:ted I ~oo~~~~gl ~~~~~:
I -Ba~~nce ---~- Co~tr~1
,
Surface
----------~-
-------
=O=P=P=O=s=it=e====~-======~~~==~======~ • ~~~
!
!
i
:
,
Less
'
More
---------[-
1
Reduced
-----------+----~---------+--------------
Surface
- -
Control
Effectivenes
Stick
Force
!
i
I
---
Increased
------~~
Servo
Pilot
Spring
Pilot @ High
Speed
Trim
Trim Control
1----
+
Opposite
Less
Opposite @
High Speed
Less@
High
Speed
Opposite
Zero'd
Reduced
Reduced
~ _ _ _ _ _ _~_ _ONLY
_ _ _ _~ _ _ _ _ _ _ _~ _ _ _ _~_ _ _ _ _ _ _ _ _ i
11 - 22
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
CONTROLS
SELF ASSESSMENT QUESTIONS
1.
An elevon is:
a)
b)
c)
d)
2.
When rolling at a steady rate the:
a)
b)
c)
d)
3.
lateral control about the lateral axis.
longitudinal control about the lateral axis.
lateral control about the longitudinal axis.
directional control about the normal axis.
Aileron reversal would be most likely to occur:
a)
b)
c)
d)
6.
the rudder.
the ailerons.
the elevators.
the flaps.
Ailerons give:
a)
b)
c)
d)
5.
up going wing experiences an increase in effective angle of attack
rate of roll depends only on aileron deflection
down going wing experiences an increase in effective angle of attack
effective angle of attack of the up going and down going wings are equal
The control surface which gives longitudinal control is:
a)
b)
c)
d)
4.
an all moving tailplane that has no elevator
the correct name for a V - tail
a surface that extends into the airflow from the upper surface of the wing to reduce the
lift
a combined aileron and elevator fitted to an aircraft that does not have conventional
horizontal stabiliser (tailplane)
with a rigid wing at high speed.
with a flexible wing at high speed.
with a rigid wing at low
with a flexible wing at low speed.
If the ailerons are deflected to 10 0, compared to 50, this will cause:
a)
b)
c)
d)
an increased angle of bank.
an increased rate of roll.
no change to either bank angle or roll rate.
a reduction in the adverse yawing moment.
11 - 23
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
7.
Yawing is a rotation around:
a)
b)
c)
d)
8.
elevator moves up.
elevator moves down.
elevator down.
elevator moves up.
roll to starboard
pitch nose up
roll first to starboard and then to port
roll to port
a yawing moment to the left but no rolling moment.
a rolling moment to the left.
a rolling moment to the right.
a yawing moment to the right but no rolling moment.
What should be the feel on a 'full and free' check of the controls:
a)
b)
c)
d)
12.
left aileron moves up, right aileron moves down,
left aileron moves down, right aileron moves up,
left aileron moves up, right aileron moves down,
left aileron moves down, right aileron moves up,
Due to the AC of the fin being above the longitudinal axis, if the rudder is moved to the right,
the force acting on the fin will give:
a)
b)
c)
d)
11.
the
the
the
the
The secondary effect of yawing to port is to:
a)
b)
c)
d)
10.
the normal axis obtained by elevator.
the lateral axis obtained by rudder.
the longitudinal axis obtained by ailerons.
the normal axis obtained by rudder.
If the control column is moved forward and to the left:
a)
b)
c)
d)
9.
CONTROLS
a gradual stiffening of the controls.
rebound on reaching the stops.
a solid stop.
controls should not be moved to the stops.
The purpose of control locks on a flying control system is:
a)
b)
c)
d)
to enable any free movement in the control system to be detected.
to prevent structural damage to the controls in gusty conditions when the aircraft is on
the ground.
to keep the control surface rigid to permit ground handling.
as a security measure.
11 - 24
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
13.
An irreversible control:
a)
b)
c)
d)
14.
the stalling angle of that wing is increased
the stalling angle of that wing is decreased
the stalling angle is not affected but the stalling speed is decreased
When rudder is used to give a coordinated tum to the left:
a)
b)
c)
18.
the control column must be pushed forward
the control column must be pulled backwards
the control wheel must be rotated
the incidence of the tailplane must be decreased because the negative camber will make
it effective in the reverse sense
If an aileron is moved downward:
a)
b)
c)
17.
to increase the feel in the control circuit
to correct for adverse aileron yaw
to allow for up-float in flight to bring the aileron into the streamlined position
to give a higher CL max for take-off
The tailplane shown has inverted camber. To cause the aircraft to pitch nose up the control
column must be:
a)
b)
c)
d)
16.
may be moved by operating the cockpit control but not by the aerodynamic loads acting
on the control surface.
has less movement in one direction than the other.
may be moved either by the cockpit control or by a load on the control surface.
is when the control locks are engaged.
Ailerons may be rigged slightly down (drooped):
a)
b)
c)
d)
15.
CONTROLS
the left pedal is moved forward, and the rudder moves right
the right pedal is moved forward and the rudder moves left
the left pedal is moved forward and the rudder moves left
The higher speed of the upper wing in a steady banked tum causes it to have more lift than the
lower wing. This may be compensated for by:
a)
b)
c)
use of the rudder control
operating the ailerons slightly in the opposite sense once the correct angle of bank has
been reached
increasing the nose up pitch by using the elevators
11 - 25
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
19.
The purpose of a differential aileron control is to:
a)
b)
c)
d)
20.
b)
c)
d)
an arm projecting upward from the control surface to which the control cables are
attached.
a proj ection of the outer edge of the control surface forward of the hinge line.
a rod projecting forward from the control surface with a weight on the end.
a projection of the leading edge of the control surface below the wing undersurface.
An aileron could be balanced aerodynamically by:
a)
b)
c)
d)
24.
to get the aircraft into balance.
to prevent flutter of the flying control.
to reduce the control load to zero.
to make the control easier to move.
A horn balance on a control surface is:
a)
23.
the up going aileron causes an increase in induced drag.
the down going aileron causes an increase in induced drag.
both cause an increase in induced drag.
induced drag remains the same, the up going aileron causes a smaller increase in profile
drag than the down going aileron.
The purpose of aerodynamic balance on a flying control is:
a)
b)
c)
d)
22.
give a yawing moment which opposes the turn
reduce the yawing moment which opposes the turn
give a pitching moment to prevent the nose from dropping in the turn
improve the rate of roll
When displacing the ailerons from the neutral position:
a)
b)
c)
d)
21.
CONTROLS
making the up aileron move through a larger angle than the down aileron.
attaching a weight to the control surface forward of the hinge.
having the control hinge set back behind the control surface leading edge.
having springs in the control circuit to assist movement.
Control overbalance results in:
a)
b)
c)
d)
a sudden increase in stick force
a sudden reduction then reversal of stick force
a sudden loss of effectiveness of the controls
a gradual increase in stick force with increasing IAS
11 - 26
© Oxford Aviation Services Limited
CONTROLS
PRINCIPLES OF FLIGHT
25.
A control surface is mass balanced by:
a)
b)
c)
d)
26.
If the control wheel is turned to the right, a balance tab on the port aileron should:
a)
b)
c)
d)
27.
will move up relative to the control surface.
will move down relative to the control surface.
will only move if the trim wheel is operated.
moves to the neutral position.
The purpose of a spring tab is:
a)
b)
c)
d)
30.
trim the aircraft
reduce the load required to move the controls at all speeds
reduce the load required to move the controls at high speeds only
give more feel to the controls
When the control column is pushed forward a balance tab on the elevator:
a)
b)
c)
d)
29.
move up relative to the aileron
move down relative to the aileron
not move unless the aileron trim wheel is turned.
move to the neutral position
The purpose of an anti-balance tab is to:
a)
b)
c)
d)
28.
fitting a balance tab.
attaching a weight acting forward of the hinge line.
attaching a weight acting on the hinge line.
attaching a weight acting behind the hinge line.
to maintain a constant tension in the trim tab system.
to increase the feel in the control system.
to reduce the pilot's effort required to move the controls against high air loads.
to compensate for temperature changes in cable tension.
The purpose of a trim tab is:
a)
b)
c)
d)
to assist the pilot in initiating movement of the controls.
to zero the load on the pilots controls in the flight attitude required.
to provide feel to the controls at high speed./
to increase the effectiveness of the controls.
11 - 27
© Oxford Aviation Services Limited
CONTROLS
PRINCIPLES OF FLIGHT
31.
To re-trim after failure of the right engine on a twin-engine aircraft:
a)
b)
c)
d)
32.
To trim an aircraft which tends to fly nose heavy with hands off, the top of the elevator trim
wheel should be:
a)
b)
c)
d)
33.
greater than that from an elevator.
the same as that from an elevator.
less than that from an elevator.
Following re-trimming for straight and level flight because of forward CG movement:
a)
b)
c)
d)
35.
moved forward to raise the nose and this would cause the elevator trim tab to move
down, and the elevator to move up.
moved backwards to raise the nose, and this would cause the elevator trim tab to move
down, and the elevator to move up.
moved backwards to raise the nose, and this would cause the elevator trim tab to move
up, and the elevator to move up.
be moved backwards to raise the nose, and this would cause the elevator trim tab to
move up and cause the nose to rise.
To achieve the same degree of longitudinal trim, the trim drag from a variable incidence
trimming tailplane would be:
a)
b)
c)
34.
the rudder trim tab will move right and the rudder left.
the trim tab will move left and the rudder right.
the trim tab will move left and the rudder remain neutral.
the trim tab will move right and the rudder remain neutral.
nose up pitch authority will be reduced
nose down pitch authority will be reduced
longitudinal stability will be reduced
tailplane down load will be reduced
An aircraft has a tendency to fly right wing low with hands off. It is trimmed with a tab on the
left aileron. The trim tab will:
a)
b)
c)
d)
move up, causing the left aileron to move up and right aileron to move down.
move down, causing the left aileron to move up, right aileron remains neutral.
move down causing the left aileron to move up, and right aileron to move down.
move up causing the left wing to move down, ailerons remain neutral.
11 - 28
© Oxford Aviation Services Limited
CONTROLS
PRINCIPLES OF FLIGHT
36.
An aircraft takes offwith the elevator control locks still in position. It is found to be nose heavy:
a)
b)
c)
d)
37.
On a servo tab operated elevator, if the pilot's control column is pushed forward in flight:
a)
b)
c)
d)
38.
the control surfaces and servo tabs are free.
the control surfaces are free but there could be locks on the servo tabs.
there could be locks on the control surfaces and on the servo tabs.
the servo tabs are free but there could be locks on the control surfaces.
In a servo operated aileron control system, turning the cockpit control wheel to the right in flight
will cause the servo tab on the left aileron:
a)
b)
c)
d)
40.
the servo tab will move down causing the elevator to move up.
the elevator will move down causing the servo tab to move up.
the elevator will move up causing the servo tab to move down.
the servo tab will move up causing the elevator to move down.
If a cockpit control check is made on an aircraft with servo operated controls, and it is found that
the cockpit controls move fully and freely in all directions:
a)
b)
c)
d)
39.
backward movement of the trim wheel would increase nose heaviness.
it would not be possible to move the trim wheel.
backward movement of the trim wheel would reduce nose heaviness.
operating the trim wheel would have no effect.
to move up and the left aileron to move down
to move down and the left aileron to move down
to move down and the left aileron to move up
to move up and the right aileron to move down
Spoilers on the upper surface of the wing may be used on landing:
a)
b)
c)
d)
to give a nose down pitching moment
to reduce the lift and so put more weight on the wheels, making the brakes more
effective
to cause drag and increase the lift from the flaps
to reduce the touchdown speed
11 - 29
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
41.
Wing mounted spoiler surfaces may be used as:
a)
b)
c)
d)
42.
reinforce the boundary layer
create turbulence at the wing root
increase the camber at the wing root
decrease lift on the upper wing surface when deployed asymmetrically
On an aircraft fitted with roll control spoilers, a roll to port is achieved by:
a)
b)
c)
d)
44.
air brakes
lift dumpers
lateral control
all of the above
Spoilers, when used for roll control will:
a)
b)
c)
d)
43.
CONTROLS
deflecting the port spoiler up and starboard down
deflecting the starboard spoiler down
deflecting the port spoiler up
deflecting the port spoiler down
In a fully power operated flying control system control feel is provided by:
a)
b)
c)
d)
the friction in the control cable system.
an artificial feel unit (Q - Feel)
the aerodynamic loads on the control surface.
the mass balance weights.
11 - 30
© Oxford Aviation Services Limited
CONTROLS
PRINCIPLES OF FLIGHT
ANSWERS
I No I A I B I c I DII
1
D
REF
III
No
I A I B I c I D II
23
C
24
B
3
C
25
B
4
C
26
B
27
6
B
28
7
D
8
9
D
10
13
A
29
C
31
B
A
B
33
C
12
D
32
B
11
A
30
C
34
B
C
A
35
A
14
36
C
C
A
15
B
37
D
16
B
38
D
17
39
C
A
18
B
40
19
B
41
D
20
B
42
D
21
22
D
B
B
43
44
11 - 31
I
C
2
5
REF
C
B
© Oxford Aviation Services Limited
CHAPTER 12 - FLIGHT MECHANICS
Contents
Page
INTRODUCTION ...................................................... 12 - 1
STRAIGHT HORIZONTAL STEADY FLIGHT
TAILPLANE AND ELEVATOR .................................... 12 - 2
BALANCE OF FORCES ........................................... 12 - 3
STRAIGHT STEADY CLIMB ............................................ 12 - 4
CLIMB ANGLE ................................................. 12 - 5
EFFECT OF WEIGHT, ALTITUDE AND TEMPERATURE
POWER ON DESCENT .................................................. 12 - 6
EMERGENCY DESCENT ......................................... 12 - 7
GLIDE ... , ................................................. , .... , ..... 12 - 8
ANGLE OF DESCENT IN THE GLIDE
EFFECT OF WEIGHT ...................................... 12 - 9
EFFECT OF WIND ....... , ............................... 12 - 10
EFFECT OF CONFIGURATION
RATE OF DESCENT IN THE GLIDE
TURNING
EFFECT OF WEIGHT ON TURNING ............................... 12 - 11
RADIUS AND RATE OF TURN ... , .......... , .... , ............... 12 - 14
LOAD FACTOR IN THE TURN ................................... 12 - 16
'G' LIMIT ON TURNING ........................................ 12 - 17
STALL LIMIT ON TURNING
THRUST LIMIT ON TURNING
MINIMUM TURN RADIUS
TURN CO-ORDINATION ........................................ 12 - 18
FLIGHT WITH ASYMMETRIC THRUST .................................. 12 - 20
YA WING MOMENT
CRITICAL ENGINE ............................................. 12 - 22
BALANCING THE YAWING MOMENT ............................ 12 - 23
ROLLANDYAWMOMENTS .................................... 12-24
MINIMUM CONTROL AIRSPEED ................................. 12 - 27
YMCA
FACTORS AFFECTING V MCA
••••• " • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • •
YMca • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • . • • • • • • • •
FACTORS AFFECTING YMca
V MCL • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • .
FACTORS AFFECTING V MCL • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • • . • • •
SUMMARY OF MINIMUM CONTROL SPEEDS
PERFORMANCE WITH ONE ENGINE INOPERATIVE ......................
SINGLE ENGINE ANGLE OF CLIMB
SINGLE ENGINE RATE OF CLIMB
CONCLUSIONS
SELF ASSESSMENT QUESTIONS .......................................
ANSWERS ....................................................
12 - 28
12 - 29
12 - 30
12 - 31
12 - 32
12 - 35
12 - 43
PRINCIPLES OF FLIGHT
12.1
FLIGHT MECHANICS
INTRODUCTION
Flight Mechanics is the study of the forces acting on an aircraft in flight, and the response of the
aircraft to those forces. For an aircraft to be in steady (unaccelerated) flight, the following
conditions must exist:a)
The forces acting upward must exactly balance the forces acting downward,
b)
the forces acting forward must exactly balance the forces acting backward, and
c)
the sum of all moments must be zero.
This condition is known as equilibrium.
12.2
STRAIGHT HORIZONTAL STEADY FLIGHT
In straight and level flight there are four forces acting on the aircraft; LIFT, WEIGHT, THRUST
and DRAG, as shown in Fig. 12.1.
Weight acts through the aircraft centre of gravity (CG), vertically downwards towards the centre
of the earth. Alternatively, weight can be defined as acting parallel to the force of gravity.
Lift acts through the centre of pressure (CP), normal (at 90 °) to the flight path.
For the purposes of this chapter (although not strictly true), thrust acts forwards, parallel to the
flight path and drag acts backwards, parallel to the flight path.
AERODYNAMIC
DRAG
L
THRUST
REQUIRED
TO BALANCE
AERODYNAMIC DRAG
w
Figure 12.1
Forces in Level Flight
12 - 1
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
FLIGHT MECHANICS
For an aircraft to be in steady level flight a condition of equilibrium must exist. This
unaccelerated condition of flight is achieved with the aircraft trimmed with lift equal to weight
and the throttles set for thrust to equal drag. It can be said that for level flight the opposing
forces must be equal.
The LID ratio of most modern aircraft is between 10 and 20 to 1. That is, lift is 10 to 20 times
greater than drag.
The lines of action of thrust and drag lie very close together, so the moment of this couple is very
small, and can be neglected for this study. The position of the CP and CG are variable and under
most conditions oflevel flight are not coincident. The CP moves forward with increasing angle
of attack and the CG moves with reduction in fuel. Generally, the CP is forward of the CG at
low speed, giving a nose up pitching moment and behind the CG at high speed, giving a nose
down pitching moment.
12.3
TAILPLANE AND ELEVATOR
The function of the tailplane is to maintain equilibrium by supply the force necessary to counter
any pitching moments arising from CP and CG movement. With the CP behind the CG during
normal cruise, as illustrated in Fig. 12.2, the tailplane must supply a down force.
~
MOMENT DUE TO
LIFT I WEIGHT COUPLE
TAIL
MOMENT
TAIL
DOWN
FORCE
w
Figure 12.2
Tailplane Maintains Equilibrium
12 - 2
© Oxford Aviation Services Limited
FLIGHT MECHANICS
PRINCIPLES OF FLIGHT
12.4
BALANCE OF FORCES
If the tailplane is producing a balancing force, this will add to or subtract from the lift force.
For a down load:
For an up load:
-
Lift
tailplane force
Lift + tailplane force
STALL ANGLE
ANGLE
OF
ATTACK
CONSTANT
LIFT
lAS
Figure 12.3
Variation of angle of attack with lAS
Weight
Weight
For steady level flight at a
constant weight, the lift force
required will be constant. At
a steady speed the wing will
give this lift at a given angle
of attack. However if the
speed is changed the angle of
attack must change to
maintain the same lift. As
the lift changes with the
square of the speed, but in
direct proportion to the angle
of attack, the angle of attack
will vary as shown in Fig.
12.3 to give a constant lift.
For steady level flight at a constant speed, the thrust must equal the drag. Drag increases with
speed (above V md) and so to maintain a higher speed, the thrust must be increased by opening
the throttle.
To fly at the speed at point
A, Fig. 12.4, requires a thrust
of T I and to fly at point B
requires a thrust ofT 2 • If the
thrust is increased from T I to
THRUST
T 2 when the aircraft is at
AND
point A, the thrust will be
DRAG
greater then the drag, and the
aircraft will accelerate in
proportion to the 'excess'
thrust AC until it reaches
point B, where the thrust and
the
drag are again equal. If
lAS
T2 is the thrust available with
the throttle fully open, then
the speed at B is the
Figure 12.4 Balance of Thrust and Drag
maximum speed achievable
in level flight.
12 - 3
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
12.5
FLIGHT MECHANICS
STRAIGHT STEADY CLIMB
Consider an aircraft in a straight steady climb along a straight flight path inclined at an angle (y)
to the horizontal. y (gamma) is the symbol used for climb angle. The forces on the aircraft
consist of Lift, normal to the flight path; Thrust and Drag, parallel to it; and Weight, parallel to
the force of gravity. This system of forces is illustrated in Fig. 12.5.
THRUST
REQUIRED
TO BALANCE
AERODYNAMIC DRAG
\
\
~ATH
I
\
\
I
\
/
\
\ :----
\
Y\
~ w cosy
/
/
/
/
- - -7 - -
BACKWARDS
COMPONENT
OF WEIGHT ---~ W sinY
EXTRA THRUST
REQUIRED TO
BALANCE
BACKWARDS
COMPONENT
OF WEIGHT
Figure 12.5 Forces in a Steady Climb
Weight is resolved into two components: one opposite Lift (W cos y), and the other acting in
the same direction as Drag (W sin y), backwards along the flight path. The requirements for
equilibrium are: Thrust must equal the sum of Drag plus the backwards component of Weight;
and Lift must equal its opposing component of Weight. For equilibrium at a greater angle of
climb, the Lift required will be less, and the backwards component of Weight will be greater.
L
T
=
=
W c'Osy
D + W siny
In a straight steady climb, Lift is less than Weight because Lift only has to support a proportion
of the weight, this proportion decreasing as the climb angle increases. (In a vertical climb no lift
is required). The remaining proportion of Weight is supported by engine Thrust.
12 - 4
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
FLIGHT MECHANICS
It can be seen that for a straight steady climb the Thrust required is greater than Drag. This is
to balance the backward component of Weight acting along the flight path.
Sin y
=
T- D
W
The ability of an aircraft to climb depends upon EXCESS THRUST, available after opposing
aerodynamic drag. The smaller the Drag for a given Thrust, the greater the ability to climb.
Drag will be less with flaps up, giving a larger climb angle (improved climb gradient).
12.5.1
CLIMB ANGLE
Climb angle depends on "excess Thrust" ( T - D ) and the Weight. As both Thrust and Drag
vary with lAS, excess Thrust will be greatest at one particular speed. This is the speed for
maximum angle of climb, V x . (see Fig. 12.27 for the propeller case).
DRAG
THRUST
AND
DRAG
THRUST (JET)
MAXIMUM
DIFFERENCE BE1W EEN
THRUST AND DRAG
lAS
Figure 12.6
Variation of excess thrust with speed (JET)
The variation of Thrust with speed will depend on the type of engine. For a jet engine, where
Thrust is fairly constant with speed, Vx will be near to VMD, but for a propeller engined aircraft
V x will usually be below V M D'
12.5.2
EFFECT OF WEIGHT, ALTITUDE AND TEMPERATURE.
The Drag of an aircraft at a given lAS is not affected by altitude or temperature, but higher
Weight will increase Drag and reduce excess Thrust and consequently the climb angle.
Thrust available from the engine decreases with increasing altitude and increasing temperature,
which also reduces excess Thrust. Climb angle therefore decreases with increasing Weight,
altitude and temperature.
12 - 5
© Oxford Aviation Services Limited
FLIGHT MECHANICS
PRINCIPLES OF FLIGHT
12.6
POWER ON DESCENT
I
ENGINE THRUST
I
~
I
\
\
FLIGHT PATH
\
/
\
Figure 12.7 Forces in a Power - On Descent
Figure 12.7 illustrates the disposition of forces in a steady Power-On descent. The force of
Weight is split into two components. One component (W cos y) acts perpendicular to the flight
path and is balanced by Lift, while the other component (W sin y) acts forward along the flight
path and 'adds' to the Thrust to balance Drag. If the nose of the aircraft is lowered with a
constant Thrust setting the increased component of Weight acting forward along the flight path
will cause an increase in lAS. The increased lAS will result in an increase in Drag which will
eventually balance the increased forward force of Weight and equilibrium will be re-established.
If the throttle is closed the force of Thrust is removed and a larger forward component of Weight
must be provided to balance Drag and maintain a constant lAS. This is accomplished by
lowering the nose of the aeroplane to increase the descent angle (y).
a)
In a descent Lift is less than Weight. This is because Lift only has to balance the
component of Weight perpendicular to the flight path (W cos y).
b)
In a descent Thrust is less than Drag. This is because Weight is giving a forward
component in the same direction as Thrust (W sin y).
12 - 6
© Oxford Aviation Services Limited
FLIGHT MECHANICS
PRINCIPLES OF FLIGHT
12.7
EMERGENCY DESCENT
In the event of cabin pressurisation failure at high altitude it is necessary to descend as quickly
as possible. The rate of descent can be increased by:
1.
Reducing Thrust by closing the throttles.
2.
Increasing Drag by:
a)
b)
3.
extending the speedbrakes,
lowering the landing gear (at or below V LO)'
Increasing speed by lowering the nose.
Speed can be increased in the clean configuration up to MMO or V MO depending on the
altitude, or to the gear extended limit speed (V LE) if the gear is down.
The overall rate of descent will be higher with the landing gear extended (lots of Drag), but if
the gear operating limit speed (V LO) is much less than the cruising speed the aircraft will have
to be slowed down before the gear can be lowered (perhaps taking several minutes in level
flight). So the initial rate of descent will be relatively low and the time spent at high altitude will
be extended.
If the gear is not extended, throttles can be closed, speedbrakes extended and the nose lowered
to accelerate the aircraft to MMON MO immediately, giving a higher initial rate of descent and
getting the passengers down to a lower altitude without delay.
At high altitude the limiting speed will be MMO and if an emergency descent is made at this Mach
number the lAS will be increasing. At some altitude the lAS will reach V MO and the nose must
then be raised so as not to exceed V MO for the remainder of the descent.
The rate of descent possible during an emergency descent can be quite high, so as the required
level-off altitude is approached the rate of descent should be reduced progressively so as to give
a smooth transition back to level flight.
12 - 7
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
12.8
FLIGHT MECHANICS
GLIDE
In a glide without Thrust, the Weight component along the flight path must supply the propulsive
force and balance Drag. In a glide there are only three forces acting on the aircraft, Lift, Weight
and Drag.
-
TOTAL
REACTION ---~~..
/
/
~
-
____
I
/
I
~
"
L =W cosy
'\
\
/
\
/
~
/
\
\
I
----:t}T
I
\
\
\
\
\
\
I
~
I
I
I
I
I
I
I
I
I
I
'-
FLIGHT PATH
/
/
I
'----
/
/
FORWARD COMPONENT
, , \~W___~_/_~7E1GHT
(W sin y)
Figure 12.8 Forces in the Glide
Fig. 12.8 shows the disposition of forces in a steady glide. The forward component of Weight
(W sin y) is a product of descent angle (y); the greater the descent angle, the greater the forward
component of weight (compare with Fig. 12.7). The forward component of weight must balance
Drag for the aircraft to be in a steady glide. It follows that if Drag is reduced and Lift remains
constant, the required balance of forces can be achieved at a smaller descent angle.
12.8.1
ANGLE OF DESCENT IN THE GLIDE
Glide angle is a function ONLY of the LID ratio. The descent (glide) angle will be least when
the LID ratio is the greatest. LID ratio is a maximum at the optimum angle of attack, and this
also corresponds to the minimum drag speed (V MO), Fig. 12.9. At speeds above or below V MO
the glide angle will be steeper.
Maximum distance in a glide can be achieved when the aircraft is flown at LID MAX (V MO)'
12 - 8
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
12.8.2
FLIGHT MECHANICS
EFFECT OF WEIGHT
LID MAX is independent of weight. Provided the aircraft is flown at its optimum angle of attack,
the glide angle and glide distance will be the same whatever the weight. The speed
corresponding to the optimum angle of attack, (V MO) will however change with weight. V MO
increases as weight increases.
/t~
/
/
/ ~ -+ i - ~~L
U~
/
I
I
/
/
/
/
:y
~
'" \
:
:
/
/
~
/
\
I
I
I
I
I
I
I
\
\
I
\
\
\
I
\
/
/
/
"
/
\
'"
~
/
I
-:YJ[-
/
/
/
~ - - - I---:§
,'
L ~/ /
/
FLIGHT PATH
/
/
/
/
~! W
Figure 12.8a Increased Weight: no effect on glide range
As illustrated in Fig. 12.8a, a higher weight will give an increased forward component of weight
and the aircraft will accelerate towards the resultant higher V MO. As the aircraft accelerates, Lift
increases and Drag will increase until it balances the increased forward component of Weight.
Equilibrium is now re-established at the same LID MAX, but a higher lAS.
At a higher weight the aircraft will glide the same distance, but at a higher speed and
consequently have an increased RATE of descent.
12 - 9
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
12.8.3
FLIGHT MECHANICS
EFFECT OF WIND
The glide angle will determine the distance that the aircraft can glide for a given change of
height.
GLIDE DISTANCE
=
LIFT (L)
DRAG (D)
HEIGHT LOST x
This distance would only be achieved in still air. If there is a wind the ground speed will change,
and so the distance over the ground will change. In a headwind the ground distance will be
decreased and in a tailwind it will be increased.
12.8.4
EFFECT OF CONFIGURATION
The maximum LID ratio of an aircraft will be obtained in the clean configuration. Extension of
flaps, spoilers, speedbrakes or landing gear etc. will reduce LlD MAX and give a steeper glide
angle, thus reducing glide range.
12.9
RATE OF DESCENT IN THE GLIDE
Minimum rate of descent in the glide is obtained at the lAS which produces minimum Power
Required (VMP) . Flying at V MP in a glide will enable the aircraft to stay airborne for as long as
possible. As shown in Fig. 12.9, V MP is a slower lAS than V MD. Wind speed and direction has
no effect on rate of descent. A frequently used method of showing the relationship of V MD and
V MP is by use of the 'whole aeroplane CL/C D polar' curve, illustrated in Fig. 12.10.
DRAG
J
UDMAX 0IMD)
I.
/
1· 32 VMD
/
i
v MP V MD
lAS
Figure 12.9
12.10
Figure 12.10
TURNING
For an aircraft to change direction, a force is required to deflect it towards the centre of the tum.
This is called the centripetal force, Fig. 12.11. Banking the aircraft inclines the Lift. It's the
horizontal component of Lift which causes the aircraft to turn. If the aircraft is banked and
the angle of attack kept constant, the vertical component of lift will be too small to balance the
weight and the aircraft will start to descend.
12 - 10
© Oxford Aviation Services Limited
FLIGHT MECHANICS
PRINCIPLES OF FLIGHT
As the angle of bank increases, the angle of attack must be increased to bring about a greater
total lift. The vertical component must be large enough tu maintain level flight, while the
horizontal component is large enough to produce the required centripetal force.
12.10.1
EFFECT OF WEIGHT ON TURNING
In a steady level turn, if thrust is ignored, lift provides a force to balance weight and centripetal
force to turn the aircraft. If the same TAS and angle of bank can be obtained, the radius of
turn is basically independent of weight or the aircraft type.
Not all aircraft can reach the same angle of bank at the same TAS. If weight increases, the
vertical component of lift required increases, but the centripetal force to maintain the same
radius of turn also increases in the same proportion. The lift required, although it is greater, has
the same inclination to the vertical as before and the bank angle is the same, Fig. 12.12.
Lift I
I
I
I
I
I
Weight
Lift~, -
Figure 12.11
I "
Forces in a turn
I
. . Increased
T Weight
Figure 12.12
12 - 11
Bank angle and weight
© Oxford Aviation Services Limited
FLIGHT MECHANICS
PRINCIPLES OF FLIGHT
CENTRE
OF TURN
CENTRIPETAL FORCE
L
sin
<p
r
w
Figure 12.13 Forces acting in a steady turn
In a steady horizontal tum, Fig. 12.13, the conditions of equilibrium can be expressed in the
form:
L cos <l> = W
(Eq 12. 14)
L sin <l> =
(Eq 12. 15)
Where (L) is the wing lift in Newtons, (W) is the weight of the aircraft in Newtons, (V) the true
air speed in mis, (r) the radius of tum in metres, <p the angle of bank and (g) the acceleration of
gravity constant of9·81 mls.
Dividing equation 12.15 by equation 12.14 we get:
tan <l>
=
v2
(Eq 12. 16)
rg
which is the basic turning equation relating (V), (r) and <p o Once two of these variables are
known, the other two can be determined. From equation 12.16, the radius of tum is given by:
turn radius =
9
(Eq 12. 17)
tan <l>
12 - 12
© Oxford Aviation Services Limited
FLIGHT MECHANICS
PRINCIPLES OF FLIGHT
and the corresponding rate of tum (= V / r ) by:
rate of turn
=
g
tan <P radians / second
V
(Eq 12. 18)
Rate of tum is the rate of change of heading or angular velocity of the tum. It may be expressed
as degrees per minute, or by a Rate Number.
Rate 1 tum is 180 0 per minute (3 0 per second)
Rate 2 tum is 360 0 per minute (6 0 per second)
Rate of tum is directly proportional to TAS and inversely proportional to the tum radius.
Rate of turn
=
TAS
Radius
For example: at a speed of 150 kt TAS (77 m/s), an aircraft performing a tum with a radius of
1480 metres would have a rate of tum of:
77
1480
=
0.052 radians / sec
there being 2n radians in a circle,
0.052 x 57.3
=
360
6.286
=
57.3 0 per radian
3 0 per second (Rate one)
a)
at a constant TAS, increasing the angle of bank decreases the tum radius and increases
the rate of tum.
b)
to maintain a constant rate of tum, increasing speed requires an increased bank angle.
c)
at a constant bank angle, increasing speed increases the tum radius and decreases the
rate of tum.
In a constant rate turn,
the angle of bank
is dependent upon TAS
12 - 13
© Oxford Aviation Services Limited
FLIGHT MECHANICS
PRINCIPLES OF FLIGHT
12.10.2
RADIUS AND RATE OF TURN
Two variables determine the rate of turn and radius oftum:a)
Bank angle (<p). A steeper bank reduces turn radius and increases the rate of turn, but
produces a higher load factor.
b)
True air speed (TAS): Reducing speed reduces turn radius and increases the rate of
turn, without increasing the load factor.
The radius of turn at any given bank angle (<p), varies directly with the square of the T AS:
radius
=
9 tan <f>
If speed is doubled, the turn radius will be four times greater, at a constant bank angle.
To appreciate the relationship between radius of turn and rate of turn at double the speed,
consider:
rate of tu rn =
rate of tu rn =
v
radius
v
(x 2)
radius (x 4 )
=
1
2
If speed is doubled, the rate of turn will be half of its previous value, at a constant bank
angle.
Because the rate of turn varies with TAS at any given bank angle, slower aeroplanes require less
time and area to complete a turn than faster aeroplanes with the same bank angle, Fig.12.14.
A specific angle of bank and T AS will produce the same rate and radius of turn regardless of
weight, CG position, or aeroplane type. It can also be seen from Fig.12.14 that increasing speed
increases the turn radius and decreases the rate of turn. The load factor remains the same
because the bank angle has not changed. /
To increase the rate and decrease the radius of turn, steepen the bank and / or decrease the speed.
A given TAS and bank angle will produce a specific rate and radius of turn in any aeroplane.
In a co-ordinated level turn, an increase in airspeed will increase the radius and decrease the rate
of turn. Load factor is directly related to bank angle, so the load factor for a given bank
angle is the same at any speed.
12 - 14
© Oxford Aviation Services Limited
FLIGHT MECHANICS
PRINCIPLES OF FLIGHT
\
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80
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300- - 0·3
200 -
0.2
0
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TAS, Kts
1
1
I
1
Figure 12.14 (For illustration purposes only). This chart will work for any aeroplane. The
example shows that for a turn at 130 kts TAS and a bank angle of20 °, the radius will be 4,200
ft and the rate of turn will be 3 0 per second. At 260 kts T AS the radius will be 16,800 ft and the
rate of turn will be 1.5 per second.
0
12 - 15
© Oxford Aviation Services Limited
FLIGHT MECHANICS
PRINCIPLES OF FLIGHT
12.10.3
LOAD FACTOR IN THE TURN
When an aircraft is in a banked turn lift must be increased so as to maintain the vertical
component of lift equal to weight, Fig. 12.15.
~---
INCREASED LIFT
w
30° BANK ANGLE
60° BANK ANGLE
Figure 12.15
Increased lift required in a turn
Thisrelationship may be expressed as:
Load factor (n)
=
L
W
=
1
cos <P
=
sec
<p
(Eq 12.19)
Refer to Page 7 - 24 and 7 - 25 for the full trigonometrical explanation.
Fig. 12.16 shows the relationship between load factor and bank angle. This chart will be
effective for any aircraft. It can be seen that load factor (n) increases with bank angle at an
increasing rate.
Load factor in the turn is a function ONLY of bank angle.
Constant bank angle,
constant load factor
12 - 16
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
FLIGHT MECHANICS
I
9
8
:Cl
0:::
0
7
6
I-
5
«
LL
4
0
3
0
«
0
.....I
J
L
IJ
-
-- -
f--
-
-
2
1/
1--
-
."
I
0
o
I
I
I
10
20
30
40
50
60
70
80
90
BANK ANGLE IN DEGREES
Figure 12.16
12.10.4
Relationship between 'g' and bank angle
'g' LIMIT ON TURNING
For each aircraft there is a design limit load factor. For modern high speed jet transport aircraft
the positive limit load factor is 2·5 g. From Fig. 12.16 it can be seen that this would occur at a
bank angle of 67° and this will determine a turn radius, depending on the T AS. This will be the
minimum radius permissible at that 'g' ifthe strength limit is not to be exceeded.
12.10.5
STALL LIMIT ON TURNING
If speed is kept constant, but the bank angle increased, the angle of attack must also be increased
to provide the increased lift required. Eventually the stalling angle will be reached and no
further increase in bank angle (and decrease in turn radius) is possible. Because the stalling
speed varies with weight, this boundary will be a function of weight.
12.10.6
THRUST LIMIT ON TURNING
During a turn lift must be greater than during level flight, and this will result in increased
induced drag. To balance this additional drag, more thrust is required in a turn than for level
flight at the same speed. The greater the bank angle the greater will be the thrust required, and
eventually the throttle will be fully open. No further increase in bank angle (and decrease in turn
radius) is then possible. The relative positions of the thrust boundary and the strength boundary
will depend on the limit load factor and thrust available.
12.10.7
MINIMUM TURN RADIUS
If the thrust available is adequate, the minimum radius of turn occurs at the intersection of the
stall limit and the strength limit. The speed at this point is V A the maximum manoeuvring speed.
The heavier the aircraft, the greater the minimum radius of turn.
12 - 17
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
12.10.8
FLIGHT MECHANICS
TURN CO-ORDINATION
Adverse aileron yaw, engine torque, propeller gyroscopic precession, asymmetric thrust and
spiral slipstream all give the possibility of unco-ordinated flight. Unco-ordinated flight exists
when the aircraft is sideslipping. Indication of sideslip is given to the pilot by the inclinometer
portion (ball) of the tum co-ordinator, Fig. 12.17. The miniature aeroplane indicates rate of
turn.
MINIATURE AEROPLANE
' - - - - LEVEL INDEX
RATE ONE TURN INDEX
'------INCLINOMETER
Figure 12.17
Turn co-ordinator
Co-ordinated flight is maintained by keeping the ball centred between the reference lines with
rudder. To do this, apply rudder pressure on the side where the ball is deflected. The simple
rule, "step on the ball," is a useful way to remember which rudder to apply.
If aileron and rudder are co-ordinated during a tum, the ball will remain centred and there will
be no sideslip. If the aircraft is sideslipping, the ball moves away from the centre of the tube.
Sideslipping towards the centre of the tum, moves the ball to the inside of the tum. Sideslipping
towards the outside of the tum, moves the ball to the outside of the tum. To correct for these
conditions and maintain co-ordinated flight, "step on the ball." Bank angle may also be varied
to help restore co-ordinated flight from a sideslip. The following illustrations give examples.
12 - 18
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
FLIGHT MECHANICS
Fig. 12.18 shows the aircraft in a rate 1 co-ordinated
tum to the right.
Figure 12.18
Fig. 12.19 shows the aircraft in an unco-ordinated tum
to the right; it will be sideslipping towards the centre
of the tum (slipping tum). Using "step on the ball,"
the tum can be co-ordinated by applying right rudder
pressure to centre the ball. The ball can also be
centred by decreasing the bank angle.
Figure 12.19
Fig. 12.20 also shows the aircraft in an unco-ordinated
tum to the right; it will be sideslipping towards the
outside of the tum (skidding tum). Using "step on the
ball," the turn can be co-ordinated by applying left
rudder pressure. The ball can also be centred by
increasing the bank angle.
Figure 12.20
12 - 19
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
12.11
FLIGHT MECHANICS
FLIGHT WITH ASYMMETRIC THRUST
INTRODUCTION
When an engine fails on a multi engine aircraft there will be a decrease in thrust and an increase
in drag on the side with the failed engine:
a)
airspeed will decay
b)
the nose will drop and
c)
most significantly, there will be an immediate yawing moment towards the failed
(dead) engine.
Fig. 12.21 shows the forces and moments acting on an aircraft following failure of the left (port)
engine. The aircraft has a yawing moment towards the dead engine. The pilot has applied
rudder to stop the yaw. The vital action when an engine fails is to STOP THE YAW!
12.11.1
YAWING MOMENT
The yawing moment is the product of thrust from the operating engine, multiplied by the
distance between the thrust line and the CG (thrust arm) plus the drag from the failed engine,
multiplied by the distance between the engine centre line and the CG. The strength of the
yawing moment will depend on:
a)
how much thrust the operating engine is developing (throttle setting and density altitude)
b)
the distance between the thrust line and the CG (thrust arm)
c)
how much drag is being produced by the failed engine.
The rudder moment, which balances the yawing moment, is the product of the rudder force
multiplied by the distance between the fin CP and the CG (rudder arm). This statement will be
modified by factors yet to be introduced. Thus, at this preliminary stage, the ability of the pilot
to counteract the yawing moment due to asymmetric thrust will depend on:
d)
rudder displacement (affecting rudder force)
e)
CG position (affecting rudder arm)
t)
the IAS (affecting rudder force)
12 - 20
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PRINCIPLES OF FLIGHT
FLIGHT MECHANICS
Assume the rudder is at full deflection, CG is at the rear limit (shortest rudder arm) and the IAS
(dynamic pressure) is just sufficient for the rudder force to give a rudder moment equal to the
yawing moment - there will be no yaw. But, any decrease in IAS will cause the aircraft to yaw
uncontrollably towards the failed engine. The uncontrollable yaw to the left, in this example,
will cause the aircraft to roll uncontrollably to the left due to greater lift on the right wing. The
aircraft will enter a spiral dive to the left (impossible to stop with the flight controls alone); if
near the ground, disaster will result. In these extreme circumstances near the ground the ONL Y
way to regain control of the aeroplane is to close the throttle(s) on the operating engine(s). This
removes the yawing moment and the aircraft can be force-landed under control.
Thus there is a minimum lAS at which directional control can be maintained following
engine failure on a multi engine aircraft. This minimum IAS is called V Me (minimum control
speed).
...
YAWING MOMENT
L--------.... __
THRUST
ARM
RUDDER
ARM
RUDDER . ._
FORCE
......
T
.....
..
111
______ - - .....'",
RUDDER MOMENT
Figure 12.21
Asymmetric thrust
12 - 21
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FLIGHT MECHANICS
PRINCIPLES OF FLIGHT
12.11.2
CRITICAL ENGINE
One of the factors influencing the yawing moment following engine failure on a multi engine
aircraft is the length of the thrust arm (distance from the CG to the thrust line of the operating
engine).
In the case of a propeller engine aircraft the length of the thrust arm is determined by the
asymmetric effect of the propeller. At a positive angle of attack the thrust line of a clockwise
rotating propeller, when viewed from the rear, is displaced to the right of the engine centre line.
This is because the down going blade generates more thrust than the up going blade (Chapter
16). If both engines rotate clockwise, the starboard (right) engine will have a longer thrust arm
than the port (left) engine.
If the left engine fails the thrust of the right engine acts through a longer thrust arm and will give
a bigger yawing moment; a higher lAS (VMd would be necessary to maintain directional control.
So at a given lAS the situation would be more critical if the left engine failed, Fig. 12.22.
The critical engine is the engine, the failure of which would give the biggest yawing
moment.
To overcome the disadvantage of having a critical engine on smaller twins, their engines may
be designed to counter-rotate. This means that the left engine rotates clockwise and the right
engine rotates anti-clockwise, giving both engines the smallest possible thrust arm. Larger turbo
- props (e.g. King Air etc. and larger) rotate in the same direction. In the case of a four engined
jet aircraft the critical engine is either of the outboard engines.
NB:
If all the propellers on a multi engine aircraft rotate in the same direction, they are sometimes
called 'co-rotating' propellers.
CRITICAL
ENGINE
GREATER
THRUST
ARM
Figure 12.22
Critical engine
12 - 22
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PRINCIPLES OF FLIGHT
12.11.3
FLIGHT MECHANICS
BALANCING THE YAWING MOMENTS AND FORCES
Although the moments are balanced in Fig. 12.21 the forces are not balanced. The unbalanced
side force from the rudder can be balanced in two ways:
a)
with the wings level and
b)
by banking slightly towards the live engine (preferred method).
Rudder to Stop Yaw - Wings Level: Rudder is used to prevent yaw and the wings are
maintained level with aileron. Yawing towards the live engine gives a sideslip force on the keel
surfaces behind the CG opposite to the rudder force, Fig. 12.23. If the sideslip angle is too
large the fin could stall. The tum indicator will be central and so will the slip indicator.
NB:
Asymmetric thrust is the exception to the rule of co-ordinated flight being indicated to the pilot
by the ball centred in the inclinomet~r.
This method of balancing the side force from the rudder gives reduced climb performance
because of the excessive parasite drag generated, so is not the recommended method for critical
situations, such as engine failure just after take-off or go-around. This technique can be used for
initial control following an engine failure in the cruise.
-
YAWING MOMENT
........ ------- .........
......
I
l
SIDE FORCE
FROM RUDDER
- ------...
~
..... ..,.
RUDDER MOMENT
Figure 12.23
SIDE FORCE
FROM SIDESLIP
......
Wings level method
12 - 23
The advantage of the 'wings
level' method of balancing
the forces is the strong visual
horizontal references
available to the pilot, both
inside and outside the
aircraft.
The disadvantages are that if
the sideslip angle is too large
the fin could stall and the
ability to climb is reduced
due to excessive parasite
drag.
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
FLIGHT MECHANICS
Lift
SIDEWAYS
COMPONENT
OF LIFT
Weight
Figure 12.24 Max. 5 0 bank towards live engine
Rudder to Stop Yaw - Bank Towards Live Engine: It is more aerodynamically efficient to
balance the rudder sideforce by banking towards the live engine, Fig. 12.24, so that lift gives a
sideways component opposite to the rudder force. The angle of bank must not exceed 5°, to
prevent excessive loss of vertical lift component.
Banking towards the live engine also reduces the side force on the fin from sideslip, which
effectively reduces the yawing moment and gives more rudder authority to stop the yaw.
The cockpit indication will be the tum needle central with the slip indicator (ball) one half
diameter displaced towards the live engine. The 'ball' is not centred, but the aircraft is not
sideslipping. This method produces minimum drag and gives the best ability to climb.
12.11.4
ROLL AND YAW MOMENTS WITH ASYMMETRIC THRUST
The rolling and yawing moments and the power of the flight controls to balance them will
determine the controllability of an aircraft with aSymmetric thrust. Rolling and yawing moments
with asymmetric thrust are affected by:
a)
Thrust on the live engine: The greater the thrust, the greater the yawing moment from
the live engine. The further the engine is mounted out on the wing (increased thrust
arm), the larger the yawing moment. Thrust is greatest at low speed and full throttle.
b)
Altitude: Thrust reduces with increasing altitude and / or increasing temperature (high
density altitude). The worst case for engine failure is low density altitude, i.e.
immediately after take-off on a cold day at a sea level airport.
12 - 24
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FLIGHT MECHANICS
PRINCIPLES OF FLIGHT
?
MILLING PROPELLER
DRAG
STATIONARY PROPELLER
FINE PITCH
STOP
/
~
1
FEATHERED
------_______
_ __________
POSITION
~-----~I----~I----~I----~I------ ----~I--~.~
o
15
30
45
60
90
PROPELLER BLADE ANGLE
Figure 12.25
c)
Propeller drag
Drag from the dead engine and propeller: Drag from the dead engine always adds
to the yawing moment. The size of the contribution depends upon whether the propeller
is windmilling, stopped or feathered, Fig.12.25. This effect will be absent on an aircraft
powered by jet engines.
i)
Drag from a windmilling propeller is high. It is being driven by the relative
airflow, and is generating both drag and torque.
ii)
If a propeller is stationary it is generating drag, but no torque. Drag from a
stationary propeller is less than from one which is windmilling.
iii)
A feathered propeller generates the least drag. There is no torque because it is
not rotating and the parasite drag is a minimum because the blades are edge on
to the relative airflow.
The drag on the dead engine can also be reduced by closing the cowl flap.
NB:
d)
Asymmetric blade effect (also known as 'P Factor): Described in Para. 12.26. If
both engines rotate clockwise, the right engine has a longer thrust arm. Failure of the
left engine will give a larger yawing moment. This effect will be absent on an aircraft
with counter rotating propellers, contra-rotating propellers or jet engines.
e)
CG position: The aircraft rotates about the CG. The fore and aft CG location has no
effect on the yawing moment from a failed engine, but will influence the rudder arm,
hence the rudder moment. CG on the aft limit will give the smallest rudder arm and the
least ability to oppose the yawing moment from a failed engine.
Contra-rotating propellers are mounted on the same shaft and are driven in opposite directions,
usually by the same engine
12 - 25
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FLIGHT MECHANICS
PRINCIPLES OF FLIGHT
f)
Torque reaction: When the engine turns the propeller, the equal and opposite reaction
tries to tum the engine in the other direction. Following failure of one engine on an
aircraft with propellers which rotate in the same direction (usually clockwise when
viewed from the rear), the torque tries to roll the aircraft to the left. Failure of the left
engine therefore gives the biggest rolling moment to the left. With counter rotating
engines, both the asymmetric blade effect (P Factor) and the torque reaction are
minimised and there is no longer a critical engine. This effect will be absent on an
aircraft powered by jet engines.
g)
Difference in lift due to slipstream: Engine failure on one side will give a loss of
induced lift from the propeller slipstream on that side. Total lift will reduce giving a
tendency to descend, but more importantly, there will be a rolling moment towards the
dead engine; a greater rolling moment towards the dead engine will occur if the trailing
edge flaps are deployed because of the higher initial CL • This effect will be absent on an
aircraft powered by jet engines.
h)
Rolling moment due to sideslip: If the aircraft is flying with yaw to balance the rudder
force, there will be a sideslip. In Fig. 12.23 the aircraft is sideslipping to the left. The
dihedral of the left wing ( with the dead engine) will cause the lift of the left wing to
increase, which will compensate some of the lift loss due to the loss of the propeller
slipstream.
i)
Weight: Any weight increase will require a higher angle of attack at a given speed.
j)
(i)
This will increase the asymmetric blade effect (P Factor) and give a bigger
yawing moment.
(ii)
The fin and rudder will be masked to a greater extent by disturbed airflow from
the wing and fuselage, making the rudder and fin less effective; consequently
the available rudder moment will be reduced.
Airspeed: The effectiveness of the flying controls depends upon dynamic pressure,
assuming full control displacement. An accurate measure of dynamic pressure at low
airspeeds is given by the Calibrated Air Speed (CAS). CAS is lAS corrected for
position error. At low airspeed / high CL the pressures sensed by the pitot / static system
are affected by the high angle of attack, so must be compensated to make the lAS reflect
a more accurate measure of dynamic pressure. A higher lAS means more control
effectiveness and consequently a larger available rudder moment to balance the yawing
moment from the failed engine. A lower lAS will reduce the available rudder moment,
if the other parameters remain the same. lAS is the vital element in control of the
aircraft with asymmetric thrust.
12 - 26
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PRINCIPLES OF FLIGHT
12.11.5
FLIGHT MECHANICS
MINIMUM CONTROL AIRSPEED
It has been shown that when a multi engine aircraft suffers an engine failure several variables
affect both the yawing moment and the rudder moment which is used to oppose it. It has also
been shown that there is a minimum lAS (VMe), below which it is impossible for the pilot to
maintain directional control with asymmetric thrust.
Airworthiness Authorities, in this case the JAA, have laid down conditions which must be
satisfied when establishing the minimum airspeeds for inclusion in the flight Manual of a new
aircraft type. As in most other cases, the conditions under which the minimum control airspeeds
are established are 'worst case'. A factor of safety is built into these speeds to allow for aircraft
age and average pilot response time.
Because there are distinct variations in the handling qualities of the aircraft when in certain
configurations, minimum control airspeed (VMe) has three separate specifications:
a)
V MeA
Minimum control speed - airborne
b)
V MeG
Minimum control speed - on the ground
c)
V MeL
Minimum control speed - in the landing configuration
12.11.6
VMCA (JAR 25.149 paraphrased)
V MeA is the calibrated airspeed, at which, when the critical engine is suddenly made
inoperative, it is possible to maintain control of the aeroplane with that engine still inoperative,
and maintain straight flight with an angle of bank of not more than 5°.
V MeA may not exceed 1·13 V SR with:a)
b)
c)
d)
e)
f)
g)
maximum available take-off power or thrust on the engines;
the aeroplane trimmed for take-off;
the most unfavourable CG position;
maximum sea level take-off weight;
the aeroplane in its most critical take-off configuration (but with gear up); and
the aeroplane airborne and the ground effect negligible; and
If applicable, the propeller of the inoperative engine;
(i)
windmilling; or
(ii)
feathered, if the aeroplane has an automatic feathering device.
The rudder forces required to maintain control at V MeA may not exceed 150 lb nor may it be
necessary to reduce power or thrust on the operative engines.
NB:
There is no performance requirement, just directional control.
12 - 27
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FLIGHT MECHANICS
PRINCIPLES OF FLIGHT
12.11.7
FACTORS AFFECTING VMCA
ANGLE OF BANK: Banking towards the live engine reduces the rudder deflection required
and so allows a lower V MCA' 50 maximum is stipulated because larger bank angles would
significantly reduce the vertical component oflift; the angle of attack would have to be increased
with the added penalty of higher induced drag.
CG POSITION: Because the aircraft rotates around the CO, the position of the CO directly
affects the length of the rudder arm and thus the power of the rudder and fin to maintain
directional stability and control. The 'worst case' is with the CO at the aft limit. If the
requirements can be met in this configuration, the ability to maintain directional control will be
enhanced at any other CO location.
AILERON EFFECTIVENESS: At low airspeed, dynamic pressure is low which reduces the
effectiveness of all the flying controls for a given angle of displacement. This affect on the
rudder has already been discussed, but the ailerons will be affected in a similar way. In Fig.
12.23 and 12.24 (right roll input) the wings are maintained either level or at the required bank
angle with the ailerons. At reduced airspeed, greater right roll aileron displacement must be used
to keep the wings in the required position. The 'down' aileron on the left side will add to the
yawing moment because of its increased induced drag. At low IAS (increased C L ) the large
angle of down aileron could stall that wing and give an uncontrollable roll towards the dead
engine. V MCA must be high enough to prevent this unwelcome possibility.
FLAP POSITION: Flap position affects lift / drag ratio, nose down pitching moment and the
stalling speed. With asymmetric thrust, flaps reduces climb performance, increases the margin
above stall, but does not directly affect V MCA. However, if take-off flap is used, the difference
in lift between the two wings due to propeller slipstream is further increased. This increases the
rolling moment, requires increased aileron deflection and indirectly increases V MCA.
UNDERCARRIAGE: The undercarriage increases drag and reduces performance. The
increased keel surface in front of the CO decreases directional stability slightly, thus the fin and
rudder are opposed in sideslip conditions and this will slightly increase V MCA.
ALTITUDE AND TEMPERATURE: YMCA is affected by the amount of thrust being
developed by the operating engine. As altitude and/ or temperature increases, the thrust from
an un supercharged engine will decrease. Therefore, V MCA decreases with an increase in altitude
and/or temperature.
RELATIONSHIP BETWEEN Vs AND VMCA : Vs is constant with increasing altitude, so can
be represented by a straight line in Fig. 12.26. (It was shown in Chapter 7 that stall speed does
increase at higher altitudes, but for this study, we are only dealing with lower altitudes).
Fig. 12.26 shows that at about 3000 ft, V s and V MCA typically correspond. So above this altitude,
the stall speed is higher than V MCA' If the aircraft is slowed following an engine failure with full
power on the operating engine the aircraft can stall before reaching V MCA. The margin above
loss of control is reduced; in this case by stalling.
12 - 28
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PRINCIPLES OF FLIGHT
FLIGHT MECHANICS
I
I
/
Vs AND V MCA COINCIDE
Figure 12.26
12.11.8
Vs and V MCA
VMCG (JAR25.149 paraphrased)
V MCG, the minimum control speed on the ground, is the calibrated airspeed during the take-off
run, at which, when the critical engine is suddenly made inoperative, it is possible to maintain
control of the aeroplane with the use of primary aerodynamic controls alone (without the use
of nose-wheel steering) to enable the take-off to be safely continued using normal piloting skill.
The rudder control forces may not exceed 150 pounds (68.1 kg) and, until the aeroplane becomes
airborne, the lateral control may only be used to the extent of keeping the wings level. In the
determination of V MCG, assuming that the path of the aeroplane accelerating with all engines
operating is along the centre of the runway, its path from the point at which the critical engine
is made inoperative to the point at which recovery to a direction parallel to the centreline is
completed may not deviate more than 30ft (9.144 m) laterally from the centre-line at any point.
As for V MCA, this must be established with:
12.11.9
a)
maximum available take-off power or thrust on the engines;
b)
the aeroplane trimmed for take-off;
c)
the most unfavourable CG position;
d)
maximum sea level take-off weight.
FACTORS AFFECTING VMCG
ALTITUDE AND TEMPERATURE: V MCG is affected by the amount of thrust being
developed by the operating engine. As altitude and/or temperature increases, the thrust from an
unsupercharged engine will decrease. Therefore, V MCG decreases with an increase in altitude
and/or temperature.
12 - 29
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PRINCIPLES OF FLIGHT
FLIGHT MECHANICS
NOSE-WHEEL STEERING: Nosewheel steering is designed for taxiing - making large and
sharp turns at low speed, turning off the runway and parking. When taking-off on wet, icy or
slippery runways, the nosewheel begins to hydroplane between 70 and 90 knots (depending on
tyre pressure and depth of water or slush) and has very little steering effect. Once the aircraft
is moving, the nosewheel doesn't do much except tum sideways and skid.
V MCG is established during flight testing, usually on a dry runway. If nosewheel steering were
used by the test pilot it would give a false, low speed at which it was possible to maintain
directional control on the ground after the critical engine is suddenly made inoperative. At this
speed on a slippery runway - even if nosewheel steering were used by a line pilot, it would not
give the required assistance in maintaining directional control following an engine failure and
the aircraft would depart the side of the runway. The regulations ensure that limits are
established in a "worst case" set of circumstances, in order to give the maximum safety factor
during normal operations.
RUDDER ARM: When the aircraft is on the ground it rotates about the main undercarriage,
which is aft of the CG. Therefore the rudder arm is shorter when the aircraft is on the ground.
It will be found that on most aircraft V MCG is higher than V MCA12.11.10
V MeL (JAR 25.149 paraphrased)
V MCU the minimum control speed during approach and landing with all engines operating, is the
calibrated airspeed at which, when the critical engine is suddenly made inoperative, it is possible
to maintain control ofthe aeroplane with that engine still inoperative, and maintain straight flight
with an angle of bank of not more than 50.
V MCL must be established with:a)
the aeroplane in the most critical configuration for approach and landing with all
engines operating;
b)
the most unfavourable CG;
c)
the aeroplane trimmed for approach with all engines operating;
d)
the most unfavourable weight;
e)
for propeller aeroplanes, the propeller of the inoperative engine in the position it
achieves without pilot action; and
f)
go-around power or thrust setting on the operating engines( s).
12 - 30
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FLIGHT MECHANICS
PRINCIPLES OF FLIGHT
In demonstrating V MCL:a)
the rudder force may not exceed 150 Ib;
b)
the aeroplane may not exhibit hazardous flight characteristics or require exceptional
piloting skill, alertness or strength;
c)
lateral control must be sufficient to roll the aeroplane, from an initial condition of
steady flight, through an angle of 20 in the direction necessary to initiate a turn
away from the inoperative engine(s), in not more than 5 seconds.
0
12.11.11
FACTORS AFFECTING VMCL
AILERON EFFECTIVENESS: At low airspeed, dynamic pressure is low which reduces the
effectiveness of all the flying controls for a given angle of displacement. This affect on the
rudder has already been discussed, but the ailerons will be affected in a similar way. At reduced
airspeed, greater aileron displacement must be used to obtain the required roll response. The
'down' aileron on the left side will also add to the yawing moment because of its increased
induced drag and may stall the wing at low lAS (high Cd. Adequate aileron effectiveness is
clearly very important when considering V MCL. because this minimum control speed contains a
roll requirement, not just directional control.
SUMMARY OF MINIMUM CONTROL SPEEDS
JAR 25.149 sets out the criteria to be used when establishing the minimum control speeds for
certification of a new aircraft. The speeds so established will be included in the aircraft's Flight
Manual.
From careful study of the above extracts, several things can be noted:a)
nose wheel steering may not be used when establishing V MCG. Its use would artificially
decrease V MCG. In service, when operating from a slippery runway, nose wheel steering
would be ineffective so it might be impossible to directionally control the aircraft when
at or above the stated V MCG.
b)
V MCL includes a roll requirement, not merely directional control, as with the other
speeds.
c)
The thrust developed by an engine depends on the air density, and so thrust will
decrease with increasing altitude and temperature. The yawing moment due to
asymmetric thrust will therefore decrease with altitude and temperature, and so control
can be maintained at a lower lAS. V MC therefore decreases with increasing altitude and
temperature (higher density altitude).
12 - 31
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FLIGHT MECHANICS
PRINCIPLES OF FLIGHT
12.11.12
PERFORMANCE WITH ONE ENGINE INOPERATIVE
It was shown in paragraph 12.5 that an aircraft's ability to climb depends upon the excess thrust
available, after aerodynamic drag is balanced. If a twin engined aircraft loses an engine, total
thrust is reduced by 50%, but the excess thrust (the thrust, minus aerodynamic drag) is reduced
by more than 50%, Fig. 12.27. The ability to climb may be reduced as much as 80%.
12.11.13
SINGLE ENGINE ANGLE OF CLIMB
Angle of climb is determined by excess thrust available. Climb angle will be a maximum
when the aircraft is flown at the lAS where excess thrust is a maximum (maximum thrust to drag
ratio). Since thrust decreases with forward speed and total drag increases below and above the
minimum drag speed (V1MD), the best angle of climb is achieved at a speed below V IMD but a safe
margin above the stall speed. The airspeed for maximum angle of climb is V x for all engines
operating and V XSE for best single engine angle of climb.
12.11.14
SINGLE ENGINE RATE OF CLIMB
Rate of climb is determined by excess power available. Power is the rate of doing work and
work is force times distance moved, so power is force times distance moved in a given time, i.e.
thrust or drag times TAS (thrust or drag because they are both forces and TAS because it is the
only speed there is!). Although thrust reduces with forward speed, total power available
increases to a point because of the speed factor. Similarly, power required is a measure of drag
times TAS, so excess power available determines the available rate of climb. The airspeed for
best rate of climb is V Y for all engines operating and V YSE for best single engine rate of climb.
Vy and VYSE are higher than Vx and VXSE and provide a safer margin above both stall and
V MeA- Under most circumstances V Y and V YSE are the best speeds to use. On small twin engine
aircraft V YSE is marked on the Air Speed Indicator by a blue radial line and is called 'blue line
speed'.
12.11.15
CONCLUSIONS
At a given altitude, airspeed and throttle position, excess thrust depends on the amount of drag
being generated, and this will depend on configuration, weight and whether turns are required
to be made. The control surface deflections required to balance asymmetric thrust will also
cause an increase in drag. It is essential therefore that after losing an engine, particularly during
take-off or during a go-around, drag is reduced and no turns are made until well away from
the ground.
Drag can be reduced by feathering the propeller of the inoperative engine, raising the
undercarriage, carefully raising the flaps, closing the cowl flap on the inoperative engine and
banking the aircraft no more than 50 towards the operating engine. Flying at V YSE (blue line
speed) with maximum continuous thrust on the operating engine will provide maximum climb
performance and optimum control over the aeroplane.
12 - 32
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FLIGHT MECHANICS
PRINCIPLES OF FLIGHT
THRUST
/
BOTH ENGINES
- - ONE ENGINE
POWER
AVAI LABLE
POWER
POWER
AVA I LABLE
65
VMCA ~
Figure 12.27
70
85
lAS
Excess thrust and excess power
12 - 33
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
FLIGHT MECHANICS
SELF ASSESSMENT QUESTIONS
1.
In straight and level powered flight the following principal forces act on an aircraft:
a)
b)
c)
d)
2.
For an aircraft in level flight, if the wing CP is aft of the CG and there is no thrust/drag couple,
the tailplane load must be:
a)
b)
c)
d)
3.
weight acts vertically toward the centre of the Earth
lift acts perpendicular to the chord line and must be greater than weight
thrust acts forward parallel to the relative wind and is greater than drag
lift acts in the opposite direction to the aircraft weight
The horizontal stabilizer usually provides a download in level flight because:
a)
b)
c)
d)
5.
upward
downward
zero
forward
When considering the forces acting upon an aeroplane in straight-and-Ievel flight at constant
airspeed, which statement is correct?
a)
b)
c)
d)
4.
thrust, lift, weight.
thrust, lift, drag, weight.
thrust, lift, drag.
lift, drag, weight.
the main plane lift is always positive
the lift/weight and thrust/drag couples combine to give a nose down pitch
the lift produced is greater than required at high speed
this configuration gives less interference
The reason a light general aviation aircraft tends to nose down during power reduction is that
the:
a)
b)
c)
d)
thrust line acts horizontally and above the force of drag
centre of gravity is located forward of the centre of pressure
centre of pressure is located forward of the centre of gravity
force of drag acts horizontally and above the thrust line
12 - 35
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FLIGHT MECHANICS
PRINCIPLES OF FLIGHT
6.
To give the best obstacle clearance on take off, take off should be made with:
a)
b)
c)
d)
7.
equal to the aerodynamic drag.
greater than the aerodynamic drag.
less than the aerodynamic drag.
equal to the weight component along the flight path.
wind speed
the aircraft weight
excess engine power
excess airspeed
Assume that after take-off a tum is made to a downwind heading. In regard to the ground, the
aeroplane will climb at:
a)
b)
c)
d)
11.
amount by which the lift exceeds the weight.
amount by which the thrust exceeds the drag.
amount by which the thrust exceeds the weight.
angle of attack of the wing.
A constant rate of climb in an aeroplane is determined by:
a)
b)
c)
d)
10.
the
the
the
the
In a climb at a steady speed, the thrust is:
a)
b)
c)
d)
9.
partially extended and at best rate of climb speed (Vy).
partially extended and at best angle of climb speed (Vx).
retracted and at best rate of climb speed (Vy).
retracted and at best angle of climb speed (Vx).
The angle of climb is proportional to:
a)
b)
c)
d)
8.
flaps
flaps
flaps
flaps
a greater rate into the wind than downwind
a steeper angle downwind than into the wind
the same angle upwind or downwind
a steeper angle into the wind than downwind
What effect does high density altitude have on aircraft performance?
a)
b)
c)
It increases takeoff performance
It increases engine performance
It reduces climb performance
12 - 36
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FLIGHT MECHANICS
PRINCIPLES OF FLIGHT
12.
During a steady climb the lift force is:
a)
b)
c)
d)
13.
In a steady climb the wing lift is:
a)
b)
c)
d)
14.
1 ft in 10ft
1 ft in 20 ft
1 ft in 40 ft
1 ft in 200 ft
To cover the greatest distance when gliding the gliding speed must be:
a)
b)
c)
d)
17.
lift, weight, thrust.
lift, drag, weight.
drag, thrust, weight.
lift and weight only.
For a glider having a maximum LID ratio of20 : 1, the flattest glide angle that could be achieved
in still air would be:
a)
b)
c)
d)
16.
equal to the weight
greater than the weight
equal to the weight component perpendicular to the flight path
equal to the vertical component of weight
During a glide the following forces act on an aircraft:
a)
b)
c)
d)
15.
less than the weight.
exactly equal to the weight.
equal to the weight plus the drag.
greater than the weight.
near to the stalling speed.
as high as possible within VNE limits.
about 30% faster than Vmd.
the one that gives the highest LID ratio.
If the weight of an aircraft is increased the maximum gliding range:
a)
b)
c)
d)
decreases.
increases.
remains the same, and rate of descent is unchanged.
remains the same, but rate of descent increases.
12 - 37
© Oxford Aviation Services Limited
FLIGHT MECHANICS
PRINCIPLES OF FLIGHT
18.
When gliding into a headwind, the ground distance covered will be:
a)
b)
c)
d)
19.
During a 'power-on' glide the forces acting on an aircraft are:
a)
b)
c)
d)
20.
32:1
16:1
8:1
4:1
During a tum the lift force may be resolved into two forces, these are:
a)
b)
c)
d)
23.
increase and glide angle will be steeper.
increase, but glide angle will remain the same.
decrease.
remain the same.
An aircraft has a LID ratio of 16: 1 at 50 kt in calm air. What would the approximate GLIDE
RATIO be with a direct headwind of25 kt?
a)
b)
c)
d)
22.
lift, drag and weight.
lift, thrust and weight.
lift, drag, thrust and weight.
lift and weight only.
If airbrakes are extended during a glide, and speed maintained, the rate of descent will:
a)
b)
c)
d)
21.
less than in still air.
the same as in still air but the glide angle will be steeper.
the same as in still air but the glide angle will be flatter.
greater than in still air.
a force opposite to thrust and a force equal and opposite to weight.
centripetal force and a force equal and opposite drag.
centripetal force and a force equal and opposite weight.
centrifugal force and a force equal and opposite thrust.
In a tum at a constant lAS, compared to straight and level flight at the same lAS:
a)
b)
c)
d)
the same power is required because the lAS is the same.
more power is required because the drag is greater.
more power is required because some thrust is required to give the centripetal force.
less power is required because the lift required is less.
12 - 38
© Oxford Aviation Services Limited
FLIGHT MECHANICS
PRINCIPLES OF FLIGHT
24.
In a tum at a given TAS and bank angle:
a)
b)
c)
d)
25.
As bank angle is increased in a tum at a constant lAS, the load factor will:
a)
b)
c)
d)
26.
increase.
decrease but bank angle will increase.
decrease but bank angle will decrease.
remain the same.
An aircraft has a stalling speed in level flight of 70 kt lAS. In a 60° balanced turn the stalling
speed would be:
a)
b)
c)
d)
29.
insufficient rate of yaw.
too much bank.
too much nose up pitch.
insufficient bank
For a tum at a constant lAS if the radius of tum is decreased the load factor will:
a)
b)
c)
d)
28.
increase in direct proportion to bank angle.
increase at an increasing rate.
decrease.
remain the same.
Skidding outward in a tum is caused by:
a)
b)
c)
d)
27.
only one radius of tum is possible.
the radius can be varied by varying the pitch.
the radius can be varied by varying the yaw.
two different radii are possible, one to the right and one to the left.
76 kt.
84 kt.
99 kt.
140 kt..
An increase in airspeed while maintaining a constant load factor during a level, coordinated tum
would result in:
a)
b)
c)
d)
an increase in centrifugal force
the same radius of tum
a decrease in the radius of tum
an increase in the radius of tum
12 - 39
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
30.
How can the pilot increase the rate of tum and decrease the radius at the same time?
a)
b)
c)
31.
Compensate for increase in induced drag.
Increase the horizontal component of lift equal to the vertical component.
Compensate for loss of vertical component of lift.
To stop the nose from dropping below the horizon and the airspeed increasing.
the heavier aircraft would have a higher" g" load
the lighter aircraft would have a higher "g" load
they would both have the same" g" load
For a multi-engined aircraft, V MeG is defined as the minimum control speed on the ground with
one engine inoperative. The aircraft must be able to:
a)
b)
c)
d)
35.
2 G'S
6 G'S
3 G's
Two aircraft of different weight are in a steady tum at the same bank angle:
a)
b)
c)
34.
9 G'S
Why must the angle of attack be increased during a tum to maintain altitude?
a)
b)
c)
d)
33.
shallow the bank and increase airspeed
steepen the bank and increase airspeed
steepen the bank and decrease airspeed
If an aircraft with a gross weight of 2,000 kg were subj ected to a total load of 6,000 kg in flight,
the load factor would be:
a)
b)
c)
d)
32.
FLIGHT MECHANICS
abandon the take off.
continue the take off or abandon it.
continue the take off using primary controls only.
continue the take off using primary controls and nosewheel steering.
What criteria determines which engine is the "critical" engine of a twin-engine aeroplane?
a)
b)
c)
d)
the
the
the
the
one with the centre of thrust farthest from the centerline of the fuselage
one with the centre of thrust closest to the centerline of the fuselage
one designated by the manufacturer which develops most usable thrust
failure of which causes the least yawing moment
12 - 40
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
36.
FLIGHT MECHANICS
Following failure ofthe critical engine, what performance should the pilot ofa light, twin-engine
aeroplane be able to maintain at V MeA?
a)
b)
c)
Heading, altitude, and ability to climb 50 ftlmin
Heading only
Heading and altitude
12 - 41
© Oxford Aviation Services Limited
FLIGHT MECHANICS
PRINCIPLES OF FLIGHT
I No I A I B I c I DII
REF
III
No
1
B
22
2
B
23
3
24
A
4
B
25
5
B
26
6
D
27
7
B
28
8
B
29
9
D
11
12
C
A
13
C
A
B
D
A
C
D
C
D
32
C
33
C
34
C
14
B
35
B
15
B
36
B
D
37
17
D
38
18
39
A
19
20
21
C
40
41
A
B
42
12 - 43
I
B
31
16
REF
C
30
C
10
I A I B I c I D II
V/ / V
V/ / /
------------------------V VV ~~
V VV /
V V V / ------------V V V V -------------------------
© Oxford Aviation Services Limited
CHAPTER 13 - HIGH SPEED FLIGHT
Contents
Page
INTRODUCTION ...................................................... 13 - 1
SPEED OF SOUND
MACH NUMBER ...................................................... 13 - 2
EFFECT ON MACH NUMBER OF CLIMBING AT A CONSTANT IAS
VARIATION OF TAS WITH ALTITUDE AT A CONSTANT MACH NUMBER ... 13 - 4
TEMPERATURE ON MACH NUMBER AT A CONSTANT FL & IAS
SUBDIVISION OF AERODYNAMIC FLOW ................................ 13 - 5
PROPAGATIONOFPRESSUREWAVES .................................. 13-6
NORMAL SHOCK WAVES .............................................. 13 - 8
CRITICAL MACH NUMBER
PRESSURE DISTRIBUTION AT TRANSONIC MACH NUMBERS ............. 13 - 10
PROPERTIES OF A NORMAL SHOCK WAVE ............................. 13 - 12
OBLIQUE SHOCK WAVES ............................................. 13 - 13
EFFECTS OF SHOCKWAVE FORMATION ................................ 13 - 14
LIFT
LIFT CURVE SLOPE AND C LMAX .................................. 13 - 16
DRAG ........................................................ 13 -17
CL/CoDRAGPOLARCURVE .................................... 13 -18
CENTRE OF PRESSURE
CP MOVEMENT ............................................... 13 - 20
FL YING CONTROLS
CONTROL BUZZ ............................................... 13 - 21
BUFFET
FACTORS WHICH AFFECT THE BUFFET BOUNDARIES ................... 13 - 22
STALL SPEED
LOAD FACTOR ................................................ 13 - 23
MACH NUMBER ............................................... 13 - 24
ANGLE OF ATTACK
PRESSURE ALTITUDE .......................................... 13 - 25
WEIGHT
CG POSITION
THE BUFFET MARGIN ................................................
USE OF THE BUFFET ONSET CHART
1·3 G ALTITUDE
BUFFET RESTRICTED SPEED LIMITS
AERODYNAMIC CEILING
LOAD FACTOR AND BANK ANGLE AT WHICH BUFFET OCCURS
DELAYING THE EFFECTS OF COMPRESSIBILITY ........................
THIN WING SECTIONS
SWEEPBACK
DISADVANTAGES OF SWEEP .............................
VORTEX GENERATORS ........................................
AREA RULE ...................................................
MACH TRIM ...................................................
SUPERCRITICAL AEROFOIL ....................................
ADVANTAGES ..........................................
DISADVANTAGES
AERODYNAMIC HEATING ............................................
MACH ANGLE .......................................................
MACH CONE .........................................................
AREA (ZONE) OF INFLUENCE
BOW WAVE
EXPANSION WAVES ..................................................
SONIC BANG ........................................................
METHODS OF IMPROVING CONTROL AT TRANSONIC SPEEDS
SUPERSONIC WAVE CHARACTERISTICS SUMMARY ..............
SWEEPBACK - FACT SHEET ....................................
SELF ASSESSMENT QUESTIONS .......................................
ANSWERS ....................................................
13 - 26
13 - 28
13
13
13
13
13
13
- 30
- 31
- 32
- 33
- 34
- 35
13 - 36
13 - 37
13 - 38
13 - 39
13 - 41
13
13
13
13
- 42
- 43
- 45
- 51
PRINCIPLES OF FLIGHT
13.1
HIGH SPEED FLIGHT
INTRODUCTION
During the preceding study oflow speed aerodynamics it was assumed that air is incompressible,
that is, there is no change in air density resulting from changes of pressure.
At any speed there are changes in air density due to 'compressibility', but if the speed is low the
changes are sufficiently small to be ignored. As speed increases however, the changes in air
density start to become significant.
When an aircraft moves through the air infinitesimally small pressure disturbances, or waves,
are propagated outward from the aircraft in all directions, but only the waves travelling ahead
of the aircraft are significant for the study of high speed flight. These pressure waves 'signal'
the approach of the aircraft and make the air change direction (upwash) and divide to allow
passage of the aircraft.
13.2
SPEED OF SOUND
For the study of high speed flight we are interested in the speed at which the infinitesimally
small pressure disturbances (waves) travel through the atmosphere. Pressure waves 'propagate'
from their source, that is, each air molecule is rapidly vibrated in tum and passes-on the
disturbance to its neighbour. The speed of propagation of small pressure waves depends
upon the temperature of the air ONLY. The lower the temperature, the lower the speed of
propagation. Sound is pressure waves and the speed of any pressure wave through the
atmosphere, whether audible or not, has become known as 'the speed of sound'.
The speed of sound at 15 0 C is 340 metres per second, or approximately 661 kt.
It can be shown that:
Where a
=
a
=
speed of sound
y = a constant (1' 4 for air)
(Eq 13.1)
.JyRT
R = the gas constant
T = absolute temperature
Since y and R are constants, the speed of sound is proportional only to the square root of
the absolute temperature. For example, at 15 D C (288 K):
a
= .J
=
a
oc
1.4 x 287 x 288
(R
=
287 J I kg K)
340 m/s
The speed of sound changes
with Temperature ONLY
r=r
13 - 1
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
13.3
HIGH SPEED FLIGHT
MACH NUMBER
As the speed of an aircraft increases, there is a decrease in the distance between the aircraft
and the influence of the advancing pressure waves. The aircraft begins to catch up the
pressure waves, so the air has less time to move from the aircraft's path and upwash has a more
acute angle.
At higher speeds there is also a change in the flow and pressure patterns around the aircraft.
Ultimately lift and drag, manoeuvrability and the stability and control characteristics will all be
changed.
These effects are due to the compressibility of air, where density can change along a streamline,
and the associated conditions and the characteristics which arise are due to 'compressibility'.
It is vitally important that the flight crew know the speed of the aircraft in relation to the
potential effects of' compressibility'. If the aircraft speed through the air (TAS) and the speed
of sound in the air through which it is flying (the local speed of sound) is known, this will give
an indication of the degree off compressibility. This relationship is known as the Mach number
and Mach number is a measure of compressibility. (e.g. M 0·5 is half the local speed of
sound).
Mach number (M) is the ratio of the true airspeed (V) to the local speed of sound (a)
(Eq 13.2)
Equation 13.2 is a good formula to remember because it allows several important relationships
to be easily understood.
13.4
EFFECT ON MACH NUMBER OF CLIMBING AT A CONSTANT lAS
It is known that temperature decreases with increasing altitude, so the speed of sound
will decrease as altitude is increased.
It is also known that if altitude is increased at a constant lAS, the TAS increases.
Therefore, the Mach number will increase if altitude is increased at a constant lAS.
This is because (V) gets bigger and (a) gets smaller.
From a practical point of view: climbing at a constant lAS makes the distance between the
aircraft and the influence of the advancing pressure waves decrease, which begins to change the
flow and pressure patterns around the aircraft.
The lower the temperature
the lower the speed of sound
13 - 2
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
HIGH SPEED FLIGHT
The International Standard Atmosphere assumes that temperature decreases from 15°C at sea
level to - 56·5°C at 36,089 ft (11,000 m), then remains constant. The speed of sound will
therefore decrease with altitude up to the tropopause, and then remain constant, Fig. 13.1.
1
50
I
i
.:=
0
0
0
..-><
I
40
STRATOSPHERE
-
W
-
-- -
--
-
-
-
0
:J
II-
-
-
I
--l
I
I SA CONDITIONS
w
0:::
:J
-
-
-
-
-
-
~
\
1
30
«
-
20
TROPOSPHERE
1
1\
\
en
en
w
~
0:::
0...
10
".
,
'~ SPEED
\
OF SOUND
11
o
400
500
600
700
SPEED OF SOUND - kt
Figure 13.1
Variation of speed of sound with altitude
Chapter 14 will fully describe V MO and MMo, the high speed (generally speaking) operational
limit speeds. It has been stated that as an aircraft climbs at a constant lAS its Mach number will
be increasing. It is clear that it is possible to exceed the maximum operating Mach number
(MMo) in a climb at a constant lAS.
As the climb continues an altitude will be reached at which the flight crew must stop flying at
a constant lAS and fly at a constant Mach num@er, to avoid accidentally exceeding MMo. The
altitude at which this changeover takes place will depend on the outside air temperature.
The lower the outside air temperature, the lower the changeover altitude.
13 - 3
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
13.5
HIGH SPEED FLIGHT
VARIATION OF TAS WITH ALTITUDE AT A CONSTANT MACH NUMBER
If M
=
TAS
When descending
at a constant Mach number
lAS will be increasing
a
then TAS
=
M x a
It can be seen from the equation that if an aircraft is flown at a constant Mach number:
a)
as altitude decreases the temperature will rise, local speed of sound will increase and
TAS will increase.
b)
as altitude increases the temperature will drop, local speed of sound will decrease and
T AS will decrease (up to the tropopause and then remain constant).
When climbing at a constant TAS
Mach number will be increasing,
up to the tropopause, and
then remain constant
13.6
INFLUENCE OF TEMPERATURE ON MACH NUMBER AT A CONSTANT FLIGHT
LEVEL AND lAS
An aircraft normally operates at Indicated Air Speeds and the Mach number can be expressed
in terms of lAS:
M
lAS
=
constant
for lAS in knots:
M
lAS
=
(Eq 13.3)
/P
V p;;
(Eq 13.4)
661
where,
p
=
pressure at altitude
Po
=
pressure at Sea Level
This shows that at a constant pressure altitude (Flight Level), the Mach number is independent
of temperature for a constant lAS.
This is because the speed of sound and the T AS for a given lAS, both change as
13 - 4
.;-T
© Oxford Aviation Services Limited
HIGH SPEED FLIGHT
PRINCIPLES OF FLIGHT
13.7
SUBDIVISIONS OF AERODYNAMIC FLOW
MO' 4
MO' 75
:d-
LOW
I
HIGH
SUBSONIC
M l' 2
TRANSONIC -
____~--- SUPERSONIC
---t~~
I
I
I
SOME ML < 1'0
ALL ML > 1'0
ALL ML < 1'0
OTHER ML > l' 0
I
I
I~
---
- _ . COMPRESSIBLE FLOW -
-
-
-
I
I
Mcrit
M 1· 0
M FS ---~
(not to scale)
(Aircraft Mach number)
about MO·7 to MO·8
depending on individual aircraft
and angle of attack
Figure 13.2
Classification of airspeed
Fig. 13.2 shows the flow speed ranges with their approximate Mach number values, where :
M Fs = Free Stream Mach number: The Mach number of the flow sufficiently remote from
an aircraft to be unaffected by it. (In effect, the Mach number of the aircraft through the
air). This is the Mach number shown on the aircraft Mach meter.
ML
Local Mach number: When an aircraft flies at a certain M Fs , the flow over it is
accelerated in some places and slowed down in others.
Local Mach number (M L), the boundary layer flow speed relative to the surface of the aircraft,
is subdivided as follows :Subsonic
Less than Mach 1·0 «M 1'0)
Sonic
Exactly Mach 1·0 (M 1-0)
Supersonic
Greater than Mach 1-0 (>M 1-0)
13 - 5
© Oxford Aviation Services Limited
HIGH SPEED FLIGHT
PRINCIPLES OF FLIGHT
13.8
PROPAGATION OF PRESSURE WAVES
KEY
= POSITION
OF OBJECT WHEN
PRESSURE WAVE GENERATED
• = POSITION
(
OF OBJECT WHEN
PRESSURE WAVE REACHES
RADIUS r
I~
===== ~
M
=
( =
(a)
MACH NUMBER OF OBJECT
PRESSURE WAVE EXPANDING
FROM SOURCE AT LOCAL
SPEED OF SOUND
Fig. 13.3 shows a series of sketches
which illustrate the basic idea of
pressure wave formation ahead of an
object moving at various Mach
numbers and of the airflow as it
approached the object. Pressure
waves are propagated continuously,
but for clarity just one is considered.
(
=~
I,~
\\
(b)
If we assume a constant local speed
of sound; as the object's Mach
number increases, the object gets
closer to the 'leading edge' of the
pressure wave and the air receives
less and less warning of the approach
ofthe object.
(c)
The greater the Mach number of the
object, the more acute the upwash
angle and the fewer the number of air
particles that can move out of the path
of the object. Air will begin to build
up in front of the object and the
density of the air will increase.
(
bi~~~
'-
~
~====
\~
(d)
AIRFLOW
PRESSURE WAVE
Figure 13.3
13 - 6
When the object' s speed has reached
the local speed of sound (d), the
pressure wave can no longer warn the
air particles ahead of the object
because the object is travelling
forward at the same speed as the
wave.
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
HIGH SPEED FLIGHT
Therefore, the free stream air particles are not aware of anything until the particles that are piled
up right in front of the object collide with them. As a result of these collisions, the air pressure
and density increase accordingly.
As the object's speed increases to just above M 1·0, the pressure and density of the air just ahead
of it are also increased. The region of compressed air extends some distance ahead of the obj ect,
the actual distance depends on the speed and size of the obj ect and the temperature of the air.
At one point the free air stream particles are completely undisturbed, having received no advance
warning of the approach of a fast moving object, and then are suddenly made to undergo drastic
changes in velocity, pressure, temperature and density. Because of the sudden nature of these
changes, the boundary line between the undisturbed air and the region of compressed air is called
a "shock wave", a stylised sketch of which is shown in Fig. 13.3a.
At supersonic speeds
there is no upwash or
downwash
\.
SHOCK WAVE
(STYLISED)
SUBSONIC
SUPERSONIC
AIRFLOW
APPROXltvIA TELY M 1· 3
AIRFLOW
/
/
/
/
Figure 13. 3a Stylised shock wave
13 - 7
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
13.9
HIGH SPEED FLIGHT
NORMAL SHOCK WAVES
(Nonnal meaning perpendicular to the upstream flow). In addition to the fonnation of a shock
wave described overleaf, a shock wave can be generated in an entirely different manner when
there is no object in the supersonic airflow. (We have now returned to the wind tunnel analogy
of a stationary aircraft and moving air). Whenever supersonic airflow is slowed to subsonic
speed without a change in direction, a "normal" shock wave will form as a boundary
between the supersonic and subsonic region. This means that some 'compressibility effects'
will occur before the aircraft as a whole reaches Mach 1·0.
NORI'v1AL
SHOCK
AIR BEING ACCELERATED
TO SUPERSONIC SPEED
~
LOCAL I'v1ACH NUMBER > 1
LOCAL I'v1ACH NUMBER < 1
(PRESSURE WAVES ABLE
TO TRAVEL FORWARD)
Figure 13.4
13.10
Shock wave at subsonic free stream Mach number
CRITICAL MACH NUMBER
An aerofoil generates lift by accelerating air over the top surface. At small angles of attack the
highest local velocity on an aircraft will usually be located at the point of maximum thickness
on the wing. For example, at a free stream speed ofMO'84, maximum local velocity on the wing
might be as high as M1'05 in cruising level flight. At increased angles of attack the local
velocity will be greater and further forward, also if the thickness/chord ratio were greater the
local speed will be higher.
As the free stream speed increases the maximum speed on the aerofoil will reach the local speed
of sound first. The Free Stream Mach number at which the local velocity first reaches
Mach 1·0 (sonic) is called the Critical Mach number (M CRIT)'
Increased thickness/chord and increased angle
of attack cause greater accelerations over the
top surface of the wing, so the critical Mach
number will decrease with increasing
thickness/chord ratio or angle of attack.
Critical Mach number is
the highest speed at which
no parts of the aircraft
are supersonic
13 - 8
© Oxford Aviation Services Limited
HIGH SPEED FLIGHT
PRINCIPLES OF FLIGHT
13.10.1
ACCELERATING BEYOND M eRIT
At speeds just above the critical Mach number there will be a small region of supersonic airflow
on the upper surface, terminated by a shock wave, Fig. 13 .5.
AREA OF SUPERSONIC FLOW
NORMAL
SHOCK
WAVE
)
SUBSONIC FLOW
~_F_LO_W
________________
Figure 13.5
==:-===~~~
Mixed supersonic and subsonic airflow at transonic speed
As the aircraft speed is further increased the region of supersonic flow on the upper surface
extends and the shockwave marking the end of the supersonic region, moves rearwards. A
similar sequence of events will occur on the lower surface although the shockwave will usually
form at a higher aircraft speed because the lower surface usually has less curvature so the air is
not accelerated so much.
When the aircraft speed reaches Mach 1·0 the airflow is supersonic over the whole of both upper
and lower surfaces, and both the upper and lower shock waves will have reached the trailing
edge. At a speed just above Mach 1·0 the other shockwave previously described and illustrated
in Fig 13.3a, the bow wave, forms ahead of the leading edge.
The bow shock wave is initially separated (detached) from the leading edge by the build up of
compressed air at the leading edge, but as speed increases it moves closer to the leading edge.
For a sharp leading edge the shock eventually becomes attached to the leading edge. The Mach
number at which this occurs depends upon the leading edge angle. For a sharp leading edge with
a small leading edge angle the bow wave will attach at a lower Mach number than one with a
larger leading edge angle.
Fig. 13 .7 overleaf, shows the development of shockwaves on an aerofoil section at a small
constant angle of attack as the airspeed is increased from subsonic to supersonic.
A shockwave forms at the rear ]
of an area of supersonic flow
I
At M CRrT there is no shockwave
)
because there is no supersonic flow
.....
13 - 9
© Oxford Aviation Services Limited
HIGH SPEED FLIGHT
PRINCIPLES OF FLIGHT
13.11
PRESSURE DISTRIBUTION AT TRANSONIC MACH NUMBERS
Refer to Fig 13.7. The solid blue line represents upper surface pressure and the dashed blue line
the lower surface. Decreased pressure is indicated upwards. The difference between the full line
and the dashed line shows the effectiveness oflift production; if the dashed line is above the full
line the lift is negative in that area. Lift is represented by the area between the lines, and the
Centre of Pressure (CP) by the centre of the area.
During acceleration to supersonic flight, the pressure distribution is irregular.
M 0·75 the subsonic picture. Separation has started near the trailing edge and there is practically
no net lift over the rear third of the aero foil section; the CP is well forward. Fig. 13.6 shows
that CL is quite good and is rising steadily; CD' on the other hand, is beginning to rise.
M 0·81 A shock wave has appeared on the top surface; notice the sudden increase of pressure
(shown by the falling line) caused bydecreasing flow speed at the shock wave. The CP has
moved back a little, but the area is still large. Fig. 13.6 shows that lift is good, but drag is now
rising rapidly.
M 0·89 The pressure distribution shows very clearly why there is a sudden drop in lift coefficient
before the aerofoil as a whole reaches the speed of sound; on the rear portion of the aerofoil the
lift is negative because the suction on the top surface has been spoilt by the shock wave, while
there is still quite good suction and high-speed flow on the lower surface. On the front portion
there is nearly as much suction on the lower surface as on the upper. The CP has now moved
well forward again. Fig. 13.6 shows that drag is still increasing rapidly.
M 0·98 This shows the important results of the shock waves moving to the trailing edge, and no
longer spoiling the suction or causing separation. The speed of the flow over the surfaces is
nearly all supersonic, the CP has moved aft again, and owing to the good suction over nearly all
the top surface, with rather less on the bottom, the lift coefficient has actually increased. The
drag coefficient is just about at its maximum, as shown Fig. 13 .6.
M 1·4 The aerofoil is through the transonic region. The bow wave has appeared. The lift
coefficient has fallen again because the pressure on both surfaces are nearly the same; and for
the first time since the critical Mach number, the drag coefficient has fallen considerably.
MO' 81
0·5
1· 0
0'5
Mach number
Figure 13.6
l' 0
1· 5
Mach number
Changes in Lift and Drag in the transonic region
13 - 10
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PRINCIPLES OF FLIGHT
HIGH SPEED FLIGHT
M 0'81
MO'75
/
,
---
!
Figure 13.7
Pressure distribution in the transonic region
13 - 11
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HIGH SPEED FLIGHT
PRINCIPLES OF FLIGHT
13.12
PROPERTIES OF A NORMAL SHOCKWAVE
~
NORfv1AL
SHOCK WAVE
~
SUPERSONIC
SUBSONIC
Figure 13.8
Normal shock wave formation
/
STRONG OBLlQU=E_ _ _-==~~~=::::::=~
SHOCK WAVE
A
NORfv1AL SHOCK WAVE ~=~~=:
WEAK OBLlQUE= _ _~~~~===
SHOCK WAVEfv1ACH LINE --~
When a shock wave is perpendicular
(normal) to the upstream flow,
streamlines pass through the shock
wave with no change of direction. A
supersonic airstream passing through
a normal shock wave will also
experience the following changes:-
\
Figure 13.9 Normal and oblique shock
waves
1.
The airstream is slowed to subsonic; the local Mach number behind the wave is
approximately equal to the reciprocal of the Mach number ahead ofthe wave e.g., if the
Mach number ahead of the wave is 1'25, the Mach number of the flow behind the wave
will be approximately 0·80. (The greater the Mach number above M 1·0 ahead of the
wave, the greater the reduction in velocity).
2.
static pressure increases
3.
temperature increases
4.
density increases
5.
The energy of the airstream [total pressure (dynamic plus static)] is greatly reduced.
Minimum energy loss through a normal shock wave will occur when the Mach number of
the airflow in front of the shock wave is small, but supersonic.
13 - 12
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PRINCIPLES OF FLIGHT
13.13
HIGH SPEED FLIGHT
OBLIQUE SHOCK WAVES
An oblique shock wave is a slightly different type of shock wave.
Referring to Fig. 13.9, at 'A' the air is travelling at supersonic speed, completely unaware of the
approaching object.
The air at 'B' has piled up and is subsonic, trying to slip around the front of the object and merge
with the airflow.
Through the shock wave supersonic air from' A' slows immediately, increasing in pressure and
density as it does so. As previously pointed out, a rise in temperature also occurs. The centre
part of the shock wave, lying perpendicular or normal to the direction of the airstream, is the
strong normal shock wave.
Notice that 'above' and 'below' the normal shock wave, the shock wave is no longer
perpendicular to the upstream flow, but is at an oblique angle; the airstream strikes the oblique
shock wave and is deflected.
Like the normal shock wave, the oblique shock wave in this region is strong. The airflow will
be slowed down; the velocity and Mach number of the airflow behind the wave are reduced, but
the flow is still supersonic. The primary difference is that the airstream passing though the
oblique shock wave changes direction. (The component of airstream velocity normal to the
shockwave will always be subsonic downstream, otherwise no shock wave).
The black dashed lines in Fig. 13.9 outline the area of subsonic flow created behind the strong
shock wave.
Particles passing through the wave at 'C' do not slow to subsonic speed. They decrease
somewhat in speed and emerge at a slower but still supersonic velocity. At 'C' the shock wave
is a weak oblique shock wave. Further out from this point the effects ofthe shock wave decrease
until the air is able to pass the object without being affected. Thus the effects ofthe shock wave
disappears, and the line cannot be properly called a shock wave at all; it is called a 'Mach line'.
SHOCK WAVE SUMMARY
1.
The change from supersonic to subsonic flow is always sudden and accompanied by
rapid and large increases in pressure, temperature and density across the shock wave that
is formed. A normal shock wave marks the change from supersonic to subsonic flow.
2.
If the shockwave is oblique, that is, at an angle to the upstream flow, the airflow is
deflected as it passes through the shock, and may remain supersonic downstream of the
shock wave. However the component of velocity normal to the shockwave will always
be subsonic downstream of the shock.
13 - 13
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HIGH SPEED FLIGHT
PRINCIPLES OF FLIGHT
13.14
EFFECTS OF SHOCKWAVE FORMATION
The formation and development of shockwaves on the wing have effects on lift, drag, stability
and control. Many of these effects are caused by shock induced separation. As the air flows
through the shock wave the sudden rise in pressure causes the boundary layer to thicken and
often to separate. This increases the depth of the turbulent wake behind the wing.
EFFECT OF SHOCK WAVES ON LIFT
At low subsonic speeds the lift coefficient C L is assumed to be constant at a given angle of
attack. With increasing Mach number however, it will vary as shown in Figure 13 .10.
MO'81
MO·98
1
- - + -1
1
- SUBSONIC - - -I
,
- - - - - - +- - - - - -
1
1
M 0'89
1--- SUPERSONIC
••------ TRANSONIC -~I
0·4
1· 2
MCRIT
1
1·4
FREE STREAM MACH NUMBER
Figure 13.10
-
M FS
Variation of CL with Mach number at constant
a
At high subsonic speed C L increases. This is the result of the changing pattern of streamlines.
At low speeds the streamlines begin to diverge well ahead of the aerofoil, Fig. 13 .3 and 13.11 .
At high subsonic speeds they do not begin to deflect until closer to the leading edge, causing
greater acceleration and pressure drop around the leading edge. It will be remembered from
Chapter 7 that this phenomena causes the stall speed to increase at high altitudes. See also
paragraph 13.16.
13 - 14
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HIGH SPEED FLIGHT
PRINCIPLES OF FLIGHT
HIGH SPEED
LOW SPEED
Figure 13.11
Streamlines at low and high subsonic speeds
At speeds above M CRIT a shockwave will have formed on the upper surface. This may cause
boundary layer separation aft ofthe shock wave, causing loss of lift (above M 0·81 , as shown
in Figs. 13.7 and 13.10).
SHOCKWAVE
SEPARATED AIRFLOW
Figure 13.12
Shock stall
This is known as the shock stall because it results from a separated boundary layer just as the
low speed stall does. The severity of the loss oflift depends on the shape ofthe wing sections.
Wings not designed for high speeds may have a severe loss of lift at speeds above M CRIT (Fig.
13 .10), but wings designed specifically for high speed flight, with sweep back, thinner sections
and less camber will have much less variation of lift through the transonic region.
Separated airflow caused by a shock stall can cause severe damage to the airframe, particularly
the empennage. This will be fully discussed in paragraph 13.15.
The lower end of the transonic region is where most modem high speed jet transport aircraft
operate and a small shock wave will exist on the top surface of the wing in the cruise.
13 - 15
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PRINCIPLES OF FLIGHT
HIGH SPEED FLIGHT
EFFECT OF SHOCK W AVES ON LIFT CURVE SLOPE AND C L MAX
At a constant angle of attack, the increase of CL as speed increases from about M 004 into the low
end of the transonic region gives a steeper lift curve slope, i.e. the change ofC L per degree angle
of attack will increase. However, because of earlier separation resulting from the formation of
the shock wave, CLMAX and the stalling angle will be reduced. Figures. 13 .13 and 13.14 illustrate
these changes.
HIGH
SUBSONIC SPEEDS
~
INCOMPRESSIBLE
FLOW
~--'---------~-----------~
ANGLE OF ATTACK
Figure 13.13
Effect of Mach number on Lift curve
Figure 13.14 Effect of Mach number on CLMAX
13 - 16
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
HIGH SPEED FLIGHT
EFFECT OF SHOCK WAVES ON DRAG
As speed increases above M e RIT shock waves begin to form and drag increases more rapidly than
it would have done without the shock waves. The additional drag is called wave drag, and is due
to energy drag and boundary layer separation.
Energy Drag: Energy drag stems from the irreversible nature of the changes which occur as an
airflow crosses a shock wave. Energy has to be used to provide the temperature rise across the
shock wave and this energy loss is drag on the aircraft. The more oblique the shock waves are,
the less energy they absorb, but because they become more extensive laterally and affect more
air, the energy drag rises progressively as M FS increases.
Boundary Layer Separation: In certain stages of shock wave movement there is a considerable
flow separation, as shown in Figs. 13.7 and 13.12. This turbulence represents energy lost to the
flow and contributes to the drag. As M FS increases through the transonic range the shock waves
move to the trailing edge and the separation decreases; hence the drag coefficient decreases .
MO'98
0·4
1· 2
MCRIT
FREE STREAM rv1ACH NUMBER
Figure 13.15
M FS
Variation of CD with Mach number
The change in drag characteristics is shown by the CD curve for a basic section at a constant
angle of attack in Fig. 13.15. The 'hump' in the curve from M 0·89 to M 1·2 is caused by:
a)
The drag directly associated with the trailing edge shock waves (energy loss).
b)
Separation of the boundary layer.
c)
The formation of the bow shock wave above M 1·0.
13 - 17
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PRINCIPLES OF FLIGHT
HIGH SPEED FLIGHT
EFFECT OF SHOCK WAVES ON THE CL / Co DRAG POLAR CURVE
Although the curve of CL / CD is unique at low speeds, at transonic speeds when compressibility
becomes significant, the curve will change. Figure 13.16 shows the variation ofC L / CD with
Mach number. The point at which the tangent from the origin touches the curve corresponds to
the maximum CL / CD or maximum L / D. In the transonic region, the LID ratio is reduced.
L~
D MAX
CD
Figure 13.16
Effect of Mach number on CL/C o polar
EFFECT OF SHOCK WAVES ON THE CENTRE OF PRESSURE
The centre of pressure of an aerofoil is determined by the pressure distribution around it. As the
speed increases through the transonic region, the pressure distribution changes and the centre
of pressure will move. It was shown in Fig 13 .7 that above M CRIT the upper surface pressure
continues to drop on the wing until the shock wave is reached. This means that a greater
proportion of the "suction" pressure will comes from the rear of the wing, and the centre of
pressure is further aft. The rearward movement of the CP however is irregular, as the pressure
distribution on the lower surface also changes. ' The shockwave on the lower surface usually
forms at a higher free stream Mach number than the upper surface shock, but reaches the trailing
edge first. The overall effect on the CP is shown in Figure 13.17. As the aircraft accelerates
to supersonic speed the overall movement of the CP is aft to the 50% chord position.
13 - 18
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HIGH SPEED FLIGHT
PRINCIPLES OF FLIGHT
1·4
1· 0
MeRIT
50%
100%
PERCENTAGE CHORD
Figure 13.17
CP movement in the transonic region
The wing root usually has a thicker section than the wing tip so will have a lower M CRlT and
shock induced separation will occur at the root first The CP will move towards the tip, and if
the wing is swept, this CP movement will also be rearward. This effect will be discussed in
detail later.
Official US Navy Photograph
Low pressure area in front of shock wave
13 - 19
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
HIGH SPEED FLIGHT
EFFECT OF SHOCK WAVES ON CP MOVEMENT
Rearward CP movement with increasing Mach number in the transonic region produces a nose
down pitching moment. This is known as "Mach Tuck", "High Speed Tuck" or "Tuck under"
A further factor contributing to the nose down pitching moment is decreased downwash at the
tail resulting from reduced lift at the wing root. If the tailplane is situated in the downwash, its
effective angle of attack is increased, giving an increase in the nose down pitching moment.
For a stable aircraft a push force is required on the stick to produce an increase in speed, but as
a result of Mach tuck the push force required may dec ease with speed above M CRIT giving an
unstable stick force gradient, Fig. 13.l8.
PUSH
UNSTABLE STICK
FORCE GRADIENT
STICK
/
FORCE
0 4---~=---------~--------~---------
PULL
MCRIT
Figure 13.18
rv1ACH NUMBER
Reduction in stick force with increasing Mach number
THE EFFECT OF SHOCK WAVES ON FLYING CONTROLS
A conventional trailing edge control surface works by changing the camber of the aerofoil to
increase or decrease its lift. Deflecting a control surface down will reduce M C RIT' If the
control is moved down at high subsonic speed and a shock wave forms on the aerofoil ahead of
the control surface, shock induced separation could occur ahead of the control, reducing its
effectiveness. At low speed, movement of a control surface modifies the pressure distribution
over the whole aerofoil. If there is a shock wave ahead of the control surface, movement of the
control cannot affect any part of the aerofoil ahead of the shock wave, and this will also
reduce control effectiveness.
Conventional trailing edge control surfaces may suffer from greatly reduced effectiveness in the
transonic speed region and may not be adequate to control the changes of moment affecting the
aircraft at these speeds. This can be overcome by incorporating some or all of the following into
the design: an all moving (slab) tailplane (Fig. 11.2), roll control spoilers (Para. 11.17(c)),
making the artificial feel unit in a powered flying control system sensitive to Mach number or
by fitting vortex generators (Page 13-31).
13 - 20
© Oxford Aviation Services Limited
HIGH SPEED FLIGHT
PRINCIPLES OF FLIGHT
CONTROL BUZZ
If a shock wave is situated near to a control hinge a control movement may cause the shock wave
to move over the hinge, resulting in rapid changes of hinge moment which can set up an
oscillation of the control surface called control buzz.
13.15
BUFFET
In the same way that separated airflow prior to a low speed stall can cause airframe buffet, shock
induced separation (shock stall) at high speed can also cause buffeting.
Aerodynamic buffet is a valuable stall warning, but can damage the aircraft structure. Because
of the higher dynamic pressure when an aircraft is operating in the transonic speed region, any
shock induced buffet will have a greater potential for severe airframe damage. High speed buffet
must be completely avoided.
The aircraft must therefore be operated in such a manner that a (safety) margin exists before
aerodynamic buffet will occur.
If the variables which affect both high speed and low speed stall are considered it will be
possible to identify the conditions under which buffeting will occur and a chart can be drawn to
show all the factors involved. This is called a "Buffet Onset" chart (illustrated in Fig. 13.24)
which is used by flight crews to ensure their aircraft is operated at all times with a specified
minimum buffet margin.
In Chapter 7 it was shown that stall speed is affected by several factors. In this study of low
speed stall combined with high speed buffet, the factors to be considered are:a)
Load factor (Bank angle)
b)
Mach number
c)
Angle of Attack
d)
Pressure altitude
e)
Weight
f)
CG position
13 - 21
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
13.16
HIGH SPEED FLIGHT
FACTORS WHICH AFFECT THE BUFFET BOUNDARIES
STALL SPEED
As altitude is increased at a constant EAS, TAS will increase and outside air temperature will
decrease, causing the local speed of sound to decrease. Mach number is proportional to TAS and
inversely proportional to the local speed of sound (a):
M
=
TAS
a
Therefore, if altitude is increased at a constant EAS, Mach number will increase. At low speed
C Lmax is fairly constant, but above M 004 CLmax decreases as shown in Fig. 13.19. Refer also to
Fig. 13.11 for the reason why C LMAX starts to decreases at speeds above M 04.
C Lmax
Figure 13.19
From the Ig stall speed formula:
L
S
It can be seen that as C Lmax decreases with increasing altitude, the Ig stall speed will increase.
ALT
ALT 1
~ 19 Stall Speed
EAS
Figure 13.20
13 - 22
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PRINCIPLES OF FLIGHT
HIGH SPEED FLIGHT
Fig. 13 .20 shows the variation with altitude of stalling speed at constant load factor (n). Such
a curve is called the stall boundary for the given load factor, in which altitude is plotted against
equivalent airspeed. At this load factor (lg), the aircraft cannot fly at speeds to the left of this
boundary. It is clear that over the lower range of altitude, stall speed does not vary with altitude.
This is because at these low altitudes, V s is too low for compressibility effects to be present.
Eventually, V s has increased with altitude to such an extent that these effects are important, and
the rise in stalling speed with altitude is apparent.
As altitude increases, stall speed is initially constant then increases.
An altitude (Alt\ in Fig. 13.20) is eventually reached when there is only one speed at which the
aircraft can fly, since increasing or decreasing speed or banking the aircraft will result in a stall.
In the case of a 1 g manoeuvre, this altitude is called the' Aerodynamic Ceiling'. If the aircraft
were allowed to 'drift up ' to this altitude, the aircraft will stall. Not a pleasant prospect for a
modern high speedjet transport aircraft. This state of difficulty is also called 'coffin corner'.
Refer also to Fig. 13.23 .
NB:
The recovery in C L MAX at supersonic speeds is such that it may still be possible to operate above
this ceiling if enough thrust is available to accelerate the aircraft to supersonic speeds at this
altitude.
FL
CONSTANT
tvlACH NUMBER
/
\
Cl
Cl
'"
Cl
t?
Cl
N
Cl
10
N
"
EAS
Figure 13.21
LOAD FACTOR
Because load factor increases the stall speed, curves like the one sketched in Fig. 13 .20 can be
drawn for all values of load factor up to the maximum permissible 'g' , and together they
constitute the set of stalling boundaries for the given aircraft. Such a set of curves is shown in
Fig. 13.21. Superimposed on these curves are dashed lines representing lines of constant Mach
number, showing how high Mach number is achieved even at relatively low EAS at high
altitudes.
13 - 23
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
HIGH SPEED FLIGHT
Stall boundaries set a lower limit to the operating speed, according to the load factor. In the case
of a high-speed aircraft, there is also an upper limit which is due to the approach of shock stall
and the associated buffet which occurs if the aircraft enters the transonic speed range. The limits
associated with these effects give the buffet boundaries.
FL
STALL
BOUNDARY
EAS
Figure 13.22
MACH NUMBER
For a given aircraft there is a Mach number which, even at low angle of attack, cannot be
exceeded because of the onset of shock stall. Fig. 13.21 shows the EAS corresponding to this
Mach number falling as altitude increases, so the range of operating speeds is reduced at both
ends.
ANGLE OF ATTACK
However, there is a further effect which makes the buffet boundary a more severe limit than that
suggested by a curve of constant Mach number. As the EAS associated with a given Mach
number falls with increased altitude, so the required CL , and hence angle of attack increases.
This results in a reduction in the Mach number at which buffeting occurs, which results in a
further reduction in the permissible airspeed. This effect is made worse as the high angle of
attack stall is approached, and by the time the buffet boundary intersects the stall boundary the
limiting Mach number may be well below its value at a lower angle of attack, as Fig. 13.22
illustrates.
Also an increase in load factor (bank angle) requires an increase in lift at a given EAS, hence an
increase in angle of attack and a further reduction in limiting Mach number.
Thus the greater the load factor (bank angle or gust), the more severe the limitation due to
buffeting.
There is a set of buffet boundaries for various load factors (bank angles), just as there is a set
of stall boundaries.
The restrictions on speed and 'g' can be summarised in the form ofa single diagram in which
load factor is plotted against EAS, shown in Fig. 13.23.
13 - 24
© Oxford Aviation Services Limited
HIGH SPEED FLIGHT
PRINCIPLES OF FLIGHT
9~
"COFFIN CORNER"
rvlAXlMUM
PERMISSIBLE
o
9 - -
/
Vs
SEA LEVEL
ENVELOPE
EAS
Figure 13.23
PRESSURE ALTITUDE
At sea level there is a stall speed below which the aircraft cannot fly. As load factor increases,
so does the stall speed (proportional to the square root of the load factor). The curve of' g'
against EAS modifies the low speed stall boundary. It will continue to rise until the'limit load
factor' is reached (Chapt. 14). The 'limit load factor' must never be exceeded. At the high
speed end, when g = 1, there is a limiting speed which must not be exceeded because of shock
induced buffet. As the load factor increases, so does the C L at given speed, and the limiting
Mach number falls, slowly at first and then more rapidly. This defines a buffet boundary, which
eventually intersects the boundary of maximum permissible' g' , to constitute an overall envelope
like the outer curve depicted in Fig. 13.23 . Thus the aircraft may operate at any combination of
speed and load factor within this envelope, but not outside it.
At altitude the situation is similar. However, at altitude the equivalent stalling speed increases
with' g' rather more rapidly than at sea level, because of the Mach number effect on CL MAX .
Also, the buffet boundary becomes much more severe.
Above a certain altitude the buffet boundary may intersect the stall boundary at a value of 'g'
lower than the structural limit, as shown in Fig. 13.23. This 'point' is another representation of
"coffin corner".
WEIGHT
The weight ofthe aircraft also affects the envelope. An increase in weight results in an increase
in stall speed, and the stall boundary is moved to the right. It also results in an increase in angle
of attack at any given speed, so that the Mach number at which buffeting occurs is reduced, and
the buffet boundary is moved to the left. Finally, increase in weight implies a reduction in the
maximum permissible 'g'. Thus all the boundaries are made more restrictive by an increase in
weight.
CGPOSITION
Forward movement ofthe CG increases stall speed so the buffet boundaries will be affected in
a similar way to that due to weight increase.
13 - 25
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PRINCIPLES OF FLIGHT
13.17
HIGH SPEED FLIGHT
THE BUFFET MARGIN
It has been stated that an altitude can eventually be reached where there is only one speed at
which the aircraft can fly. In the case of a 1 g manoeuvre, this altitude is called the
'Aerodynamic Ceiling'. Operating an aircraft at its aerodynamic ceiling would leave no safety
margin. In 1 g flight the aircraft would be constantly on the point of stall. It could not be
manoeuvred nor experience the smallest gust without stalling. Regulations require an aircraft
to be operated with a minimum buffet margin of O· 3 g.
13.18
USE OF THE BUFFET ONSET CHART (Fig. 13.24)
1·3 g Altitude (1 g + 0·3 g = 1·3 g): At this altitude a 'g' increment of 0·3 can be sustained
without buffet occurring. Using the data supplied:Follow the vertical solid red line upwards from 1·3 g to the 110 tons line, then
horizontally to the 30% CG vertical line, then parallel to the CG reference line, again
horizontally to the M 0·8 vertical line. The altitude curve must now be 'parallelled' to
read-off the Flight Level of 405. The 1·3 g altitude is 40,500 ft.
If the aircraft is operated above FL 405 at this mass and CG a gust, or bank angle of less than
40 0, could cause the aircraft to buffet. (40 ° of bank at high altitude is excessive, a normal
operational maximum at high altitude would be 10° to 15°).
Buffet restricted speed limits: Using the data supplied:Follow the vertical dashed red line upwards from 1 g to the 110 tons line, then
horizontally to the 30% CG vertical line, then parallel to the CG reference line. Observe
the FL 350 curve. The curve does not reach the horizontal dashed red line at the high
speed end because M 0·84 (MMo)is the maximum operating speed limit. At the low
speed end of the dashed red line, the FL 350 curve is intersected at M 0·555. Thus
under the stated conditions, the low speed buffet restriction is M 0·555 and there is no
high speed buffet restriction because MMO is the maximum operating Mach number
which may not be exceeded under any circumstances.
Aerodynamic ceiling: at 150 tons can be determined by:Following the vertical dashed red line from 1 g to the 150 tons line, then following the
solid red line horizontally to M 0·8 (via the CG correction). The altitude curve gives an
aerodynamic ceiling of FL 390.
Load factor and bank angle at which buffet occurs: Using the data supplied:From M 0·8, follow the dashed blue line to obtain 54 ° bank angle or 1·7 g.
13 - 26
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HIGH SPEED FLIGHT
PRINCIPLES OF FLIGHT
BUFFET ONSET
clean configuration
REF
FLiG HT
LEVEL
150
NOTE :
FOR MAC H NUMBERS
ABOVE OR EQUAL TO
.82 THERE IS NO C.G.
VARIATION EFFECT
FROM REFERENCE
VALUE
VMO LIMIT
L
140
130 60
120 80
100
11 0 -
120
100
-
..
140 ~
.
90
160
17'
..
80
180
200
220
-
WEIGHT (I)
240
250
-n
r
.
'.
270
290
f
-
310
330
..
350
370
390
,
41 0
.50
~,O
.55
.60
.65
.70
.75
MAC H NUMBER
.85 15 20 25 30 35 40
.80
(
i 5H~~H:r l*
4550 55
L
1 12 1.4 1.6 1.8
CG%
60
•
65 ..
,I
2 2.2 2.4
LOAD FACTOR
MMO
RESULTS : BUFFET ONSET AT :
DATA : M = .80
- - FL=350
WEIGHT = 110 tons
CG = 30 %
M=0.80 WITH 54° BANK ANGLE , OR AT 1.7g
LOW SPEED (1 g) : M = 0 .555
HIGH SPEED : ABOVE M 0 .84 ( MMO )
1.3g ALTITUDE = FL405
Figure 13.24
Example of a buffet onset chart
13 - 27
© Oxford Aviation Services Limited
HIGH SPEED FLIGHT
PRINCIPLES OF FLIGHT
13.19
DELAYING OR REDUCING THE EFFECTS OF COMPRESSIBILITY
To maximise revenue, airlines require their aircraft to fly as fast and as efficiently as possible.
It has been shown that the formation of shock waves on the wing results in many undesirable
characteristics and a massive increase in drag. Up to speeds in the region of M CRIT the effects
of compressibility are not too serious. It is therefore necessary to increase M CRIT as much as
possible. Many methods have been adopted to delay or reduce the effects of compressibility to
a higher Mach number, as detailed below.
THIN WING SECTIONS
On a low tic ratio wing, the flow acceleration is reduced, thus raising the value ofM CRIT' For
example if MCRIT for a 15% tic wing is M 0'75, then MCRIT for a 5% tic wing will be
approximately M 0·85.
The use of a low tic ratio wing section has some disadvantages:
a)
The lift produced by a thin wing will be less, giving higher take-off and landing speeds
and increased distances.
b)
A thin wing requires disproportionally wider main spars for the same strength and
stiffness. This increases structural weight.
c)
Limited stowage space is available in a thin wing for:
i)
fuel
ii)
high lift devices and their actuating mechanism and
iii)
the main undercarriage and its actuating mechanism.
SWEEPBACK (see Page 13-43 for Sweepback Fact Sheet)
One of the most commonly used methods of increasing MCRIT is to sweep the wing back.
Forward sweep gives a similarly effect but wing bending and twisting creates such a problem
that sweep back is more practical for ordinary applications.
A simplified method of visualising the effect of sweepback is shown in Fig. 13.25. The swept
wing shown has the free stream velocity broken down to a component of velocity perpendicular
to the leading edge and a component parallel to the leading edge.
The component of velocity perpendicular to the leading edge is less than the free stream velocity
(by the cosine of the sweep angle) and it is this velocity component which determines the
magnitude of the pressure distribution. MCRIT will increase since the velocity component
affecting the pressure distribution is less than the free stream velocity.
13 - 28
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PRINCIPLES OF FLIGHT
HIGH SPEED FLIGHT
~
ELOCITY COMPONENT
PARALLEL TO LEADING
EDGE
FREE STREAM
VELOCITY
(
SWEEP ANGLE
VELOCITY COMPONENT
PERPENDICULAR TO
LEADING EDGE
Figure 13.25
Effect of sweep back
Alternatively, it can be considered
that compared to a straight wing, a
swept back wing of the same aero foil
section has a smaller effective
thickness chord ratio. Sweeping the
wing back increases the effective
aerodynamic chord for the same
dimensional thickness, Fig. 13.26.
SAME
CHORD
.
FREE
STREAM
FLOW
The local velocity will be lower for a
given free stream velocity. In this
way, the M eRIT of a swept wing will
be higher than a straight wing.
SA~
CHORD
II
EFFECTIVE
AERODYNAMIC .
CHORD
INCREASED
i
Figure 13.26
13 - 29
Sweeping the wing back has nearly
the same aerodynamic advantages as
a thin wing, without suffering
reduced strength and fuel capacity.
Unfortunately, there are some
disadvantages. It was explained in
Chapter 7 that swept back wings tend
to tip stall, leading to pitch-up and
possibly super stall. Swept back
wings also increase the magnitude of
high speed tuck.
© Oxford Aviation Services Limited
HIGH SPEED FLIGHT
PRINCIPLES OF FLIGHT
Another advantage of sweepback is the reduced lift curve slope. This is illustrated by the lift
curve comparison in Fig. 13.27 for the straight and swept wing.
STRAIGHT
{
SWEPT
ANGLE OF ATTACK
Figure 10.27 Effect of sweep back on sensitivity to gusts
Any reduction of lift curve slope makes the wing less sensitive to changes in angle of attack due
to a gust or turbulence. Since the swept wing has the lower lift curve slope, a given vertical gust
will increase the C L , and hence the load factor, by a smaller amount than would occur if the wing
were straight. (It is mentioned on Page 10 - 61 that sweeping the fin back reduces the tendency
for fin stall at high sideslip angles).
DISADV ANT AGES OF SWEEP
a)
Reduced C LMAX
i)
gives a higher stall speed and increased take-off and landing distances.
ii)
Maximum lift angle of attack is increased, which complicates the problem of
landing gear design (possibility of tail-strike) and reduced visibility from the
flight deck during take-off and landing. The contribution to stability of a given
tail surface area is also reduced (see Page 10 - 65 for a detailed description).
b)
A sweptback wing has an increased tendency to tip stall resulting in pitch-up at the stall
and possible deep stall problems (see Page 7 - 15 for a detailed description).
c)
Reduced effectiveness of trailing edge control surfaces and high lift devices because
their hinge line is swept. To produce a reasonable CLMAX on a swept wing the hinge line
of the inboard flaps may be made straight. Leading edge high lift devices are also used
to improve the low speed characteristics.
13 - 30
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
HIGH SPEED FLIGHT
VORTEX GENERATORS
It has been shown that most of the unfavourable characteristics associated with compressibility
are due to boundary layer separation behind the shock wave (shock stall).
Flow separation occurs because the boundary layer loses kinetic energy as it flows against the
adverse pressure gradient. Shock wave formation increases the adverse pressure gradient so the
loss of kinetic energy in the boundary layer will be greater.
Increasing the kinetic energy of the boundary layer will reduce flow separation. A very simple
device called vortex generators are used to re-energises the boundary layer.
Vortex generators are small plates, vanes, blades or wedges mounted in spanwise rows along the
wing surface, as illustrated in Fig. 13.28. (see also page 7 - 14)
Each vortex generator
produces a vortex at its tip
which will induce high
energy air from the free
stream flow to mix with the
boundary layer, thus
increasing its kinetic energy
and helping it flow through
the shock wave with much
less separation.
Figure 13.28 Vortex generators (blade type)
13 - 31
Vortex generators are usually
located on the upper wing
surface, particularly ahead of
control surfaces, but may be
used anywhere where
separation is causing high
drag, or reduced control
effectiveness. It should be
noted that vortex generators
may also be used on subsonic
aircraft, to prevent separation
caused by high adverse
pressure gradients due to the
contours of the surface.
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
HIGH SPEED FLIGHT
AREA RULE
On Page 6 - 4 it was stated that in addition to the drag of individual components there is an extra
drag due to interference between these components, principally between wing and fuselage. This
is especially important at high speed. Experiments have shown that a large part of the transonic
drag rise for a complete aircraft is due to interference. Interference drag at transonic speeds may
be minimised by ensuring that the cross-sectional area distribution along the aircraft' s
longitudinal axis follows a certain smooth pattern.
«w
0:::
«
TAIL
FUSELAGE
NOSE
TAIL
Figure 13.29
Without Area Rule
With some early high speed aircraft designs this was not the case. The area increased rapidly
in the region of the wing, again in the vicinity of the tail and decreased elsewhere, giving an area
distribution like the one illustrated in Fig. 13.29.
On later aircraft, the fuselage was waisted, i.e., the area was reduced in the region of the wing
attachment, and again near the tail, so that there was no "hump" in the area distribution, giving
a distribution like the one illustrated in Fig. 13.30. There is an optimum area distribution, and
the minimisation of transonic interference drag requires that the aircraft should be designed to
fit this distribution as closely as possible. This requirement is known as the 'transonic area rule' .
In practice, no aircraft has this optimum distribution, but any reasonably smooth area distribution
helps to reduce the transonic drag rise.
«
w
0:::
«
WING
TAIL
FUSELAGE
NOSE
TAIL
Figure 13.30 Area Rule
13 - 32
© Oxford Aviation Services Limited
HIGH SPEED FLIGHT
PRINCIPLES OF FLIGHT
MACH TRIM
It was stated on page 13 - 20 that as speed increases beyond MeRIT' shock wave formation at the
root of a swept back wing will generate a nose down pitching moment because lift forward of
the CG is reduced and downwash at the tailplane is reduced.
At high Mach numbers an aircraft will tend to become speed unstable. Instead of an increasing
push force being required as speed increases, a pull force becomes necessary to prevent the
aircraft accelerating further. This is potentially very dangerous. A small increase in Mach
number will give a nose down pitching moment which will tend to further increase the Mach
number. This in tum leads to a further increase in the nose down pitching moment. This
unfavourable high speed characteristic, known as "Mach Tuck", "High Speed Tuck" or "Tuck
Under" would restrict the maximum operating speed of a modem high speed jet transport
aircraft.
Some improvement can be made by mounting the tailplane on top of the fin, where it is clear of
the downwash, but it has been shown that this can produce a deep stall problem.
To maintain the required stick force gradient at high Mach numbers, a Mach trim system
must be fitted. This device, sensitive to Mach number, may:
a)
deflect the elevator up
b)
decrease the incidence ofthe variable incidence trimming tailplane or
c)
move the CG rearwards by transferring fuel from the wings to a rear trim tank
by an amount greater than that required merely to compensate for the trim change. This ensures
the required stick force gradient is maintained in the cruise at high Mach numbers.
Whichever method of trim is used by a particular manufacturer, a Mach trim system will adjust
longitudinal trim and operates only at high Mach numbers.
PUSH
STICK
FORCE
O ~~~--------+-----~.------
PULL L-__________~______________
M CRIT
Figure 13.31
I'v1ACH NUMBER
Effect of Mach Trim
13 - 33
© Oxford Aviation Services Limited
HIGH SPEED FLIGHT
PRINCIPLES OF FLIGHT
SUPERCRITICAL AEROFOIL
A fairly recent design development, used to increase efficiency when operating in the transonic
speed region, is the 'supercritical aerofoil'.
LARGE
ESS
TH
Ir
FLAT UPPER
SURFACE
S - SHAPED CAMBER LINE
THICK TRAILING
EDGE
BLUNT
NOS
~'7%
12 %
CONVENTIONAL
SECTION
Figure 13.32
SUPERCRITICAL
SECTION
Supercritical aerofoil shape
A supercritical aerofoil shape, illustrated in Fig. 13.32, differs from a conventional section by
having:
a)
a blunt nose,
b)
large thickness,
c)
an S - shaped camber line,
d)
a relatively flat upper surface and
e)
a thick trailing edge.
Because the airflow does not achieve the same increase of speed over the flattened upper surface
compared to a conventional section, the formation of shock waves is delayed to a higher M FS and
are much smaller and weaker when they do form.
Because the shock waves are smaller and weaker there is not such a sharp pressure rise on the
rear of the section and this gives a much more even 'loading' on the wing.
13 - 34
© Oxford Aviation Services Limited
HIGH SPEED FLIGHT
PRINCIPLES OF FLIGHT
THE ADVANTAGES OF A SUPERCRITICAL AEROFOIL
a)
Because of the delayed formation of shock waves and their weaker nature, less sweep
angle is required for a given cruising Mach number, thus reducing some of the problems
associated with sweepback.
b)
The greater thickness gives increased stiffness and strength for a given structural weight.
This also allows a higher aspect ratio to be used which reduces induced drag.
c)
The increased section depth gives more storage space for fuel.
This type of wing section can be used to increase performance in one of two ways:
i)
Increased Payload: By using existing cruise speeds, the fuel consumption
would be reduced, thus allowing an increase in payload with little or no drag
increase over a conventional wing at the same speed.
ii)
Increased Cruising Speed: By retaining existing payloads, the cruise Mach
number could be increased with little or no increase in drag.
THE DISADVANTAGES OF A SUPERCRITICAL AEROFOIL
a)
The aerofoil front section has a negative camber to give optimum performance at cruise
Mach numbers, but this is less than ideal for low speed flight. C LMAX will be reduced,
requiring extensive and complex high lift devices at the leading edge, which may
include Krueger flaps, variable camber flaps, slats and slots.
b)
The trailing edge of the aerofoil has large positive camber to produce the 'aft loading'
required, but which also gives large negative (nose down) pitching moments.
i)
This must be balanced by the tailplane, causing trim drag and
ii)
Shock induced buffet may cause severe oscillations.
13 - 35
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
13.20
HIGH SPEED FLIGHT
AERODYNAMIC HEATING
Air is heated when it is compressed or when it is subjected to friction. An aircraft will have
compression at the stagnation point, compression through a shock wave, and friction in the
boundary layer.
500
400
()
0
ui
300
0:::
:::>
I-
~
200
I-
100
w
a.
:2:
w
0
-40
-100
0
2
3
4
MACH NUMBER
Figure 13.33
Surface temperature rise
with Mach number
So when an aeroplane moves through the air it's skin temperature will increase. This occurs at
all speeds, but only becomes significant from a skin temperature point of view at higher Mach
numbers.
It can be seen from Fig. 13 .33 that the temperature rise at M 1·0 is approximately 40 °C. Again
from a skin temperature point of view, this rise in temperature does not become significant until
speeds in the region of M 2·0 are reached, which is the approximate limit speed for aircraft
manufactured from conventional aluminium alloys. Above this speed the heat treatment of the
structure would be changed and the fatigue life shortened. For speeds above Mach 2 '0, Titanium
or "stainless steel" must be used.
13 - 36
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
13.21
HIGH SPEED FLIGHT
MACH ANGLE
Reference to Fig. 13.7 will show that as the Mach number increases the shock waves become
more acute. To illustrate why the angle of the shock waves change it is necessary to consider
the meaning and significance ofthe Mach angle ').1' (mu).
If the T AS of the aircraft is greater than the local speed of sound, the source of pressure waves
is moving faster than the disturbance it creates.
E
MACH LlNE ~
OR WAVE
~
LOCAL SPEED
OF SOUND
B
... DIRECTION
OF FLIGHT
D~--~~-----+--r---+-~--------r--e
VELOCITY OF AIRCRAFT,
V
Figure 13.34
Mach angle
Consider a point moving at velocity ' V' in the direction' A ' to 'D', as in Fig. 13 .34. A pressure
wave propagated when the point is at 'A' will travel spherically outwards at the local speed of
sound; but the point is moving faster, and by the time it has reached 'D' , the wave from 'A' and
other pressure waves sent out when the point was at 'B ' and 'C' will have formed circles as
shown, and it will be possible to draw a common tangent 'DE' to these pressure waves. The
tangent represents the limit to which all the pressure waves have reached when the point has
reached ' D ' .
' AE' represents the local speed of sound (a) and 'AD' represents the TAS (V)
=
M
The angle ' ADE' , or
).1
TAS
as illustrated, M
a
=
2.6
is called the Mach angle and by simple trigonometry:
sin
).1
=
a
TAS
=
1
M
The greater the Mach number, the more acute the Mach angle
13 - 37
).1 .
At M 1'0,
).1
is 90 0 •
© Oxford Aviation Services Limited
HIGH SPEED FLIGHT
PRINCIPLES OF FLIGHT
13.22
MACH CONE
In three dimensions, the disturbances propagating from a moving point source expand outward
as spheres, not circles. If the speed of the source (V) is greater than the local speed of sound (a),
these spheres are enclosed within a Mach cone, whose semi vertical angle is ~.
rvlACH CONE ~
v
Figure 13.35
Mach cone at approximately M 5·0
It can be seen from Fig. 13.35 that the Mach angle
(~)
continues to decrease with increasing
Mach number. The Mach angle is inversely proportional to the Mach number.
13.23
AREA (ZONE) OF INFLUENCE
When travelling at supersonic speeds the Mach cone represents the limit of travel of the pressure
disturbances created by an aircraft, anything forward ofthe Mach cone cannot be influenced
by the disturbances. The space inside the Mach cone is called the area or zone of influence.
A finite body such as an aircraft will produce a similar pattern of waves but the front will be an
oblique shock wave and the wave angle will be greater than the Mach angle because the initial
speed of propagation of the shock waves will be greater than the free stream speed of sound.
13.24
BOW WAVE
Consider a supersonic stream approaching the leading edge of an aerofoil. In order to flow
around the leading edge the air would suddenly have to turn through a right angle (see Fig. 13.3).
At supersonic speeds this is not possible in the distance available. The free stream velocity will
suddenly decelerate to below supersonic speed and a normal shock wave will form ahead of the
wing at the junction of supersonic and subsonic airflow. Behind the shock wave the airflow is
subsonic and is able to flow around the leading edge. Within a short distance the flow again
accelerates to supersonic speed, as illustrated in Fig. 13 .36.
13 - 38
© Oxford Aviation Services Limited
HIGH SPEED FLIGHT
PRINCIPLES OF FLIGHT
/
OBLIQUE
BOW WAVE
I'
SHOCK
NORMAL
SHOCK
Figure 13.36
Bow wave
The shock wave ahead of the leading edge is called a bow wave and is normal only in the
vicinity of the leading edge. Further away from the leading edge ("above" and "below") it
becomes oblique. It can be seen in Fig. 13 .36 that the trailing edge shock waves are no longer
normal because the free stream mach number is greater than 1·0; they are also now oblique.
13.25
EXPANSION W AVES
In the preceding paragraphs it has been shown that supersonic flow is able to turn a corner by
decelerating to subsonic speed when it meets an object. A shock wave forms at the junction of
the supersonic and subsonic flow , the generation of which is wasteful of energy (wave drag).
There is another way a supersonic flow is able to turn a corner. Consider first a convex corner
with a subsonic flow, as illustrated in Fig. 13.37.
SUBSONIC
. _ --
-
FLOW
Figure 13. 37 Subsonic flow at a convex corner
With subsonic airflow the adverse pressure gradient would be so steep that the airflow would
instantly separate at the "corner".
13 - 39
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
HIGH SPEED FLIGHT
EXPANSION
WAVE
~~
j
SUPERSONIC
VELOCITY
FLOW
UP
PRESSURE,
DENSITY
AND
TEMPERATURE
DOWN
Figure 13.38
Supersonic flow at a convex corner with expansion wave
Fig. 13 .38 shows that a supersonic airflow can follow a convex comer because it expands upon
reaching the comer. The velocity INCREASES and the other parameters, pressure, density and
temperature DECREASE. Supersonic airflow behaviour through an expansion wave is exactly
opposite to that through a shock wave.
M FS A BOV E 1· 2
/
/
/
/
/
/
/
/
/
EXPANSION WAVES
/
/
/
/
/
\
\
\
\
\
\
Figure 13.39
\
\
\
/
\
\
\
/
\
\
'
/
/
\
\
Expansion waves in a supersonic flow
13 - 40
© Oxford Aviation Services Lim ited
HIGH SPEED FLIGHT
PRINCIPLES OF FLIGHT
Fig. 13.39 shows a series of expansion waves in a supersonic airflow. After passing through the
bow shock wave, the compressed supersonic flow is free to expand and follow the surface
contour. As there are no sudden changes to the airflow, the expansion waves are NOT shock
waves. A supersonic airflow passing through an expansion wave will experience the following
changes:-
13.26
a)
The airflow is accelerated; the velocity and Mach number behind the expansion wave
are greater.
b)
The flow direction is changed to follow the surface.
c)
The static pressure of the airflow behind the expansion wave is decreased.
d)
The density of the airflow behind the expansion wave is decreased.
e)
Since the flow change is gradual there is no "shock" and no loss of energy in the airflow.
An expansion wave does not dissipate airflow energy.
SONIC BANG
The intensity of shock waves reduces with distance from the aircraft, but the pressure waves can
be of sufficient magnitude to create a disturbance on the ground. Thus, "sonic bangs" are a
consequence of supersonic flight. The pressure waves move with aircraft groundspeed over
the earth surface.
13.27
METHODS OF IMPROVING CONTROL AT TRANSONIC SPEEDS
It has been seen that control effectiveness may decrease in the transonic region if conventional
control surfaces are used. Some improvement in control effectiveness may be obtained by
placing vortex generators ahead of control surfaces.
However, alternative forms of control such as:
a)
an all moving (slab) tailplane
b)
or roll control spoilers give better control in the transonic speed region.
These types of control are explained in Flying Controls Chapter 11. Control surface buzz is
sometimes remedied by fitting narrow strips along the trailing edge of the control surface, or may
be prevented by including dampers in the control system or by increasing the stiffness of the
control circuit.
Because of the high control loads involved at high speeds and the variation in loads through the
transonic region, the controls will normally be fully power operated with artificial feel.
13 - 41
© Oxford Aviation Services Limited
HIGH SPEED FLIGHT
PRINCIPLES OF FLIGHT
The table in Fig. 13.40 is provided to summarise the characteristics of the three principal wave
forms encountered with supersonic flow.
Supersonic Wave Characteristics
TYPE OF WAVE
OBLIQUE Shock wave
NORMAL Shock wave
I
-
FLOW DIRECTION
CHANGE
TURNED INTO A PRECEDING
FLOW
EFFECT ON VELOCITY
and MACH NUMBER.
BEHIND WAVE
DECREASED BUT STILL
SUPERSONIC.
Figure 13.40
/ /
11/
INCREASE
A PLANE OF DISCONTINUITY,
NORMAL TO FLOW DIRECTION.
NO CHANGE
TURNED AWAY FROM
PRECEDING FLOW .
DECREASED TO SUBSONIC
INCREASED TO HIGHER
SUPERSONIC.
GREAT INCREASE
DECREASE .
----------------,--------" - - - - - - - - - , - - - - -
EFFECT ON
TEMPERATURE
/
~
A PLANE OF DISCONTINUITY ,
INCLINED MORE THAN 90'
FROM FLOW DIRECTION.
EFFECT ON ENERGY
OF AIRFLOW.
/
_ -•• 111/
DEFINrTlON
EFFECT ON STATIC
PRESSURE and DENSITY .
EXPANSION wave
----.-~--
DECREASE
GREAT DECREASE
INCREASE
INCREASE
NO CHANGE (NO SHOCK) .
DECREASE.
Characteristics of the three principal wave forms
13 - 42
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
HIGH SPEED FLIGHT
SWEEPBACK - FACT SHEET
Sweep Angle: The angle between the line of25% chords and a perpendicular to the root chord
Purpose of Sweep back: To increase MeRIT'
A SWEPT WING INCREASES THE CRITICAL MACH NUMBER (MeRIT)' All other
effects from a swept wing are by-products, most of them disadvantages. However, the benefits
from a higher MeRIT outweigh the associated disadvantages.
BY - PRODUCTS OF SWEEPBACK
1.
Increased tendency to stall at the tip first - minimised by fitting wing fences , vortilons
or saw tooth leading edge.
a)
b)
c)
Tip stall can lead to pitch - up, a major disadvantage.
Pitch - up can give the tendency for a swept wing aircraft to Super Stall.
Aircraft that show a significant tendency to Pitch - up at the stall MUST be
fitted with a stall prevention device; a stick Pusher.
Close to the stall, ailerons and coordinated use of rudder should be used to maintain
wings level because the use of rudder alone would give excessive rolling moments.
(V SRis adjusted so that adequate roll control exists from the use of ailerons close to the
stall).
2.
When compared to a straight wmg of the same section, a swept WlOg
aerodynamically efficient.
(i)
(ii)
(iii)
IS
less
At a given angle of attack C L is less.
C LMAX is less and occurs at a higher angle of attack.
The lift curve has a smaller gradient (change in C L per degree change in alpha
is less).
HIGH ASPECT
RATIO
LOW ASPECT
RATIO
(or sweepback)
13 - 43
© Oxford Aviation Services Limited
HIGH SPEED FLIGHT
PRINCIPLES OF FLIGHT
a)
Swept wings must be fitted with complex high lift devices, both leading and
trailing edge, to give a reasonable take - off and landing distance.
(i)
The least efficient type of leading edge device is used on the inboard
part of the swept wing to help promote root stall.
b)
Because of the higher stalling angle of attack, the fin or vertical stabiliser is
swept to delay fin stall to a greater sideslip angle.
c)
A swept wing must be flown at a higher angle of attack than a straight wing to
give the required lift coefficient, this is most noticeable at low speeds.
d)
One of the few advantages of a swept wing is that it is less sensitive to changes
in angle of attack due to gust or turbulence; a smaller change in Load Factor for
a given gust will result.
3.
A swept wing makes a small positive contribution to static directional stability.
4.
A swept wing makes a significant positive contribution to static lateral stability.
5.
At speeds in excess of MeRIT a swept wing generates a nose down pitching moment; a
phenomena known as Mach Tuck, High Speed Tuck or Tuck Under. This must be
counteracted by a Mach Trim System which adjusts the aircraft's longitudinal trim.
6.
The hinge line of trailing edge 'flap' type control surfaces are not at right angles to the
airflow, which reduces the efficiency of the controls.
13 - 44
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
HIGH SPEED FLIGHT
SELF ASSESSMENT
1.
Identify which of the following is the correct formula for Mach number:
a)
b)
2.
M
=
constant
lAS
a
M
d)
M = TAS x a
a
What is the result of a shock-induced separation of airflow occurring symmetrically near the
wing root of a sweptwing aircraft?
a severe nose-down pitching moment or "tuck under"
a high-speed stall and sudden pitch up
severe porpoising
pitch-up
Mach number is:
the ratio of the aircraft's TAS to the speed of sound at sea level.
the ratio of the aircraft's TAS to the speed of sound at the same atmospheric conditions.
the ratio of the aircraft's lAS to the speed of sound at the same atmospheric conditions.
the speed of sound.
For an aircraft climbing at a constant lAS the Mach number will:
a)
b)
c)
d)
5.
=
M a
TAS =
a)
b)
c)
d)
4.
TAS
c)
a)
b)
c)
d)
3.
QUESTIO~S
increase.
decrease.
remain constant.
initially show an increase, then decrease.
The term 'transonic speed' for an aircraft means:
a)
b)
c)
d)
speeds where the airflow is completely subsonic.
speeds where the airflow is completely supersonic.
speeds where the airflow is partly subsonic and partly supersonic.
speeds between M 0.4 and M 1.0
13 - 45
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
6.
At M 0·8 a wing has supersonic flow between 200/0 chord and 600/0 chord. There will be a
shockwave:
a)
b)
c)
d)
7.
will start to increase.
will start to decrease.
will remain constant.
is directly proportional to the Mach number.
As air flows through a shockwave:
a)
b)
c)
d)
11.
will decrease if angle of attack is increased.
will increase if angle of attack is increased.
will not change with changes of angle of attack.
is only influenced by changes in temperature.
At speeds just above the critical Mach number, the lift coefficient:
a)
b)
c)
d)
10.
static pressure increases, density decreases, temperature increases.
static pressure increases, density increases, temperature increases.
static pressure decreases, density increases, temperature decreases.
static pressure decreases, density decreases, temperature decreases.
For a wing section of given thickness, the critical Mach number:
a)
b)
c)
d)
9.
at 20% chord only.
at 20% chord and 60% chord.
at 60% chord only.
forward of 20% chord.
As air flows through a shockwave:
a)
b)
c)
d)
8.
HIGH SPEED FLIGHT
its speed increases.
its speed decreases.
its speed remains the same.
it changes direction to flow parallel with the Mach cone.
If an aeroplane accelerates above the Critical Mach number, the first high Mach number
characteristic it will usually experience is:
a)
b)
c)
d)
a nose up pitch or "Shock Stall".
a violent and sustained oscillation in pitch (porpoising).
Dutch roll and/or spiral instability.
a nose down pitching moment (Mach, or high speed tuck).
13 - 46
© Oxford Aviation Services Limited
HIGH SPEED FLIGHT
PRINCIPLES OF FLIGHT
12.
High speed buffet is caused by:
a)
b)
c)
d)
13.
The "area rule" applied to high speed aircraft requires:
a)
b)
c)
d)
14.
d)
moves the centre of gravity to maintain stable lateral stick forces in the transonic region.
automatically compensates for pitch changes while flying in the transonic speed region.
prevents the aircraft from exceeding its critical Mach number.
switches out the trim control to prevent damage in the transonic region.
What is the movement of the centre of pressure when the wingtips ofa sweptwing aeroplane are
shock-stalled first?
a)
b)
c)
d)
17.
because the effect of the elevator is reversed above the critical Mach number.
because shock wave formation on the elevator causes excessive stick forces.
because shock wave formation ahead of the elevator causes separation and loss of
elevator effectiveness.
because it would be physically impossible for a pilot to control the aircraft in pitch with
a conventional tailplane and elevator configuration.
Mach Trim is a device which:
a)
b)
c)
d)
16.
that the cross sectional area shall be as small as possible.
that the variation of cross sectional area along the length of the aircraft follows a smooth
pattern.
that the maximum cross sectional area of the fuselage should occur at the wing root.
that the fuselage and the wing area be of a ratio of 3 : 1.
An all-moving tailplane is used in preference to elevators on high speed aircraft:
a)
b)
c)
15.
the shock waves striking the tail.
the high speed airflow striking the leading edge of the wing.
wing flutter caused by the interaction of the bottom and top surface shock waves.
the airflow being detached by the shock wave and the turbulent flow striking the tail.
outward and forward
inward and aft
outward and aft
inward and forward
The airflow behind a normal shock wave will:
a)
b)
c)
d)
always be subsonic and in the same direction as the original airflow.
always be supersonic and in the same direction as the original airflow.
may be subsonic or supersonic.
always be subsonic and will be deflected from the direction of the original airflow.
13 - 47
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
18.
As airflow passes through a normal shock wave, which of the following changes in static
pressure (i), density (ii), and Mach number (iii) will occur:
a)
b)
c)
d)
19.
d)
c)
d)
(iii)
< 1.0
< 1.0
> 1.0 or < 1.0
< 1.0
have its centre of pressure at 50 % chord.
have its centre of pressure at 25% chord.
give a larger proportion of lift from the lower surface than from the upper surface, and
have its centre of pressure at 50 % chord.
give approximately equal lift from the upper and lower surfaces, and have its
aerodynamic centre at 50% chord.
a shock wave which forms on the nose of the aircraft at M CRIT'
the shape formed when the shock waves on the upper and lower wing surface meet at
the trailing edge.
a shock wave that forms immediately ahead of an aircraft which is travelling faster than
the speed of sound.
the shape of a shock wave when viewed vertically.
When an aircraft is flying at supersonic speed, where will the area of influence of any pressure
disturbance due to the presence of the aircraft be located?
a)
b)
c)
d)
22.
(ii)
increase
decrease
decrease
increase
A bow wave is:
a)
b)
21.
(i)
decrease
Increase
Increase
Increase
An aerofoil travelling at supersonic speed will:
a)
b)
c)
20.
HIGH SPEED FLIGHT
Within the Mach Cone.
In front of the Mach Cone.
In front of the bow wave.
In front of the Mach Cone only when the speed exceeds M 1.0
The temperature of the airflow as it passes through an expansion wave:
a)
b)
c)
d)
Increases.
decreases.
is inversely proportional to the square root of the Mch number.
remains the same.
13 - 48
© Oxford Aviation Services Limited
HIGH SPEED FLIGHT
PRINCIPLES OF FLIGHT
23.
The influence of weight (wing loading) on the formation of shockwaves is:
a)
b)
c)
d)
24.
a higher wing loading will increase M CRIT
low wing loading will give a higher M CRIT
wing loading does not influence M CRIT
wing loading and M CRIT are directly proportional
What influence does an oblique shock wave have on the streamline pattern (i), variation of
pressure (ii), temperature(iii), density (iv) and velocity (v)?
(i)
a)
b)
c)
d)
25.
increase
decrease
decrease
decrease
mcrease
decrease
decrease
increase
(v)
decrease
Increase
Increase
decrease
shock waves interfering with the smooth airflow into the engine intakes.
flying faster than M MO.
the conversion of mechanical energy into thermal energy by the shock wave.
flying faster than V MO
What is the effect of a shock wave on control surface efficiency?
a)
b)
c)
d)
27.
increase
decrease
decrease
increase
(iv)
Wave drag is caused by:
a)
b)
c)
d)
26.
parallel to surface
normal to wave
parallel to wave
parallel to chord
(iii)
(ii)
Increase in efficiency, due to increased velocity.
Increase in efficiency, due to the extra leverage caused by the shock wave.
Decrease in efficiency, due to the bow wave.
loss of efficiency, due to control deflection no longer modifying the total flow over the
wmg.
At what speed does an oblique shock wave move over the earth surface?
a)
b)
c)
d)
Aircraft ground speed
The TAS of the aircraft plus the wind speed
The TAS of the aircraft less the wind speed
The TAS relative to the speed of sound at sea level
13 - 49
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
HIGH SPEED FLIGHT
ANSWERS
I No I A I B I c I D II
1
A
2
A
3
4
A
C
6
C
7
B
A
9
B
10
B
11
D
12
D
13
B
14
C
15
B
16
17
D
A
18
19
D
A
20
21
C
A
22
B
23
B
24
A
25
C
26
27
I
B
5
8
REF
D
A
13 - 51
© Oxford Aviation Services Limited
CHAPTER 14 - LIMITATIONS
Contents
Page
OPERATING LIMIT SPEEDS ....................... , .....................
LOADS AND SAFETY FACTORS
LOADS ON THE STRUCTURE
LOAD FACTOR ........................................................
THE MANOEUVRE ENVELOPE ..........................................
THE C LMAX BOUNDARY
DESIGN MANOEUVRING SPEED VA .....................................
EFFECT OF AIRCRAFT WEIGHT ON V A
DESIGN CRUISING SPEED Vc ...........................................
DESIGN DIVE SPEED V D
NEGATIVE LOAD FACTORS
THE NEGATIVE STALL ................. , ..............................
MANOEUVRE BOUNDARIES
OPERATIONAL SPEED LIMITS ..........................................
14 - 1
14 - 2
14 - 3
14 - 5
14 - 6
14-7
14 - 8
VMo/M Mo
VNE
V NO
GUST LOADS ......................................................... 14 - 9
EFFECT OF A VERTICAL GUST ON THE LOAD FACTOR .................. 14 - 10
EFFECT OF THE GUST ON STALLING ................................... 14 - 11
OPERATIONAL ROUGH AIR SPEED ..................................... 14 - 12
LANDING GEAR SPEED LIMITATIONS .................................. 14 - 14
VLO
V LE
FLAP SPEED LIMIT ................................................... 14 - 15
VFE
AEROELASTICITY (AERO - ELASTIC COUPLING) ........................
FLUTTER ............................................................
CONTROL SURFACE FLUTTER ..................................
TORSIONAL AILERON FLUTTER
FLEXURAL AILERON FLUTTER
AILERON REVERSAL .................................................
LOW SPEED
HIGH SPEED
SELF ASSESSMENT QUESTIONS .......................................
14 - 16
14 - 19
14 - 20
14 - 22
14 - 25
PRINCIPLES OF FLIGHT
14.1
LIMITATIONS
OPERATING LIMIT SPEEDS
In service an aircraft must observe certain speed limitations. These may be maximum speeds or
minimum speeds, but in each case they are set to give safe operation in the prevailing conditions.
The limits may be set by various considerations, the main ones being:a)
strength of the aircraft structure
b)
stiffness of the aircraft structure
c)
adequate control of the aircraft
Strength is the ability of the structure to withstand a load, and stiffness is the ability to withstand
deformation.
14.2
LOADS AND SAFETY FACTORS.
Limit load:
The maximum load to be expected in service
Ultimate load:
The failing load of the structure
Factor of safety:
The ratio of ultimate load to limit load
For aircraft structures the factor of safety is 1·5
The safety factor on aircraft structures is much lower than the safety factors used in other forms
of engineering because of the extreme importance of minimum weight in aircraft structures. To
keep the weight as low as possible the safety factor must be kept to a minimum. Because of this
it is extremely important not to exceed the limitations set on the operation of the aircraft, as the
safety margin can easily be exceeded and structural damage occur.
14.3
LOADS ON THE STRUCTURE
The airframe structure must obviously be strong enough to take the loads acting upon it in
normal level flight, that is the forces due to lift, drag, thrust and weight. However the aircraft
is also required to manoeuvre and to fly in turbulent air. Under these conditions the loads on the
aircraft will be increased so it must also be strong enough to withstand whatever manoeuvres are
specified for the aircraft and by the gusts which are required to be considered.
The structure should also have sufficient stiffness to ensure that phenomena such as aileron
reversal, flutter, and divergence do not occur within the permitted speed range of the aircraft.
14 - 1
© Oxford Aviation Training
PRINCIPLES OF FLIGHT
14.4
LIMITATIONS
LOAD FACTOR
The loads which must be considered are given in the design requirements of an aircraft. They
are given in terms ofload factor (n), colloquially known as 'g'.
Load factor (n) =
Lift
Weight
In level flight, since lift equals weight, the load factor is 1·0 (lg). If the aircraft is performing
a manoeuvre such that, for example the lift is twice the weight, the load factor is 2·0 (2g).
The limit load is given in terms of load factor to make the requirement general to all aircraft.
However, it should be appreciated that failure of the structure will occur at some particular
applied load. For example, if the structure fails at 10,000 lb load, an aircraft weighing 4,000 lb
will reach this load at a load factor of2·5. However, if the aircraft weighs 5,000 lb the failing
load is reached at a load factor of 2 '0, i.e. it takes less' g' to overstress a heavy aircraft than a
light one.
Limit load factors are based on the maximum weight of the aircraft.
C
.......
3
POSITIVE
CLmax
2
~
0:::
A
C
0
~
1
0
«
u.
0
«
0
E
0
SPEED
(EAS)
--l
-1
VA
NEGATIVE CL
max
I
VD
Figure 14.1 The Manoeuvre Envelope
14 - 2
© Oxford Aviation Training
LIMITATIONS
PRINCIPLES OF FLIGHT
14.5
THE MANOEUVRE ENVELOPE (V - n Diagram)
The maximum load factors which must be allowed for during manoeuvres are shown in an
envelope of load factor against speed (EAS). Fig.14.1 shows a typical manoeuvre envelope or
V - n diagram.
The limit load factors will depend on the design category of the aircraft.
The JAR regulations state that:
a)
For normal category aircraft, the positive limit load factor may not be less than 2'5 and
need not be more than 3' 8.
(So that structural weight can be kept to an absolute minimum a manufacturer will not
design an aircraft to be any stronger than the minimum required by the regulations).
The positive limit load factor for modern high speed jet transport aircraft is 2·5.
b)
For utility category aircraft the positive limit load factor is 4·4
c)
For aerobatic category aircraft the positive limit load factor is 6'0
The negative limit load factor may not be less than:
14.6
d)
- 1·0 for normal category aircraft
e)
-1,76 for utility category aircraft
f)
- 3·0 for aerobatic category aircraft
THE CLMAX BOUNDARY
The line OA in Fig.14.1 is determined by the CLMAX of the aircraft. In theory, the lift, and hence
the load factor for a given weight, depends on the angle of attack of the wing and the airspeed.
The maximum possible lift will occur at the angle of attack where CL is a maximum. At this
angle of attack the lift will increase with speed as shown by the line OA.
For level (lg) flight the speed at CLMAX will be the stalling speed (V s), represented by point S
in Fig. 14.1.
At Point A, the load factor reaches its positive limit.
14 - 3
© Oxford Aviation Training
PRINCIPLES OF FLIGHT
LIMITATIONS
It can be seen from Fig.14.2 that at speeds below point A the wing cannot produce a lift force
equal to the limit load factor, whereas at speeds above point A the limit load factor can be
exceeded. Manoeuvres at speeds above point A therefore have the potential to cause permanent
deformation to the structure or structural failure if the ultimate load is exceeded.
This does not mean that any manoeuvre at a speed greater than point A will always cause
structural damage; manoeuvres may be performed safely provided that the limit load factor is
not exceeded.
,------- PERIv1ANENT DEFORIv1ATION OF STRUCTURE POSSIBLE
/
4
/
/
/
/
/
/
/
/
/
STRUCTURAL FAILURE
3
0::
o
I()
~
o
2
«o
....J
o
SPEED (EAS)
Figure 14.2 Loads Imposed During Manoeuvres
There is of course a safety factor on the airframe of 1·5 so complete failure of the structure will
not occur at the load factor of2·5 but at 2·5 x 1·5 = 3·75.
However, permanent deformation of the structure may occur at load factors between 2·5 and
3·75, so it is not safe to assume that the load factor may be increased above the limiting value
just because there is a safety factor.
14 - 4
© Oxford Aviation Training
PRINCIPLES OF FLIGHT
14.7
LIMITATIONS
DESIGN MANOEUVRING SPEED, VA
The highest speed at which sudden, full elevator deflection (nose up) can be made without
exceeding the design limit load factor.
C LMAX WING FlAPS UP
3
2 -
1-
-
-
-
l
-
A
I
-
-
I A1
- 1- - -
I
I
I
-I-
Vc
I
O ~~~----~------~-------------+-------7---
-1
Figure 14.2a Design Manoeuvring Speed VA
When establishing VA the aeroplane is assumed to be flying in steady level flight, at point Al
in Fig. 14.2a, and the pitch control is suddenly moved to obtain extreme positive pitch
acceleration (nose up). V A is slower than the speed at the intersection of the C LMAX line and the
positive limit load factor line (point A) to safeguard the tail structure because of the higher load
on the tailplane during the pitch manoeuvre (Ref. paragraph 10.15).
Line OA in Fig. 14.2a represents the variation of stalling speed with load factor. Stalling speed
increases with the square root of the load factor, therefore;
For example an aircraft with a Ig stalling speed of 60kt and limit load factor of2'5 would have
a VA of:
60~ = 95 kt
14.8
EFFECT OF AIRCRAFT WEIGHT ON VA •
The Ig stalling speed depends on the weight of the aircraft. The line OA is drawn for the
maximum design weight, so for lower weights the stalling speed will be less.
For the same limit load factor VA will therefore decrease. For the example considered above,
if V A is 95 kt at 2500 lb weight, then at 2000 lb weight it will be:
95
NB:
~
2000
2500
=
85 kt
20% decrease in weight has given approximately 10% decrease in V A'
14 - 5
© Oxford Aviation Training
LIMITATIONS
PRINCIPLES OF FLIGHT
3
POSITIVE
CLmax
..-..
c
'-'
A
C
~
2
0:::
0
I-
0
1
«LL
0
«
0
E
0
SPEED
(EAS)
....J
-1
VA
NEGATIVE CL
max
Vc
~----------------------------- I
VD
Figure 14.1
14.9
(Repeat)
DESIGN CRUISING SPEED Vc
Point 'C' in Figure 14.1 (repeated above) is the design cruise speed V c. This is a speed selected
by the designer and used to assess the strength requirements in the cruise. Its value is determined
by the requirements JAR 25 .335 and JAR 23 .335. It must give adequate spacing from VB (Ref.
paragraph 14.15) and V D to allow for speed upsets. For example JAR 25 requires Vc to be at
least 43 kt above VB' and not greater than 0.8 V D' JAR 23 has similar requirements. V c need
not exceed the maximum speed in level flight at maximum continuous power (VH) or in JAR 23,
0.9 V H at sea level
14.10
DESIGN DIVE SPEED VD
Point 'D' in figure 14.1 is the design dive speed YD' This is the maximum speed which has to
be considered when assessing the strength of the aircraft. It is based on the principle of an upset
occurring when the aircraft is flying V c, resulting in a shallow dive, during which the speed
increases, until recovery is effected. If the resulting speed is not suitable because of buffet or
other high speed effects, a demonstrated speed may be used. This is called V DF the flight
demonstrated design dive speed.
14.11
NEGATIVE LOAD FACTORS.
In normal flying and manoeuvres it is not likely that very large negative ' g' forces will be
produced, however some negative 'g' forces may occur during manoeuvres and the aircraft must
be made strong enough to withstand them.
14 - 6
© Oxford Aviation Training
LIMITATIONS
PRINCIPLES OF FLIGHT
14.12
THE NEGATIVE STALL.
If the angle of attack of the wing is 'increased' in the negative direction, it wi II eventually reach
an angle at which it will stall. (If the wing section is symmetrical this angle will be the same
as the positive stall angle, but for a cambered wing, the angle and the negative C LMAX will usually
be lower). The line OR in Fig. 14.1 represents the negative C LM AX boundary. For large aircraft
a limit load factor of - 1 must be considered up to V co From Ve to V D the negative load factor
varies linearly from - 1 to o.
14.13
MANOEUVRE BOUNDARIES.
Taking into account the limiting values of positive and negative load factor, and the maximum
speed to be considered, the aircraft is therefore safe to operate within the boundaries shown in
Figure 14.3.
3
A
c
0:::
2
0
I-
0
<{
u..
L
1
0
E
<{
0
-J
0
SPEED
(EAS)
-1
Figure 14.3
Manoeuvre boundaries
Line SL represents level Ig flight. Line SA shows the load factors that could be produced by
pitching the wing to its stalling angle. Line ACD is the limit set by the maximum positive ' g'
which the airframe is required to withstand. Line OR shows the negative load factors that could
be produced with the wing at its negative stalling angle, and line RFE is the negative 'g' limit.
The design speeds (Vc and V D), already defined, are used for the purpose of assessing the
strength requirements of the aircraft in various flight conditions. These speeds are not scheduled
in the aircraft's Flight Manual, but the operational speed limits which are scheduled, are related
to them.
14 - 7
© Oxford Aviation Training
PRINCIPLES OF FLIGHT
14.14
LIMITATIONS
OPERATIONAL SPEED LIMITS.
The maximum airspeed at which an aircraft is permitted to fly is V MO for 'large aircraft' (JAR
25) or V NE for other aircraft (JAR 23) other than turbine engined aircraft. (For certification, a
large aircraft is defined as one of more then 5,700 kg Maximum Certificated Take-off Mass).
Maximum Operating Speed (Large Aircraft) V MO / MMo: V MO is a speed that may not be
deliberately exceeded in any regime offlight (climb, cruise or descent). V MO must not be greater
than V c and must be sufficiently below V D to make it highly improbable that V D will be
inadvertently exceeded in operations.
Because V MO is an Indicated Air Speed, as altitude increases the Mach number corresponding
to V MO will increase. There will be additional limitations on the aircraft because of
compressibility effects. In a climb V MO will be superceded by MMo (maximum operating Mach
number) at about 24,000 to 29,000 ft, depending on atmospheric conditions. There is an audible
warning to alert flight crews of an inadvertent overspeed beyond V MO / MMo,
When climbing at a constant lAS
it is possible to exceed MMO
When descending at a constant Mach number
it is possible to exceed V MO
Never Exceed Speed (Small Aircraft) V NE: V NE is set below V D to allow for speed upsets to
be recovered. (VNE = 0·9 V D)' V NE will be shown by a radial red line on the airspeed indicator
at the high speed end of the yellow arc.
Maximum Structural Cruise Speed (Small Aircraft) VNO: VNO is the normal operating cruise
speed limit and must be not greater than the lesser of V c or 0·89 V NE'
On the airspeed indicator V NO is the upper limit of the green arc.
From VNO to V NE there will be a yellow arc, which is the caution range. You may fly at
speeds within the yellow arc only in smooth air, and then only with caution.
14 - 8
© Oxford Aviation Training
LIMITATIONS
PRINCIPLES OF FLIGHT
14.15
GUST LOADS
The structural weight of an aircraft must be kept to a minimum while maintaining the required
strength. The following gust strengths were first formulated in the late 1940's and their
continued effectiveness has been verified by regular examination of actual flight data recorder
traces.
/"
+ 50 ft/sec
+ 66 ft/sec
../
./
GUST
LOAD
+ 25 ft/sec
FACTOR
1·0
o
.....
~---------;
SPEED
~---~.---"
EA.SL
-25 ft/sec
-66 ft/sec
- 50 ft/sec
-------
""
Vc
Figure 14.4
Aircraft are designed to be strong enough to withstand a 66 ft/sec vertical gust at VB (the design
speed for maximum gust intensity). If an aircraft experienced a 66 ft/sec vertical gust while
flying at VB, it would stall before exceeding the limit load factor. In turbulence an aircraft would
receive maximum protection from damage by flying at V B'
V B is quite a low airspeed and it would take some time for an aircraft to slow from Ve (the
design cruising speed) to V B if it flew into turbulence. Therefore, another design strength
requirement is for the aircraft also to be strong enough to withstand a vertical gust of 50 ft/sec
(EAS) at Ve.
Protection is also provided for the remote possibility of a vertical gust during a momentary upset
to a speed of V D (the design diving speed). The aircraft must also be strong enough to withstand
a vertical gust of 25 ftlsec at VD' (VB' Ve and V D are design speeds and are not quoted in an
aircraft's flight manual).
In practice, a slightly higher speed than VB is used for turbulence is penetration. This speed is
V RA/MRA (the rough-air speed). V RA I MRA will give adequate protection from over-stressing
the aircraft plus give maximum protection from an inadvertent stall.
14 - 9
© Oxford Aviation Training
PRINCIPLES OF FLIGHT
14.16
LIMITATIONS
EFFECT OF A VERTICAL GUST ON THE LOAD FACTOR
Vertical gusts will affect the load factor (n) by changing the angle of attack of the wing, Fig.
14.5.
INCREASE IN LIFT
(C L )
AIRCRAFT TAS,
V
INCREASE IN ANGLE
OF ATTACK
EFFECTIVE AIRFLOW
GUST
VELOC ITY
Figure 14.5
The following example illustrates the effect of a vertical gust on the load factor (n).
An aircraft is flying straight and level at a C L of 0-42. A 1 0 change in angle of attack will change
the C L by 0·1. If the aircraft is subject to a vertical gust which increases the angle of attack by
3 0 , what load factor will the aircraft experience?
Load factor =
LIFT
WEIGHT
In straight and level flight:
n
=
1
or
0.42
0.42
a 3° increase in angle of attack will give:
the CL will increase by 0.3:
n
=
0.72
0.42
=
3 x 0.1
= 0.3
0.42 + 0.3 = 0.72
1.7
a gust which increases the angle of attack by 3° will increase the load factor to 1.7
14 - 10
© Oxford Aviation Training
PRINCIPLES OF FLIGHT
LIMITATIONS
For a given gust speed and aircraft TAS, the increment in the load factor depends on the increase
in CL per change in angle of attack due to the gust (the slope of the lift curve). If the lift curve
has a steep slope, the 'g' increment will be greater. Factors which affect the lift curve are aspect
ratio and wing sweep.
HIGH ASPECT
RATIO
LOW ASPECT
RATIO
(or sweepback)
Figure 14.6
Wings having a low aspect ratio, or sweep, will have a lower lift curve slope, and so will give
a smaller increase in 'g' when meeting a given gust at a given TAS.
High wing loading reduces the 'g' increment in a gust. This is because the lift increment
produced is a smaller proportion of the original lift force for the more heavily loaded aircraft.
For a given aircraft the only variables for load factor increment in a gust are the aircraft T AS and
the gust speed.
14.17
EFFECT OF THE GUST ON STALLING
If anaerofoil encounters an upgust, it will experience an increase in angle of attack. For a given
gust velocity the increment in angle increases as the aircraft T AS decreases. If the angle of
attack is already large (low speed) the increment due to the gust could cause the wing to stall.
There is thus a minimum speed at which it is safe to fly if a gust is likely to be met, so as not to
stall in the gust.
14 - 11
© Oxford Aviation Training
LIMITATIONS
PRINCIPLES OF FLIGHT
14.18
OPERATIONAL ROUGH-AIR SPEED (VRA / M RA)
For flight in turbulence an air speed must chosen to give protection against two possibilities:
stalling and overstressing the aircraft structure. Turbulence is defined by a gust of a defined
value. If this defined gust is encountered the aircraft speed must be:
a)
high enough to avoid stalling and
b)
low enough to avoid damage to the structure.
These requirements are fulfilled by calculating the stall speed in the gust and then building in
sufficient strength for this speed.
The key is the chosen value of the gust, as this will dictate the strength required and therefore
the aircraft weight. The gust velocity is associated with the design speed, VB, and the vertical
value of the gust is 66 ft. per second. Encountering a gust before the pilot is able to slow the
aircraft, plus the possibility of hitting a gust if the aircraft is 'upset' at high speed must also be
taken into consideration. Because these probabilities are lower however, progressively lower
values of gust velocity are chosen at the higher speeds. These values are 50 ft. per second at the
design cruise speed V c and 25 ft. per second at the design dive speed V D.
The design gust values of 66, 50 and 25 ft. per second for gusts at the design speeds of V B, V c
and V D have existed since the early 1940's. In the UK they were established as a result of the
earliest "Flight Data Recorder" results. Modem flight recorder results and sophisticated design
analyses continue to support the original boundaries of the design gust envelope.
Generally, design for strength is based on calculating the increase in load on the aircraft as a
function of an instantaneous increase in angle of attack on the wing (page 14 - 11).
On large aircraft, additional allowances have to be made for several reasons:a)
The greater dynamic response due to increased structural flexibility.
b)
The possible implications of the smaller margin between actual cruise speed and design
cruise speed.
c)
The significance, in the more advanced designs, of the effects of build-up of gusts and
unsteady flow generally.
d)
The frequency of storm penetrations.
e)
The implications of the limited slow-down capabilities.
14 - 12
© Oxford Aviation Training
PRINCIPLES OF FLIGHT
LIMITATIONS
All design speeds, and design gust values, are EAS. But, remember: the increase in angle of
attack due to a gust is a function of the T AS of the aircraft and the T AS of the gust.
The choice of rough-air speed to be used operationally must be consistent with the strength of
the aircraft. At the same time the aircraft must comply with both minimum stability and control
criteria. There is also the important consideration of what maximum speed reduction can be
achieved in a slow-down technique. A typical chart ofthe speeds to which the rough air speed
is related, is shown below in Fig. 14.7. The illustration is drawn for a single (mid) weight. Line
AB is the 19 stall speed.
B
50
0
0
0
..- 40
x
L E
R-
I--
LL
LlJ
30
0
::J
I-I-....J
K
20
«
10
o
c
A
100
o
200
G
300
J
400
SPEED - KNOTS EAS
Figure 14.7
Line CE is the stall speed in a 66 ft. per sec. gust.
(This assumes the 66 ft. per sec. gust up to maximum altitude. Note that point E would represent
an extremely high true air speed gust value).
Line GHI is the VmolMmo line.
Line JKL is the VDF/MDF line.
Line MN is an example of a maximum strength speed line for a 66 ft. per sec. gust.
Line RS is the maximum altitude at which the aeroplane can sustain 1·5 g without too much
buffet.
14 - 13
© Oxford Aviation Training
LIMITATIONS
PRINCIPLES OF FLIGHT
At all speeds above the line CE the aeroplane will sustain a 66 fps gust without stalling and at
all speeds below the line MN the aeroplane is strong enough to withstand a 66 fps gust. The
rough-air speed therefore should lie somewhere between these two speeds, and the line OP gives
equal protection between accidentally stalling and overstressing the aircraft.
The line MN is a curious shape because different parts of the structure become critical at
different altitudes. This line is actually the lowest speed boundary of a collection of curves at
the higher speed end of the chart.
Because of the obvious attraction of a single speed at all altitudes up to that at which the roughair speed becomes a rough-air Mach number, the line could be adjusted slightly so as to avoid
any variations with altitude. As turbulence is generally completely random, this halfway speed
would give equal protection against the 50-50 probability of being forced too fast or too slow.
It has been stated that the diagram is drawn for a mid weight. The effect of weight change in
terms of the lower and upper limits to rough-air speed is, of course, significant, but selfcancelling. At low weights the stall line for a 66 ft. per sec. gust falls to lower speeds and the
maximum strength speed line increases to higher speeds. There is therefore no point in
attempting a sophisticated variation of V RA with weight.
The maximum altitude limit does, however, vary significantly with weight, and also varies for
the level of manoeuvre capability chosen. A 0·5 g increment to buffet is not too much protection
in severe turbulence. A lower altitude will therefore be required for a higher level of protection,
and, for a given level of protection, a lower altitude will be required for higher weights.
14.19
LANDING GEAR SPEED LIMITATIONS
The landing gear will normally be retracted as soon as possible after take-off, to reduce drag and
increase the climb gradient. There is no normal requirement for the gear to be operated at high
lAS so the retract and extend mechanism together with the attachment points to the structure are
sized for the required task. To design the gear for operation at high lAS would unnecessarily
increase structural weight.
VLO: the landing gear operating speed is the speed at which it is safe both to extend and to
retract the landing gear. If the extension speed is not the same as the retraction speed, the two
speeds must be designated as V LO (EXT) and V LO (~ET).
When the gear is retracted or extended the doors must open first. The doors merely streamline
the undercarriage bay and are not designed to take the aerodynamic loads which would be placed
on them at high lAS. Consequently V LO is usually lower than V LE.
VLE: the landing gear extended speed. There may be occasions when it is necessary to ferry the
aircraft with the gear down, and to do this a higher permissible speed would be convenient. V LE
is the speed at which it is safe to fly the aircraft with the landing gear secured in the fully
extended position. Because the undercarriage doors are closed, V LE is normally higher than V LO.
14 - 14
© Oxford Aviation Training
PRINCIPLES OF FLIGHT
14.20
LIMITATIONS
FLAP SPEED LIMIT
Flaps are designed to reduce take-off and landing distances and are used when airspeed is
relatively low. The flaps, operating mechanism and attachment points to the structure are not
designed to withstand the loads which would be applied at high airspeeds (dynamic pressure).
CL(ING
3
F~PS ~;:/~~ C L _
WING
F~PS
UP
2
/
_~
/
/
/
1
1
1----
1
I
1
VF
Vc
Vo
-1
Figure 14.8
Flaps increase C L MAX and decrease stall speed, so when flaps are deployed it is necessary to
provide additional protection to avoid exceeding the structural limit load. It can be seen from
the V -n diagram in Fig. 14.8 that it is possible for a greater load to be applied to the structure at
quite moderate airspeeds with flaps down. The limit load factor with flaps deployed is reduced
from 2·5 to 2 to give additional protection to the flaps and also the wing structure. If flaps are
deployed in turbulence a given vertical gust can generate a much larger lift force, which will
subject the structure to a larger load, possibly exceeding the ability of the structure to withstand
it and the structure could fail.
VFE: the Wing Flaps Extended Speed is the maximum airspeed at which the aircraft should be
flown with the flaps in a prescribed extended position. (Top of the white arc on the ASI).
Extending flaps for turbulence penetration in the cruise would reduce the stall speed and
increases the margin to stall, but the margin to structural limitations will be reduced by
a greater amount. Flaps must only be used as laid down in the aircraft Flight Manual.
14 - 15
© Oxford Aviation Training
PRINCIPLES OF FLIGHT
14.21
LIMITATIONS
AERO ELASTICITY (AERO - ELASTIC COUPLING)
Aerodynamic forces acting on the aircraft produce distortion of the structure, and this distortion
produces corresponding elastic forces in the structure ("winding up the spring"). Structural
distortion produces additional aerodynamic loading and this process is continued until either an
equilibrium condition is reached, or structural failure occurs.
This interaction between the aerodynamic loads and the elastic deformation of the airframe is
known as aeroelasticity, or aero-elastic coupling.
At low airspeeds, the aerodynamic forces are relatively small, and the resultant distortion of the
structure produces only negligible effects. At higher speeds, aerodynamic loads and the
consequent distortion are correspondingly greater. Aerodynamic force is proportional to y2, but
structural torsional stiffness remains constant. This relationship implies that at some high
speed, the aerodynamic force buildup may overpower the resisting torsional stiffness and
'divergence' will occur. The aircraft must be designed so the speed at which divergence occurs
is higher than the design speeds Y oIM D •
Definitions:
Elasticity: No structure is perfectly rigid. The structure of an aircraft is designed to be
as light as possible. This results in the aircraft being a fairly flexible structure, the
amount of flexibility depending on the design configuration of the aircraft. e.g. aspect
ratio, sweepback, taper ratio. etc.
Backlash: The possibility of movement of the control surface without any movement
of the pilof s controls.
Mass distribution: The position of the CG ofa surface in relation to its torsional axis.
Mass balance: A mass located to change the position of the CG ofa surface in relation
to its torsional axis.
Divergence: The structure will continue to distort until it breaks.
Flutter: The rapid and uncontrolled oscillation of a surface resulting from imbalance.
Flutter normally leads to a catastrophic failure of the structure.
14 - 16
© Oxford Aviation Training
PRINCIPLES OF FLIGHT
LIMITATIONS
t
4
3
2
_~
_ _-+-_
C:::=
-III---.~1-~
AC -
Jl
L
FLEXURAL
-
AXIS
Figure 14.9
Refer to Fig. 14.9 which represents the view of a wingtip, and consider a vertical gust increasing
the angle of attack of the wing. The additional lift force will bend the wing tip upwards from
position 1 to 2 and the increase in lift acting through the AC, which is forward of the flexural
axis, will twist the wing tip nose up; this increases the angle of attack further. The wing tip will
rapidly progress to position 3 and 4. The wing is being wound up like a spring and can break
if distorted too much.
How far the structure is distorted depends on:
a)
the flexibility of the structure
b)
the distance between the AC and the flexural axis and
c)
the dynamic pressure (lAS).
Methods of delaying divergence to a higher speed:
a)
The structure can be made stiffer, but this will increase weight.
b)
A better solution is to move the flexural axis closer to the AC. This can easily be
accomplished by mounting a mass forward of the AC. Instead of using a large piece of
lead, as in control surface mass balance, the engines can be mounted forward of the
leading edge and this will move the flexural axis closer to the AC. (also see flutter,
paragraph 14.23).
14 - 17
© Oxford Aviation Training
PRINCIPLES OF FLIGHT
LIMITATIONS
TRAILING EDGE
WING ROOT
LEADING EDGE
c
u
Figure 14.10
Typical flutter mode
14 - 18
© Oxford Aviation Training
PRINCIPLES OF FLIGHT
14.22
LIMITATIONS
FLUTTER.
Flutter involves:
a)
aerodynamic forces
b)
inertia forces and
c)
the elastic properties of a surface.
The distribution of mass and stiffness in a structure determine certain natural frequencies and
modes of vibration. If the structure is subject to a 'forcing' frequency near these natural
frequencies, a resonant condition can result giving an unstable oscillation which can rapidly lead
to destruction.
An aircraft is subject to many aerodynamic excitations (gusts, control inputs, etc.) and the
aerodynamic forces at various speeds have characteristic properties for rate of change of force
and moment. The aerodynamic forces may interact with the structure and may excite (or
negatively damp) the natural modes of the structure and allow flutter. Flutter must not occur
within the normal flight operating envelope and the natural modes must be damped if possible
or designed to occur beyond V o/M o. A typical flutter mode is illustrated in Fig. 14.10.
Since the problem is one of high speed flight, it is generally desirable to have very high natural
frequencies and flutter speeds well above the normal operating speeds. Any change of stiffness
or mass distribution will alter the modes and frequencies and thus allow a change in the flutter
speeds. If the aircraft is not properly maintained and excessive play and flexibility (backlash)
exist, flutter could occur at flight speeds well below the operational limit speed (V MO I MMO) '
Wing flutter can be delayed to a higher speed, for a given structural stiffness (weight), by
mounting the engines on pylons beneath the wing forward of the leading edge, Fig. 14.11. The
engines act as 'mass balance' for the wing by moving the flexural axis forward, closer to the AC.
AC
Figure 14.11
FLEXURAL AXIS
MOVED FORWARD
Wing mass balanced by podded engines
14 - 19
© Oxford Aviation Training
PRINCIPLES OF FLIGHT
14.23
LIMITATIONS
CONTROL SURFACE FLUTTER
Control surface flutter can develop as a result of an oscillation of the control surface coupled
with an oscillation in bending or twisting of the wing, tailplane or fin. A control surface
oscillation can result from backlash (free play) in the control system, or from a disturbance
(gust). Flutter can develop if the CG of the control surface is behind the hinge line, so that the
inertia of the control surface causes a moment around the hinge.
Torsional Aileron Flutter: Figure 14.12 illustrates the sequence for a half cycle, which is
described below.
1.
The aileron is displaced downwards, exerting an upwards force on the aileron hinge.
2.
The wing twists about the torsional axis, the trailing edge rising, taking the aileron hinge
up with it, but the aileron surface lags behind due to the CG being aft of the hinge line.
3.
The inherent stiffness of the wing has arrested the twisting motion (the spring is now
wound up), but the air loads on the aileron, the stretch of the control circuit, and its
upwards momentum, cause the aileron to 'flick' upwards, placing a down load on the
trailing edge of the wing.
4.
The energy stored in the twisted wing and the reversed aerodynamic load of the aileron
cause the wing to twist in the opposite direction. The cycle is then repeated.
Torsional aileron flutter can be prevented either by mass balancing the ailerons with attachment
of a mass ahead of the hinge line to bring the CG onto, or slightly ahead of the hinge line, or by
making the controls irreversible (fully powered controls with no manual reversion).
Flexural Aileron Flutter: is generally similar, but is caused by the movement of the aileron
lagging behind the rise and fall of the outer portion of the wing as it flexes (wing tips up and
down), thus tending to increase the oscillation. This type of flutter can also be prevented by
mass balancing the ailerons. The positioning of the mass balance 'weight' is important, the
nearer the wing tip the smaller the mass required. On many aircraft the mass is distributed along
the whole length of the aileron in the form of a leading edge 'spar', thus increasing the stiffness
of the aileron and preventing a concentrated mass starting torsional vibrations in the aileron
itself.
Mass balancing must also be applied to elevators and rudders to prevent their inertia and the
'springiness' of the fuselage starting similar motions. Mass balancing may even be applied to
tabs.
The danger of all forms of flutter is that the speed and amplitude of each cycle is greater than
its predecessor, so that in a second or two the structure may be bent beyond its elastic limit and
fail. Decreasing speed if flutter is detected is theoretically the only means of preventing
structural failure, but the rate of divergence is so rapid that slowing down is not really a practical
solution.
14 - 20
© Oxford Aviation Training
PRINCIPLES OF FLIGHT
LIMITATIONS
HINGE LINE
TORSIONAL AXIS
CG
1
2
3
4
Figure 14.12
Torsional aileron flutter
14 - 21
© Oxford Aviation Training
LIMITATIONS
PRINCIPLES OF FLIGHT
14.24
AILERON REVERSAL
Figure 14.13
Low speed aileron reversal
LOW SPEED: It was described in paragraph 7.6 that if an aileron is lowered when flying at
high angles of attack, that wing could possibly stall, Fig. 14.13. In that case the wing will drop
instead of rising as intended. Hence the tenn low speed aileron reversal.
-- E
------- T======ELASTIC WING
FLEXURAL
AXIS
Figure 14.14
HIGH SPEED: Aileron reversal can also occur at high speed when the wing twists as a result
of the loads caused by operating the ailerons. In Figure 14.14 the aileron has been deflected
downwards to increase lift and raise the wing. Aerodynamic forces act upwards on the aileron,
and as this is behind the flexural axis of the wing, it will cause a nose down twisting moment on
the wing structure. This will reduce the angle of attack of the wing which will reduce its lift.
If the twisting is sufficient, the loss of lift due to decreased angle of attack will exceed the gain
of lift due to increased camber, and the wing will drop instead of lifting.
14 - 22
© Oxford Aviation Training
PRINCIPLES OF FLIGHT
LIMITATIONS
SPOILER SURFACES
OUTBOARD AILERONS
(LOW SPEED ONLY)
~
INBOARD AILERONS -- /
(HIGH SPEED AND LOW SPEED)
Figure 14.15
Inboard and outboard ailerons and roll control spoilers
High speed aileron reversal can be delayed to a speed higher than V D / MDby having inboard and
outboard ailerons and/or roll control spoilers. The inboard ailerons, Fig. 14.l5, are mounted
where the wing structure is naturally stiffer, and work at all speeds. The outboard ailerons work
only at low speed, being deactivated when the flaps are retracted.
On most high speed j et transport aircraft roll control spoilers assist the ailerons. Because they
are mounted further forward and on a stiffer part of the wing, roll control spoilers do not distort
the wing structure to the same degree as ailerons.
14 - 23
© Oxford Aviation Training
PRINCIPLES OF FLIGHT
LIMITATIONS
SELF ASSESSMENT QUESTIONS
1.
If an aircraft is flown at its design manoeuvring speed V A :
a)
b)
c)
d)
2.
The speed VNE is:
a)
b)
c)
d)
3.
normal operations.
abrupt manoeuvres.
flight in smooth air.
flight in rough air.
The maximum allowable airspeed with flaps extended ( V FE
because:
a)
b)
c)
d)
5.
the airspeed which must not be exceeded except in a dive
the maximum airspeed at which manoeuvres approaching the stall may be carried out
the maximum airspeed at which an aircraft may be flown
the maximum speed, above which flaps should not be extended
Maximum structural cruising speed VNO is the maximum speed at which an aeroplane can be
operated during:
a)
b)
c)
d)
4.
it is possible to subject the aircraft to a load greater than its limit load during high 'g'
manoeuvres.
it is only possible to subject the aircraft to a load greater than its limit load during
violent increases in incidence, i.e. when using excessive stick force to pull-out ofa dive.
it is not possible to exceed the limit load.
it is possible to subject the aircraft to a load greater than its limit load at high TAS.
)
is lower than cruising speed
they are used only when preparing to land
the additional lift and drag created would overload the wing and flap structure at higher
speeds
flaps will stall if they are deployed at too high an airspeed
too much drag is induced
Why is V L E greater than V LO on the maj ority of large jet transport aircraft?
a)
b)
c)
d)
VLO is used when the aircraft is taking off and landing when the lAS is low.
extending the gear at too high an airspeed would cause excessive parasite drag.
flying at too high an airspeed with the gear down would prevent retraction of the
forward retracting nose gear.
VLO is a lower lAS because the undercarriage doors are vulnerable to aerodynamic loads
when the gear is in transit, up or down.
14 - 25
© Oxford Aviation Training
PRINCIPLES OF FLIGHT
6.
The phenomenon of flutter is described as:
a)
b)
c)
d)
7.
d)
d)
the down-going aileron increasing the semi-span angle of attack beyond the critical.
flow separation ahead of the aileron leading edge.
uneven shock wave formation on the top and bottom surface of the aileron, with the
attendant movement in control surface CP, causing the resultant force to act in the
opposite direction from that intended.
dynamic pressure acting on the aileron twisting the wing in the opposite direction,
possibly causing the aircraft to bank in a direction opposite to that intended.
Controls are mass balanced in order to:
a)
b)
c)
d)
10.
They give increased ground clearance in roll.
They give improved longitudinal mass distribution.
The wing structure can be lighter because the engine acts as a mass balance and also
relieves wing bending stress.
They enable a longer undercarriage to be used which gives an optimum pitch attitude
for take-off and landing.
Aileron reversal at high dynamic pressures is caused by:
a)
b)
c)
9.
rapid oscillatory motion involving only rotation of the control surfaces, associated with
the shock waves produced around the control surfaces.
oscillatory motion of part or parts of the aircraft relative to the remainder of the
structure.
rapid movement of the airframe caused by vibration from the engines.
reversal of the ailerons caused by wing torsional flexibility.
What is the purpose of fitting the engines to an aircraft using wing mounted pylons?
a)
b)
c)
8.
LIMITATIONS
eliminate control flutter.
aerodynamically assist the pilot in moving the controls.
provide equal control forces on all three controls.
return the control surface to neutral when the controls are released.
If an aircraft weight is reduced by 15%, VA win:
a)
b)
c)
d)
Not change.
Increase by 15%.
Increase by 7.5%.
Decrease by 7.5%.
14 - 26
© Oxford Aviation Training
LIMITATIONS
PRINCIPLES OF FLIGHT
11.
Which of the following statements is correct?
1 -
It is a design requirement that control reversal speeds must be higher than any speed to
be achieved in flight.
2
-
3
4
12.
Flying control surfaces are aerodynamically balanced to prevent flutter.
-
An aircraft is not a rigid structure.
5
Aeroelasticity affects are inversely proportional to lAS.
6 -
Control reversal speed is higher if the aircraft is fitted with outboard ailerons which are
locked-out as the aircraft accelerates; the inboard ailerons alone controlling the aircraft
in roll at higher speeds.
a)
b)
c)
d)
All the above statements are correct.
1,2,3and6.
1,2,4and6.
1,3,5 and 6.
V Lois defined as:
a)
b)
c)
d)
13.
The airframe must be made strong and stiff enough to ensure that the wing torsional
divergence speed is higher, by a substantial safety margin, than any speed which will
ever be achieved in any condition in flight.
maximum landing gear operating speed.
maximum landing gear extended speed.
maximum leading edge flaps extended speed.
maximum flap speed.
If flutter is experienced during flight, the preferable action would be:
a)
b)
c)
d)
immediately increase speed beyond V MO / MMO' by sacrificing altitude if necessary.
immediately close the throttles, deploy the speed brakes and bank the aircraft.
rapidly pitch-up to slow the aircraft as quickly as possible.
reduce speed immediately by closing the throttles, but avoid rapid changes in attitude
and/or configuration.
14 - 27
© Oxford Aviation Training
LIMITATIONS
PRINCIPLES OF FLIGHT
ANSWERS
I No I A I B I c I D II
1
C
2
C
3
B
D
5
B
6
7
C
D
8
A
D
10
11
12
13
I
A
4
9
REF
C
A
D
14 - 29
© Oxford Aviation Training
CHAPTER 15 - WINDSHEAR
Contents
Page
INTRODUCTION ...................................................... 15 - 1
MICROBURST
WINDSHEAR ENCOUNTER DURING APPROACH .......................... 15 - 3
EFFECTS OF WINDSHEAR .............................................. 15 - 4
"ENERGY GAIN" DUE TO INCREASE IN HEADWIND
"ENERGY LOSS" DUE TO DOWNDRAUGHT ........................ 15 - 5
"ENERGY LOSS" DUE TO LOSS OF HEADWIND
"TYPICAL" RECOVERY FROM WINDSHEAR .............................. 15 - 6
WINDSHEAR REPORTING .............................................. 15 - 7
VISUAL CLUES
CONCLUSIONS ........................................................ 15 - 8
SELF ASSESSMENT QUESTIONS ........................................ 15 - 9
ANSWERS .................................................... 15 - 15
PRINCIPLES OF FLIGHT
15.1
WINDSHEAR
INTRODUCTION (Ref: AIC 33/1997)
Windshear is a sudden drastic shift in wind speed and/or direction that occurs over a short
distance at any altitude in a vertical and/or horizontal plane. It can subject an aircraft to sudden
updraughts, downdraughts, or extreme horizontal wind components, causing sudden loss of lift
or violent changes in vertical speeds or altitudes. Windshear will cause abrupt displacement
from the flight path and require substantial control action to counteract it.
A windshear encounter is a very dynamic event which can strike suddenly and with devastating
effect which has been beyond the recovery powers of experienced pilots flying modem and
powerful aircraft. An encounter may cause alarm, a damaged undercarriage, or a total
catastrophe. The first and most vital defence is avoidance.
The most powerful examples of windshear are associated with thunderstorms (cumulonimbus
clouds), but windshear can also be experienced in association with other meteorological features
such as the passage of a front, or a marked low-level temperature inversion. The meteorological
features of windshear will be dealt with fully elsewhere.
15.2
MICROBURST
Microbursts are associated with thunderstorms and are one of the most dangerous sources of
windshear. Microbursts are small-scale intense downdraughts which, on reaching the surface,
spread outward in all directions from the downdraught centre. This causes the presence of both
vertical and horizontal wind shear that can be extremely hazardous to all types and sizes of
aircraft, especially when within 1,000 feet of the ground.
A microburst downdraught is typically less than 1 mile in diameter as it descends from the cloud
base to about 1,000 to 3,000 feet above the ground. In the transition zone near the ground, the
downdraught changes to a horizontal outflow that can extend to approximately 2Y2 miles (4 km)
in diameter.
a)
Downdraughts can be as strong as 6,000 feet per minute.
b)
Horizontal winds near the surface can be as strong as 45 knots resulting in a 90
knot shear as the wind changes to or from a headwind across the microburst.
c)
These strong horizontal winds occur within a few hundred feet of the ground. An
individual microburst seldom lasts longer than 15 minutes from the time it strikes
the ground until dissipation.
These are maximum values but they do indicate how it is possible for large and powerful aircraft
to become uncontrollable when they meet such examples of the micro burst.
15 - 1
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
WINDSHEAR
A microburst intensifies for about 5 minutes after it first strikes the ground, with the maximum
intensity winds lasting approximately 2 to 4 minutes. Sometimes microbursts are concentrated
into a line structure and, under these conditions, activity may continue for as long as an hour.
Once micro burst activity starts, multiple microbursts in the same general area are not uncommon
and should be expected.
STRONG DOWNDRAUGHT
Increasing Headwind
Increasing Tailwind
tn
¢
"'IIIIIIIII_ _ _ _"
Outflow
Figure 15.1
A microburst encounter during take-off
During takeoff into a microburst, shown in Fig. 15.1, an aircraft first experiences a headwind
which increases performance without a change in pitch and power (1).
This is followed by a decreasing headwind and performance, and a strong downdraft (2).
Performance continues to deteriorate as the wind shears to a tailwind in the downdraft (3).
The most severe downdraft will be encountered between positions 2 and 3, which may result in
an uncontrollable descent and impact with the ground (4).
15 - 2
© Oxford Aviation Services Limited
WINDSHEAR
PRINCIPLES OF FLIGHT
15.3
WINDSHEAR ENCOUNTER DURING APPROACH
The Power setting and vertical velocity required to maintain the glide slope should be closely
monitored. If any windshear is encountered, it may be difficult to stay on the glide path at
normal power and descent rates. If there is ever any doubt that you can regain a reasonable
rate of descent, and land without abnormal manoeuvres, you should apply full power and
go-around or make a missed approach.
Windshear can vary enormously in its impact and effect. Clearly some shears will be more
severe and consequently more dangerous than others.
When countering the effects of winds hear, it is best to assume 'worse case'. It is impossible to
predict at the first stages of a windshear encounter how severe it will be, and it is good
advice to suggest that recovery action should anticipate the worst.
WIND SHEAR
From
To
Headwind Calm or Tailwind
From
Tailwind
To
Calm or Headwind
INDICATIONS
Indicated Airspeed
Decrease
Increase
Pitch Attitude
Decrease
Increase
Tends to Sink
Balloons
Increases
Decreases
Increase
Decrease
Up to Glideslope
Down to Glideslope
Reduce Power
Increase Power
Aircraft
Groundspeed
ACTIONS
Power
Fly
Be prepared to
To Stay on Glide Path
Figure 15.2
Increase Rate of Descent
(Due to faster groundspeed)
Decrease Rate of Descent
(Due to slower groundspeed)
Indications and recovery actions for windshear encounter during approach
Referring to Fig. 15.2, this table gives guidance should you encounter wind shear during a
stabilized landing approach. Approaches should never be attempted into known wind shear
conditions.
15 - 3
© Oxford Aviation Services Limited
WINDSHEAR
PRINCIPLES OF FLIGHT
15.4
EFFECTS OF WINDSHEAR
The relationship of an aeroplane in a moving air mass to its two reference points must be fully
understood. One reference is the air mass itself and the other is the ground.
On passing through a shear line, the change of airspeed will be sudden, but the inertia of the
aircraft will at first keep it at its original ground speed. The wind is a form of energy and when
it shears, an equivalent amount of energy is lost or gained.
a)
A rapid increase in headwind (or loss of tailwind) are both 'energy gains', and will
temporarily improve performance, Fig. 15.3.
b)
Down-draughts or a sudden drop of headwind (or increase in tailwind) are the main
danger at low altitude because they give an 'energy loss', Fig. 15.4 and 15.5 .
I
'ENERGY GAIN' - Rapid increase in headwind .
~~~ 60 kt
I
10 kt
•••
Vertical Speed :
Ground Speed:
lAS :
Vertical Speed:
Ground Speed :
lAS :
200 ft/m in R.O.C .
130 kt
190 kt
700 ft/min R.O.D.
130 kt
140 kt
GLIDE SLOPE
SHEAR
LINE
\
. . . . . . . . . . . .iawu. . . . ::.S
. . . . . .IH•...::
. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ..
Figure 15.3
"Energy Gain" due to increase in headwind
15 - 4
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
WINDSHEAR
! 'ENERGY LOSS' - Effect of downdraught. !
10 kt
•••
Vertical Speed :
Ground Speed:
lAS :
1500 ft/min R.O .D.
130 kt
GLIDE
110 kt
~. . .
Vertical Speed :
Ground Speed:
lAS :
SLOPE
700 ftlmin R.O.D.
130 kt
140 kt
SHEAR
LINE
I
Figure 15.4
"Energy Loss" due to downdraught
! 'ENERGY LOSS' - Loss of headwind. !
10kt
•••
Vertical Speed :
Ground Speed:
lAS:
~.~
.
'---=-
Vertical Speed :
Ground Speed:
lAS:
1000 ft/min R.O.D.
130 kt
110 kt
700 ftlmin R.O.D.
130 kt
140 kt
GLIDE SLOPE
20 kt
.........
SHEAR
LINE
\
... £&£~~~~~mmmm"~~~"Em"""~
Figure 15.5
"Energy Loss" due to loss of headwind
15 - 5
© Oxford Aviation Services Limited
WINDSHEAR
PRINCIPLES OF FLIGHT
15.5
"TYPICAL" RECOVERY FROM WINDSHEAR
The combination of increasing headwind, followed by down-draught, followed by increasing
tailwind should be considered, as this is the sequence which might be encountered in a
micro burst on the approach, or following take-off.
a)
The presence of thunderstorms should be known and obvious, so the increase in speed
caused by the rising headwind should be seen as the forerunner of a down-burst or
microburst; any hope of a stabilised approach should be abandoned and a missed
approach carried out as the only safe course of action.
b)
The initial rise in airspeed and rise above the approach path (balloon) should be seen as
a bonus and capitalised. Without hesitation, increase to go-around power, being
prepared to go to maximum power if necessary, select a pitch angle consistent with
a missed approach, typically about 15 0 and hold it against turbulence and buffeting.
c)
The next phase may well see the initial advantages of increased airspeed and rate of
climb being rapidly eroded. The down-draught now strikes, airspeed may be lost and
the aircraft may start to descend, despite the high power and pitch angle. It will be
impossible to gauge the true angle of attack, so there is a possibility that the stick shaker
(if fitted) may be triggered; only then should the attempt to hold the pitch angle
normally be relaxed.
d)
the point at which a down-draught begins to change to increasing tailwind may well be
the most critical period. The rate of descent may lessen, but the airspeed may still
continue to fall; the height loss may have cut seriously into ground obstacle clearance
margins. Given that maximum thrust is already applied, as an extreme measure if
the risk of striking the ground or an obstacle still exists, it may be necessary to
increase the pitch angle further and deliberately raise the nose until stick shaker
is feit, decrease back-pressure on the pitch control to try and hold this higher pitch
angle, until the situation eases with the aircraft beginning to escape from the effects of
the microburst.
When there is an indefinite risk of shear, it may be possible to use a longer runway, or one that
points away from an area of potential threat. It may also be an option to rotate at a slightly
higher speed, provided this does not cause undue-tyre stress or any handling problems. The high
power setting and high pitch angle after rotate, already put the aircraft into a good configuration
should a microburst then be encountered. The aircraft is however very low, there is little safety
margin and the ride can be rough. If there is still extra power available, it should be used
without hesitation. Ignore noise abatement procedures and maintain the high pitch angle,
watching out for stick shaker indications as a signal to decrease back-pressure on the pitch
control.
15 - 6
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
WINDSHEAR
In both approach and take-off cases, Vital Actions are:
15.6
a)
Use the maximum power available as soon as possible.
b)
Adopt a pitch angle of around 150 and try and hold that attitude. Do not chase
airspeed.
c)
Be guided by stick shaker indications when holding or increasing pitch attitude,
easing the back pressure as required to attain and hold a slightly lower attitude.
WINDSHEAR REPORTING
If you encounter a wind shear on an approach or departure, you are urged to promptly report it
to the controller. An advanced warning of this information can assist other pilots in avoiding
or coping with a wind shear on approach or departure. The recommended method for wind shear
reporting is to state the loss or gain of airspeed and the altitudes at which it was encountered.
If you are unable to report wind shear in specific terms, you are encouraged to make reports in
terms of the effect upon your aircraft.
15.7
VISUAL CLUES
You can see thunderstorms and hence receive a mental trigger to 'think windshear'. Once
alerted, lookout for tell-tale signs such as:
a)
Divergent wind sleeves or smoke;
b)
Strong shafts of rain or hail, also 'virga'; (intense precipitation which falls in shafts
below a cumulonimbus cloud and evaporates in the dry air beneath);
c)
Divergent wind patterns indicated by grass, crops or trees being beaten down or lashed;
d)
Rising dust or sand.
To observe and recognise any of the above will suggest that windshear danger is very close, if
not imminent; nevertheless a few seconds of advance warning may make all the difference, if
the warning is heeded and those seconds put to good use.
15 - 7
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WINDSHEAR
PRINCIPLES OF FLIGHT
15.8
CONCLUSIONS
Most pilots will experience windshear in some form or other; for most it may be no more than
a very firm landing or a swing on take-off or landing requiring momentary use of, perhaps, full
rudder for correction; they will probably put it down to 'gusts'. Some few pilots will experience
more authentic examples of windshear which will stretch their skills to the limit. A very small
number may find their skills inadequate. There is no sure way of knowing in advance the
severity of windshear which will be encountered, so it is better not to put one's skills to the test,
rather than find them inadequate. Windshear, particularly when linked with thunderstorms, has
caused disaster in the past and may well cause disaster again, but it will not harm those who
understand its power and have the good sense to avoid it. An inadvertent encounter on the
approach is most likely to de-stabilise it to such an extent that a missed approach is the
only safe course and the sooner that decision is made, the safer it is likely to be. Other
encounters must be treated on their merits, but any hint of 'energy loss' should be met with a
firm and positive response in line with the guidance put forward.
Recognise -
that windshear is a hazard.
and
Recognise Avoid
the signs which may indicate its presence.
- windshear by delay or diversion.
Prepare
for the inadvertent encounter by a speed 'margin' if' energy loss'
windshear is suspected.
Recover
know the techniques recommended for your aircraft and use them
without hesitation if windshear is encountered.
15 - 8
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
WINDSHEAR
SELF ASSESSMENT QUESTIONS
1.
Take-off EPR is being delivered by all engines and the take-off is proceeding normally, the
undercarriage has just retracted. Which initial indications may be observed when a headwind
shears to a down draught?
a)
b)
c)
d)
2.
80 knots.
40 knots.
90 knots.
45 knots.
What is the expected duration of an individual micro burst?
a)
b)
c)
d)
5.
6,000 ft/min.
7,000 ft/min.
8,000 ft/min.
10,000 ft/min.
An aircraft that encounters a headwind of 45 knots, within a micro burst, may expect a total shear
across the micro burst of
a)
b)
c)
d)
4.
constant. Vertical Speed: decreases. Pitch Attitude: decreases.
increases. Vertical Speed: decreases. Pitch Attitude: constant.
decreases. Vertical Speed: constant. Pitch Attitude: constant.
decreases. Vertical Speed: decreases. Pitch Attitude: decreases.
Maximum downdrafts in a microburst encounter may be as strong as
a)
b)
c)
d)
3.
Indicated Air Speed:
Indicated Air Speed:
Indicated Air Speed:
Indicated Air Speed:
Two minutes with maximum winds lasting approximately 1 minute.
Seldom longer than 15 minutes from the time the burst strikes the ground until
dissipation.
One micro burst may continue for as long as 2 to 4 hours.
For as long as 1 hour.
Which wind-shear condition results in a loss of airspeed?
a)
b)
c)
d)
Decreasing headwind or tailwind.
Increasing headwind and decreasing tailwind.
Decreasing headwind and increasing tailwind.
Increasing headwind or tailwind.
15 - 9
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
6.
Which performance characteristics should be recognized during takeoff when encountering a
tailwind shear that increases in intensity?
a)
b)
c)
d)
7.
Tailwind which suddenly increases in velocity.
Sudden decrease in a headwind component.
Sudden increase in a headwind component.
Calm wind which suddenly shears to a tailwind.
Which INITIAL cockpit indications should a pilot be aware of when a constant tailwind shears
to a calm wind?
a)
b)
c)
d)
9.
Loss of, or diminished climb ability.
Increased climb performance immediately after takeoff.
Decreased takeoff distance.
Improved ability to climb.
Which condition would INITIALLY cause the indicated airspeed and pitch to increase and the
sink rate to decrease?
a)
b)
c)
d)
8.
WINDSHEAR
Altitude increases; pitch and indicated airspeed decrease.
Altitude, pitch, and indicated airspeed increase.
Altitude, pitch, and indicated airspeed decrease.
Altitude decreases; pitch and indicated airspeed increase.
What is the recommended technique to counter the loss of airspeed and resultant lift from wind
shear?
a)
b)
c)
d)
Maintain, or increase, pitch attitude and accept the lower-than-normal airspeed
indications.
Lower the pitch attitude and regain lost airspeed.
Avoid overstressing the aircraft, pitch to stick shaker, and apply maximum power.
Accelerate the aircraft to prevent a stall by sacrificing altitude.
15 - 10
© Oxford Aviation Services Limited
WINDSHEAR
PRINCIPLES OF FLIGHT
10.
Which of the following would be acceptable techniques to minimise the effects ofa windshear
encounter?
2 3
4
5
6
-
a)
b)
c)
d)
11.
To prevent damage to the engines, avoid the use of maximum available thrust.
Increase the pitch angle until the stick shaker activates, then decrease back pressure to
maintain that angle of pitch.
maintain a constant airspeed.
Use maximum power available as soon as possible.
Keep to noise abatement procedures.
Wait until the situation resolves itself before taking any action.
1, 3, 5 and 6
2,3 and 5
2, 3, 4, 5 and 6
2 and 4
Which of the following statements about windshear is true?
1 2
-
3 -
4
5
a)
b)
c)
d)
-
Windshear can subject your aircraft to sudden up-draughts, down draughts, or extreme
horizontal wind components.
Windshear will cause abrupt displacement from the flight path and require substantial
control action to counteract it.
Windshear only affects small single and twin engine aircraft. Large, modem, powerful,
fast gas turbine engine powered aircraft will not suffer from the worst affects of a
microburst.
Microbursts are associated with cumulonimbus clouds.
Windshear can strike suddenly and with devastating effect which has been beyond the
recovery powers of experienced pilots flying modem and powerful aircraft.
1, 2, 3, 4 and 5
1,2 and 4
1, 2, 4 and 5
2,3,4 and 5
15 - 11
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
12.
A microburst is one of the most dangerous sources of winds hear associated with thunderstorms.
They are:
a)
b)
c)
d)
13.
small-scale intense up-draughts, which suck warm moist air into the cumulonimbus
cloud.
small-scale shafts of violent rain, which can cause severe problems to gas turbine
engmes.
large-scale, violent air, associated with air descending from the 'anvil' of a thunder
cloud.
small-scale (typically less than 1 mile in diameter) intense down-draughts which, on
reaching the surface, spread outward in all directions from the down-draught centre.
Thrust is being managed to maintain desired indicated airspeed and the glide slope is being
flown. Which of the following is the recommended procedure when you observe a 30 kt loss of
airspeed and the descent rate increases from 750 ftlmin to 2,000 ft/min?
a)
b)
c)
d)
14.
WINDSHEAR
Increase power to regain lost airspeed and pitch-up to regain the glide slope - continue
the approach and continue to monitor your flight instruments.
Decrease the pitch attitude to regain airspeed and then fly-up to regain the glide slope.
Apply full power and execute a go-around; report windshear to A TC as soon as
practicable.
Wait until the airspeed stabilises and the rate of descent decreases, because microbursts
are quite small and you will soon fly out of it.
Which of the following statements are correct?
1
2
3
4
5
A rapid increase in headwind is an 'energy gain' .
A rapid loss of tailwind is an 'energy gain' .
A shear from a tailwind to calm is an 'energy gain'.
A shear from calm to a headwind is an 'energy gain'.
A shear from headwind to calm is an 'energy loss' .
a)
b)
c)
d)
1,2 and 4
1, 2, 3,4 and 5
1,4 and 5
4 and 5 only
15 - 12
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
15.
Which of the following statements are correct?
2
3
4
5
-
a)
b)
c)
d)
16.
A downdraught is an 'energy gain'.
A rapid loss of tailwind is an 'energy loss' .
A shear from a tailwind to calm is an 'energy loss' .
A shear from calm to a headwind is an 'energy gain' .
A down draught is an 'energy loss' .
1,3 and 4
1, 2, 3 and 5
1,4 and 5
4 and 5 only
Which of the following sequences might be encountered when flying into a microburst?
a)
b)
c)
d)
17.
WINDSHEAR
Increased headwind, followed by down-draught, followed by increased tailwind on the
approach, or following take-off.
Increased headwind, followed by down-draught, followed by increased tailwind on the
approach. Increased tailwind, followed by down-draught, followed by increased
headwind following take-off.
Increased headwind, followed by down-draught, followed by increased tailwind on takeoff. Increased tailwind, followed by down-draught, followed by increased headwind on
the approach.
Increased tailwind, followed by down-draught, followed by increased headwind on takeoff. Increased headwind, followed by down-draught, followed by increased tailwind on
the approach.
Which of the following statements is correct when considering windshear?
2 3
4 5 -
a)
b)
c)
d)
Recognise that windshear is a hazard to all sizes and types of aircraft.
Recognise the signs which may indicate its presence.
Avoid windshear by delaying departure or by diverting if airborne.
Prepare for the inadvertent encounter by a speed 'margin' if 'energy loss' windshear is
suspected.
Know the techniques for recovery recommended for your aircraft and use them without
any hesitation if windshear is encountered.
2,4 and 5
3,4 and 5
1,2, and 5
1,2,3,4and5
15 - 13
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
WINDSHEAR
ANSWERS
I No I A I B I c I D I
1
2
A
C
4
B
5
C
A
7
C
8
B
9
C
10
D
11
C
12
D
13
C
14
B
15
16
17
I
D
3
6
REF
D
A
D
15 - 15
© Oxford Aviation Services Limited
CHAPTER 16 - PROPELLERS
Contents
Page
INTRODUCTION ...................................................... 16 - 1
PROPELLER DEFINITIONS
BLADE ANGLE
GEOMETRIC PITCH
BLADE TWIST .... " ............................................ 16 - 2
EFFECTIVE PITCH
PROPELLER SLIP
THE HELIX ANGLE
ANGLE OF ATTACK
FIXED PITCH PROPELLER
AERODYNAMIC FORCES ON THE PROPELLER ........................... 16 - 4
THRUST
TORQUE
PROPELLER TWISTING MOMENTS .......... , ..................... , ..... 16 - 5
CENTRIFUGAL TWISTING MOMENT (CTM)
AERODYNAMIC TWISTING MOMENT (ATM)
PROPELLER EFFICIENCY .............................................. 16 - 6
VARIATION OF PROPELLER EFFICIENCY WITH SPEED
VARIABLE PITCH PROPELLERS ........................................ 16 - 7
ADJUSTABLE PITCH
TWO PITCH
CONSTANT SPEED PROPELLER
WIND MILLING ................................................. 16 - 9
FEATHERING ................................................. 16 - 10
POWER ABSORPTION ................................................. 16 - 11
SOLIDITY
MOMENTS AND FORCES GENERATED ................................. 16 - 12
TORQUE REACTION
GYROSCOPIC EFFECT ..................................... , .... 16 - 13
SPIRAL SLIPSTREAM EFFECT ... , ............................... 16 - 14
ASYMMETRIC BLADE EFFECT .................. " .............. 16 - 15
EFFECT OF ATMOSPHERIC CONDITIONS
SELF ASSESSMENT QUESTIONS ....................................... 16 - 17
ANSWERS .................................................... 16 - 21
PRINCIPLES OF FLIGHT
16.1
PROPELLERS
INTRODUCTION
A propeller converts shaft power from the engine into thrust. It does this by accelerating a mass
of air rearwards. Thrust from the propeller is equal to the mass of air accelerated rearwards
multiplied by the acceleration given to it. A mass is accelerated rearwards and the equal and
opposite reaction drives the aircraft forwards .
16.2
DEFINITIONS
The propeller blade is an aerofoil and the definitions for chord, camber, thickness/chord ratio
and aspect ratio are the same as those given previously for the wing. Additionally the following
must be considered.
BLADE ANGLE or PITCH
/
/
/
The angle between the blade chord and the
plane of rotation. Blade angle decreases from
the root to the tip of the blade (twist) because
rotational velocity of the blade increases from
root to tip. For reference purposes, the blade
angle is measured at a point 75% of the blade
length from the root.
/
."
/
/,\
BLADE ANGLE
OR
PITCH
PLANE OF ROTATION
Figure 16.1
(
/
Blade angle
/'
Figure 16.2
GEOMETRIC
PITCH
GEOMETRIC PITCH
The geometric pitch is the distance the
' propeller would travel forward in one
complete revolution if it were moving through
the air at the blade angle. (It might help to
imagine the geometric pitch as a screw thread,
but do not take this "screw" analogy any
further).
Geometric pitch
16 - 1
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
PROPELLERS
BLADE TWIST
Sections near the tip of the propeller are at a greater distance from the propeller shaft and travel
through a greater distance. Tip speed is therefore greater. The blade angle must be decreased
towards the tip to give a constant geometric pitch along the length of the blade.
The blade angle determines the geometric pitch of the propeller. A small blade angle is called
"fine pitch", a large blade angle is called "coarse pitch".
EFFECTIVE PITCH
SLIP
GEOMETRIC
PITCH
In flight the propeller does not move
through the air at the geometric pitch,
the distance it travels forward in each
revolution depends on the aircraft's
forward speed. The distance which it
actually moves forward in each
revolution is called the "effective
pitch" or "advance per revolution".
PROPELLER SLIP
EFFECTIVE
PITCH
_~" L"_~
__
__
~~
)
/ )"
"""~~"
~
_____ ___
The difference between the
Geometric and the Effective Pitch is
called the Slip.
/. /
-,:;-",//~
HELIX
ANGLE
THE HELIX ANGLE
Figure 16.3
Effective pitch & Slip
The angle that the actual path of the
propeller makes to the plane of
rotation.
ANGLE OF ATTACK
The path of the propeller through the air determines the direction of the relative airflow. The
angle between the blade chord and the relative airflow is the angle of attack (a), Fig. 16.4. The
angle of attack (a) is the result of propeller rotational velocity (RPM) and aircraft forward
velocity (TAS).
FIXED PITCH PROPELLER
Fig. 16.5 shows a "fixed pitch" propeller at constant RPM. Increasing TAS decreases the angle
of attack of the propeller. Fig. 16.6 shows a "fixed pitch" propeller at a constant TAS .
Increasing RPM increases the angle of attack of the propeller.
16 - 2
© Oxford Aviation Services Limited
PROPELLERS
PRINCIPLES OF FLIGHT
12
RESULTANT PATH
OF BLADE ELEMENT
(RELA~WE AIRFLON
TAS OF AIRCRAFT
+
INDUCED FLOW
BLADE ANGLE
OR
PITCH
\
/
PLANE OF ROTATION
--.-.l
PROPELLER (RPM)
Fig'ure 16.4
Angle of attack
TAS
INCREASED
CONSTANT (RPM)
Figure 16.5
Angle of attack decreased by higher T AS
CONSTANT TAS
/~------------------~--------~--~~
INCREASED (RPM)
Figure 16.6 Angle of attack increased by higher RPM
16 - 3
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
16.3
PROPELLERS
AERODYNAMIC FORCES ON THE PROPELLER
A propeller blade has an aerofoil section, and when moving through the air at an angle of attack
will generate aerodynamic forces in the same way as a wing. The shape of the section will
generate a pressure differential between the two surfaces. The surface which has the greater
pressure is called the "pressure face" or "thrust face". When the propeller is giving forward
thrust, the thrust face is the rear, (flat) surface. The pressure differential will generate an
aerodynamic force, the total reaction, which may be resolved into two components, thrust and
propeller torque.
16.3.1
THRUST
A component at right angles to the plane of rotation. The thrust force will vary along the length
of each blade, reducing at the tip where the pressures equalise and towards the root where the
rotational velocity is low. Thrust will cause a bending moment on each blade, tending to bend
the tip forward. (Equal and opposite reaction to "throwing" air backwards).
16.3.2
TORQUE (Propeller)
The equal and opposite reaction to being rotated, which generates a turning moment about the
aircraft longitudinal axis. Propeller torque also gives a bending moment to the blades in, but in
the opposite direction to, the plane of rotation.
TOTAL
REACTION
THRUST
/
/
/
/
/
/
/
/
/
/
TORQUE
PLANE OF ROTATION
Figure 16.7
Thrust and Torque
16 - 4
© Oxford Aviation Services Limited
PROPELLERS
PRINCIPLES OF FLIGHT
16.4
CENTRIFUGAL TWISTING MOMENT (CTM)
Components 'A' and 'B', of the centrifugal force acting on the blade, produce a moment around
the pitch change axis which tends to 'fine' the blade off.
s ....
PITCH
CHANGE
AXIS
Figure 16.8 Centrifugal Turning Moment (CTM)
16.4.1 AERODYNAMIC TWISTING MOMENT (ATM)
Because the blade CP is in front of the pitch change axis, aerodynamic force generates
a moment around the pitch change axis acting in the direction of coarse pitch.
TOTAL
REACTION
' PITCH
CHANGE
AXIS
Figure 16.9 Aerodynamic Twisting Moment (ATM)
The ATM partially offsets the CTM during normal engine operations, but the CTM is
dominant. However, when the propeller is windmilling the ATM acts in the same
direction as the CTM (See Fig. 16.15), and will reinforce it.
16 - 5
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PRINCIPLES OF FLIGHT
16.5
PROPELLERS
PROPELLER EFFICIENCY
The efficiency of the propeller can be measured from the ratio, Power out / Power in. The power
extracted (out) from a propeller "Thrust Power" is the product of Force (Thrust) x Velocity
(TAS). The power into the propeller "Shaft Power" is engine torque (Force) x Rotational
Velocity (RPM). The efficiency of the propeller can be expressed as:
Propeller Efficiency
16.5.1
=
Thrust Power
Shaft Power
VARIATION OF PROPELLER EFFICIENCY WITH SPEED
Fig. 16.5 illustrated that for a fixed pitch propeller, increasing TAS at a constant RPM reduces
the blade angle of attack. This will decrease thrust. The effect of this on propeller efficiency
is as follows:
a)
At some high forward speed the blade will be close to zero lift angle of attack and thrust,
and therefore Thrust Power, will be zero. From the above 'equation' it can be seen that
propeller efficiency will also be zero.
b)
There will be only one speed at which a fixed pitch propeller is operating at its most
efficient angle of attack and where the propeller efficiency will be maximum, Fig. 16.10.
c)
As TAS is decreased, thrust will increase because blade angle of attack is increased.
Thrust is very large, but the TAS is low so propeller efficiency will be low. Thus no
useful work is being done when the aircraft is, for instance, held against the brakes at
full power prior to take-off. The efficiency of a fixed pitch prop' varies with forward
speed
Ifblade angle can be varied as TAS and/or RPM is changed, the propeller will remain efficient
over a much wider range of aircraft operating conditions, as illustrated in Fig. 16.10.
>()
z
w
(3
u:::
LL
W
AIRCRAFT FORWARD SPEED
Figure 16.10
Efficiency improved by varying blade angle
16 - 6
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
16.6
PROPELLERS
VARIABLE PITCH PROPELLERS
Adjustable pitch propellers:' These are propellers which can have their pitch adjusted on the
ground by mechanically re-setting the blades in the hub. In flight they act as fixed pitch
propellers.
Two pitch propellers: These are propellers which have a fine and coarse pitch setting which
can be selected in flight. Fine pitch can be selected for take off, climb and landing and coarse
pitch for cruise. They will usually also have a feathered position.
(Variable pitch) Constant speed propellers: Modem aircraft have propellers which are
controlled automatically to vary their pitch (blade angle) so as to maintain a selected RPM. A
variable pitch propeller permits high efficiency to be obtained over a wider range ofTAS, giving
improved take-off and climb performance and cruising fuel consumption.
CONSTANT SPEED PROPELLER
OPEN
INCR
MD<TURE
THROTTLE
RPM
I
I
CLOSE
DECR
Figure 16.11
Fig. 16.11 illustrates a 'typical' set of engine and propeller controls for a small piston engine
aircraft with a variable pitch propeller. Throttle, prop' and mixture are shown in the take-off.
(all forward) position.
"Pulling back" on the prop' control will decrease RPM.
"Pushing forward" on the prop' control will increase RPM.
NB:
A reasonable analogy is to think ofthe prop' control as an infinitely variable "gear change".
Forward (increase RPM) is first gear.
Back (decrease RPM) is fifth gear.
16 - 7
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
PROPELLERS
FINE PITCH
("small" blade angle)
Fig. 16.12 shows conditions during the early
stages of take-off roll. The RPM is set to
maximum and the TAS is low. The angle of
attack is optimum and maximum available
efficiency is obtained. As the aircraft
continues to accelerate the TAS will increase,
which decreases the angle of attack of the
blades. Less thrust will be generated and less
propeller torque. This gives less resistance
for the engine to overcome and RPM wouLd
tend to increase. The constant speed unit
(CSU) senses the RPM increase and increases
pitch to maintain the blade angle of attack
constant.
AT THE START OF THE
TAKE - OFF ROLL.
LOW FORWARD SPEED ,
HIGH RPM
Figure 16.12 Low TAS, high RPM
TAS
COARSE PITCH
("large" blade angle)
Fig. 16.13 shows the conditions at
high forward speed in level flight.
As the TAS increased, the CSU
continually increased the blade angle
(coarsened the pitch) to maintain a
constant blade angle of attack.
RPM
HIGH FORWARD SPEED,
HIGH RPM
Figure 16.13 High TAS , high RPM
16 - 8
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PRINCIPLES OF FLIGHT
PROPELLERS
TAS
RPM
Tt r
Fig. 16.14 shows conditions when the engine
and prop ' have been set for cruise conditions.
Optimum throttle and RPM setting are listed
in the aircraft Flight Manual.
The
recommended procedure is to reduce the
throttle first, then RPM.
Whatever configuration into which the aircraft
is placed, climb, descent or bank, the CSU
will adjust the blade angle (prop ' pitch) to
maintain the RPM which has been set. At
least it will try to maintain constant RPM.
There are exceptions, which will be discussed.
CRUISE SETTING
Figure 16.14
FINE PITCH
("small" blade angle)
TAS
WINDMILLING
If a loss of engine torque occurs (the
throttle is closed or the engine fails), the
prop ' will "fine off' in an attempt to
maintain the set RPM.
RPM
DRAG
TOTAL
REACTION
The relative airflow will impinge on the
front surface of the blade and generate drag
and " negative propeller torque". The
propeller will now drive the engine, as
shown in Fig. 16.15 .
The drag generated by a windmilling
propeller is very high.
STEADY GLIDE ,
THROTTLE CLOSED ,
NO SHAFT POWER ,
PROPELLER WINDMILLING.
Figure 16.15
16 - 9
Windmilling
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
PROPELLERS
FEATHERING
Following an engine failure on a twin engine aeroplane the increased drag from the windmilling
propeller will seriously degrade climb performance, limit range and add to the yawing moment
caused by the failed engine which will affect controllability. (Please review Paragraphs 12.24
to 12.39 for detailed information). Also, by continuing to tum a badly damaged engine, eventual
seizure of the engine or an engine fire might result.
ZERO LIFT
ANGLE OF ATTACK
)
I
I
Windmilling drag is one of the "ingredients" of the
yawing moment from a failed engine. Feathering the
propeller of a failed engine will also reduce the
yawing moment and consequently, V Me.
I
I
I
I
Figure 16.16
By turning the blades to their zero lift angle of attack,
no propeller torque is generated and the propeller will
stop, reducing drag to a minimum, as shown in Fig.
16.16. This will improve climb performance because
the ability to climb is dependent on excess thrust after
balancing aerodynamic drag.
Feathered
A single engine aeroplane fitted with a constant
speed propeller does not have a "feathering"
capability, as such. However, following engine
failure, drag can be reduced to a minimum by
"pulling" the RPM (prop') control to the fully
coarse position, as shown in Fig. 16.17.
In a steady glide with no shaft power from the
engine (throttle closed), if the propeller pitcli is
increased by pulling back the prop' lever, the
aircraft Lift/Drag ratio will increase. This will
decrease the rate of descent. The RPM would
decrease because of the reduction in negative
propeller torque.
The opposite will be true if the propeller pitch is
decreased.
16 - 10
COARSE PITCH
("large" blade angle)
STEADY GLIDE,
THROTTLE CLOSED
PROP' LEVER "PULLED - BACK"
Figure 16.17
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
16.7
PROPELLERS
POWER ABSORPTION
A propeller must be able to absorb all the shaft power developed by the engine and also operate
with maximum efficiency throughout the required performance envelope of the aircraft. The
critical factor is tip velocity. If tip velocity is too high the blade tips will approach the local
speed of sound and compressibility effects will decrease thrust and increase rotational drag.
Supersonic tip speed will considerably reduce the efficiency of a propeller and greatly
increase the noise it generates.
This imposes a limit on propeller diameter and RPM, and the T AS at which it can be used.
Other limitations on propeller diameter are the need to maintain adequate ground clearance and
the need to mount the engines of a multi-engine aircraft as close to the fuselage as possible to
minimise the thrust arm. Increasing the propeller diameter requires the engine to be mounted
further out on the wing to maintain adequate fuselage clearance. To keep V Me within acceptable
limits the available rudder moment would have to be increased. Clearly, increasing the propeller
diameter to increase power absorption is not the preferred option.
SOLIDITY
To increase power absorption several
characteristics of the propeller can be
adjusted. The usual method is to increase the
'solidity' of the propeller. Propeller solidity
is the ratio of the total frontal area of the
blades to the area of the propeller disc. It can
be seen from Fig. 16.18 that an increase in
solidity can be achieved by:
a)
Increasing the chord of each blade.
This increases the solidity, but blade
aspect ratio is reduced, making the
propeller less efficient.
b)
Increasing the number of blades.
Power absorption is increased without
increasing tip speed or reducing the
aspect ratio. Increasing the number
of blades beyond a certain number
(five or six) will reduce overall
efficiency.
PROPELLER DISC
Figure 16.18
Solidity of a propeller
Thrust is generated by accelerating air rearwards. Making the disk too solid will reduce the mass
of air that can be drawn through the propeller and accelerated. To increase the number of blades
efficiently, two propellers rotating in opposite directions on the same shaft are used. These are
called: contra-rotating propellers.
16 - 11
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
16.8
PROPELLERS
MOMENTS AND FORCES GENERA TED BY A PROPELLER
Due to its rotation a propeller generates yawing, rolling and pitching moments. These are due
to several different causes :
NB:
a)
Torque reaction
b)
Gyroscopic precession
c)
Spiral (asymmetric) slipstream effect
d)
Asymmetric blade effect
The majority of modern engines are fitted with propellers which rotate clockwise when
viewed from the rear, so called "right-hand" propellers. The exceptions are small twin piston
engine aircraft which often have the propeller of the right engine rotating anti-clockwise to
eliminate the disadvantage of having a "critical engine" (see Paragraph 12.26) plus some older
aircraft.
16.8.1 TORQUE REACTION
Because the propeller rotates clockwise, the equal and opposite reaction (torque) will
give the aircraft an anti-clockwise rolling moment about the longitudinal axis. During
take-off this will apply a greater down load to the left wheel, Fig. 16.19, causing more
rolling resistance on the left wheel making the aircraft want to yaw to the left. In flight,
torque reaction will also make the aircraft want to roll to the left. Torque reaction will
be greatest during high power, low airspeed (lAS) flight conditions. Low lAS will
reduce the power of the controls to counter the "turning" moment due to torque.
TORQU~
Figure 16.19
~~~ELLER
~~TATION
Torque reaction giving left turn during take - off
16 - 12
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
PROPELLERS
Torque reaction can be eliminated by fitting contra-rotating propellers. Torque
from the two propellers, rotating in opposite directions on the same shaft, will cancel
each other out. Co-rotating propellers on a small twin will not normally give a torque
reaction until one engine fails. A left "turning" tendency would then occur. Counterrotating propellers on a small twin will reduce the torque reaction following an engine
failure. (Review Paragraph 12.28(f)).
16.8.2 GYROSCOPIC EFFECT
A rotating propeller has the properties of a gyroscope - rigidity in space and precession.
The characteristic which produces "gyroscopic effect" is precession. Gyroscopic
precession is the reaction that occurs when a force is applied to the rim of a rotating
disc. When a force is applied to the rim of a propeller the reaction occurs 90 ahead in
the direction of rotation, and in the same direction as the applied force. As the aircraft
is pitched up or down or yawed left or right, a force is applied to the rim of the spinning
propeller disc.
0
NB:
Gyroscopic effect only occurs when the aircraft pitches and/or yaws.
For example, if an aircraft with a clockwise rotating propeller is pitched nose up,
imagine that a forward force has been applied to the bottom of the propeller disc. The
force will "emerge" at 90 in the direction of rotation, i.e. a right yawing moment.
Gyroscopic effect can be easily determined when the point of application of the
imagined forward force on the propeller disc is considered.
0
0
Pitch down - forward force on the top, force emerges 90 clockwise, left yaw.
0
Left yaw - forward force on the right, force emerges 90 clockwise, pitch up.
0
Right yaw - forward force on the left, force emerges 90 clockwise, pitch down.
Gyroscopic effect will be cancelled if the propellers are contra rotating.
16 - 13
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
PROPELLERS
16.8.3 SPIRAL SLIPSTREAM EFFECT
As the propeller rotates it produces a backward flow of air, or slipstream, which rotates
around the aircraft, as illustrated in Fig. 16.18. This spiral slipstream causes a change
in airflow around the fin (vertical stabiliser). Due to the direction of propeller rotation
(clockwise) the spiral slipstream meets the fin at an angle from the left, producing a
sideways force on the fin to the right.
Spiral slipstream effect gives the aircraft a yawing moment to the left.
The amount of rotation given to the air will depend on the throttle and RPM setting.
Spiral slipstream effect can be reduced by:
a)
the use of contra or counter rotating propellers,
b)
a small fixed tab on the rudder,
c)
the engine thrust line inclined slightly to the right,
d)
or offsetting the fin slightly.
PROPELLER ROTATION
/
SPIRAL SLIPSTREAM
LEFT YAW
Figure 16.18
Spiral slipstream effect
16 - 14
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
PROPELLERS
16.8.4 ASYMMETRIC BLADE EFFECT
In general the propeller shaft will be inclined upwards from the direction of flight due
to the angle of attack of the aircraft. This gives the down going propeller blade a greater
effective angle of attack than the up going blade. The down going (right) blade will
generate more thrust. The difference in thrust on the two sides of the propeller disc
will give a yawing moment to the left with a clockwise rotating propeller in a noseup attitude.
Asymmetric blade effect will be greatest at full power and low airspeed (high angle of
attack).
16.9
EFFECT OF ATMOSPHERIC CONDITIONS
Changes of atmospheric pressure or temperature will cause a change of air density. This will
affect:
a)
the power produced by the engine at a given throttle position and
b)
the resistance to rotation of the propeller (its drag).
An increase in air density will increase both the engine power and the propeller drag. The
change in engine power is more significant than the change in propeller drag.
16.9.1 ENGINE AND PROPELLER COMBINED
If the combined effect of an engine and propeller is being considered, it is the engine
power change which will determine the result. For an engine driving a fixed pitch
propeller:
a)
if density increases, RPM will increase.
b)
if density decreases, RPM will decrease.
16.9.2 ENGINE ALONE
If the shaft power required to drive the propeller is being considered, then it is only the
propeller torque which needs to be taken into account. To maintain the RPM of a fixed
pitch propeller:
a)
if density increases, power required will increase,
b)
if density decreases, power required will decrease.
16 - 15
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
PROPELLERS
SELF ASSESSMENT QUESTIONS
1.
As a result of gyroscopic precession, it can be said that:
a)
b)
c)
d)
2.
A propeller rotating clockwise as seen from the rear, creates a spiralling slipstream that tends
to rotate the aeroplane to the:
a)
b)
c)
d)
3.
prevents the portion of the blade near the hub from stalling during cruising flight.
permits a relatively constant angle of attack along its length when in cruising flight.
permits a relatively constant angle of incidence along its length when in cruising flight.
minimises the gyroscopic effect.
The Geometric Pitch of a propeller is:
a)
b)
c)
d)
5.
right around the normal axis, and to the left around the longitudinal axis
right around the normal axis, and to the right around the longitudinal axis
left around the normal axis, and to the left around the longitudinal axis
left around the normal axis, and to the right around the longitudinal axis
The reason for variations in geometric pitch (twisting) along a propeller blade is that it:
a)
b)
c)
d)
4.
any pitching around the longitudinal axis results in a yawing moment
any yawing around the normal axis results in a pitching moment
any pitching around the lateral axis results in a rolling moment
any rolling around the longitudinal axis results in a pitching moment
the distance it would move forward in one revolution if there were no slip.
the angle the propeller shaft makes to the plane of rotation.
the distance the propeller actually moves forward in one revolution.
the angle the propeller chord makes to the relative airflow.
Propeller 'slip' is:
a)
b)
c)
d)
the air stream in the wake of the propeller.
the amount by which the distance covered in one revolution falls short of the geometric
pitch.
the increase in rpm which occurs during take-off.
the change of blade angle from root to tip.
16 - 17
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
6.
The distance a propeller actually advances in one revolution is:
a)
b)
c)
d)
7.
airspeed and RPM.
airspeed and altitude.
altitude and RPM.
torque and blade angle.
Which statement is true regarding propeller efficiency? Propeller efficiency is the:
a)
b)
c)
d)
11.
actual distance a propeller advances in one revolution.
ratio of thrust horsepower to shaft horsepower.
ratio of geometric pitch to effective pitch.
ratio of TAS to rpm.
A fixed-pitch propeller is designed for best efficiency only at a given combination of:
a)
b)
c)
d)
10.
angle of attack and chord line.
angle of attack and line of thrust.
chord line and plane of rotation.
thrust line and propeller torque.
Propeller efficiency is the:
a)
b)
c)
d)
9.
twisting.
effective pitch.
geometric pitch.
blade pitch.
Blade angle of a propeller is defined as the angle between the:
a)
b)
c)
d)
8.
PROPELLERS
difference between the geometric pitch of the propeller and its effective pitch.
actual distance a propeller advances in one revolution.
ratio of thrust horsepower to shaft horsepower.
ratio between the rpm and number of blade elements.
Which statement best describes the operating pLinciple of a constant-speed propeller?
a)
b)
c)
d)
As throttle setting is changed by the pilot, the prop governor causes pitch angle of the
propeller blades to remain unchanged.
The propeller control regulates the engine RPM and in tum the propeller RPM.
A high blade angle, or increased pitch, reduces the propeller drag and allows more
engine power for takeoffs.
As the propeller control setting is changed by the pilot, the RPM of the engines remains
constant as the pitch angle of the propeller changes.
16 - 18
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
12.
When does asymmetric blade effect cause the aeroplane to yaw to the left?
a)
b)
c)
d)
13.
When at high angles of attack.
When at high airspeeds.
When at low angles of attack.
In the cruise at low altitude.
The left turning tendency of an aeroplane caused by asymmetric blade effect is the result of the:
a)
b)
c)
d)
14.
PROPELLERS
gyroscopic forces applied to the rotating propeller blades acting 90° in advance of the
point the force was applied.
clockwise rotation of the engine and the propeller turning the aeroplane counterclockwise.
propeller blade descending on the right, producing more thrust than the ascending blade
on the left.
the rotation of the slipstream striking the tail on the left.
With regard to gyroscopic precession, when a force is applied at a point on the rim of a spinning
disc, the resultant force acts in which direction and at what point?
a)
b)
c)
d)
0
In the same direction as the applied force, 90 ahead in the plane of rotation.
In the opposite direction of the applied force, 90 ahead in the plane of rotation.
In the opposite direction of the applied force, at the point of the applied force.
In the same direction as the applied force, 90 ahead of the plane of rotation when the
propeller rotates clockwise, 90 retarded when the propeller rotates counter-clockwise.
0
0
0
15.
The angle of attack of a fixed pitch propeller:
a)
b)
c)
d)
16.
depends on forward speed only.
depends on forward speed and engine rotational speed.
depends on engine rotational speed only.
is constant for a fixed pitch propeller.
Counter-rotating propellers are:
a)
b)
c)
d)
propellers which rotate counter clockwise.
propellers which are geared to rotate in the opposite direction to the engine.
two propellers driven by separate engines, rotating in opposite directions.
two propellers driven by the same engine, rotating in opposite directions.
16 - 19
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
17.
PROPELLERS
If engine rpm is to remain constant on an engine fitted with a variable pitch propeller, an
increase in engine power requires:
a)
b)
c)
d)
a decrease in blade angle.
a constant angle of attack to be maintained to stop the engine from overspeeding.
an increase in blade angle.
the prop control lever to be advanced.
16 - 20
© Oxford Aviation Services Limited
PRINCIPLES OF FLIGHT
PROPELLERS
ANSWERS
I No I A I B I c I D II
1
D
3
B
A
5
B
6
B
7
C
8
9
B
A
10
C
11
12
B
A
13
14
15
I
B
2
4
REF
C
A
B
16
C
17
C
16 - 21
© Oxford Aviation Services Limited
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