CAE-Oxford-Aviation-Academy-080-Principles-of-Flight-ATPL-Ground-Training-Series-2014.pdf

PRINCIPLES OF FLIGHT ATPL GROUND TRAINING SERIES

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Introduction

© CAE Oxord Aviation Academy (UK) Limited 2014

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All Rights Reserved

I n t r o d u c t i o n

This text book is to be used only or the purpose o private study by individuals and may not be reproduced in any orm or medium, copied, stored in a retrieval system, lent, hired, rented, transmitted or adapted in whole or in part without the prior written consent o CAE Oxord Aviation Academy.

Copyright in all documents and materials bound within these covers or attached hereto, excluding that material which is reproduced by the kind permission o third parties and acknowledged as such, belongs exclusively to CAE Oxord Aviation Academy. Certain copyright material is reproduced with the permission o the International Civil Aviation Organisation, the United Kingdom Civil Aviation Authority and the European Aviation Saety Agency (EASA).

This text book has been written and published as a reerence work to assist students enrolled on an approved EASA Air Transport Pilot Licence (ATPL) course to p repare themselves or the EASA ATP ATPL L theoretical knowledge examinations. Nothing in the content o this book is to be interpreted as constituting instruction or advice relating to practical flying.

Whilst every effort has been made to ensure the accuracy o the inormation contained within this book, neither CAE Oxord Aviation Academy nor the distributor gives any warranty as to its accuracy or otherwise. Students preparing or the EASA ATPL (A) theoretical knowledge examinations should not regard this book as a substitute or the EASA ATPL (A) theoretical knowledge knowledge training syllabus published in the current edition edition o ‘Part-FCL 1’ (the Syllabus). The Syllabus constitutes the sole authoritative definition o the subject matter to be studied in an EASA ATPL (A) theoretical knowledge training programme. No student should prepare or, or is currently entitled to enter himsel/hersel or the EASA ATPL (A) theoretical knowledge examinations without first being enrolled in a training school which has been granted approval by an EASA authorised national aviation authority to deliver EASA ATPL (A) training. CAE Oxord Aviation Academy excludes all liability or any loss or damage incurred or suffered as a result o any reliance on all or part o this book except or any liability or death or personal injury resulting rom CAE Oxord Aviation Aviatio n Academy’s negligence or any other liability which may not legally be excluded.

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Introduction

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Introduction

Textbook Series

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Book

Title

1

010 Air Law

2

020 Aircraf General Knowledge 1

Subjec t

Airrames & Systems Fuselage, Wings & Stabilising Suraces Landing Gear Flight Controls Hydraulics Air Systems & Air Conditioning Anti-icing & De-icing Fuel Systems Emergency Equipment

3


Electrics – Elec tronics Direct Current Alternating Current

4


Powerplant Piston Engines Gas Turbines

5


Instrumentation Flight Instruments Warning & Recording Automatic Flight Control Power Plant & System Monitoring Instruments

iv

6

030 Flight Perormance & Planning 1

Mass & Balance Perormance

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030 Flight Per ro ormance & Plan ann ning 2

Flight Planning & Monitoring

8

040 Hu Human Pe Per ro ormance & Limitations

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050 Meteorology

10

060 Navigation 1

General Navigation

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060 Navigation 2

Radio Navigation

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070 Op Operational Pr Procedures

13

080 Principles o Flight

14

090 Communications

VFR Communications IFR Communications

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Introduction

Contents

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n o i t c u d o r t n I

ATPL Book 13 Principles of Flight 1. Over view and Definitions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .1 2. The Atmosphere . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 23 3. Basic Aerodynamic Theory . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .41 4. Subsonic Air flow . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .51 5. Lif . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .69 69 6. Drag . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

109

7. Stalling . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

143

8. High Lif Devices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

207

9. Air rame Contamination . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

231

10. Stability and Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

239

11. Controls . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 331 12. Flight Mechanics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

363

13. High Speed Flight . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

405

14. Limitations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

457

15. Windshear . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

487

16. Propellers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

501

17. Revision Questions. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

523

18. Index . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

593

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Introduction

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Chapter

1 Overview and Definitions

Overview Overvie w . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

3

General De D efinitions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

7

Glossary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

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List o Symb Symbols ols . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .14 Greek Symbols. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14 Others . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

15

Sel-assessment Questions Questions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16 Answers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

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Overview and Definitions

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O v e r v i e w a n d D e fi n i t i o n s

Figure 1.1

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Overview and Definitions Overview

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s n o i t i n fi e D d n a w e i v r e v O

The primary requirements o an aircraf are as ollows: • • • • •

a wing to to generate generate a lif orce; a uselage to house the payload; tail suraces to add stability; control suraces to change change the direction o flight; and engines to make make it go orward.

The process o lif generation is airly straightorward straightorward and easy to to understand. Over the years aircraf designers, aerodynamicists and structural engineers have refined the basics and, by subtle changes o shape and configuration, have made maximum use o the current understanding o the physical properties o air to produce aircraf best suited to a particular role. Aircraf come in different shapes and sizes, sizes, each usually designed or a specific task. All aircraf share certain eatures, but to obtain the perormance required by the operator, the designer will configure each type o aeroplane in a specific way. As can be seen rom the illustrations on the acing page, the position o the eatures shared by all types o aircraf i.e. wings, uselage, tail suraces and engines varies rom type to type. Why are wing plan shapes different? diff erent? Why are wings mounted sometimes on top o the uselage instead o the bottom? Why are wings mounted in that position and at that angle? Why is the horizontal stabilizer mounted sometimes high on top o the fin rather than on either side o the rear uselage? Every eature has a purpose and is never included merely or reasons o st yle.

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Overview and Definitions An aeroplane, like like all bodies, has mass. With the aircraf stationary on the ground it has only the orce orce due to to the accelerat acceleration ion o gravity acting upon it. This orce, orce, its WEIGHT, WEIGHT, acts vertically vertically downward at all times.

1


W Figure 1.2 The orce o weight

Beore an aeroplane can leave the ground and fly, the orce o weight must be balanced by a orce which acts upwards. This orce is called LIFT. The lif orce must be increased until it is the same as the aeroplane’s weight.

L

W Figure 1.3 The orces o weight & lif

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Overview and Definitions To generate a lif orce, the aeroplane must be propelled orward through the air by a orce called THRUST, provided by the engine(s).

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L

W Figure 1.4 The orces o weight, lif & thrust

From the very moment the aeroplane begins to move, air resists its orward motion with a orce called DRAG.

L

W Figure 1.5 The orces o weight, lif, thrust & drag

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Overview and Definitions When an aeroplane is moving there are our main orces acting upon it:

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WEIGHT, LIFT, THRUST and DRAG.


These are all closely interrelated, interrelated, i.e.: The greater the weight - the greater the lif requirement. The greater the lif - the greater the drag. The greater the drag drag - the greater greater the thrust required, and so on ... Air has properties which change with altitude. altitude. Knowledge o these variables, variables, together with their effect effect on an aeroplane, aeroplane, is a prerequisite prerequisite or a ull understanding o the principles o flight. The structural and aerodynamic design design o an aeroplane is a masterpiece masterpiece o compromise. An improvement in one area requently leads to a loss o efficiency in another. An aeroplane does not ‘grip’ the air as a car does the road. An aeroplane is ofen not pointing in the same direction in which it is moving.

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Overview and Definitions General Definitions

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Mass Unit - Kilogram (kg) - ‘The quantity o matter in a body.’ The mass o a body is a measure o 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.) • The larger the mass, the greater the FORCE required to start or stop it in the same distance. • Mass has a big influence on the time and/or distance required to change the direction o a body.

Force Unit - newton (N) - ‘A push or a pull’. That which causes or tends to cause a change in motion o a body. There are our orces acting on an aircraf in flight - pushing or pulling in different directions.

Weight Unit - newton (N) - ‘The orce due to gravity’. ( F =

× g )

m

where (m) is the mass o the object and ( g) is the acceleration due to the gravity constant, which has the value o 9.81 m/s . ( A 1 kg mass ‘weighs’ 9.81 newtons ) 2

I the mass o a B737 is 60 000 kg and F =

m

× g

it is necessary to generate: [60 000 kg × 9.81 m/s ] 2

588 600 N o lif orce.

Centre of Gravity (CG) The point through which the weight o an aircraf acts. • An aircraf in flight rotates around its CG. • The CG o an aircraf must remain within certain orward and af limits, or reasons o both stability and control.

Work Unit - Joule (J) - A orce is said to do work on a body when it moves the body in the direction in which the orce is acting. The amount o work done on a body is the product o the orce applied to the body and the distance moved by that orce in the direction in which it is acting. I a orce is exerted and no movement takes place, no work has been done. • Work = Force × Distance (through which the orce is applied) • I a orce o 10 newtons moves a body 2 metres along its line o action, it does 20 newton metres (Nm) o work. [10 N × 2 m = 20 Nm] • A newton metre, the unit o work, is called a joule (J). 7

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Overview and Definitions Power

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Unit - Watt (W) - Power is simply the rate o doing work (the time taken to do work).


• Power (W) =

Force (N) × Distance (m) Time (s)

• I a orce o 10 N 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 o power. So the power used in this example is 4 watts.

Energy Unit - Joule (J) - Mass has energy i it has the ability to do work. The amount o energy a body possesses is measured by the amount o work it can do. The unit o energy will, thereore, be the same as those o work, joules.

Kinetic Energy Unit - Joule (J) - ’The energy possessed by mass because o its motion’. ’A mass that is moving can do work in coming to rest’. KE = ½mV joules 2

The kinetic energy o a 1 kg mass o air moving at 52 m/s (100 knots) is 1352 joules; it possesses 1352 joules o kinetic energy. [ 0.5 × 1 × 52 × 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).

Newton’s First Law of Motion ’A body will remain at rest or in uniorm motion in a straight line unless acted on by an external orce’. To move a stationary object or to make a moving object change its direction, a orce must be applied.

Inertia ‘The opposition which a body offers to a change in motion’. A property o all bodies, inertia is a quality, but it is measured in terms o mass, which is a quantity. • The larger the mass, the greater the orce required or the same result. • A large mass has a lot o inertia. • Inertia reers to both stationary and moving masses.

Newton’s Second Law of Motion ’The acceleration o a body rom a state o rest, or uniorm motion in a straight line, is proportional to the applied orce and inversely proportional to the mass’.

Velocity Unit - Metres per second (m/s). - ‘Rate o change o displacement’

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Overview and Definitions Acceleration

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Unit - Metres per second per second (m/s ) - ‘Rate o change o velocity’. 2

A orce o 1 newton acting on a mass o 1 kg will produce an acceleration o 1 m/s Acceleration =


2

Force Mass

• For the same mass; the bigger the orce, the greater the acceleration. • For the same orce; the larger the mass, the slower the acceleration.

Momentum Unit - Mass × Velocity (kg-m/s) - ‘The quantity o motion possessed by a body’. The tendency o a body to continue in motion afer being placed in motion. • A body o 10 kg mass moving at 2 m/s has 20 kg-m/s o momentum. • 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’ • I a orce accelerates a mass in one direction, the body supplying the orce will be subject to the same orce in the opposite direction.

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Overview and Definitions Glossary

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Aerooil - A body so shaped as to produce aerodynamic reaction normal to the direction o its

motion through the air without excessive drag. Af - To the rear, back or tail o the aircraf. Air brake - Any device primarily used to increase drag o an aircraf at will. Ambient - Surrounding or pertaining to the immediate environment. Amplitude - Largeness; abundance; width; range; extent o repetitive movement (rom

extreme to extreme). Attitude - The nose-up or nose-down orientation o an aircraf relative to the horizon. Boundary Layer - The thin layer o air adjacent to a surace, in which the viscous orces are

dominant. Buffeting - An irregular oscillation o any part o an aircraf produced and maintained directly

by an eddying flow. Cantilever (wing) - A wing whose only attachment to the uselage is by fittings at the wing root:

it has no external struts or bracing. The attachments are aired-in to preserve the streamline shape Control Lock (Gust lock) - A mechanical device designed to saeguard, by positive lock, the

control suraces and flying control system against damage in high winds or gusts when the aircraf is parked. Control Reversal - At high speed: the displacement o a control surace producing a moment

on the aircraf in a reverse sense because o excessive structural distortion. At low speed: the displacement o an aileron increasing the angle o attack o 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 (MCRIT) - The ree stream Mach number at which the peak velocity on the

surace o a body first becomes equal to the local speed o sound.

Damping - To slow down the rate; to diminish the amplitude o vibrations or cycles. Geometric Dihedral - The angle between the horizontal datum o an aeroplane and the plane

o a wing or horizontal stabilizer semi-span. Divergent - Inclined or turned apart. Divergence - A disturbance which increases continually

with time. Eddy - An element o air having intense vorticity. Effective Angle o Attack (αe) - The angle between the chord line and the mean direction o a

non-uniorm disturbed airstream.

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Overview and Definitions Equilibrium - A condition that exists when the sum o all moments acting on a body is zero AND the sum o all orces acting on a body is zero.

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Fairing - A secondary structure added to any part to reduce its drag. Feel - The sensations o orce and displacement experienced by the pilot rom the aerodynamic

orces on the control suraces. Fence - A projection rom the surace o the wing and extending chordwise to modiy the wing

surace pressure distribution. Fillet - A airing at the junction o two suraces to improve the airflow. Flight Path - The path o the Centre o Gravity (CG) o an aircraf. Fluid - A substance, either gaseous or liquid, that will conorm to the shape o the container

that holds it. Free Stream Velocity - The velocity o the undisturbed air relative to the aircraf. Gradient (Pressure) - Rate o change in pressure with distance. Gust - A rapid variation, with time or distance, in the speed or direction o air. Gust Lock - See control lock. Instability - The quality whereby any disturbance rom steady motion tends to increase. Laminar Flow - Flow in which there is no mixing between adjacent layers. Load Factor - The ratio o the weight o an aircraf to the load imposed by lif. The correct

symbol or load actor is (n), but it is colloquially known as (g). Load Factor =

Lif Weight

Mach Number (M) - The ratio o the True Airspeed to the speed o sound under prevailing

atmospheric conditions. M =

TAS Local Speed o Sound (a)

Magnitude - Largeness; size; importance. Moment (N-m) - The moment o a orce about a point is the product o the orce 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 or housing engines (usually). Normal - Perpendicular; at 90°. Oscillation - Swinging to and ro 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.

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Overview and Definitions Pitot Tube - A tube, with an open end acing up-stream, wherein at speeds less than about our

1

tenths the speed o 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 rom a uselage or wing. Propagate - To pass on; to transmit; to spread rom one to another. Relative Airflow , (Relative Wind), (Free Stream Flow) - The direction o airflow produced

by the aircraf moving through the air. The relative airflow flows in a direction parallel and opposite to the direction o flight. Thereore, the actual flight path o the aircraf determines the direction o the relative airflow. Also, air in a region where pressure, temperature and relative velocity are unaffected by the passage o the aircraf through it. Scale - I a 1/10th scale model is considered, all the linear dimensions are 1/10th o the real

aircraf, but the areas are 1/100th; and i the model is constructed o the same materials, the mass is 1/1000th o the real aircraf. So the model is to scale in some respects, but not others. Schematic - A diagrammatic outline or synopsis; an image o the thing; representing something

by a diagram. Separation - Detachment o the airflow rom a surace with which it has been in contact. Shock Wave - 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 o velocity relative to the shock wave is supersonic upstream and subsonic downstream. Side-slip - Motion o an aircraf, relative to the relative airflow, which has a component o

velocity along the lateral axis. Slat - An auxiliary, cambered aerooil positioned orward o the main aerooil so as to orm a

slot. Spar - A principal spanwise structural member o a wing, tailplane, fin or control surace. Speed - Metres per second (m/s) is used in most ormulae, but nautical miles per hour or knots

(kt) are commonly used to measure the speed o an aircraf. There are 6080 f in 1 nautical mile and 3.28 f in 1 metre. Speed o Sound (a) - Sound is pressure waves which propagate spherically through the

atmosphere rom their source. The speed o propagation varies ONLY with the temperature o the air. The lower the temperature, the lower the speed o propagation. On a ’standard’ day at sea level the speed o sound is approximately 340 m/s (660 kt TAS). Stability - The quality whereby any disturbance o 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 surace. Static Vent - A small aperture in a plate fixed to orm part o the uselage and located

appropriately or measuring the ambient static pressure. Throat - A section o minimum area in a duct.

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Overview and Definitions True Airspeed (TAS) or (V) - The speed at which the aircraf is travelling through the air.

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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 o fluid particles to flow over each other. All fluids have the property

o viscosity. A fluid with high viscosity would not flow very easily. The viscosity o 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 o fluid in circulatory motion, having a core o intense vorticity, the strength

o the vortex being given by its circulation. Vortex Generator - A device, ofen a small vane attached to a surace, to produce one or

more discrete vortices which trail downstream adjacent to the surace, promote mixing in the boundary layer and delay boundary layer separation. (Increases the kinetic energy o the boundary layer). Vorticity - Generally, rotational motion in a fluid, defined, at any point in the fluid, as twice the

mean angular velocity o a small element o fluid surrounding the point. Wake - The region o air behind an aircraf in which the total pressure has been changed by the

presence o the aircraf. Wash-out - Decrease in angle o incidence towards the tip o a wing or other aerooil. Wing Loading - Ratio o aircraf weight to wing area. Wing Loading =

Aircraf Weight Wing Area

Zoom - Using kinetic energy to gain height.

13

1

Overview and Definitions List of Symbols

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The ollowing symbols are used throughout these notes. However, no universal defining standard or their use exists. Other books on the subject may use some o these symbols with different definitions. Every effort has been made to employ symbols that are widely accepted and that conorm to the Learning Objectives. a AC AR b C c CD CG CP CL CM D Di F g K L L/D M m n p Q or q S T t/c V VS W

speed o sound aerodynamic centre aspect ratio span Centigrade chord length drag coefficient centre o gravity centre o pressure lif coefficient pitching moment coefficient drag induced drag orce acceleration due to gravity, also used or load actor Kelvin lif lif to drag ratio Mach number mass load actor pressure dynamic pressure area; wing area temperature thickness-chord ratio ree stream speed (TAS) stall speed weight

Greek Symbols α β γ Δ μ ρ σ φ

14

(alpha) angle o attack (beta) sideslip angle (gamma) angle o climb or descent (delta) increment in (mu) Mach angle (rho) density (sigma) relative density (phi) angle o bank

1

Overview and Definitions Others ∝

proportional to

=

is approximately equal to

• •

1


Note: The Greek symbol γ (gamma) has been used in these notes to denote angle o climb and descent. The Learning Objectives use θ (theta). Evidence exists that a question in the exam uses γ (gamma) or angle o climb and descent. The notes have been amended to use γ , but consider either γ or θ to indicate angle o climb and descent.

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Questions Self-assessment Questions

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Q u e s t i o n s

Aircraf (1)

Mass: 2000 kilograms (kg) Engine thrust: 4000 newtons (N) V1 speed: 65 knots (kt) Take-off run to reach V 1: 750 metres (m) Time taken to reach V 1: 30 seconds (s) Aircraf (2)

Mass: 2000 kilograms (kg) Engine thrust: 8000 newtons (N) V1 speed: 130 knots (kt) Take-off run to reach V 1: 1500 metres (m) Time taken to reach V 1: 40 seconds (s) where 1 nautical mile = 6080 f and 1 metre = 3.28 f At V1 both aircraf experience an engine ailure and take-off is abandoned. a. b. c. d. e. . g. h. i. j. k. l. m.

16

How much work was done to aircraf (1) getting to V1? How much power was used to get aircraf (1) to V1? How much work was done to aircraf (2) getting to V1? How much power was used to get aircraf (2) to V1? How much momentum does aircraf (1) possess at V1? How much momentum does aircraf (2) possess at V1? How many times greater is the momentum o aircraf (2)? How much kinetic energy does aircraf (1) possess at V 1? How much kinetic energy does aircraf (2) possess at V 1? How many times greater is the kinetic energy o aircraf (2)? State the mass and velocity relationship o both aircraf and compare to their momentum and kinetic energy. Which has the greater effect on kinetic energy, mass or velocity? What must be done with the kinetic energy so the aircraf can be brought to a stop?

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Questions 1.

An aircraf’s mass is a result o:

a. b. c. d. 2.

the orce on mass due to gravity. the action o a alling mass. how much matter the object contains. the rate o mass per unit area.

About which point does an aircraf rotate?

a. b. c. d. 8.

kilogram. newton. watt. kilowatt.

Weight is the result o:

a. b. c. d. 7.

mass-kilogram. newton-metre. joule. newton.

The unit o 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 o an applied input.

The unit o orce is the:

a. b. c. d. 5.

joule. watt. newton. kilogram.

The definition o a orce is:

a. b. c. d. 4.

s n o i t s e u Q

its weight. how big it is. how much matter it contains. its volume.

The unit o mass is the:

a. b. c. d. 3.

1

The wings. The main undercarriage. The centre o gravity. The rudder.

I a orce is applied to a mass and the mass does not move:

a. b. c. d.

work is done even though there is no movement o the mass. work is done only i the mass moves a long way. power is exerted, but no work is done. no work is done.

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Questions 9.

1

The unit o work is called the:

a. b. c. d.

Q u e s t i o n s

10.

The unit o power is called the:

a. b. c. d. 11.

pascal. joule. watt. kilogram.

joule. newton-metre. watt. metre per second.

I a orce o 20 newtons moves a mass 5 metres: 1 - the work done is 100 Nm 2 - the work done is 100 joules 3 - the work done is 4 joules 4 - the work done is 0.25 joules The correct statements are:

a. b. c. d. 12.

I a orce o 50 newtons is applied to a 10 kg mass and the mass moves 10 metres and a orce o 50 newtons 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.

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the rate o orce applied. the rate o movement per second. the rate o doing work. the rate o applied orce.

I a orce o 500 newtons 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 o power is:

a. b. c. d. 14.

1 only. 1 and 3. 1 and 2. 2 only.

the energy a mass possesses due to its position in space. the energy a mass possesses when a orce has been applied. the energy a mass possesses due to the orce o gravity. the energy a mass possesses because o its motion.

1

Questions 16.

The unit o kinetic energy is the:

a. b. c. d. 17.

1

s n o i t s e u Q

joule. metre per second. watt. newton-metre per second.

When considering kinetic energy: 1 - a moving mass can apply a orce by being brought to rest. 2 - kinetic energy is the energy possessed by a body because o its motion. 3 - i a body’s kinetic energy is increased, a orce must have been applied. 4 - kinetic energy = ½ m V 2 joules. The combination o correct statements is:

a. b. c. d. 18.

The property o 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 o its motion. the opposition which a body offers to a change in motion. that every action has an equal and opposite reaction. the quantity o motion possessed by a body.

Considering Newton’s first law o motion: 1 - a body is said to have energy i it has the ability to do work. 2 - the amount o energy a body possesses is measured by the amount o work it can do. 3 - a body will tend to remain at rest, or in uniorm motion in a straight line, unless acted upon by an external orce. 4 - to move a stationary object or to make a moving object change its direction, a orce 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.

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Questions 20.

1

Considering Newton’s second law o motion: 1 - every action has an equal and opposite reaction. 2 - i the same orce is applied, the larger the mass the slower the acceleration. 3 - i two orces are applied to the same mass, the bigger the orce the greater the acceleration. 4 - the acceleration o a body rom a state o rest, or uniorm motion in a straight line, is proportional to the applied orce and inversely proportional to the mass.

Q u e s t i o n s

The combination o true statements is:

a. b. c. d. 21.

Newton’s third law o motion states:

a. b. c. d. 22.

the energy possessed by a mass is inversely proportional to its velocity. every orce has an equal and opposite inertia. or every orce there is an action. every action has an equal and opposite reaction.

The definition o velocity is the:

a. b. c. d. 23.

1 only. 1, 2, 3 and 4. 2, 3, and 4. 3 and 4.

rate o change o acceleration. rate o change o displacement. the quantity o motion possessed by a body. the acceleration o a body in direct proportion to its mass.

When considering acceleration: 1 - acceleration is the rate o change o velocity. 2 - the units o acceleration are metres per second. 3 - the units o acceleration are kilogram-metres per second. 4 - the units o acceleration are seconds per metre per metre. The combination o correct statements is:

a. b. c. d. 24.

The definition o momentum is:

a. b. c. d.

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4 only. 1 and 4. 1 only. 1 and 2.

the quantity o mass possessed by a body. the quantity o inertia possessed by a body. the quantity o motion possessed by a body. the opposition which a body offers to a change in velocity.

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Questions 25.

A orce o 24 newtons moves a 10 kg mass 60 metres in 1 minute. The power used is:

1

s n o i t s e u Q

1 - 24 watts. 2 - 240 watts. 3 - orce times distance moved in one second. 4 - orce times the distance the mass is moved in one second. Which o 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 o motion possessed by a body. 2 - momentum is the tendency o a body to continue in motion afer being placed in motion. 3 - a mass o 2000 kg moving at 55 m/s has 110 000 kg-m/s o momentum. 4 - a large mass moving at 50 m/s will have less momentum than a small mass moving at 50 m/s. The correct combination o statements is:

a. b. c. d.

1 and 3. 1, 2, 3 and 4. 1, 2 and 3. 2, 3 and 4.

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Answers

Answers

1

A n s w e r s

Aircraf number (1) V1 speed o 65 knots = 33.5 m/s Aircraf number (2) V1 speed o 130 knots = 67 m/s a. b. c. d. e . g. h. i. j. k.

3 000 000 joules 100 000 watts 12 000 000 joules 300 000 watts 67 000 kg-m/s 134 000 kg-m/s twice 1 122 250 joules 4 489 000 joules our times greater same mass, speed doubled, momentum doubled, but kinetic energy our times greater. velocity has a greater effect on kinetic energy than mass. it must be dissipated by the braking systems.

l. m.

22

1 c

2 d

3 c

4 d

5 b

6 a

7 c

8 d

9 b

10 c

11 c

12 a

13 c

14 a

15 d

16 a

17 b

18 b

19 a

20 c

21 d

22 b

23 c

24 c

25 b

26 c

Chapter

2 The Atmosphere

Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 25 The Physical Properties o Air . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .25 Static Pressure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 25 Temperature. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .26 Air Density. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .26 International Standard Atmosphere (ISA) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .26 Dynamic Pressure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .27 Key Facts. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .29 Measuring Dynamic Pressure. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .30 Relationships between Airspeeds . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 31 Airspeed . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .32 Errors and Corrections . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 32 V Speeds . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .33 Summary. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .34 Questions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

35

Answers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

40

23

2

The Atmosphere

2

T h e A t m o s p h e r e

24

2

The Atmosphere Introduction The atmosphere is the medium in which an aircraf operates. It is the properties o the atmosphere, changed by the shape o the wing, that generate the required lif orce.

2

e r e h p s o m t A e h T

The most important property is air density (the “thickness” o air).

KEY FACT: I air density decreases, the mass o air flowing over the aircraf in a given time will decrease. Not usually considered during the study o Principles o Flight, keeping the idea o mass flow (kg/s) in the ‘back o your mind’ can aid general understanding. A given mass flow will generate the required lif orce, but a decrease in air density will reduce the mass flow. To maintain the required lif orce i density decreases, the speed o the aircraf through the air must be increased. The increased speed o airflow over the wing will maintain the mass flow and lif orce at its required value.

The Physical Properties of Air Air has substance! Air has mass; not very much i compared to other matter, but nevertheless a significant amount. A mass o moving air has considerable kinetic energy; or example, when moving at 100 knots the kinetic energy o 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 o the lower pressure). The viscosity o air is so low that very small orces are able to move the molecules in relation to each other. When considering the portion o atmosphere in which most aircraf operate (up to 40 000 f), with increasing altitude the characteristics o air undergo a gradual transition rom those at sea level. Since air is compressible, the lower layers contain much the greater part o the whole mass o the atmosphere. Pressure alls steadily with increasing altitude, but temperature alls steadily only to about 36 000 f, above which it then remains constant through the stratosphere.

Static Pressure The unit or static pressure is N/m , the symbol is lower case ‘p’. 2

• Static pressure is the result o the weight o the atmosphere pressing down on the air beneath. • Static pressure will exert the same orce per square metre on all suraces o an aeroplane. The lower the altitude, the greater the orce per square metre. • It is called static pressure because o the air’s stationary or static presence. • An aircraf always has static pressure acting upon it. Newtons per square metre is the SI unit or pressure. 1 N/m 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). 2

25

2

The Atmosphere Static pressure at a particular altitude will vary rom day to day, and is about 1000 hPa at sea level. In those countries that measure static pressure in inches o mercury (inHg), sea level static pressure is about 30 inHg.

2


Temperature The unit or temperature is °C, or K. It is degrees Celsius (or centigrade) when measured relative to the reezing point o water, or Kelvin when measured relative to absolute zero. (0°C is equivalent to 273 K). Temperature decreases with increasing altitude up to about 36 000 f and then remains constant.

Air Density The unit or density is kg/m and the symbol is the Greek letter ρ [rho]. 3

• Density is ‘mass per unit volume’ (The ‘number’ o air particles in a given space). • Density varies with static pressure, temperature and humidity. • Density decreases i static pressure decreases. • Density decreases i temperature increases. • Density decreases i humidity increases. Air Density is proportional to pressure and inversely proportional to temperature. This is shown in the ideal gas law ormula below. P = constant, more useully it can be said that ρ ∝ Tρ

P T

where p = pressure, T = temperature, and ρ = density Density decreases with increasing altitude because o decreasing static pressure. However, with increasing altitude temperature also decreases, which would tend to increase density, but the effect o decreasing static pressure is dominant.

International Standard Atmosphere (ISA) The values o temperature, pressure and density are never constant in any given layer o the atmosphere. To enable accurate comparison o aircraf perormance and the calibration o pressure instruments, a ‘standard’ atmosphere has been adopted. The standard atmosphere represents the mean or average properties o the atmosphere. Europe uses the standard atmosphere defined by the International Civil Aviation Organization (ICAO). The ICAO standard atmosphere assumes the ollowing mean sea level values: Temperature 15°C Pressure 1013.25 hPa Density 1.225 kg/m

3

26

2

The Atmosphere The temperature lapse rate is assumed to be uniorm at a rate o 2°C per 1000 f (1.98°C) rom mean sea level up to a height o 36 090 f (11 000 m) above which the lapse rate becomes zero and the temperature remains constant at -56.5°C.

2


ICAO Standard Atmosphere

Altitude (f)

Temperature (°C)

Pressure (hPa) (p)

Density (kg/m ) (ρ)

Relative Density (σ)

50 000

- 56.5

116.0

0.186

0.15

45 000

- 56.5

147.5

0.237

0.19

40 000

- 56.5

187.6

0.302

0.25

35 000

- 54.3

238.4

0.386

0.31

30 000

- 44.4

300.9

0.458

0.37

25 000

- 34.5

376.0

0.549

0.45

20 000

- 24.6

465.6

0.653

0.53

15 000

- 14.7

571.8

0.771

0.63

10 000

- 4.8

696.8

0.905

0.74

5000

5.1

843.1

1.056

0.86

Sea Level

15

1013.25

1.225

1.0

3

NOTE : High Density Altitude means that the conditions that actually exist at the airport o take- off or landing represent those o a higher altitude in the International Standard Atmosphere i.e. less air density.

Dynamic Pressure The unit or dynamic pressure is N/m and the symbol is lower case ‘q’ or upper case ‘Q’. 2

• Because air has mass, air in motion must possess kinetic energy , and will exert a orce per square metre on any object in its path. (KE = ½ m V ) 2

• It is called DYNAMIC pressure because the air is moving in relation to the object being considered, in this case an aircraf. • Dynamic pressure is proportional to the density o the air and the square o the speed o the air flowing over the aircraf. An aircraf immersed in moving airflow will thereore experience both static AND dynamic pressure. (Remember, static pressure is always present).

27

2

The Atmosphere The kinetic energy o one cubic metre o air moving at a stated speed is given by the ormula: Kinetic Energy = ½ ρ V joules where ρ is the local air density in kg/m and V is the speed in m/s 2

2

3


I this cubic metre o moving air is completely trapped and brought to rest by means o an open-ended tube the total energy will remain constant , but by being brought completely to rest the kinetic energy will become pressure energy which, or all practical purposes, is equal to: Dynamic Pressure = ½ ρ V

2

N/m

2

Consider air flowing at 52 m/s (100 kt) with a density o 1.225 kg/m

3

(100 kt = 100 NM/h = 100 × 6080 f/h = 608 000 ÷ 3.28 = 185 366 ÷ 60 ÷ 60 m/s = 52 m/s) Dynamic pressure = 0.5 × 1.225 × 52 × 52 = 1656 N/m (16.56 hPa) 2

I speed is doubled, dynamic pressure will be our times greater

0.5 × 1.225 × 104 × 104 = 6625 N/m

2

(66.25 hPa)

I the cross-sectional area o the tube is 1 m a orce o ½ ρ V 2

2

newtons will be generated.

(Force = Pressure × Area) Dynamic pressure ( ½ ρ V ) is common to ALL aerodynamic orces and determines the air loads imposed on an aeroplane moving through the air. 2

The symbol or dynamic pressure ( ½ ρ V ) is q or Q 2

Q = ½ ρ V2

28

2

The Atmosphere Key Facts 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 o static and dynamic pressure, in this context, is known as ‘Total’ pressure.

2


(Dynamic + Static pressure can also be reerred to as Stagnation or Pitot pressure). Total Pressure = Static Pressure + Dynamic Pressure This can be re-arranged to show that: Total Pressure - Static Pressure = Dynamic Pressure The significance o dynamic pressure to the understanding o Principles o Flight cannot be overemphasized. Because dynamic pressure is dependent upon air density and the speed o the aircraf through the air, it is necessary or students to ully appreciate the actors which affect air density. • Temperature - increasing temperature decreases air density. Changes in air density due to air temperature are significant during all phases o flight. • Static pressure - decreasing static pressure decreases air density. Changes in air density due to static pressure are significant during all phases o flight. • Humidity - increasing humidity decreases air density. (The reason increasing humidity decreases air density is that the density o water vapour is about 5/8 that o dry air). Humidity is most significant during take-off and landing. Increasing altitude will decrease air density because the effect o decreasing static pressure is more dominant than decreasing temperature.

29

2

The Atmosphere Measuring Dynamic Pressure

2

All aerodynamic orces acting on an aircraf are determined by dynamic pressure, so it is essential to have some means o measuring dynamic pressure and presenting that inormation to the pilot.


A sealed tube, open at the orward end, is located where it will collect air when the aircraf is moving. The pressure in the tube (pitot tube) is Dynamic + Static and, in this context, is called “Pitot” pressure (because the air is inside the pitot tube). Some way o ‘removing’ the static pressure rom the pitot pressure must be ound. A hole (vent) in a surace parallel to the airflow will sense static pressure. Reerring to the diagram below, i the pressure rom the pitot tube is ed to one side o a diaphragm mounted in a sealed case, and static pressure is ed 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 o 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 o the speed o the aircraf through the air when the air density is 1.225 kg/m . 3

This is not a problem because the pilot needs an indication o dynamic pressure, and this is what the instrument provides. The instrument is made in such a way that it indicates the square root o the dynamic pressure in nautical miles per hour (knots) or statute miles per hour (mph). So, i this “Indicated Airspeed ” is doubled, the speed o the aircraf through the air will also be doubled.

The Airspeed Indicator is a pressure gauge Airflow

PITOT TUBE

PITOT PRESSURE (Static + Dynamic)

Needle indicates changes in DYNAMIC PRESSURE

Airflow

STATIC VENT

STATIC PRESSURE

Figure 2.1 Schematic o the airspeed indicator (ASI)

30

2

The Atmosphere Relationships between Airspeeds Indicated Airspeed: (IAS). The speed registered on the Airspeed Indicator.

2


Calibrated Airspeed: (CAS). An accurate measure o dynamic pressure when the aircraf is

flying slowly. The position o the pitot tube(s) and static vent(s), together with the aircraf’s configuration (flaps, landing gear etc.) and attitude to the airflow (angle o attack and sideslip) will affect the pressures sensed, particularly the pressures sensed at the static vent(s) . Under the influence o the above conditions a alse dynamic pressure (IAS) will be displayed. When IAS is corrected or this ‘position’ or ‘pressure’ error, as it is called, the resultant is Calibrated Airspeed. (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 Airspeed : (EAS). An accurate measure o dynamic pressure when the aircraf is

flying ast. Air entering the pitot tube(s) is compressed, which gives a alse dynamic pressure (IAS) reading, but only becomes significant at higher speeds. At a given air density, the amount o compression depends on the speed o the aircraf through the air. When the IAS is corrected or ‘position’ AND ‘compressibility’ error, the resultant is Equivalent Airspeed. True Airspeed: (TAS) or (V). The speed o the aircraf through the air. THE ONLY SPEED

THERE IS - All the other, so called, speeds are pressures. EAS Where, б is Relative Density √б The Airspeed Indicator is calibrated or ‘standard’ sea level density, so it will only read TAS i the density o the air through which the aircraf is flying is 1.225 kg/m . Thus at 40 000 f where the ‘standard’ density is one quarter o the sea-level value, to maintain the same EAS the aircraf TAS =

3

will have to move through the air twice as ast! The Speed o Sound: (a)

Sound is ‘weak’ pressure waves which propagate spherically through the atmosphere rom their source. The speed at which pressure waves propagate is proportional to the square root o the absolute temperature o the air . The lower the temperature, the lower the speed o propagation. On a ‘standard’ day at sea level the speed o sound is approximately 340 m/s (660 kt TAS). At higher aircraf True Airspeeds (TAS) and/or higher altitudes, it is essential to know the speed o the aircraf in relation to the local speed o sound. This speed relationship is known as the Mach Number (M). M =

TAS (a)

Where (a) is the Local Speed o Sound

I the True Airspeed o the aircraf is our tenths the speed at which pressure waves propagate through the air mass surrounding the aircraf, the Mach meter will register M 0.4 Critical Mach Number: (MCRIT) The critical Mach number is the Mach number o the aircraf

when the speed o the airflow over some part o the aircraf (usually the point o maximum thickness on the aerooil) first reaches the speed o sound.

31

2

The Atmosphere Airspeed

2

This inormation is to reinorce that contained in the preceding paragraphs.


The airspeed indicator is really a pressure gauge, the ‘needle’ o which responds to changes in dynamic pressure (½ ρ V ). 2

The Airspeed Indicator is a Pressure Gauge Calibration o the airspeed indicator is based on standard sea level density (1.225 kg/m ). The “airspeed” recorded will be different rom the actual speed o the aircraf through the air unless operating under standard sea level conditions (unlikely). The actual speed o the aircraf relative to the ree stream is called true airspeed (TAS), and denoted by (V). The ‘speed’ recorded by the airspeed indicator calibrated as above, i there are no other errors, is called equivalent airspeed (EAS). 3

It may seem to be a drawback that the instrument records equivalent rather than true airspeed, but the true airspeed may always be determined rom it. Also, many o the handling characteristics o an aircraf depend mainly on the dynamic pressure, i.e. on the equivalent airspeed, so it is ofen more useul to have a direct reading o EAS than TAS.

Errors and Corrections An airspeed indicator is, however, also subject to errors other than that due to the difference between the density o the air through which it is flying and standard sea level density. • Instrument Error : This error may arise rom the imperections in the design and manuacture o the instrument, and varies rom one instrument to another. Nowadays this type o error is usually very small and or all practical purposes can be disregarded. Where any instrument error does exist, it is incorporated in the calibrated airspeed correction chart or the particular aeroplane. • Position Error (Pressure Error) : This error is o 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 aircraf where the flow is affected by the presence o the aircraf, changes in configuration (flaps and maybe gear) and proximity to the ground (ground effect). I so, the static pressure recorded will be the local and not the ree stream value. The pitot pressure may be under-recorded because o incorrect alignment - the tube(s) may be inclined to the airstream instead o acing directly into it (changes in angle o attack, particularly at low speeds). The magnitude o the consequent errors will generally depend on the angle o attack and, hence, the speed o the aircraf. • Compressibility Error : At high speeds, the dynamic pressure is not simply ½ ρ V , but exceeds it by a actor determined by Mach number. Thus the airspeed indicator will over-read. 2

Because o 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 rectiy the instrument and position errors are determined experimentally. In flight, using special instruments, measurements are taken over the whole range o speeds and configurations, rom which a calibration curve is obtained which gives the corrections appropriate to each indicated airspeed. The compressibility error correction may be obtained by calculation. 32

2

The Atmosphere The indicated airspeed, afer correction or instrument, position (pressure) and compressibility errors, gives the equivalent airspeed ½ ρ V . 2

2

V Speeds


These include: V S , V 1 , V R , V 2 , V MD , V MC , V YSE and many others - these are all Calibrated Airspeeds because they relate to aircraf operations at low speed. However, the appropriate corrections are made and these speeds are supplied to the pilot in the Flight Manual as IAS. VMO - The maximum operating IAS is, however, an EAS because it is a high speed, but again it is supplied to the pilot in the Flight Manual as an IAS.

33

2

The Atmosphere Summary

2

Dynamic pressure (Q) is affected by changes in air density.


Q = ½ ρ V2

Air density decreases i atmospheric pressure decreases. Air density decreases i air temperature increases. Air density decreases i humidity increases. With the aircraf on the ground: Taking off rom 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 (IAS). For the purpose o general understanding: A constant IAS will give constant dynamic pressure. Increasing altitude decreases air density because o decreasing static pressure. With the aircraf airborne: As altitude increases, a higher TAS is required to maintain a constant dynamic pressure. Maintaining a constant IAS will compensate or changes in air density. There is only one speed, the speed o the aircraf through the air, the TAS. All the other, so called, speeds are pressures. The Airspeed Indicator is a pressure gauge. Aircraf ‘V’ speeds are CAS, except V MO which is an EAS, but all are presented to the pilot in the Flight Manual as IAS.

34

2

Questions Questions 1.

When considering air:

2

s n o i t s e u Q

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 o air is very high. 5 - moving air has kinetic energy. The correct combination o all true statements is:

a. b. c. d. 2.

Why do the lower layers contain the greater proportion o the whole mass o the atmosphere?

a. b. c. d. 3.

1, 2, 3 and 5. 2, 3 and 4. 1 and 4. 1, 3, and 5.

Because air is very viscous. Because air is compressible. Because o greater levels o humidity at low altitude. Because air has very little mass.

With increasing altitude, up to about 40 000 f, the characteristics o air change: 1 - temperature decreases continuously with altitude. 2 - pressure alls steadily to an altitude o about 36 000 f, where it then remains constant. 3 - density decreases steadily with increasing altitude. 4 - pressure alls steadily with increasing altitude. The combination o true statements is:

a. b. c. d. 4.

3 and 4. 1, 2 and 3. 2 and 4. 1 and 4.

When considering static pressure: 1 - in aviation, static pressure can be measured in hectopascals. 2 - the SI unit or static pressure is N/m 2. 3 - static pressure is the product o the mass o air pressing down on the air beneath. 4 - reerred to as static pressure because o 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.

2, 4 and 5. 1, 2, 3, 4 and 5. 1, 3 and 5. 1 and 5.

35

2

Questions 5.

When considering air density: 1 - density is measured in millibars. 2 - density increases with increasing altitude. 3 - i temperature increases, the density will increase. 4 - as altitude increases, density will decrease. 5 - temperature decreases with increasing altitude, and this will cause air density to increase

2

Q u e s t i o n s

The combination o correct statements is:

a. b. c. d. 6.

Air density is:

a. b. c. d. 7.

4 only. 4 and 5. 5 only. 2, 3 and 5.

mass per unit volume. proportional to temperature and inversely proportional to pressure. independent o both temperature and pressure. dependent only on decreasing pressure with increasing altitude.

When considering the ICAO International Standard Atmosphere and comparing it with the actual atmosphere, which o the ollowing statements is correct? 1 - Temperature, pressure and density are constantly changing in any given layer o the actual atmosphere. 2 - A requirement exists or a hypothetical ’standard’ atmosphere. 3 - The values given in the International Standard Atmosphere exist at the same altitudes in the actual atmosphere. 4- The International Standard Atmosphere was designed or the calibration o pressure instruments and the comparison o aircraf perormance 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 o the ollowing statements is correct? 1 - The temperature lapse rate is assumed to be uniorm at 2°C per 1000 f (1.98°C) up to a height o 11 000 f. 2 - Sea level temperature is assumed to be 15°C. 3 - Sea level static pressure is assumed to be 1.225 kg/m 3. 4 - Sea level density is assumed to be 1013.25 hPa.

a. b. c. d.

36

1, 2, 3 and 4. No statements are correct. 1, 3 and 4. 2 only.

2

Questions 9.

A moving mass o air possesses kinetic energy. An object placed in the path o such a moving mass o air will be subject to which o the ollowing?

a. b. c. d. 10.

b. c. d.

b. c. d.

on a part o the aircraf structure where the airflow is undisturbed, in a surace at right angles to the airflow direction. on a part o the structure where the airflow is undisturbed, in a surace parallel to the airflow direction. at the stagnation point. at the point on the surace where the airflow reaches the highest speed.

The inputs to an Airspeed Indicator are rom:

a. b. c. d. 15.

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.

14.

density times speed squared. hal the density times the indicated airspeed squared. hal the true airspeed times the density squared. hal the density times the true airspeed squared.

A tube acing 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 o 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.

s n o i t s e u Q

Dynamic pressure is:

a.

11.

2

Dynamic pressure. Static pressure. Static pressure and dynamic pressure. Dynamic pressure minus static pressure.

a static source. pitot pressure. a pitot and a static source. pitot, static and density.

The deflection o the pointer o the Airspeed Indicator is proportional to:

a. b. c. d.

dynamic pressure. static pressure. the difference between static and dynamic pressure. static pressure plus dynamic pressure.

37

2

Questions 16.

Calibration o the Airspeed Indicator is based upon the density:

a. b. c. d.

2

Q u e s t i o n s

17.

at the altitude at which the aircraf 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 o airspeed: 1 - True Airspeed (TAS) is read directly rom the Airspeed Indicator. 2 - Equivalent Airspeed is Indicated Airspeed corrected or position error. 3 - Indicated Airspeed is not a speed at all, it is a pressure. 4 - True Airspeed is the speed o the aircraf through the air. Which o 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 o Airspeed: 1 - Calibrated Airspeed is Indicated Airspeed corrected or position error. 2 - Equivalent Airspeed is Indicated Airspeed corrected or position error & compressibility. 3 - Position error, which causes alse Indicated Airspeed readings, is due to variations in the pressures sensed at the pitot and static ports. 4 - The Airspeed Indicator is calibrated to read True Airspeed when the ambient density is that o the ICAO International Standard Atmosphere at sea level. The combination o correct statements is:

a. b. c. d. 19.

The speed o sound:

a. b. c. d. 20.

is dependent upon the True Airspeed and the Mach number o the aircraf. is inversely proportional to the absolute temperature. is proportional to the square root o the absolute temperature o the air. is directly proportional to the True Airspeed o the aircraf.

Mach number is:

a. b. c. d.

38

none o the statements are correct. 1, 2 and 4. 2 and 3. 1, 2, 3 and 4.

the aircraf True Airspeed divided by the local speed o sound. the speed o sound in the ambient conditions in which the aircraf is flying. the True Airspeed o the aircraf at which the relative airflow somewhere on the aircraf first reaches the local speed o sound. the Indicated Airspeed divided by the local speed o sound sea level.

2

Questions 21.

An aircraf’s critical Mach number is:

a. b. c. d.

the speed o the airflow when the aircraf first becomes supersonic. the speed o the aircraf when the airflow somewhere reaches the speed o sound. the Indicated Airspeed when the aircraf first becomes supersonic. the aircraf’s Mach number when airflow over it first reaches the local speed o sound.

2

s n o i t s e u Q

39

2

Answers

Answers 2

A n s w e r s

40

1 d

2 b

3 a

4 b

5 a

6 a

7 d

8 d

9 c

13 b

14 c

15 a

16 b

17 c

18 d

19 c

20 a

21 d

10 b

11 d

12 c

Chapter

3 Basic Aerodynamic Theory

The Principle o Continuity . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .43 Bernoulli’s Theorem. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .44 Streamlines and the Streamtube. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 45 Summary. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .46 Questions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

47

Answers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

50

41

3

Basic Aerodynamic Theory

3

B a s i c A e r o d y n a m i c T h e o r y

42

3

Basic Aerodynamic Theory The Principle of Continuity One o the undamental laws o the universe is ENERGY and MASS can neither be created nor destroyed, only changed rom one orm to another. To demonstrate the effect this basic Principle o Continuity has on aerodynamic theory, it is instructive to consider a streamline flow o air through a tube which has a reduced cross-sectional area in the middle.

3

y r o e h T c i m a n y d o r e A c i s a B

The air mass flow, or mass per unit time, through the tube will be the product o the crosssectional area (A), the airflow velocity (V) and the air density ( ρ). Mass flow will remain a constant value at all points along the tube. The Equation o Continuity is: A × V × ρ = Constant

Because air is a compressible fluid, any pressure change in the flow will affect the air density. However, at low subsonic speeds (< M 0.4) density changes will be insignificant and can be disregarded. The equation o continuity can now be simplified to: A × V = constant, or: Velocity (V)

=

Constant Area (A)

Airflow

Cross-sectional Area (A)

1m

½ m

Velocity (V)

52 m/s (100 kt)

104 m/s (200 kt)

Mass Flow (Constant)

3

3

52 m /s

3

3

52 m /s

3

1m

52 m/s (100 kt) 3

52 m /s

Figure 3.1 The principle o continuity

Because the mass flow must remain constant, it can be seen rom the equation o continuity that the reduction in the tube’s cross-sectional area results in an increase in velocity, and vice versa. The equation o continuity enables the velocity changes o airflow around a given shape to be predicted mathematically, (< M 0.4).

43

3

Basic Aerodynamic Theory Bernoulli’s Theorem “In the steady flow o an ideal fluid the sum o the pressure energy and the kinetic energy remains constant”.

3

Note: An ideal fluid is both incompressible and has no viscosity.


This statement can be expressed as: Pressure + Kinetic energy = Constant or: p + 1/2 ρ V2 = Constant

Consider a mass o air: Static Pressure 101 325 N/m , Density 1.225 kg/m and Velocity 52 m/s, its dynamic pressure will be: 1656 N/m . [Q = ½ × 1.225 × 52 × 52] 2

3

2

Pressure (101 325 N/m ) + Kinetic energy (1656 N/m ) = Constant (102 981 N/m ) 2

2

2

Figure 3.2 Bernoulli’s Theorem

Because the velocity o air at the throat has doubled, its dynamic p ressure has risen by a value o our, and the static pressure has decreased. The significant point is that: Static Pressure + Dynamic Pressure is a constant. This constant can be reerred to either as: TOTAL PRESSURE, STAGNATION PRESSURE or PITOT PRESSURE.

It can be seen that flow velocity is dependent on the shape o the object over which it flows. And rom Bernoulli’s theorem, it is evident that an increase in velocity will cause a decrease in static pressure, and vice versa.

44

3

Basic Aerodynamic Theory The tubes illustrated above are used only to demonstrate the principle o continuity and Bernoulli’s theorem and are o no practical use in making an aeroplane fly. But an aerodynamic orce to oppose the weight o an aircraf can be generated by using a specially shaped body called an aerooil.

3

y r o e h T c i m a n y d o r e A c i s a B

Figure 3.3 Typical aerooil section

The airflow velocity over the top surace o a lifing aerooil will be greater than that beneath, so the pressure differential that results will produce a orce per unit area acting upwards. The larger the surace area, the bigger the orce that can be generated. In the next section we see that the flow over the top o the aerooil looks very like the tube on the opposite page and the principle o continuity and Bernoulli’s theorem still apply.

Streamlines and the Streamtube A streamline is the path traced by a particle o 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.

Figure 3.4 Streamlines & a streamtube

A streamtube is an imaginary tube made o streamlines. There is no flow into or out o the streamtube through the “walls”, only a flow along the tube. With this concept it is possible to visualize the airflow around an aerooil being within a tube made up o streamlines.

45

3

Basic Aerodynamic Theory Summary At flow speeds o less than about M 0.4, pressure changes will not affect air density.

3

Continuity:


• Air accelerates when the cross-sectional area o a streamline flow is reduced. • Air decelerates when the cross-sectional area increases again. Bernoulli:

• I a streamline flow o air accelerates, its kinetic energy will increase and its static pressure will decrease. • When air decelerates, the kinetic energy will decrease and the static pressure will increase again. By harnessing the principle o continuity and Bernoulli’s theorem an aerodynamic orce can be generated.

46

3


I the cross-sectional area o an airflow is mechanically reduced:

a. b. c. d. 2.

d.

Bernoulli’s theorem. the principle o continuity. Newton’s second law o motion. the Magnus effect.

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 o the ollowing statements is correct? 1. 2. 3. 4.

a. b. c. d. 5.

s n o i t s e u Q

I the velocity o an air mass is increased:

a. b. c.

4.

3

The statement, Pressure plus Kinetic energy equals constant, reers to:

a. b. c. d. 3.

the velocity o the airflow remains constant and the kinetic energy increases. the velocity o 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 o the airflow increases.

A resultant decrease in static pressure is indicated by streamlines shown close together. An increase in velocity is indicated by streamlines shown close together. Accelerating airflow with a resultant decreasing static pressure is indicated by converging streamlines. Diverging streamlines indicate decelerating airflow with a resultant increasing static pressure.

2 and 4. 1, 3 and 4. 2, 3 and 4. 1, 2, 3 and 4.

I the pressure on one side o a surace is lower than on the other side:

a. b. c. d.

a orce per unit area will exist, acting in the direction o the lower pressure. no orce will be generated, other than drag. a orce will be generated, acting in the direction o the higher pressure. the pressure will leak around the sides o the surace, cancelling out any pressure differential.

47

3

Questions 6.

When considering a streamtube, which o the ollowing statements is correct? 1.

4.

Different sizes o stream tube are manuactured to match the wingspan o the aircraf to which they will be fitted. A streamtube is a concept to aid understanding o aerodynamic orce generation. There is no flow into or out o the streamtube through the “walls”, only flow along the tube. A streamtube is an imaginary tube made up o streamlines.

a. b. c. d.

1 only. 1 and 3. 2, 3 and 4. 1, 2 and 3.

3

2.

Q u e s t i o n s

3.

7.

In accordance with the principle o continuity: 1. 2. 3. 4.

air accelerates when the cross-sectional area o a streamline flow is reduced. when air accelerates the density o air in a streamline flow is increased. air decelerates when the cross-sectional area o a streamline flow is increased. changes in cross-sectional area o a streamline flow will affect the air velocity.

Which o the preceding statements are true?

a. b. c. d. 8.

1, 2, 3 and 4. 1 and 4. 3 and 4. 1, 3 and 4.

In accordance with Bernoulli’s theorem: 1. 2. 3. 4.

i a streamline flow o air decelerates, its kinetic energy will decrease and the static pressure will increase. i a streamline flow o air accelerates, its kinetic energy will increase and the static pressure will decrease. i a streamline flow o air is accelerated, the dynamic pressure will increase and the static pressure will increase. i a streamline flow o air is decelerated, its dynamic pressure will decrease and the static pressure will increase.

The combination o correct statements is:

a. b. c. d. 9.

The statement, “Energy and mass can neither be created nor destroyed, only changed rom one orm to another”, reers to:

a. b. c. d.

48

1, 2, 3 and 4. 3 only. 1, 2 and 4. 3 and 4.

Bernoulli’s theorem. the equation o kinetic energy. the principle o continuity. Bernoulli’s principle o continuity.

3

Questions

3

s n o i t s e u Q

49

3

Answers

Answers 1 d

3

A n s w e r s

50

2 a

3 c

4 d

5 a

6 c

7 d

8 c

9 c

Chapter

4 Subsonic Airflow

Aerooil Terminology . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .52 Basics about Airflow . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .54 Two Dimensional Airflow. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .54 Summary. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .62 Questions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

63

Answers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

68

51

4

Subsonic Airflow

MAXIMUM THICKNESS LOCATION OF MAX. THICKNESS

MAXIMUM LEADING EDGE

4

CAMBER

RADIUS

S u b s o n i c A i r fl o w

MEAN CAMBER LINE CHORD LINE LEADING TRAILING EDGE

EDGE

CHORD

LOCATION OF MAX. CAMBER

LIFT

TOTAL REACTION

ANGLE OF ATTACK

DRAG

RELATIVE AIRFLOW

AIRCRAFT FLIGHT PATH

Figure 4.1

Aerofoil Terminology Aerofoil A shape capable o producing lif with relatively high efficiency.

Chord Line A straight line joining the centres o curvature o the leading and trailing edges o an aerooil.

Chord The distance between the leading and trailing edges measured along the chord line.

Angle of Incidence The angle between the wing root chord line and the longitudinal axis o the aircraf. (This angle is fixed or the wing, but may be variable or the tailplane).

52

4

Subsonic Airflow Mean Line or Camber Line A line joining the leading and trailing edges o an aerooil, equidistant rom the upper and lower suraces.

Maximum Camber The maximum distance o the mean line rom the chord line. Maximum camber is expressed as a percentage o the chord, with its location as a percentage o the chord af o the leading edge. When the camber line lies above the chord line the aerooil is said to have positive camber, and i the camber line is below the chord line, it is said to have negative camber. A symmetrical aerooil has no camber because the chord line and camber line are coincidental.

4

w o fl r i A c i n o s b u S

Thickness/Chord Ratio The maximum thickness or depth o an aerooil section expressed as a percentage o the chord, with its location as a percentage o the chord af o the leading edge. The thickness and thickness distribution o the aerooil section have a great influence on its airflow characteristics.

Leading Edge Radius The radius o curvature o the leading edge. The size o the leading edge radius can significantly affect the initial air flow characteristics o the aerooil section.

Relative Airflow (Relative Wind or Free Stream Flow): Relative Airflow has three qualities. • DIRECTION - air parallel to and in the opposite direction to the flight path o the aircraf, in act the path o the CG; the direction in which the aircraf is pointing is irrelevant. • CONDITION - air close to, but unaffected by the presence o, the aircraf; its pressure, temperature and velocity are not affected by the passage o the aircraf through it. • MAGNITUDE - The magnitude o the Relative Airflow is the TAS. I Airflow does not possess all three o these qualities, it is reerred to as EFFECTIVE AIRFLOW.

Total Reaction The resultant o all the aerodynamic orces acting on the aerooil section.

Centre of Pressure (CP) The point on the chord line through which lif is considered to act.

Lift The aerodynamic orce which acts at 90° to the Relative Airflow.

Drag The aerodynamic orce which acts parallel to and in the same direction as the Relative Airflow (or opposite to the aircraf flight path).

Angle of Attack (α or alpha) (can also be reerred to as Aerodynamic Incidence). The angle between the chord

line and the Relative Airflow. The angle between the chord line and the effective airflow is reerred to as the EFFECTIVE ANGLE OF ATTACK .

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Subsonic Airflow Basics about Airflow When considering airflow velocity, it makes no difference to the pressure pattern i the aircraf is moving through the air or the air is flowing over the aircraf: it is the relative velocity which is the important actor. To promote a ull understanding, reerences will be made to both wind tunnel experiments, where air is flowing over a stationary aircraf, and aircraf in flight moving through ‘stationary’ air.

4


Three dimensional airflow : Three dimensional flow is the true airflow over an aircraf and

consists o 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 aerooil section along the entire

span with no spanwise pressure differential or flow.

Two Dimensional Airflow This CONCEPT, Figure 4.2 and Figure 4.3, is used to illustrate the basic principles o aerodynamic orce generation. As Airflows towards an aerooil it will be turned towards the lower pressure at the upper surace; this is termed upwash. Afer passing over the aerooil, the airflow returns to its original position and state; this is termed downwash.

Figure 4.2 INCREASED LOCAL VELOCITY DOWNWASH

UPWASH

Figure 4.3

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Subsonic Airflow Influence of Dynamic Pressure Figure 4.4 shows an aerooil section at a representative angle o attack subject to a given

dynamic pressure (IAS). “I the static pressure on one side o a body is reduced more than on the other side, a pressure differential will exist”. Figure 4.5 shows the same aerooil section at the same angle o attack, but subject to a higher

4

dynamic pressure (IAS). “I the dynamic pressure (IAS) 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 surace area will produce an upward acting orce. “I the dynamic pressure (IAS) is increased, the upward orce will increase”.

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Subsonic Airflow Influence of Angle of Attack At a constant dynamic pressure (IAS), increasing the angle o attack (up to about 16°) will likewise increase the pressure differential, but it will also change the pattern o pressure distribution. The aerooil profile presented to the airflow will determine the distribution o velocity and hence the distribution o pressure on the surace. This profile is determined by the aerooil geometry, i.e. thickness and distribution (fixed), camber and distribution (assumed to be fixed or now) and by the angle o attack (variable).

4


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 o attack increases rom -4°, the leading edge stagnation point moves rom the upper surace around the leading edge to the lower surace. It is at the ront stagnation point where the flow divides to pass over and under the section. The pressure at the stagnation point (stagnation pressure) is Static + Dynamic. The flow over the top o the section accelerates rapidly around the nose and over the leading portion o the surace - inducing a substantial decrease in static pressure in those areas. The rate o acceleration increases with increase in angle o attack, up to about 16°. (Anything which changes the accurately manuactured profile o the leading portion o the surace can seriously disrupt airflow acceleration in this critical area e.g. ice, snow, rost, dirt or dents). The pressure reduces continuously rom the stagnation value through the ree stream value to a position on the top surace where a peak negative value is reached. From there onwards the flow continuously slows down again and the pressure increases back to the ree stream value in the region o the trailing edge. At angles o 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 ree stream value in the region o the trailing edge. The pressure differential between the leading edge stagnation point and the lower pressure at the trailing edge creates a orce acting backward which is called ‘orm (pressure) drag’. (This will be discussed in more detail later).

Angle of Attack (-4°) The decrease in pressure above and below the section are equal and no differential exists. There will, thus, be no lif orce. ( Figure 4.6 ). This can be called the “zero lif angle o attack”.

Figure 4.6

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4

Subsonic Airflow Angles o Attack (0° to 8°)

Compared to ree stream static pressure, there is a pressure decrease over the upper surace and a lesser decrease over most o the lower surace. For a cambered aerooil there will be a small amount o lif even at small negative angles (-4° to 0°).

4


Angles o attack (0° to 16°)

Increasing the angle o attack increases the lif orce because the acceleration o the airflow over the top surace is increased by the reduction in effective cross-sectional area o the local streamtube. The reduced pressure ‘peak’ moves orward as the angle o attack increases. The greatest contribution to overall lif comes rom the upper surace. Pressure Gradient

This is a change in air pressure over distance. The greater the difference in pressure between two points, the steeper the gradient. A avourable gradient is when air pressure is alling in the direction o airflow. An adverse pressure gradient is when air pressure is rising in the direction o airflow, such as between the point o minimum pressure on the top surace and the trailing edge. The higher the angle o attack, the steeper the pressure gradient. At angles o attack higher than approximately 16° , the extremely steep adverse pressure gradient prevents air that is flowing over the top surace rom ollowing the aerooil contour, and the previously smooth streamline flow will separate rom the surace, causing the low pressure area on the top o the section to suddenly collapse. Any pressure differential remaining is due to the pressure increase on the lower surace only. This condition is known as the stall and will be described in detail in Chapter 7 .

Figure 4.7

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Subsonic Airflow Centre of Pressure (CP) The whole surace o the aerooil contributes to lif, but the point along the chord where the distributed lif is effectively concentrated is termed the Centre o Pressure ( Figure 4.8). The location o the CP is a unction o camber and section lif coefficient, i.e. angle o attack.

4


Figure 4.8

Movement of the Centre of Pressure As the angle o attack increases rom 0° to 16° the upper ‘suction’ peak moves orward ( Figure 4.7 ), so the point at which the lif is effectively concentrated, the CP, will move orward. The CP moves orward and the magnitude o the lif orce increases with increase in angle o attack until the stall is reached when the lif orce decreases abruptly and the CP generally moves back along the chord ( Figure 4.9). Note that the CP is at its most orward location just beore the stall (CL MAX).

Aerodynamic Force Coefficient A coefficient is a dimensionless number expressing degree o magnitude. An aerodynamic orce coefficient is a common denominator or all A/C o whatever weight, size and speed. An aerodynamic orce coefficient is a dimensionless ratio between the average aerodynamic pressure and the airstream dynamic pressure. By this definition a lif coefficient (C L ) is the ratio between lif divided by the wing planorm area and dynamic pressure and a drag coefficient (C D) is the ratio between drag divided by the wing planorm area and dynamic pressure. The use o the coefficient o an aerodynamic orce is necessary since the orce coefficient is: • An index o the aerodynamic orce independent o area, density and velocity. It is derived rom the relative pressure and velocity distribution. • Influenced only by the shape o the surace and angle o attack since these actors determine the pressure distribution.

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4

Subsonic Airflow

C L M AX K C A T T A F O

4


E L G N A

0

10%

20%

30%

40%

50%

60%

70%

80%

90%

100%

CP POSITION (Percentage chord, aft of leading edge)

TRAILING EDGE

LEADING EDGE

Figure 4.9 CP movement with angle o attack

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Subsonic Airflow Development of Aerodynamic Pitching Moments The distribution o pressure over a surace is the source o aerodynamic moments as well as orces. There are two ways to consider the effects o changing angle o attack on the pitching moment o an aerooil. • Changes in the magnitude o lif acting through a moving CP, or more simply:

4

• Changes in the magnitude o lif 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 o the lif orce effectively take place’, AND: ‘The point about which the pitching moment will remain constant at ‘normal’ angles o attack’. A nose-down pitching moment exists about the AC which is the product o a orce (lif at the CP) and an arm (distance rom the CP to the AC). Since an increase in angle o attack will increase the lif orce, but also move the CP towards the AC (shortening the lever arm), the moment about the AC remains the same at any angle o attack within the “normal” range. L1

1

M CP

AC

L2

d1

2

M AC

CP

d2

Figure 4.10

When considering subsonic airflows o less than M 0.4, the AC is located at the 25% chord point or any aerooil regardless o camber, thickness and angle o attack. The aerodynamic centre (AC) is an aerodynamic reerence point, the most direct application being to the longitudinal stability o an aircraf, which will be discussed in Chapter 10.

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Subsonic Airflow Pitching Moment for a Symmetrical Aerofoil Note the change in pressure distribution with angle o attack or the symmetrical aerooil in Figure 4.11. When at zero angle o attack, the upper and lower surace orces are equal and located at the same point. With an increase in angle o attack, the upper surace orce increases while the lower surace orce decreases. A change in the magnitude o lif has taken place with no change in the CP position - a characteristic o symmetrical aerooils. Thus, the pitching moment about the AC or a symmetrical aerooil will be zero at ‘normal’ angles o attack - one o the big advantages o symmetrical aerooils.

4


SYMMETRICAL AEROFOIL AT ZERO LIFT

AC

SYMMETRICAL AEROFOIL AT POSITIVE LIFT

AC

Figure 4.11

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Subsonic Airflow Summary Airflow pattern, and ultimately lif and drag, will depend upon: • • • •

4


Angle o attack - airflow cross-sectional area change Aerooil shape (thickness & camber) - airflow cross-sectional area change Air density - mass flow o air (decreases with increased altitude) Velocity - mass flow o air (changes with aircraf TAS)

The lif orce is the result o the pressure differential between the top and bottom suraces o an aerooil; the greatest contribution to overall lif comes rom the top surace. Anything (ice in particular, but also rost, snow, dirt, dents and even water droplets) which changes the accurately manuactured profile o the leading portion o the upper surace can seriously disrupt airflow acceleration in that area, and hence the magnitude o the lif orce will be affected. An increase in dynamic pressure (IAS) will increase the lif orce, and vice versa. An increase in angle o attack will increase the lif orce, and vice versa, (0° to 16°) The centre o pressure (CP) o a cambered aerooil moves orward as the angle o attack increases. The CP o a symmetrical aerooil does not move under the influence o angle o attack (within the confines o ‘normal range’). Throughout the normal range o angles o attack, the aerooil nose-down pitching moment about the aerodynamic centre (AC) will remain constant. The AC is located at the quarter chord position or subsonic flow o less than M 0.4. The coefficient o lif (C L ) is the ratio between lif per unit wing area and dynamic pressure. As the angle o attack increases rom -4°, the leading edge stagnation point moves rom the upper surace around the leading edge to the lower surace. The greatest positive pressure occurs at the leading edge stagnation point, where the relative flow velocity is zero. Form (pressure) drag is the result o the pressure differential between the leading edge and trailing edge o the aerooil. An increase in dynamic pressure (IAS) will increase orm drag, and vice versa. The coefficient o drag (C D ) is the ratio between drag per unit wing area and dynamic pressure.

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With reerence to aerooil section terminology, which o the ollowing statements are true? 1. 2. 3. 4.

a. b. c. d. 2.

b. c. d.

d.

1, 2, 3 and 4. 1, 2 and 4. 2, 3 and 4. 2 and 4.

the aerodynamic orce which acts perpendicular to the chord line o the aerooil. the aerodynamic orce that results rom the pressure differentials about an aerooil. the aerodynamic orce which acts perpendicular to the upper surace o the aerooil. the aerodynamic orce which acts at 90° to the relative airflow.

negative air pressure below and a vacuum above the surace. vacuum below the surace and greater air pressure above the surace. higher air pressure below the surace and lower air pressure above the surace. higher air pressure at the leading edge than at the trailing edge.

On an aerooil section, the orce o lif acts perpendicular to, and the orce o drag acts parallel to the:

a. b. c. d. 5.

s n o i t s e u Q

An aerooil section is designed to produce lif resulting rom a difference in the:

a. b. c.

4.

4

The definition o lif is:

a.

3.

The chord line is a line joining the centre o curvature o the leading edge to the centre o the trailing edge, equidistant rom the top and bottom surace o the aerooil. The angle o incidence is the angle between the chord line and the horizontal datum o the aircraf. The angle between the chord line and the relative airflow is called the aerodynamic incidence or angle o attack. The thickness/chord ratio is the maximum thickness o the aerooil as a percentage o the chord; the location o maximum thickness is measured as a percentage o the chord af o the leading edge.

flight path. longitudinal axis. chord line. aerooil section upper surace.

When the angle o attack o a symmetrical aerooil is increased, the centre o pressure will:

a. b. c. d.

have very limited movement. move af along the aerooil surace. remain unaffected. move orward to the leading edge.

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4

Questions 6.

Why does increasing speed also increase lif?

a. b. 4

c. d.

Q u e s t i o n s

7.

The point on an aerooil section through which lif acts is the:

a. b. c. d. 8.

greater than atmospheric pressure. equal to atmospheric pressure. less than atmospheric pressure. non existent.

The angle o attack o an aerooil section directly controls:

a. b. c. d.

64

chord line. camber. mean camber line. longitudinal axis.

At zero angle o attack, the pressure along the upper surace o a symmetrical aerooil section would be:

a. b. c. d. 12.

incidence. lif. attack. sweepback

A line drawn rom the leading edge to the trailing edge o an aerooil section and equidistant at all points rom the upper and lower contours is called the:

a. b. c. d. 11.

the angle o attack. the angle o incidence. dihedral. sweepback.

The angle between the chord line o an aerooil section and the relative wind is known as the angle o:

a. b. c. d. 10.

midpoint o the chord. centre o gravity. centre o pressure. aerodynamic centre.

The angle between the chord line o the aerooil section and the longitudinal axis o the aircraf is known as:

a. b. c. d. 9.

The increased impact o the relative wind on an aerooil’s lower surace creates a greater amount o air being deflected downward. The increased speed o the air passing over an aerooil’s upper surace decreases the static pressure, thus creating a greater pressure differential between the upper and lower surace. The increased velocity o the relative wind overcomes the increased drag. Increasing speed decreases drag.

the amount o airflow above and below the section. the angle o incidence o the section. the distribution o positive and negative pressure acting on the section. the angle relative to the horizontal datum

4

Questions 13.

When the angle o attack o a positively cambered aerooil is increased, the centre o pressure will:

a. b. c. d. 14.

b. c. d.

ormed by the longitudinal axis o the aeroplane and the chord line o the section. between the section chord line and the relative wind. between the aeroplane’s climb angle and the horizon. ormed by the leading edge o the section and the relative airflow.

4. 5.

Relative airflow, ree stream flow, relative wind and aircraf flight path are parallel. Aircraf flight path, relative airflow, relative wind and ree stream flow are parallel, but the aircraf flight path is opposite in direction. The pressure, temperature and relative velocity o the ree stream flow are unaffected by the presence o the aircraf. The relative wind is produced by the aircraf moving through the air. The direction o flight is parallel with and opposite to the relative airflow.

a. b. c. d.

5 only. 3, 4 and 5. 1 and 2. 1, 2, 3, 4 and 5.

2. 3.

Which o the ollowing statements are correct? 1. 2. 3. 4. 5.

a. b. c. d. 17.

s n o i t s e u Q

Which o the ollowing statements is true? 1.

16.

4

The term “angle o attack’’ is defined as the angle:

a.

15.

have very little movement. move orward along the chord line. remain unaffected. move back along the chord.

Maximum camber is the maximum distance between the top and bottom surace o an aerooil section. The thickness/chord ratio is expressed as a percentage o the chord. It is easier or air to flow over a well-rounded leading edge radius than a sharp leading edge. Two dimensional airflow assumes a wing with the same aerooil section along its entire span, with no spanwise pressure differential. Air flowing towards the lower pressure o the upper surace is called upwash.

1, 2, 3, 4 and 5. 2, 3 and 4. 2, 3, 4 and 5. 1 and 5.

When considering an aerooil section at a constant angle o attack, which o the ollowing statements is true?

a. b. c. d.

I the static pressure on one side is reduced more than on the other side, a pressure differential will exist. I dynamic pressure is increased, the pressure differential will decrease. The pressure differential will increase i the dynamic pressure is decreased Dynamic pressure and pressure differential are not related.

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4

Questions 18.

Considering an aerooil section subject to a constant dynamic pressure, which o the ollowing statements is correct?

a. 4

b.

Q u e s t i o n s

c. d. 19.

When considering the effect o changing angle o attack on the pitching moment o an aerooil, which o the ollowing statements is correct? 1. 2. 3. 4.

a. b. c. d. 20.

At ‘normal’ angles o attack the pitching moment is nose-up. The pitching moment about the aerodynamic centre (AC) is constant at normal angles o attack. The aerodynamic centre (AC) is located approximately at the 25% chord point. The moment about the aerodynamic centre (AC) is a product o the distance between the aerodynamic centre (AC) and the centre o pressure (CP) and the magnitude o the lif orce.

1, 2, 3 and 4. 4 only. 3 and 4. 2, 3 and 4.

Ice contamination o the leading portion o the aerooil has which o the ollowing consequences? 1. 2. 3. 4.

a. b. c. d.

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I the angle o attack is increased rom 4° to 14°, the pressure differential will not change but lif will be greater due to increased dynamic pressure acting on the lower surace. Up to about 16°, increasing the angle o attack will increase the pressure differential between the top and bottom surace o the aerooil. Changing the angle o attack does not affect the pressure differential, only changes in dynamic pressure affect the pressure differential. Up to about 16°, increasing the angle o attack decreases the pressure differential between the top and bottom surace o the aerooil section.

The profile o the leading portion o the surace can be changed, preventing normal acceleration o the airflow and substantially reducing the magnitude o the lif orce. Form (pressure) drag will be increased because o the increased rontal area o the aerooil section. Loss o lif will have a greater effect than an increase in orm (pressure) drag. At ‘normal’ angles o attack lif can be lost entirely i enough ice accumulates.

1, 2, 3 and 4 1, 3 and 4 1, 2 and 3 3 and 4

4

Questions

4

s n o i t s e u Q

67

4

Answers

Answers

4

A n s w e r s

68

1 c

2 d

3 c

4 a

5 c

6 b

7 c

8 b

13 b

14 b

15 d

16 c

17 a

18 b

19 d

20 a

9 c

10 c

11 c

12 c

Chapter

5 Lift

Aerodynamic Force Coefficient . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .71 The Basic Lif Equation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .72 Review: . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .75 The Lif Curve . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .76 Interpretation o the Lif Curve . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .76 Velocity - Dynamic Pressure Relationship . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .79 Density Altitude . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .79 Aerooil Section Lif Characteristics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .79 Introduction to Drag Characteristics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .80 Lif/Drag Ratio . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .80 Effect o Aircraf Weight on Minimum Flight Speed . . . . . . . . . . . . . . . . . . . . . . . . .

82

Condition o the Surace . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .82 Flight at High Lif Conditions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .82 Three Dimensional Airflow . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .85 Wing Terminology . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .85 Wing Tip Vortices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .86 Wake Turbulence: (Re: AIC P 072/2010) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 88 Ground Effect . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 91 Conclusion . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .96 Summary. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .98 Answers rom page 77 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .99 Answers rom page 78 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 100 Questions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

101

Answers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

108

69

5

Lift

5

L i f t

70

5

Lift Aerodynamic Force Coefficient The aerodynamic orces o both lif and drag depend on the combined effect o many variables. The important actors are: • Airstream velocity (V) • Air density ( ρ)

}

Dynamic Pressure ( ½ ρ V ) 2

• Shape or profile o the surace • Angle o attack

}

5

t f i L

Pressure Distribution (C L or CD)

• Surace area (S) • Condition o the surace • Compressibility effects (to be considered in later chapters)

Dynamic Pressure The dynamic pressure (½ ρ V ) o the airflow is a common denominator o aerodynamic orces and is a major actor since the magnitude o a pressure distribution depends on the energy given to the airflow (KE = ½ m V ). 2

2

Pressure Distribution Another major actor is the relative pressure distribution existing on the surace. The distribution o velocities, with resulting pressure distribution, is determined by the shape or profile o the surace and the angle o attack (C L or CD).

Surface Area Since aerodynamic orces are the result o various pressures distributed on a surace, the surace area (S) is the remaining major actor - the larger the surace area or a given pressure differential, the greater the orce generated. Thus, any aerodynamic orce can be represented as the product o three major actors: • The dynamic pressure o the airflow (½ρ V ) • The coefficient o orce determined by the relative pressure distribution (C L or CD), 2

and • The surace area o the object (S) The relationship o these three actors is expressed by the ollowing equation: F = Q CF S where F Q CF S

= = = =

aerodynamic orce (Lif or Drag) dynamic pressure (½ ρ V ) coefficient o aerodynamic orce (C L or CD) surace area 2

71

5

Lift The Basic Lift Equation Lif is defined as the net orce generated normal (at 90°) to the relative airflow or flight path o the aircraf. The aerodynamic orce o lif results rom the pressure differential between the top and bottom suraces o the wing. This lif orce can be defined by the ollowing equation:

5

L = 1 /2 ρ V2 CL S

L i f t

Correct interpretation o the lif ormula is a key element in the complete understanding o Principles o Flight.

Figure 5.1

Note: For the sake o clarity; during this initial examination o the lif ormula it is stated that C L is determined by angle o attack. This is true, but C L is also influenced by the shape or profile o the surace and other actors which will be amplified in later sections.

• An aircraf spends most o its time in straight and level flight. • How much lif is required? The same as the weight. • Consider that at any moment in time weight is constant, so lif must be constant. • While generating the required lif orce, the less drag the better because drag has to be balanced by thrust, and thrust costs money. • The value o lif divided by drag is a measure o aerodynamic efficiency. This has a maximum value at one particular angle o attack. For a modern wing this is about 4°. I this “optimum” angle o attack is maintained, maximum aerodynamic efficiency will be achieved. Note: Maximum CL and minimum CD are not obtained at best L/D. • Lif is generated by a pressure differential between the top and bottom surace o the wing. Pressure is reduced by the air accelerating over the top surace o the wing. The wing area must be big enough to generate the required lif orce.

72

5

Lift • Air gets thinner as altitude increases. I the speed o the aircraf through the air is kept constant as altitude is increased, the amount o air flowing over the wing in a given time would decrease - and lif would decrease. • For a constant lif orce as altitude is increased, a constant mass flow must be maintained. As air density decreases with altitude, the speed o the wing through the air (the true airspeed (TAS) must be increased.

5

I you reer to the ICAO Standard Atmosphere chart on page 27 , the air density at 40 000 f is only one quarter o the sea level value. We can use this as an example to illustrate the relationship between the changes in TAS that are required as air density changes with altitude.

t f i L

TO KEEP LIFT CONSTA NT AT 40 000 ft , TAS MUST BE DOUBLED

× ×

4

2 KEEP CONSTANT TO MAINTAIN L/D MAX L =

½

V

2

CL

S

FIXED AREA

CONSTANT CONSTANT DYNAMIC PRESSURE (IAS)

1

4 Figure 5.2

For this example we will assume the optimum angle o attack o 4° is maintained or aerodynamic efficiency and that the wing area is constant. At 40 000 f the air density is /4 o the sea level value, so the speed o the aircraf through the air must be doubled to maintain dynamic pressure (hence lif) constant. TAS is squared because essentially we are considering the kinetic energy o the airflow (KE = ½ m V ). 1

2

73

5

Lift The lif ormula can also be used to consider the relationship between speed and angle o attack at a constant altitude (air density).

CL

IF SPEED IS DOUBLED,

MUST BE REDUCED

TO ¼ OF ITS PREVIOUS VALUE 5

L i f t

×

4 1

×

2

L =

½

V

2

CL

S

4

FIXED AREA

CONSTANT DYNAMIC PRESSURE FOUR TIMES GREATER (IAS DOUBLED)

CONSTANT ALT ITUDE

Figure 5.3

As speed is changed, angle o attack must be adjusted to keep lif constant. As an example: i IAS is doubled, TAS will double, and the square unction would increase dynamic pressure (hence lif) by a actor o our. As the aircraf is accelerated, the angle o attack must be decreased so that the C L reduces to one quarter o its previous value to maintain a constant lif orce. It is stated on page 27 that IAS will vary approximately as the square root o the dynamic pressure. The proportionality between IAS and dynamic pressure is: I AS

Q

For the sake o simplicity and to promote a general understanding o this basic principle (though no longer true when considering speeds above M 0.4), it can be said that TAS will change in proportion to IAS, at constant altitude, (double one, double the other, etc). The lif ormula can be transposed to calculate many variables which are o interest to a proessional pilot. For example: i speed is increased in level flight by 30% rom the minimum level flight speed, we can calculate the new C L as a percentage o C LMAX :

74

5

Lift

L = ½ρ V

2

CL S

transposed becomes:

CL =

L ½ρ V S

As density, lif and wing area are constant, this can be written :

2

CL ∝

1 V

2

30% above minimum level flight speed can be written as 1.3V The proportional change in C L is thereore

1 (1.3)

2

=

5

1 = 0.59 = 59% 1.69

t f i L

While maintaining level flight at a speed 30% above minimum level flight speed, the C L would be 59% o C LMAX

Review: Lif must balance weight in straight and level flight, so at any moment in time, weight and the lif required is constant. • To maintain constant lif i density varies because o altitude change, the TAS must be changed. • I altitude is increased, density decreases, so TAS must be increased. • I altitude is decreased, density increases, so TAS must be decreased. Maintaining a constant IAS will compensate or density changes. • To maintain constant lif i speed is changed at a constant altitude (density), the angle o attack must be adjusted. • I speed is increased, angle o attack must be decreased, (i speed is doubled, angle o attack must be decreased to make C L one quarter o its previous value). • I speed is decreased, angle o attack must be increased, (i speed is halved, angle o attack must be increased to make C L our times its previous value). • Generally, a cruise speed is chosen so the aircraf operates at its optimum angle o attack (L/D MAX - approximately 4°).

75

5

Lift The Lift Curve Figure 5.4 shows the lif curve o an aerooil section, with lif coefficient (C L) plotted against

angle o attack. It is evident that the section is symmetrical because no lif is produced at zero angle o attack.

The lif curve is a convenient way to illustrate the properties o various configurations and will be used extensively throughout these notes.

5

Lif coefficient increases with angle o attack up to a maximum (C LMAX), which corresponds to the “Critical” angle o attack. Continuing to increase the angle o attack beyond this point makes it impossible or the airflow to maintain its previous smooth flow over the contour o the upper surace, and lif will reduce. This phenomena, stall, will be discussed in detail later.

L i f t

Interpretation of the Lift Curve • To generate a constant lif orce, any adjustment in dynamic pressure must be accompanied by a change in angle o attack. (At C L less than CLMAX). • For a constant lif orce, each dynamic pressure requires a specific angle o attack. • Minimum dynamic pressure is determined by the maximum lif coefficient (C LMAX), which occurs at a specific angle o attack (approximately 16°). • The angle o attack or CLMAX is constant. (This is true or a given configuration). • I more lif is required due to greater operating weight, a greater dynamic pressure is required to maintain a given angle o attack. • The greater the operating weight, the higher the minimum dynamic pressure. To use the lif ormula with specific values, it is necessary to convert each item to SI units .

The mass o the aircraf is 60 000 kg. To convert to a weight, the mass must be multiplied by the acceleration o gravity (9.81 m/s ). The wing area is 105 m . Density is the ICAO Standard Atmosphere sea level value o 1.225 kg/m . 2

2

3

The speed resulting rom the calculation will be in m/s. There are 6 080 f in one nautical mile and 3.28 f in one metre. The lif ormula:

L = ½ ρ V2 CL S

when transposed to calculate speed becomes:

76

V

=

L ½ ρ CL S

5

Lift

CL Knots

1.532

C LMAX 5

t f i L STALL

0.863

0.552 0.384

ANG LE OF A TTAC K ( DEGREES )

Figure 5.4 Typical lif curve

Please answer the ollowing questions : (Answers are provided on page 99)

a.

How many newtons o lif are required or straight and level flight?

b.

Calculate the airspeed in knots or each highlighted coefficient o lif.

c.

What is the lowest speed at which the aircraf can be flown in level flight?

d.

What coefficient o lif must be used to fly as slowly as possible in level flight?

e.

Does each angle o attack require a particular speed?

.

As speed is increased, what must be done to the angle o attack to maintain level flight?

g.

At higher altitude air density will be lower; what must be done to maintain the required lif orce i the angle o attack is kept constant?

h.

At a constant altitude, i speed is halved, what must be done to the angle o attack to maintain level flight?

77

5

Lift

CAMBERED WITH 12% THICKNESS

CL

CAMBER GIVES INCREASE IN LMAX

C

5

T N E I C I F F E O C

L i f t

SYMMETRICAL

T F I L

WITH 12% THICKNESS GREATER THICKNESS

N O I T C E S

GIVES 70% INCREASE IN C

LMAX

SYMMETRICAL WITH 6% THICKNESS

0

SECTION ANGLE OF ATTACK (DEGREES)

Figure 5.5

Using the above graph, please answer the ollowing questions: (Answers on page 100)

78

a.

Why does the cambered aerooil section have a significantly higher C LMAX?

b.

For the same angle o attack, why do the symmetrical aerooil sections generate less lif than the cambered aerooil section?

c.

Why does the cambered aerooil section o 12% thickness generate a small amount o lif at slightly negative angles o attack?

d.

For a given angle o attack, the symmetrical aerooil section o 6% thickness generates the smallest amount o lif. In what way can this be a avourable characteristic?

e.

What are the disadvantages o the symmetrical aerooil section o 6% thickness?

5

Lift Velocity - Dynamic Pressure Relationship It is very important to understand the relationship between the velocity used in the orce equations and dynamic pressure. The velocity in the orce equation is the speed o the aircraf relative to the air through which it is moving - the True Airspeed (TAS). At a given angle o attack: “For a constant lif orce a constant dynamic pressure must be maintained”. When an aircraf 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 = ½ ρ V ).

5

t f i L

2

Density Altitude Air density at the time o take-off and landing can significantly affect aircraf perormance. I air density is low, a longer take-off run will be needed. Air density is a product o pressure, temperature and humidity. Humidity reduces air density because the density o water vapour is about 5/8 that o 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. I the temperature is above standard, the density o the air will be less, perhaps a lot less, with no direct indication o this act visible to the pilot. I the temperature is 25°C, it would be 10°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 aircraf will need a higher TAS or a given dynamic pressure, and, hence, a longer take-off run to achieve the required IAS. To remember what “high density altitude” means, think o it as “HIGH density ALTITUDE”.

Aerofoil Section Lift Characteristics Figure 5.5 shows aerooil sections with different thickness and camber combinations producing specific CL against α plots.

• An increase in the thickness o a symmetrical aerooil gives a higher C LMAX. • The introduction o camber also has a beneficial effect on C LMAX. The importance o maximum lif coefficient is obvious: The greater the C LMAX , the lower the minimum flight speed (stall speed). However, thickness and camber necessary or a high C LMAX will produce increased orm drag and large twisting moments at high speed. So a high C LMAX is just one o the requirements or an aerooil section. The major point is that a high C LMAX will give a low minimum flight speed (IAS). I an aerooil section o greater camber is used to give a lower minimum flight speed, the efficient cruise speed will be lower due to o the generation o excessive drag. It is better to use an aerooil section that is efficient at high cruise speed, with the ability to temporarily increase the camber o the wing when it is necessary to fly slowly. This can be achieved by the use o adjustable hinged sections o the wing leading and trailing edges ( Flaps).

79

5

Lift Introduction to Drag Characteristics Drag is the aerodynamic orce 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 orces, drag orces may be expressed in the orm o a coefficient which is independent o dynamic pressure and surace area.

5

D = Q CD S

L i f t

Drag is the product o dynamic pressure, drag coefficient and surace area. C D is the ratio o drag per unit wing area to dynamic pressure. I the CD o a representative wing were plotted against angle o attack, the result typically would be a graph similar to that shown in Figure 5.6 . At low angles o attack C D is low and small changes in angle o attack create only small changes in C D. But at higher angles o attack, the rate o change in C D per degree o angle o attack increases; CD change with angle o attack is exponential. Beyond the stalling angle o attack (CLMAX ), a urther large increase in C D takes place.

Lift/Drag Ratio An appreciation o the efficiency o lif production is gained rom studying the ratio between lif and drag, a high L/D ratio being more efficient. The proportions o CL and CD can be calculated or each angle o attack. Figure 5.7 shows that the L/D ratio increases with angle o attack up to a maximum at about 4°; this is called the “optimum” angle o attack. The L/D ratio then decreases with increasing angle o attack until CLMAX is reached. Note: The plot o lif, the plot o drag and the plot o L/D ratio shown in Figure 5.7 are all at different scales and no conclusions should be drawn rom the intersection o plots.

The maximum lif/drag ratio (L/D MAX ) o a given aerooil section will occur at one specific angle o attack. I the aircraf is operated in steady level flight at the optimum angle o attack, drag will be least while generating the required lif orce. Any angle o attack lower or higher than that or L/D MAX reduces the L/D ratio and consequently increases drag or the required lif. Assume the L/D MAX o Figure 5.7 is 12.5. In steady flight at a weight o 588 600 N and IAS to give the required lif at 4° angle o attack, the drag would be 47 088 N (588 600 ÷ 12·5). Any higher or lower speed would require a different angle o attack to generate the required lif orce. Any angle o attack other than 4° will generate more drag than 47 088 N. O course, this same ‘aircraf’ could be operated at a different weight and the same L/D MAX o 12.5 could be obtained at the same angle o attack. But a change in weight requires a change in IAS to support the new weight at the same angle o attack. The lower the weight, the lower IAS required to stay at the L/D MAX angle o attack, and vice versa. For a given configuration (flaps, gear, spoilers and air rame contamination) and at speeds less than M 0.4, changes in weight will not change L/D MAX.

80

5

Lift

5

t f i L

Figure 5.6

L

CD

D C LMAX

L D MAX

STALL

4

16 OPTIMUM ANGLE OF ATTACK

ANGLE OF ATTACK (DEGREES)

Figure 5.7

81

5

Lift The design o an aircraf has a great effect on the L/D ratio. Typical values are listed below or various types.

5

L i f t

Aircraf Type

L/D MAX

High perormance sailplane

rom 25 to 60

Modern jet transport

rom 12 to 20

Propeller powered trainer

rom 10 to 15

Effect of Aircraft Weight on Minimum Flight Speed A given aerooil section will always stall at the same angle o attack, but aircraf weight will influence the IAS at which this occurs. Modern large jet transport aircraf may have just over hal their maximum gross take-off weight made up o uel. So stall speed can vary considerably throughout the flight.

Condition of the Surface Surace irregularities, especially near the leading edge, have a considerable effect on the characteristics o aerooil sections. C LMAX, in particular, is sensitive to leading edge roughness. Figure 5.8 illustrates the effect o a rough leading edge compared to a smooth surace. In general, CLMAX decreases progressively with increasing roughness o the leading edge. Roughness urther downstream than about 20 percent o the chord rom the leading edge has little effect on C LMAX or the lif curve slope. Frost, snow and even rainwater can significantly increase surace roughness. Dirt or slush picked up rom contaminated parking areas, taxiways and runways can also have a serious effect. In-flight icing usually accumulates at the leading edge o aerooils and will severely increase surace roughness causing a significant decrease in CLMAX.

Flight at High Lift Conditions The aerodynamic lif characteristics o an aircraf are shown by the curve o lif coefficient versus angle o attack in Figure 5.9, or a specific aircraf in the clean and flap down configurations. A given aerodynamic configuration experiences increases in lif coefficient with increases in angle o attack until the maximum lif coefficient is obtained. A urther increase in angle o attack produces stall and the lif coefficient then decreases.

Effect of High Lift Devices The primary purpose o high lif devices (flaps, slots, slats, etc) is to reduce take-off and landing distance by increasing the C LMAX o the aerooil section and so reduce the minimum speed. The effect o a “typical” high lif device is shown by the lif curves o Figure 5.9. The principal effect o the extension o flaps is to increase C LMAX and reduce the angle o attack or any given lif coefficient. The increase in C LMAX afforded by flap deflection reduces the stall speed by a certain proportion. (High lif devices will be ully covered later).

82

5

Lift

C LMAX Basic Smooth Wing

CL

Take-off C L

T N E I T C I F I F L F E O C

5

t f i L

Wing with Frost, Dirt, Water or Slush Wing with Ice

ANGLE OF ATTACK

Figure 5.8

CL T N E I C I F F E O C

C LMAX

FLAPS

FLAPS DOWN

C LMAX

T F I L

CLEAN

CLEAN CONFIGURATION

ANGLE OF ATTACK

Figure 5.9

83

5

Lift

b

S = WING AREA, sq. m (b × c)

c

b = SPAN, m c = AVERAGE CHORD, m

5

c

L i f t

b

AR = ASPECT RATIO AR = b c AR = b

2

S

c C R = ROOT CHORD, m

b C T = TIP CHORD, m CT

CR

=

TAPER RATIO

SWEEP ANGLE, degrees

CR

CT

MAC = MEAN AERODYNAMIC CHORD, m

MAC

Figure 5.10 Wing terminology

84

5

Lift Three Dimensional Airflow So ar we have considered only two dimensional airflow. This has been a oundation or an appreciation o the actual pattern o airflow over an aircraf. Even minute pressure differences will modiy airflow direction by inducing air to flow towards any region o lower pressure. Three dimensional airflow modifies the effective angle o attack, increases drag, alters stalling characteristics and can influence the control and stability o the aircraf. From now on, instead o just an aerooil section, the entire wing will be considered.

5

t f i L

Wing Terminology Wing Area (S): The plan surace area o the wing. Although a portion o the area may

be covered by uselage or engine nacelles, the pressure carryover on these suraces allows legitimate consideration o the entire plan area. Wingspan (b): The distance rom tip to tip. Average Chord (c): The mean geometric chord. The product o the span and the average

chord is the wing area (b × c = S). Aspect Ratio (AR) : The proportion o the span and the average chord (AR = b/c). I the planorm

has curvature and the average chord is not easily determined, an alternative expression is b /S. The aspect ratio o the wing determines the aerodynamic characteristics and structural weight. Typical aspect ratios vary rom 35 or a high perormance sailplane to 3 or a jet fighter. The aspect ratio o a modern high speed jet transport is about 12. 2

Root Chord (CR): The chord length at the wing centre line. Tip Chord (CT): The chord length at the wing tip. Taper Ratio (C T / CR): The ratio o the tip chord to the root chord. The taper ratio affects the

lif distribution and the structural weight o the wing. A rectangular wing has a taper ratio o 1.0 while the pointed tip delta wing has a taper ratio o 0.0 Sweep Angle: Usually measured as the angle between the line o 25% chords and a perpendicular

to the root chord. The sweep o a wing causes definite changes in compressibility, maximum lif, and stall characteristics. Mean Aerodynamic Chord (MAC): A rectangular wing o this chord and the same span would

have broadly similar pitching moment characteristics. The MAC is located on the reerence axis o the aircraf and is a primary reerence or longitudinal stability considerations.

85

5

Lift Wing Tip Vortices

5

L i f t

Figure 5.11

Air flowing over the top surace o 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 o flow, inducing a span wise vector towards the root on the upper surace and, generally, towards the tip on the lower surace, Figure 5.11. “Conventionally”, an aircraf is viewed rom the rear. An anticlockwise vortex will be induced at the right wing tip and a clock-wise vortex at the lef wing tip, Figure 5.12, Figure 5.13 & Figure 5.14. At higher angles o attack (lower IAS) the decreased chordwise vector will increase the effect o the resultant spanwise flow, making the vortices stronger.

Figure 5.12

Figure 5.13

Figure 5.14

86

Induced Downwash (Figure 5.15) Trailing vortices create certain vertical velocity components in the airflow in the vicinity o the wing, both in ront o and behind it. These vertical velocities cause a downwash over the wing resulting in a reduction in the effective angle o attack. The stronger the vortices, the greater the reduction in effective angle o attack. Because o this local reduction in effective angle o attack, the overall lif generated by a wing will be below the value that would be generated i there were no spanwise pressure differential. It is the production o lif itsel which reduces the magnitude o the lif orce being generated. To replace the lif lost by the increased downwash, the aircraf must be flown at a higher angle o attack. This increases drag. This extra drag is called induced drag. The stronger the vortices, the greater the induced drag.

5

Lift

Upwash Increased

Vertical Velocities in the vicinity of the wing are a function of tip vortex strength

Downwash Increased 5

t f i L

EFFECTIVE AIRFLOW

Angular deflection of effective airflow is a function of both vortex strength and True Airspeed (TAS).

V

Induced Downwash

Relative Airflow V

Induced Drag ( D )

i

Lift With Normal Downwash

Lift Inclined Rearwards because of Decreased Effective Angle of Attack

i

Effective Airflow e

i Relative Airflow

e = effect ive angle of attack

i

= induced angle of attack

Figure 5.15

Wing tip vortices, in particular their influence on upwash and downwash, have a significant effect on several important areas o aircraf aerodynamics, stability and control. Some o these effects will be examined now and throughout the remaining chapters.

87

5

Lift Wake Turbulence: (Ref: AIC P 072/2010) Trailing wing tip vortices extend behind aircraf or a considerable distance and can present an extreme hazard to any aircraf unortunate enough to encounter them. Maximum tangential airspeed in the vortex system may be as high as 90 m/s (300 f/sec) immediately behind a large aircraf. Wake turbulence cannot be detected, so it is important or pilots to be aware o the potential distribution and duration o trailing vortices, plus modifications made to the “classic” vortex system by surace wind speed and direction.

5

L i f t

Aircraft Wake Vortex Characteristics Wake vortex generation begins when the nose wheel lifs off the runway on take-off and continues until the nose wheel touches down on landing. Wake vortices exist behind every aircraf, including helicopters, when in flight, but are most severe when generated by heavy aircraf. They present the greatest danger during the take-off, initial climb, final approach and landing phases o flight - in other words, at low altitude where large numbers o aircraf congregate. A wake turbulence encounter is a hazard due to potential loss o control and possible structural damage, and i the experience takes place near the ground, there may be insufficient time and/or altitude to recover rom an upset.

Figure 5.16

The characteristics o trailing vortices are determined by the “generating” aircraf’s: • Gross weight - the higher the weight, the stronger the vortices. • Wingspan - has an influence upon the proximity o the two trailing vortices. • Airspeed - the lower the speed, the stronger the vortices. • Configuration - vortex strength is greatest with aircraf in a “clean” configuration (or a given speed and weight). • Attitude - the higher the angle o attack, the stronger the vortices. As a general rule, the larger the “generating” aircraf relative to the aircraf encountering the wake turbulence, the greater the hazard. There is also evidence that or a given weight and speed a helicopter produces a stronger vortex than a fixed-wing aircraf.

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Lift Distribution of Trailing Vortices Typically the two trailing vortices remain separated by about three quarters o the aircraf’s wingspan, and in still air they tend to drif slowly downwards and level off, usually between 500 and 1000 f below the flight path o the aircraf. Behind a large aircraf the trailing vortices can extend as much as nine nautical miles.

5

t f i L

Figure 5.17

Figure 5.18

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Lift Vortex Movement near the Ground Figure 5.19 shows that i the generating aircraf is within 1000 f o the ground, the two vortices

will “touch down” and move outwards at about 5 kt rom the track o the generating aircraf at a height approximately equal to hal the aircraf’s wingspan.

5

L i f t

1000 ft

5 kt Drift

5 kt Drift

STILL AIR - (viewed from the rear)

Figure 5.19

In a crosswind, i the surace wind is light and steady, the wake vortex system “in contact” with the ground will drif with the wind. Figure 5.20 shows the possible effect o a crosswind on the motion o a vortex close to the ground. With parallel runways, wake turbulence rom an aircraf operating on one runway can be a potential hazard to aircraf operating rom the other.

5 kt Wind

Zero Drift (5 kt - 5 kt)

10 kt Drift (5 kt + 5 kt)

5 kt CROSSW IND - (Viewed from the rear)

Figure 5.20

The Decay Process of Trailing Vortices Atmospheric turbulence has the greatest influence on the decay o wake vortices, the stronger the wind, the quicker the decay.

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Lift 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 o an encounter to a lower level, and may minimize the magnitude o the upset i an encounter does occur. Particular care should be exercised when ollowing any substantially heavier aircraf, especially in conditions o light wind. The majority o serious incidents, close to the ground, occur when winds are light.

5

Wake Turbulence Avoidance

t f i L

I the location o wake vortices behind a preceding or crossing aircraf are visualized, appropriate flight path control will minimize the probability o a wake turbulence encounter. Staying above and/or upwind o a preceding or crossing aircraf will usually keep your aircraf out o the generating aircraf’s wake vortex. Unortunately, deviating rom published approach and departure requirements in order to stay above/upwind o the flight path o a preceding aircraf may not be advisable. Maintaining proper separation remains the best advice or avoiding a wake turbulence encounter.

Ground Effect When landing and taking off, the closeness o the wing to the ground prevents ull development o the trailing vortices, Figure 5.21, making them much weaker. Upwash and downwash are reduced, causing the effective angle o attack o the wing to increase, (re: Figure 5.15). Thereore, when an aircraf is “ in ground effect ” lif will generally be increased and induced drag (CDi) will be decreased. In addition, the reduced downwash will affect both longitudinal stability because o CP movement, and the pitching moment because o changes to the effec tive angle o attack o the tailplane, (Re: Figure 5.23).

Figure 5.21

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Lift The Impact of Ground Effect The influence o ground effect depends on the distance o the wing above the ground. A large reduction in C Di will take place only when the wing is very close to the ground, (within hal the wingspan). For a representative aircraf with a 40 m span, (Re. Figure 5.22): • At a height o 40 m, the reduction in CDi is only 1.4%. • At a height o 10 m, the reduction in C Di is 23.5%, but • At a height o 4 m, the reduction in CDi is 47.6%

5

L i f t

60 Percent Reduction in Induced Drag Coefficient

C Di

50

C L Constant

40 30 20 10 0 0

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 (h/b)

Figure 5.22

The height o the wing above the ground when the aircraf is in the landing attitude is influenced by its mounting position on the uselage. From the graph in Figure 5.22 it can be seen that the last ew metres makes a big difference to the reduction o C Di. In general, it can be said that a low wing aircraf will experience a greater degree o ground effect than an aircraf with a high mounted wing.

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Lift High and Low Tail Characteristics While ground effect may possibly change the aerodynamic characteristics o the tailplane in its own right, a low mounted tailplane will have its effective angle o attack modified by the changing downwash angle behind the wing. A high mounted tailplane may be outside the influence o the changing downwash angle and not suffer the same disadvantages.

Normal Downwash

5

t f i L

Down load on Tailplane

Downwash Decreased by Ground Effect

Decreased Down load on Tailplane

Figure 5.23

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Lift

REDUCED DOWNWASH ANGLE

" NORMAL" DOWNWASH

INCREASED UPFORCE

5

L i f t

POSITIVE CAMBER

Tailplane (Every illustration)

NEGATIVE CAMBER

DECREASED DOW NFORCE

SYMMETRICAL DECREASED DOWNFORCE

Figure 5.24

Influence of Tailplane Camber on Pitching Moment It can be seen rom Figure 5.24 that the type o tailplane camber does not influence the pitching moment generated when downwash rom the wing changes. Decreased downwash will always result in an aircraf nose-down pitching moment. The opposite will be true o increased downwash. Downwash will change not only because o ground effect, but also when flaps are operated and when a shock wave orms on the wing at speeds higher than M CRIT, so appreciation o this phenomena is a key element towards a ull understanding o Principles o Flight.

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Lift

5

t f i L

Figure 5.25

Tailplane Angle of Attack Angle o attack is the angle between the chord line and the relative airflow. The relative airflow has three characteristics: • Magnitude - the speed o the aircraf through the air; the True Airspeed (TAS) • Direction - parallel to and in the opposite direction to the aircraf flight path, and • Condition - unaffected by the presence o the aircraf. Air flowing off the wing trailing edge (downwash) cannot be defined as relative air flow because it does not conorm to the definitions. Neither is it possible to think strictly o a tailplane angle o attack. Airflow which has been influenced by the presence o the aircraf (direction o flow and dynamic pressure) must be thought o as Effective Airflow. And the angle between the chord line and the effective airflow must be thought o as Effective Angle o Attack . Consider Figure 5.25. Airflow rom direction A gives the tailplane zero (effective) angle o attack. Airflow rom direction E, F or G would be an increase in (effective) angle o attack. I airflow rom direction G is now considered, flow rom F, E, A, B, C or D would be a decrease in (effective) angle o attack. The term “negative angle o attack” is not used .

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Lift Conclusion Increasing downwash (G to D) gives a decrease in tailplane (effective) angle o attack and decreasing downwash (D to G) gives an increase in tailplane (effec tive) angle o attack. It is necessary to understand the effect o changing downwash on tailplane angle o attack, but it is vital to understand the influence o downwash on aircraf pitching moment.

5

Entering Ground Effect

L i f t

Consider an aircraf entering ground effect, assuming that a constant C L and IAS is maintained. As the aircraf descends into ground effect the ollowing changes will take place: • The decreased downwash will give an increase in the effective angle o attack, requiring a smaller wing angle o attack to produce the same lif coefficient. I a constant pitch attitude is maintained as ground effect is encountered, a “floating” sensation may be experienced due to the increase in C L and the decrease in C Di (thrust requirement), • Figure 5.15 & Figure 5.26 . The decrease o 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 aircraf a tendency to climb above the desired glide path, “balloon”, i a reduced throttle setting is not used. • I airspeed is allowed to decay significantly during short finals and the resulting sink-rate arrested by increasing the angle o attack, upon entering ground effect the wing could stall, resulting in a heavy landing. • 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 o attack o the tailplane, Figure 5.23. The down load on the tail is reduced, producing a nose-down pitching moment. • Due to the changes in the flowfield around the aircraf there will be a change in position error which may cause the ASI to misread. In the majority o cases, local pressure at the static port will increase and cause the ASI and altimeter to under-read.

Figure 5.26

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Lift Leaving Ground Effect The effects o climbing out o ground effect will generally be the opposite to those o entering. Consider an aircraf climbing out o ground effect while maintaining a constant C L and IAS. As the aircraf climbs out o ground effect the ollowing changes will take place: • The CL will reduce and the C Di (thrust requirement) will increase. The aircraf will require an increase in angle o attack to maintain the same C L.

5

• The increase in downwash will generally produce a nose-up pitching moment. The pitch input rom the pilot may need to be reduced (less elevator back-pressure).

t f i L

• Position error changes may cause the ASI to misread. In the majority o cases, local pressure at the static port will decrease and cause the ASI and altimeter to over-read. • It is possible to become airborne in ground effect at an airspeed and angle o attack which would, afer leaving ground effect, cause the aircraf to settle back on to the runway It is thereore vitally important that correct speeds are used or take-off. • The nose-up pitching moment may induce an inadvertent over rotation and tail strike.

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Lift Summary Three major actors influence production o the required lif orce:

• Dynamic Pressure (IAS). • Pressure Distribution (section profile & angle o attack). • Wing Area (S). 5

To provide a constant lif orce, each IAS corresponds to a particular angle o attack. L i f t

The angle o attack at C LMAX is constant. A higher aircraf weight requires an increase in lif orce to balance it; an increased IAS is needed to provide the greater lif at the same angle o attack. As altitude increases, a constant IAS will supply the same lif orce at a given angle o attack. A thinner wing will generate less lif at a given angle o attack, and have a higher minimum speed. A thinner wing can fly aster beore shock wave ormation increases drag. A thinner wing requires high lif devices to have an acceptably low minimum speed. The Lif/Drag ratio is a measure o aerodynamic efficiency. Contamination o the wing surace, particularly the ront 20% o the chord, will seriously decrease aerodynamic perormance. Wing tip vortices:

• • • • • •

Decrease overall lif production. Increase drag. Modiy the downwash which changes the effective angle o attack o the tailplane. Generate trailing vortices which pose a serious hazard to aircraf that encounter them. Affect the stall characteristics o the wing. Change the lif distribution.

The sudden ull effects o vortices or their absence must be anticipated during take-off and landing.

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Lift Answers from page 77 CL 150

Knots

CLMAX

1.532

5

t f i L

200 kt

STALL

0.863

250 kt 0.552

300 kt 0.384

ANG LE OF A TTAC K ( DEGREES )

Figure 5.27

a.

How many newtons o lif are required or straight and level flight? 588 600 N.

b.

Calculate the airspeed in knots or each highlighted coefficient o lif. As above.

c.

What is the lowest speed at which the aircraf can be flown in level flight? 150 kt.

d.

What coefficient o lif must be used to fly as slowly as possible in level flight? CLMAX

e.

Does each angle o attack require a particular speed? Yes.

.

As speed is increased, what must be done to the angle o attack to maintain level flight? Angle o attack must be decreased.

g.

At higher altitude air density will be lower; what must be done to maintain the required lif orce? Increase the True Airspeed (TAS).

h.

At a constant altitude, i speed is halved, what must be done to the angle o attack to maintain level flight? Increased so that C L is our times greater.

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Lift Answers from page 78 CAMBERED WITH 12% THICKNESS

CL

CAMBER GIVES INCREASE IN LMAX

C

5 T N E I C I F F E O C

L i f t


T F I L

GREATER THICKNESS

N O I T C E S

GIVES 70% INCREASE IN C LMAX


0

SECTION ANGLE OF ATTACK (DEGREES)

Figure 5.28

a.

Why does the cambered aerooil section have a significantly higher C LMAX?

When compared to a symmetrical section o the same thickness: at approximately the same stall angle, the cross-sectional area o the streamtube over the top sur ace is smaller with a more gradual section change. This allows greater acceleration o the air over the top surace, and a bigger pressure differential .

b.

For the same angle o attack, why do the symmetrical aerooil sections generate less lif than the cambered aerooil section? Angle o attack is the angle between the chord line and the relative airflow. At the same angle o attack, the cross-sectional area o the symmetrical section upper surace streamtube is larger .

c.

Why does the cambered aerooil section o 12% thickness generate a small amount o lif at slightly negative angles o attack? At small negative angles o attack, a cambered aerooil is still providing a reduced cross-sectional area streamtube over the top surace, generating a small pressure differential.

d.

For a given angle o attack, the symmetrical aerooil section o 6% thickness generates the smallest amount o lif. In what way can this be a avourable characteristic? At the high speeds at which modern high speed jet transport aircraf operate, a thin wing can generate the required lif orce with minimum drag caused by the ormation o shock waves. (This will be ully explained in later chapters).

e.

What are the disadvantages o the symmetrical aerooil section o 6% thickness? It will give a high minimum speed, requiring complex high lif devices to enable the aircraf to use existing runways.

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To maintain altitude, what must be done as Indicated Airspeed (IAS) is reduced?

a. b. c. d.

Decrease angle o attack to reduce the drag. Increase angle o attack to maintain the correct lif orce. Deploy the speed brakes to increase drag. Reduce thrust. 5

2.

I more lif orce is required because o greater operating weight, what must be done to fly at the angle o attack which corresponds to C LMAX?

a. b. c. d. 3.

2. 3. 4.

To generate a constant lif orce, any adjustment in IAS must be accompanied by a change in angle o attack. For a constant lif orce, each IAS requires a specific angle o attack. Minimum IAS is determined by C LMAX. The greater the operating weight, the higher the minimum IAS.

a. b. c. d.

1, 2 and 4. 4 only. 2, 3 and 4. 1, 2, 3 and 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.

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 aircraf leaves ground effect?

a. b. c. d. 6.

Increase the angle o attack. Nothing, the angle o attack or CLMAX is constant. It is impossible to fly at the angle o attack that corresponds to C LMAX. Increase the Indicated Airspeed (IAS).

Which o the ollowing statements is correct? 1.

4.

s n o i t s e u Q

A decrease in parasite drag permitting a lower angle o attack. An increase in induced drag and a requirement or a higher angle o attack. An increase in dynamic stability. A decrease in induced drag requiring a smaller angle o attack.

I the angle o attack and other actors remain constant and airspeed is doubled, lif will be:

a. b. c. d.

two times greater. our times greater. the same. one quarter.

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Questions 7.

What true airspeed and angle o attack should be used to generate the same amount o lif as altitude is increased?

a. b. c. d. 5

8.

Q u e s t i o n s

How can an aeroplane produce the same lif in ground effect as when out o ground effect?

a. b. c. d. 9.

d.

rise rom the surace to traffic pattern altitude. sink below the aircraf generating the turbulence. accumulate and remain or a period o time at the point where the takeoff roll began. dissipate very slowly when the surace wind is strong.

How does the wake turbulence vortex circulate around each wing tip, when viewed rom the rear?

a. b. c. d.

102

using high power settings. operating at high airspeeds. developing lif. operating at high altitude.

Wing tip vortices created by large aircraf tend to:

a. b. c.

13.

Heavy, slow, gear and flaps up. Heavy, ast, 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 aircraf is created only when that aircraf is:

a. b. c. d. 12.

lif and airspeed, but not drag. lif, gross weight, and drag. lif, airspeed, and drag. lif and drag, but not airspeed.

Which flight conditions o a large jet aeroplane create the most severe flight hazard by generating wing tip vortices o the greatest strength?

a. b. c. d. 11.

A lower angle o attack. A higher angle o attack. The same angle o attack. The same angle o attack, but a lower IAS.

By changing the angle o attack o a wing, the pilot can control the aeroplane’s:

a. b. c. d. 10.

A higher true airspeed or any given angle o attack. The same true airspeed and angle o attack. A lower true airspeed and higher angle o attack. A constant angle o attack and true airspeed.

Inward, upward, and around the wing tip. Counterclockwise. Outward, upward, and around the wing tip. Outward, downward and around the wing tip.

5

Questions 14.

Which statement is true concerning the wake turbulence produced by a large transport aircraf?

a. b. c. d. 15.

b. c. d.

remain below the flight path o the jet aeroplane. climb above and stay upwind o the jet aeroplane’s flight path. lif off at a point well past the jet aeroplane’s flight path. remain below and downwind o the jet aeroplane’s flight path.

Light quartering headwind. Light quartering tailwind. Direct tailwind. Strong, direct crosswind.

I you take off behind a heavy jet that has just landed, you should plan to lif off:

a. b. c. d. 19.

The downwind vortex will tend to remain on the runway longer than the upwind vortex. A crosswind will rapidly dissipate the strength o both vortices. A crosswind will move both vortices clear o the runway. The upwind vortex will tend to remain on the runway longer than the downwind vortex.

What wind condition prolongs the hazards o wake turbulence on a landing runway or the longest period o time?

a. b. c. d. 18.

s n o i t s e u Q

To avoid the wing tip vortices o a departing jet aeroplane during take-off, the pilot should:

a. b. c. d. 17.

5

What effect would a light crosswind have on the wing tip vortices generated by a large aeroplane that has just taken off?

a.

16.

Wake turbulence behind a propeller-driven aircraf is negligible because jet engine thrust is a necessary actor in the ormation o vortices. Vortices can be avoided by flying 300 f below and behind the flight path o the generating aircraf. The vortex characteristics o any given aircraf may be altered by extending the flaps or changing the speed. Vortices can be avoided by flying downwind o, and below the flight path o the generating aircraf.

prior to the point where the jet touched down. at the point where the jet touched down and on the upwind edge o the runway. beore the point where the jet touched down and on the downwind edge o the runway. beyond the point where the jet touched down.

The adverse effects o ice, snow or rost on aircraf perormance and flight characteristics include decreased lif and:

a. b. c. d.

increased thrust. a decreased stall speed. an increased stall speed. an aircraf will always stall at the same indicated airspeed.

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Questions 20.

Lif on a wing is most properly defined as the:

a. b. c. d. 5

21.

Q u e s t i o n s

Which statement is true relative to changing angle o attack?

a. b. c. d. 22.

c. d.

camber line. longitudinal axis. chord line. flight path.

Which statement is true, regarding the opposing orces acting on an aeroplane in steady-state level flight?

a. b. c. d.

104

negative air pressure below and a vacuum above the wing’s surace. vacuum below the wing’s surace and greater air pressure above the wing’s surace. higher air pressure below the wing’s surace and lower air pressure above the wing’s surace. higher pressure at the leading edge than at the trailing edge.

On a wing, the orce o lif acts perpendicular to, and the orce o 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. our times greater than at the lower speed. one quarter as much.

An aircraf wing is designed to produce lif resulting rom a difference in the:

a. b.

25.

angle o incidence o the wing. distribution o pressures acting on the wing. amount o airflow above and below the wing. dynamic pressure acting in the airflow.

In theory, i the angle o attack and other actors remain constant and the airspeed is doubled, the lif produced at the higher speed will be:

a. b. c. d. 24.

A decrease in angle o attack will increase pressure below the wing, and decrease drag. An increase in angle o attack will decrease pressure below the wing, and increase drag. An increase in angle o attack will increase drag. An increase in angle o attack will decrease the lif coefficient.

The angle o attack o a wing directly controls the:

a. b. c. d. 23.

differential pressure acting perpendicular to the chord o the wing. orce acting perpendicular to the relative wind. reduced pressure resulting rom a laminar flow over the upper camber o an aerooil, which acts perpendicular to the mean camber. orce acting parallel with the relative wind and in the opposite direction.

Thrust is greater than drag and weight and lif are equal. These orces are equal. Thrust is greater than drag and lif is greater than weight. Thrust is slightly greater than Lif, but the drag and weight are equal.

5

Questions 27.

At higher elevation airports the pilot should know that indicated airspeed:

a. b. c. d. 28.

will be unchanged, but ground speed will be aster. will be higher, but ground speed will be unchanged. should be increased to compensate or the thinner air. should be higher to obtain a higher landing speed.

An aeroplane leaving ground effect will: 5

a. b. c. d. 29.

b. c. d.

There is a corresponding indicated airspeed required or every angle o attack to generate sufficient lif to maintain altitude. An aerooil will always stall at the same indicated airspeed; thereore, an increase in weight will require an increase in speed to generate sufficient lif to maintain altitude. At lower airspeeds the angle o attack must be less to generate sufficient lif to maintain altitude. The lif orce must be exactly equal to the drag orce.

At a given Indicated Airspeed, what effect will an increase in air density have on lif 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.

Which is true regarding the orce o lif in steady, unaccelerated flight?

a.

31.

s n o i t s e u Q

I the same angle o attack is maintained in ground effect as when out o ground effect, lif will:

a. b. c. d. 30.

experience a reduction in ground riction and require a slight power reduction. require a lower angle o attack to maintain the same lif coefficient. experience a reduction in induced drag and require a smaller angle o attack experience an increase in induced drag and require more thrust.

Lif will increase but drag will decrease. Lif and drag will increase. Lif and drag will decrease. Lif and drag will remain the same.

I the angle o attack is increased beyond the critical angle o attack, the wing will no longer produce sufficient lif to support the weight o the aircraf:

a. b. c. d.

unless the airspeed is greater than the normal stall speed. regardless o 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.

105

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Questions 33.

Given that: Aircraf A. Wingspan: 51 m Average wing chord: 4 m Aircraf B. Wingspan: 48 m Average wing chord: 3.5 m

5

Q u e s t i o n s

Determine the correct aspect ratio and wing area:

a. b. c. d. 34.

Aspect ratio o the wing is defined as the ratio o the:

a. b. c. d. 35.

b. c. d.

Increase the angle o attack to compensate or the decreasing dynamic pressure. Maintain a constant angle o attack until the desired airspeed is reached, then increase the angle o attack. Increase angle o attack to produce more lif than weight. Decrease the angle o attack to compensate or the decrease in drag.

Take-off rom an airfield with a low density altitude will result in:

a. b. c. d.

106

wingspan to the wing root. square o the chord to the wingspan. wingspan to the average chord. square o the wing area to the span.

What changes to aircraf control must be made to maintain altitude while the airspeed is being decreased?

a.

36.

aircraf A has an aspect ratio o 13.7, and has a larger wing area than aircraf B. aircraf B has an aspect ratio o 13.7, and has a smaller wing area than aircraf A. aircraf B has an aspect ratio o 12.75, and has a smaller wing area than aircraf A. aircraf A has an aspect ratio o 12.75, and has a smaller wing area than aircraf B.

a longer take-off run. a higher than standard IAS beore lif off. a higher TAS or the same lif off IAS. a shorter take-off run because o the lower TAS required or the same IAS.

5

Questions

5

s n o i t s e u Q

107

5

Answers

Answers

5

A n s w e r s

108

1 b

2 d

3 d

4 a

5 b

6 b

7 a

8 a

9 c

10 a

11 c

12 b

13 c

14 c

15 d

16 b

17 b

18 d

19 c

20 b

21 c

22 b

23 c

24 c

25 d

26 b

27 a

28 d

29 b

30 a

31 d

32 b

33 b

34 c

35 a

36 d

Chapter

6 Drag

Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 111 Parasite Drag . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 112 Induced Drag . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 116 Methods o Reducing Induced Drag . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 122 Effect o Lif on Parasite Drag . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 123 Aeroplane Total Drag . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 124 The Effect o Aircraf Gross Weight on Total Drag . . . . . . . . . . . . . . . . . . . . . . . . . 126 The Effect o Altitude on Total Drag. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 127 The Effect o Configuration on Total Drag . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 127 Speed Stability. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 128 Power Required (Introduction) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 130 Summary. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 132 Questions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

134

Annex A . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 139 Annex B . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 139 Annex C . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 140 Answers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

142

109

6

Drag

TOTAL DRAG

INDUCED DRAG

PARASITE DRAG 6

D r a g

SKIN FRICTION DRAG

FORM DRAG

INTERFERENCE

PROFILE DRAG

Figure 6.1

110

DRAG

6

Drag Introduction Drag is the orce which resists the orward motion o the aircraf. 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 aircraf is moving through the air or the air is flowing past the aircraf: it is the relative velocity which is the important actor.

6

g a r D

Figure 6.2

Every part o an aeroplane exposed to the airflow produces different types o resistance to orward motion which contribute to the Total Drag. Total Drag is sub-divided into two main types: PARASITE DRAG - independent o lif generation, and INDUCED DRAG - the result o lif generation.

Parasite drag is urther sub-divided into: • Skin Friction Drag • Form (Pressure) Drag, and • Intererence Drag NOTE: Skin Friction and Form Drag are together known as PROFILE DRAG.

Induced drag will be considered later. We will first consider the elements o parasite drag.

111

6

Drag Parasite Drag I an aircraf were flying at zero lif angle o attack, the only drag present would be parasite drag. Parasite drag is made-up o ‘Skin Friction’,’Form’ and ‘Intererence’ drag.

Skin Friction Drag Particles o air in direct contact with the surace are accelerated to the speed o the aircraf 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 aircraf because the viscosity o air is low. As distance rom the surace increases, less and less acceleration o the layers o air takes place. Thereore, over the entire surace there will exist a layer o air whose relative velocity ranges rom zero at the surace to a maximum at the boundary o the air affected by the presence o the aircraf. The layer o air extending rom the surace to the point where no viscous effect is detectable is known as the boundary layer. In flight, the nature o the boundary layer will determine the maximum lif coefficient, the stalling characteristics, the value o orm drag, and to some extent the high speed characteristics o an aircraf.

6

D r a g

TRANSITION POINT TURBULENT BOUNDARY LAYER LAMINA R BOUNDARY LAYER

Figure 6.3

Consider the flow o air across a flat surace, as in Figure 6.3. The boundary layer will exist in two orms, either laminar or turbulent. In general, the flow at the ront will be laminar and become turbulent some distance back, known as the transition point. The increased rate o change in velocity at the surace in the turbulent flow will give more skin riction than the laminar flow. A turbulent boundary layer also has a higher level o kinetic energy than a

laminar layer. Forward movement o the transition point will increase skin riction because there will be a greater area o turbulent flow. The position o the transition point is dependent upon: • Surace condition - The thin laminar layer is extremely sensitive to surace irregularities. Any roughness on the skin o a leading portion o an aircraf will cause transition to turbulence at that point and the thickening, turbulent boundary layer will spread out anwise downstream causing a marked increase in skin riction drag.

112

6

Drag • Adverse pressure gradient (Figure 6.4) - A laminar layer cannot exist when pressure is rising in the direction o flow. On a curved surace, such as an aerooil, the transition point is usually at, or near to, the point o maximum thickness. Because o the adverse pressure gradient existing on a curved surace, the transition point will be urther orward than i the surace was flat.

TRANSITION VELOCITY DECREASING PRESSURE INCREASING (in the direction of flow)

6

ADVERSE PRESSURE GRADIENT

g a r D

LAMINAR FLOW TURBULENT FLOW

REVERSE FLOW SEPARATION

Figure 6.4

NOTE : The vertical scale o the boundary layer in the above sketch is greatly exaggerated. Typically, boundary layer thickness is rom 2 millimetres at the leading edge, increasing to about 20 millimetres at the trailing edge.

Form (Pressure) Drag Form (pressure) drag results rom the pressure at the leading edge o a body being greater than the pressure at the trailing edge. Overall, skin riction causes a continual reduction o boundary layer kinetic energy as flow continues back along the surace. The adverse pressure gradient behind the transition point will cause an additional reduction in kinetic energy o the boundary layer. I the boundary layer does not have sufficient kinetic energy in the presence o the adverse pressure gradient, the lower levels o the boundary layer stop moving (stagnate). The upper levels o the boundary layer will overrun at this point (separation point) and the boundary layer will separate rom the surace at the separation point. See Figure 6.4. Also, surace flow af o the separation point will be orward, toward the separation point - a flow reversal. Because o separation, there will be a lower pressure at the trailing edge than the leading edge. An aerodynamic orce will act in the direction o the lower pressure - orm drag. Separation will occur when the boundary layer does not have sufficient kinetic energy in the presence o a given adverse pressure gradient .

113

6

Drag Loss o kinetic energy in the boundary layer can be caused by various actors. • As angle o 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 orward. Eventually, boundary layer separation will occur so close to the leading edge that there will be insufficient wing area to provide the required lif orce, C L will decrease and stall occurs. • When a shock wave orms on the upper surace, the increase o static pressure through the shock wave will create an extreme adverse pressure gradient. I the shock wave is sufficiently strong, separation will occur immediately behind the shock wave. This will be explained ully in Chapter 13 - High Speed Flight .

6

D r a g

Laminar and Turbulent Separation Separation has been shown to be caused by the air flow meeting an adverse pressure gradient, but it is ound 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 preerable to the laminar one, but rom the point o view o drag the laminar flow is preerable.

Streamlining Each part o an aircraf will be subject to orm (pressure) drag. To reduce orm 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 o a body, reducing the curvature o the suraces and thus the adverse pressure gradient. Fineness ratio is the measure o streamlining. It has been ound that the ideal fineness ratio is 3:1, as illustrated in Figure 6.5. NOTE : The addition o airings and fillets (see Glossary, Page 10 ) at the junction o components exposed to the airflow is also reerred to as “Streamlining” .

Depth

Length

Figure 6.5

Profile Drag The combination o skin riction and orm drag is known as profile drag. It can be considered that these drags result rom the “profile” (or cross-sectional area) o the aircraf presented to the relative airflow.

114

6

Drag Interference Drag When considering a complete aircraf, parasite drag will be greater than the sum o the parts. Additional drag results rom boundary layer ‘intererence’ at wing/uselage, wing/engine nacelle and other such junctions. Filleting is necessary to minimize intererence drag.

Factors Affecting Parasite Drag • Indicated Airspeed • Parasite Drag varies directly with the square o the Indicated Airspeed (IAS). 6

• I IAS is doubled, the Parasite Drag will be our times greater - i IAS is halved, the Parasite Drag will be one quarter o its previous value.

g a r D

• Configuration Parasite Drag varies directly in proportion to the rontal area presented to the airflow; this is known as ‘Parasite Area’. I flaps are deployed, the undercarriage lowered, speed brakes selected or roll control spoilers operated, ‘Parasite Area’ is increased and Parasite Drag will increase. • Airrame Contamination Contamination by ice, rost, snow, mud or slush will increase the Parasite Drag Coefficient, and in the case o severe airrame icing, the Parasite Area.

The Parasite Drag Formula DP = ½ ρ V CDp S 2

where, DP

= Parasite Drag

½ ρ V

= Dynamic Pressure (Q)

CDp

= Parasite Drag Coefficient

S

= Area (Parasite Area)

2

115

6

Drag Induced Drag Induced drag is an undesirable by-product o lif. Wing tip vortices modiy upwash and downwash in the vicinity o the wing which produces a rearward component to the lif vector known as induced drag. The lower the IAS, the higher the angle o attack - the stronger the vortices. The stronger the vortices - the greater the induced drag .

6

Wing Tip Vortices Airflow over the top surace o a wing is at a lower pressure than that beneath. The trailing edge and the wing tips are where the airflows interact, Figure 6.6 . The pressure differential modifies the directions o flow, inducing a spanwise vector towards the root on the upper surace and towards the tip on the lower surace. “Conventionally”, an aircraf is viewed rom the rear. An anti-clockwise vortex will be induced at the right wing tip and a clock-wise vortex at the lef wing tip, Figure 6.7 . At higher angles o attack (lower IAS) the decreased chordwise vector will increase the resultant spanwise flow, making the vortices stronger.

D r a g

UPPER SURFACE (Lower Pressure)

Figure 6.6

Figure 6.7

Induced Downwash Wing tip vortices create certain vertical velocity components in the airflow in the vicinity o the wing, both in ront o and behind it, Figure 6.9. These vertical velocities strengthen upwash and downwash which reduces the effective angle o attack. The stronger the vortices, the greater the reduction in effective angle o attack.

Due to the localized reduction in effective angle o attack, the overall lif generated by a wing will be below the value that would be generated i there were no spanwise pressure differential. It is the production o lif itsel which reduces the magnitude o the lif orce being generated. To replace the lif lost by the increased upwash and downwash, the wing must be flown at a higher angle o attack than would otherwise be necessary. This increases drag. This extra drag is called Induced drag, Figure 6.10.

116

6

Drag

EFFECT IVE AIRFLOW RELATIVE AIRFLOW

Tip vortices increase upwash over outer portions of span

6

g a r D

Tip vortices increase downwash over outer portions of span

INCREASED DOW NWA SH AND UPWASH REDUCES EFFECTIVE ANGLE OF ATTACK OVER OUTER PORTIONS OF SPAN

Figure 6.8

Upwash Increased

Vertical Velocities in the vicinity of the wing are a function of tip vortex strength

Downwash Increased

EFFECTIVE AIRFLOW

Angular deflection of effective airflow is a function of both vortex strength and True Airspeed (TAS). V

Induced

Relative Airflow

Downwash V

Figure 6.9

117

6

Drag

e = effective angle of attack

Induced Drag ( D )

i

i = induced angle of attack

Lift With Normal Downwash

Lift Inclined Rearwards because of Decreased Effective Angle of Attack

6

D r a g

i

Effective Airflow

e

i Relative Airflow

Figure 6.10

Factors that Affect Induced Drag: The size o the lif orce - Because induced drag is a component o the lif orce, the greater the lif, the greater will be the induced drag. Lif must be equal to weight in level flight so induced drag will depend on the weight o the aircraf. Induced drag will be greater at higher aircraf weights. Certain manoeuvres require the lif orce to be greater than the aircraf weight. The relationship o lif to weight is known as the ‘Load Factor’ (or ‘g’). For example, lif is greater than weight during a steady turn so induced drag will be higher during a steady turn than in straight and level flight. Thereore, induced drag also increases as the Load Factor increases. Induced drag will increase in proportion to the square o the lif orce. The speed o the aircraf - Induced drag decreases with increasing speed (or a constant lif

orce). This is because, as speed increases, the downwash caused by the tip vortices becomes less significant, the rearward inclination o the lif is less, and thereore induced drag is less. Induced drag varies inversely as the square o the speed . (Reer to page 121or a detailed explanation). The aspect ratio o the wing - The tip vortices o a high aspect ratio wing affect a smaller

proportion o the span so the overall change in downwash will be less, giving a smaller rearward tilt to the lif orce. Induced drag thereore decreases as aspect ratio increases (or a given lif orce). The induced drag coefficient is inversely proportional to the aspect ratio.

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6

Drag From the previous three actors it is possible to develop the ollowing equation: C CDi = L AR It can be seen that the relationship or the induced drag coefficient, (C Di), emphasizes the need o a high aspect ratio wing or aeroplane configurations designed to operate at the higher lif coefficients during the major portion o their flight, i.e. conventional high speed jet transport aircraf. 2

The effect o aspect ratio on lif and drag characteristics is shown in Figure 6.11 and Figure 6.12. The basic aerooil section properties are shown on these plots, and these properties would be typical only o a wing planorm o extremely high (infinite) aspect ratio. When a wing o some finite aspect ratio is constructed o this basic section, the principal differences will be in the lif and drag characteristics - the moment characteristics remain essentially the same.

6

g a r D

The effect o increasing aspect ratio on the lif curve, Figure 6.11, is to decrease the wing angle o attack necessary to produce a given lif coefficient. Higher aspect ratio wings are more sensitive to changes in angle o attack, but require a smaller angle o attack or maximum lif.

AR = 12

1.4 WING

CL

1.2

AR = 18

AR = 5

AR = 2

BASIC SECTION INFINITE AR

1.0

0.8

0.6

( NO SWEEPBACK )

0.4

0.2

0 5

10

15

20

25

WING ANGLE OF ATTACK

Figure 6.11

119

6

Drag From Figure 6.12 it can be seen that at any lif 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 lif. At high lif coefficients (low IAS), the induced drag is very high and increases very rapidly with lif coefficient.

BASIC SECTION INFINITE AR

1.4

6

AR = 18

WING D r a g

CL

AR = 12

AR = 5

AR = 2

1.2

1.0

0.8

0.6

( LOW MACH NUMBER )

0.4

0.2

0 0.05

0.10

0.15

W ING DRAG COEFFICIENT

0.20

0.25

CD

Figure 6.12

The lif and drag curves or a high aspect ratio wing, Figure 6.11 and Figure 6.12, show continued strong increase in C L with α up to stall and large changes in C D only at the point o stall. Continuing to increase aspect ratio is restricted by the ollowing considerations. Very high aspect ratio wings will experience the ollowing: • Excessive wing bending moments : which can be reduced by carrying uel in the wings and mounting the engines in pods beneath the wing. • Reduced rate o 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 o attack. The increased effective angle o attack is due to the resultant o the orward TAS o the wing and the angular TAS o the tip. The higher the aspect ratio, the greater the vertical TAS o the tip or a given roll rate, leading to a greater increase in effective angle o attack. The higher the effective angle o 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.

120

6

Drag The Induced Drag Coefficient (CDi ) Di = ½ ρ V

CDi

2

S

This equation would seem to imply that induced drag (D i) increases with speed, but the induced drag coefficient (CDi) is proportional to CL2 and inversely proportional to wing aspect ratio. As speed increases, to maintain a constant lif orce CL must be reduced. Thus, with an increase in speed, CDi decreases: C CDi = L AR 2

6

g a r D

The ollowing example illustrates the change in C Di with speed, which leads to the change in D i. I an aircraf’s speed is increased rom 80 kt (41 m/s) to 160 kt (82 m/s), the dynamic pressure will be our times greater. (Sea level ISA density is used in the example, but any constant density will give the same result). Q = ½ ρ V

2

Q = 0.5 × 1.225 × 41 × 41 = 1029.6 Q = 0.5 × 1.225 × 82 × 82 = 4118.4 Reerring to the lif ormula: L = Q CL S I dynamic pressure is our times greater because speed is doubled, C L must be reduced to a quarter o its previous value to maintain a constant lif orce. Applying 1/4 o the previous C L to the C Di ormula: C CDi = L AR 2

CDi =

(¼) AR

2

because AR is constant

C Di

=

(¼)

2

=

/16

1

I /16 o the previous CDi is applied to the induced drag ormula: 1

Di = (Q × 4) ×

/16

1

=

¼

Conclusion: I speed is doubled in level flight: dynamic pressure will be our times greater, C L

must be decreased to ¼ o its previous value, C Di will be /16 o its previous value and Di will be reduced to ¼ o its previous value. 1

I speed is halved in level flight: dynamic pressure will be ¼ o its previous value, C L will need to be our times greater, C Di will be 16 times greater, giving our times more D i

121

6

Drag Methods of Reducing Induced Drag Induced drag is low at high speeds, but at low speeds it comprises over hal the total drag. Induced drag depends on the strength o the trailing vortices, and it has been shown that a high aspect ratio wing reduces the strength o the vortices or a given lif orce. However, very high aspect ratios increase the wing root bending moment, reduce the rate o roll and give reduced ground clearance in roll during take-off and landing; thereore, aspect ratio has to be kept within practical limits. The ollowing list itemizes other methods used to minimize induced drag by weakening the wing tip vortices.

6

• 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 itsel will cause parasite drag, and at higher speeds there may be no overall saving in drag.

D r a g

• 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. • Winglets: These are small vertical aerooils which orm part o the wing tip ( Figure 6.13). Shaped and angled to the induced airflow, they generate a small orward orce (i.e. “negative drag”, or thrust). Winglets partly block the air flowing rom the bottom to the top surace o the wing, reducing the strength o the tip vortex. In addition, the small vortex generated by the winglet interacts with and urther reduces the strength o the main wing tip vortex. • Wing tip shape: The shape o the wing tip can affect the strength o the tip vortices, and designs such as turned down or turned up wing tips have been used to reduce induced drag.

W inglet

Figure 6.13

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6

Drag Effect of Lift on Parasite Drag The sum o drag due to orm, riction and intererence is termed “parasite” drag because it is not directly associated with the development o lif. While parasite drag is not directly associated with the production o lif, in reality it does vary with lif. The variation o parasite drag coefficient, C Dp , with lif coefficient, C L , is shown or a typical aeroplane in Figure 6.14.

CDp

1.4

CL

6

1.4

CD p min

1.2

1.2

CL

1.0 CD i

0.8

2

=

CL

AR

1.0 0.8

0.6

0.6

0.4

0.4

CDp min

0.2

g a r D

+ CD C D = C Dp i min

0.2 0

0 0

0.05

0.10

0.15

0

0.05

0.15

CD

CD Figure 6.14

0.10

Figure 6.15

However, the part o parasite drag above the minimum at zero lif is included with the induced drag coefficient . Figure 6.15.

Effect o Configuration Parasite drag, Dp , is unaffected by lif, but is variable with dynamic pressure and area. I all other actors are held constant, parasite drag varies significantly with rontal area. As an example, lowering the landing gear and flaps might increase the parasite area by as much as 80%. At any given IAS this aeroplane would experience an 80% increase in parasite drag. Effect o Altitude In most phases o flight the aircraf will be flown at a constant IAS, the dynamic pressure and, thus, parasite drag will not vary. The TAS would be higher at altitude to provide the same IAS. Effect o Speed The effect o speed alone on parasite drag is the most important. I all other actors are held constant, doubling the speed will give our times the dynamic pressure and, hence, our times the parasite drag, (or one quarter as much parasite drag at hal the original speed). This variation o parasite drag with speed points out that parasite drag will be o greatest importance at high IAS and o much lower significance at low dynamic pressures. To illustrate this act, an aeroplane in flight just above the stall speed could have a parasite drag which is only 25% o the total drag. However, this same aeroplane at maximum level flight speed would have a parasite drag which is very nearly 100% o the total drag. The predominance o parasite drag at high flight speeds emphasizes the necessity or great aerodynamic cleanliness (streamlining) to obtain high speed perormance.

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6

Drag Aeroplane Total Drag The total drag o an aeroplane in flight is the sum o induced drag and parasite drag. Figure 6.16 illustrates the variation o total drag with IAS or a given aeroplane in level flight at a particular weight and configuration.

6

TOTAL DRAG

DRAG

D r a g

L

D MAX

Parasite Drag

Induced Drag

V MD

IAS

Figure 6.16

Figure 6.16 shows the predominance o induced drag at low speed and parasite drag at high

speed. Because o 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 drags are equal . The speed or minimum drag is an important reerence or many items o

aeroplane perormance. Range, endurance, climb, glide, manoeuvre, landing and take-off perormance are all based on some relationship involving the aeroplane total drag curve. Since flying at VMD incurs the least total drag or lif-equal-weight flight, the aeroplane will also be at L/DMAX angle o attack (approximately 4°). It is important to remember that L/D MAX is obtained at a specific angle o attack and also that the maximum Lif/Drag ratio is a measure o aerodynamic efficiency. NOTE : I an aircraf is operated at the L/DMAX angle o attack, drag will be a minimum while generating the required lif orce. Any angle o attack lower or higher than that or L/DMAX increases the drag or a given lif orce; greater drag requires more thrust, which would be inefficient, and expensive. It must also be noted that i IAS is varied, L/D will vary.

124

6

Drag Figure 5.7 illustrated L/D ratio plotted against angle o attack. An alternative presentation o L/D is a polar diagram in which C L is plotted against C D, as illustrated in Figure 6.17 .

CL

6

g a r D

L D MAX

CD Figure 6.17

The CL / CD , whole aeroplane polar diagram in Figure 6.17 shows CL increasing initially much more rapidly than CD but that ultimately CD increases more rapidly. The condition or maximum Lif/Drag ratio may be ound rom the drag polar by drawing the tangent to the curve rom the origin. NOTE : This is a very common method o displaying L/D ratio, so the display in Figure 6.17 should become well known.

125

6

Drag The Effect of Aircraft Gross Weight on Total Drag The effect o variation in aircraf gross weight on total drag can be seen rom Figure 6.18. As uel is consumed, gross weight will decrease. As the aircraf weight decreases, less lif is required (lower CL) which will reduce induced drag. Total drag will be less and VMD will occur at a lower IAS. I an aircraf is operated at a higher gross weight, more lif will be required. I more lif is generated, induced drag will be higher, total drag will be greater and V MD will occur at a higher IAS. I an aircraf is manoeuvred so that the load actor is increased, the result will be similar to that caused by an increase in gross weight, i.e. induced drag will increase.

6

D r a g

DRAG Decreased TOTAL DRAG at lower weight

Parasite Drag

Less Induced Drag at lower weight

V

Decreased MD because of lower weight

I AS

Figure 6.18

126

6

Drag The Effect of Altitude on Total Drag Aircraf usually operate within limits o Indicated Airspeed (IAS), so it is relevant to consider the variation o drag with IAS. I an aircraf is flown at a constant IAS, dynamic pressure will be constant. As density decreases with increasing altitude, TAS must be increased to maintain the constant IAS (Q = ½ ρ V ). I the aircraf is flown at a constant IAS, drag will not vary with altitude. 2

The Effect of Configuration on Total Drag

6

Extension o the landing gear, air brakes, or flaps will increase parasite drag but will not substantially affect induced drag. The effect o increasing parasite drag is to increase total drag at any IAS but to decrease the speed V MD compared to the clean aircraf, ( Figure 6.19).

g a r D

Figure 6.19

127

6

Drag Speed Stability For an aircraf to be in steady flight, the aircraf must be in equilibrium - there can be no out o balance orces or moments. When an aircraf is trimmed to fly at a steady speed, thrust and drag are equal. Thereore, when an aircraf 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 IAS as in the graph o Figure 6.16 , the term ‘THRUST REQUIRED’ can be substituted or drag.

6

For an aircraf in steady flight, i there 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 o thrust available. This phenomena is illustrated in Figure 6.20.

D r a g

DRAG Thrust Excess

or

Thrust Required

Thrust Av ailable

Thrust Deficiency Thrust Excess

Thrust Deficiency

A

B

Non Stable IAS

Stable IAS region

Region

V MD

Neutral IAS Region

Figure 6.20

128

IAS

6

Drag I an aircraf is established in steady flight at point ‘A’ in Figure 6.20, lif is equal to weight and the thrust available is set to match the thrust required. I the aircraf is disturbed to some airspeed slightly greater than point ‘A’, a thrust deficiency will exist and, i the aircraf is disturbed to some airspeed slightly lower than point ‘A’, a thrust excess will exist. This relationship provides a tendency or the aircraf to return to the equilibrium o point ‘A’ and resume the original trim speed. Steady flight at speeds greater than V MD is characterized by a relatively strong tendency o the aircraf to maintain the trim speed quite naturally; the aircraf is speed stable . Speed stability is an important consideration, particularly at speeds at and below V MD, most ofen encountered during the approach to landing phase o flight.

6

g a r D

I an aircraf is established in steady flight at point ‘B’ in Figure 6.20, lif is equal to weight and the thrust available is set to match the thrust required. I the aircraf is disturbed and goes aster than the trim speed, there will be a decrease in drag giving an excess o thrust which will cause the aircraf to accelerate. I a disturbance slows the aircraf below the trim speed, there will be an increase in drag which will give a thrust deficiency causing the aircraf to slow urther. This relationship is basically unstable because the variation o excess thrust to either side o point ‘B’ tends to magniy any original disturbance. Steady flight at speeds less than VMD is characterized by a tendency or the aircraf to drif away rom the trim speed and the aircraf is speed unstable. I a disturbance reduces speed, it will naturally continue to reduce. I a disturbance increases speed, it will continue to accelerate until the thrust and drag are once more balanced. For this reason, the pilot must closely monitor IAS during the approach phase o flight. Any tendency or the aircraf to slow down must be countered immediately by a ‘generous’ application o thrust to quickly return to the desired trim speed. Consider Figure 6.19. I an aircraf maintains a constant IAS in the speed unstable region, the addition o parasite drag by selecting undercarriage down or by deploying flaps has the benefit o reducing V MD which can improve speed stability by moving the speed stable region to the lef. At speeds very close to V MD an aircraf usually exhibits no tendency towards either speed stability or speed instability - the neutral IAS region.

129

6

Drag 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 reerred to as ‘Thrust Required’ and you will now see that a similar relationship exists with ‘Power Required’ - they are both impor tant airrame considerations. • Thrust is a FORCE (a push or a pull), used to oppose Drag,

6

but Power is the RATE o doing WORK, or

D r a g

and

POWER =

WORK TIME

WORK = FORCE × DISTANCE

so POWER must be

FORCE × DISTANCE TIME

For Power Required: Which Force? Drag. Distance divided by time is speed. Which speed? The only speed there is - the speed o the aircraf through the air, True Airspeed (TAS). Thereore: POWER REQUIRED = DRAG × TAS • I an aircraf climbs at a constant IAS, Drag will remain constant, but TAS must be increased - so power required will increase . It is necessary to consider power required when studying Principles o Flight because Work must be done on the aircraf to “raise” it to a higher altitude when climbing. Logically, maximum work can be done on the aircraf in the minimum time when the power available rom the engine(s) is greatest and the power required by the airrame is least. For easy reerence, associate the word POWER with the word RATE. e.g. minimum rate o descent is achieved in a steady glide when the aircraf is flown at the minimum p ower required speed (VMP ). These and other considerations will be examined more ully during the study o Aircraf Perormance in Book 6 and Flight Mechanics in Chapter 12 o this Book.

130

6

Drag

POWER REQUIRED (kW)

6

(DRAG × TAS)

g a r D

V MP

THRUST REQUIRED L / D MAX

or DRAG (kN)

TAS (kt)

V MD

Figure 6.21

Figure 6.21 is drawn or sea level conditions where TAS = IAS and is valid or one particular

aircraf, or one weight, only in level flight, and shows how a graph o TAS against ‘Power Required’ has been constructed rom a TAS/Drag curve by multiplying each value o drag by the appropriate TAS and converting it to kilowatts. The speed or minimum power required is known as V MP and is an Indicated Airspeed (IAS). Note that the speed corresponding to minimum power required (V MP), is slower than the speed or minimum drag (V MD).

Effect o Altitude An aircraf flying at V MD will experience constant drag at any altitude because V MD is an IAS. At altitude the TAS or a given IAS is higher, but the power required also increases because Power Required = Drag × TAS. So the ratio o TAS to Power Required is unaffected and V MP will remain slower than V MD.

This inormation primarily concerns aircraf perormance, but the relationship o speed or minimum power required (VMP) and speed or minimum drag (V MD) is important or the study o rate and angle o descent in a steady glide, outlined in Chapter 12.

131

6

Drag Summary Parasite Drag is made up o:

Skin riction drag. Form (Pressure) drag. (Skin riction drag plus Form drag is known as Profile drag.) 6

Intererence drag.

D r a g

Parasite Drag varies directly as the square o the Indicated Airspeed (IAS) - Double the speed,

our times the parasite drag. Halve the speed, one quarter the parasite drag. The designer can minimize parasite drag by:

Streamlining.

Filleting. The use o laminar flow wing sections.

Flight crews must ensure the airrame, and the wing in particular, is not contaminated by ice, snow, mud or slush. Induced Drag

Spanwise airflow generates wing tip vortices. The higher the CL (the lower the IAS), the stronger the wing tip vortices. Wing tip vortices strengthen downwash. Strengthened downwash inclines wing lif rearwards. The greater the rearward inclination o wing lif, the greater the induced drag. Induced Drag varies inversely as the square o the Indicated Airspeed (IAS) - Halve the speed,

16 times the induced drag coefficient (C Di) and our times the induced drag (D i). Double the speed, one sixteenth the C Di and one quarter the D i. The designer can minimize induced drag by: Using a high aspect ratio wing planorm. Using a tapered wing planorm with wing twist and/or spanwise camber variation, or incorporation o wing end plates, tip tanks, winglets or various wing tip shapes.

132

6

Drag Total Drag

Total drag is the sum o Parasite drag and Induced drag. Total drag is a minimum when Parasite drag and Induced drag are equal. At low IAS Induced drag is dominant. At high IAS Parasite drag dominates. The IAS at which Parasite and Induced drags are equal is called minimum drag speed (V MD).

6

g a r D

As gross weight decreases in flight, Induced drag decreases, Total drag decreases and V MD decreases. At a constant IAS, altitude has no effect on Total drag, but TAS will increase as density decreases with increasing altitude. Configuration changes which increase the “Parasite Area”, such as undercarriage, flaps or speed brakes, increase Parasite drag, increase Total drag and decrease V MD.

Speed Stability

An aircraf flying at a steady IAS higher than V MD with a fixed throttle setting will have speed stability. An aircraf flying at a steady IAS at V MD or slower with a fixed throttle setting will usually NOT have speed stability. I an aircraf flying at a steady IAS and a fixed throttle setting within the non-stable IAS region encounters a disturbance which slows the aircraf, the aircraf will tend to slow urther; IAS will tend to continue to decrease and Total drag increase.

Power Required

VMP the Indicated Airspeed or minimum ‘Power Required’ is slower than the minimum drag speed (VMD). Maximum TAS/Power ratio (1.32VMP) occurs at V MD .

133

6


What is the effect on total drag o an aircraf i the airspeed decreases in level flight below that speed or maximum L/D?

a. b. c. d. 6

2.

Q u e s t i o n s

By changing the angle o attack o a wing, the pilot can control the aeroplane’s:

a. b. c. d. 3.

decreases because o lower parasite drag. increases because o increased parasite drag. increases because o increased induced drag. decreases because o lower induced drag.

(Reer to annex ‘A’) At the airspeed represented by point B, in steady flight, the aeroplane will:

a. b. c. d.

134

twice as great. hal as great. our times greater. one quarter as much.

As airspeed decreases in level flight below that speed or maximum lif/drag ratio, total drag o an aeroplane:

a. b. c. d. 6.

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, i the airspeed o an aeroplane is doubled while in level flight, parasite drag will become:

a. b. c. d. 5.

lif and airspeed, but not drag. lif, gross weight, and drag. lif, airspeed, and drag. lif 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 o increased induced drag. Drag decreases because o lower induced drag. Drag increases because o increased parasite drag. Drag decreases because o lower parasite drag.

have its maximum L/D ratio. have its minimum L/D ratio. be developing its maximum coefficient o lif. be developing its minimum coefficient o drag.

6

Questions 7.

Which statement is true relative to changing angle o attack?

a. b. c. d. 8.

On a wing, the orce o lif acts perpendicular to, and the orce o drag acts parallel to the:

a. b. c. d. 9.

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 o the airspeed.

a minimum. less than induced drag. greater than induced drag. equal to induced drag.

30 : 1 15 : 1 25 : 1 7.5 : 1

Which is true regarding aerodynamic drag?

a. b. c. d. 13.

flight path. longitudinal axis. chord line. longitudinal datum.

An aircraf has a L/D ratio o 15:1 at 50 kt in calm air. What would the L/D ratio be with a direct headwind o 25 kt?

a. b. c. d. 12.

s n o i t s e u Q

The best L/D ratio o an aircraf occurs when parasite drag is:

a. b. c. d. 11.

6

That portion o the aircraf’s total drag created by the production o lif is called:

a. b. c. d. 10.

A decrease in angle o attack will increase pressure below the wing, and decrease drag. An increase in angle o attack will decrease pressure below the wing, and increase drag. An increase in angle o attack will increase drag. A decrease in angle o attack will decrease pressure below the wing and increase drag.

Induced drag is a by-product o lif and is greatly affected by changes in airspeed. All aerodynamic drag is created entirely by the production o lif. Induced drag is created entirely by air resistance. Parasite drag is a by-product o lif.

At a given True Airspeed, what effect will increased air density have on the lif and drag o an aircraf?

a. b. c. d.

Lif will increase but drag will decrease. Lif and drag will increase. Lif and drag will decrease. Lif and drag will remain the same.

135

6

Questions 14.

I the Indicated Airspeed o an aircraf is increased rom 50 kt to 100 kt, parasite drag will be:

a. b. c. d. 15. 6

I the Indicated Airspeed o an aircraf is decreased rom 100 kt to 50 kt, induced drag will be:

a. b. c. d.

Q u e s t i o n s

16.

induced drag. orm drag. parasite drag. intererence drag.

Which relationship is correct when comparing drag and airspeed?

a. b. c. d.

136

decrease in induced drag only. increase in induced drag. increase in parasite drag. decrease in parasite drag only.

The resistance, or skin riction, due to the viscosity o the air as it passes along the surace o a wing is a type o:

a. b. c. d. 20.

manoeuvrability. controllability. stability. instability.

As Indicated Airspeed increases in level flight, the total drag o an aircraf becomes greater than the total drag produced at the maximum lif/drag speed because o the:

a. b. c. d. 19.

varies with Indicated Airspeed. varies depending upon the weight being carried. varies with air density. remains constant regardless o Indicated Airspeed changes.

The tendency o an aircraf to develop orces which restore it to its original condition, when disturbed rom a condition o steady flight, is known as:

a. b. c. d. 18.

two times greater. our times greater. hal as much. one quarter as much.

The best L/D ratio o an aircraf in a given configuration is a value that:

a. b. c. d. 17.

our times greater. six times greater. two times greater. one quarter as much.

Parasite drag varies inversely as the square o the airspeed. Induced drag increases as the square o the airspeed. Parasite drag increases as the square o the lif coefficient divided by the aspect ratio. Induced drag varies inversely as the square o the airspeed.

6

Questions 21.

I the same angle o attack is maintained in ground effect as when out o ground effect, lif will:

a. b. c. d. 22.

Which statement is true regarding aeroplane flight at L/Dmax?

a. b. c. d. 23.

square o the chord to the wingspan. wingspan to the wing root. area squared to the chord. wingspan to the mean chord.

poor control qualities at low airspeeds. increased drag at high angles o attack. a lower stall speed. reduced bending moment on its attachment points.

increased drag, especially at a low angle o attack. decreased drag, especially at a high angle o attack. increased drag, especially at a high angle o attack. decreased drag, especially at low angles o attack.

(Reer to annex ‘B’) Which aircraf has the highest aspect ratio?

a. b. c. d. 27.

s n o i t s e u Q

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.

6

A wing with a very high aspect ratio (in comparison with a low aspect ratio wing) will have:

a. b. c. d. 25.

Any angle o attack other than that or L/Dmax increases parasite drag. Any angle o attack other than that or L/Dmax increases the lif/drag ratio. Any angle o attack other than that or L/Dmax increases total drag or a given aeroplane’s lif. Any angle o attack other than that or L/Dmax increases the lif and reduces the drag.

Aspect ratio o a wing is defined as the ratio o the:

a. b. c. d. 24.

decrease, and parasite drag will decrease. increase, and induced drag will decrease. decrease, and parasite drag will increase. increase and induced drag will increase.

3. 4. 2. 1.

(Reer to annex ‘B’) Which aircraf has the lowest aspect ratio?

a. b. c. d.

4. 2. 3. 1.

137

6

Questions 28.

(Reer to annex ‘B’) Consider only aspect ratio (other actors remain constant). Which aircraf will generate greatest lif?

a. b. c. d. 29. 6

(Reer to annex ‘B’) Consider only aspect ratio (other actors remain constant). Which aircraf will generate greatest drag?

a. b. c. d.

Q u e s t i o n s

30.

32.

Increases. Increases then decreases. Decreases. Decreases then increases.

(Reer to annex ‘C’), the whole aircraf C L against C D polar. Point ‘B’ represents: 1. 2. 3. 4.

Best Lif/Drag ratio. The critical angle o attack. Recommended approach speed. Never exceed speed (V NE ).

a. b. c. d.

1 and 2. 1 only. 2 and 3. 4 only.

I the Indicated Airspeed o an aircraf in level flight is increased rom 100 kt to 200 kt, by what actor will (i) TAS (ii) C Di (iii) D i change?

a. b. c. d.

138

1. 4. 3. 2.

What happens to total drag when accelerating rom C LMAX to maximum speed?

a. b. c. d. 31.

1. 2. 3. 4.

(i)

(ii)

(iii)

2 0 4 2

1/4 4 1/16 1/16

1/16 16 1/4 1/4

6

Questions Annex A

6

s n o i t s e u Q

Annex B Aircraf 1.

Span 22.5 metres Chord 4 metres

Aircraf 2.

Wing Area 90 square metres Span 45 metres

Aircraf 3.

Span 30 metres Chord 3 metres

Aircraf 4.

Wing Area 90 square metres Span 40 metres

139

6

Questions Annex C

6

Q u e s t i o n s

140

6

Questions

6

s n o i t s e u Q

141

6

Answers

Answers 1 a

2 c

3 b

4 c

5 c

6 a

7 c

8 a

9 a

10 d

11 b

12 a

13 b

14 a

15 b

16 d

17 c

18 c

19 c

20 d

21 b

22 c

23 d

24 c

25 b

26 c

27 d

28 b

29 a

30 d

31 b

32 d

6

A n s w e r s

142

Chapter

7 Stalling

Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 145 Cause o the Stall . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 145 The Lif Curve . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 146 Stall Recovery . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 146 Aircraf Behaviour Close to the Stall . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 147 Use o Flight Controls Close to the Stall . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 147 Stall Recognition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 148 Stall Speed . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 148 Stall Warning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 150 Artificial Stall Warning Devices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 151 Basic Stall Requirements (EASA and FAR) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 154 Wing Design Characteristics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 154 The Effect o Aerooil Section . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 154 The Effect o Wing Planorm . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 156 Key Facts 1. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 162 Super Stall (Deep Stall). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 166 Super Stall Prevention - Stick Pusher . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 167 Factors That Affect Stall Speed. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 168 1g Stall Speed . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 168 Effect o Weight Change on Stall Speed. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 169 Composition and Resolution o Forces . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 170 Using Trigonometry to Resolve Forces . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 170 Lif Increase in a Level Turn. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 171 Effect o Load Factor on Stall Speed . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 172 Effect o High Lif Devices on Stall Speed . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 173 Effect o CG Position on Stall Speed . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 174 Effect o Landing Gear on the Stall Speed . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 175 Effect o Engine Power on Stall Speed. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 175 Continued Overlea

143

7

Stalling Effect o Mach Number (Compressibility) on Stall Speed . . . . . . . . . . . . . . . . . . . 177 Effect o Wing Contamination on Stall Speed. . . . . . . . . . . . . . . . . . . . . . . . . . 179 Warning to the Pilot o Icing-induced Stalls . . . . . . . . . . . . . . . . . . . . . . . . . . . 181 Stabilizer Stall Due to Ice . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 182 Effect o Heavy Rain on Stall Speed . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 182 Stall and Recovery Characteristics o Canards. . . . . . . . . . . . . . . . . . . . . . . . . . 182 Spinning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

183

Primary Causes o a Spin . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 183

7

Phases o a Spin . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 184

S t a l l i n g

The Effect o Mass and Balance on Spins . . . . . . . . . . . . . . . . . . . . . . . . . . . .

185

Spin Recovery . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 185 Special Phenomena o Stall . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 187 High Speed Buffet (Shock Stall) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 189 Answers to Questions on Page 173 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 191 Key Facts 2. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 192 Questions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

195

Key Facts 1 (Completed) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Key Facts 2 (Completed) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Answers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

206

Note: Throughout this chapter reerence will be made to EASA Certification Specifications (CS23, CS25) stall requirements etc, but it must be emphasised that these reerences are or training purposes only and are not subject to amendment action.

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Stalling Introduction Stalling is a potentially hazardous manoeuvre involving loss o height and loss o control. A pilot must be able to clearly and unmistakably identiy an impending stall so that it can be prevented. Different types o aircraf exhibit various stall characteristics, some less desirable than others. Airworthiness authorities speciy minimum stall qualities that an aircraf must possess.

Cause of the Stall The CL o an aerooil increases with angle o attack up to a maximum (C LMAX ). Any urther increase above this stalling angle, or critical angle o attack , will make it impossible or the airflow to smoothly ollow the upper wing contour, and the flow will separate rom the surace, causing CL to decrease and drag to increase rapidly. Since the C LMAX o an aerooil corresponds to the minimum steady flight speed (the 1g stall speed), it is an important point o reerence.

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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. Figure 7.1 shows that at low angles o attack

virtually no flow separation occurs beore the trailing edge, the flow being attached over the rear part o the surace in the orm o a turbulent boundary layer. As angle o attack increases, the adverse pressure gradient increases, reducing the kinetic energy, and the boundary layer will begin to separate rom the surace at the trailing edge. Further increase in angle o attack makes the separation point move orward and the wing area that generates a pressure differential becomes smaller. At angles o attack higher than approximately 16°, the extremely steep adverse pressure gradient will have caused so much separation that insufficient lif is generated to balance the aircraf weight.

Figure 7.1

It is important to remember that the angle o attack is the angle between the chord line and the relative airflow. Thereore, i the angle o attack is increased up to or beyond the critical angle, an aeroplane can be stalled at any airspeed or flight attitude.

An aeroplane can be stalled at any airspeed or attitude

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Stalling The Lift Curve

CLMAX CL 7

Stall

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0

4

8

12

Angle of Attack in Degrees

16

(α)

Figure 7.2

Figure 7.2 shows that as the angle o attack increases rom the zero lif value, the curve is linear

over a considerable range. As the effects o separation begin to be elt, the slope o the curve begins to all off. Eventually, lif reaches a maximum and begins to decrease. The angle at which it does so is called the stalling angle or critical angle o attack, and the corresponding value o lif coefficient is C LMAX. A typical stalling angle is about 16°.

Stall Recovery To recover rom a stall or prevent a ull stall, the angle o attack must be decreased to reduce the adverse pressure gradient . This may consist o merely releasing back pressure,

or it may be necessary to smoothly move the pitch control orward, depending on the aircraf design and severity o the stall. (Excessive orward movement o the pitch control, however, may impose a negative load on the wing and delay recovery). For most modern jet transport aircraf it is usually sufficient to lower the nose to the horizon or just below while applying maximum authorized power to minimize height loss. On straight wing aircraf the rudder should be used to prevent wing drop during stall and recovery. On swept wing aircraf it is recommended that the ailerons be used to prevent wing drop, with a small amount o smoothly applied co-ordinated rudder. (The rudder on modern high speed jet transport aircraf is very powerul, 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 actor and damage the aircraf structure. As angle o attack reduces below the critical angle, the adverse pressure gradient will decrease, airflow will re-attach, and lif and drag will return to their normal values.

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Stalling Aircraft Behaviour Close to the Stall Stall characteristics vary with different types o aircraf. However, or modern aircraf during most normal manoeuvres, the onset o stall is gradual. The first indications o a stall may be provided by any or all o the ollowing: • unresponsive flight controls, • a stall warning or stall prevention device, or • aerodynamic buffet. The detailed behaviour o various aircraf types will be discussed later.

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Use of Flight Controls Close to the Stall

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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 eel unresponsive or “mushy”. I an accidental stall does occur, it is vitally important that the stall and recovery should occur without too much wing drop. Moving a control surace modifies the chord line and, hence, the angle o attack. An aircraf being flown close to the stall angle may have one wing that produces slightly less lif than the other; that wing will tend to drop. Trying to lif a dropping wing with aileron will increase its angle o attack, Figure 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 aircraf the rudder should be used to yaw the aircraf just enough to increase the speed o a dropping wing to maintain a wing’s level attitude. Swept wing aircraf basic stall requirements are designed to enable the ailerons to be used successully up to ”stall recognition” ( Page 148 and Page 154), but small amounts o rudder can be used i smoothly applied and co-ordinated with the ailerons.

15º

22º

Figure 7.3

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Stalling Stall Recognition The aeroplane is considered stalled when the behaviour o the aeroplane gives the pilot a clear and distinctive indication o an acceptable nature that the aeroplane is stalled. Acceptable indications o a stall, occurring either individ ually or in combination, are: (1)

A nose-down pitch that cannot be readily arrested;

(2)

Buffeting, o a magnitude and severity that is a strong and effective deterrent to urther speed reduction; or

(3)

The pitch control reaches the af stop and no urther increase in pitch attitude occurs when the control is held ull af or a short time beore recovery is initiated.

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S t a l l i n g

Stall Speed It is necessary to fly at slow speeds (high angles o attack) during take-off and landing in order to keep the required runway lengths to a reasonable minimum. There must be an adequate saety margin between the minimum speed allowed or normal operations and the stall speed.

Prototype aircraf are stalled and stall speeds established or inclusion in the Flight Manual during the flight testing that takes place beore type certification. “Small” aircraf (CS-23) use V S0 and VS1 on which to base the stall speed. For “Large” aircraf (CS-25) a reerence stall speed, V SR , is used. • The reerence stall speed (VSR ) is a calibrated airspeed defined by the aircraf manuacturer. VSR may not be less than a 1g stall speed. V SR is expressed as: VSR

≥

VCLMAX √ nZW

Where: VCLMAX =

Calibrated airspeed obtained when the load actor corrected lif coefficient is first a maximum during the manoeuvre prescribed in the starred bullet point on page 149. In addition, when the manoeuvre is limited by a device that abruptly pushes the nose down at a selected angle o attack (e.g. a stick pusher), V CLMAX may not be less than the speed existing at the instant the device operates.

nZW

148

=

Load actor normal to the flight path at VCLMAX

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Stalling Note: On aircraf without a stick pusher, VSR can be considered to be the same as the 1g stall speed (VS1g ). But it is impossible to fly at speeds less than that at which the stick pusher activates, so or aircraf fitted with a stick pusher, V SR will be 2 knots or 2% greater than the speed at which the stick pusher activates. (See Figure 7.4 and Figure 7.5 or an illustration o the designations o stall speed and stall warning).

From the “sample” aeroplane on Page 76 , the speed at C LMAX was 150 kt. This can be considered as that aeroplane’s V CLMAX . At 1g, VSR would thereore be 150 kt. • VCLMAX is determined with: • Zero thrust at the stall speed.

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• Propeller pitch controls (i applicable) in the take-off position. • The aeroplane in other respects (such as flaps and landing gear) in the condition existing in the test or perormance standard in which V SR is being used. • The weight used when V SR is being used as a actor to determine compliance with a required perormance standard. • The centre o gravity position that results in the highest value o reerence stall speed; and • The aeroplane trimmed or straight flight at a speed selected by the manuacturer, but not less than 1.13VSR and not greater than 1.3V SR. • *Starting rom the stabilized trim condition, apply the longitudinal control to decelerate the aeroplane so that the speed reduction does not exceed one knot per second. • In addition to the requirements above, when a device that abruptly pushes the nose down at a selected angle o attack (e.g. a stick pusher) is installed, the reerence 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. VSR will vary with each o the above conditions. Additional actors which affect V SR are load actor, thrust in excess o zero and wing contamination. All these effects will be detailed later.

Density altitude does not affect indicated stall speed

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Stalling Stall Warning Having established a stall speed or each configuration, there must be clear and distinctive warning, sufficiently in advance o the stall , or the stall itsel 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 urnished either through the inherent aerodynamic qualities o the aeroplane or by a device that will give clearly distinguishable indications under expected conditions o flight. However, a visual stall warning device that requires the attention o the crew within the cockpit is not acceptable by itsel. I a warning device is used, it must provide a warning in each o the aeroplane configurations prescribed in sub-paragraph (a) o this paragraph at the speed prescribed in sub-paragraphs (c) and (d) o 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 Stall Recognition, on page 148, by not less than 5 knots or 5% CAS, whichever is the greater. Once initiated, stall warning must continue until the angle o attack is reduced to approximately that at which stall warning began.

(d)

In addition to the requirements o sub-paragraph (c) o this paragraph, when the speed is reduced at rates not exceeding one knot per second, in straight flight with engines idling and CG position specified on page 149, VSW, in each normal configuration, must exceed VSR 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 on page 148 - Stall Recognition) when recovery is initiated not less than one second afer the onset o stall warning in slow-down turns with at least 1.5g load actor normal to the flight path and airspeed deceleration rates o at least 2 knots per second, with the flaps and landing gear in any normal position, with the aeroplane trimmed or straight flight at a speed o 1.3VSR , and with the power or thrust necessary to maintain level flight at 1.3VSR .

()

Stall warning must also be provided in each abnormal configuration o the high lif devices that is likely to be used in flight ollowing system ailures (including all configurations covered by Flight Manual procedures).

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150

7

Stalling

V

C LMAX VSR V SW

VS1g 5 kt or 5%

CAS 1, 2 & 3 on page 148 7

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Figure 7.4 Aircraf without stick pusher

STICK PUSH VSR

VSW

2 kt or

3 kt or

2%

3% CAS

V

C LMAX

Figure 7.5 Aircraf with stick pusher

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 fi tted to all aircraf. Artificial stall warning on small aircraf is usually given by a buzzer or horn. The artificial stall warning device used on modern large aircraf is a stick shaker, in conjunction with lights and a noisemaker.

Stick Shaker A stick shaker represents what it is replacing; it shakes the stick and is a tactile warning. I the stick shaker activates when the pilot’s hands are not on the controls, when the aircraf is on autopilot, or example, a very quiet stick shaker could not unction as a stall warning so a noisemaker is added in parallel. The stick shaker is a pair o simple electric motors, one clamped to each pilot’s control column, rotating an out o balance weight. When the motor runs, it shakes the stick.

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Stalling An artificial stall warning device can receive its signal rom a number o different types o detector switch, all activated by changes in angle o attack.

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FLAPPER SW ITCH ( activated by movement of stagnation point )

STAGNATION POINT ( has moved downwards and backwards around leading edge ) Figure 7.6 Flapper switch

Flapper Switch (Leading Edge Stall Warning Vane) Figure 7.6 . As angle o attack increases, the stagnation point moves downwards and backwards

around the leading edge. The flapper switch is so located that, at the appropriate angle o attack, the stagnation point moves to its underside and the increased pressure lifs and closes the switch.

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Stalling

AS ANGLE OF ATTACK INCREASES, VANE ROTATES RELATIVE TO FUSELAGE

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VANE

FUSELAGE SKIN

Figure 7.7 Angle o attack vane

Angle of Attack Vane Figure 7.7 . Mounted on the side o the uselage, the vane streamlines with the relative airflow

and the uselage rotates around it. The stick shaker is activated at the appropriate angle o attack.

Angle of Attack Probe Also mounted on the side o the uselage, it consists o slots in a probe, which are sensitive to changes in angle o relative airflow. All o these sense angle o attack and, thereore, automatically take care o changes in aircraf mass; the majority also compute the rate o change o angle o attack and give earlier warning in the case o aster rates o approach to the stall. The detectors are usually datum compensated or configuration changes and are always heated or anti-iced. There are usually sensors on both sides to counteract any sideslip effect.

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Stalling Basic Stall Requirements (EASA and FAR) • It must be possible to produce and to correct roll and yaw by unreversed use o aileron and rudder controls, up to the time the aeroplane is stalled. No abnormal nose-up pitching may occur. The longitudinal control orce must be positive up to and throughout the stall. In addition, it must be possible to promptly prevent stalling and to recover rom a stall by normal use o the controls . • For level wing stalls, the roll occurring between the stall and the completion o the recovery may not exceed approximately 20°.

7

• For turning flight stalls, the action o the aeroplane afer 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 o the aeroplane. The maximum bank angle that occurs during the recovery may not exceed:

S t a l l i n g

• Approximately 60 degrees in the original direction o the turn, or 30 degrees in the opposite direction, or deceleration rates up to 1 knot per second; and • Approximately 90 degrees in the original direction o the turn, or 60 degrees in the opposite direction, or deceleration rates in excess o 1 knot per second.

Wing Design Characteristics It has been shown that stalling is due to airflow separation, characterized by a loss o lif, and an increase in drag, that will cause the aircraf to lose height. This is generally true, but there are aspects o aircraf behaviour and handling at or near the stall which depend on the design o the wing aerooil section and planorm.

The Effect of Aerofoil Section Shape o the aerooil section will influence the manner in which it stalls. With some sections, stall occurs very suddenly and the drop in lif is very marked. With others, the approach to stall is more gradual, and the decrease in lif is less disastrous. In general, an aeroplane should not stall too suddenly, and the pilot should have adequate warning, in terms o handling qualities, o the approach o a stall. This warning generally takes the orm o buffeting and general lack o response to the controls. I a particular wing design stalls too suddenly, it will be necessary to provide some sort o artificial pre-stall warning device or even a stall prevention device. A given aerooil section will always stall at the same angle o attack

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7

Stalling Features o aerooil section design which affect behaviour near the stall are: • • • •

leading edge radius, thickness-chord ratio, camber, and particularly the amount o camber near the leading edge, and chordwise location o the points o maximum thickness and maximum camber.

Generally, the sharper the nose (small leading edge radius), the thinner the aerooil section, or the urther af the position o maximum thickness and camber, the more sudden will be the stall. e.g. an aerooil section designed or efficient operation at higher speeds, Figure 7.8. The stall characteristics o the above listed aerooil sections can be used to either encourage a stall to occur, or delay stalling, at a particular location on the wingspan.

7

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1. ROUNDED LEADING EDGE

CL

2. HIGHER THICKNESS-CHORD RATIO 3. MAX. THICKN ESS AND CA MBER MORE FW D.

1. SHARP LEADING EDGE 2. LOW T HICKNESS-CHO RD RATIO 3. AFT MAXIMUM THICKNESS AND CAMBER

Figure 7.8

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Stalling The Effect of Wing Planform On basic wing planorms, airflow separation will not occur simultaneously at all spanwise locations.

STRONG TIP VORTICES DECREASE EFFECTIVE ANGLE OF ATTACK AT W ING TIP, THUS DELAYING TIP STALL.

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S t a l l i n g

CP

CP MOVES REARWARDS, AIRCRAFT NOSE DROPS.

Figure 7.9 Rectangular wing

The Rectangular Wing Figure 7.9. On a rectangular wing, separation tends to begin at the root and spreads out

towards the tip. Reduction in lif initially occurs inboard near the aircraf CG, and i it occurs on one wing beore the other, there is little tendency or the aircraf to roll . The aircraf loses height, but in doing so it remains more or less wings level. Loss o lif is elt ahead o the centre o gravity o the aircraf and the CP moves rearwards, so the nose drops and angle o attack is reduced . Thus, there is a natural tendency or the aircraf to move away rom

the high angle o attack which gave rise to the stall. The separated airflow rom the root immerses the rear uselage and tail area, and aerodynamic buffet can provide a warning o the approaching stall. Being located outside o the area o separated airflow, the ailerons tend to remain effective when the stalling process starts. All o these actors give the most desirable kind o response to a stall: • • • •

aileron effectiveness, nose drop, aerodynamic buffet, and absence o violent wing drop.

Unortunately, a rectangular wing has unacceptable wing bending characteristics and is not very aerodynamically efficient, so most modern aircraf have a tapered and/or swept planorm.

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Stalling

W ING TIP IS UNABLE TO SUPPORT TIP VORTICES, CAUSING THEM TO FORM CLOSER TO THE ROOT.

CP

THIS GIVES A DECREASED EFFECTIVE ANGLE OF ATTACK AT THE W ING ROOT, THUS DELAYING THE ROOT STALL. 7

g n i l l a t S

Figure 7.10 Tapered wing

The Tapered Wing Figure 7.10. Separation tends to occur first in the region o the wing tips, reducing lif in those

areas. I 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 or an aircraf to enter a spin). There would be no buffet on the tail, no strong nose-down pitching moment and very little, i any, aileron effectiveness. To give avourable stall characteristics, a tapered wing must be modified using one or more o the ollowing: • Geometric twist ( washout), a decrease in incidence rom root to tip. This decreases the angle o attack at the tip, and the root will tend to stall first. • The aerooil section may be varied throughout the span such that sections with greater thickness and camber are located near the tip. The higher CLMAX o such sections delays stall so that the root will tend to stall first.

AILERON

SLOT

A A SECTION

A-A

Figure 7.11 Leading edge slot

• Leading edge slots, Figure 7.11, towards the tip re-energize (increase the kinetic energy o) the boundary layer. They increase local C LMAX and are useul, both or delaying separation at the tip and retaining aileron effectiveness. The unction o slats and slots will be ully described in Chapter 8.

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7

Stalling

7

S t a l l i n g

STALL STRIP

Figure 7.12 Stall strip

• Another method or improving the stall pattern is by orcing a stall to occur rom the root. An aerooil section with a smaller leading edge radius at the root would promote airflow separation at a lower angle o attack but decrease overall wing efficiency. The same result can be accomplished by attaching stall strips (small triangular strips), Figure 7.12, to the wing leading edge. At higher angles o attack, stall strips promote separation, but they will not effect the efficiency o the wing in the cruise.

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Stalling

VORTEX GENERATORS

7

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Figure 7.13 Vortex generators

• Vortex generators, Figure 7.13, are rows o small, thin aerooil shaped blades which project vertically (about 2.5 cm) into the airstream. They each generate a small vortex which causes the ree stream flow o high energy air to mix with and add kinetic energy to the boundary layer. This re-energizes the boundary layer and tends to delay separation.

159

7

Stalling

LATERAL AXIS

OUTBOARD SUCTION

CP

PRESSURES TEND TO DRAW BOUNDARY LAYER TOWARDS TIP. CP MOVES FORWA RD A ND

7

CREATES AN UNSTABLE NOSE-UP PITCHING MOMENT

S t a l l i n g

Figure 7.14

Sweepback Figure 7.14. A swept wing is fitted to allow a higher maximum speed, but it has an increased tendency to stall first near the tips. Loss o lif at the tips moves the CP orward, giving a nose-up pitching moment .

Effective lif production is concentrated inboard and the maximum downwash now impacts the tailplane, Figure 7.15, adding to the nose-up pitching moment.

Pitch-up As soon as a swept wing begins to stall, both or ward CP movement and increased downwash at the tailplane cause the aircraf nose to rise rapidly, urther increasing the angle o attack. This is a very undesirable and unacceptable response at the stall and can result in complete loss o control in pitch rom which it may be very difficult, or even impossible, to recover. This phenomenon is known as pitch-up, and is a very dangerous characteristic o many high speed, swept wing aircraf .

TIP STALL

UNSTALLED

CP STALLED

MAXIMUM DOWNWASH

Figure 7.15 Pitch-up

160

STALLED

7

Stalling The tendency o a swept-back wing to tip stall is due to the induced spanwise flow o the boundary layer rom root to tip . The ollowing design eatures can be incorporated to

minimize this effect and give a swept wing aircraf more acceptable stall characteristics:

WING FENCE

7

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Figure 7.16

Wing ences (boundary layer ences), Figure 7.16 , are thin metal ences which generally extend

rom the leading edge to the trailing edge on the top surace and are intended to prevent outward drif o the boundary layer.

VORTILON

SAW TOOTH

ENGINE PYLON

Figure 7.17 Vortilon

Figure 7.18 Saw tooth

Vortilons, Figure 7.17 , are also thin metal ences, but are smaller than a ull chordwise ence.

They are situated on the underside o the wing leading edge. The support pylons o pod mounted engines on the wing also act in the same way. At high angles o attack a small but intense vortex is shed over the wing top surace which acts as an aerodynamic wing ence. Saw tooth leading edges, Figure 7.18, will also generate a strong vortex over the wing upper

surace at high angles o attack, minimizing spanwise flow o the boundary layer. (Rarely used on modern high speed jet transport aircraf).

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Stalling Key Facts 1 Self Study The ollowing our pages contain a revision aid to encourage students to become amiliar with any new terminology, together with the key elements o “stalling”. Insert the missing words in these statements, using the oregoing paragraphs or reerence. Stalling involves loss o ________ and loss o _________.

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S t a l l i n g

A pilot must be able to clearly and unmistakably ___________ a stall. A stall is caused by airflow _____________. Separation can occur when either the boundary layer has insufficient _________ energy or the _________ ___________ gradient becomes too great. Adverse pressure gradient increases with increase in angle o ________. Alternative names or the angle o attack at which stall occurs are the _______ angle and the __________ angle o attack. The coefficient o lif at which a stall occurs is ________. A stall can occur at any ___________ or flight ___________. A typical stalling angle is approximately ____°. To recover rom a stall the angle o ________ must be ___________. Maximum power is applied during stall recovery to minimize _________ loss. On small aircraf, the _________ should be used to prevent wing _______ at the stall. On swept wing aircraf, the _______ should be used to prevent wing _____ at the stall. Recover height lost during stall recovery with moderate _______ pressure on the _________ control.

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7

Stalling The first indications o a stall may be _____________ flight controls, stall _________ device or aerodynamic ________. At speeds close to the stall, __________ must be used with caution to ______ a dropping wing. Acceptable indications o a stall are: (1) (2) (3)

a nose ______ pitch that can not be readily arrested. severe ___________. pitch control reaching _____ stop and no urther increase in _______ attitude occurs.

7

g n i l l a t S

Reerence stall speed (V SR ) is a CAS defined by the ________ ___________. VSR may not be _____ than a ____ stall speed. When a device that abruptly pushes the _____ _____ at a selected angle o ______ is installed, VSR may not be _____ than ___ knots or ___ %, whichever is ______, above the speed at which the ________ operates. Stall warning with sufficient _______ to prevent inadvertent stalling must be ______ and ______________ to the pilot in straight and turning flight. Acceptable stall warning may consist o the inherent ____________ qualities o the aeroplane or by a ___________ that will give clearly distinguishable indications under expected conditions o flight. Stall warning must begin at a speed exceeding the stall speed by not less than __ knots or __ % CAS, whichever is the greater. Artificial stall warning on a small aircraf is usually given by a ______ or ________. Artificial stall warning on a large aircraf is usually given by a _______ shaker, in conjunction with ________ and a noisemaker. An artificial stall warning device can be activated by a _________ switch, an angle o ________ vane or an angle o attack _______. Most angle o attack sensors compute the ______ o change o angle o attack to give _________ warning in the case o accelerated rates o stall approach.

163

7

Stalling EASA required stall characteristics, up to the time the aeroplane is stalled, are:

7

S t a l l i n g

a.

It must be possible to produce and correct ____ by unreversed use o the ________ and ________.

b.

No abnormal nose-up ________ may occur.

c.

Longitudinal control orce must be ________.

d.

It must be possible to promptly prevent ________ and recover rom a stall by normal use o the ________.

e.

There should be no excessive ____ between the stall and completion o recovery.

.

For turning flight stalls, the action o the aeroplane afer the stall may not be so _______ or _______ as to make it difficult, with normal piloting _____, to effect prompt _________ and to regain _______ o the aeroplane.

An aerooil section with a small leading edge ______ will stall at a _______ angle o attack and the stall will be more ______. An aerooil section with a large thickness-chord ratio will stall at a ______ angle o attack and will stall more ______. An aerooil section with camber near the ________ attack.

______ will stall at a higher angle o

A rectangular wing planorm will tend to stall at the ____ first. A rectangular wing planorm usually has ideal stall characteristics; these are: a.

Aileron _____________ at the stall.

b.

Nose _____ at the stall.

c.

Aerodynamic _______ at the stall.

d.

Absence o violent wing _____ at the stall.

To give a wing with a tapered planorm the desired stall characteristics, the ollowing devices can be included in the design:

164

a.

________ (decreasing incidence rom root to tip).

b.

An aerooil section with ________ thickness and camber at the tip.

c.

Leading edge ______ at the tip.

d.

Stall _______ fitted to the wing inboard leading edge.

e.

_______ generators which re-energize the _________ layer at the tip.

7

Stalling A swept-back wing has an increased tendency to tip stall due to the spanwise flow o boundary layer rom root to tip on the wing top surace. Methods o delaying tip stall on a swept wing planorm are: a.

Wing _______, thin metal ences which generally extend rom the leading edge to the trailing edge on the wing top surace.

b.

_________, also thin metal ences, but smaller and are situated on the underside o the wing leading edge.

c.

Saw _____ leading edge, generates vortices over wing top surace at high angles o attack.

d.

Engine _______ o pod mounted wing engines also act as vortilons.

e.

_______ generators are also used to delay tip stall on a swept wing.

7

g n i l l a t S

Tip stall on a swept wing planorm gives a tendency or the aircraf to _____-___ at the stall. This is due to the ___ moving orwards when the wing tips stall ______. KEY FACTS 1, WITH WORD INSERTS CAN BE FOUND ON page 201.

165

7

Stalling Super Stall (Deep Stall) A swept-back wing tends to stall first near the tips. Since the tips are situated well af o the CG, the loss o lif at the tips causes the pitch attitude to increase rapidly and urther increase the angle o attack. Figure 7.19.

7 PITCH - UP

S t a l l i n g

TIP STALL

Figure 7.19 Pitch-up

This “automatic” increase in angle o attack, caused by pitch-up, stalls more o the wing. Drag will increase rapidly, lif 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 o attack, Figure 7.20.

DOWNWA RD INCLINED FLIGHT PATH

TAILPLANE IMMERSED IN SEPARATED A IRFLOW FROM STALLED WING

Figure 7.20 Super stall

Separated airflow rom the stalled wing will immerse a high-set tailplane in low energy turbulent air, Figure 7.20. Elevator effectiveness is greatly reduced making it impossible or the pilot to decrease the angle o attack. The aeroplane will become stabilized in what is known as the “super stall” or “deep stall” condition.

166

7

Stalling Clearly, the combination o a swept-back wing and a high mounted tailplane (‘T’ - Tail) are the actors involved in the “super or deep stall”. O the two: THE SWEPT-BACK WING IS THE MAJOR CONTRIBUTORY FACTOR.

It has been shown that the tendency or a swept-back wing to pitch-up can be reduced by design modifications (wing ences, vortilons and saw tooth leading edges) which minimize the root-to-tip spanwise flow o the boundary layer. These devices delay tip stall. Vortex generators are also requently 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 modiying the aerooil section at the root, fitting stall strips and by fitting less efficient leading edge flaps (Kruger flaps) to the inboard section o the wing.

7

g n i l l a t S

Aircraf 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, uselage mounted engines. The only contribution rear mounted engines make is that they are the reason the designer placed the tailplane on top o the fin in the first place. In and o itsel, mounting the engines on the rear uselage does not contribute to super stall.

Super Stall Prevention - Stick Pusher An aircraf design which exhibits super stall characteristics must be fitted with a device to prevent it rom ever stalling . This device is a stick pusher. Once such an aircraf begins to stall it is too late; the progression to super stall is too ast or a human to respond, and the aircraf cannot then be un-stalled. A stick pusher is a device, attached to the elevator control system, which physically pushes the control column orward, reducing the angle o attack beore super stall can occur. The orce o the push is typically about 80 lb. 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 o a malunction. Once dumped, the pusher cannot normally be reset in flight. Once actuated, the stick pusher will automatically disengage once the angle o attack reduces below a suitable value.

167

7

Stalling Factors that Affect Stall Speed Page 148 details the CAS at which an aircraf stalls (V SR). We know that stalling is caused by exceeding the critical angle o attack. Stalling has nothing to do with the speed o the aircraf; the critical angle o attack can be exceeded at any aircraf speed . However, it has been shown that i an aircraf 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 reerence stall speed (VSR) that the recommended take-off, manoeuvre, approach and landing speeds are based, to give an adequate margin rom the stall during normal operations (1.05V SR, 1.1VSR, 1.2VSR, 1.3VSR etc).

7

Factors which can affect V SR are:

S t a l l i n g

• • • • • • • •

Changes in weight. Manoeuvring the aircraf (increasing the load actor). Configuration changes (changes in CLMAX and pitching moment). CG position. Engine thrust and propeller slipstream. Mach number. Wing contamination. Heavy rain.

1g Stall Speed In straight and level flight the weight o the aircraf is balanced by the lif. Load Factor (n) or ‘g’ =

Lif Weight

While (n) is the correct symbol or load actor, the relationship between lif and weight has or years been popularly known as ‘g’. (1g corresponds to the orce acting on us in every day lie). I more lif is generated than weight, the load actor or ‘g’ will be greater than one; the orce acting on the aircraf and everything in it, including the pilot, will be greater. I Lif = Weight, the load actor will be one and rom the lif ormula: L = ½ ρ V CL S 2

it can be seen that lif will change whenever any o the other actors in the ormula change. We consider density ( ρ) and wing area (S) constant or this example . I the engine is throttled back, drag will reduce speed (V) and, rom the ormula, it can be seen that lif would decrease. To keep lif constant and maintain 1g flight at a reduced speed, C L must be increased by increasing the angle o attack.

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7

Stalling Any urther reduction in speed would need a ur ther increase in angle o attack, each succeeding lower CAS corresponding to a greater angle o attack. Eventually, at a certain CAS, the wing reaches its stalling angle (C LMAX), beyond which any urther increase in angle o attack, in an attempt to maintain lif, will precipitate a stall. We can transpose the lif ormula to show this relationship:VS1g

√

=

L ½ ρ CLMAX S

Density altitude does not affect indicated stall speed

Effect of Weight Change on Stall Speed 7

At CLMAX or 1g flight, a change in weight requires a change in lif and it can be seen rom the VS1g ormula that, or instance, an increase in weight (lif) will increase V S1g

g n i l l a t S

The relationship between basic stalling speeds at two different weights can be obtained rom the ollowing ormula: VS1g new

=

VS1g old

√

new weight old weight

The angle o attack at which stall occurs will NOT be affected by the weight. (Provided that the appropriate value o CLMAX is not affected by speed - as it will be at speeds greater than M0.4, see page 177 ). To maintain a given angle o attack in level flight, it is necessary to change the dynamic pressure (CAS) i the weight is changed . As an example: at a weight o 588 600 N an aircraf stalls at 150 kt CAS. What is the V S1g stall speed at a weight o 470 880 N? VS1g new = =

150

√

470880 588600

Weight does not affect stall angle

134 knots CAS

It should be noted that a 20% reduction in weight has resulted in an approximate 10% reduction in stall speed. (As a “rule o 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.

169

7

Stalling Composition and Resolution of Forces A orce is a vector quantity. It has magnitude and direction, and it can be represented by a straight line passing through the point at which it is applied, its length representing the magnitude o the orce, and its direction corresponding to that in which the orce is ac ting.

FORCE

FORCE

VECTOR

VECTOR

FORCE

7

FORCE

S t a l l i n g

VECTOR

FORCE

FORCE VECTOR

Figure 7.21 The resolution o a orce into two vectors and the addition o vectors to orm a resultant

As vector quantities, orces can be added or subtracted to orm a resultant orce, or they can be resolved - split into two or more component parts by the simple process o drawing the vectors to represent them . Figure 7.21.

Using Trigonometry to Resolve Forces I one o the angles and the length o one o the sides o a right angled triangle are known, it is possible to calculate the length o the other sides using trigonometry. This technique is used when resolving a orce into its horizontal and vertical components.

Hypotenuse Opposite Ad jacen t

TAN

=

Opp Ad j

SIN

=

Figure 7.22

170

Opp Hyp

COS

=

Ad j Hyp

7

Stalling Lift Increase in a Level Turn

45º

A DJACENT

1

LIFT INCREASE REQUIRED

L HYPOTENUSE 7

g n i l l a t S

W Figure 7.23

Figure 7.23 shows an aircraf in a level 45° bank turn. Weight always acts vertically downwards.

For the aircraf to maintain altitude, the UP orce must be the same as the DOWN orce. Lif is inclined rom the horizontal by the bank angle o 45° and can be resolved into two components, or vectors; one vertical and one horizontal. It can be SEEN rom the illustration that in a level turn, lif must be increased in order to produce an upwards orce vector equal to weight. We know the vertical orce must be equal to the weight, so the vertical orce can be represented by 1. The relationship between the vertical orce and lif can be ound using trigonometry, where φ (phi) is the bank angle: cos φ =

ADJ (1) HYP (L)

transposing this ormula gives, L =

1 cos φ

In this case φ = 45 degrees L =

1 0.707

= 1.41

This shows that: In a 45° bank, LIFT must be greater than weight by a actor o 1.41 Another way o saying the same thing: in a level 45° bank turn, lif must be increased by 41%.

171

7

Stalling Effect of Load Factor on Stall Speed It has been demonstrated that to bank an aircraf and maintain altitude, lif has to be greater than weight. And that additional lif in a turn is obtained by increasing the angle o attack. To consider the relationship between lif and weight we use Load Factor. LOAD FACTOR (n) or ‘g’ =

(a) (b)

7

LIFT WEIGHT

Increasing lif in a turn, increases the load actor. As bank angle increases, load actor increases .

In straight and level flight at C LMAX it would be impossible to turn AND maintain altitude. Trying to increase lif would stall the aircraf. I a turn was started at an IAS above the stall speed, at some bank angle C L would reach its maximum and the aircraf would stall at a speed higher than the 1g stall speed.

S t a l l i n g

The increase o lif in a level turn is a unction o the bank angle only . Using the ollowing

ormula, it is possible to calculate stall speed as a unction o bank angle or load actor. V St is the stall speed in a turn VSt =

VS

√

1 cos φ

Load actor does not affect stall angle

Using our example aeroplane: the 1g stall speed is 150 knots CAS, so what will be the stall speed in a 45° bank? VSt =

150

√

1 0.707

= 178 knots CAS

In a 60° bank the stall speed will be: VSt =

150

√

1 0.5

= 212 knots CAS

Stall speed in a 45° bank is 19% greater than V S1g and in a 60° bank the stall speed is 41% greater than V S1g, and since these are ratios, this will be true or any aircraf.

172

7

Stalling As bank angle is increased, stall speed will increase at an increasing rate. While operating at high CL, during take-off and landing in particular, only moderate bank angles should be used to manoeuvre the aircraf. For a modern high speed jet transport aircraf, the absolute maximum bank angle which should be used in service is 30° (excluding emergency manoeuvres). The normal maximum would be 25°, but at higher altitude the normal maximum is 10° to 15°. I the 1g stall speed is 150 kt, calculate the stall speed in a 25° and a 30° bank turn. (Answers on page 191). I the stall speed in a 15° bank turn is 153 kt CAS and it is necessary to calculate the stall speed in a 45° bank turn, you would need to calculate the 1g stall speed first, as ollows: VSt =

VS1g =

VS1g

√

153 1.02

1 cos 15°

=

transposition gives V S1g =

7

VSt

√

g n i l l a t S

1 cos 15°

150 kt CAS

Effect of High Lift Devices on Stall Speed Modern high speed jet transport aircraf have swept wings with relatively low thickness/chord ratios (e.g. 12% or an A310). The overall value o CLMAX or these wings is airly 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 o C LMAX. In addition to decreasing the stall speed, these high lif devices will usually alter the stalling characteristics. The devices include: a)

leading edge flaps and slats

b)

trailing edge flaps

From the 1g stall ormula:

VS1g =

√

L ½ ρ CLMAX S

it can be seen that an increase in C LMAX will reduce the stall speed. It is possible, with the most modern high lif devices, to increase C LMAX by as much as 100%. High lif devices will be ully described in Chapter 8. High lif devices decrease stall speed, hence minimum flight speed, so provide a shorter take-off and landing run - this is their sole purpose.

173

7

Stalling Effect of CG Position on Stall Speed CS-25.103(b) states that V CLMAX is determined with the CG position that results in the highest value o reerence stall speed.

L

7

S t a l l i n g

CP

TAIL DOWNLOAD

W

Figure 7.24

I the CG is in ront o the CP, Figure 7.24, 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. Lif must be increased to maintain an upwards orce equal to the increased downwards orce. From the 1g stall ormula it can be seen that C LMAX will be divisible into the increased lif orce more times.

VS1g =

√

L ½ ρ CLMAX S

Forward movement o the CG increases stall speed.

174

7

Stalling Effect of Landing Gear on the Stall Speed

L

7

CP

g n i l l a t S

TAIL DOWNLOAD PROFILE DRAG FROM GEAR

W

Figure 7.25

From Figure 7.25 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. Lif must be increased to balance balance the increased downwards downwards orce. orce. CG movement due to the direction in which the undercarriage extends will have an insignificant influence on stall speed. By ar the greater greater influence is the increased increased profile drag o the gear when it is extended. Extending the undercarriage increases stall speed .

Effect of Engine Power on Stall Speed CS-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 o the aircraf is entirely supported by lif. I thrust is applied close to to the stall, the nose high attitude o the aircraf produces a vertical component o thrust, Figure 7.27 , which assists in supporting the weight and less lif is required. Aircraf with propellers will have have an additional effect caused by the propeller slipstream. The most important impor tant actors affecting this relationship are engine type (propeller or jet), thrust to weight ratio ratio and inclination o the thrust vector vector at C LMAX.

175

7

Stalling

INDUCED FLOW FROM PROPELLER SLIPSTREAM

7

S t a l l i n g

Figure 7.26

Propeller

greater than the the ree stream flow, flow, Figure 7.26 . The slipstream velocity behind the propeller is greater 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 generat generates es much more lif than at zero thrust. The lif o the aeroplane at a given angle o o attack and airspeed will be greatly greatly affected. I the aircraf is in the the landing flare, reducing power suddenly will cause a significant reduction in lif and a heavy landing could result. On the other hand, a potentially potentially heavy landing can be avoided by a judicious ‘blast’ rom the engines. Jet The typical jet aircraf does not experience the induced flow velocities encountered encountered in propeller driven aeroplanes, thus the only significant actor is the vertical component o thrust, Figure 7.27 . Since this vertical component contributes contributes to supporting supporting the weight o the aircraf, less aerodynamic lif is required to hold the aeroplane aeroplane in flight. I the thrust is large and is given a large inclination at maximum maximum lif angle, the effect on stall speed can be very large. large. Since there is very little induced flow rom the jet, the angle o attack at stall is essentially the same poweron as power-off.

VERTICAL COMPONENT OF THRUST

Figure 7.27

Power-on stall speed is less than power-off . This will be shown shown to be significant during the study o windshear in Chapter 15.

176

7

Stalling Effect of Mach Number (Compressibility) on Stall Speed As an aircraf flies aster, aster, the streamline pattern around the wing changes. Faste Fasterr than about our tenths tenths the speed o sound (M 0.4) these these changes start to become significant. significant. This phenomena is known known as compressibility. compressibility. This will be discussed ully in Chapter 13. Pressure waves, generated by the passage o a wing through the air, propagate ahead o the wing at the speed o sound. These pressure waves waves upwash air ahead o the wing towards the lower pressure on the top surace.

7

g n i l l a t S

HIGH SPEED

LOW SPEED

Figure 7.28

Figure 7.28 shows that at low speed, the streamline pattern is affected ar ahead o the wing

and the air has a certain distance in which to upwash. upwash. As speed increases, the wing gets closer closer to its leading pressure wave, and the streamline pattern is affected a shorter shor ter 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 separation occurs at a reduced angle o attack. attack. Above M0.4 C LMAX decreases as shown in Figure 7.29.

C LMAX

1 0

0 4

M

Figure 7.29

177

7

Stalling Reerring to the 1g stall speed ormula: VS1g =

√

L ½ ρ CLMAX S

I CLMAX decreases, V S1g will increase. To maintain a constant EAS as altitude increases, TAS is increased. increased. Also, outside air temperature decreases with increasing increasing altitude, causing the local speed o sound to decrease. Mach number is proportional to TAS and inversely proportional to the local speed o sound (a): M =

7

S t a l l i n g

TAS a

Thereore, at a constant EAS, Mach number will increase as altitude increases. A lt

1g Stall Speed

EA S Figure 7.30

Figure 7.30 shows the variation o stalling speed with altitude at constant constant load actor (n). Such

a curve is called the stalling boundary or the given load actor, in which altitude is plotted against equivalent airspeed. At this load actor (1g), the aircraf aircraf cannot fly at speeds speeds to the lef o this boundary. It is clear that over the the lower range o altitude, stall speed speed does not vary with altitude. This is because at these these low altitudes, the Mach Mach number at V S is less than M 0.4, too low or compressibility effects to be present. Eventu Eventually ally (approximately (app roximately 30 000 f), Mach number at VS 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 rom earlier earlier,, the V S1g o 150 kt is equal to M 0.4 at approximately 29 000 f using u sing ISA values. As altitude increases, stall speed is initially constant then increases, due to compressibility.

178

7

Stalling Effect of Wing Contamination on Stall Speed Reer to: AIC 106/2004 “Frost Ice and Snow on Aircraf”, Aircraf”, and AIC 98/1999 98/1999 “T “Turbo-Prop urbo-Prop and other Propeller Propeller Driven Aeroplanes: Aeroplanes: Icing Induced Stalls”. Any contamination on the wing, but particularly ice, rost or snow, will drastically alter the aerodynamic contour and affect the nature o the boundary layer. layer. ICE

The ormation o ice on the leading edge o the wing will produce:

7

a) Large changes in the local contour, contour, leading to severe severe local adverse pressure gradients. b) High surace riction and a considerable reduction reduction o boundary layer layer kinetic energy. energy.

g n i l l a t S

These cause a large decrease in C LMAX and can increase stall speed by approximately 30% with no change in angle o attack . The added weight o the ice will also increase the stall speed, but the major actor is the reduction in C LMAX. FROST The effect o rost is more subtle. subtle. The accumulation o a hard coat o rost on the wing

upper surace will produce p roduce a surace texture o considerable roughness. Tests have shown that ice, snow or rost, with the thickness and surace roughness similar to medium or coarse sandpaper on the leading edge and upper surace o a wing can reduce lif 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 surace roughness increases skin riction and reduces the kinetic energy o the boundary layer. Separation will occur at an angle o attack and lif coefficients lower than or the clean smooth wing. SNOW The effect o snow can be similar similar to rost in that that it will increase increase surace roughness. I

there is a coating o snow on the aircraf it must be removed beore flight. Not only will the snow itsel increase increase skin riction drag, it may obscure airrame 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 o take-off. It is very important impor tant that the holdover time o any de-icing or anti-icing fluid applied to the airrame air rame is known. I this time will be exceeded beore take-off, take-off, the aircraf must be treated again.

179

7

Stalling While the reduction in C LMAX due to rost ormation is not usually as great as that due to ice ormation, 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 LMAX. However However,, kinetic energy o the boundary layer is an important actor influencing separation o the airflow and this energy is reduced reduced by an increase in surace roughness. The general effects effects o ice and rost ormation on CLMAX is typified by the illustrations in Figure 7.31. Ice, rost and snow change the aerooil aerooil section, decrease the stall angle and increase the stall speed

7

S t a l l i n g LEADING EDGE ICE FORMATION FORMATIO N

UPPER SURFACE FROST

BASIC SMOOTH WING C LMAX

CL W ING W ITH FROST FROST

W ING W ITH ICE ICE

ANGLE OF ATTACK

Figure 7.31

The increase in stall speed due to ice ormation is not easy to quantiy, as the accumulation and shape o the ice ormation is impossible to predict. Even a little ice is too much . Ice or rost must never be allowed to remain on any aerodynamic suraces in flight, nor must ice, rost, snow or other contamination be allowed to remain on the aircraf immediately beore flight.

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Stalling Warning Warn ing to the Pilot of Icing-induced Stalls There have been recent cases involving loss o control in icing conditions due to undetected stalling at speeds significantly above the normal stalling speed, accompanied by violent roll oscillations. Control o an aeroplane can be lost as a result o an icing-induced stall, the onset o which can be so insidious* as to be difficult to detect. The ollowing advice is offered on the recognition o, and the recovery rom, insidious icinginduced wing-stalls:

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a)

Loss o perormance in icing conditions may indicate a serious build-up o airrame icing (even i this cannot be seen) which causes a gradual loss o lif and a significant increase in drag;

b)

this build-up o ice can cause the aeroplane to stall at approximately 30% above the normal stall speed;

c)

the longitudinal characteristics o 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 o attack will be below that required to trigger the switch), so should not not be relied upon to give a warning o this condition. condition. Airrame buffet, however, however, may assist in identiying the onset o 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;

)

a combination o rolling oscillation and onset o high drag can cause the aeroplane to enter a high rate o descent unless prompt p rompt recovery action is taken;

g)

i 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 o attack). attack). The de-icing system system should also be activated. I the aeroplane is fitted fitted with an anti-icing system this this should have been activated prior to entry into icing conditions in accordance with the Flight Manual/ Operations Manual procedures procedures and recommendations. I the anti-icing system has not not been in use, then then it should be immediately activated. activated. Consideration should should also be given to leaving icing conditions by adjusting track and/or altitude i possible.

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*Insidious - advancing advancing imperceptibly: imperceptibly: without warning

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Stalling Stabilizer Stall Due to Ice The tailplane is an aerooil, and because it is thinner than the wing, it is likely to experience icing beore the wing does. The effect will be the same as or the the wing; the stall will occur at a lower angle o attack. The tailplane is normally operating at a negative negative angle o attack, producing a down load, so i the tailplane stalls and the down load is lost, the nose o the aircraf will drop and longitudinal control will be lost. Stalling o an ice contaminated tailplane could be precipitated by extension o the wing flaps. Lowering the flaps increases the downwash, and this increases the negative angle o attack o the tailplane. I the tailplane has ice contamination, contamination, this could be sufficient to cause it to stall. Recovery procedure procedure in this situation situation would be to retract retract the flaps again, thus reducing reducing the downwash.

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Effect of Heavy Rain on Stall Speed Weight Heavy rain will orm a film o water on an aircraf and increase its weight slightly, maybe as much as 1 - 2% this in itsel will increase stall speed. Aerodynamic Effect The film o water will distort the aerooil, roughen the surace and alter the airflow pattern on the whole aircraf. CLMAX will decrease causing stall speed to increase. Drag The film o water will increase increase intererence intererence drag, profile drag and orm drag. In light rain, drag may increase increase by 5%, in moderate moderate by 20% and in heavy rain by up to 30%. This obviously increases thrust required. Impact An additional consideration, while not affecting stall speed, is the effect o the impact impac t o heavy rain on the aircraf. aircraf. Momentum will be lost and airspeed will decrease, requiring increased increased thrust. At the same time, heavy rain will also be driving the aircraf downwards. The volume o rain in any given situation will vary, but an aircraf on final fi nal approach which suddenly enters a torrential downpour o heavy rain will be subject to a loss o momentum and a decrease in altitude, similar to the effect o microburst windshear. ( Chapter 15).

Stall and Recovery Characteristics of Canards With the conventional rear tailplane configuration the wing stalls beore the tailplane, and longitudinal control and stability are maintained at the stall. On a canard layout layout i the wing stalls first, stability is lost, but i the oreplane stalls first then control is lost and the maximum value o CL is reduced.

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Stalling Spinning When an aircraf is accidentally or deliberately stalled, the motion o the aircraf may in some cases develop into into a spin. The important characteristics o a spin are: are: a)

the aircraf is descending along a steep helical path about a vertical spin axis,

b)

the angle o attack o both wings is well above the stall angle,

c)

the aircraf has a high rate o rotation about the vertical spin axis,

d)

viewed rom above, the aircraf execute executess a circular path about the spin axis, and the radius o the helix is usually less than the semi-span o the wing,

e)

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the aircraf may be in the “erect” or “inverted” position in the spin.

The spin is one o the most complex complex o all flight manoeuvres. A spin may be defined as as an aggravated aggravat ed stall resulting in autorotation, which means the rotation is stable and will continue due to aerodynamic orces orces i nothing intervenes. During the spin, the wings remain unequally stalled.

Primary Causes of a Spin A stall must occur beore a spin can take take place. place. A spin occurs when one wing stalls more than than the other, other, Figure 7.32. The wing that is more more stalled will drop and the nose o the aircraf aircraf will yaw in the direction o the lower wing. The cause o an accidental spin is exceeding the critical angle o attack while perorming a manoeuvre with either too much or not enough rudder input or the amount o aileron being used (crossed-controls). (crossed-controls). I the correct stall recovery is not initiated initiated promptly, the stall could develop into a spin. Co-ordinated use o the flight controls is important, especially during flight at low airspeed and high angle o o attack. Although most pilots are able to maintain co-ordinated flight during routine manoeuvres, this ability ofen deteriorates when distractions occur and their attention is divided between important tasks. Distractions that have have caused problems include include preoccupation with situations inside or outside the flight deck, manoeuvring to avoid other aircraf and manoeuvring to clear obstacles during take-off, climb, approach or landing. A spin may also develop i orces on the aircraf are unbalanced in other ways, or example, rom yaw orces due to an engine ailure on a multi-engine aircraf, or i the CG is laterally displaced by an unbalanced uel load.

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Stalling

STALL

UPGOING SEMI - SPAN

CL

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CD

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DOWNGO ING SEMI - SPAN

ANGLE OF ATTACK

Figure 7.32

Phases of a Spin There are three phases o a spin. 1.

The incipient spin is the first phase, and exists rom the time the aeroplane stalls and rotation starts until the spin is ully developed.

2.

A ully developed spin exists rom the time the angular rotation rates, airspeed and vertical descent rate are stabilized stabilized rom one turn to the next.

3.

The third phase, spin recovery, begins when the anti-spin orces overcome the pro-spin orces.

I an aircraf is near the critical angle o attack, and more lif is lost rom one wing than the other,, that wing will drop. Its relative airflow will be inclined upwards, other upwards, increasing its effective angle o attack. As the aeroplane rolls around its CG, the rising wing wing has a reduced effective angle o attack and remains less less stalled than the other. other. This situation o unbalanced lif tends to increase as the aeroplane yaws towards the low wing, accelerating the high, high , outside wing and slowing the inner, inner, lower wing. As with any stall, the nose drops, and as inertia inertia orces begin to take effect, the spin usually stabilizes at a steady rate o rotation and descent. It is vitally important that recovery rom an unintentional spin is begun as soon as possible, since many aeroplanes will not easily recover rom a ully  ully developed spin, and others continue or several several turns beore recovery recovery inputs become effective. effective. Recovery rom an incipient spin normally requires less altitude and time than than the recovery rom a ully developed developed spin. Every aeroplane spins differently, and an individual aeroplane’s spin characteristics vary depending on configuration, loading and other actors.

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Stalling The Effect of Mass and Balance on Spins Both the total mass o the aircraf and its distribution influence the spin characteristics o the aeroplane. Higher masses generally mean slower slower initial spin rates, but as the spin progresses, spin rates may may tend to increase. increase. The higher angular momentum extends the time and altitude necessary or recovery rom 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 o the spin itsel. a)

CG towards the orward limit makes an aircraf more stable, and control orces 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 i the controls are released, but the stall speed will be higher.

b)

CG towards the af limit decreases longitudinal static stability and reduces pitch control orces, which tends to to make the aeroplane easier easier to stall. Once a spin is entered, entered, the urther af the CG, the flatter the spin attitude.

c)

I the CG is outside the af limit, or i power is not reduced promptly, the spin is more likely to go flat. A flat spin is characterized characterized by a near level pitch and roll attitude with the spin axis near the CG. Although the altitude lost in each turn o a flat spin may be less than in a normal spin, the extreme yaw rate (ofen exceeding 400° per second) results in a high descent descent rate. rate. The relative airflow in a flat spin is nearly straight straight up, keeping the wings wings at high angles o attack. attack. More importantly, the upward flow over the tail may render render the elevator elevator and rudder ineffective, making recovery recovery impossible.

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Spin Recovery Recovery rom a simple stall is achieved by reducing the angle o attack which restores the airflow over the wing; spin recovery additionally involves stopping the rotation. The extremely complex aerodynamics o a spin may dictate vastly different recovery procedures or different aeroplanes, so no universal spin recovery procedure can exist or all aeroplanes .

The recommended recovery procedure or some aeroplanes is simply to reduce power to idle and release pressure on the controls. At the other extreme, the design o some aircraf is such that recovery rom  rom a developed spin requires definite control movements, precisely timed to coincide with certain points po ints in the rotation, or several turns.

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Stalling The ollowing is a general recovery procedure or erect spins. Always reer to to the Flight Manual or the particular par ticular aircraf being flown and ollow the manuacturer’s recommendations. 1.

Move the throttle or throttles to idle. This minimizes altitude loss and reduces the possibility o a flat spin spin developing. It also eliminates eliminates possible asymmetric asymmetric thrust in multi-engine aeroplanes. aeroplanes. Engine torque and gyroscopic propeller effect can increase increase the angle o attack attack or the rate o o rotation in single-engine aeroplanes, aeroplanes, aggravating aggravating the spin.

2.

Neutralize the ailerons. Aileron position is ofen a contributory actor to flat spins, or Neutralize to higher rotation rates in normal spins.

3.

Apply ull rudder against the spin. Spin direction is most reliably determined rom the turn co-ordinator co-ordinator.. Do not use the the ball in the slip indicator; indicator; its indications are not reliable reliable and may be affected by its location within the flight deck.

4.

Move the elevat elevator or control briskly to approximately the neutral position. Some aircraf merely require a relaxation o back pressure, while others require ull orward pitch control travel.

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The above our items can be accomplished simultaneously. 5.

Hold the recommended control positions until rotation stops.

6.

As rotation stops, neutraliz neutralize e the rudder rudder.. I rudder deflection is maintained afer rotation stops, the aircraf may enter a spin in the other direction!

7.

Recover rom 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 actor and damage the aircraf structure.

b)

Recovering too slowly rom 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 or closing the throttles during spin recovery recovery c)

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Add power as you resume normal flight, being careul to observe power and RPM limitations.

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Stalling Special Phenomena of Stall Crossed-control Stall A crossed-control stall can occur when flying at high angles o 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 rom neutral.

Crossed-control stalls can occur with little or no warning; one wing will stall a long time beore the other and a quite violent wing drop can occur. The “instinctive” reaction to stop the wing drop with aileron must be resisted. The rudder should be used to keep the aircraf in balanced, co-ordinated flight at all times (ball in the middle), especially at low airspeeds/high angles o attack.

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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 rom a dive. It is called an “accelerated stall” because it occurs at a load actor greater than 1g. An accelerated stall is usually more violent than a 1g stall and is ofen unexpected because o the relatively high airspeed. Secondary Stall A secondary stall may be triggered while attempting to recover rom a stall. This usually happens as a result o trying to hasten the stall recovery: either by not decreasing the angle o attack enough at stall warning or by not allowing sufficient time or the aircraf to begin flying again beore attempting to regain lost altitude. With ull power still applied, relax the back pressure and allow the aeroplane to fly beore reapplying moderate back pressure to regain lost height. Large Aircraf During airline “type” conversion training on large aircraf, ull stalls are not practised. To amiliarize pilots with the characteristics o their aircraf, only the approach to stall (stick shaker activation) is carried out.

(a)

Jet Aircraf (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 minimize 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 (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 minimize 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 jet and propeller aircraf is the rapidly changing propeller torque and slipstream that will be evident during power application. It is essential or 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 o power. Yaw must be prevented during a stall and recovery.

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Small Aircraf (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 minimize height loss, prevent wing drop with rudder and raise the gear i applicable.

(ii)

Power-on stall and recovery in a single-engine propeller aircraf has additional complications. At the high nose attitude and low airspeed associated with a power-on stall, there will be considerable “turning effects” rom the propeller. (These are ully detailed in Chapter 16 ).

To maintain co-ordinated flight during the approach to, and recovery rom, a power-on stall, the pilot o a single-engine propeller aircraf must compensate or the turning effects o the propeller with the correct combination o rudder and aileron. It is essential to maintain coordinated flight (ball in the middle) when close to the stall AND during recovery. Any yawing tendency could easily develop into a spin. When the aircraf nose drops at the stall, gyroscopic effect will also be apparent, increasing the nose lef 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 turn into a spin. It is essential that correct stall recovery action is taken at the first indication o a stall. (Forward movement o the pitch control; neutralize the roll control; and prevent wing drop with the rudder).

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Stalling Stall and Recovery in a Climbing and Descending Turn When an aircraf is in a level co-ordinated turn at a constant bank angle, the inside wing is moving through the air more slowly than the outside wing and consequently generates less lif. I the ailerons are held neutral, the aircraf has a tendency to continue to roll in the direction o bank (over-banking tendency). Rather than return the ailerons to neutral when the required degree o bank angle is reached, the pilot must hold aileron opposite to the direction o bank; the lower the airspeed, the greater the aileron input required. The inner (lower) wing may have a greater effective angle o attack due to the lowered aileron and may reach the critical angle o attack first. The rudder must be used at all times to maintain co-ordinated flight (ball in the middle).

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In a climbing turn, airspeed will be lower and in a single-engine propeller aircraf, the rolling and yawing orces generated by the propeller and its slipstream will add their own requirements or unusual rudder and aileron inputs. E.g. or an aircraf with a clockwise rotating propeller in a climbing turn to the lef at low speed it may be necessary or the pilot to be holding a lot o right roll aileron and right rudder. I an aircraf in this situation were to stall, the gross control deflections could make the aircraf yaw or roll violently. Correct co-ordination o the controls is essential, in all phases o flight, to prevent the possibility o an accidental spin.

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Conclusions In whatever configuration, attitude or power setting a stall warning occurs, the correct pilot action is to decrease the angle o attack below the stall angle to un-stall the wing, apply maximum allowable power to minimize altitude loss and prevent any yaw rom developing to minimize the possibility o spinning (pretty much, in that order). “Keep the ball in the middle”.

High Speed Buffet (Shock Stall) When explaining the basic Principles o Flight, we consider air to be incompressible at speeds less than our tenths the speed o sound (M 0.4). That is, pressure is considered to have no effect on air density. At speeds higher than M 0.4 it is no longer practical to make that assumption because density changes in the airflow around the aircraf begin to make differences to the behaviour o the aircraf.

SHOCKWAVE

SEPARATED A IRFLOW

Figure 7.33

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Stalling At high altitude, a large high speed jet transport aircraf will be cruising at a speed marginally above its critical Mach number, and it will have a small shock wave on the wing. I such an aircraf overspeeds, the shock wave will rapidly grow larger, causing the static pressure to increase sharply in the immediate vicinity o the shock wave. The locally increased adverse pressure gradient will cause the boundary layer to separate immediately behind the shock wave, Figure 7.33. This is called a ‘shock stall’. The separated airflow will engul the tail area in a very active turbulent wake and cause severe airrame buffeting - a very undesirable phenomenon. High speed buffet (shock stall) can seriously damage the aircraf structure, so an artificial warning device is installed that will alert the pilot i the aircraf exceeds its maximum operational speed limit (VMO /MMO)* by even a small margin. The high speed warning is aural (“clacker”, horn or siren) and is easily distinguishable rom the “low speed” high angle o attack “stick shaker” warning.

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We have seen that approaching the critical angle o attack can cause airrame buffeting (“low speed” buffet) and we have now shown that flying too ast will also cause airrame buffeting (“high speed” buffet). ANY airrame 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 f), the margin between the high angle o attack stall warning and the high speed warning may be as little as 15 kt. *VMO is the maximum operating Indicated Airspeed, M MO is the maximum operating Mach number. (These will be ully discussed in Chapter 14). Note: It is operationally necessary to fly as ast as economically possible and designers are constantly trying to increase the maximum speed at which aircraf 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 (V MO /M MO ) gives a speed margin into which the aircraf can momentarily overspeed and be recovered by the pilot beore any undesirable characteristics occur. (Tuck, loss o control effectiveness and several stability problems - these will all be detailed in later chapters).

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Stalling Answers to Questions on Page 173 Stall speed in a 25° and 30° bank i V S1g = 150 kt CAS. (with % comparisons)

25° = 158 kt CAS (5% increase in stall speed above V S1g) [lif 10% greater] 30° = 161 kt CAS (7% increase in stall speed above V S1g) [lif 15% greater] 45° = 178 kt CAS (19% increase in stall speed above V S1g) [lif 41% greater] 7

60° = 212 kt CAS (41% increase in stall speed above V S1g) [lif 100% greater]

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Stalling Key Facts 2 Sel Study Insert the missing words in these statements, using the oregoing paragraphs or reerence. The swept-back wing is the major contributory actor to _______ stall. An aircraf design with super stall tendencies must be fitted with a stick ______. Factors which can affect V SR are:

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a.

Changes in _______.

b.

Manoeuvring the aircraf (increasing the ______ _______ ).

c.

Configuration changes (changes in _______ and _________ moment).

d.

Engine _______ and propeller ______________.

e.

_______ number.

.

Wing _______________.

g.

Heavy ________.

In straight and level flight the load actor is _____. At a higher weight, the stall speed o an aircraf will be _________. I the weight is decreased by 50%, the stall speed will ___________ by approximately ____%. Load actor varies with ______ ______. The increase in stall speed in a turn is proportional to the square root o the ______ _______. High lif devices will __________ the stall speed because C LMAX is __________. Forward CG movement will __________ stall speed due to the increased tail ________ load. Lowering the landing gear will increase stall speed due to the increased tail ________ load. Increased engine power will decrease stall speed due to propeller _________ and/or the __________ inclination o thrust. The effect o increasing Mach number on stall speed begin at M ______. The effects o compressibility increases stall speed by decreasing _________.

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Stalling The ormation o ice on the leading edge o the wing can ___________ stall speed by ____ %. Frost ormation on the wing can __________ stall speed by ____ %. An aircraf must be ree o all _____, _____ and ____ immediately beore ______. Airrame contamination ________ stall speed by reducing ______, increasing the adverse _________ _________ and/or reducing the _______ energy o the boundary layer. Indications o an icing-induced stall can be loss o aircraf ___________, _____ oscillations or _____ drop and high rate o _______. Artificial stall warning will be _______, but aerodynamic _______ may assist in identiying the onset o wing stall.

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Very heavy ____ can ________ the stall speed due to the film o water altering the ___________ contour o the wing. A _______ must occur beore a spin can take place. In a steady spin, _____ wings are stalled, one more than the other. A spin may also develop i orces on the aircraf are unbalanced in other ways, or example, rom yaw orces due to an ______ ailure on a multi-engine aircraf, or i the ___ is laterally displaced by an unbalanced _____ load. The ollowing is a general recovery procedure or erect spins: 1.

Move the throttle or throttles to _____.

2.

___________ the ailerons.

3.

Apply ull _______ against the spin.

4.

Move the _________ control briskly to approximately the neutral position.

5.

______ the recommended control positions until rotation stops.

6.

As rotation stops, neutralize the _______.

7.

Recover rom the resulting dive with ________ back pressure on the ______ control.

A crossed-control stall can be avoided by maintaining the ___ o the slip indicator in the ______. A stall can occur at any ______ or flight _________ i the ________ angle o attack is exceeded. A secondary stall can be triggered either by not ___________ the angle o ______ enough at stall warning or by not allowing sufficient ___ or the aircraf to begin _____ again beore attempting to ________ lost altitude.

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Stalling An added complication during an accidental stall and recovery o a single engine-propeller aircraf is due to the _______ and ______ orces generated by the _________. It is essential to maintain balanced, co-ordinated flight, particularly at ____ airspeed, high angles o _______. In whatever configuration, attitude, or power setting a stall warning occurs, the correct pilot action is to ________ the angle o attack below the _____ angle to un-stall the ____, apply maximum allowable ______ to minimize altitude loss and prevent any ____ rom developing to minimize the possibility o ________. “Keep the _____ in the middle”. I a large shock wave orms 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”.

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KEY FACTS 2, WITH WORD INSERTS CAN BE FOUND ON page 204.

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An aeroplane will stall at the same:

a. b. c. d. 2.

A typical stalling angle o attack or a wing without sweepback is:

a. b. c. d. 3.

b. c. d.

remain the same. decrease. increase. the position o the CG does not affect the stall speed.

lif and drag will both decrease. lif will decrease and drag will increase. lif will increase and drag will decrease. lif and drag will both increase.

will occur at smaller angles o attack flying downwind than when flying upwind. is dependent upon the speed o the airflow over the wing. is a unction o speed and density altitude. will remain constant regardless o gross weight.

An aircraf whose weight is 237 402 N stalls at 132 kt. At a weight o 356 103 N it would stall at:

a. b. c. d. 7.

s n o i t s e u Q

The angle o attack at which an aeroplane stalls:

a.

6.

7

I the angle o attack is increased above the stalling angle:

a. b. c. d. 5.

4°. 16°. 30°. 45°.

I the aircraf weight is increased without change o C o G position, the stalling angle o attack will:

a. b. c. d. 4.

angle o attack and attitude with relation to the horizon. airspeed regardless o the attitude with relation to the horizon. angle o attack regardless o the attitude with relation to the horizon. indicated airspeed regardless o altitude, bank angle and load actor.

88 kt. 162 kt. 108 kt. 172 kt.

For an aircraf with a 1g stalling speed o 60 kt IAS, the stalling speed in a steady 60° turn would be:

a. b. c. d.

43 kt. 60 kt. 84 kt. 120 kt.

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Questions 8.

For an aircraf in a steady turn the stalling speed would be:

a. b. c. d. 9.

Formation o ice on the wing leading edge will:

a. b. c. d.

7

Q u e s t i o n s

10.

simultaneously across the whole span. at the centre o the span. at the root. at the tip.

Sweepback on a wing will:

a. b. c. d.

196

cause the wing to stall first at the root. cause the wing to stall at the tip first. delay wing root stall. re-energize the boundary layer at the wing root.

On a highly tapered wing without wing twist the stall will commence:

a. b. c. d. 15.

increasing the adverse pressure gradient. increasing the surace roughness o the wing top surace. distortion o the leading edge by ice build-up. increasing the kinetic energy o the boundary layer.

A stall inducer strip will:

a. b. c. d. 14.

changes in air density. variations in aeroplane loading. variations in flight altitude. changes in pitch attitude.

Stalling may be delayed to a higher angle o attack by:

a. b. c. d. 13.

wing loading. lif/drag ratio. aspect ratio. load actor.

The stalling speed o an aeroplane is most affected by:

a. b. c. d. 12.

not affect the stalling speed. cause the aircraf to stall at a higher speed and a higher angle o attack. cause the aircraf to stall at a higher speed and a lower angle o attack. cause the aircraf to stall at a lower speed.

Dividing lif 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 o attack. at a higher speed than in level flight and at the same angle o attack.

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 o attack.

7

Questions 16.

The purpose o a boundary layer ence on a swept wing is:

a. b. c. d. 17.

A wing with washout would have:

a. b. c. d. 18.

c. d.

increases rom root to tip. increases rom tip to root. is constant across the span. is greatest at centre span, less at root and tip.

will stall at the tip first due to the increase in spanwise flow. could drop a wing at the stall due to the lack o any particular stall inducing characteristics. will pitch nose down approaching the stall due to the orward movement o the centre o pressure. will stall evenly across the span.

increasing leading edge camber. delaying separation. reducing the effective angle o attack. reducing spanwise flow.

A rectangular wing, when compared to other wing planorms, has a tendency to stall first at the:

a. b. c. d. 22.

s n o i t s e u Q

Slots increase the stalling angle o attack by:

a. b. c. d. 21.

7

A wing o constant thickness which is not swept-back:

a. b.

20.

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. 19.

to re-energize the boundary layer and prevent separation. to control spanwise flow and delay tip stall. to generate a vortex over the upper surace o the wing. to maintain a laminar boundary layer.

wing root providing adequate stall warning. wing tip providing inadequate stall warning. wing tip providing adequate stall warning. leading edge, where the wing root joins the uselage.

Vortex generators are used:

a. b. c. d.

to reduce induced drag. to reduce boundary layer separation. to induce a root stall. to counteract the effect o the wing tip vortices.

197

7

Questions 23.

A stick shaker is:

a. b. c. d. 24.

A stall warning device must be set to operate:

a. b. c. d.

7

Q u e s t i o n s

25.

the angle o attack o both wings will be positive. the angle o attack o both wings will be negative. the angle o attack o 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 rom a spin, the elevators should be:

a. b. c. d.

198

a device to prevent an aircraf rom stalling. a type o trim system. a device to assist the pilot to move the controls at high speed. a device which automatically compensates or pitch changes at high speed.

In a developed spin:

a. b. c. d. 30.

angle o attack only. angle o attack, and in some systems rate o change o angle o attack. airspeed only. airspeed and sometimes rate o change o airspeed.

A stick pusher is:

a. b. c. d. 29.

on the upper surace at about mid chord. on the lower surace at about mid chord. at the leading edge on the lower surace. at the leading edge on the upper surace.

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 o the stall warning vane. on top o the leading edge because o the extremely high angle o attack.

A wing mounted stall warning detector vane would be situated:

a. b. c. d. 27.

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.

Just beore the stall the wing leading edge stagnation point is positioned:

a. b. c. d. 26.

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.

moved up to increase the angle o attack. moved down to reduce the angle o attack. set to neutral. allowed to float.

7

Questions 31.

High speed buffet (shock stall) is caused by:

a. b. c. d. 32.

In a 30° bank level turn, the stall speed will be increased by:

a. b. c. d. 33.

d.

b. c. d.

36.

7

s n o i t s e u Q

Water increases the viscosity o air. Heavy rain can block the pitot tube, giving alse airspeed indications. The extra weight and distortion o the aerodynamic suraces by the film o water. The impact o heavy rain will slow the aircraf.

I the tailplane is supplying a down load and stalls due to contamination by ice:

a.

35.

7%. 30%. 1.07%. 15%.

Heavy rain can increase the stall speed o an aircraf or which o the ollowing reasons?

a. b. c.

34.

the boundary layer separating in ront o a shock wave at high angles o attack. the boundary layer separating immediately behind the shock wave. the shock wave striking the tail o the aircraf. the shock wave striking the uselage.

the wing will stall and the aircraf will pitch-up due to the weight o the ice behind the aircraf CG. the increased weight on the tailplane due to the ice ormation will pitch the aircraf nose up, which will stall the wing. because it was supplying a down load the aircraf will pitch nose up. the aircraf will pitch nose down.

Indications o 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. airrame buffet. violent wing drop. extremely high rate o descent while in a ‘normal’ flight attitude.

a. b. c. d.

1, 2, 4 and 5. 1, 3 and 5. 1, 4 and 6. 3, 4, 5 and 6.

I a light single-engine propeller aircraf is stalled, power-on, in a climbing turn to the lef, which o the ollowing is the preerred recovery action?

a. b. c. d.

Elevator stick orward, ailerons stick neutral, rudder to prevent wing drop. Elevator stick neutral, rudder neutral, ailerons to prevent wing drop, power to idle. Elevator stick orward, ailerons and rudder to prevent wing drop. Elevator stick neutral, rudder neutral, ailerons stick neutral, power to idle.

199

7

Questions 37.

I the stick shaker activates on a swept wing jet transport aircraf immediately afer take-off while turning, which o the ollowing statements contains the preerred course o action?

a. b. c. d.

7

Q u e s t i o n s

200

Decrease the angle o attack. Increase thrust. Monitor the instruments to ensure it is not a spurious warning. Decrease the bank angle.

7

Answers Key Facts 1 (Completed) Correct Statements Stalling involves loss o height and loss o control. A pilot must be able to clearly and unmistakably identiy 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.

7

s r e w s n A

Adverse pressure gradient increases with increase in angle o attack. Alternative names or the angle o attack at which stall occurs are the stall angle and the critical angle o attack. The coefficient o lif at which a stall occurs is CLMAX. A stall can occur at any airspeed or flight attitude. A typical stalling angle is approximately 16°. To recover rom a stall the angle o attack must be decreased. Maximum power is applied during stall recovery to minimize height loss. On small aircraf, the rudder should be used to prevent wing drop at the stall. On swept wing aircraf 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 o 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 lif a dropping wing. Acceptable indications o a stall are: (1)

a nose-down pitch that can not be readily arrested.

(2)

severe buffeting.

(3)

pitch control reaching af stop and no urther increase in pitch attitude occurs.

Reerence stall speed (V SR ) is a CAS defined by the aircraf manuacturer. VSR may not be less than a 1g stall speed. When a device that abruptly pushes the nose down at a selected angle o attack is installed, VSR may not be less than 2 knots or 2 %, whichever is greater, above the speed at which the device operates.

201

7

Answers

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 o the inherent aerodynamic qualities o the aeroplane or by a device that will give clearly distinguishable indications under expected conditions o 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 aircraf is usually given by a horn or buzzer. 7

Artificial stall warning on a large aircraf is usually given by a stick shaker, in conjunction with lights and a noisemaker.

A n s w e r s

An artificial stall warning device can be activated by a flapper switch, an angle o attack vane or an angle o attack probe. Most angle o attack sensors compute the rate o change o angle o attack to give earlier warning in the case o accelerated rates o stall approach. EASA required stall characteristics, up to the time the aeroplane is stalled, are: a.

It must be possible to produce and correct yaw by unreversed use o the ailerons and rudder.

b.

No abnormal nose-up pitching may occur.

c.

Longitudinal control orce must be positive.

d.

It must be possible to promptly prevent stalling and recover rom a stall by normal use o the controls.

e.

There should be no excessive roll between the stall and completion o recovery.

.

For turning flight stalls, the action o the aeroplane afer 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 o the aeroplane.

An aerooil section with a small leading edge radius will stall at a smaller angle o attack and the stall will be more sudden. An aerooil section with a large thickness-chord ratio will stall at a higher angle o attack and will stall more gently. An aerooil section with camber near the leading edge will stall at a higher angle o attack. A rectangular wing planorm will tend to stall at the root first. A rectangular wing planorm usually has ideal stall characteristics; these are:

202

a.

aileron effectiveness at the stall.

b.

nose drop at the stall.

c.

aerodynamic buffet at the stall.

d.

absence o violent wing drop at the stall.

7

Answers To give a wing with a tapered planorm the desired stall characteristics, the ollowing devices can be included in the design: a.

washout (decreasing incidence rom root to tip).

b.

an aerooil section with greater thickness and camber at the tip.

c.

leading edge slots at the tip.

d.

stall strips fitted to the wing inboard leading edge.

e.

vortex generators which re-energize the boundary layer at the tip.

7

s r e w s n A

A swept-back wing has an increased tendency to tip stall due to the spanwise flow o boundary layer rom root to tip on the wing top surace. Methods o delaying tip stall on a swept wing planorm are: a.

wing ences, thin metal ences which generally extend rom the leading edge to the trailing edge on the wing top surace.

b.

vortilons, also thin metal ences, but smaller and are situated on the underside o the

wing leading edge. c.

saw tooth leading edge, generates vortices over wing top surace at high angles o attack.

d.

engine pylons o pod mounted wing engines also act as vortilons.

e.

vortex generators are also used to delay tip stall on a swept wing.

Tip stall on a swept wing planorm gives a tendency or the aircraf to pitch-up at the stall. This is due to the CP moving orwards when the wing tips stall first.

203

7

Answers

Key Facts 2 (Completed) Correct Statements The swept-back wing is the major contributory actor to super stall. An aircraf design with super stall tendencies must be fitted with a stick pusher. Factors which can affect V SR are:

7

A n s w e r s

a.

changes in weight.

b.

manoeuvring the aircraf (increasing the load actor).

c.

configuration changes (changes in CLMAX and pitching moment).

d.

engine thrust and propeller slipstream.

e.

Mach number.

.

wing contamination.

g.

heavy rain.

In straight and level flight the load actor is one. At a higher weight, the stall speed o an aircraf will be higher. I the weight is decreased by 50%, the stall speed will decrease by approximately 25%. Load actor varies with bank angle. The increase in stall speed in a turn is proportional to the square root o the load actor. High lif devices will decrease the stall speed because CLMAX 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 o thrust. The effect o increasing Mach number on stall speed begin at M 0.4. The effects o compressibility increases stall speed by decreasing CLMAX. The ormation o ice on the leading edge o the wing can increase stall speed by 30%. Frost ormation on the wing can increase stall speed by 15%. An aircraf must be ree o all snow, rost and ice immediately beore flight. Airrame contamination increases stall speed by reducing CLMAX, increasing the adverse pressure gradient and/or reducing the kinetic energy o the boundary layer.

204

7

Answers Indications o an icing-induced stall can be loss o aircraf perormance, roll oscillations or wing drop and high rate o descent. Artificial stall warning will be absent, but aerodynamic buffet may assist in identiying the onset o wing stall. Very heavy rain can increase the stall speed due to the film o water altering the aerodynamic contour o the wing. A stall must occur beore a spin can take place. In a steady spin, both wings are stalled, one more than the other. A spin may also develop i orces on the aircraf are unbalanced in other ways, or example, rom yaw orces due to an engine ailure on a multi-engine aircraf, or i the CG is laterally displaced by an unbalanced uel load.

7

s r e w s n A

The ollowing is a general recovery procedure or erect spins: 1.

move the throttle or throttles to idle.

2.

neutralize the ailerons.

3.

apply ull rudder against the spin.

4.

move the elevator control briskly to approximately the neutral position.

5.

hold the recommended control positions until rotation stops.

6.

as rotation stops, neutralize the rudder.

7.

recover rom the resulting dive with gradual back pressure on the pitch control.

A crossed-control stall can be avoided by maintaining the ball o the slip indicator in the middle. A stall can occur at any speed or flight attitude i the critical angle o attack is exceeded. A secondary stall can be triggered either by not decreasing the angle o attack enough at stall warning or by not allowing sufficient time or the aircraf to begin flying again beore attempting to regain lost altitude. An added complication during an accidental stall and recovery o a single-engine propeller aircraf is due to the rolling and yawing orces generated by the propeller. It is essential to maintain balanced, co-ordinated flight, particularly at low airspeed, high angles o attack. In whatever configuration, attitude, or power setting a stall warning occurs, the correct pilot action is to decrease the angle o attack below the stall angle to un-stall the wing, apply maximum allowable power to minimize altitude loss and prevent any yaw rom developing to minimize the possibility o spinning. “Keep the ball in the middle”. I a large shock wave orms 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’.

205

7

Answers

Answers

7

A n s w e r s

1 c

2 b

3 a

4 b

5 d

6 b

7 c

8 d

9 c

10 d

11 b

12 d

13 a

14 d

15 b

16 b

17 b

18 a

19 b

20 b

21 a

22 b

23 c

24 c

25 b

26 c

27 b

28 a

29 a

30 b

31 b

32 a

33 c

34 d

35 d

36 a

37 a

206

Chapter

8 High Lift Devices

Purpose o High Lif Devices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 209 Take-off and Landing Speeds. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 209 CLMAX Augmentation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 209 Flaps . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

209

Trailing Edge Flaps . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 209 Plain Flap . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 209 Split Flap . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 210 Slotted and Multiple Slotted Flaps . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 210 The Fowler Flap . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 211 Comparison o Trailing Edge Flaps . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 211 CLMAX and Stalling Angle . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 212 Drag . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

213

Lif / Drag Ratio . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 214 Pitching Moment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 214 Centre o Pressure Movement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 214 Change o Downwash . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 214 Overall Pitch Change . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 215 Aircraf Attitude with Flaps Lowered . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 215 Leading Edge High Lif Devices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 216 Leading Edge Flaps . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 216 Effect o Leading Edge Flaps on Lif . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 217 Leading Edge Slots . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 218 Leading Edge Slat . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 218 Automatic Slots . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 220 Disadvantages o the Slot . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 220 Drag and Pitching Moment o Leading Edge Devices . . . . . . . . . . . . . . . . . . . . . . . 220 Trailing Edge Plus Leading Edge Devices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 220 Continued Overlea

207

8

High Lift Devices Sequence o Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 221 Asymmetry o High Lif Devices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 221 Flap Load Relie System. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 222 Choice o Flap Setting or Take-off, Climb and Landing . . . . . . . . . . . . . . . . . . . .

222

Management o High Lif Devices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 224 Flap Extension Prior to Landing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 226

8

H i g h L i f t D e v i c e s

208

Questions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

227

Annexes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

229

Answers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

230

8

High Lift Devices Purpose of High Lift Devices Aircraf are fitted with high lif devices to reduce the take-off and landing distances . This

permits operation at greater weights rom given runway lengths and enables greater payloads to be carried.

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 perormance regulations. For both take-off and landing, one o the requirements is or a sae margin above the stalling speed (1.2V S1 or take-off and 1.3V S0 or landing). The stalling speed is determined by the C LMAX o the wing, and so to obtain the lowest possible distances, the C LMAX , must be as high as possible.

8

s e c i v e D t f i L h g i H

CLMAX Augmentation One o the main actors which determine the C LMAX o an aerooil section is the camber. It has been shown earlier that increasing the camber o an aerooil section increases the C L at a given angle o 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 aerooil section to optimise cruise and modiy the section or take-off and landing by the use o flaps.

Flaps A flap is a hinged portion o the trailing or leading edge which can be deflected downwards and so produce an increase o camber. For low speed aerooils the flaps will be on the trailing edge only, but on high speed aerooils 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.

Trailing Edge Flaps The basic principle o the flap has been adapted in many ways. The more commonly used types o trailing edge flap are considered below.

Plain Flap The plain flap, illustrated in Figure 8.1, has a simple construction and gives a good increase in CLMAX, although with airly high drag. It is used mainly on low speed aircraf and where very short take-off and landing is not required.

Figure 8.1 Plain flap

209

8

High Lift Devices Split Flap The flap orms part o the lower surace o the wing trailing edge, the upper surace contour being unaffected when the flap is lowered.

8


Figure 8.2 Split flap

The split flap gives about the same increase in lif as the plain flap at low angles o attack but gives slightly more at higher angles as the upper surace camber is not increased, and so separation is delayed. The drag, however, is higher than or the plain flap due to the increased depth o the wake.

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 o the slot is to direct higher pressure air rom the lower surace over the flap and re-energize the boundary layer. This delays the separation o the airflow on the upper sur ace o the flap. The slotted flap gives a bigger increase in CLMAX than the plain or split flap and much less drag, but it has a more complex construction.

210

8

High Lift Devices The Fowler Flap The Fowler flap, Figure 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 8


Triple Slotted Fowler Flap

Figure 8.4

Because o the combined effects o increased area and camber, the Fowler flap gives the greatest increase in lif o the flaps considered and also gives the least drag because o the slot and the reduction o thickness : chord ratio. However, the change o pitching moment is greater because o the rearward extension o the chord.

Comparison of Trailing Edge Flaps Figure 8.5 shows a comparison o the lif FOW LER FLAP

CL

SLOTTED FLAP

SPLIT FLAP

curves or the flaps considered above, or the same angle o flap deflection. It should be noted, however, that the different types o flap do not all give their greatest increase in lif at the same flap angle.

PLAIN FLAP

BASIC SECTION

Figure 8.5

211

8

High Lift Devices Figure 8.6 shows the variation o the lif increment with flap angle and the variation o drag

increment with flap angle. It can be seen that the increment in lif is decreasing and the increment in drag is increasing as flaps are deployed. It is important to note that any amount o flap increases drag.

CL

5º

CD

TO 10º

8


0º

TO 5º

5º

0º 0º

5º

TO 10º

TO 5º

10º

0º

FLAP ANGLE

5º

10º

FLAP ANGLE

Figure 8.6

CLMAX and Stalling Angle It can be seen rom Figure 8.5 that with the flap lowered C LMAX is increased, but the stalling angle is reduced. This is because lowering the flap increases the effective angle o attack.

REDUCED ANGLE OF ATTACK OF BASIC SECTION EFFECTIVE ANGLE OF ATTACK

Figure 8.7

It is conventional to plot the C L ~ α curve using the angle o attack or the basic section. Consequently, as shown in Figure 8.7 , at the stalling angle o attack or the section with flap lowered, the basic wing section is at a reduced angle.

212

8

High Lift Devices Drag Figure 8.8 shows a comparison o the drag polar curves or the various types o flap. It can

be seen that or a given flap deflection the drag produced by the different types o flap varies considerably, the split flap giving the highest drag and the Fowler flap the least.

FOW LER

CL

SLOTTED 8


SPLIT PLA IN

BASIC SECT ION

CD 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 orce and so the flap drag is beneficial. The addition o drag during approach also improves speed stability. As in the case o the lif increments, the drag increments with increasing flap angle are not constant: the increments in drag get larger as the flap angle increases.

213

8

High Lift Devices Lift / Drag Ratio Lowering flap increases both the lif and the drag, but not in the same proportion. Although the lif is the larger orce, the proportional increase in the drag is greater, and so the maximum obtainable lif / drag ratio decreases. The maximum lif / drag ratio occurs where the tangent rom the origin o the drag polar touches the curve, and the gradient o the tangent line is a measure o the maximum lif / drag ratio ( Figure 8.9).

CL 8


L D

RATIO

CD Figure 8.9 L/D ratio

The lif / drag ratio is a measure o aerodynamic efficiency and affects the aircraf’s perormance in areas such as range, climb angle and glide angle. With flaps lowered, range will be decreased, climb angle reduced and glide angle increased.

Pitching Moment Flap movement, up or down, will usually cause a change o pitching moment. This is due to Centre o Pressure (CP) movement and downwash at the tailplane.

Centre of Pressure Movement Moving a trailing edge flap will modiy the pressure distribution over the whole chord o the aerooil, but the greatest changes will occur in the region o the flap. When flap is lowered, the Centre o Pressure will move rearwards giving a nose-down pitching moment, Figure 8.10a. In the case o a Fowler flap, rearward movement o the flap will also cause the CP to move af, resulting in an even greater increase in the nose-down pitching moment.

Change of Downwash Tailplane effective angle o attack is determined by the downwash rom the wing. I the flaps are lowered, the downwash will increase and the tailplane angle o attack will decrease, causing a nose-up pitching moment, Figure 8.10b.

214

8

High Lift Devices

W ING

TAILPLANE

DOW NWASH CP

8


INCREASED DOW NWASH CP

NOSE-DOW N

PITCHING MOMENT

NOSE-UP

PITCHING MOMENT

Figure 8.10 (a)

(b)

Overall Pitch Change The resultant aircraf pitching moment will depend upon which o the two effects is dominant. The pitching moment will be influenced by the type o flap, the position o the wing and the relative position o the tailplane, and may be nose-up, nose-down or almost zero. For example, on flap extension, a tailplane mounted on top o the fin will be less influenced by the change o downwash, resulting in an increased aircraf nose-down pitching moment.

Aircraft Attitude with Flaps Lowered When the aircraf is in steady flight the lif must be equal to the weight. I the flaps are lowered but the speed kept constant, lif will increase, and to maintain it at its original value, the angle o attack must be decreased. The aircraf will thereore fly in a more nose-down attitude i the flaps are down. On the approach to landing this is an advantage as it gives better visibility o the landing area.

215

8

High Lift Devices Leading Edge High Lift Devices There are two orms o leading edge high lif device commonly in use: the leading edge flap and the leading edge slot or slat.

Leading Edge Flaps On high speed aerooil sections the leading edge may have very little camber and have a small radius. This can give flow separation just af o the leading edge at quite low angles o attack. This can be remedied by utilizing a leading edge flap, which increases the leading edge camber.

8


Figure 8.11 Krueger flap

Krueger Flap The Krueger flap is part o the lower surace o the leading edge, which can be rotated about its orward edge as shown in Figure 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.

216

8

High Lift Devices

RETRACTED

EXTENDED 8


Figure 8.12 Variable camber leading edge flap

Variable Camber Leading Edge Flap To improve efficiency by giving a better leading edge profile, the camber o a leading edge flap may be increased as it is deployed. Unlike trailing edge flaps, which can be selected to intermediate positions, leading edge flaps are usually either ully extended (deployed) or retracted (stowed).

Effect of Leading Edge Flaps on Lift The main effect o the leading edge flap is to delay separation and so increase the stalling angle and the corresponding C LMAX. However, there will be some increase o lif at lower angles o attack due to the increased camber o the aerooil section. Figure 8.13 shows the effect o these flaps on the lif curve.

CL

W ITH LEAD ING EDGE FLAP

BASIC W ING SECTION

Figure 8.13

217

8

High Lift Devices Leading Edge Slots A leading edge slot is a gap rom the lower surace to the upper surace o the leading edge, and it may be fixed, or created by moving part o the leading edge (the slat) orwards.

C LMAX W ING

8


(Given Adverse Pressure Gradient)

SLAT

SLAT OPEN - Boundary Layer Re - energized (Same Adverse Pressure Gradient)

Figure 8.14 Leading edge slat

Leading Edge Slat A slat is a small auxiliary aerooil attached to the leading edge o the wing, Figure 8.14. When deployed, the slat orms a slot which allows passage o air rom 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 orming a convergent duct. When slats are deployed, the boundary layer is re-energized

I Kinetic Energy is added to the boundary layer, boundary layer separation will be delayed to a much higher angle o attack. At approximately 25°, the increased adverse pressure gradient will once again overwhelm the Kinetic Energy o the boundary layer and separation will occur. I 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 or high speed flight and opened or low speeds, usually in conjunction with the trailing edge flaps and actuated by the same selector on the flight deck. The graph at Figure 8.15 shows the comparative figures or a slatted and un-slatted wing o the same basic dimensions.

218

8

High Lift Devices

WING PLUS SLATS

2.0

CL 1.5

C LMAX

1.0 WING 0.5

5

10

15

20 25

8

30


Angle of Attack

Figure 8.15

The effect o the slat is to prolong the lif curve by delaying boundary layer separation until a higher angle o attack. When operating at high angles o attack, the slat itsel is generating a high lif coefficient because o its marked camber. The action o the slat is to flatten the marked peak o the low pressure envelope at high angles o attack and to change it to one with a more gradual pressure gradient. The flattening o the lif 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 ormer ‘suction’ peak, and so it retains much o its Kinetic Energy, thus enabling it to penetrate almost the ull chord o the wing beore separating. Figure 8.16 shows the alleviating effect o the slat on the low pressure peak and that, although flatter, the area o the low pressure region, which is proportional to its strength, is unchanged or even increased. The ‘suction’ peak does not move orward, so the effect o the slot on pitching moment is insignificant.

No Slat

W ith Slat

Figure 8.16

219

8

High Lift Devices Automatic Slots On some aircraf the slots are not controlled by the pilot, but operate automatically. Their movement is caused by the changes o pressure which occur around the leading edge as the angle o attack increases. At low angles o attack the high pressures around the stagnation point keep the slat in the closed position. At high angles o attack the stagnation point has moved underneath the leading edge and ‘suction’ pressures occur on the upper surace o the slat. These pressures cause the slat to move orward and create the slot. This system is used mainly on small aircraf as a stall protection system. On larger aircraf, the position o the slats is selected when required by the pilot, their movement being controlled electrically or hydraulically.

8

Disadvantages of the Slot


The slot can give increases in C LMAX o 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 o attack to give its C LMAX. In flight this means that the aircraf will have a very nose-up attitude at low speeds, and on the approach to land, visibility o the landing area could be restricted.

Drag and Pitching Moment of Leading Edge Devices Compared to trailing edge flaps, the changes o drag and p itching moment resulting rom the operation o leading edge devices are small.

Trailing Edge Plus Leading Edge Devices Most large transport aircraf employ both trailing edge and leading edge devices. Figure 8.17 shows the effect on the lif curve o both types o device.

220

8

High Lift Devices

FLAP + SLOT

CL

BASIC + SLOT

TRAILING EDGE FLAP

BASIC SECTION 8


Figure 8.17

Sequence of Operation For some aerooils the sequence o flap operation is critical. Lowering a trailing edge flap increases both the downwash and the upwash. For a high speed aerooil, an increase o upwash at the leading edge when the angle o attack is already airly high could cause the wing to stall. The leading edge device must thereore be deployed beore the trailing edge flap is lowered. When the flaps are retracted, the trailing edge flap must be retracted beore the leading edge device is raised.

Asymmetry of High Lift Devices Deployment o high lif devices can produce large changes o lif, drag and pitching moment. I the movement o the devices is not symmetrical on the two wings, the unbalanced orces 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 i one side should ail, movement on the other side is automatically stopped. However, on less sophisticated systems a ailure o the operating mechanism could lead to an asymmetric situation. The difference in lif 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 o asymmetry and the control power available.

221

8

High Lift Devices Flap Load Relief System On large high speed jet transport aircraf, a device is fitted in the flap operating system to prevent the flaps deploying i the aircraf speed is too high. The pilot can select the flaps, but they will not extend until the airspeed is below the flap extend speed (V FE). I 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.

Choice of Flap Setting for Take-off, Climb and Landing Take-off Take-off distance depends upon unstick speed and rate o acceleration to that speed. 8


a)

Lowest unstick speed will be possible at the highest C LMAX and this will be achieved at a large flap angle, Figure 8.18.

b)

But large flap angles also give high drag, Figure 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 unstick 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. I leading edge devices are fitted, they will be used or take-off as they increase the C LMAX or any trailing edge flap setting.

Climb Afer take-off, a minimum climb gradient is required in the take-off configuration. Climb gradient is reduced by flap, so i climb gradient is limiting, a lesser flap angle may be selected even though it gives a longer take-off distance.

Landing Landing distance will depend on touchdown speed and deceleration. The lowest touchdown speed will be given by the highest C LMAX, obtained at a large flap angle, Figure 8.18. Large flap angle will also give high drag, Figure 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 CLMAX.

222

8

High Lift Devices

CL 30° 20°

FLAPS UP

8


LANDING

TAKE-OFF

ANGLE OF ATTACK

Figure 8.18

30° 20°

CD FLAPS UP

LANDING

TAKE-OFF

ANGLE OF ATTACK

Figure 8.19

223

8

High Lift Devices Management of High Lift Devices To take ull advantage o the capabilities o flaps, the flight crew must properly manage their retraction and extension.

Flap Retraction after Take-off With reerence to Figure 8.20, assume the aircraf has just taken off with flaps extended and is at point ‘A’ on the lif curve. I the flaps are retracted, with no change made to either angle o attack or IAS, the coefficient o lif will reduce to point ‘C’ and the aircraf will sink.

8


1.

From point ‘A’ on the lif curve the aircraf should be accelerated to point ‘B’.

2.

From point ‘B’, as the flaps are retracted the angle o attack should be increased to point ‘C’ to maintain the coefficient o lif constant.

The pilot should not retract the flaps until the aircraf has sufficient IAS. O course, this same actor must be considered or any intermediate flap position between extended and retracted. (Reer to Page 76 or a review o the Interpretation o the Lif Curve i necessary.) As the configuration is altered rom the flaps down to the flaps up or “clean” configuration, three important changes take place: • Changes o pressure distribution on the wing generate a nose-up pitching moment. But reduced wing downwash increasing the tailplane effective angle o attack generates a nosedown pitching moment. The resultant, actual, pitching moment experienced by the aircraf will depend upon which o these two pitching moments is dominant. • With reerence to Figure 8.21, the retraction o flaps (‘B’ to ‘C’) causes a reduction o drag coefficient. This drag reduction improves the acceleration o the aircraf. • Flap retraction usually takes place in stages, and movement o the flaps between stages will take a finite period o time. It has been stated that as flaps are retracted, an increase in angle o attack is required to maintain the same lif coefficient. I aircraf acceleration is low throughout the flap retraction speed range, the angle o attack must be increased an appreciable amount to prevent the aircraf rom sinking. This situation is typical afer take-off when gross weight and density altitude are high. However, most modern jet transport aircraf have enough acceleration throughout the flap retraction speed range that the resultant rapid gain in airspeed requires a much less noticeable increase in angle o attack.

224

8

High Lift Devices

CL

A

FLAPS EXTENDED

FLA PS RETRACTED

B C

8


ANGLE OF ATTACK

Figure 8.20

CL FLAPS EXTENDED

FLAPS RETRACTED

C B

CD

Figure 8.21

225

8

High Lift Devices Flap Extension Prior to Landing With reerence to Figure 8.22, assume the aircraf is in level flight in the terminal area prior to landing and is at point ‘A’ on the lif curve. I the flaps are extended, with no change made to angle o attack, the coefficient o lif will increase to point ‘C’ and the aircraf will gain altitude (balloon). 1.

From point ‘A’, as the flaps are extended the angle o attack should be decreased to point ‘B’ to maintain the coefficient o lif constant.

2.

From point ‘B’ on the lif curve the aircraf should be decelerated to point ‘C’.

(Reer to Page 76 or a review o the Interpretation o the Lif Curve i necessary.)

8


C

CL

FLAPS EXTENDED

FLAPS RETRACTED

B A

ANGLE OF ATTACK

Figure 8.22 Deployment o flaps or landing

226

8


With the flaps lowered, the stalling speed will:

a. b. c. d. 2.

When flaps are lowered the stalling angle o attack o the wing:

a. b. c. d. 3.

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.

increase. decrease. remain the same but will occur at a higher angle o attack. remain the same but will occur at a lower angle o attack.

increases the angle o descent without increasing the airspeed. decreases the angle o descent without increasing power. eliminates floating. permits approaches at a higher indicated airspeed.

Lowering flaps sometimes produces a pitch moment change due to:

a. b. c. d. 7.

s n o i t s e u Q

Lowering the flaps during a landing approach:

a. b. c. d. 6.

8

When a leading edge slot is opened, the stalling speed will:

a. b. c. d. 5.

remains the same, but C LMAX increases. increases and C LMAX increases. decreases, but C LMAX increases. decreases, but C LMAX remains the same.

With ull flap, the maximum lif/drag ratio:

a. b. c. d. 4.

increase. decrease. increase, but occur at a higher angle o attack. remain the same.

decrease o the angle o incidence. movement o the centre o pressure. movement o the centre o gravity. increased angle o attack o the tailplane.

Which type o flap would give the greatest change in pitching moment?

a. b. c. d.

Split. Plain. Fowler. Plain slotted.

227

8

Questions 8.

A split flap is:

a. b. c. d. 9.

I the flaps are lowered in flight, with the airspeed kept constant, to maintain level flight the angle o attack:

a. b. c. d.

8

Q u e s t i o n s

10.

slotted Krueger flap. slotted variable camber flap. slotted slat. slotted Fowler flap.

With reerence to Annex B , the type o flap illustrated is a:

a. b. c. d.

228

reduced. increased. the same as or a landing with flaps. the same as or a landing with flaps but with a steeper approach.

With reerence to Annex A , the type o flap illustrated is a:

a. b. c. d. 14.

does not change. increase towards the tip. increases towards the root. increases in speed but has no change o direction.

I a landing is to be made without flaps, the landing speed must be:

a. b. c. d. 13.

the lif would not change until the aircraf is airborne. the lif would increase when the flaps are lowered. the lif would decrease. the acceleration would increase.

When flaps are lowered the spanwise flow on the upper surace o 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.

I flaps are lowered during the take-off run:

a. b. c. d. 11.

a flap divided into sections which open to orm slots through the flap. a flap manuactured in several sections to allow or wing flexing. a flap which can move up or down rom the neutral position. a flap where the upper surace contour o the wing trailing edge is fixed and only the lower surace contour is altered when the flaps are lowered.

slat. Fowler flap. Krueger flap. variable camber flap.

8

Questions Annexes

8

s n o i t s e u Q

229

8

Answers

Answers

8

A n s w e r s

230

1 b

2 c

13 d

14 d

3 b

4 b

5 a

6 b

7 c

8 d

9 a

10 b

11 c

12 b

Chapter

9 Airframe Contamination

Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 233 Types o Contamination . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 233 Effect o Frost and Ice on the Aircraf . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 233 Effect on Instruments. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 234 Effect on Controls . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 234 Water Contamination. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 234 Airrame Aging . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 234 Questions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

235

Answers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

238

231

9

Airframe Contamination

9

A i r f r a m e C o n t a m i n a t i o n

232

9

Airframe Contamination Introduction The airrame may become contaminated by ice, rost or water either whilst it is in flight or when standing on the ground. The meteorological conditions that cause ice and rost to orm are dealt with elsewhere, but the effect is an accumulation o ice or rost on the surace o the aircraf which will affect its perormance and handling.

Types of Contamination a)

Frost

Frost can orm on the surace o the aircraf either when it is standing on the ground when the temperature alls below 0°C, or in flight, i the aircraf, afer flying in a region where the temperature is below 0°C, moves into a warmer layer o air. It consists o a airly thin coating o crystalline ice. b)

9

n o i t a n i m a t n o C e m a r f r i A

Ice

The main orms o icing are clear ice, rime ice and rain ice. Clear ice (glaze ice) is a translucent layer o ice with a smooth surace, caused by large supercooled water droplets striking the leading edges o the airrame. As there is some delay in reezing, there is some flow back along the surace behind the leading edge. Rime ice orms when small supercooled water droplets strike the leading edges and reeze almost immediately so that there is no flow back. It is a white opaque ormation. Rain ice is caused by rain which becomes supercooled by alling rom an inversion into air which is below 0°C. It does not reeze immediately and orms considerable flow back, and it builds up very quickly.

Effect of Frost and Ice on the Aircraft The ormation o ice and rost on the airrame will: • modiy the profile o the aerooil. • increase the roughness o the aircraf surace. • increase the weight o the aircraf. The main effect o rost will be to increase the surace roughness and this will increase the energy loss in the boundary layer. The skin riction drag will increase and the boundary layer will have an earlier separation, giving a reduced C LMAX. Take-off with rost on the wings could result in a stall afer lif-off i the normal take-off speed is used. Tests have shown that rost, ice or snow with the thickness and sur ace roughness o medium or coarse sandpaper reduces lif by as much as 30% and increases drag by 40%.

Ice will normally orm on and behind the leading edges o wings and tailplane and can result in severe distortion o the leading edge profile. This will give a large increase in drag and a substantial decrease in C LMAX.

233

9

Airframe Contamination The reduced CLMAX o the wing will give a higher stalling speed and the decreased C LMAX o the tailplane could cause it to stall when the aircraf is flying at low speed, particularly i the wing downwash is increased as a result o flap extension. Tailplane stall will result in loss o longitudinal control. Clear ice and rain ice especially can add considerable weight to the airrame, and this will in turn give a higher stalling speed, as well as increased induced drag. The margin o thrust to drag will be decreased, reducing the ability to climb. Increased power will be required to maintain height, resulting in increased uel consumption. Ice ormation on propeller blades can upset the balance o the propeller and cause severe vibration, particularly i pieces o ice break off rom one blade. Pieces o ice shed rom propellers can also cause damage to the uselage.

Effect on Instruments

9

Formation o ice on static vents and pitot heads could cause errors in the readings o pressure instruments and, eventually, ailure to show any reading.

A i r f r a m e C o n t a m i n a t i o n

Effect on Controls Any moveable surace could become jammed by ice orming in the gaps around the control or by pieces o ice breaking off and becoming jammed in the control gaps. The controls could become difficult to operate or immovable.

Water Contamination I the wings are contaminated with water due to heavy rain, the boundary layer may become turbulent urther orward on the wing, particularly i the section is o the 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 o the aircraf manuacturer or aircraf operator when taking off and landing in heavy rain.

Airframe Aging Over a period o years the condition o the airrame will deteriorate due to small scratches, minor damage, repairs and general accumulation o dir t and grease. The overall effect o this will be to increase the drag o the aircraf (mainly skin riction drag) with a consequent increase in uel consumption. The cost o operating the aircraf will thereore increase with the age o the airrame. The normal deterioration o the airrame is allowed or in the perormance charts o the aeroplane.

234

9


Afer an aircraf 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.

s n o i t s e u Q

Increased angle o attack or stalls. Increased stall speed. Increased pitch down tendencies. Decreased speed or stalling.

the aircraf to stall at an angle o attack that is lower than normal. no problems to pilots. drag actors so large that sufficient speed cannot be obtained or take-off. the aircraf to stall at an angle o attack that is higher than normal.

I it is suspected that ice may have ormed on the tailplane and longitudinal control difficulties are experienced ollowing flap selection, the prudent action to take would be:

a. b. c. d. 6.

9

Frost covering the upper surace o an aircraf wing will usually cause:

a. b. c. d. 5.

relative humidity is low and temperature rises. pressure is high and humidity alls. relative humidity is high and temperature is low. relative pressure is high and temperature is high.

Which is an effect o ice, snow or rost ormation on an aeroplane?

a. b. c. d. 4.

snow should be removed but smooth ice may be lef. all snow and ice should be removed. loose snow may be lef but ice must be removed. providing the contamination is not too thick, it may be lef in place.

immediately decrease the flap setting. allow the speed to increase. select a greater flap deflection because this will increase C LMAX. reduce the angle o attack.

When considering in-flight airrame contamination with rost or ice, which o the ollowing statements is correct?

a. b. c. d.

Build-up can be identified by the ice detection equipment fitted to the aircraf. The pilot can visually identiy 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 o the accumulation o airrame icing may not exist. Due to the high speed o modern aircraf, significant airrame contamination with rost, ice or snow will not occur.

235

9

Questions 7.

9

Q u e s t i o n s

236

In the event o an icing-induced wing stall, which o the ollowing indications will reliably be available to the flight crew? 1. 2. 3. 4. 5. 6.

Activation o the stall warning device (horn or stick shaker). The aircraf pitching nose down. Loss o elevator effectiveness. Airrame buffet. A roll control problem (increasing roll oscillation or violent wing drop). A high rate o descent.

a. b. c. d.

1, 2, 3, 4, 5 and 6. 1, 3 and 4. 1, 4 and 6. 5 and 6.

9

Questions

9

s n o i t s e u Q

237

9

Answers

Answers 1 b

9

A n s w e r s

238

2 c

3 b

4 a

5 a

6 c

7 d

Chapter

10 Stability and Control

Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 241 Static Stability . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 241 Aeroplane Reerence Axes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 244 Static Longitudinal Stability . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 244 Neutral Point . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 249 Static Margin . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 250 Trim and Controllability. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 251 Key Facts 1. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 254 Graphic Presentation o Static Longitudinal Stability . . . . . . . . . . . . . . . . . . . . . . . . 256 Contribution o the Component Suraces . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 259 Power-off Stability . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 264 Effect o CG Position . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 265 Power Effects . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 266 High Lif Devices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 268 Control Force Stability . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 269 Manoeuvre Stability . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

274

Stick Force Per ‘g’ . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 275 Tailoring Control Forces. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 277 Longitudinal Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 279 Manoeuvring Control Requirement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 279 Take-off Control Requirement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 280 Landing Control Requirement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 281 Dynamic Stability . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 282 Longitudinal Dynamic Stability . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 286 Long Period Oscillation (Phugoid). . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 287 Short Period Oscillation. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 288 Directional Stability and Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 290 Continued Overlea

239

10

Stability and Control Sideslip Angle . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .291 Static Directional Stability . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .292 Contribution o the Aeroplane Components. . . . . . . . . . . . . . . . . . . . . . . . . . .293 Lateral Stability and Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .301 Static Lateral Stability . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .302 Contribution o the Aeroplane Components . . . . . . . . . . . . . . . . . . . . . . . . . . 304 Lateral Dynamic Effects. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .309 Spiral Divergence . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .309 Dutch Roll . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .309 Pilot Induced Oscillations (PIO) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .310 High Mach Numbers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .311

1 0

Mach Trim . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .311 Key Facts 2. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .312

S t a b i l i t y a n d C o n t r o l

Summary. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .315 Questions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

318

Key Facts 1 (Completed) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .323 Key Facts 2 (Completed) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .326 Answers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

240

330

Stability and Control

10

Introduction Stability is the tendency o an aircraf to return to a steady state o flight without any help rom the pilot, afer being disturbed by an external orce. An aircraf must have the ollowing qualities: • Adequate stability to maintain a uniorm flight condition. • The ability to recover rom various disturbing influences. • Sufficient stability to minimize the workload o the pilot. • Proper response to the controls so that it may achieve its design perormance with adequate manoeuvrability. There are two broad categories o stability, static and dynamic. Dynamic stability will be considered later.

0 1

l o r t n o C d n a y t i l i b a t S

Static Stability An aircraf is in a state o equilibrium (trim) when the sum o all orces is zero and the sum o all moments is zero; there are no accelerations and the aircraf will continue in steady flight. I equilibrium is disturbed by a gust, or deflection o the controls, the aircraf will experience accelerations due to an unbalance o moments or orces. The type o static stability an aircraf possesses is defined by its initial tendency , ollowing the removal o some disturbing orce. • Positive static stability (or static stability) exists i an aircraf is disturbed rom equilibrium and has the tendency to return to equilibrium. • Neutral static stability exists i an aircraf is subject to a disturbance and has neither the tendency to return nor the tendency to continue in the displacement direction. • Negative static stability (or static instability) exists i an aircraf has a tendency to continue in the direction o disturbance. Examples o the three types o static stability are shown in Figure 10.1, Figure 10.2 and Figure 10.3

241

10

Stability and Control Figure 10.1 illustrates the condition o positive static stability (or static stability). The ball is

displaced rom equilibrium at the bottom o the trough. When the disturbing orce is removed, the initial tendency o the ball is to return towards the equilibrium condition. The ball may roll back and orth through the point o equilibrium but displacement to either side creates the initial tendency to return. POSIT IV E STAT IC STA BILITY

Tendency to Return to Equilibrium

1 0


Equilibrium Figure 10.1

Figure 10.2 illustrates the condition o neutral static stability . The ball encounters a new

equilibrium at any point o displacement and has no tendency to return to its original equilibrium. Equilibrium Encountered at any Point of Displacement

NEUTRA L STAT IC STA BILITY Figure 10.2

242


10

Figure 10.3 illustrates the condition o negative static stability (or static instability).

Displacement rom equilibrium at the hilltop gives a tendency or greater displacement. Tendency to Continue in Displacement Direction

Equilibrium

0 1


NEGAT IV E STAT IC STA BILITY Figure 10.3

The term “static” is applied to this orm o stability since any resulting motion is not considered. Only the initial tendency to return to equilibrium is considered in static stability. The static longitudinal stability o an aircraf is assessed by it being displaced rom some trimmed angle o attack. I the aerodynamic pitching moments created by this displacement tend to return the aircraf to the equilibrium angle o attack, the aircraf has positive static longitudinal stability .

243

10

Stability and Control Aeroplane Reference Axes In order to visualize the orces and moments on the aircraf, it is necessary to establish a set o reerence axes passing through the centre o gravity . Figure 10.4 illustrates a conventional right hand axis system. The longitudinal axis passes through the CG rom nose to tail. A moment about this axis is a rolling moment, L, and 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 aircraf 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.

1 0


Lateral Axis Positive Pitching Moment, M

Longitudinal Axis

Positive Rolling Moment,

Centre Of Gravity

Positive Yawing Moment, N

L

Normal Axis

Figure 10.4

Static Longitudinal Stability Longitudinal stability is motion about the lateral axis. To avoid conusion, consider the axis about which the particular type o 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.

244


10

• An aircraf will exhibit static longitudinal stability i it tends to return towards the trim angle o attack when displaced by a gust OR a control input. • It is essential that an aircraf has positive static longitudinal stability. I it is stable, an aeroplane is sae and easy to fly since it seeks and tends to maintain a trimmed condition o flight. It also ollows that control deflections and control “eel” (stick orce) must be logical, both in direction and magnitude. • I the aircraf is neutrally stable, it tends to remain at any displacement to which it is disturbed. • Neutral static longitudinal stability usually defines the lower limit o 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 aircraf has no tendency to return to trim ollowing a disturbance - generally, this would not be acceptable. • The aircraf which is unstable will continue to pitch in the disturbed direction until the displacement is resisted by opposing control orces.

0 1


• The aeroplane with negative static longitudinal stability is inherently divergent rom any intended trim condition. I it is at all possible to fly the aircraf, it cannot be trimmed and illogical control orces and deflections are required to provide equilibrium with a change o attitude and airspeed - clearly, this would be totally unacceptable.

245

10

Stability and Control For the study o stability it is convenient to consider the changes in magnitude o lif orce due to changes in angle o attack, acting through a stationary point, the aerodynamic centre (AC). It will be remembered that the location o the AC is at the quarter chord (or 25% af o the leading edge). It should be noted that the pitching moment about the AC is negative (nosedown) and that this negative (nose-down) pitching moment about the AC does not change with changes in angle o attack. Figure 10.5. L1

1

M

CP

AC

d1

1 0


L2

2

M

AC

CP

d2

MOMENT (M) REMAINS THE SAME AT "NORMAL" ANGLES OF ATTACK BECAUSE L 1 × d1 at

α1

=

L 2 × d2 at

α2

Figure 10.5 Aerodynamic centre (AC)

The pitching moment about the AC remains constant as the angle o attack is increased because the magnitude o the lif orce increases but acts through a smaller arm due to the CP moving orward. It is only at the AC (25% chord) that this will occur. I a point in ront o, or to the rear o the AC were considered, the pitching moment would change with angle o attack. For the study o stability, we will consider the lif to act at the AC. The AC is a stationary point located at the 25% chord only when the airflow is subsonic .

246


10

L

L x

AC

Flight Path

wing

Momentary Relative Airflow due to Gust

0 1


Figure 10.6 A wing alone is unstable

A wing considered alone is statically unstable because the AC is in ront o the CG, Figure 10.6 . A vertical gust will momentarily increase the angle o attack and increase lif (∆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 o attack urther, an unstable pitching moment. The wing on its own would rotate nose-up about the CG, Figure 10.7 .

L

AN AIRCRAFT ROTATES AROUND ITS CG

L

AC

x

CG

UNSTABLE (NOSE-UP) PITCHING MOMENT ABOUT THE CG Figure 10.7

Now consider a wing together with a tailplane. The tailplane is positioned to generate a stabilizing pitching moment about the aircraf CG. The same vertical gust will increase the angle o attack o the tailplane and increase tailplane lif (∆Lt), which, when multiplied by arm ‘y’, will generate a negative (nose-down) pitching moment about the aircraf CG.

247

10

Stability and Control I the tail moment is greater than the wing moment, the sum o the moments will not be zero and the resultant nose-down moment will give an angular acceleration about the CG. The nosedown angular acceleration about the CG will return the aircraf towards its original position o equilibrium. The greater the tail moment relative to the wing moment, the greater the rate o acceleration towards the original equilibrium position. (Too much angular acceleration is not good).

L

L

x

y Lt A IRCRAFT CG

Flight Path

1 0

Lt

AC wing

AC tail




=

Change in angle of att ack due to gust

AC

=

Aerodynamic Centre

L

=

L

=

Lt

=

Tailplane lift

Lt

=

Change in tailplane 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 unction o the change in wing lif multiplied by arm ‘x’. The tail moment is a unction o the change in tailplane lif multiplied by arm ‘y’, Figure 10.8. The length o both arms is dependent upon CG position. I the CG is considered in a more orward position, the tail arm is larger and the wing arm is smaller. A more orward CG position increases static longitudinal stability. I the nose-down (negative) tail moment is greater than the nose-up (positive) wing moment, the aircraf will have static longitudinal stability .

248


10

INCREASED

L

L

DECREASED

x

y Lt Lt

AC

AC

NEUTRAL POINT 0 1

POSITION OF CG W HEN


TAIL MOMENT AND W ING MOMENT ARE EQUAL

Figure 10.9 Neutral point

Neutral Point I you consider the CG moving rearwards rom a position o static longitudinal stability: • the tail arm ‘y’ will decrease and the wing arm ‘x’ will increase; consequently, • the (negative) tail moment will decrease and the (positive) wing moment will increase, Figure 10.9. Eventually the CG will reach a position at which the tail moment is the same as the wing moment. I a vertical gust displaces the aircraf nose-up, the sum o the moments will be zero and there will be no angular acceleration about the CG to return the aircraf towards its original position o equilibrium. Because there is no resultant moment, either nose-up or nose-down , the aircraf will remain in its new position o equilibrium; the aircraf will have neutral static longitudinal stability. See page 245. The position o the CG when the sum o the changes in the tail moment and wing moment caused by the gust is zero is known as the neutral point, Figure 10.9.

249

10

Stability and Control Static Margin We have established that with the CG on the neutral point the aircraf 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. I the CG is positioned just orward o 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 o attack will generate a small nose-down angular acceleration about the CG, which will gently return the aircraf towards its original position o trim (equilibrium). The urther orward the CG, the greater the nose-down angular acceleration about the CG the greater the degree o static longitudinal stability. STATIC MA RGIN

1 0

L


L

x

y Lt Lt

AC

AC

A FT CG LIMIT

NEUTRAL POINT

Figure 10.10 Static margin & af CG limit

The neutral point is an important point o reerence in the study o static longitudinal stability. In practice, the CG will never be allowed to move so ar af that it reached the neutral point. The aircraf would be much too sensitive to the controls. It has been stated that the urther orward the CG is rom the neutral point, the greater the static longitudinal stability. The distance the CG is orward o the neutral point will give a measure o the static longitudinal stability; this distance is called the static margin, Figure 10.10. The greater the static margin, the greater the static longitudinal stability. A certain amount o static longitudinal stability is always required, so the af CG limit will be positioned some distance orward o the neutral point. The distance between the neutral point and the af CG limit gives the required minimum static stability margin .

250


10

Trim and Controllability An aircraf is said to be trimmed (in trim) i all moments in pitch, roll, and yaw are equal to zero. The establishment o trim (equilibrium) at various conditions o flight may be accomplished by: • • • • •

pilot effort trim tabs variable incidence trimming tailplane moving uel between the wing tanks and an af located trim tank, or bias o a surace actuator (powered flying controls)

The term controllability reers to the ability o the aircraf to respond to control surace displacement and achieve the desired condition o flight. Adequate controllability must be available to perorm take-off and landing and accomplish the various manoeuvres in flight. A contradiction exists between stability and controllability. A high degree o stability gives reduced controllability . The relationship between static stability and controllability is demonstrated by the ollowing our illustrations.

0 1


POSIT IV E STAT IC STA BILITY

Figure 10.11

Degrees o static stability are illustrated by a ball placed on various suraces. Positive static stability is shown by the ball in a trough, Figure 10.11; i the ball is displaced rom equilibrium at the bottom o the trough, there is an initial tendency to return to equilibrium. I it is desired to “control” the ball and maintain it in the displaced position, a orce must be supplied in the direction o displacement to balance the inherent tendency to return to equilibrium. This same stable tendency in an aircraf resists displacement rom trim equally, whether by pilot effort on the controls (stick orce) or atmospheric disturbance .

251

10


INCREASED POSITIVE STAT IC STA BILITY

Figure 10.12 1 0

The effect o increased static stability (orward CG movement) on controllability is illustrated by the ball in a steeper trough, Figure 10.12. A greater orce is required to “control” the ball to the same position o displacement when the static stability is increased. In this manner, a large degree o static stability tends to make the aircraf less controllable. It is necessary to achieve the proper proportion between static stability and controllability during the design o an aircraf because too much static stability (orward CG position) reduces controllability. The orward CG limit is set to ensure minimum controllability, Figure 10.13.


STATIC MARGIN

L

L x

y Lt Lt

AC

AC

FWD CG LIMIT NEUTRA L POINT

HIGH STICK FORCE

A FT CG LIMIT LOW STICK FORCE

Figure 10.13

252


10

NEUT RA L STAT IC STA BILITY

Figure 10.14

The effect o reduced static stability on controllability is shown by the ball on a flat surace, Figure 10.14. I neutral static stability exists (CG on the neutral point), the ball may be displaced rom equilibrium and there is no tendency to return. A new point o equilibrium is obtained and no orce 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 o displacement, aerodynamic damping. For this reason, decreased static stability (af CG movement) increases controllability . I the stability o the aircraf is too low, control deflections may create exaggerated displacements o the aircraf.

0 1


NEGAT IV E STAT IC STA BILITY

Figure 10.15

The effect o static instability on controllability (CG af o the neutral point) is shown in Figure 10.15 by the ball on a hill. I the ball is displaced rom equilibrium at the top o the hill, the initial tendency is or the ball to continue in the displaced direction. In order to “control” the ball at this position o displacement, a orce must be applied opposite to the direction o displacement. This effect would be apparent during flight by an unstable “eel” to the aircraf. I the controls were deflected to increase the angle o attack, the aircraf would need to be ‘held’ at the higher angle o attack by a push orce to keep the aircraf rom continuing in the nose-up direction. The pilot would be supplying the stability by his attempt to maintain the equilibrium; this is totally unacceptable!

253

10

Stability and Control Key Facts 1 Sel Study (Insert the missing words, with reerence to the preceding paragraphs). Stability is the ________ o an aircraf to return to a _____ state o flight, afer being disturbed by an external ______, without any help rom the _____. There are two broad categories o stability: ________ and ________ . An aircraf is in a state o __________ (trim) when the sum o all orces is ____ and the sum o all ________ is zero. The type o static stability an aircraf possesses is defined by its ______ tendency, ollowing the removal o some disturbing orce. The three different types o static stability are:

1 0


a)

_________ static stability exists i an aircraf is disturbed rom equilibrium and has the tendency to return to equilibrium.

b)

______ static stability exists i an aircraf 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 i an aircraf has a tendency to continue in the direction o disturbance.

The longitudinal axis passes through the ____ rom _____ to _____. The normal axis passes “vertically” through the ___ at __° to the ___________ axis. The lateral axis is a line passing through the ___, parallel to a line passing through the ____ tips. The three reerence axes all pass through the _______ ___ ________. Lateral stability involves motion about the __________ axis (_______). Longitudinal stability involves motion about the ______ axis (_______). Directional stability involves motion about the _______ axis (_______). We consider the changes in __________ o lif orce due to changes in angle o ________, acting through a __________ point; the ___________ ______. The aerodynamic centre (AC) is located at the ___% chord position. The _________ pitching moment about the AC remains ________ at normal angles o attack. A wing on its own is statically ________ because the ___ is in ront o the ___. An upward vertical gust will momentarily ________ the angle o attack o the wing. The ________ lif orce magnitude acting through the ___ will increase the ______ pitching moment about the ___. This is an ________ pitching moment. The ________ is positioned to generate a _________ pitching moment about the aircraf ___.

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10

I the tail moment is greater than the wing moment, the sum o the moments will not be ____ and the resultant nose _____ moment will give an angular _________ about the ____. The ______ the tail moment relative to the wing moment, the _______ the rate o return _______ the original __________ position. The tail moment is increased by moving the aircraf ___ orwards, which _________ the tail arm and decreases the _____ arm. I the nose-down (_______) tail moment is greater than the nose-up (_______) wing moment, the aircraf will have _______ __________ stability. The position o the CG when changes in the sum o the tail m oment and wing moment due to a disturbance is zero is known as the ______ _____. The urther orward the ___, the ______ the nose-down angular __________ about the ___ the ______ the degree o _____ __________ stability.

0 1

The _______ the ___ is orward o the ________ point will give a measure o the _____ longitudinal stability; this distance is called the static ______.


The greater the static margin, the ______ the _______ ___________ stability. The ____ CG limit will be positioned some distance _______ o the _____ _____. The distance between the ___ ___ limit and the neutral point gives the required _________ static stability ________. An aircraf is said to be _______ i all ________ in pitch, roll, and yaw are equal to _____. Trim ( __________ ) is the unction o the _______ and may be accomplished by: a)

______ effort

b)

trim _____,

c)

moving _____ between the wing ______ and an af located _____ tank, or

d)

bias o a surace _______ ( ________ flying controls).

The term ____________ reers to the ability o the aircraf to respond to control surace displacement and achieve the desired ________ o flight. A high degree o stability tends to reduce the ____________ o the aircraf. The stable tendency o an aircraf resists displacement rom ___ equally, whether by ____ effort on the controls ( ______ orce) or _____. I the CG moves orward, static longitudinal stability ________ and controllability _________ (stick orce ________). I the CG moves af, static longitudinal stability __________ and controllability ________ (stick orce ________ ).

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10

Stability and Control With the CG on the orward limit, static longitudinal stability is _______, controllability is ____ and stick orce is _____. With the CG on the af limit, static longitudinal stability is _____, controllability is _______ and stick orce is ____. The af CG limit is set to ensure a _________ degree o static longitudinal stability. The wd CG limit is set to ensure a _________ degree o controllability under the worst circumstance. KEY FACTS 1 WITH THE MISSING WORDS INSERTED CAN BE FOUND AT THE END OF THIS CHAPTER.

Graphic Presentation of Static Longitudinal Stability Static longitudinal stability depends upon the relationship o angle o attack and pitching moment. It is necessary to study the pitching moment contribution o each component o the aircraf. In a manner similar to all other aerodynamic orces, the pitching moment about the lateral axis is studied in the coefficient orm.

1 0


M = CM Q S (MAC) or M Q S (MAC)

CM = where: M

=

pitching moment about the CG (positive i in a nose-up direction)

Q

=

dynamic pressure

S

=

wing area

MAC

=

mean aerodynamic chord

=

pitching moment coefficient

CM

The pitching moment coefficients contributed by all the various components o the aircraf are summed up and plotted versus lif coefficient (angle o attack). Study o the plots o C M versus CL is a convenient way to relate the static longitudinal stability o an aeroplane.

256


+

10

A TRIM CM = 0

x

LIFT COEFFICIENT CL

y

Figure 10.16 Graph A

0 1


Graph A illustrates the variation o pitching moment coefficient (C M) with lif coefficient (C L) or an aeroplane with positive static longitudinal stability. Evidence o 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 C M = 0, and i the aeroplane is disturbed to some different CL, the pitching moment change tends to return the aircraf to the point o trim. I the aeroplane were disturbed to some higher C L (point y), a negative or nose-down pitching moment is developed which tends to decrease angle o attack back to the trim point. I the aeroplane were disturbed to some lower C L (point x), a positive or nose-up pitching moment is developed which tends to increase the angle o attack back to the trim point. Thus, positive static longitudinal stability is indicated by a negative slope o C M versus CL. The degree o static longitudinal stability is indicated by the slope o the curve (red line).

+

B

STABLE TRIM NEUTRAL

CM

CL

UNSTABLE

Figure 10.17 Graph B

Graph B provides comparison o 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 i the curve had zero slope. I neutral stability existed, the aeroplane could be disturbed to some higher or lower lif coefficient without change in pitching moment coefficient.

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10

Stability and Control 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 o the C M versus CL curve (blue line) would be unstable. I the unstable aeroplane were subject to any disturbance rom equilibrium at the trim point, the changes in pitching moment would only magniy the disturbance. When the unstable aeroplane is disturbed to some higher CL a positive change in C M occurs which would illustrate a tendency or continued, greater displacement. When the unstable aeroplane is disturbed to some lower C L a negative change in CM takes place which tends to create continued displacement.

+

C STABLE UNSTABLE

1 0

CM CL


LESS STABLE NEUTRAL

Figure 10.18 Graph C

Ordinarily, the static longitudinal stability o a conventional aeroplane configuration does not vary with lif coefficient. In other words, the slope o CM versus CL does not change with C L. However, i: • the aeroplane has sweepback, • there is a large contribution o “power effect” on stability, or • there are significant changes in downwash at the horizontal tail, noticeable changes in static stability can occur at high lif coefficients (low speed). This condition is illustrated by graph C. The curve o CM versus CL o this illustration shows a good stable slope at low values o 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 lif coefficient, the static stability o the aeroplane is depicted by the slope o the curve o CM versus CL.

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10

Contribution of the Component Surfaces The net pitching moment about the lateral axis is due to the contribution o each o the component suraces acting in their appropriate flow fields. By studying the contribution o each component, their effect on static stability may be appreciated. It is necessary to recall that the pitching moment coefficient is defined as: CM =

M Q S (MAC)

Thus, any pitching moment coefficient (CM) - regardless o source - has the common denominator o dynamic pressure (Q), wing area (S), and wing mean aerodynamic chord (MAC). This common denominator is applied to the pitching moments contributed by the: • uselage and nacelles, • horizontal tail, and • power effects as well as pitching moments contributed by the wing.

0 1


Wing The contribution o the wing to stability depends primarily on the location o the aerodynamic centre (AC) with respect to the aeroplane centre o 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 lif coefficient . All changes in lif coefficient

effectively take place at the wing aerodynamic centre. Thus, i the wing experiences some change in lif coefficient, the pitching moment created will be a direct unction o the relative location o the AC and CG. Note: The degree o positive camber o the wing has no effect on longitudinal stability. The pitching moment about the AC is always negative regardless o angle o attack .

Stability is given by the development o restoring moments. As the wing AC is orward o the CG, the wing contributes an unstable pitching moment to the aircraf, as shown in Figure 10.19.

259

10


CHANGE IN LIFT

CG A ERODYNAMIC CENTRE

1 0


CG AFT OF AC

CM

UNSTABLE SLOPE

CL

Figure 10.19 Unstable wing contribution

Since the wing is the predominating aerodynamic surace o an aeroplane, any change in the wing contribution may produce a significant change in the aeroplane stability.

260


10

Figure 10.20

Fuselage and Nacelles

0 1

In most cases, the contribution o the uselage and nacelles is destabilizing. A symmetrical body in an airflow develops an unstable pitching moment when given an angle o attack. In act, an increase in angle o attack produces an increase in the unstable pitching moment without the development o lif. Figure 10.20 illustrates the pressure distribution which creates this unstable moment on the body. An increase in angle o attack causes an increase in the unstable pitching moment but a negligible increase in lif.


Horizontal Tail The horizontal tail usually provides the greatest stabilizing influence o all the components o the aeroplane.

L L

y

x

Lt Flight Path

AC

Lt

wing


tail

Momentary Relative Airflow due to Gust Figure 10.21

To appreciate the contribution o the horizontal tail to stability, inspect Figure 10.21. I the aeroplane is given an increase in angle o attack (by a gust OR control displacement), an increase in tail lif will occur at the aerodynamic centre o the tail. An increase in lif at the horizontal tail produces a negative (stabilizing) moment about the aircraf CG.

261

10

Stability and Control For a given vertical gust velocity and aircraf TAS, the wing moment is essentially determined by the CG position. BUT, the tail moment is determined by the CG position AND the effectiveness o the tailplane. For a given moment arm (CG position), the effectiveness o the tailplane is dependent upon: • Downwash rom the wing. • Dynamic pressure at the tailplane. • Longitudinal dihedral. Downwash rom the wing and dynamic pressure at the tailplane will be discussed in due course, but the effect o longitudinal dihedral is shown below.

Longitudinal Dihedral This is the difference between tailplane and wing incidence. For longitudinal static stability the tailplane incidence is smaller. As illustrated in Figure 10.22, this will generate a greater percentage increase in tailplane lif than wing lif or a given vertical gust.

1 0

This guarantees that the positive contribution o the tailplane to static longitudinal stabilit y will be sufficient to overcome the sum o the destabilizing moments rom the other components o the aeroplane.


L = 100% L t = 200%

4º INCIDENCE

2º INCIDENCE

AC AC

4º INCREASE IN ANGLE OF ATTACK DUE TO VERTICAL GUST

Figure 10.22

262


10

DOW NWASH AT HORIZONTAL TAIL

Figure 10.23

Downwash It should be appreciated that the flow at the horizontal tail does not have the same flow direction or dynamic pressure as the ree stream. Due to the wing wake, uselage boundary layer and power effects, the dynamic pressure at the horizontal tail may be greatly different rom the dynamic pressure o the ree stream. In most instances, the dynamic pressure at the tail is usually less and this reduces the efficiency o the tail.

0 1


When the aeroplane is given a change in angle o attack, the horizontal tail does not experience the same change in angle o attack as the wing, Figure 10.23. Because o the increase in downwash behind the wing, the horizontal tail will experience a smaller change in angle o attack, e.g. i a 10° change in wing angle o attack causes a 4° increase in downwash at the horizontal tail, the horizontal tail experiences only a 6° change in angle o attack. In this manner, the downwash at the horizontal tail reduces the contribution to stability. Any actor which alters the rate o change o 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 .

263

10

Stability and Control Power-off Stability When the aerodynamic stability o a configuration is o interest, power effects are neglected and the stability is considered by a build-up o the contributing components. Figure 10.24 illustrates a typical build-up o the components o a conventional aeroplane

configuration. I the CG is arbitrarily set at 30 percent MAC, the contribution o the wing alone is destabilizing, as indicated by the positive slope o C M versus CL. The combination o the wing and uselage increases the instability. The contribution o the tail alone is highly stabilizing rom the large negative slope o the curve. The contribution o the tail must be sufficiently stabilizing so that the complete configuration will exhibit positive static stability at the anticipated CG locations.

TYPICAL BUILD-UP OF COMPONENTS 1 0


W ING + FUSELAGE

CM

W ING ONLY

AEROPLANE CG @ 30% MAC

TAILPLANE ONLY Figure 10.24

264

CL


CM

10

50% MAC

40% MAC (NEUTRAL POINT)

30% MAC

CL

20% MAC 0 1


10% MAC

Figure 10.25

Effect of CG Position A variation o 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. I the incidence o all suraces remains fixed, the effect o CG position on static longitudinal stability is typified by the chart in Figure 10.25. As the CG is gradually moved af, the aeroplane static stability decreases, then becomes neutral then unstable. The CG position which produces zero slope and neutral static stability is reerred to as the “neutral point”. The neutral point may be imagined as the effective aerodynamic centre o the entire aeroplane configuration, i.e. with the CG at the neutral point, all changes in net lif effectively occur at that point and no change in pitching moment results. The neutral point defines the most af CG position without static instability.

265

10

Stability and Control Power Effects The effects o power may cause significant changes in trim lif 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 .

DESTABILIZING 1 0

Figure 10.26


The effects o power are considered in two main categories. First, there are the direct effects resulting rom the orces created by the propulsion unit. Next, there are the indirect effects o the slipstream and other associated flow which alter the orces and moments o the aerodynamic suraces. The direct effects o power are illustrated in Figure 10.26 . The vertical location o the thrust line defines one o the direct contributions to stability. I the thrust line is below the CG, as illustrated, a thrust increase will produce a positive or nose-up moment and the effect is destabilizing.

NORMAL FORCE DUE TO MOMENT CHANGE

Figure 10.27

A propeller located ahead o the CG contributes a destabilizing effect. As shown in Figure 10.27 , a rotating propeller inclined to the relative airflow causes a deflection o the airflow. The momentum change o the slipstream creates a normal orce at the plane o the propeller. As this normal orce will increase with an increase in aeroplane angle o attack, the effect will be destabilizing when the propeller is ahead o the CG. The magnitude o the unstable contribution depends on the distance rom the CG to the propeller and is largest at high power and low dynamic pressure.

266


10

W ING, NACELLE AND FUSELAGE MOMENTS AFFECTED BY SLIPSTREAM

DYNAMIC PRESSUR AT TAIL AFFECTED BY SLIPSTREAM

W ING LIFT AFFECTED BY SLIPSTREAM

0 1

Figure 10.28


The indirect effects o power are o greatest concern in the propeller powered aeroplane rather than the jet powered aeroplane. As shown in Figure 10.28, the propeller powered aeroplane creates slipstream velocities on the various suraces which are different rom the flow field typical o power-off flight. Since the various wing, nacelle and uselage suraces are partly or wholly immersed in this slipstream, the contribution o these components to stability can be quite different rom the power-off flight condition. Ordinarily, the change o uselage and nacelle contribution with power is relatively small. The added lif on the portion o the wing immersed in the slipstream requires that the aeroplane operate at a lower angle o attack to produce the same effective lif coefficient. Generally, this reduction in angle o attack to effect the same C L reduces the tail contribution to stability. However, the increase in dynamic pressure at the tail tends to increase the effectiveness o the tail and may be a stabilizing effect. The magnitude o this contribution due to the slipstream velocity on the tail will depend on the CG position and trim lif coefficient.

DOW NWASH AT TA IL A FFECTED BY SLIPST REAM DIRECT ION

Figure 10.29

The deflection o the slipstream shown in Figure 10.29 by the normal orce at the propeller tends to increase the downwash at the horizontal tail and reduce the contribution to stability.

267

10


FLOW INDUCED BY JET FAN EXHAUST

Figure 10.30 1 0

Essentially the same destabilizing effect is produced by the flow induced at the exhaust o turbo-jet/an engines, Figure 10.30. Ordinarily, the induced flow at the horizontal tail o a jet aeroplane is slight and is destabilizing when the jet passes underneath the horizontal tail. The magnitude o 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 o static stability at high power, high C L and low dynamic pressure. It is generally true that any aeroplane will experience the lowest level o static longitudinal stability under these conditions. Because o the greater magnitude o both direct and indirect power effects, the propeller powered aeroplane usually experiences a greater effect than the jet powered aeroplane.

High Lift Devices An additional effect on stability can be rom the extension o high lif devices. High lif devices tend to increase downwash at the tail and reduce the dynamic pressure at the tail, both o which are destabilizing. However, high lif devices may prevent an unstable contribution o the wing at high CL. While the effect o high lif devices depends on the aeroplane configuration, the usual effect is destabilizing. Hence, the aeroplane may experience the most critical orward neutral point during the power approach or overshoot/missed approach. During this condition o flight, the static stability is usually the weakest and particular attention must be given to precise control o the aeroplane. The power-on neutral point may set the most af limit o CG position.

268


10

Control Force Stability The static longitudinal stability o an aeroplane is defined by the tendency to return to equilibrium upon displacement. In other words, a stable aeroplane will resist displacement rom trim or equilibrium. The control orces o the aeroplane should reflect the stability o the aeroplane and provide suitable reerence to the pilot or precise control o the aeroplane.

EFFECT OF ELEVATO R DEFLECT ION

ELEVATOR DEFLECTION

CM

+

0 1


TRIM FOR 10º UP

TRIM FOR 0 º

C

L

CG @ 20% MAC

Figure 10.31

The effect o elevator deflection on pitching moments is illustrated by the graph o Figure 10.31. I the elevators o the aeroplane are held at zero deflection, the resulting line o C M versus CL or 0° depicts the static stability and trim lif coefficient. I the elevators are held at a deflection o 10° up (aircraf trimmed at a lower speed), the aeroplane static stability is unchanged but the trim lif coefficient is increased.

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10

Stability and Control As the elevator is held in various positions, equilibrium (trim) will occur at various lif coefficients, and the trim C L can be correlated with elevator deflection as shown in Figure 10.32.

TR IM

ELEVAT VAT OR DEFLECTION C L V ERSUS ELE

CG LOCATION 10% MAC

20% MAC UP 30% MAC

1 0

C S t a b i l i t y a n d C o n t r o l

L

40% MAC (NEUTRAL POINT) DOW DO WN

Figure 10.32

When the CG position o the aeroplane is fixed, each elevator position corresponds to a particular trim lif coefficient. As the CG is moved af, the slope o o this line decreases, and the decrease in stability is evident by a given g iven control displacement causing a greater greater change in trim lif coefficient. This is evidence that decreasing decreasing stability causes increased controllability and, o course, increasing stability decreases controllability.

I the CG is moved af until the line o trim C L versus elevator deflection has zero slope, neutral static stability is obtained.

A change in elevat elevator or position does not alter the tail contribution to stability

270


10

TR IM A IRSPEED V ERSUS ELEVAT ELEVAT OR DEFLECTION

STABLE

UNSTABLE

UP

EQUIVALENT AIRSPEED DOW DO WN

0 1


Figure 10.33

Since each value o lif coefficient corresponds to a particular value o dynamic pressure required to support an aeroplane in level flight, flight, trim airspeed can be correlate correlated d with elevator elevator deflection as in the graph o Figure 10.33. I the CG location is ahead o the neutral point and control position is directly related to surace deflection, the aeroplane will give evidence o stick position stability . In other words, the aeroplane will require the stick to be moved af to increase the angle o attack and trim at a lower airspeed and to be moved orward to decrease the angle o attack and trim at a higher airspeed. It is highly desirable to have an aeroplane aeroplane demonstrate this eature. eature. I the aeroplane were to to have stick position instability, the aeroplane would require the stick to be moved af to trim at a higher airspeed or to be moved orward to trim at a lower airspeed.

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10

Stability and Control There is an increment o orce dependent on the trim tab setting which varies with the dynamic pressure or the square o equivalent airspeed. Figure 10.34 indicates the variation o stick orce with airspeed and illustrates the effect o tab setting on stick orce.

EFFECT OF TRIM TAB SETTING

PULL

EAS EA S

0

CG @ 20% MAC 1 0

1

2

3

PUSH


Figure 10.34

In order to trim the aeroplane at point (1) a certain amount o up elevator is required and zero stick orce is obtained with the use o the trim tab. To trim the aeroplane or higher speeds corresponding to points (2) and (3), ( 3), less and less aircraf nose-up tab is required. Note that when the aeroplane is properly trimmed, a push orce is required to increase airspeed and a pull orce is required to decrease decrease airspeed. In this manner, manner, the aeroplane would have positive stick orce stability with a stable “eel” or airspeed.

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10

EFFECT OF CG POSIT POSIT ION

CG POSITION 10% MAC

PULL

TRIM SPEED

20% MAC

30% MAC

0

EA S

40% MAC

50% MAC

PUSH

0 1

Figure 10.35


I the CG o the aeroplane were varied while maintaining trim at a constant airspeed, the effect o CG position on stick orce stability could be appreciated. appreciated. As illustrated in Figure 10.35, moving the CG af decreases the slope o the line o stick orce orce through the trim speed. Thus, on decreasing stick-orce stability it is evident that smaller stick orces are necessary to displace the aeroplane aeroplane rom the trim speed. When the stick orce orce gradient gradient (or slope) becomes zero zero,, the CG is at the neutral neutral point and neutral stability exists. I the CG is af o the neutral point, stick orce instability will exist, e.g. the aeroplane will require a push orce at a lower speed or a pull orce at a higher speed. It should be noted that the stick orce gradient is low at low airspeeds, and when the aeroplane is at low speeds and high power and has a CG position near the af limit, the “eel” or airspeed will be weak.

EFFECT OF CONTROL SYSTEM FRICTION

PULL

TRIM SPEED BAND

EA S

0

PUSH

FRICTION FORCE BAND

Figure 10.36

Control system riction can create very undesirable effects on control orces. Figure 10.36 illustrates that that the control orce versus versus airspeed relationship is a band rather than than a line. A wide riction orce band can completely mask the stick orce stability when the stick orce stability is low. Modern flight control systems require require precise maintenance maintenance to minimize minimize the riction orce band and preserve proper eel to the aeroplane.

273

10

Stability and Control Manoeuvre Stability When the pilot pitches the aircraf, it rotates about the CG and the tailplane is subject to a pitching velocity, in this example, example, downwards. downwards. Due to the pitching pitching velocity in manoeuvring flight, the longitudinal stability o the aeroplane is slightly greater than in steady flight conditions.

CHAN GE IN TA TA IL LIFT

1 0


TAS

RELATIVE AIRFLOW FROM ANGULAR ROTATION PITCHING VELOCITY

INCREA SE IN TAIL ANGLE OF ATTACK DUE TO PITCHING VELOCITY

Figure 10.3 10.37 7 Aerodynamic Aerodynamic damping damping

shows that the tailplane experiences an upwards component o airflow due to Figure 10.37 shows its downwards pitching velocity. The vector addition o this this vertical component to to the TAS TAS provides an increase in effective angle o attack o the tail, which creates an increase in tail lif, 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, or a given pitching velocity, decreases the angle o attack due to pitching velocity.

274


10

MANOEUVRE MARGIN

L

L x

y Lt Lt

AC

AC

FW D CG LIMIT NEUTRAL POINT

HIGH

0 1

STICK FORCE


A FT CG LIMIT

MANOEU MANOE UV RE POINT

LOW STICK FORCE

Figure 10.38 Manoeuvre point

The pitching moment rom 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 urther af than or 1g flight, as shown in Figure 10.38. In most cases the manoeuvre point will not be a critical item; i the aeroplane demonstrates static stability in 1g flight, it will definitely have stability in manoeuvring flight.

Stick Force Per ‘g’ The most direct appreciation o the manoeuvring stability o an aeroplane is obtained rom a plot o stick orce versus load actor such as shown in Figure 10.39. The aeroplane with positive manoeuvring stability should demonstrat d emonstrate e a steady increase in stick orce with increase in load actor or “g”. The manoeuvring stick orce gradient - or stick stick orce per “g” - must be positive but should be o the proper magnitude. The stick orce gradient gradient must not be excessively high or the aeroplane will be difficult and tiring to manoeuvre. Also, the stick orce gradient gradient must not be too low or the aeroplane may be overstressed inadvertently when light control orces exist. INCREASING ALTITUDE AT A CONSTANT IAS DECREASES AERODYNAMIC DAMPING

275

10


30 MANOEUVRING STICK

20

FORCE GRADIENT

10

1

2

3

4 5 6 LOAD FACTOR, (n) or (g)

7

8

1 0


CG POSITION % MAC LOW LO W ALTITUDE 10 20

HIGH ALTITUDE

30

40

LOAD FACTOR

LOAD FACTOR

Figure 10.39

When the aeroplane has high static stability, the manoeuvring stability will be high and a high stick orce gradient will result, Figure 10.39. A possibility exists that the orward orward CG limit could be set to prevent prevent an excessively high manoeuvring stick orce orce gradient. As the CG moves af, the stick orce orce gradient gradient decreases decreases with decreasing decreasing manoeuvring stability and the lower lower limit o stick orce gradient may be reached. When asked asked to calculate ‘stick orce orce per g’, remember that the aircraf aircraf is at 1g to start with. So 1g must be subtracted rom the ‘g’ limit beore dividing by the pull orce. The pitch damping o the aeroplane is related related to air density. At high altitudes, the high TAS TAS reduces the change in tail angle o attack or a given pitching velocity and reduces the pitch damping. Thus, a decrease in manoeuvring stick orce stability can be expected expected with increased altitude.

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10

Tailoring Control Forces Control orces should reflect the stability o the aeroplane but, at the same time, should be o a tolerable magnitude. magnitude. A manual flying control system may may employ an infinite variety o techniques to to provide satisactory control control orces throughout the speed, CG CG and altitude range o the aircraf.

EFFECT OF ST ST ICK CENTR ING SPR SPRING ING

STICK CENTRING SPRING

0 1


Figure 10.40

Stick Centring Spring I a spring is added to the control system as shown in Figure 10.40, it will tend to centre the stick and provide a orce increment increment depending on stick displacement. When the control system has a fixed gearing between stick position and surace deflection, the centring centring spring spring will provide a contribution to to stick orce stability according to stick stick position. The contribution to stick orce stability will be largest at low flight speeds where relatively relatively large control deflections are are required. The contribution will be smallest at high airspeed because o the smaller control control deflections required. Thus, the stick centring spring spring will increase the airspeed and manoeuvring stick orce stability, but the contribution decreases at high airspeeds. A variation o this device would be a spring stiffness which would be controlled to vary with dynamic pressure (Q - Feel). In that case, the contribution o o the spring to stick orce stability would not diminish with speed.

Down Spring A down spring added to a control system is a means o increasing airspeed stick orce stability without a change in aeroplane static stability. As shown in Figure 10.4 10.41 1, a down spring consists o a long pre-loaded spring attached to the control system which tends tends to rotate rotate the elevators elevators down (aircraf nose-down). The effect o the down spring is to contribute an increment o pull orce independent o control deflection or airspeed.

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10


EFFECT OF DOW N SPRING

PRELOADED SPRING

Figure 10.4 10.41 1

When the down spring is added to the control system o an aeroplane and the aeroplane is re-trimmed or the original speed, the airspeed stick orce gradient is increased and there is a stronger eel or airspeed. The down spring would provide a “synthetic” improvement improvement to to an aeroplane deficient deficient in airspeed stick orce orce stability. Since the orce orce increment increment rom the down spring is unaffected by stick s tick position or normal acceleration, the manoeuvring stick orce stability would be unchanged.

1 0


Bobweight The bobweight is an effective device or improving stick orce orce stability. As shown in Figure 10.42, the bobweight consists o an eccentric mass attached to the control system which, in unaccelerated unaccelerat ed flight, contributes an increment increment o pull orce identical identical to the down spring. spring. In act, a bobweight added to the control system o an aeroplane produces an effect id entical to the down spring. The bobweight will increase the airspeed stick orce gradient and increase the eel or airspeed. The bobweight also has an effect on the manoeuvring stick orce gradient since the bobweight mass is subjected to the same acceleration acceleration as the aeroplane. Thus, the bobweight will provide an increment o stick orce in direct proportion propor tion to the manoeuvring acceleration o the aeroplane (load actor applied). This will prevent the the pilot applying too much ‘g’ during manoeuvres; manoeuvres; the more you pull back, the more resistance the bobweight adds to the control system. EFFECT OF BOBWE IGHT

BOBW BOB W EIGHT

Figure 10.42

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10

Longitudinal Control To be satisactory, an aeroplane must have adequate controllability as well as adequate stability. An aeroplane with high static longitudinal stability will exhibit great resistance to to displacement rom equilibrium. Hence, the most critical conditions conditions o controllability will occur when the aeroplane has high static stability, i.e. the lower limits o controllability will set the upper limits o static stability . (Fwd. CG limit). There are three principal conditions o flight which provide the critical requirements o longitudinal control power power (manoeuvring, take-off take-off and landing). Any one or combination combination o these conditions can determine d etermine the overall longitudinal control power and set a limit to the orward CG position.

10% MA MA C

UP

0 1

18% MAC

MAXIMUM DEFLECTION


MOST FORWARD CG FOR MA NOEUV NOEUVRING RING CONTRO LLABILITY 20% MAC 30% MAC

CL DOW DO WN

C LMAX

CG POSITION

Figure 10.43

Manoeuvring Control Requirement The aeroplane should have sufficient longitudinal control power to attain the maximum usable lif coefficient or the limit load actor during manoeuvres. As shown in Figure 10.43, orward movement o the CG increases the longitudinal stability o an aeroplane and requires larger control deflections to produce produce changes in trim lif coefficient. coefficient. For the example shown, the maximum effective deflection o the elevator is not capable o trimming the aeroplane at C LMAX or CG positions ahead o 18 percent MAC. This particular control requirement can be most critical or an aeroplane in supersonic flight. Supersonic flight is usually accompanied by large increases in static longitudinal stability (due to af CP movement) movement) and a reduction in the effectiveness o control suraces. In order to to cope with these trends, powerul all-moving suraces must be used to attain limit load actor or maximum usable CL in supersonic flight. This requirement is so important that once satisfied, the supersonic configuration usually has sufficient longitudinal control power or all other conditions o flight.

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10

Stability and Control Take-off Control Requirement At take-off, the aeroplane must have sufficient elevator control power to assume the take-off attitude prior to reaching take-off speed.

LIFT TAIL LOAD

ROLLING FRICTION

1 0

WEIGHT


Figure 10.44

Figure 10.44 illustrates the principal orces acting on an aeroplane aeroplane during take-off take-off roll. When

the aeroplane is in the three point attitude at some speed less than the stall speed, the wing lif will be less than the weight o the aeroplane. aeroplane. As the elevators elevators must be capable o rotating the aeroplane to to the take-off take-off attitude, the critical condition condition will be with zero zero load on the nose wheel and the net o lif and weight supported on the main gear. Rolling riction resulting rom the normal orce on the main gear creates an adverse nose-down moment. Also, the CG ahead o the main gear contributes a nose-down moment. To balance these two nose-down moments, the horizontal tail must be capable o producing a nose-up moment big enough to attain the take-off 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 efficiency o the surace. The jet aeroplane does not experience a similar magnitude o this effect since the induced velocities rom the jet are relatively small compared to the slipstream velocities rom a propeller.

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10

Landing Control Requirement At landing, the aeroplane must have sufficient control power to ensure adequate control at specified landing speeds. The most critical requirement requirement will exist when the CG CG is in the most orward position, flaps are ully extended, and power is set at idle. idle. This configuration will provide the most stable condition which is most demanding o controllability. The landing control requirement has one particular difference rom the manoeuvring control requirement o ree ree flight. As the aeroplane approaches the the surace, there will be a change in the three-dimensional flow over the aeroplane due to to ground ground effect. effect. A wing in proximity proximity to to the ground plane will experience a decrease in tip vor tices and downwash at a given lif coefficient. The decrease in downwash at the tail tends to increase the static stability and produce a nosedown moment rom the reduction reduction in down load on the tail. Thus, the aeroplane just off the runway surace, Figure 10.45, will require additional control deflection to trim at a given lif coefficient, and the landing control requirement may be critical in the design o longitudinal control power. 0 1


REDUCED DOW DOW NWASH DUE TO GROUND EFFECT EFFECT

Figure 10.45

As an example o ground effect, a typical typi cal propeller powered aeroplane may require as much as 15° more up elevator to trim at C LMAX in ground effect than in ree flight. In some cases the effectiveness o the elevator is adversely affected by the use o trim tabs. I trim is used to excess in trimming stick orces, the effectiveness o the elevator may be reduced which would hinder landing or take-off control.

Each o the three principal conditions requiring adequate longitudinal control are critical or high static stability. I the orward CG limit is exceeded, exceeded, the aeroplane may may encounter encounter a deficiency o controllability in any o these conditions. The orward CG limit is set by the minimum permissible controllability. The af CG limit is set by the minimum permissible stability.

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10

Stability and Control Dynamic Stability While static stability is concerned with the initial tendency o an aircraf to return to equilibrium, dynamic stability is defined by the resulting motion with time . I an aircraf is disturbed rom equilibrium, the time history o the resulting motion indicates its dynamic stability. In general, an aircraf will demonstrate positive dynamic stability i the amplitude o motion decreases with time. The various conditions o possible dynamic behaviour are illustrated in the ollowing six history diagrams. The nonoscillatory modes shown in diagrams A, B and C depict the time histories possible without cyclic motion. Initial Disturbance

A

SUBSIDENCE (or Dead Beat Return)

1 0


TIME

(Positive Static) (Positive Dynamic) Figure 10.46 Chart A

Chart A illustrates a system which is given an initial disturbance and the motion simp ly subsides without oscillation; the mode is termed “subsidence” or “dead beat 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.

B

DIVERGENCE

TIME (Negative Static) (Negative Dynamic)

Figure 10.47 Chart B

Chart B illustrates the mode o “divergence” by a non-cyclic increase o amplitude with time. The initial tendency to continue in the displacement direction is evidence o static instability and the increasing amplitude is proo o dynamic instability .

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10

C NEUTRAL STATIC STABILITY (Neutral Static) (Neutral Dynamic) TIME

Figure 10.48 Chart C 0 1

Chart C illustrates the mode o pure neutral stability. I the original disturbance creates a displacement which then remains constant, the lack o tendency or 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 eature common to each o 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.

D DAMPED OSCILLATION

T IM (Positive Static) (Positive Dynamic)

Figure 10.49 Chart D

Chart D illustrates the mode o a damped oscillation where the amplitude decreases with time. The reduction o amplitude with time indicates there is resistance to motion and that energy is being dissipated . Dissipation o energy or damping is necessary to provide positive dynamic stability.

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10


E UNDAMPED OSCILLATION

(Positive Static) (Neutral Dynamic) TIME

1 0

Figure 10.50 Chart E


I there is no damping in the system, the mode o chart E is the result, an undamped oscillation. Without damping, the oscillation continues with no reduction o 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 aircraf must have sufficient damping to rapidly dissipate any oscillatory motion which would affect the sae operation o the aircraf. When natural aerodynamic damping cannot be obtained, artificial damping must be provided to give the necessary positive dynamic stability.

F DIVERGENT OSCILLATION (Positive Static) (Negative Dynamic)

TIME

Figure 10.51 Chart F

Chart F illustrates the mode o 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.

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10

Divergent oscillation results when energy is supplied to the motion rather than dissipated by positive damping. An example o divergent oscillation occurs i a pilot unknowingly makes control inputs which are near the natural requency o the aeroplane in pitch; energy is added to the system, negative damping exists, and Pilot Induced Oscillation (PIO) results. The existence o static stability does not guarantee the existence o dynamic stability. However, the existence o dynamic stability implies the existence o static stability. IF AN AIRCRAFT IS STATICALLY UNSTABLE, IT CANNOT BE DYNAMICALLY STABLE Any aircraf must demonstrate the required degrees o static and dynamic stability. I the aircraf were allowed to have static instability with a rapid rate o divergence, it would be very difficult, i not impossible to fly. In addition, positive dynamic stability is mandatory in certain areas to prevent objectionable continued oscillations o the aircraf. 0 1


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10

Stability and Control Longitudinal Dynamic Stability The considerations o longitudinal dynamic stability are concerned with the time history response o the aeroplane to disturbances, i.e. the variation o displacement amplitude with time ollowing a disturbance. From previous definition: • dynamic stability will exist when the amplitude o motion decreases with time, and • dynamic instability will exist i the amplitude increases with time. An aeroplane must demonstrate positive dynamic stability or the major longitudinal motions. In addition, the aeroplane must demonstrate a certain degree o longitudinal stability by reducing the amplitude o motion at a certain rate. The required degree o dynamic stability is usually specified by the time necessary or the amplitude to reduce to one-hal the original value: the time to damp to hal-amplitude.

1 0

The aeroplane in ree flight has six degrees o reedom: rotation in roll, pitch, and yaw and translation in the horizontal, vertical and lateral directions. In the case o longitudinal dynamic stability, the degrees o reedom can be limited to pitch rotation, plus vertical and horizontal translation.


Since the aeroplane is usually symmetrical rom l ef to right, there will be no need to consider coupling between longitudinal and lateral / directional motions.

Thus, the principal variables in the longitudinal motion o an aeroplane will be: • The pitch attitude o the aeroplane. • The angle o attack (which will differ rom the pitch attitude by the inclination o the flight path). • True airspeed (TAS) The longitudinal dynamic stability o an aeroplane generally consists o two basic modes o oscillation:• long period oscillation (phugoid) • short period motion While the longitudinal motion o the aeroplane may consist o a combination o these modes, the characteristics o each mode are sufficiently distinct that each oscillatory tendency may be studied separately.

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10

Long Period Oscillation (Phugoid) The first mode o dynamic longitudinal stability consists o a long period oscillation reerred to as the phugoid. The phugoid or long period oscillation involves noticeable variations in: • • • •

pitch attitude, altitude and airspeed, but nearly constant angle o attack (not much change in load actor).

The phugoid is a gradual interchange o potential and kinetic energy about some equilibrium airspeed and altitude. Figure 10.52 illustrates the characteristic motion o the phugoid.

ANGLE OF ATTACK AT EACH

0 1

INSTANT ALONG THE FLIGHT PATH IS ESSENTIALLY


CONSTANT

LONG PERIOD

TIME

Figure 10.52 Long period oscillation (phugoid)

The period o oscillation in the phugoid is between 1 and 2 minutes. Since the pitch rate is quite low and only negligible changes in angle o attack take place, damping o the phugoid is weak. However, such weak damping does not necessarily have any great consequence. Since the period o oscillation is so great, long period oscillation is easily controlled by the pilot . Due to the nature o the phugoid, it is not necessary to make any specific aerodynamic provisions to counteract it.

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10

Stability and Control Short Period Oscillation The second mode o dynamic longitudinal stability is the short period oscillation. Short period oscillation involves significant changes in angle o attack (load actor), with approximately constant speed, height and pitch attitude; it consists o rapid pitch oscillations during which the aeroplane is constantly being restored towards equilibrium by its static stabilit y and the amplitude o the short period oscillations being decreased by pitch damping .

MOTION OCCURS AT ESSE NTIA LLY CONSTA NT SPEED

TIME TO DAMP TO HALF AMPLITUDE

1 0


TIME

SHORT PERIOD

Figure 10.53 Short period oscillation

Short period oscillation at high dynamic pressures with large changes in angle o attack could produce severe ‘g’ loads (large changes in load actor). Shown in Figure 10.53, 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 orcibly damp an oscillation may actually reinorce the oscillation (PIO) and produce instability. Short period oscillation is not easily controlled by the pilot.

I short period oscillation occurs, release the controls; the aeroplane is designed to d emonstrate 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 reinorce the oscillation to produce ailing flight loads. Modern, large high speed jet transport aircraf are fitted with pitch dampers, which automatically compensate or any dynamic longitudinal instability.

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10

O the two modes o dynamic longitudinal stability, the short period oscillation is o 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 o the oscillations are decreased by pitch damping, so the problems o dynamic stability can become acute under the conditions o flight where reduced aerodynamic damping occurs. High altitude, and consequently low density (high TAS), reduces aerodynamic damping, as detailed on page 274. DYNAMIC STABILITY IS REDUCED AT HIGH ALTITUDE DUE TO REDUCED AERODYNAMIC DAMPING

0 1


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10

Stability and Control Directional Stability and Control The directional stability o 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 o static directional stability would be the development o yawing moments which tend to restore the aeroplane to equilibrium.

Definitions The axis system o 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 orm so that static stability can be evaluated independent o weight, altitude, speed, etc. The yawing moment, N, is defined in the coefficient orm by the ollowing equation: N = Cn Q S b

1 0

or


Cn =

N Q S b

where: N

=

yawing moment

Q

=

dynamic pressure

S

=

wing area

b

=

wingspan

Cn

=

yawing moment coefficient (positive to the right)

The yawing moment coefficient, C n, is based on the wing dimensions S and b as the wing is the characteristic surace o the aeroplane.

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10

Sideslip Angle The sideslip angle relates the displacement o the aeroplane centre line rom the relative airflow. Sideslip angle is provided the symbol β (beta) and is positive when the relative wind is displaced to the right o the aeroplane centre line. Figure 10.54 illustrates the definition o sideslip angle.

RELATIVE AIRFLOW

A YAW TO THE LEFT GIVES A SIDESLIP TO THE RIGHT SIDESLIP ANGLE

0 1


N, YAW ING MOMENT

Figure 10.54 Sideslip angle ( β )

The sideslip angle, β, is essentially the “directional angle o attack” o the aeroplane and is the primary reerence in directional stability as well as lateral stability considerations. Static directional stability o the aeroplane is appreciated by response to sideslip.

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10


YAW ING MOMENT COEFFICIENT, C n STABLE

Cn

NEUTRAL

SIDESLIP ANGLE,

UNSTABLE

1 0


Cn

Figure 10.55

Static Directional Stability Static directional stability can be illustrated by a graph o yawing moment coefficient, C n, versus sideslip angle, β, such as shown in Figure 10.55. When the aeroplane is subject to a positive sideslip angle, static directional stability will be evident i a positive yawing moment coefficient results. Thus, when the relative airflow comes rom the right (+ β ) 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 o C n versus β has a positive slope, and the degree o stability will be a unction o the slope o this curve. I the curve has zero slope, there is no tendency to return to equilibrium, and neutral static directional stability exists. When the curve o Cn versus β has a negative slope, the yawing moments developed by sideslip tend to diverge rather than restore, and static directional instability exists.

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10

NEUTRAL

Cn

STABLE

UNSTABLE

Figure 10.56 0 1

Figure 10.56 illustrates the act that the instantaneous slope o the curve o C n versus β will


describe the static directional stability o the aeroplane.

• At small angles o sideslip, a strong positive slope depicts strong directional stability. • Large angles o sideslip produce zero slope and neutral stability. • At very high sideslip, the negative slope o the curve indicates directional instability. This decay o directional stability with increased sideslip is not an unusual condition. However, directional instability should not occur at the angles o sideslip o ordinary flight conditions. Static directional stability must be in evidence or all the critical conditions o flight. Generally, good directional stability is a undamental quality directly affecting the pilots’ impression o an aeroplane.

Contribution of the Aeroplane Components. Because the contribution o each component depends upon and is related to the others, it is necessary to study each separately.

Fuselage The uselage is destabilizing, Figure 10.57 . It is an aerodynamic body and a condition o sideslip can be likened to an “angle o attack”, so that an aerodynamic side orce is created. This side orce acts through the uselage aerodynamic centre (AC), which is close to the quarter-length point. I this aerodynamic centre is ahead o aircraf centre o gravity, as is usually the case, the effect is destabilizing.

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10


FUSELAGE (Plan View)

FORCE

AC

FLIGHT PATH

UNSTABLE YAWING MOMENT 1 0

Figure 10.57


Dorsal and Ventral Fins To overcome the instability in the uselage it is possible to incorporate into the overall design dorsal or ventral fins. A dorsal fin is a small aerooil, o very low aspect ratio, mounted on top o the uselage near the rear. A ventral fin is mounted below. Such fins are shown in Figure 10.58.

DORSAL FIN

FUSELAGE (Side View)

VENTRAL FIN

Figure 10.58

I the aircraf is yawed to the right, the dorsal and ventral fins will create a side orce to the right. The line o action o this orce is well af o the aircraf CG, giving a yawing moment to the lef (a stabilizing effect). However, at small angles o yaw they are ineffective.

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10

The side orce created by dorsal and ventral fins at small sideslip angles will be very small because: • the dorsal and ventral fins are at a low angle o attack, • they have a small surace area, and • their aspect ratio is very low, resulting in a small lif curve slope. Figure 10.59.

CF

HIGH ASPECT RATIO

LOW ASPECT RATIO (or sweepback) 0 1


Figure 10.59

When fitted with dorsal and ventral fins, a uselage 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 aircraf is likely to experience in service. The effectiveness o dorsal and ventral fins increases with increasing sideslip angle, so that the combination o a uselage 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 destabilizing in this mode, as will be demonstrated later. For this reason, the dorsal fin is much more common.

DORSAL FIN

Figure 10.60

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10


RELATIVE AIRFLOW

TAIL MOMENT AR M CHANGE IN FIN LIFT

1 0

Figure 10.61


Fin The fin (vertical stabilizer) is the major source o directional stability or the aeroplane. As shown in Figure 10.61, in a sideslip the fin will experience a change in angle o attack. The change in lif (side orce) on the fin creates a yawing moment about the centre o gravity which tends to yaw the aeroplane into the relative airflow. The magnitude o the fin contribution to static directional stability depends on both the change in fin lif and the fin moment arm. Clearly, the fin moment arm is a powerul actor. The contribution o the fin to directional stability depends on its ability to produce changes in lif, or side orce, or a given change in sideslip angle. The contribution o the fin is a direct unction o its area. The required directional stability may be obtained by increasing the fin area. However, increased surace area has the obvious disadvantage o increased parasite drag. The lif curve slope o the fin determines how sensitive the surace is to change in angle o attack. While it is desirable to have a high lif curve slope or the fin, a high aspect ratio surace is not necessarily practical or desirable - bending, lower stalling angle ( Figure 10.59), hangar roo clearance, etc. The stall angle o the surace must be sufficiently great to prevent stall and subsequent loss o effectiveness at expected sideslip angles. (Sweepback or low aspect ratio increases the stalling angle o attack o the fin). The flow field in which the fin operates is affected by other components o the aeroplane as well as power effects. The dynamic pressure at the fin could depend on the slipstream o a propeller or the boundary layer o the uselage. Also, the local flow direction at the fin is influenced by the wing wake, uselage crossflow, induced flow o the horizontal tail or the direction o slipstream rom a propeller. Each o these actors must be considered as possibly affecting the contribution o the fin to directional stability. A high mounted tailplane (‘T’ - tail) makes the fin more effective by acting as an “end plate”.

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The side orce on the fin may still be relatively small compared to that on the uselage, which is destabilizing, but because its line o ac tion is ar af o the CG, the yawing moment it creates is relatively large and gives overall stability to the uselage-fin combination. The principle behind the effect o the fin as a stabilizer is just the same as in the case o 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 o sideslip. It remains effective until the angle o sideslip is such that the fin angle o attack approaches its stalling angle, but above this value, the side orce on the fin decreases with increasing sideslip angle, and the fin ceases to be effective as a stabilizer. It is at this point that the dorsal or ventral fin becomes important. Because it stalls at a very much higher angle o attack, it takes over the stabilizing role o the fin at large angles o sideslip.

Wing and Nacelles The contribution o the wing to static directional stability is usually small: • The contribution o a straight wing alone is usually negligible. • Sweepback produces a stabilizing effect , which increases with increase in C L (i.e. at lower IAS).

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• Engine nacelles on the wings produce a contribution that will depend on such actors as their size and position and the shape o the wing planorm. On a straight wing, they usually produce a destabilizing effect. A swept wing provides a stable contribution depending on the am ount o sweepback, but the contribution is relatively weak when compared with other components. Consider a sideslipping swept wing, as illustrated in Figure 10.62.

V

V

NORMAL

V

NORMAL

Figure 10.62

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Stability and Control The inclination o the orward, right, wing to the relative airflow is greater than that o the rearward wing, so there is more lif and, hence, more induced drag, on the right side, (the influence o increased lif on the orward wing will be explained when lateral static stability is considered). The result o this discrepancy in drag on the two sides o the wing is a yawing moment to the right, which tends to eliminate the sideslip. This is a stabilizing effect, and may be important i the sweepback angle is quite large. Figure 10.63 illustrates a typical build-up o the directional stability o an aeroplane by separating the contribution o the uselage and fin. As shown by the graph o C n versus β, the

contribution o the uselage is destabilizing, but the instability decreases at large sideslip angles. The contribution o the fin alone is highly stabilizing up to the point where the surace begins to stall. The contribution o the fin must be large enough so that the complete aeroplane (wing-uselage-fin combination) exhibits the required degree o stability .

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AEROPLANE W ITH DORSAL FIN STALL

Cn


ADDED

FIN ALONE

COMPLETE AEROPLANE

FUSELAGE ALONE

Figure 10.63

The dorsal fin has a powerul effect on preserving the directional stability at large angles o sideslip which would produce stall o the fin. The addition o a dorsal fin to the aeroplane will reduce the decay o directional stability at high sideslip in two ways: • The least obvious but most important effect is a large increase in the uselage stability at large sideslip angles. • In addition, the effective aspect ratio o the fin is reduced which increases the stall angle or the surace. By this twoold effect, the addition o the dorsal fin is a very useul device. The decreased lif curve slope o a swept-back fin will also decrease the tendency or the fin to stall at high sideslip angles.

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Power Effect The effects o power on static directional stability are similar to the power effects on static longitudinal stability. The direct effect is confined to the normal orce at the propeller plane and, o course, is destabilizing when the propeller is located ahead o the CG. In addition, the air in the slipstream behind a propeller spirals around the uselage, and this results in a sidewash at the fin (rom the lef with a clockwise rotating propeller). The indirect effects o power induced velocities and flow direction changes at the fin (spiral slipstream effect) are quite significant or the propeller driven aeroplane and can produce large directional trim changes. As in the longitudinal case, the indirect effects are negligible or the jet powered aeroplane. The contribution o the direct and indirect power effects to static directional stability is greatest or the propeller powered aeroplane and usually slight or the jet powered aeroplane. In either case, the general effect o power is destabilizing and the greatest contribution will occur at high power and low dynamic pressure .

Critical Conditions

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The most critical conditions o static directional stability are usually the combination o several separate effects. The combination which produces the most critical condition is much dependent upon the type o aeroplane. In addition, there exists a coupling o lateral and directional


effects such that the required degree o static directional stability may be determined by some o these coupled conditions .

Centre of Gravity Position Centre o gravity position has a relatively negligible effect on static directional stability. The usual range o CG position on any aeroplane is set by the limits o longitudinal stability and control. Within this limiting range o CG position, no significant changes take place in the contribution o the vertical tail, uselage, nacelles, etc. Hence, static directional stability is essentially unaffected by the variation o CG position within the longitudinal limits.

LOW ANGLE OF ATTACK

HIGH ANGLE OF ATTACK

SIDESLIP ANGLE,

Figure 10.64

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Stability and Control High Angle of Attack When the aeroplane is at a high angle o attack a decrease in static directional stability can be anticipated. As shown by Figure 10.64, a high angle o attack reduces the stable slope o the curve o Cn versus β. The decrease in static directional stability is due in great part to the reduction in the contribution o the fin. At high angles o attack, the effectiveness o the fin is reduced because o increase in the uselage boundary layer at the fin location. The decay o directional stability with angle o attack is most significant or an aeroplane with sweepback since this configuration requires a high angle o attack to achieve high lif coefficients.

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Figure 10.65

Ventral Fin Ventral fins may be added as an additional contribution to directional stability, Figure 10.65. Landing clearance requirements may limit their size, require them to be retractable or require two smaller ventral fins to be fitted instead o one large one. The most critical demands o static directional stability will occur rom some combination o the ollowing effects: • • • •

high angle o sideslip high power at low airspeed high angle o attack 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 o high Mach numbers may produce the critical conditions or the typical transonic, jet powered aeroplane. In addition, the coupling o lateral and directional effects may require prescribed degrees o directional stability.

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Lateral Stability and Control The static lateral stability o an aeroplane involves consideration o rolling moments due to sideslip. I an aeroplane has avourable rolling moment due to a sideslip, a lateral displacement rom wing level flight produces a sideslip, and the sideslip creates a rolling moment tending to return the aeroplane to wing level flight. By this action, static lateral stability will be evident. O course, a sideslip will produce yawing moments depending on the nature o the static directional stability, but the consideration o static lateral stability will involve only the relationship o rolling moments and sideslip.

Definitions The axis system o 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 orm so that lateral stability can be evaluated independent o weight, altitude, speeds, etc. The rolling moment, L, is defined in the coefficient orm by the ollowing equation:

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L = Cl Q S b


or Cl =

L Q S b

where: L = rolling moment (positive to right) Q = dynamic pressure S = wing area b = wingspan Cl = rolling moment coefficient (positive to the right) The angle o sideslip, β, has been defined previously as the angle between the aeroplane centre line and the relative wind and is positive when the relative wind is to the right o the centre line.

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Stability and Control Static Lateral Stability Static lateral stability can be illustrated by a graph o rolling moment coefficient, C l, versus sideslip angle, β, such as shown in Figure 10.67 . When the aeroplane is subject to a positive sideslip angle, lateral stability will be evident i a negative rolling moment coefficient results. Thus, when the relative airflow comes rom the right (+ β), a rolling moment to the lef (-C l) should be created which tends to roll the aeroplane to the lef. Lateral stability will exist when the curve o Cl versus β has a negative slope and the degree o stability will be a unction o the slope o this curve. I the slope o the curve is zero, neutral lateral stability exists; i the slope is positive, lateral instability is present.

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RELATIVE A IRFLOW

l,

ROLLING MOMENT

Figure 10.66

ROLLING MOMENT COEFFICIENT

C

l UNSTABLE

NEUTRAL

SIDESLIP ANGLE, STABLE

Figure 10.67

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It is desirable to have static lateral stability (avourable roll due to sideslip), Figure 10.68. However, the required magnitude o lateral stability is determined by many actors. Excessive roll due to sideslip complicates crosswind take-off and landing and may lead to undesirable oscillatory coupling with the directional motion o the aeroplane. In addition, high lateral stability may combine with adverse yaw to hinder rolling perormance. Generally, good handling qualities are obtained with a relatively light, or weak positive, lateral stability.

STABL E ROLL DUE TO SIDESLIP

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NEUTRAL

UNSTABLE ROLL DUE TO SIDESLIP

Figure 10.68 Static lateral stability

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Stability and Control Contribution of the Aeroplane Components In order to appreciate the development o lateral stability in an aeroplane, each o the components which contribute must be inspected. There will be intererence between the components, which will alter the contribution to stability o each component on the aeroplane.

EFFECTIVE INCREASE IN LIFT DUE TO SIDESLIP

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EFFECTIVE DECREASE IN LIFT DUE TO SIDESLIP


Figure 10.69 Geometric dihedral

Wing The principal surace contributing to the lateral stability o an aeroplane is the wing. The effect o *geometric dihedral is a powerul contribution to lateral stability. As shown in Figure 10.69, a wing with geometric dihedral will develop stable rolling moments with sideslip. I the relative wind comes rom the side, the wing into the wind is subject to an increase in angle o attack and develops an increase in lif. The wing away rom the wind is subject to a decrease in angle o attack and develops a decrease in lif. The changes in lif 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 powerul in producing lateral stability it is taken as a common denominator o the lateral stability contribution o all other components. Generally, the contribution o wing position, flaps, power, etc., is expressed as “ DIHEDRAL EFFECT”. *Geometric Dihedral: The angle between the plane o each wing and the horizontal, when the aircraf is unbanked and level; positive when the wing lies above the horizontal, as in Figure 10.69. Negative geometric dihedral is used on some aircraf, and is known as anhedral.

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Wing Position The contribution o the uselage alone is usually quite small; depending on the location o the resultant aerodynamic side orce on the uselage. However, the effect o the wing - uselage - tail combination is significant since the vertical placement o the wing on the uselage can greatly affect the combination. A wing located at the mid wing position will generally exhibit a “dihedral effect” no different rom that o the wing alone. Figure 10.70 illustrates the effect o wing position on static lateral stability.

• A low wing position gives an unstable contribution. The direction o relative airflow decreases the effective angle o attack o the wing into wind and increases the effective angle o attack o the wing out o wind - tending to increase the rolling moment. • A high wing location gives a stable contribution. The direction o relative airflow increases the effective angle o attack o the wing into wind and decreases the effective angle o attack o the wing out o wind, tending to decrease the rolling moment.

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SIDESLIP HIGH WING POSITION

LOW WING POSITION

Figure 10.70 Wing - uselage intererence effect

The magnitude o “dihedral effect” contributed by the vertical position o the wing is large and may require a noticeable dihedral angle or 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 Sweepback The contribution o sweepback to “dihedral effect” is important because o the nature o the contribution. As shown in Figure 10.71 and Figure 10.72, i the wing is at a positive lif coefficient, the wing into the wind has less sweep and an increase in lif, and the wing out o the wind has more sweep and a decrease in lif; 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-orward wing would give a negative dihedral effect).

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INCREASED

EFFECTIVE

SWEEP DECREASED EFFECTIVE SWEEP

Figure 10.71 The effect o sweepback

The contribution o sweepback to “dihedral effect” is proportional to the wing lif coefficient as well as the angle o sweepback. Because high speed flight requires a large amount o sweepback, an excessively high “dihedral effect” will be present at low speeds (high C L). An aircraf with a swept-back wing requires less geometric dihedral than a straight wing.

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LEAD ING W ING IN SIDESLIP DIFFERENCE IN

CL

LIFT ON THE TWO WINGS

ZERO SIDESLIP

TRA ILING W ING IN SIDESLIP

0

0 1

HIGH SPEED


LOW SPEED

Figure 10.72 Effect o speed on ‘Dihedral Effect’ o swept wing

The fin can provide a small “dihedral effect” contribution, Figure 10.73. I the fin is large, the side orce 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. SMALL STABILIZING ROLLING MOMENT IN SIDESLIP

RELATIVE AIRFLOW

Figure 10.73 Fin contribution

The ventral fin, being below the aircraf CG, has a negative influence on lateral static stability, as illustrated in Figure 10.74. SMALL DESTABILIZING ROLLING MOMENT IN SIDESLIP

VENTRAL FIN RELATIVE AIRFLOW

Figure 10.74 Ventral fin contribution

Generally, the “dihedral effect” should not be too great since high roll due to sideslip can create certain problems.

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Stability and Control Excessive “dihedral effect” can lead to “Dutch roll” difficult rudder coordination in rolling manoeuvres, or place extreme demands or lateral control power during crosswind take-off and landing. I the aeroplane demonstrates satisactory “dihedral effect” during cruise, certain exceptions can be tolerated when the aeroplane is in the take-off and landing configuration. Since the effects o flaps and power are destabilizing and reduce the “dihedral effect”, a certain amount o negative “dihedral effect” may be possible due to these sources.

REDUCED ARM

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Figure 10.75 Partial span flaps reduce lateral stability

The deflection o flaps causes the inboard sections o the wing to become relatively more effective and these sections have a small spanwise moment arm, Figure 10.75. Thereore, the changes in wing lif due to sideslip occur closer inboard and the dihedral effect is reduced. The effect o power on “dihedral effect” is negligible or the jet aeroplane but considerable or 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 p ower effects are combined, e.g. the propeller driven aeroplane in a power-on approach. With certain exceptions during the conditions o landing and take-off, 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 orces and displacements required to maintain sideslip. Positive stick orce stability will be evident by stick orces required in the direction o the controlled sideslip.

Conclusion The designer is aced with a dilemma. An aircraf is given sweepback to increase the speed at which it can operate, but a by-product o sweepback is static lateral stability. A swept-back wing requires much less geometric dihedral than a straight wing. I a requirement also exists or the wing to be mounted on top o the uselage, an additional “dihedral effect”is present. A high mounted and swept-back wing would give excessive “dihedral effect”, so anhedral is used to reduce “dihedral effect” to the required level.

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Lateral Dynamic Effects Previous discussion has separated the lateral and directional response o 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 o an aeroplane are: • Rolling moment due to sideslip, or “dihedral effect” (static lateral stability). • Yawing moment due to sideslip, or static directional stability.

Spiral Divergence Spiral divergence will exist when static directional stability is very larg e when compared to the “dihedral effect”.

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The character o spiral divergence is not violent. The aeroplane, when disturbed rom the equilibrium o level 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 o divergence in the spiral motion is usually so gradual that the pilot can control the tendency without difficulty.

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 o the oscillatory nature. When a yaw is introduced, the strong “dihedral effect” will roll the aircraf due to the lif increase on the wing into wind. The increased induced drag on the rising wing will yaw the aircraf in the opposite direction, reversing the coupled oscillations. Aircraf with a tendency to Dutch roll are fitted with a Yaw Damper. This automatically displaces the rudder proportional to the rate o yaw to damp-out the oscillations. I the Yaw Damper ails in flight, it is recommended that the ailerons be used by the pilot to damp-out Dutch roll. Because o the response lag, i the pilot uses the rudder, pilot induced oscillation (PIO) will result and the Dutch roll may very quickly become divergent, leading to loss o control.

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Stability and Control Dutch roll is objectionable, and spiral divergence is tolerable i the rate o divergence is low. For this reason, the “dihedral effect” should be no more than that required or satisactory lateral stability. I 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 o high static directional stabilit y and minimum necessary “dihedral effect”, most aeroplanes demonstrate a mild spiral tendency. As previously mentioned, a weak spiral tendency is o little concern to the pilot and certainly preerable to Dutch roll. The contribution o sweepback to the lateral dynamics o an aeroplane is significant. Since the “dihedral effect” rom sweepback is a unction o lif coefficient, the dynamic characteristics may vary throughout the flight speed range. When the swept wing aeroplane is at low C L the “dihedral effect” is small and the spiral tendency may be apparent. When the swept wing aeroplane is at high C L the “dihedral effect” is increased and the Dutch roll oscillatory tendency is increased.

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Pilot Induced Oscillation (PIO) Certain undesirable motions may occur due to inadvertent action on the controls. These can occur about any o the axes, but the most important condition exists with the short period longitudinal motion o 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 o producing damaging flight loads and loss o control o 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 urnish negative damping to the oscillatory motion, and dynamic instability will exist. Since short period motion is o relatively high requency, the amplitude o the pitching oscillation can reach dangerous proportions in an unbelievably shor t time. When pilot induced oscillation is encountered, the most effective solution is an immediate release o the controls. Any attempt to orcibly 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 o its inherent dynamic stability.

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High Mach Numbers Generally, flight at high Mach numbers will take place at high altitude, hence the effect o high altitude must be separated or study. Aerodynamic damping is due to moments created by pitching, rolling, or yawing o the aircraf. These moments are derived rom the changes in angles o attack o the tail, wing and fin suraces with angular rotation (see Figure 10.38). Higher TAS common to high altitude flight reduces the angle o attack changes and reduces aerodynamic damping. In act, aerodynamic damping is proportional to the square root o the relative density, similar to the proportion o True Airspeed to Equivalent Airspeed. Thus, at an ISA altitude o 40 000 f, aerodynamic damping would be reduced to one-hal the ISA sea level value.

Mach Trim As speed increases beyond the Critical Mach number (M CRIT), shock wave ormation at the root o a swept-back wing will:

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• reduce lif orward o the CG, and • reduce downwash at the tailplane. Together, these actors will generate a nose-down pitching moment. At high Mach numbers, an aircraf will become unstable with respect to speed; instead o an increasing push orce being required as speed increases, a pull orce becomes necessary to prevent the aircraf accelerating urther. This is potentially dangerous. A small increase in Mach number will give a nose-down pitch which will urther increase the Mach number. This in turn leads to a urther increase in the nose-down pitching moment. This unavourable high speed characteristic, known as “Mach Tuck”, “High Speed Tuck” or “Tuck Under” would restrict the maximum operating speed o a modern high speed jet transport aircraf. To maintain the required stick orce gradient at high Mach numbers, a Mach trim system must be fitted . This device, sensitive to Mach number, may:

• deflect the elevator up, • decrease the incidence o the variable incidence trimming tailplane, or • move the CG rearwards by transerring uel rom the wings to a rear trim tank. by an amount greater than that required merely to compensate or the trim change. This ensures the required stick orce gradient is maintained in the cruise at high Mach numbers. Whichever method o trim is used by a particular manuacturer, a Mach trim system will adjust longitudinal trim and operates only at high Mach numbers .

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Stability and Control Key Facts 2 Sel Study (Insert the missing words, with reerence to the preceding paragraphs). Positive static longitudinal stability is indicated by a _______ slope o C M versus CL. The degree o _______ longitudinal stability is indicated by the ______ o the curve. The net pitching moment about the ________ axis is due to the contribution o each o the component _________ acting in their appropriate _____ fields. In most cases, the contribution o the uselage and nacelles is ____________. (Page 259) Noticeable changes in static stability can occur at high C L (low speed) i:

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a)

the aeroplane has __________.

b)

there is a large contribution o ‘______ effect’.

c)

there are significant changes in __________ at the horizontal tail.

The horizontal tail usually provides the ________ stabilizing influence o all the components o the aeroplane. __________ decreases static longitudinal stability. I the thrust line is below the CG, a thrust increase will produce a _______ or nose ___ moment and the effect is ___________. High lif devices tend to _________ downwash at the tail and _____ the dynamic pressure at the tail, both o which are ___________. An increase in TAS, or a given pitching velocity, _________ aerodynamic damping. The aeroplane with positive manoeuvring stability should demonstrate a steady _______ in stick orce with ________ in load actor or “__”. The stick orce gradient must not be excessively ____ or the aeroplane will be difficult and tiring to manoeuvre. Also, the stick orce gradient must not be too _____ or the aeroplane may be overstressed inadvertently when light control orces exist. When the aeroplane has high static stability, the manoeuvring s tability will be _____ and a ____ stick orce gradient will result. The ________ CG limit could be set to prevent an excessively high manoeuvring stick orce gradient. As the CG moves af, the stick orce gradient _________ with ____________ manoeuvring stability and the _______ limit o stick orce gradient may be reached. At high altitudes, the high TAS ________ the change in tail angle o attack or a given pitching velocity and _________ the pitch damping. Thus, a decrease in manoeuvring stick orce stability can be expected with _________ altitude. A flying control system may employ _______ springs, _____ springs or ____ weights to provide satisactory control orces throughout the speed, CG and altitude range o an aircraf. While static stability is concerned with the initial tendency o an aircraf to return to equilibrium, dynamic stability is defined by the resulting _______ with _____.

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10

An aircraf will demonstrate positive dynamic stability i the _________ o motion ________ with time. When natural aerodynamic damping cannot be obtained, _________ damping must be provided to give the necessary positive dynamic stability. The longitudinal dynamic stability o an aeroplane generally consists o two basic modes o oscillation: a)

_____ period (phugoid)

b)

______ period

The phugoid oscillation occurs with nearly constant ______ o _______. The period o oscillation is so great, the pilot is easily able to counteract ____ _____ oscillation. Short period oscillation involves significant changes in ______ o _______.

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Short period oscillation is ____ _______ controlled by the pilot. The problems o dynamic stability can become acute at _____ altitude because o _________ aerodynamic ________. To overcome the directional instability in the uselage it is possible to incorporate into the overall design _______ or ________ fins. The _____ is the major source o directional stability or the aeroplane. A ___ - tail makes the fin more effective by acting as an “____ plate”. Because the _______ fin stalls at a very much higher angle o attack, it takes over the stabilizing role o the fin at large angles o sideslip. ___________ produces a directional stabilizing effect, which increases with increase in C L. _________ fins increase directional stability at _____ angles o attack. Landing clearance requirements may limit their size, require them to be retractable, or require two smaller ventral fins to be fitted instead o one large one. Generally, good handling qualities are obtained with a relatively _____, or ____ positive, lateral stability. The principal surace contributing to the lateral stability o an aeroplane is the _____. The effect o geometric _________ is a powerul contribution to lateral stability. A low wing position gives an ________ contribution to static lateral stability. A _____ wing location gives a stable contribution to static lateral stability. The magnitude o “dihedral effect” contributed by the vertical position o the wing is _____ and may require a noticeable dihedral angle or the _____ wing configuration. A high wing position, on the other hand, usually requires ___ geometric ________ at all. The ______ back wing contributes a positive “dihedral effect”.

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Stability and Control An aircraf with a swept-back wing requires _____ geometric dihedral than a straight wing. The fin contribution to purely lateral static stability, is usually very ______. Excessive “dihedral effect” can lead to “______ roll,” difficult rudder coordination in ________ manoeuvres, or place extreme demands or _______ control power during crosswind take-off and landing. Deploying partial span flaps gives a _________ dihedral effect. A swept-back wing requires much less geometric dihedral than a straight wing. I a requirement also exists or the wing to be mounted on top o the uselage, an additional “dihedral effect” is present. A high mounted and swept-back wing would give excessive “dihedral effect”, so _________ is used to reduce “dihedral effect” to the required level. 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.

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Spiral divergence will exist when static directional stability is very _____ when compared to the “dihedral effect”.


The rate o divergence in the spiral motion is usually so ________ that the pilot can control the tendency without _________. Dutch roll will occur when the “dihedral effect” is ______ when compared to static directional stability. Aircraf which Dutch roll are fitted with a _____ Damper. This automatically displaces the rudder proportional to the _____ o yaw to damp-out the oscillations. I the Yaw Damper ails in flight, it is recommended that the ________ be used by the pilot to damp-out Dutch roll. I the pilot uses the ________, pilot induced oscillation (PIO) will result and the Dutch roll may very quickly become _________, leading to loss o _______. When the swept wing aeroplane is at low C L the “dihedral effect” is small and the ______ tendency may be apparent. When the swept wing aeroplane is at high C L the “dihedral effect” is increased and the ______ _____ oscillatory tendency is increased. When pilot induced oscillation is encountered, the most effective solution is an immediate _______ o the controls. Any attempt to orcibly damp the oscillation simply _________ the excitation and _________ the oscillation. Higher TAS common to high altitude flight _______ the _____ o ______ changes and reduces aerodynamic ________. Mach Tuck is caused by ___ o lif in ront o the ___ and _______ downwash at the tail due to the ormation o a __________ on a swept-back wing at _____ Mach numbers. The Mach trim system will adjust ___________ _____ to maintain the required _____ _____ gradient and operates only at _____ Mach numbers. KEY FACTS 2 WITH THE MISSING WORDS INSERTED CAN BE FOUND ON page 326.

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Summary Self Study Stability is the inherent quality o an aircraf to correct or conditions that may disturb its

equilibrium and to return to, or continue on its original flight path. An aircraf can have two basic types o stability: static and dynamic, and three condition o each type: positive, neutral, and negative. Static stability describes the initial reaction o an aircraf afer it has been disturbed rom

equilibrium about one or more o its three axes. Positive static stability is the condition o stability in which restorative orces are set-up that

will tend to return an aircraf to its original condition a nytime it’s disturbed rom a condition o equilibrium. I an aircraf has an initial tendency to return to its original attitude o equilibrium, it has positive static stability . (statically stable). 0 1

An aircraf with neutral static stability produces orces that neither tend to return it to its original condition, nor cause it to depart urther rom this condition. I an aircraf tends to remain in its new, disturbed state, it has neutral static stability . (statically neutral).


I an aircraf has negative static stability , anytime it is disturbed rom a condition o equilibrium, orces are set up that will tend to cause it to depart urther rom its original condition. Negative static stability is a highly undesirable characteristic as it can cause loss o control. When an aircraf continues to diverge, it exhibits negative static stability . (statically unstable). Most aeroplanes have positive static stability in pitch and yaw, and are close to being neutrally statically stable in roll. When an aircraf exhibits positive static stability about any o its three axes, the term “ dynamic stability” describes the long term tendency o the aircraf. When an aircraf is disturbed rom 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 o oscillations. Positive dynamic stability is a condition in which the orces o static stability decrease with

time. Positive dynamic stability is desirable. I oscillations become smaller with time, an aircraf has positive dynamic stability . (dynamically stable). Neutral dynamic stability causes an aircraf to hunt back and orth around a condition o

equilibrium, with the corrections getting neither larger or smaller. (dynamically neutral). Neutral dynamic stability is undesirable. I an aircraf diverges urther away rom its original attitude with each oscillation, it has negative dynamic stability . Negative dynamic stability causes the orces o static stability to increase with time. (dynamically unstable). Negative dynamic stability is extremely undesirable. The overall design o an aircraf contributes to its stability (or lack o it) about each o its three axes o motion.

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Stability and Control The vertical stabilizer ( fin) is the primary source o directional stability (yaw). The horizontal stabilizer ( tailplane) is the primary source o longitudinal stability (pitch). The wing is the primary source o lateral stability (roll). CG location also affects stability.

I the CG is close to its af limit, an aircraf will be less stable in both pitch and yaw. As the CG is moves orward, stability increases. Even though an aeroplane will be less stable with an af CG, it will have some desirable aerodynamic characteristics due to reduced aerodynamic loading o the horizontal tail surace. This type o an aeroplane will have a slightly lower stall speed and will cruise aster or a given power setting. Manoeuvrability is the quality o an aircraf that permits it to be manoeuvred easily and to

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withstand the stresses imposed by those manoeuvres.


Controllability is the capability o an aircraf to respond to the pilot’s control, especially with

regard to flight path and attitude. An aircraf is longitudinally stable i it returns to a condition o level flight afer a disturbance in pitch, caused by either a gust or displacement o the elevator by the pilot. The location o the CG and the effectiveness o the tailplane determines the longitudinal stability, and thus the controllability o an aircraf. Increasing stability about any axis: • decreases manoeuvrability and controllability, and • increases stick (or pedal) orces. Phugoid oscillation is a long-period oscillation in which the pitch attitude, airspeed and

altitude vary, but the angle o attack remains relatively constant. It is a gradual interchange o potential and kinetic energy about some equilibrium airspeed and altitude. An aircraf experiencing longitudinal phugoid oscillation is demonstrating positive static stability, and it is easily controlled by the pilot . An aircraf will return towards wing level afer a wing drop i it has static lateral stability . The wing o most aircraf has a positive geometric dihedral angle (dihedral). This is the angle produced by the wing tips being higher than the wing root. I the lef wing drops in flight, an aircraf will momentarily begin to slip to the lef, and the effective angle o attack o the lef wing will increase and the effective angle o attack o the right wing will decrease. The change in angle o attack o 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 aircraf a higher MCRIT. An aircraf with a swept-back wing will not require as much geometrical dihedral as a straight wing. Some aircraf have the wing mounted on top o the uselage or various reasons. Also as a by-product, a high mounted wing will give a “dihedral effect” due to the direction o airflow around the uselage and wing during a sideslip. An aircraf with a high mounted wing does not require as much geometric dihedral.

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An aircraf 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 . Static directional stability is the tendency o the nose o an aircraf to yaw towards the relative

airflow. It is achieved by the keel surace behind the CG being larger than that in ront o the CG. A swept-back wing also provides a measure o static directional stability. Too much static directional stability could result in dynamic instability - Spiral Instability . Interaction between static lateral stability and static directional stability . I a wing drops

and the aircraf begins to slip to the side, directional stability will cause the nose to yaw into the relative airflow. 0 1

“Dihedral effect” tends to roll an aircraf when a wing drops, and directional stability causes the nose to yaw into the direction o the low wing.


These two orces interact (coupled motion): • An aircraf 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 aeroplane will begin to turn. The increased speed o the wing on the outside o the turn will increase the angle o bank, and the reduction in the vertical component o lif will orce the nose to a low pitch angle. This will cause the aircraf to enter a descending spiral. • An aircraf with strong “dihedral effect” and weak directional stability will have a tendency towards dutch roll instability. A Mach trim system maintains the required stick orce gradient at high Mach numbers by adjusting the longitudinal trim. The Mach trim system only operates at high Mach numbers.

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An aeroplane which is inherently stable will:

a. b. c. d. 2.

Afer a disturbance in pitch an aircraf oscillates in pitch with increasing amplitude. It is:

a. b. c. d.

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3.

Q u e s t i o n s

b. c. d.

changes in lif produce a wing pitching moment which acts to reduce the change o lif. changes in lif produce a wing pitching moment which acts to increase the change o lif. changes in lif give no change in wing pitching moment. when the aircraf sideslips the CG causes the nose to turn into the sideslip thus applying a restoring moment.

The longitudinal static stability o an aircraf:

a. b. c. d.

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anhedral. dihedral. sweepback. uselage and fin area.

I the wing AC is orward o the CG:

a.

7.

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 aircraf 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 ollowing the removal o any disturbing orce.

is reduced by the effects o wing downwash. is increased by the effects o wing downwash. is not affected by wing downwash. is reduced or nose-up displacements, but increased or nose-down displacements by the effects o wing downwash.

Questions 8.

To ensure some degree o longitudinal stability in flight, the position o the CG:

a. b. c. d. 9.

d.

s n o i t s e u Q

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.

will be increased because o the effective increase o dihedral. will be increased because o increased lif. will be reduced because the centre o lif o each semi-span is closer to the wing root. will not be affected.

Dihedral gives a stabilizing rolling moment by causing an increase in lif:

a. b. c. d. 14.

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With flaps lowered, lateral stability:

a. b. c.

13.

the CG and the AC. the AC and the neutral point. the CG and the neutral point. the CG and the CG datum point.

I a disturbing orce causes the aircraf to roll:

a. b. c. d. 12.

very small orces are required on the control column to produce pitch. longitudinal stability is reduced. very high stick orces are required to pitch because the aircraf is very stable. stick orces are the same as or an af CG.

The static margin is equal to the distance between:

a. b. c. d. 11.

must always coincide with the AC. must be orward o the Neutral Point. must be af o the Neutral Point. must not be orward o the af CG limit.

When the CG is close to the orward limit:

a. b. c. d. 10.

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on the up-going wing when the aircraf rolls. on the down-going wing when the aircraf rolls. on the lower wing i the aircraf is sideslipping. on the lower wing whenever the aircraf 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. c. d.

greater longitudinal stability. the same degree o 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 uselage when the aircraf is sideslipping increasing the effective angle o attack o the lower wing.

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Questions 15.

At a constant IAS, what effect will increasing altitude have on damping in roll?

a. b. c. d. 16.

Sweepback o the wings will:

a. b. c. d. 17.

b.

Q u e s t i o n s

c. d. 18.

increases rudder effectiveness. must be disengaged beore making a turn. augments stability. increases the rate o yaw.

A wing which is inclined downwards rom root to tip is said to have:

a. b. c. d.

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use roll inputs. use yaw inputs. move the CG. reduce speed below MMO.

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.

To correct Dutch roll on an aircraf with no automatic protection system:

a. b. c. d. 20.

increased downwash rom the wing will cause the elevators to be more responsive. due to the increased angle o attack o the wing the air will flow aster 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 o longitudinal stability.

Following a lateral disturbance, an aircraf with Dutch roll instability will:

a. b. c. d. 19.

not affect lateral stability. decrease lateral stability. increases lateral stability at high speeds only. increases lateral stability at all speeds.

At low orward speed:

a. 1 0

It remains the same. It increases because the TAS increases. It decreases because the ailerons are less effective. It decreases because the density decreases.

wash out. taper. sweep. anhedral.

Questions 22.

The lateral axis o an aircraf is a line which:

a. b. c. d. 23.

The relationship o thrust and lif to weight and drag. The effectiveness o the horizontal stabilizer, rudder, and rudder trim tab. The location o the CG with respect to the AC. the size o the pitching moment which can be generated by the elevator.

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 o the main planes to the horizontal datum.

Stability around the normal axis:

a. b. c. d. 28.

s n o i t s e u Q

Dihedral angle is:

a. b. c. d. 27.

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by sweeping the wings. by giving the wings anhedral. by reducing the size o the fin. by longitudinal dihedral.

What determines the longitudinal static stability o an aeroplane?

a. b. c. d. 26.

loss o longitudinal stability, and the nose to pitch up at slow speeds. excessive upward orce on the tail, and the nose to pitch down. excessive load actor in turns. high stick orces.

The tendency o an aircraf to suffer rom Dutch roll instability can be reduced:

a. b. c. d. 25.

passes through the wing tips. passes through the centre o pressure, at right angles to the direction o the airflow. passes through the quarter chord point o the wing root, at right angles to the longitudinal axis. passes through the centre o gravity, parallel to a line through the wing tips.

Loading an aircraf so that the CG exceeds the af limits could result in:

a. b. c. d. 24

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is increased i the keel surace behind the CG is increased. is given by the lateral dihedral. depends on the longitudinal dihedral. is greater i the wing has no sweepback.

I the Centre o Gravity o an aircraf is ound to be within limits or take-off:

a. b. c. d.

the C o G will be within limits or landing. the C o G or landing must be checked, allowing or uel consumed. the C o G will not change during the flight. the flight crew can adjust the CG during flight to keep it within acceptable limits or landing.

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Questions 29.

The ailerons are deployed and returned to neutral when the aircraf has attained a small angle o bank. I the aircraf then returns to a wings-level attitude without urther control movement it is:

a. b. c. d. 30.

The property which tends to decreases rate o displacement about any axis, but only while displacement is taking place, is known as:

a. b. c. d. 31.

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stability. controllability. aerodynamic damping. manoeuvrability.

I an aircraf is loaded such that the stick orce required to change the speed is zero:

a. b. c. d.

Q u e s t i o n s

neutrally stable. statically and dynamically stable. statically stable, dynamically neutral. statically stable.

the CG is on the neutral point. the CG is behind the neutral point. the CG is on the manoeuvre point. the CG is on the orward CG limit.

Answers

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Key Facts 1 (Completed) Stability is the tendency o an aircraf to return to a steady state o flight, afer being disturbed by an external orce, without any help rom the pilot. There are two broad categories o stability: static and dynamic. An aircraf is in a state o equilibrium (trim) when the sum o all orces is zero and the sum o all moments is zero. The type o static stability an aircraf possesses is defined by its initial tendency, ollowing the removal o some disturbing orce. The three different types o static stability are: • Positive static stability exists i an aircraf is disturbed rom equilibrium and has the tendency to return to equilibrium.

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s r e w s n A

• Neutral static stability exists i an aircraf is subject to a disturbance and has neither the tendency to return nor the tendency to continue in the displacement direction. • Negative static stability exists i an aircraf has a tendency to continue in the direction o disturbance. The longitudinal axis passes through the CG rom nose to tail. The normal axis passes “vertically” through the CG at 90° to the longitudinal axis. The lateral axis is a line passing through the CG, parallel to a line passing through the wing tips. The three reerence axes all pass through the centre o gravity. Lateral stability involves motion about the longitudinal axis ( rolling). Longitudinal stability involves motion about the lateral axis ( pitching). Directional stability involves motion about the normal axis ( yawing). We consider the changes in magnitude o lif orce due to changes in angle o attack, acting through a stationary point; the aerodynamic centre . The aerodynamic centre (AC) is located at the 25% chord position. The negative pitching moment about the AC remains constant at normal angles o attack. A wing on its own is statically unstable because the AC is in ront o the CG. An upward vertical gust will momentarily increase the angle o attack o the wing. The increased lif orce magnitude acting through the AC will increase the positive pitching moment about the CG. This is an unstable pitching moment. The tailplane is positioned to generate a stabilizing pitching moment about the aircraf CG.

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Answers

I the tail moment is greater than the wing moment, the sum o the moments will not be zero and the resultant nose- down moment will give an angular acceleration about the CG. The greater the tail moment relative to the wing moment, the greater the rate o return towards the original equilibrium position. The tail moment is increased by moving the aircraf CG orwards, which increases the tail arm and decreases the wing arm. I the nose-down ( negative) tail moment is greater than the nose-up ( positive) wing moment, the aircraf will have static longitudinal stability. The position o the CG when changes in the sum o the tail m oment and wing moment due to a disturbance is zero, is known as the neutral point. The urther orward the CG, the greater the nose-down angular acceleration about the CG the greater the degree o static longitudinal stability.

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The distance the CG is orward o the neutral point will give a measure o the static longitudinal stability; this distance is called the static margin.

A n s w e r s

The greater the static margin, the greater the static longitudinal stability. The af CG limit will be positioned some distance orward o the neutral point. The distance between the af CG limit and the neutral point gives the required minimum static stability margin. An aircraf is said to be trimmed i all moments in pitch, roll, and yaw are equal to zero. Trim (equilibrium) is the unction o the controls and may be accomplished by: a)

pilot effort,

b)

trim tabs,

c)

moving uel between the wing tanks and an af located trim tank, or

d)

bias o a surace actuator (powered flying controls).

The term controllability reers to the ability o the aircraf to respond to control surace displacement and achieve the desired condition o flight. A high degree o stability tends to reduce the controllability o the aircraf. The stable tendency o an aircraf resists displacement rom trim equally, whether by pilot effort on the controls ( stick orce) or gusts. I the CG moves orward, static longitudinal stability increases and controllability decreases (stick orce increases). I the CG moves af, static longitudinal stability decreases and controllability increases (stick orce decreases).

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Answers

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With the CG on the orward limit, static longitudinal stability is greatest, controllability is least and stick orce is high. With the CG on the af limit, static longitudinal stability is least, controllability is greatest and stick orce is low. The af CG limit is set to ensure a minimum degree o static longitudinal stability. The wd CG limit is set to ensure a minimum degree o controllability under the worst circumstance.

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s r e w s n A

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Answers

Key Facts 2 (Completed) Positive static longitudinal stability is indicated by a negative slope o CM versus CL. The degree o static longitudinal stability is indicated by the slope o the curve. The net pitching moment about the lateral axis is due to the contribution o each o the component suraces acting in their appropriate flow fields. In most cases, the contribution o the uselage and nacelles is destabilizing. Noticeable changes in static stability can occur at high C L (low speed) i:

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a)

the aeroplane has sweepback,

b)

there is a large contribution o ‘ power effect’, or

c)

there are significant changes in downwash at the horizontal tail,

The horizontal tail usually provides the greatest stabilizing influence o all the components o the aeroplane. ( page 259).

A n s w e r s

Downwash decreases static longitudinal stability.

I the thrust line is below the CG, a thrust increase will produce a positive or nose-up moment and the effect is destabilizing. High lif devices tend to increase downwash at the tail and reduce the dynamic pressure at the tail, both o which are destabilizing. An increase in TAS, or a given pitching velocity, decreases aerodynamic damping. The aeroplane with positive manoeuvring stability should demonstrate a steady increase in stick orce with increase in load actor or “ g”. The stick orce gradient must not be excessively high or the aeroplane will be difficult and tiring to manoeuvre. Also, the stick orce gradient must not be too low or the aeroplane may be overstressed inadvertently when light control orces exist. When the aeroplane has high static stability, the manoeuvring stability will be high and a high stick orce gradient will result. The orward CG limit could be set to prevent an excessively high manoeuvring stick orce gradient. As the CG moves af, the stick orce gradient decreases with decreasing manoeuvring stability and the lower limit o stick orce gradient may be reached. At high altitudes, the high TAS reduces the change in tail angle o attack or a given pitching velocity and reduces the pitch damping. Thus, a decrease in manoeuvring stick orce stability can be expected with increased altitude. A flying control system may employ centring springs, down springs or bob weights to provide satisactory control orces throughout the speed, CG and altitude range o a n aircraf. While static stability is concerned with the initial tendency o an aircraf to return to equilibrium, dynamic stability is defined by the resulting motion with time.

326

Answers

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An aircraf will demonstrate positive dynamic stability i the amplitude o motion decreases with time. When natural aerodynamic damping cannot be obtained, artificial damping must be provided to give the necessary positive dynamic stability. The longitudinal dynamic stability o an aeroplane generally consists o two basic modes o oscillation: a)

long period (phugoid)

b)

short period

The phugoid oscillation occurs with nearly constant angle o attack. The period o oscillation is so great, the pilot is easily able to counteract long period oscillation. Short period oscillation involves significant changes in angle o attack.

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s r e w s n A

Short period oscillation is not easily controlled by the pilot. The problems o dynamic stability can become acute at high altitude because o reduced aerodynamic damping. To overcome the directional instability in the uselage it is possible to incorporate into the overall design dorsal or ventral fins. The fin is the major source o directional stability or the aeroplane. A T - tail makes the fin more effective by acting as an “ end plate”. Because the dorsal fin stalls at a very much higher angle o attack, it takes over the stabilizing role o the fin at large angles o sideslip. Sweepback produces a directional stabilizing effect, which increases with increase in C L. Ventral fins increase directional stability at high angles o attack.

Landing clearance requirements may limit their size, require them to be retractable, or require two smaller ventral fins to be fitted instead o one large one. Generally, good handling qualities are obtained with a relatively light, or weak positive, lateral stability. The principal surace contributing to the lateral stability o an aeroplane is the wing. The effect o geometric dihedral is a powerul contribution to lateral stability. A low wing position gives an unstable contribution to static lateral stability. A high wing location gives a stable contribution to static lateral stability. The magnitude o “dihedral effect” contributed by the vertical position o the wing is large and may require a noticeable dihedral angle or the low wing configuration. A high wing position, on the other hand, usually requires no geometric dihedral at all.

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Answers

The swept-back wing contributes a positive “dihedral effect”. An aircraf with a swept-back wing requires less geometric dihedral than a straight wing. The fin contribution to purely lateral static stability, is usually very small. Excessive “dihedral effect” can lead to “ Dutch roll,” difficult rudder coordination in rolling manoeuvres, or place extreme demands or lateral control power during crosswind take-off and landing. Deploying partial span flaps gives a reduced dihedral effect. A swept-back wing requires much less geometric dihedral than a straight wing. I a requirement also exists or the wing to be mounted on top o the uselage, an additional “dihedral effect” is present. A high mounted and swept-back wing would give excessive “dihedral effect”, so anhedral is used to reduce “dihedral effect” to the required level. 1 0

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.

A n s w e r s

Spiral divergence will exist when static directional stability is very large when compared to the “dihedral effect”. The rate o divergence in the spiral motion is usually so gradual that the pilot can control the tendency without difficulty. Dutch roll will occur when the “dihedral effect” is large when compared to static directional stability. Aircraf which Dutch roll are fitted with a Yaw Damper. This automatically displaces the rudder proportional to the rate o yaw to damp-out the oscillations. I the Yaw Damper ails in flight, it is recommended that the ailerons be used by the pilot to damp-out Dutch roll. I the pilot uses the rudder, pilot induced oscillation (PIO) will result and the Dutch roll may very quickly become divergent, leading to loss o control. When the swept wing aeroplane is at low C L the “dihedral effect” is small and the spiral tendency may be apparent. When the swept wing aeroplane is at high C L the “dihedral effect” is increased and the Dutch Roll oscillatory tendency is increased. When pilot induced oscillation is encountered, the most effective solution is an immediate release o the controls. Any attempt to orcibly damp the oscillation simply continues the excitation and amplifies the oscillation. Higher TAS common to high altitude flight reduces the angle o attack changes and reduces aerodynamic damping. Mach Tuck is caused by loss o lif in ront o the CG and reduced downwash at the tail due to the ormation o a shock wave on a swept-back wing at high Mach numbers. The Mach trim system will adjust longitudinal trim to maintain the required stick orce gradient and operates only at high Mach numbers.

328

Answers

10

0 1

s r e w s n A

329

10

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Answers

1 0

A n s w e r s

330

1 d

2 b

3 c

4 a

5 a

6 b

7 a

8 b

9 c

10 c

11 a

12 c

13 c

14 d

15 d

16 d

17 c

18 b

19 a

20 c

21 d

22 d

23 a

24 b

25 c

26 d

27 a

28 b

29 b

30 c

31 a

Chapter

11 Controls

Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 333 Hinge Moments . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 334 Control Balancing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 335 Mass Balance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 340 Longitudinal Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 340 Lateral Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 341 Speed Brakes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 346 Directional Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 348 Secondary Effects o Controls . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 350 Trimming . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 351 Questions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

355

Answers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

362

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Controls

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Important Definitions

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Pitch Angle:

The angle between the aircraf longitudinal axis and the horizon.

Roll Angle:

The angle between the aircraf lateral axis and the horizon.

Yaw Angle:

The angle between the aircraf longitudinal axis and the relative airflow.

Controls

11

Introduction All aircraf are fitted with a control system to enable the pilot to manoeuvre and trim the aircraf in flight about each o its three axes. The aerodynamic moments required to to rotate the aircraf about the axes are usually supplied by means o ‘flap’ t ype control suraces positioned at the extremities o the aircraf so that they have the longest possible moment arm about the CG. There are usually three separate separate control systems and three sets o control control suraces: • Rudder or control in yaw about the normal axis (directional control). • Elevator or control in pitch about the lateral axis (longitudinal control). • Ailerons or control in roll about the longitudinal axis (late (lateral ral control). control). Spoilers may also be used to assist or replace the ailerons ail erons or roll control. The effect o two o these controls may be combined in a single set o control suraces: • Elevons: combine the effects o elevator elevator and aileron.

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• Ruddervator: (‘V’ or butterfly tail) combines the effects o rudder and elevator elevator..

s l o r t n o C

• Tailerons : slab horizontal horizontal tail suraces that that move either either together, together, as pitch control, control, or independently or control in roll. The moment around an axis is produced by changing the aerodynamic orce on the appropriate aerooil. The magnitude o the orce is a product o the dynamic pressure (IAS ) and the angular displacement o o the control surace. Aerodynamic orce can be changed by: 2

• adjusting the camber o the aerooil. aerooil. • changing the incidence o the aerooil. • decreasing lif and increasing drag by “spoiling” “spoiling” the airflow. Changing the camber o any aerooil (wing, tailplane or fin) will change its lif. Deflecting a control surace effectively changes its camber. Figure 11.1 shows the effect on CL o movement m ovement o a control surace.

CL

ANGLE OF ATTACK

Figure 11.1 Control surace changes camber & lif

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Controls Changing the incidence o an aerooil will change its lif. The usual application o this system is or pitch control - the all moving (slab) tailplane. There is no elevator; elevator; when the the pilot makes a pitch input, the incidence o the whole tailplane changes.

Figure 11.2 All moving (slab) tailplane

SPOILER SURFACES 1 1

AILERONS

C o n t r o l s

Spoilers are a device or reducing the lif o an aerooil by disturbing the airflow over the upper surace. They assist lateral control by moving up on the side with the up-going aileron, as illustrated in Figure 11.3.

Figure 11.3 Spoilers

Hinge Moments I an aerodynamic orce acts on a control surace, it will try to rotate the control around its hinge in the direction o the orce. The moment is a product o the orce (F) times the distance (d) rom the hinge line to the control control surace CP. This is called the the hinge moment. moment. The orce is due to the surace area, the angular displacement o the control surace and the dynamic pressure.

HINGE MOMENT = F × d

F2 F

d

d

Figure 11.4 Hinge moment (eel)

To move the control surace to the required angular displacement and maintain it in that position the pilot has to overcome, then balance, the hinge moment by applying a orce (stick orce) to the cockpit control. control. The stick orce will thereore thereore depend on the size o o the hinge moment.

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Controls

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Control Balancing The aerodynamic orce on the controls will depend on the area o the control surace, its angular displacement and the IAS. For large and ast aircraf the resulting resulting aerodynamic orce orce can give hinge moments / stick orces which are too high or easy operation o 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 orm o aerodynamic balance.

Aerodynamic Balance Aerodynamic balance involves using the aerodynamic orces on the control surace to reduce the hinge moment / stick orce orce and may be done in several several ways: ways: Inset Hinge HINGE SET - BACK INTO SURFACE

F2

d

I the distance (d) is reduced, the hinge moment will be reduced. The smaller smaller the hinge moment, the smaller the stick orce and the easier it will be or the pilot to move the controls. Setting the hinge back does not reduce the effectiveness o the control, only the hinge moment.

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s l o r t n o C

I the aerodynamic orce (F2) were to move orward o the hinge, a condition known as “overbalance” would exist. As the orce moved orward, a reduction then a reversal o the stick orce orce would occur. This would be very dangerous and the designer must ensure the aerodynamic orce can never move orward o the hinge.

Figure 11.5 Inset hinge

Horn Balance HINGE AEROFOIL AEROFOI L

LINE

CONTROL SURFACE HORN

Figure 11.6 Horn balance

The principle o the horn balance is similar to that o the inset hinge, in that part o the surace is orward o the hinge line, and orces on this part o the surace sur ace give hinge moments which are in the opposite direction to the moments on the main part o the the surace. The overall moment is thereore reduced, but the control effectiveness is not.

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Controls Internal Balance

This balance works on the same principle as the inset hinge, but the aerodynamic balance area is inside the wing.

LOW PRESSURE

HIGH PRESSURE

Figure 11.7 Internal balance

Movement o the control causes pressure changes on the aerooil, and these pressure changes are elt on the balance area. For example, i the control control surace is moved down, down, pressure above the aerooil aerooil is reduced and pressure below it is increased. increased. The reduced pressure is elt on the upper surace o the balance ‘panel’, and the increased pressure on the lower surace. The pressure difference on the balance thereore gives a hinge moment which is the opposite to the hinge moment on the main control surace, and the overall overall hinge moment is reduced. reduced.

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C o n t r o l s

See page 354 or a Tab Quick Reerence Guide. Balance Tab

The preceding types o aerodynamic balance work by causing some o the dynamic pressure on the control control surace to act orward o o the hinge line. The balance tab provides provides a orce acting on the control surace trailing edge opposite to the orce on the main control surace. The balance tab moves in the opposite direction direction to the control surace. The pilot moves the surace, the surace moves the tab.

TAB FORCE PILOT INPUT

CONTROL FORCE

Figure 11.8 Balance tab

Unlike the previous types o balance, the balance tab will give some reduction in control effectiveness, as the tab orce is opposite to the control orce.

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Anti-balance Tab The anti-balance tab moves in the same direction as the control surace and increases i ncreases control

effectiveness, but it will increase increase the hinge moment moment and give heavier heavier stick orces. The pilot moves the surace, the surace sur ace moves the tab. PILOT INPUT TAB FORCE

CONTROL FORCE

Figure 11.9 Anti-balance tab

Servo Tab

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s l o r t n o C

Pilot control input deflects the servo tab only ; the aerodynamic orce on the tab then moves the control surace until an equilibrium position is reached. reached. I external control locks are ar e fitted to the control surace sur ace on the ground, the cockpit control will still be ree to move; thereore, you must physically check any central locks have been removed beore flight . Older types o

high speed jet transport aircraf (B707) successully used servo tab controls. The disadvantage o the servo tab is reduced control effectiveness at low IAS.

CONTROL "HORN" FREE TO

PILOT INPUT

PIVOT ON HINGE AXIS

Figure 11.10 Servo tab

Spring Tab

The spring tab is a modification o the servo tab, such that tab movement is proportional to the applied stick orce. orce. Maximum tab assistance is obtained at high speed when the stick orces are greatest. greatest. High dynamic pressure will prevent prevent the surace rom moving, so the spring is compressed by the pilot input and the tab moves moves the surace. The spring is not compressed at low IAS, so the pilot input deflects the control surace and the tab, increasing the surace area and control effectiveness at low speed. HORN FREE TO PIVOT ON HINGE AXIS PILOT INPUT SPRING

Figure 11.11 Spring tab

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Controls (Hydraulic) Powered Flying Controls I the required assistance or the pilot to move the controls cannot be provided by the preceding types o aerodynamic balance, balance, then power assisted assisted or ully powered controls have have to be used. used.

POW ER FLYING POWER FLYING CONTROL UNIT (PFCU)

SERVO VALVE 1 1

C o n t r o l s

Figure 11.12 Power assisted flying control

Power Assisted Controls

With a power assisted flying control, Figure 11.12, only a certain proportion o the orce required to oppose the hinge moment is provided by the pilot; the hydraulic system provides most o the orce. orce. Although the pilot does not have have to provide all the orce required, required, the natural ‘eel’ o the controls is retained and the stick orce increases as the square o the IAS, just as in a completely manual manual control.

POWE R FLYING FLYING CONTROL UNIT (PFCU)

SERVO VALVE

Figure 11.13 Fully powered flying control

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Controls

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Fully Powered Controls

For bigger and/or aster aircraf, hinge moments are so large that ully powered controls must be used. In a ully powered control system, system, none o the orce to move move the control surace is supplied by the pilot. The only orce the pilot supplies is that required required to overcome overcome system riction and to move the servo valve; all the necessary power to move the control surace is supplied by the aircraf’s ai rcraf’s hydraulic system. Figure 11.13 shows that movement o the servo valve to the lef allows hydraulic fluid to enter enter

the lef chamber o the PFCU. The body o the unit will move to the lef, its movement being transerred transerre d to the control surace. sur ace. As soon as the PFCU body reaches the position into which the pilot placed the servo valve, the PFCU body, and hence the control surace, stops moving. The unit is now locked in its new position by “incompressible” liquid trapped on both sides o the piston and will remain in that position until the servo valve is again moved m oved by the pilot. Aerodynamic loads on the control surace are unable to move the cockpit controls, so powered flying controls are known as “irreversible” controls. Artificial Feel (‘Q’ Feel) 1 1

s l o r t n o C

POWER FLYING CONTROL CONTR OL UNIT UN IT (PFC (PFCU) U)

STATIC

PITOT

SERVO VALVE

ARTIFICIAL FEEL UNI ARTIFICIAL UNIT T ( 'Q ' FEEL )

Figure 11. 11.1 14 Artifici Artificial al eel (‘Q’ eel)

With a ully powered flying control the pilot is unaware o the aerodynamic orce on the controls, so it is necessary to incorporate “artificial” eel to prevent the aircraf rom being overstressed. As shown schematically schematically in Figure 11.14, a device sensitive to dynamic pressure (½ ρ V ) or ‘Q’ is used. 2

Pitot pressure is ed to one side o a chamber and static pressure to the other, which moves a diaphragm under the influence o changing dynamic pressure with airspeed and causes “regulated” hydraulic pressure to provide a resistance or ‘eel’ on the pilot’s input controls proportional to IAS , just as in a manual control. In addition, stick orce should increase as stick displacement increases. 2

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Controls Mass Balance Mass balance is a WEIGHT attached to the control surace orward o the h inge . Most control

suraces are mass mass balanced. The purpose o this is to to prevent prevent control surace flutter. flutter. Flutter is an oscillation o the control surace which can occur o ccur due to the bending and twisting twis ting o the structure under load. I the control surace CG is behind the hinge line, inertia will cause the surace to oscillate about about its hinge line. The oscillations can be divergent divergent and cause structural ailure. A detailed explanation o flutter flutter is given in Chapter 14.

HINGE LINE

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C o n t r o l s

Figure 11.1 11.15 5 Mass balance weights weig hts

Flutter may be prevent p revented ed by adding weight to the control surace in ront o the hinge line . This brings the centre o gravity o the control orward to a position on, or slightly in ront o the hinge, but always always to to the point required required by the designers designers . This reduces reduces the inertia moments about the hinge and prevents flutter developing. Figure 11.15 illustrates some common methods o mass balancing.

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 aircraf throughout its speed range at all permitted CG positions and configurations and to give an adequate rate o pitch or manoeuvres.

Effect of Elevator Deflection Suppose that the aircraf is flying in balance at a steady speed speed with the elevator neutral. neutral. I the elevator is deflected upwards, the tail will develop a down load which will begin to pitch the aircraf nose upwards. As the angle o attack increases, increases, the tailplane down load decreases decreases and the aircraf will reach an equilibrium pitch position. It will then remain in that pitch position with the elevator elevator kept at the selected selected angle. I the elevator is returned returned to neutral, the tail will develop an upload which will begin to pitch the aeroplane down again. At a given CG position there will be a given pitch pitch attitude or each elevator elevator position.

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Controls

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Direction of the Tailplane Load The elevator elevator angle required to give balance balance depends on IAS and the CG position. At normal cruising speeds and CG positions, the elevat elevator or should ideally be approximately neutral. neutral. The tailplane will be giving a down load and, a nd, consequently, a nose-up pitching moment. This will balance the nose-down moment created by the wing with its centre o pressure airly well af. At higher than normal normal speeds, the CP will move urther rearwards giving a stronger nosedown pitch and needing a larger down-load rom the the tailplane. However However,, at higher speed the aircraf’s angle o attack will wi ll be decreased, requiring some down elevator to provide the correct tail-load. At lower than normal speeds, the CP will move orward and the wing and uselage may cause a nose-up pitching pitching moment. The tailplane will be required to to give an up-load or balance. balance. At low speed, the aircraf will be at a high h igh angle o attack, and to reach this attitude, the elevator will have been moved up.

Elevator Angle with ‘g’ When the aircraf is perorming a pitching manoeuvre, the tailplane angle o attack is increased by the effect o the rotational rotational velocity and the aerodynamic damping damping is increased. This means that a larger elevat elevator or angle will be required than or the same conditions in 1g flight. The additional elevator elevator angle required required will be proportional to to the ‘g’ being being experienced. experienced. The elevator movement available should be sufficient to allow the design limit ‘g’ to be reached.

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s l o r t n o C

The most demanding requirement or elevator up authority will be when the aircraf is being flared or landing, in ground effect with most orward CG .

Effect of Ice on the Tailplane The tailplane is an aerooil, usually symmetrical as it i s required to produce both up and down loads. It is set at an angle o incidence which is less than than that o the wing. This ensures that that it will not stall beore the wing, and so control can be maintained maintained up to the stall. It is usually affected by the downwash rom the wing and this reduces its effective angle o attack . Typically the tail will be at a negative angle o attack, attack, producing a down load or balance. balance. I ice orms on the tailplane leading edge, its aerooil shape will be distorted, and its stalling angle reduced. This could cause the tailplane to stall, particularly i the downwash downwash is increased as a result o lowering lowering flaps. With the tailplane stalled its its down load would be reduced, reduced, and the aircraf would pitch down and could not be recovered.

Lateral Control Control in roll is usually obtained obtained by ailerons or by spoilers, or by a combination o the two. The main requirement or lateral control is to achieve an adequate rate o roll. On the ground with the control wheel in the neutral position both ailerons should be slightly below alignment with the wing trailing edge, or “drooped”. When airborne, the lower pressure on the top surace will “suck” both ailerons up into a position where they are perectly aligned with the wing trailing edge, thus reducing drag.

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Controls Effect of Aileron Deflection (Aerodynamic Damping) In steady level flight with the ailerons ailerons neutral, the lif on the two wings will be equal. I the control wheel is turned to the lef, the lef aileron will move up and the right aileron down. The up aileron will decrease the lif o the lef wing which which will begin to ‘drop’. The downward movement o the wing creates a relative airflow upwards, which increases its effective angle o attack. The opposite effects will occur on the right right (up-going) wing.

HIGH TAS

LOW TAS

RELATIVE AIRFLOW FROM A NGULAR ROTATION ROTATIO N EQUAL WING TIP VELOCITY INCREAS E IN EFFECTIV EFFECTIV E A NGLE OF ATTA ATTACK CK DUE TO WING TIP DOWNWA RDS VELOCITY VELOCITY

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C o n t r o l s

Figure 11. 11.16 16 Aerodynamic damping damp ing in roll

The increased effective angle o attack o the down-going wing increases its lif, which opposes the roll. This is called aerodynamic damping. The greater the rate o roll, the greater the damping. It can also be seen rom Figure 11.16 that the greater the TAS, the smaller the increase in effective angle o attack or a given roll rate. The change in wing lif or a given aileron deflection depends on the IAS, but the change o effective angle o attack due to roll velocity depends on TAS. TAS. At high TAS TAS (constant IAS, higher altitude) the change in effective angle o attack will be reduced and a higher rate o roll will be possible. Rate o roll thereore thereore increases (aerodynamic damping decreases) decreases) with higher TAS or a given aileron deflection. The aileron is known as a rate control since a given aileron angle o deflection determines a rate o roll, not a roll displacement. Effect o Wingspan on Rate o Roll For a given rate o roll, the wing tip rolling velocity will increase with the wingspan. Aerodynamic damping will thereore be greater greater i the span is greater greater.. Under the same conditions, a short span wing will have a greater rate rate o roll than a large span wing. Adverse Aileron Yaw Yaw The ailerons produce a rolling moment by increasing the lif on one wing and decreasing it on the other. other. The increased lif on the up-going wing gives an increase in the induced drag, drag, whereas the reduced lif on the down-going wing gives a decease decease in induced 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 lef produces a yawing moment to to the right. This is known as adverse aileron yaw.

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Reducing Adverse Aileron Yaw LARGE UPWARD MOV MO V EMENT

SMALL DOWNWARD MOVEMENT

Figure 11. 11.1 17 Differential ailerons aileron s

• Differential ailerons : The aileron linkage linkage causes the up-going aileron aileron to move move through a larger angle than the down-going aileron, Figure 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.

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s l o r t n o C

Figure Figu re 11. 1.18 18 Frise Fr ise ailerons aileron s

• Frise ailerons: These have an asymmetric leading edge, as illustrated in Figure 11.18. The leading edge o the up-going aileron protrudes below the lower surace o the wing, causing high drag. The leading edge o the down-going aileron remains remains shrouded and causes less drag. • Aileron-rudder coupling: In this system the aileron and rudder controls are interconnected, interconnected, so that when the ailerons are deflected, the rudder automatically moves to counter the adverse yaw. • Roll control spoilers: I roll spoilers are used to to augment the roll rate rate obtained rom the ailerons, they will reduce the adverse yaw, as the down-going wing will have an increase in drag due to the raised spoiler.

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Controls Inboard Ailerons Ailerons are normally situated at the wing tips to give the greatest rolling moment or the orce 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 o aileron effectiveness. As IAS increases, a down-going aileron will give more wing twist (decreased wing tip incidence). Eventually, an IAS will be reached at which the decrease in tip incidence will give a larger downorce than the uporce 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 urther inboard. Unortunately, this would reduce aileron effectiveness at low speed.

SPOILER SURFACES (LIFT DUMP POSITION) OUTBOARD AILERONS (LOW SPEED ONLY)

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C o n t r o l s

INBOARD AILERONS (HIGH SPEED AND LOW SPEED)

Figure 11.19 Inboard & outboard ailerons & spoiler suraces

Alternatively, two sets o ailerons may be fitted, as illustrated in Figure 11.19: one set at the wing tips or use only at low speeds when the orces involved are low, and one set inboard or use at high speeds when the orces are greater and could cause greater structural distortion. Outboard (low speed) ailerons are “locked-out” as the flaps retract. At low speed both sets o ailerons work, but at high speed only the inboard ailerons respond to control input. Flaperons The flaps and the ailerons both occupy part o the wing trailing edge. For good take-off and landing perormance, the flaps need to be as large as possible, and or a good rate o 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 rom the drooped position to give lateral control. Another system is to use the trailing edge moveable suraces to perorm the operation o both flaps and ailerons.

344

Controls

11

Roll Control Spoilers Spoilers may be used to give lateral control, in addition to, or instead o ailerons. Spoilers consist o movable panels on the upper wing surace, hinged at their orward edge, which can be raised hydraulically, as illustrated in Figure 11.20. A raised spoiler will disturb the airflow over the wing and reduce lif.

OUTBOARD AILERONS

SPOILER SURFACES ASSISTING W ITH ROLL CONTROL

LOCKED - OUT AT HIGH SPEED

INBOARD AILERONS OPERATE AT ALL SPEEDS

1 1

s l o r t n o C

Figure 11.20 Roll control spoilers

To unction as a lateral control, the spoilers rise on the wing with the up-going aileron (downgoing wing), proportional to aileron input. On the wing with the down-going aileron, they remain flush. Unlike ailerons, spoilers cannot give an increase o lif, so a roll manoeuvre controlled by spoilers will always give a net loss o lif. However, the spoiler has several advantages compared to the aileron: • There is no adverse yaw : The raised spoiler increases drag on that wing, so the yaw is in the same direction as the roll. • Wing twisting is reduced : The aerodynamic orce on the spoilers acts urther orward than is the case with ailerons, reducing the moment which tends to twist the wing. • At transonic speed its effectiveness is not reduced by shock induced separation. • It cannot develop flutter. • Spoilers do not occupy the trailing edge, which can then be utilized or flaps.

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11

Controls Combined Aileron and Spoiler Controls On a ew aircraf, lateral control is entirely by spoilers, but in the majority o applications, the spoilers work in conjunction with the ailerons. Ailerons alone may be inadequate to achieve the required rate o 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 i there is shock induced separation. Spoilers can be used to augment the rate o roll, but they may not be required to operate over the whole speed range. On some aircraf, the spoilers are only required at low speed, and this can be achieved by making them inoperative when the flaps are retracted.

Movement o the cockpit control or 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.

Speed Brakes Speed brakes are devices to increase the drag o an aircraf when it is required to decelerate quickly or to descend rapidly. Rapid deceleration is required i turbulence is encountered at high speed, to slow down to the Rough-air Speed as quickly as possible. A high rate o descent may be required to conorm to Air Traffic Control requirements, and particularly i an emergency descent is required.

1 1

C o n t r o l s

Types o Speed Brake Ideally, the speed brake should produce an increase in drag with no loss o lif or change in pitching moment. The uselage mounted speed brake is best suited to meet these requirements, Figure 11.21.

W ING MOUNTED SPEED BRAKES

FUSELAGE MOUNTED SPEED BRAKE

Figure 11.21 Wing mounted & uselage mounted speed brakes

However, as the wing mounted spoiler gives an increase in drag, it is convenient to use the spoiler suraces as speed brakes in addition to their lateral control unction. To operate as speed brakes they are controlled by a separate lever in the cockpit and activate symmetrically . There is no speed restriction or the operation o speed brakes, but they may “blow back” rom the ully extended position at high speeds. Spoilers will still unction as a roll control whilst being used as speed brakes, by moving asymmetrically rom the selected speed brake position.

346

Controls

11

An example is illustrated in Figure 11.22. Speed brakes have been selected, and then a turn to the lef is initiated. The spoiler suraces 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 suraces on the wing with the down-going aileron will modulate towards the stowed position. The spoiler suraces on the wing with the down-going aileron may partially or ully stow, again depending on the speed brake selection and the roll input.

SPEED BRAKES

SPEED BRAKE AND ROLL INPUT 1 1

s l o r t n o C

Figure 11.22 Mixed speed brake & roll input

Effect o Speed Brakes on the Drag Curve The drag resulting rom the operation o speed brakes is profile drag, so it 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 aircraf in the clean configuration. Ground Spoilers ( Lif Dumpers) During the landing run, the decelerating orce 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 lif that the wing is producing. Lif can be reduced by operating the speed brake lever to the lif dump position, Figure 11.19. Both the wheel brake drag and the aerodynamic drag are increased, and the landing run is reduced. On many aircraf types, additional spoiler suraces are activated in the lif 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 o the leg afer take-off.

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11

Controls Directional Control Control in yaw is obtained by the rudder. The rudder is required to: • • • • •

maintain directional control with asymmetric power. correct or crosswinds on take-off and landing. correct or adverse yaw. recover rom a spin. correct or changes in propeller torque on single-engine aircraf.

Effect o Rudder Deflection I the rudder is deflected to the lef, the aircraf will begin to yaw to the lef. This will create a sideslip to the right. The sideslip airflow rom the right acting on the fixed part o the fin will cause a side load to the lef, opposing the effect o the rudder. As the yaw increases, this damping orce will increase until it balances the rudder orce. The aircraf will then stop yawing and will maintain that angle o yaw, with the rudder deflected to its original position. I the rudder is returned to the neutral position, both the fin and the rudder will give a orce to the lef which will return the aircraf to its original position with zero yaw. A given rudder angle will thereore correspond to a given yaw displacement.

1 1

C o n t r o l s

Fin Stall The sideslip angle is effectively the angle o attack o the fin, and as or any aerooil, there will be a critical angle at which it will stall. I the rudder is deflected in the direction to correct the sideslip, the stalling angle will be reduced.

DORSAL FIN

Figure 11.23

The stalling angle o an aerooil is affected by its aspect ratio, and so the stalling angle o the fin could be increased by decreasing its aspect ratio. This can be done by fitting a dorsal fin, Figure 11.23.

348

Controls

11

Asymmetric Thrust For a twin-engine aircraf, i engine ailure occurs, the thrust rom the operating engine will cause a yawing moment. This must be counteracted by the rudder. The rudder orce will vary with speed squared, and so there will be a minimum speed at which the orce will be sufficient to balance the engine yawing moment. This is the minimum control speed (V MC). Rudder Ratio Changer 450

400

350

300

250

1 1

200

s l o r t n o C

150

100

50

0

5

10

15

20

25

30

RUDDER ANGLE - DEGREES

Figure 11.24 Rudder ratio

With a simple control system, ull rudder pedal movement will provide ull rudder deflection. With high speed aircraf, while it is necessary to have large rudder deflections available at low speed, when flying at high speed, ull 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 or ull pedal movement above a certain speed, or a progressive gear change which gives a decreasing rudder deflection with ull pedal movement as speed increases, Figure 11.24.

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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. Yawing Moment Due to Roll • 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 aircraf, or instance, roll lef and at the same time, yaw right.

• I the aircraf is rolling, the down-going wing experiences an increased angle o attack and the up-going wing a decreased angle o attack, increasing the adverse yawing moment. Rolling Moment Due to Yaw • I the aircraf is yawing to the lef, the right wing has a higher velocity than the lef wing and so will give more lif. The difference in lif will give a rolling moment to the lef.

1 1

• I the rudder is deflected to the lef (to give yaw to the lef) the orce on the fin is to the right. This will give a small rolling moment to the right because the fin CP is above the aircraf CG. This effect is usually very small, but a high fin may give an adverse roll.

C o n t r o l s

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.

350

Controls

11

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. I it is necessary or a control surace to be deflected to maintain balance o the aircraf, the pilot will need to apply a orce to the cockpit control to hold the surace in its deflected position. This orce may be reduced to zero by operation o the trim controls.

The aircraf may need to be trimmed in pitch as a result o: • • • •

changes o speed. changes o power. varying CG positions. changes o configuration.

Trimming in yaw will be needed: • on a multi-engine aircraf i there is asymmetric power. • as a result o changes in propeller torque.

1 1

Trimming in roll is less likely to be needed but could be required i the configuration is asymmetric, or i there is a lateral displacement o the CG.

s l o r t n o C

Methods o Trimming Various methods o trimming are in use. The main ones are:

• • • • •

the trimming tab. variable incidence (trimming) tailplane. spring bias. CG adjustment. adjustment o the artificial eel unit.

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Controls Trim Tab A trim tab is a small adjustable surace set into the trailing edge o a main control surace. Its 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 surace in its required position, the tab is moved in the opposite direction to the control surace until the tab moment balances the control surace hinge moment.

d

F

D f 1 1

Figure 11.25

C o n t r o l s

Figure 11.25 shows ( × D) rom tab opposes (F × d) rom control surace. I the two moments

are equal, the control will be trimmed, i.e. the stick orce will be zero. Operation o the trim tab will slightly reduce the orce being produced by the main control surace. Fixed Tabs Some trim tabs are not adjustable in flight but can be adjusted on the ground, to correct a permanent out o trim condition. They are usually ound on ailerons and rudder. They operate in the same manner as the adjustable trim tab. Variable Incidence (Trimming) Tailplane This system o 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, Figure 11.26 . Stick orce is now zero.

The main advantages o a variable incidence (trimming) tailplane are: • the drag is less in the trimmed state as the aerooil is more streamlined. • trimming does not reduce the effective range o pitch control as the elevator remains approximately neutral when the aircraf is trimmed. • it is very powerul and gives an increased ability to trim or larger CG and speed range. The disadvantage o a variable incidence (trimming) tailplane is that it is more complex and is heavier than a conventional trim tab system.

352

Controls

11

ELEVATOR POSITIONED TO TRIM

A / C STRUCTURE

SCREW JACK

AFTER TRIM INPUT

Figure 11.26 Variable incidence (trimming) tailplane

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s l o r t n o C

The amount o trim required will depend on the CG position, and recommended stabilizer takeoff settings will be given in the aircraf Flight Manual. It is important that these are correctly set beore take-off as incorrect settings could give either an excessive rate o pitch when the aircraf is rotated, leading to possible tail strikes, or very heavy stick orces on rotation, leading to increased take-off distances required.

10 20 10

5

15

20

Figure 11.27 Reduced aircraf nose-up pitch authority

The disadvantage o a ”conventional” elevator and trim tab, Figure 11.27 , is that the aircraf nose-up pitch authority reduces with orward CG movement. Forward CG positions will require the elevator to be trimmed more aircraf nose-up. The illustration shows up elevator authority reduced rom 10° to 5°.

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Controls Spring Bias In the spring bias trim system, an adjustable spring orce is used to decrease the stick orce. No tab is required or this system. CG Adjustment I the flying controls are used or trimming, this results in an increase o drag due to the deflected suraces. The out o balance pitching moment can be reduced by moving the CG, thus reducing the balancing load required and thereore the drag associated with it. This will give an increase o cruise range. CG movement is usually achieved by transerring uel between tanks at the nose and tail o the aircraf. Artificial Feel Trim I the flying controls are power operated, there is no eedback o the load on the control surace to the cockpit control. The eel on the controls has to be created artificially. When a control surace is moved, the artificial eel unit provides a orce to resist the movement o the cockpit control. To remove this orce (i.e. to trim) the datum o the eel unit can be adjusted so that it no longer gives any load on the flight deck controls.

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TABS - Quick Reerence Guide

C o n t r o l s

Type o Tab Balance Anti-balance Servo Spring Trim

354

Operated by

Movement Relative to Control Surace

Stick Force

Control Effectiveness

Opposite

Less

Reduced

Same

More

Increased

Pilot

Opposite

Less

Reduced

Pilot at High Speed Trim Control ONLY

Opposite at High Speed

Less at High Speed

Reduced at High Speed

Opposite

Zeroed

Reduced

Control Surace Control Surace

Questions

11

Questions 1.

An elevon is:

a. b. c. d. 2.

When rolling at a steady rate the:

a. b. c. d. 3.

s n o i t s e u Q

lateral control about the lateral axis. longitudinal control about the lateral axis. lateral control about the longitudinal axis. directional control about the normal axis.

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.

I the ailerons are deflected to 10°, compared to 5°, this will cause:

a. b. c. d. 7.

the rudder. the ailerons. the elevators. the flaps.

Aileron reversal would be most likely to occur:

a. b. c. d. 6.

1 1

Ailerons give:

a. b. c. d. 5.

up-going wing experiences an increase in effective angle o attack. rate o roll depends only on aileron deflection. down-going wing experiences an increase in effective angle o attack. effective angle o attack o the up-going and down-going wings are equal.

The control surace which gives longitudinal control is:

a. b. c. d. 4.

an all moving tailplane that has no elevator. the correct name or a V - tail. a surace that extends into the airflow rom the upper surace o the wing to reduce the lif. a combined aileron and elevator fitted to an aircraf that does not have conventional horizontal stabilizer (tailplane).

an increased angle o bank. an increased rate o roll. no change to either bank angle or roll rate. a reduction in the adverse yawing moment.

Yawing is a rotation around:

a. b. c. d.

the normal axis controlled by elevator. the lateral axis controlled by rudder. the longitudinal axis controlled by ailerons. the normal axis controlled by rudder.

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Questions 8.

I the control column is moved orward and to the lef:

a. b. c. d. 9.

The secondary effect o yawing to port is to:

a. b. c. d. 10.

Q u e s t i o n s

11.

c. d.

b. c. d.

may be moved by operating the cockpit control but not by the aerodynamic loads acting on the control surace. has less movement in one direction than the other. may be moved either by the cockpit control or by a load on the control surace. is when the control locks are engaged.

Ailerons may be rigged slightly down (drooped):

a. b. c. d.

356

to enable any ree movement in the control system to be detected. to prevent structural damage to the controls in gusty conditions when the aircraf is on the ground. to keep the control surace rigid to permit ground handling. as a security measure.

An irreversible control:

a.

14.

A gradual stiffening o the controls. Rebound on reaching the stops. A solid stop. Controls should not be moved to the stops.

The purpose o control locks on a flying control system is:

a. b.

13.

a yawing moment to the lef but no rolling moment. a rolling moment to the lef. a rolling moment to the right. a yawing moment to the right but no rolling moment.

What should be the eel on a ‘ull and ree’ check o the controls?

a. b. c. d. 12.

roll to starboard. pitch nose-up. roll first to starboard and then to port. roll to port.

Due to the AC o the fin being above the longitudinal axis, i the rudder is moved to the right, the orce acting on the fin will give:

a. b. c. d.

1 1

the lef aileron moves up, right aileron moves down, elevator moves up. the lef aileron moves down, right aileron moves up, elevator moves down. the lef aileron moves up, right aileron moves down, elevator down. the lef aileron moves down, right aileron moves up, elevator moves up.

to increase the eel in the control circuit. to correct or adverse aileron yaw. to allow or up-float in flight to bring the aileron into the streamlined position. to give a higher CLMAX or take-off.

Questions 15.

11

The tailplane shown has inverted camber.

To cause the aircraf to pitch nose-up:

a. b. c. d. 16.

I an aileron is moved downward:

a. b. c. 17.

c.

s n o i t s e u Q

use o the rudder control. operating the ailerons slightly in the opposite sense once the correct angle o bank has been reached. increasing the nose-up pitch by using the elevators.

give a yawing moment which opposes the turn. reduce the yawing moment which opposes the turn. give a pitching moment to prevent the nose rom dropping in the turn. improve the rate o roll.

When displacing the ailerons rom the neutral position:

a. b. c. d. 21.

1 1

The purpose o a differential aileron control is to:

a. b. c. d. 20.

the lef pedal is moved orward, and the rudder moves right. the right pedal is moved orward and the rudder moves lef. the lef pedal is moved orward and the rudder moves lef.

The higher speed o the upper wing in a steady banked turn causes it to have more lif than the lower wing. This may be compensated or by:

a. b.

19.

the stalling angle o that wing is increased. the stalling angle o that wing is decreased. the stalling angle is not affected but the stalling speed is decreased.

When rudder is used to give a coordinated turn to the lef:

a. b. c. 18.

the control column must be pushed orward. the control column must be pulled backwards. the control wheel must be rotated. the incidence o the tailplane must be decreased because the negative camber will make it effective in the reverse sense.

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 o aerodynamic balance on a flying control is:

a. b. c. d.

to get the aircraf into balance. to prevent flutter o the flying control. to reduce the control load to zero. to make the control easier to move.

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Questions 22.

A horn balance on a control surace is:

a. b. c. d. 23.

An aileron could be balanced aerodynamically by:

a. b. c. d. 24.

Q u e s t i o n s

25.

trim the aircraf. 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 eel to the controls.

When the control column is pushed orward a balance tab on the elevator:

a. b. c. d.

358

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 o an anti-balance tab is to:

a. b. c. d. 28.

fitting a balance tab. attaching a weight acting orward o the hinge line. attaching a weight acting on the hinge line. attaching a weight acting behind the hinge line.

I the control wheel is turned to the right, a balance tab on the port aileron should:

a. b. c. d. 27.

a sudden increase in stick orce. a sudden reduction then reversal o stick orce. a sudden loss o effectiveness o the controls. a gradual increase in stick orce with increasing IAS.

A control surace is mass balanced by:

a. b. c. d. 26.

making the up aileron move through a larger angle than the down aileron. attaching a weight to the control surace orward o the hinge. having the control hinge set back behind the control surace leading edge. having springs in the control circuit to assist movement.

Control overbalance results in:

a. b. c. d.

1 1

an arm projecting upward rom the control surace to which the control cables are attached. a projection o the outer edge o the control surace orward o the hinge line. a rod projecting orward rom the control surace with a weight on the end. a projection o the leading edge o the control surace below the wing undersurace.

will move up relative to the control surace. will move down relative to the control surace. will only move i the trim wheel is operated. moves to the neutral position.

Questions 29.

The purpose o a spring tab is:

a. b. c. d. 30.

b. c. d.

s n o i t s e u Q

moved orward 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.

greater than that rom an elevator. the same as that rom an elevator. less than that rom an elevator.

Following re-trimming or straight and level flight because o orward CG movement:

a. b. c. d. 35.

1 1

To achieve the same degree o longitudinal trim, the trim drag rom a variable incidence trimming tailplane would be:

a. b. c. 34.

the rudder trim tab will move right and the rudder lef. the trim tab will move lef and the rudder right. the trim tab will move lef and the rudder remain neutral. the trim tab will move right and the rudder remain neutral.

To trim an aircraf which tends to fly nose heavy with hands off, the top o the elevator trim wheel should be:

a.

33.

to assist the pilot in initiating movement o the controls. to zero the load on the pilots controls in the flight attitude required. to provide eel to the controls at high speed. to increase the effectiveness o the controls.

To re-trim afer ailure o the right engine on a twin-engine aircraf:

a. b. c. d. 32.

to maintain a constant tension in the trim tab system. to increase the eel in the control system. to reduce the pilot’s effort required to move the controls against high air loads. to compensate or temperature changes in cable tension.

The purpose o a trim tab is:

a. b. c. d. 31.

11

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 aircraf has a tendency to fly right wing low with hands off. It is trimmed with a tab on the lef aileron. The trim tab will:

a. b. c. d.

move up, causing the lef aileron to move up and right aileron to move down. move down, causing the lef aileron to move up, right aileron remains neutral. move down causing the lef aileron to move up, and right aileron to move down. move up causing the lef wing to move down, ailerons remain neutral.

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11

Questions 36.

An aircraf takes off with the elevator control locks still in position. It is ound to be nose heavy and:

a. b. c. d. 37.

On a servo tab operated elevator, i the pilot’s control column is pushed orward in flight:

a. b. c. d. 38.

Q u e s t i o n s

39.

c. d.

air brakes. lif dumpers. lateral control. all o the above.

Spoilers, when used or roll control, will:

a. b. c. d.

360

to give a nose-down pitching moment. to reduce the lif and so put more weight on the wheels, making the brakes more effective. to cause drag and increase the lif rom the flaps. to reduce the touchdown speed.

Wing mounted spoiler suraces may be used as:

a. b. c. d. 42.

to move up and the lef aileron to move down. to move down and the lef aileron to move down. to move down and the lef aileron to move up. to move up and the right aileron to move down.

Spoilers on the upper surace o the wing may be used on landing:

a. b.

41.

the control suraces and servo tabs are ree. the control suraces are ree but there could be locks on the servo tabs. there could be locks on the control suraces and on the servo tabs. the servo tabs are ree but there could be locks on the control suraces.

In a servo operated aileron control system, turning the cockpit control wheel to the right in flight will cause the servo tab on the lef 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.

I a cockpit control check is made on an aircraf with servo operated controls, and it is ound that the cockpit controls move ully and reely in all directions:

a. b. c. d.

1 1

backward movement o the trim wheel would increase nose heaviness. it would not be possible to move the trim wheel. backward movement o the trim wheel would reduce nose heaviness. operating the trim wheel would have no effect.

reinorce the boundary layer. create turbulence at the wing root. increase the camber at the wing root. decrease lif on the upper wing surace when deployed asymmetrically.

Questions 43.

On an aircraf fitted with roll control spoilers, a roll to port is achieved by:

a. b. c. d. 44.

11

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 ully power operated flying control system control eel is provided by:

a. b. c. d.

the riction in the control cable system. an artificial eel unit (Q - Feel). the aerodynamic loads on the control surace. the mass balance weights.

1 1

s n o i t s e u Q

361

11

Answers

Answers

1 1

A n s w e r s

362

1 d

2 c

3 c

4 c

5 b

6 b

7 d

8 c

9 d

10 b

11 c

12 b

13 a

14 c

15 b

16 b

17 c

18 b

19 b

20 b

21 d

22 b

23 c

24 b

25 b

26 a

27 d

28 a

29 c

30 b

31 a

32 b

33 c

34 a

35 c

36 a

37 d

38 d

39 a

40 b

41 d

42 d

43 c

44 b

Chapter

12 Flight Mechanics

Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 365 Straight Horizontal Steady Flight. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 365 Tailplane and Elevator . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 366 Straight Steady Climb . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 368 Climb Angle . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 369 Effect o Weight, Altitude and Temperature. . . . . . . . . . . . . . . . . . . . . . . . . . . . . 369 Power-on Descent. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 370 Emergency Descent . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 371 Glide . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

372

Rate o Descent in the Glide . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 374 Turning. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 374 Flight with Asymmetric Thrust . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 384 Summary o Minimum Control Speeds . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 395 Questions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

398

Answers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

404

363

12

1 2

F l i g h t M e c h a n i c s

364

Flight Mechanics

Flight Mechanics

12

Introduction Flight Mechanics is the study o the orces acting on an aircraf in flight and the response o the aircraf to those orces. For an aircraf to be in steady (unaccelerated) flight, the ollowing conditions must exist: • the orces acting upward must exactly balance the orces acting downward, • the orces acting orward must exactly balance the orces acting backward, and • the sum o all moments must be zero. This condition is known as equilibrium.

Straight Horizontal Steady Flight In straight and level flight there are our orces acting on the aircraf: LIFT, WEIGHT, THRUST and DRAG, as shown in Figure 12.1. Weight acts through the aircraf centre o gravity (CG), vertically downwards towards the centre o the earth. Alternatively, weight can be defined as acting parallel to the orce o gravity.

2 1

s c i n a h c e M t h g i l F

Lif acts through the centre o pressure (CP), normal (at 90°) to the flight path. For the purposes o this chapter (although not strictly true), thrust acts orwards, parallel to the flight path and drag acts backwards, parallel to the flight path.

L

AERODYNAMIC DRAG

THRUST REQUIRED TO BALANCE AERODYNAMIC DRAG

D

T

FLIGHT PATH

W Figure 12.1 Forces in level flight

365

12

Flight Mechanics For an aircraf to be in steady level flight, a condition o equilibrium must exist. This unaccelerated condition o flight is achieved with the aircraf trimmed with lif equal to weight and the throttles set or thrust to equal drag. It can be said that or level flight the opposing orces must be equal. The L/D ratio o most modern aircraf is between 10 and 20 to 1. That is, lif is 10 to 20 times greater than drag. The lines o action o thrust and drag lie very close together, so the moment o this couple is very small and can be neglected or this study. The position o the CP and CG are variable and under most conditions o level flight are not coincident. The CP moves orward with increasing angle o attack and the CG moves with reduction in uel. Generally, the CP is orward o the CG at low speed, giving a nose-up pitching moment and behind the CG at high speed, giving a nose-down pitching moment.

Tailplane and Elevator The unction o the tailplane is to maintain equilibrium by supply the orce necessary to counter any pitching moments arising rom CP and CG movement. With the CP behind the CG during normal cruise, as illustrated in Figure 12.2, the tailplane must supply a downorce.

1 2


L

MOMENT DUE TO LIFT / WEIGHT COUPLE

TAIL MOMENT

D

T

TAIL DOWN FORCE

W Figure 12.2 Tailplane maintains equilibrium

366

Flight Mechanics

12

Balance of Forces I the tailplane is producing a balancing orce, this will add to or subtract rom the lif orce. For a down load: For an up load:

Lif - tailplane orce = Weight Lif + tailplane orce = Weight

STALL ANGLE ANGLE OF ATTACK

CONSTANT LIFT

I AS

Vs

2 1

Figure 12.3 Variation o angle o attack with IAS


For steady level flight at a constant weight, the lif orce required will be constant. At a steady speed the wing will give this lif at a given angle o attack. However, i the speed is changed, the angle o attack must change to maintain the same lif. As the lif changes with the square o the speed, but in direct proportion to the angle o attack, the angle o attack will vary as shown in Figure 12.3 to give a constant lif. For steady level flight at a constant speed, the thrust must equal the drag. Drag increases with speed (above VMD) and so to maintain a higher speed, the thrust must be increased by opening the throttle.

C THRUST

T2

B

AND DRAG

T1 A

I AS

Figure 12.4 Balance o thrust & drag

To fly at the speed at point A, Figure 12.4, requires a thrust o T 1 and to fly at point B requires a thrust o T2. I the thrust is increased rom T1 to T2 when the aircraf is at point A, the thrust will be greater then the drag, and the aircraf will accelerate in proportion to the ‘excess’ thrust AC until it reaches point B, where the thrust and the drag are again equal. I T 2 is the thrust available with the throttle ully open, then the speed at B is the maximum speed achievable in level flight.

367

12

Flight Mechanics Straight Steady Climb Consider an aircraf in a straight steady climb along a straight flight path inclined at an angle (γ) to the horizontal. γ (gamma) is the symbol used or climb angle. The orces on the aircraf consist o Lif, normal to the flight path; Thrust and Drag, parallel to it; and Weight, parallel to the orce o gravity. This system o orces is illustrated in Figure 12.5. THRUST REQUIRED TO BALANCE AERODY NAMIC DRAG

L

AERODYNAMIC DRAG

FLIGHT PATH

1 2

W cos


W

CLIMB ANGLE BACKWARDS COMPONENT OF WEIGHT

W sin

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 Lif (W cos γ) and the other acting in the same direction as Drag (W sin γ), backwards along the flight path. The requirements or equilibrium are: Thrust must equal the sum o Drag plus the backwards component o Weight; and Lif must equal its opposing component o Weight. For equilibrium at a greater angle o climb, the Lif required will be less, and the backwards component o Weight will be greater. L = W cos γ T = D + W sin γ In a straight steady climb, Lif is less than Weight because Lif only has to suppor t a proportion o the weight, this proportion decreasing as the climb angle increases. (In a vertical climb no lif is required). The remaining proportion o Weight is supported by engine Thrust.

368

Flight Mechanics

12

It can be seen that or a straight steady climb the Thrust required is greater than Drag. This is to balance the backward component o Weight acting along the flight path. T - D W The ability o an aircraf to climb depends upon EXCESS THRUST, available afer opposing aerodynamic drag. The smaller the Drag or a given Thrust, the greater the ability to climb. Drag will be less with flaps up, giving a larger climb angle (improved climb gradient). Sin γ =

Climb Angle Climb angle depends on “excess Thrust” ( T - D ) and the Weight. As both Thrust and Drag vary with IAS, excess Thrust will be greatest at one particular speed. This is the speed or maximum angle o climb, VX. (see Figure 12.28 or the propeller case).

DRAG THRUST A ND

2 1

DRAG

THRUST (JET)


MAXIMUM DIFFERENCE BETW EEN THRUST AND DRAG

V

IAS

X

Figure 12.6 Variation o excess thrust with speed (JET)

The variation o Thrust with speed will depend on the type o engine. For a jet engine, where Thrust is airly constant with speed, V X will be near to V MD, but or a propeller engined aircraf VX will usually be below V MD.

Effect of Weight, Altitude and Temperature. The Drag o an aircraf at a given IAS is not affected by altitude or temperature, but higher Weight will increase Drag and reduce excess Thrust and, consequently, the climb angle. Thrust available rom the engine decreases with increasing altitude and increasing temperature, which also reduces excess Thrust. Climb angle thereore decreases with increasing Weight, altitude and temperature.

369

12

Flight Mechanics Power-on Descent TOTAL REACTION

L = W cos

ENGINE THRUST

D FLIGHT PATH

1 2

FORWARD COMPONENT


OF WEIGHT

( W sin

)

W

Figure 12.7 Forces in a power-on descent

Figure 12.7 illustrates the disposition o orces in a steady Power-on descent. The orce o Weight is split into two components. One component (W cos γ) acts perpendicular to the flight path and is balanced by Lif, while the other component (W sin γ) acts orward along

the flight path and ‘adds’ to the Thrust to balance Drag. I the nose o the aircraf is lowered with a constant Thrust setting, the increased component o Weight acting orward along the flight path will cause an increase in IAS. The increased IAS will result in an increase in Drag which will eventually balance the increased orward orce o Weight and equilibrium will be re-established. I the throttle is closed, the orce o Thrust is removed, and a larger orward component o Weight must be provided to balance Drag and maintain a constant IAS. This is accomplished by lowering the nose o the aeroplane to increase the descent angle ( γ). • In a descent Lif is less than Weight. This is because Lif only has to balance the component o Weight perpendicular to the flight path (W cos γ). • In a descent Thrust is less than Drag. This is because Weight is giving a orward component in the same direction as Thrust (W sin γ).

370

Flight Mechanics

12

Emergency Descent In the event o cabin pressurization ailure at high altitude it is necessary to descend as quickly as possible. The rate o descent can be increased by: • Reducing Thrust by closing the throttles. • Increasing Drag by: • extending the speedbrakes, • lowering the landing gear (at or below VLO). • Increasing speed by lowering the nose. Speed can be increased in the clean configuration up to MMO or VMO depending on the altitude, or to the gear extended limit speed (V LE) i the gear is down. The overall rate o descent will be higher with the landing gear extended (lots o Drag), but i the gear operating limit speed (V LO) is much less than the cruising speed, the aircraf will have to be slowed down beore the gear can be lowered (perhaps taking several minutes in level flight). So the initial rate o descent will be relatively low and the time spent at high altitude will be extended.

2 1


I the gear is not extended, throttles can be closed, speedbrakes extended and the nose lowered to accelerate the aircraf to M MO /VMO immediately, giving a higher initial rate o descent and getting the passengers down to a lower altitude without delay. At high altitude the limiting speed will be M MO, and i an emergency descent is made at this Mach number, the IAS will be increasing. At some altitude the IAS will reach V MO, and the nose must then be raised so as not to exceed V MO or the remainder o the descent. The rate o descent possible during an emergency descent can be quite high, so as the required level-off altitude is approached, the rate o descent should be reduced progressively so as to give a smooth transition back to level flight.

371

12

Flight Mechanics Glide In a glide without Thrust, the Weight component along the flight path must supply the propulsive orce and balance Drag. In a glide there are only three orces acting on the aircraf: Lif, Weight and Drag.

TOTAL REACTION

L = W cos

D 1 2

FLIGHT PATH


FORWARD COMPONENT OF WEIGHT

( W sin

)

W Figure 12.8 Forces in the glide

Figure 12.8 shows the disposition o orces in a steady glide. The orward component o Weight (W sin γ) is a product o descent angle ( γ); the greater the descent angle, the greater the orward component o weight (compare with Figure 12.7 ). The orward component o

weight must balance Drag or the aircraf to be in a steady glide. It ollows that i Drag is reduced and Lif remains constant, the required balance o orces can be achieved at a smaller descent angle.

Angle of Descent in the Glide Glide angle is a unction ONLY o the L/D ratio . The descent (glide) angle will be least when

the L/D ratio is the greatest. L/D ratio is a maximum at the optimum angle o attack, and this also corresponds to the minimum drag speed (V MD), Figure 12.10. At speeds above or below VMD the glide angle will be steeper. Maximum distance in a glide can be achieved when the aircraf is flown at L/D

372

(VMD).

MAX

Flight Mechanics

12

Effect of Weight L/D

MAX

is independent o weight . Provided the aircraf is flown at its optimum angle o

attack, the glide angle and glide distance will be the same whatever the weight. The speed corresponding to the optimum angle o attack, (V MD), will, however, change with weight. VMD increases as weight increases.

L

D 2 1


FLIGHT PATH

W Figure 12.9 Increased weight: no effect on glide range

As illustrated in Figure 12.9, a higher weight will give an increased orward component o weight and the aircraf will accelerate towards the resultant higher V MD. As the aircraf accelerates, lif increases and drag will increase until it balances the increased orward component o Weight. Equilibrium is now re-established at the same L/D MAX, but a higher IAS. At a higher weight the aircraf will glide the same distance but at a higher speed, and consequently it will have an increased RATE o descent.

373

12

Flight Mechanics Effect of Wind The glide angle will determine the distance that the aircraf can glide or a given change o height. GLIDE DISTANCE = HEIGHT LOSS ×

LIFT (L) DRAG (D)

This distance would only be achieved in still air. I 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.

Effect of Configuration The maximum L/D ratio o an aircraf will be obtained in the clean configuration. Extension o flaps, spoilers, speedbrakes or landing gear etc. will reduce L/D MAX and give a steeper glide angle, thus reducing glide range.

Rate of Descent in the Glide Minimum rate o descent in the glide is obtained at the IAS which produces minimum Power Required (VMP). Flying at VMP in a glide will enable the aircraf to stay airborne or as long as possible. As shown in Figure 12.10, VMP is a slower IAS than V MD. Wind speed and direction has no effect on rate o descent. A requently used method o showing the relationship o V MD and VMP is by use o the ‘whole aeroplane C L /CD polar’ curve, illustrated in Figure 12.11.

1 2


CL

C LMAX V MP

DRAG

L/D MAX (V MD )

L / D MAX

1.32 V MD

V

MP

V MD

Figure 12.10

IAS

CD

Figure 12.11

Turning For an aircraf to change direction, a orce is required to deflect it towards the centre o the turn. This is called the centripetal orce, Figure 12.12. Banking the aircraf inclines the lif. It is the horizontal component o lif which causes the aircraf to turn . I the aircraf is banked and the angle o attack kept constant, the vertical component o lif will be too small to balance the weight and the aircraf will start to descend.

374

Flight Mechanics

12

As the angle o bank increases, the angle o attack must be increased to bring about a greater total lif. The vertical component must be large enough to maintain level flight, while the horizontal component is large enough to produce the required centripetal orce.

Effect of Weight on Turning In a steady level turn, i thrust is ignored, lif provides a orce to balance weight and centripetal orce to turn the aircraf. I the same TAS and angle o bank can be obtained, the radius o turn is basically independent o weight or the aircraf type . Not all aircraf can reach the same angle o bank at the same TAS. I weight increases, the vertical component o lif required increases, but the centripetal orce to maintain the same radius o turn also increases in the same proportion. The lif required, although it is greater, has the same inclination to the vertical as beore and the bank angle is the same, Figure 12.13.

Lift

2 1

φ


CENTRIPETAL FORCE

Weight Figure 12.12 Forces in a turn

Lift

φ CENTRIPETAL FORCE

Increased Weight

Figure 12.13 Bank angle & weight

375

12

Flight Mechanics

L

CENTRE OF TURN

L cos φ

φ CENTRIPETAL FORCE

L sin φ

r W 1 2

Figure 12.14 Forces acting in a steady turn


In a steady horizontal turn, Figure 12.14, the conditions o equilibrium can be expressed in the orm: L cos φ = W

(Eq 12.1)

WV rg

2

L sin φ =

(Eq 12.2)

where (L) is the wing lif in newtons, (W) is the weight o the aircraf in newtons, (V) the true airspeed in m/s, (r) the radius o turn in metres, φ the angle o bank and (g) the acceleration o gravity constant o 9.81 m/s 2. Dividing equation 2 by equation 1 we get: V rg 2

tan φ =

(Eq 12.3)

which is the basic turning equation relating (V), (r) and φ. Once two o these variables are known, the other one can be determined. From equation 3, the radius o turn is given by: Turn Radius

376

=

V2 g tan φ

Flight Mechanics

12

and the corresponding rate o turn ( = V / r ) by: Rate o Turn =

g tan φ V

radians / second

Rate o turn is the rate o change o heading or angular velocity o the turn. It may be expressed as degrees per minute, or by a Rate Number. Rate 1 turn is 180° per minute (3° per second) Rate 2 turn is 360° per minute (6° per second) Rate o turn is directly proportional to TAS and inversely proportional to the turn radius. Rate o Turn =

TAS Radius

For example: at a speed o 150 kt TAS (77 m/s), an aircraf perorming a turn with a radius o 1480 metres would have a rate o turn o: 77 1480

2 1

= 0.052 radians / sec

there being 2π radians in a circle,

360 6.286


= 57.3° per radian

0.052 × 57.3 = 3° per second (Rate 1) • At a constant TAS, increasing the angle o bank decreases the turn radius and increases the rate o turn. • To maintain a constant rate o turn, increasing speed requires an increased bank angle. • At a constant bank angle, increasing speed increases the turn radius and decreases the rate o turn.

In a Constant Rate Turn The Angle o Bank is Dependent Upon TAS

377

12

Flight Mechanics Radius and Rate of Turn Two variables determine the rate o turn and radius o turn: • Bank angle (φ). A steeper bank reduces turn radius and increases the rate o turn, but produces a higher load actor. • True airspeed (TAS): Reducing speed reduces turn radius and increases the rate o turn, without increasing the load actor. The radius o turn, at any given bank angle (φ), varies directly with the square o the TAS: V g tan φ 2

Radius

=

I speed is doubled, the turn radius will be our times greater, at a constant bank angle.

To appreciate the relationship between radius o turn and rate o turn at double the speed, consider: 1 2


Rate o Turn =

V Radius

Rate o Turn =

V (×2) Radius (×4)

=

1 2

I speed is doubled, the rate o turn will be hal o its previous value, at a constant bank angle.

Because the rate o turn varies with TAS at any given bank angle, slower aeroplanes require less time and area to complete a turn than aster aeroplanes with the same bank angle, Figure 12.15. A specific angle o bank and TAS will produce the same rate and radius o turn regardless o weight, CG position, or aeroplane type. It can also be seen rom Figure 12.15 that increasing speed increases the turn radius and decreases the rate o turn. The load actor remains the same because the bank angle has not changed . To increase the rate and decrease the radius o turn, steepen the bank and/or decrease the speed. A given TAS and bank angle will produce a specific rate and radius o turn in any aeroplane. In a co-ordinated level turn, an increase in airspeed will increase the radius and decrease the rate o turn. Load actor is directly related to bank angle, so the load actor or a given bank angle is the same at any speed .

378

Flight Mechanics

BANK

80º

ANGLE FOR RATE

70º

OF TURN

60º

90 000

90

80 000

80

70 000

70

60 000

60

50 000

50

40 000

40

30 000

30

20 000

20

10 000 9000

10 9

8000

8

7000

7

6000

6

5000

5

4000

4

3000

3

2000

2

1000 900

1.0 0.9

800

0.8

700

0.7

600

0.6

500

0.5

400

0.4

300

0.3

200

0.2

12

50º 40º 30º 20º

10º

2 1


10º 20º 30º 40º 50º 60º 70º 80º

BANK ANGLE FOR TURN RADIUS

TAS, Kts

Figure 12.15

(For illustration purposes only). This chart will work or any aeroplane. The example shows that or a turn at 130 kt TAS and a bank angle o 20°, the radius will be 4200 f and the rate o turn will be 3° per second. At 260 kt TAS the radius will be 16 800 f and the rate o turn will be 1.5° per second.

379

12

Flight Mechanics Load Factor in the Turn When an aircraf is in a banked turn, lif must be increased so as to maintain the vertical component o lif equal to weight, Figure 12.16 . INCREASED LIFT

L

CENTRIPETAL FORCE

W

1 2

60


BAN K A NGLE

30 BANK A NGLE

Figure 12.16 Increased lif required in a turn

This relationship may be expressed as: Load Factor (n) =

L W

1 = cos φ

= sec φ

Reer to Chapter 7 or the ull trigonometrical explanation. Figure 12.17 shows the relationship between load actor and bank angle. This chart will be

effective or any aircraf. It can be seen that load actor (n) increases with bank angle at an increasing rate. Load actor in the turn is a unction ONLY o bank angle

Constant Bank Angle, Constant Load Factor

380

Flight Mechanics

12

9 8 7 6 5 4 3 2 1 0 0

10

20

30

40

50

60

70

80

90

BANK ANGLE IN DEGREES

Figure 12.17 Relationship between ‘g’ & bank angle

‘g’ Limit on Turning

2 1

For each aircraf there is a design limit load actor. For modern high speed jet transport aircraf the positive limit load actor is 2.5g. From Figure 12.17 it can be seen that this would occur at a bank angle o 67°, and this will determine a turn radius, depending on the TAS. This will be the minimum radius permissible at that ‘g’ i the strength limit is not to be exceeded.


Stall Limit on Turning I speed is kept constant, but the bank angle increased, the angle o attack must also be increased to provide the increased lif required. Eventually the stalling angle will be reached, and no urther increase in bank angle (and decrease in turn radius) is possible. Because the stalling speed varies with weight, this boundary will be a unction o weight.

Thrust Limit on Turning During a turn lif 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 or level flight at the same speed. The greater the bank angle, the greater will be the thrust required, and eventually the throttle will be ully open. No urther increase in bank angle (and decrease in turn radius) is then possible. The relative positions o the thrust boundary and the strength boundary will depend on the limit load actor and thrust available.

Minimum Turn Radius I the thrust available is adequate, the minimum radius o turn occurs at the intersection o the stall limit and the strength limit. The speed at this point is V A, the maximum manoeuvring speed. The heavier the aircraf, the greater the minimum radius o turn.

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Flight Mechanics Turn Co-ordination Adverse aileron yaw, engine torque, propeller gyroscopic precession, asymmetric thrust and spiral slipstream all give the possibility o unco-ordinated flight. Unco-ordinated flight exists when the aircraf is sideslipping. Indication o sideslip is given to the pilot by the inclinometer portion (ball) o the turn co-ordinator, Figure 12.18. The miniature aeroplane indicates rate o turn.

MINIATURE AEROPLANE

LEV EL INDEX

L

2 MIN

R

RATE ONE TURN INDEX INCLINOMETER

Figure 12.18 Turn co-ordinator

1 2

Co-ordinated flight is maintained by keeping the ball centred between the reerence 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 useul way to remember which rudder to apply.


I aileron and rudder are co-ordinated during a turn, the ball will remain centred and there will be no sideslip. I the aircraf is sideslipping, the ball moves away rom the centre o the tube. Sideslipping towards the centre o the turn moves the ball to the inside o the turn. Sideslipping towards the outside o the turn moves the ball to the outside o the turn. To correct or these conditions and maintain co-ordinated flight, “step on the ball.” Bank angle may also be varied to help restore co-ordinated flight rom a sideslip. The ollowing illustrations give examples.

382

Flight Mechanics

L

2 MIN

12

R

Figure 12.19

Figure 12.19 shows the aircraf in a rate 1 co-ordinated turn to the right.

2 1

L

2 MIN


R

Figure 12.20

Figure 12.20 shows the aircraf in an unco-ordinated turn to the right; it will be sideslipping

towards the centre o the turn (slipping turn). Using “step on the ball,” the turn can be coordinated by applying right rudder pressure to centre the ball.

L

2 MIN

R

Figure 12.21

Figure 12.21 also shows the aircraf in an unco-ordinated turn to the right; it will be sideslipping

towards the outside o the turn (skidding turn). Using “step on the ball,” the turn can be coordinated by applying lef rudder pressure.

383

12

Flight Mechanics Flight with Asymmetric Thrust Introduction When an engine ails on a multi-engine aircraf there will be a decrease in thrust and a n increase in drag on the side with the ailed engine: • airspeed will decay • the nose will drop and • most significantly, there will be an immediate yawing moment towards the ailed (dead) engine. Figure 12.22 shows the orces and moments acting on an aircraf ollowing ailure o the lef

engine. The aircraf has a yawing moment towards the dead engine. The pilot has applied rudder to stop the yaw. The vital action when an engine ails is to STOP THE YAW !

Yawing Moment The yawing moment is the product o thrust rom the operating engine multiplied by the distance between the thrust line and the CG (thrust arm), plus the drag rom the ailed engine multiplied by the distance between the engine centre line and the CG. The strength o the yawing moment will depend on:

1 2


• how much thrust the operating engine is developing (throttle setting and density altitude). • the distance between the thrust line and the CG (thrust arm). • how much drag is being produced by the ailed engine. The rudder moment, which balances the yawing moment, is the result o the rudder orce multiplied by the distance between the fin CP and the CG (rudder arm). This statement will be modified by actors yet to be introduced. Thus, at this preliminary stage, the ability o the pilot to counteract the yawing moment due to asymmetric thrust will depend on: • rudder displacement (affecting rudder orce). • CG position (affecting rudder arm). • the IAS (affecting rudder orce).

384

Flight Mechanics

12

Assume the rudder is at ull deflection, CG is at the rear limit (shortest rudder arm) and the IAS (dynamic pressure) is just sufficient or the rudder orce to give a rudder moment equal to the yawing moment - there will be no yaw. But any decrease in IAS will cause the aircraf to yaw uncontrollably towards the ailed engine. The uncontrollable yaw to the lef, in this example, will cause the aircraf to roll uncontrollably to the lef due to greater lif on the right wing. The aircraf will enter a spiral dive to the lef (impossible to stop with the flight controls alone); i near the ground, disaster will result. In these extreme circumstances near the ground, the ONLY way to regain control o the aeroplane is to close the throttle(s) on the operating engine(s). This removes the yawing moment, and the aircraf can be orce-landed under control. Thus there is a minimum IAS at which directional control can be maintained ollowing engine ailure on a multi-engine aircraf. This minimum IAS is called VMC (minimum control speed). YAW ING MOMENT

THRUST 2 1


DRAG THRUST AR M

RUDDER RUDDER

ARM

FORCE

RUDDER MOMENT

Figure 12.22 Asymmetric thrust

385

12

Flight Mechanics Critical Engine One o the actors influencing the yawing moment ollowing engine ailure on a multi-engine aircraf is the length o the thrust arm (distance rom the CG to the thrust line o the operating engine). In the case o a propeller engine aircraf the length o the thrust arm is determined by the asymmetric effect o the propeller. At a positive angle o attack, the thrust line o a clockwise rotating propeller, when viewed rom the rear, is displaced to the right o the engine centre line. This is because the down-going blade generates more thrust than the up-going blade (Chapter 16 ). I both engines rotate clockwise, the starboard (right) engine will have a longer thrust arm than the port (lef) engine. I the lef engine ails, the thrust o the right engine acts through a longer thrust arm and will give a bigger yawing moment; a higher IAS (V MC) would be necessary to maintain directional control. So at a given IAS, the situation would be more critical i the lef engine ailed, Figure 12.23. The critical engine is the engine, the ailure o which would give the biggest yawing moment.

To overcome the disadvantage o having a critical engine on smaller twins, their engines may be designed to counter-rotate. This means that the lef 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 o a ourengine jet aircraf the critical engine is either o the outboard engines.

1 2


Note: I all the propellers on a multi-engine aircraf rotate in the same direction, they are sometimes called ‘co-rotating’ propellers.

LONGER CRITICAL ENGINE

Figure 12.23 Critical engine

386

THRUST A RM

Flight Mechanics

12

Balancing the Yawing Moments and Forces Although the moments are balanced in Figure 12.22, the orces are not balanced. Consequently, the aircraf is not in equilibrium and will drif, in this case, to the lef. The unbalanced side orce rom the rudder can be balanced in two ways: • with the wings level and • by banking slightly towards the live engine (preerred 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 orce on the keel suraces opposite to the rudder orce, Figure 12.24. I the sideslip angle is too large, the fin could stall . The turn indicator will be central and so will the slip indicator. Note: Asymmetric thrust is the exception to the rule o co-ordinated flight being indicated to the pilot by the ball centred in the inclinometer.

This method o balancing the side orce rom the rudder gives reduced climb perormance because o the excessive parasite drag generated so is not the recommended method or critical situations, such as engine ailure just afer take-off or go-around.

2 1


YAW ING MOMENT

L

SIDE FORCE FROM RUDDER

2 MIN

R

SIDE FORCE FROM SIDESLIP

RUDDER MOMENT

Figure 12.24 Wings level method

The only advantage o the ‘wings level’ method o balancing the orces is the strong visual horizontal reerences available to the pilot, both inside and outside the aircraf. The disadvantages are that i the sideslip angle is too large, the fin could stall plus the ability to climb is reduced due to excessive parasite drag.

387

12

Flight Mechanics

Lift RUDDER FORCE SIDEWAYS COMPONENT OF LIFT

L

2 MIN

R

Weight Figure 12.25 Maximum 5° bank towards live engine 1 2

Rudder to Stop Yaw - Bank Towards Live Engine


It is more aerodynamically efficient to balance the rudder sideorce by banking towards the live engine, Figure 12.25, so that lif gives a sideways component opposite to the rudder orce. The angle o bank must not exceed 5° , to prevent excessive loss o vertical lif component. Banking towards the live engine also reduces the side orce on the fin rom sideslip, which minimizes V MC, effectively reduces the yawing moment and gives more rudder authority to stop the yaw.

The cockpit indication will be the turn needle central with the slip indicator (ball) one hal diameter displaced towards the live engine. The ‘ball’ is not centred, but the aircraf is not sideslipping. This method produces minimum drag and gives the best ability to climb and is thereore the preerred method o putting the aircraf in equilibrium ollowing engine ailure.

Roll and Yaw Moments with Asymmetric Thrust The rolling and yawing moments and the power o the flight controls to balance them will determine the controllability o an aircraf with asymmetric thrust. Rolling and yawing moments with asymmetric thrust are affected by: • Thrust on the live engine The greater the thrust, the greater the yawing moment rom the live engine. The urther the engine is mounted out on the wing (increased thrust arm), the larger the yawing m oment. Thrust is greatest at low speed and ull throttle. • Altitude Thrust reduces with increasing altitude and/or increasing temperature (high density altitude). The worst case or engine ailure is low density altitude, e.g. immediately afer take-off on a cold day at a sea level airport.

388

Flight Mechanics

12

W INDMILLING PROPELLER

DRAG STATIONARY PROPELLER FEATHERED POSITION

FINE P ITCH STOP

0

15

30

45

60

90

PROPELLER BLADE ANGLE

Figure 12.26 Propeller drag

• Drag rom the dead engine and propeller Drag rom the dead engine always adds to the yawing moment. The size o the contribution depends upon whether the propeller is windmilling, stopped or eathered, Figure 12.26 . This effect will be absent on an aircraf powered by jet engines.

2 1


• Drag rom a windmilling propeller is high. It is being driven by the relative airflow and is generating both drag and torque. • I a propeller is stationary, it is generating drag but no torque. Drag rom a stationary propeller is less than rom one which is windmilling. • A eathered 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. • Asymmetric blade effect (also known as ‘P’ Factor) I both engines rotate clockwise, the right engine has a longer thrust arm. Failure o the lef engine will give a larger yawing moment. This effect will be absent on an aircraf with counter-rotating propellers, contra-rotating propellers or jet engines. • CG position The aircraf rotates about the CG. The ore and af CG location has no effect on the yawing moment rom a ailed engine, but will influence the rudder arm, hence the rudder moment. CG on the af limit will give the smallest rudder arm and the least ability to oppose the yawing moment rom a ailed engine. Note: Contra-rotating propellers are mounted on the same shaf and are driven in opposite directions, usually by the same engine

389

12

Flight Mechanics • Torque reaction When the engine turns the propeller, the equal and opposite reaction tries to turn the engine in the other direction. Following ailure o one engine on an aircraf with propellers which rotate in the same direction (usually clockwise when viewed rom the rear), the torque tries to roll the aircraf to the lef. Failure o the lef engine thereore gives the biggest rolling moment to the lef. With counter-rotating engines, both the asymmetric blade effect (P Factor) and the torque reaction are minimized, and there is no longer a critical engine. This effect will be absent on an aircraf powered by jet engines. • Difference in lif due to slipstream Engine ailure on one side will give a loss o induced lif rom the propeller slipstream on that side. Total lif 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 i the trailing edge flaps are deployed because o the higher initial C L. This effect will be absent on an aircraf powered by jet engines. • Rolling moment due to sideslip I the aircraf is flying with yaw to balance the rudder orce, there will be a sideslip. In Figure 12.24 the aircraf is sideslipping to the lef. The dihedral o the lef wing (with the dead engine) will cause the lif o the lef wing to increase, which will compensate some o the lif loss due to the loss o the propeller slipstream.

1 2


• Weight Any weight increase will require a higher angle o attack at a given speed. • This will increase the asymmetric blade effect (P Factor) and give a bigger yawing moment. • The fin and rudder will be masked to a greater extent by disturbed airflow rom the wing and uselage, making the rudder and fin less effective; consequently, the available rudder moment will be reduced. • Airspeed The effectiveness o the flying controls depends upon dynamic pressure, assuming ull control displacement. An accurate measure o dynamic pressure at low airspeeds is given by the Calibrated Airspeed (CAS). CAS is IAS corrected or position error. At low airspeed / high CL the pressures sensed by the pitot / static system are affected by the high angle o attack, so must be compensated to make the IAS reflect a more accurate measure o dynamic pressure. A higher IAS means more control effectiveness and consequently a larger available rudder moment to balance the yawing moment rom the ailed engine. A lower IAS will reduce the available rudder moment i the other parameters remain the same. IAS is the vital element in control o the aircraf with asymmetric thrust .

390

Flight Mechanics

12

Minimum Control Airspeed It has been shown that when a multi-engine aircraf suffers an engine ailure 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 IAS (V MC), below which it is impossible or the pilot to maintain directional control with asymmetric thrust. Airworthiness Authorities, in this case the EASA, have laid down conditions which must be satisfied when establishing the minimum airspeeds or inclusion in the Flight Manual o a new aircraf type. As in most other cases, the conditions under which the minimum control airspeeds are established are ‘worst case’. A actor o saety is built into these speeds to allow or aircraf age and average pilot response time. Because there are distinct variations in the handling qualities o the aircraf when in certain configurations, minimum control airspeed (V MC) has three separate specifications: • VMCA • VMCG • VMCL

Minimum control speed - airborne. Minimum control speed - on the ground. Minimum control speed - in the landing configuration.

V MCA (CS 25.149 paraphrased)

2 1

VMCA is the calibrated airspeed at which, when the critical engine is suddenly made inoperative, it is possible to maintain control o the aeroplane with that engine still inoperative and maintain straight flight with an angle o bank o not more than 5°.


VMCA may not exceed 1.13V SR with: • • • • • • •

maximum available take-off power or thrust on the engines. the aeroplane trimmed or take-off. the most unavourable 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 i applicable, the propeller o the inoperative engine: • windmilling • eathered, i the aeroplane has an automatic eathering device.

The rudder orces required to maintain control at V MCA may not exceed 150 lb nor may it be necessary to reduce power or thrust on the operative engines. Note: There is no perormance requirement, just directional control.

391

12

Flight Mechanics Factors Affecting V MCA Angle o Bank Banking towards the live engine reduces the rudder deflection required and so allows a lower VMCA. 5° maximum is stipulated because larger bank angles would significantly reduce the vertical component o lif; the angle o attack would have to be increased with the added penalty o higher induced drag. CG Position Because the aircraf rotates around the CG, the position o the CG directly affects the length o the rudder arm and, thus, the power o the rudder and fin to maintain directional stability and control. The ‘worst case’ is with the CG at the af limit. I the requirements can be met in this configuration, the ability to maintain directional control will be enhanced at any other CG location. Aileron Effectiveness At low airspeed, dynamic pressure is low which reduces the effectiveness o all the flying controls or a given angle o displacement. This effect on the rudder has already been discussed, but the ailerons will be affected in a similar way. In Figure 12.24 and Figure 12.25 (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 lef side will add to the yawing moment because o its increased induced drag. At low IAS (increased CL ), the large angle o 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.

1 2


Flap Position Flap position affects lif / drag ratio, nose-down pitching moment and the stalling speed. With asymmetric thrust, flaps reduce climb perormance, increase the margin above stall, but do not directly affect V MCA. However, i take-off flap is used, the difference in lif between the two wings due to propeller slipstream is urther increased. This increases the rolling moment, requires increased aileron deflection and indirectly increases V MCA. Undercarriage The undercarriage increases drag and reduces perormance. The increased keel surace in ront o the CG decreases directional stability slightly, thus the fin and rudder are opposed in sideslip conditions, and this will slightly increase V MCA. Altitude and Temperature VMCA is affected by the amount o thrust being developed by the operating engine. As altitude and/or temperature increases, the thrust rom an unsupercharged engine will decrease. Thereore, VMCA decreases with an increase in altitude and/or temperature. Relationship between V S and V MCA VS is constant with increasing altitude, so can be represented by a straight line in Figure 12.27 . (It was shown in Chapter 7 that stall speed does increase at higher altitudes, but or this study, we are only dealing with lower altitudes). Figure 12.27 shows that at about 3000 f, V S and VMCA typically correspond. So above this altitude, the stall speed is higher than V MCA. I the aircraf is slowed ollowing an engine ailure with ull power on the operating engine, the aircraf can stall beore reaching V MCA. The margin above loss o control is reduced; in this case by stalling.

392

Flight Mechanics

ALTITUDE

12

Vs AN D V MCA COINCIDE

Vs

IAS

Figure 12.27 V S and V MCA

VMCG (CS 25.149 paraphrased)

VMCG, the minimum control speed on the ground, is the calibrated airspeed during the takeoff run, at which, when the critical engine is suddenly made inoperative, it is possible to maintain control o the aeroplane using the rudder control alone ( without the use o nose wheel steering) to enable the take-off to be saely continued using normal piloting skill. The rudder control orces may not exceed 150 pounds (68.1 kg) and, until the aeroplane becomes airborne, the lateral control may only be used to the extent o keeping the wings level. In the determination o VMCG, assuming that the path o the aeroplane accelerating with all engines operating is along the centre o the runway, its path rom the point at which the critical engine is made inoperative to the point at which recovery to a direction parallel to the centre line is completed may not deviate more than 30 f (9.144 m) laterally rom the centre line at any point. As with VMCA, this must be established with: • • • •

2 1


maximum available take-off power or thrust on the engines. the aeroplane trimmed or take-off. the most unavourable CG position. maximum sea level take-off weight.

Factors Affecting V MCG Altitude and Temperature VMCG is affected by the amount o thrust being developed by the operating engine. As altitude and/or temperature increases, the thrust rom an unsupercharged engine will decrease. Thereore, VMCG decreases with an increase in altitude and/or temperature.

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Flight Mechanics Nose Wheel Steering Nose wheel steering is designed or 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 nose wheel begins to hydroplane between 70 and 90 knots (depending on tyre pressure and depth o water or slush) and has very little steering effect. Once the aircraf is moving, the nose wheel doesn’t do much except turn sideways and skid. VMCG is established during flight testing, usually on a dry runway. I nose wheel steering were used by the test pilot it would give a alse, low speed at which it was possible to maintain directional control on the ground afer the critical engine is suddenly made inoperative. At this speed on a slippery runway, even i nose wheel steering were used by a line pilot, it would not give the required assistance in maintaining directional control ollowing an engine ailure and the aircraf would depart the side o the runway. The regulations ensure that limits are established in a “worst case” set o circumstances in order to give the maximum saety actor during normal operations.

Rudder Arm When the aircraf is on the ground it rotates about the main undercarriage, which is af o the CG. Thereore the rudder arm is shorter when the aircraf is on the ground. It will be ound that on most aircraf VMCG is higher than V MCA.

1 2


VMCL (CS 25.149 paraphrased)

VMCL, 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 o the aeroplane with that engine still inoperative and maintain straight flight with an angle o bank o not more than 5°. VMCL must be established with: • the aeroplane in the most critical configuration or approach and landing with all engines operating, • the most unavourable CG, • the aeroplane trimmed or approach with all engines operating, • the most unavourable weight, • or propeller aeroplanes, the propeller o the inoperative engine in the position it achieves without pilot action and • go-around power or thrust setting on the operating engines(s).

394

Flight Mechanics

12

In demonstrating V MCL: • the rudder orce may not exceed 150 lb. • the aeroplane may not exhibit hazardous flight characteristics or require exceptional piloting skill, alertness or strength. • lateral control must be sufficient to roll the aeroplane, rom an initial condition o steady flight, through an angle o 20° in the direction necessary to initiate a turn away rom the inoperative engine(s), in not more than 5 seconds.

Factors Affecting V MCL Aileron Effectiveness At low airspeed, dynamic pressure is low which reduces the effectiveness o all the flying controls or a given angle o displacement. This effect 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 lef side will also add to the yawing moment because o its increased induced drag and may stall the wing at low IAS (high CL). Adequate aileron effectiveness is clearly very important when considering V MCL because this minimum control speed contains a roll requirement, not just directional control.

2 1


Summary of Minimum Control Speeds CS 25.149 sets out the criteria to be used when establishing the minimum control speeds or certification o a new aircraf. The speeds so established will be included in the aircraf’s Flight Manual. From careul study o the above extracts, several things can be noted: • nose wheel steering may not be used when establishing V MCG. Its use would artificially decrease V MCG. In service, when operating rom a slippery runway, nose wheel steering would be ineffective, so it might be impossible to directionally control the aircraf when at or above the stated V MCG. • VMCL includes a roll requirement , not merely directional control, as with the other speeds. • 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 thereore decrease with altitude and temperature, and so control can be maintained at a lower IAS. VMC thereore decreases with increasing altitude and temperature (higher density altitude).

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Flight Mechanics Performance with One Engine Inoperative It was shown on page 369 that an aircraf’s ability to climb depends upon the excess thrust available, afer aerodynamic drag is balanced. I a twin-engine aircraf loses an engine, total thrust is reduced by 50%, but the excess thrust (the thrust, minus aerodynamic drag) is reduced by more than 50%, Figure 12.28. The ability to climb may be reduced as much as 80%.

Single-engine Angle of Climb Angle o climb is determined by excess thrust available . Climb angle will be a maximum

when the aircraf is flown at the IAS where excess thrust is a maximum (maximum thrust to drag ratio). Since thrust decreases with orward speed and total drag increases below and above the minimum drag speed (VIMD), the best angle o climb is achieved at a speed below V IMD but a sae margin above the stall speed. The airspeed or maximum angle o climb is V X or all engines operating and V XSE or best single-engine angle o climb.

Single-engine Rate of Climb Rate o climb is determined by excess power available . Power is the rate o doing work and

work is orce times distance moved, so power is orce times distance moved in a given time, i.e. thrust or drag times TAS (thrust or drag because they are both orces and TAS because it is the only speed there is!). Although thrust reduces with orward speed, total power available increases because o the speed actor. Similarly, power required is a measure o drag times TAS, so excess power available determines the available rate o climb. The airspeed or best rate o climb is V Y or all engines operating and V YSE or best single-engine rate o climb.

1 2


VY and VYSE are higher than V X and VXSE and provide a saer margin above both stall and V MCA.

Under most circumstances V Y and VYSE are the best speeds to use. On small twin-engine aircraf VYSE is marked on the Airspeed Indicator by a blue radial line and is called ‘blue line speed’.

Conclusions At a given altitude, airspeed and throttle position, excess thrust depends on the amount o drag being generated, and this will depend on configuration, weight and whether turns are required to be made. The control surace deflections required to balance asymmetric thrust will also cause an increase in drag. It is essential, thereore, that afer losing an engine, particularly during take-off or during a go-around, drag is reduced and no turns are made until well away rom the ground. Drag can be reduced by eathering the propeller o the inoperative engine, raising the undercarriage, careully raising the flaps, closing the cowl flap on the inoperative engine and banking the aircraf no more than 5° towards the operating engine. Flying at V YSE (blue line speed) with maximum continuous thrust on the operating engine will provide maximum climb perormance and optimum control over the aeroplane.

396

Flight Mechanics

12

THRUST

THRUST REQUIRED (DRAG)

EXCESS

EXCESS

MINIMUM DRAG

THRUST AVAILABLE

VS

VX 65

70

V MCA

V XSE

V I MD

75

I AS

85 SAFETY MARGI N

2 1


BOTH ENGINES ONE ENGINE

POWER AVAI LABLE

POWER VS

POWER REQUI RED EXCESS POWER AVAI LABLE EXCESS

MI NI MUM POWER

VY V YSE

V I MD 65

70

85

I AS

V MCA SAFETY MARGI N

Figure 12.28 Excess thrust & excess power.

397

12


In straight and level powered flight the ollowing principal orces act on an aircraf:

a. b. c. d. 2.

For an aircraf in level flight, i the wing CP is af o the CG and there is no thrust/ drag couple, the tailplane load must be:

a. b. c. d. 3.

Q u e s t i o n s

4.

flaps partially extended and at best rate o climb speed (V Y). flaps partially extended and at best angle o climb speed (V X). flaps retracted and at best rate o climb speed (V Y). flaps retracted and at best angle o climb speed (V X).

The angle o climb is proportional to:

a. b. c. d.

398

thrust line acts horizontally and above the orce o drag. centre o gravity is located orward o the centre o pressure. centre o pressure is located orward o the centre o gravity. orce o drag acts horizontally and above the thrust line.

To give the best obstacle clearance on take-off, take-off should be made with:

a. b. c. d. 7.

the main plane lif is always positive. the lif/weight and thrust/drag couples combine to give a nose-down pitch. the lif produced is greater than required at high speed. this configuration gives less intererence.

The reason a light general aviation aircraf tends to nose-down during power reduction is that the:

a. b. c. d. 6.

Weight acts vertically toward the centre o the Earth. Lif acts perpendicular to the chord line and must be greater than weight. Thrust acts orward parallel to the relative wind and is greater than drag. Lif acts in the same direction to the aircraf weight.

The horizontal stabilizer usually provides a down load in level flight because:

a. b. c. d. 5.

upward. downward. zero. orward.

When considering the orces acting upon an aeroplane in straight-and-level flight at constant airspeed, which statement is correct?

a. b. c. d.

1 2

thrust, lif, weight. thrust, lif, drag, weight. thrust, lif, drag. lif, drag, weight.

the amount by which the lif exceeds the weight. the amount by which the thrust exceeds the drag. the amount by which the thrust exceeds the weight. the angle o attack o the wing.

Questions 8.

In a climb at a steady speed, the thrust is:

a. b. c. d. 9.

It increases take-off perormance. It increases engine perormance. It reduces climb perormance.

less than the weight. exactly equal to the weight. equal to the weight plus the drag. greater than the weight.

equal to the weight. greater than the weight. equal to the weight component perpendicular to the flight path. equal to the vertical component o weight.

During a glide the ollowing orces act on an aircraf:

a. b. c. d. 15.

s n o i t s e u Q

In a steady climb the wing lif is:

a. b. c. d. 14.

2 1

During a steady climb the lif orce is:

a. b. c. d. 13.

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 aircraf perormance?

a. b. c. 12.

wind speed. the aircraf weight. excess engine power. excess airspeed.

Assume that afer take-off a turn is made to a downwind heading. In regard to the ground, the aeroplane will climb at:

a. b. c. d. 11.

equal to the aerodynamic drag. greater than the aerodynamic drag. less than the aerodynamic drag. equal to the weight component along the flight path.

A constant rate o climb in an aeroplane is determined by:

a. b. c. d. 10.

12

lif, weight, thrust. lif, drag, weight. drag, thrust, weight. lif and weight only.

For a glider having a maximum L/D ratio o 20:1, the flattest glide angle that could be achieved in still air would be:

a. b. c. d.

1 f in 10 f. 1 f in 20 f. 1 f in 40 f. 1 f in 200 f.

399

12

Questions 16.

To cover the greatest distance when gliding the gliding speed must be:

a. b. c. d. 17.

I the weight o an aircraf is increased, the maximum gliding range:

a. b. c. d. 18.

19.

20.

32:1 16:1 8:1 4:1

During a turn the lif orce may be resolved into two orces; these are:

a. b. c. d.

400

increase and glide angle will be steeper. increase, but glide angle will remain the same. decrease. remain the same.

An aircraf has a L/D ratio o 16:1 at 50 kt in calm air. What would the approximate GLIDE RATIO be with a direct headwind o 25 kt?

a. b. c. d. 22.

lif, drag and weight. lif, thrust and weight. lif, drag, thrust and weight. lif and weight only.

I air brakes are extended during a glide, and speed maintained, the rate o 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.

During a ‘power-on’ descent the orces acting on an aircraf are:

a. b. c. d.

Q u e s t i o n s

decreases. increases. remains the same, and rate o descent is unchanged. remains the same, but rate o descent increases.

When gliding into a headwind, the ground distance covered will be:

a. b. c. d. 1 2

near to the stalling speed. as high as possible within VNE limits. about 30% aster than V MD. the one that gives the highest L/D ratio.

a orce opposite to thrust and a orce equal and opposite to weight. centripetal orce and a orce equal and opposite to drag. centripetal orce and a orce equal and opposite to weight. centriugal orce and a orce equal and opposite to thrust.

Questions 23.

In a turn at a constant IAS, compared to straight and level flight at the same IAS:

a. b. c. d. 24.

s n o i t s e u Q

insufficient rate o yaw. too much bank. too much nose-up pitch. insufficient bank.

increase. decrease but bank angle will increase. decrease but bank angle will decrease. remain the same.

An aircraf has a stalling speed in level flight o 70 kt IAS. In a 60° balanced turn the stalling speed would be:

a. b. c. d. 29.

2 1

For a turn at a constant IAS i the radius o turn is decreased, the load actor 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 turn is caused by:

a. b. c. d. 27.

only one radius o turn 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 lef.

As bank angle is increased in a turn at a constant IAS, the load actor will:

a. b. c. d. 26.

the same power is required because the IAS 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 orce. less power is required because the lif required is less.

In a turn at a given TAS and bank angle:

a. b. c. d. 25.

12

76 kt. 84 kt. 99 kt. 140 kt.

An increase in airspeed while maintaining a constant load actor during a level, coordinated turn would result in:

a. b. c. d.

an increase in centriugal orce. the same radius o turn. a decrease in the radius o turn. an increase in the radius o turn.

401

12

Questions 30.

How can the pilot increase the rate o turn and decrease the radius at the same time?

a. b. c. 31.

I an aircraf with a gross weight o 2000 kg were subjected to a total load o 6000 kg in flight, the load actor would be:

a. b. c. d. 32.

1 2

33.

the one with the centre o thrust arthest rom the centreline o the uselage. the one with the centre o thrust closest to the centreline o the uselage. the one designated by the manuacturer which develops most usable thrust. the ailure o which causes the least yawing moment.

Following ailure o the critical engine, what perormance should the pilot o a light, twin-engine aeroplane be able to maintain at V MCA?

a. b. c.

402

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 nose wheel steering.

What criteria determines which engine is the “critical” engine o a twin-engine aeroplane?

a. b. c. d. 36.

the heavier aircraf would have a higher “g” load. the lighter aircraf would have a higher “g” load. they would both have the same “g” load.

For a multi-engine aircraf, VMCG is defined as the minimum control speed on the ground with one engine inoperative. The aircraf must be able to:

a. b. c. d. 35.

Compensate or increase in induced drag. Increase the horizontal component o lif equal to the vertical component. Compensate or loss o vertical component o lif. To stop the nose rom dropping below the horizon and the airspeed increasing.

Two aircraf o different weight are in a steady turn at the same bank angle:

a. b. c. 34.

9g 2g 6g 3g

Why must the angle o attack be increased during a turn to maintain altitude?

a. b. c. d. Q u e s t i o n s

Shallow the bank and increase airspeed. Steepen the bank and increase airspeed. Steepen the bank and decrease airspeed.

Heading, altitude, and ability to climb 50 f/min. Heading only. Heading and altitude.

Questions

12

2 1

s n o i t s e u Q

403

12

Answers

Answers

1 2

A n s w e r s

404

1 b

2 b

3 a

4 b

5 b

6 d

7 b

8 b

9 c

10 d

11 c

12 a

13 c

14 b

15 b

16 d

17 d

18 a

19 c

20 a

21 b

22 c

23 b

24 a

25 b

26 d

27 a

28 c

29 d

30 c

31 d

32 c

33 c

34 c

35 b

36 b

Chapter

13 High Speed Flight

Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 407 Speed o Sound . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 407 Mach Number . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 408 Effect on Mach Number o Climbing at a Constant IAS . . . . . . . . . . . . . . . . . . . . . .

408

Variation o TAS with Altitude at a Constant Mach Number . . . . . . . . . . . . . . . . . . . 410 Influence o Temperature on Mach Number at a Constant Flight Level and IAS . . . . . . . . . 410 Subdivisions o Aerodynamic Flow . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 411 Propagation o Pressure Waves . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 412 Normal Shock Waves . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 414 Critical Mach Number. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 414 Pressure Distribution at Transonic Mach Numbers . . . . . . . . . . . . . . . . . . . . . . . . .

416

Properties o a Normal Shock Wave . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 418 Oblique Shock Waves . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 419 Effects o Shock Wave Formation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 420 Buffet . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 427 Factors Which Affect the Buffet Boundaries . . . . . . . . . . . . . . . . . . . . . . . . . . . . 428 The Buffet Margin . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 432 Use o the Buffet Onset Chart . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

432

Delaying or Reducing the Effects o Compressibility . . . . . . . . . . . . . . . . . . . . . . . . 434 Aerodynamic Heating . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 442 Mach Angle . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 443 Mach Cone . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 444 Area (Zone) o Influence . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 444 Bow Wave . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 444 Expansion Waves . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 445 Sonic Bang. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 447 Methods o Improving Control at Transonic Speeds . . . . . . . . . . . . . . . . . . . . . . . . 447 Continued Overlea

405

13

High Speed Flight Sweepback - Fact Sheet. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .449

1 3

H i g h S p e e d F l i g h t

406

Questions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

451

Answers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

456

High Speed Flight

13

Introduction During the preceding study o low speed aerodynamics it was assumed that air is incompressible, that is, there is no change in air density resulting rom changes o pressure. At any speed there are changes in air density due to ‘compressibility’, but i 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 aircraf moves through the air infinitesimally small pressure disturbances, or waves, are propagated outward rom the aircraf in all directions, but only the waves travelling ahead o the aircraf are significant or the study o high speed flight. These pressure waves ’signal’ the approach o the aircraf and make the air change direction (upwash) and divide to allow passage o the aircraf.

Speed of Sound For the study o high speed flight we are interested in the speed at which the infinitesimally small pressure disturbances (waves) travel through the atmosphere. Pressure waves ‘propagate’ rom their source, that is, each air molecule is rapidly vibrated in turn and passes on the disturbance to its neighbour. The speed o propagation o small pressure waves depends upon the temperature o the air ONLY . The lower the temperature, the lower the speed o propagation. Sound is pressure waves, and the speed o any pressure wave through the atmosphere, whether audible or not, has become known as ‘the speed o sound’.

3 1

t h g i l F d e e p S h g i H

The speed o sound at 15°C is 340 metres per second, or approximately 661 kt. It can be shown that: where

a = √γRT

(Eq 13.1)

a = speed o sound

R = the gas constant

γ = a constant (1.4 or air)

T = absolute temperature

Since γ and R are constants, the speed o sound is proportional only to the square root o the absolute temperature . For example, at 15°C (288 K): a = √ 1.4 × 287 × 288

(R = 287 J/kg K)

= 340 m/s

a ∝ √ T

The speed o sound changes with Temperature ONLY

407

13

High Speed Flight Mach Number As the speed o an aircraf increases, there is a decrease in the distance between the aircraf and the influence o the advancing pressure waves . The aircraf begins to catch up the pressure waves, so the air has less time to move rom the aircraf’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 aircraf. Ultimately lif and drag, manoeuvrability and the stability and control characteristics will all be changed. These effects are due to the compressibility o 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 o the aircraf in relation to the potential effects o ‘compressibility’. I the aircraf speed through the air (TAS) and the speed o sound in the air through which it is flying (the local speed o sound) is known, this will give an indication o the degree o compressibility. This relationship is known as the Mach number and Mach number is a measure o compressibility . (E.g. M 0.5 is hal the local speed o sound). Mach number (M) is the ratio o the true airspeed (V) to the local speed o sound (a)

1 3


M =

V a

(Eq 13.2)

Equation 13.2 is a good ormula to remember because it allows several important relationships to be easily understood.

Effect on Mach Number of Climbing at a Constant IAS • It is known that temperature decreases with increasing altitude, so the speed o sound will decrease as altitude is increased. • It is also known that i altitude is increased at a constant IAS, the TAS increases. • Thereore, the Mach number will increase i altitude is increased at a constant IAS . This is because (V) gets bigger and (a) gets smaller. From a practical point o view: climbing at a constant IAS makes the distance between the aircraf and the influence o the advancing pressure waves decrease, which begins to change the flow and pressure patterns around the aircraf.

The lower the temperature The lower the speed o sound

408

High Speed Flight

13

The International Standard Atmosphere assumes that temperature decreases rom 15°C at sea level to -56.5°C at 36 089 f (11 000 m), then remains constant. The speed o sound will thereore decrease with altitude up to the tropopause and then remain constant, Figure 13.1.

50

40

STRATOSPHERE

×

TROPOSPHERE 30

I SA CONDITIONS

20

10

SPEED OF SOUND 3 1


0 400

500

600

700

SPEED OF SOUND - kt

Figure 13.1 Variation o speed o sound with altitude

Chapter 14 will ully describe V MO and MMO, the high speed (generally speaking) operational

limit speeds. It has been stated that as an aircraf climbs at a constant IAS 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 IAS.

As the climb continues, an altitude will be reached at which the flight crew must stop flying at a constant IAS and fly at a constant Mach number, to avoid accidentally exceeding M MO. 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.

409

13

High Speed Flight Variation of TAS with Altitude at a Constant Mach Number I M =

TAS a

When descending at a constant Mach number IAS will be increasing

then TAS = M × a It can be seen rom the equation that i an aircraf is flown at a constant Mach number: • as altitude decreases the temperature will rise, local speed o sound will increase and TAS will increase. • as altitude increases the temperature will drop, local speed o sound will decrease and TAS 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 remains constant

1 3


Influence of Temperature on Mach Number at a Constant Flight Level and IAS An aircraf normally operates at Indicated Airspeeds and the Mach number can be expressed in terms o IAS: IAS

M =

constant

For IAS in knots:

M =

√

P P0

(Eq 13.3)

IAS 661

√

P P0

(Eq 13.4)

where: P = pressure altitude P0 = pressure at sea level This shows that at a constant pressure altitude (Flight Level), the Mach number is independent o temperature or a constant IAS. This is because the speed o sound and the TAS, or a given IAS, both change as

410

√ T

High Speed Flight

13

Subdivisions of Aerodynamic Flow

M 0 75

M0 4

M1 2

LOW HIGH SUBSONIC

ALL M L < 1 0

TRANSONIC

SOME ML < 1 0 OTHER M L > 1 0

SUPERSONIC

ALL M L > 1 0

COMPRESSIBLE FLOW

M CRIT

M1 0

about M 0 7 t o M 0 8 depending on individual aircraft and angle of attack

M FS

(not to scale)

(Aircraft Mach number)

3 1


Figure 13.2 Classification o airspeed

Figure 13.2 shows the flow speed ranges with their approximate Mach number values, where:

MFS = Free Stream Mach number : The Mach number o the flow sufficiently remote rom an aircraf to be unaffected by it. (In effect, the Mach number o the aircraf through the air). This is the Mach number shown on the aircraf Mach meter. ML = Local Mach number: When an aircraf 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 surace o the aircraf, is subdivided as ollows: Subsonic

Less than Mach 1.0 (
Sonic

Exactly Mach 1.0 (M 1.0)

Supersonic

Greater than Mach 1.0 (>M 1.0)

411

13

High Speed Flight Propagation of Pressure Waves KEY W EAK PRESSURE WAVE

=

POSITION OF OBJECT WHEN PRESSURE WAVE GENERATED

=

POSITION OF OBJECT W HEN PRESSURE WAVE REACHES

M0 2

RADIUS

r

r M

(a)

=

MACH NUMBER OF OBJECT

=

PRESSURE WAVE EXPANDING FROM SOURCE AT LOCAL SPEED OF SOUND

Figure 13.3 shows a series o sketches which M0 5

r 1 3

(b)


M 0 75

r

illustrate the basic idea o pressure wave ormation ahead o an object moving at various Mach numbers and o the airflow as it approached the object. Pressure waves are propagated continuously, but or clarity just one is considered. I we assume a constant local speed o sound, then as the object’s Mach number increases, the object gets closer to the ‘leading edge’ o the pressure wave and the air receives less and less warning o the approach o the object. The greater the Mach number o the object, the more acute the upwash angle and the ewer the number o air particles that can move out o the path o the object. Air will

(c)

begin to build up in ront o the object and the density o the air will increase . M1 0

r (d) A IRFLOW

PRESSURE WAV E

Figure 13.3

412

When the object’s speed has reached the local speed o sound (d), the pressure wave can no longer warn the air particles ahead o the object because the object is travelling orward at the same speed as the wave.

High Speed Flight

13

Thereore, the ree stream air particles are not aware o anything until the particles that are piled up right in ront o the object collide with them. As a result o 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 o the air just ahead o it are also increased. The region o compressed air extends some distance ahead o the object, the actual distance depending on the speed and size o the object and the temperature o the air. At one point the ree air stream particles are completely undisturbed, having received no advance warning o the approach o a ast moving object, and then are suddenly made to undergo drastic changes in velocity, pressure, temperature and density. Because o the sudden nature o these changes, the boundary line between the undisturbed air and the region o compressed air is called a ‘shock wave’, a stylized sketch o which is shown in Figure 13.4.

At supersonic speeds there is no upwash or downwash 3 1


SHOCK WAVE (STYLIZED)

SUBSONIC SUPERSONIC AIRFLOW APPROXIMATELY M 1 3

AIRFLOW

Figure 13.4 Stylized shock wave

413

13

High Speed Flight Normal Shock Waves (Normal meaning perpendicular to the upstream flow). In addition to the ormation o a shock wave described overlea, 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 o a stationary aircraf and moving air). Whenever supersonic airflow is slowed to subsonic speed without a change in direction, a ‘normal’ shock wave will orm as a boundary between the supersonic and subsonic region . This means that some ‘compressibility effects’

will occur beore the aircraf as a whole reaches Mach 1.0.

AIR BEING ACCELERATED TO SUPERSONIC SPEED

1 3

NORMAL SHOCK WAVE

LOCAL MACH NUMBER > 1


LOCAL MACH NUMBER < 1 (PRESSURE WAVES ABLE TO TRAVEL FORWARD)

Figure 13.5 Shock wave at subsonic ree stream Mach number

Critical Mach Number An aerooil generates lif by accelerating air over the top surace. At small angles o attack the highest local velocity on an aircraf will usually be located at the point o maximum thickness on the wing. For example, at a ree stream speed o M 0.84, maximum local velocity on the wing might be as high as M 1.05 in cruising level flight. At increased angles o attack the local velocity will be greater and urther orward. Also, i the thickness/chord ratio were greater, the local speed will be higher. As the ree stream speed increases, the maximum speed on the aerooil will reach the local speed o sound first. The ree stream Mach number at which the local velocity first reaches Mach 1.0 (sonic) is called the Critical Mach number (M CRIT ). Critical Mach number is the highest speed at which no parts o the aircraf are supersonic

414

Increased thickness/chord ratio and increased angle o attack cause greater accelerations over the top surace o the wing, so the critical Mach number will decrease with increasing thickness/chord ratio or angle o attack .

High Speed Flight

13

Accelerating beyond M CRIT

At speeds just above the critical Mach number there will be a small region o supersonic airflow on the upper surace, terminated by a shock wave, Figure 13.6 .

NORMAL SHOCK AREA OF SUPERSONIC FLOW

WAVE

SUBSONIC FLOW SUBSONIC FLOW

Figure 13.6 Mixed supersonic & subsonic airflow at transonic speeds 3 1

As the aircraf speed is urther increased, the region o supersonic flow on the upper surace extends, and the shock wave marking the end o the supersonic region moves rearwards. A similar sequence o events will occur on the lower surace although the shock wave will usually orm at a higher aircraf speed because the lower surace usually has less curvature so the air is not accelerated so much.


When the aircraf speed reaches Mach 1.0, the airflow is supersonic over the whole o both upper and lower suraces, and both the upper and lower shock waves will have reached the trailing edge. At a speed just above Mach 1.0 the other shock wave previously described and illustrated in Figure 13.4, the bow wave, orms ahead o the leading edge. The bow shock wave is initially separated (detached) rom the leading edge by the build-up o 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. Figure 13.8 on page 417 , shows the development o shock waves on an aerooil section at a

small constant angle o attack as the airspeed is increased rom subsonic to supersonic.

A shock wave orms at the rear o an area o supersonic flow

At MCRIT there is no shock wave because there is no supersonic flow

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13

High Speed Flight Pressure Distribution at Transonic Mach Numbers Reer to Figure 13.8. The solid blue line represents upper surace pressure and the dashed blue line the lower surace. Decreased pressure is indicated upwards. The difference between the ull line and the dashed line shows the effectiveness o lif production; i the dashed line is above the ull line, the lif is negative in that area. Lif is represented by the area between the lines, and the Centre o Pressure (CP) by the centre o the area. During acceleration to supersonic flight, the pressure distribution is irregular . M 0.75 This is the subsonic picture. Separation has started near the trailing edge and there is practically no net lif over the rear third o the aerooil section; the CP is well orward. Figure 13.7 shows that CL is quite good and is rising steadily; C D, on the other hand, is beginning to

rise.

M 0.81 A shock wave has appeared on the top surace; notice the sudden increase o pressure

(shown by the alling line) caused by decreasing flow speed at the shock wave. The CP has moved back a little, but the area is still large. Figure 13.7 shows that lif is good, but drag is now rising rapidly. M 0.89 The pressure distribution shows very clearly why there is a sudden drop in lif coefficient

beore the aerooil as a whole reaches the speed o sound; on the rear portion o the aerooil the lif is negative because the suction on the top surace has been spoilt by the shock wave, while there is still quite good suction and high-speed flow on the lower surace. On the ront portion there is nearly as much suction on the lower surace as on the upper. The CP has now moved well orward again. Figure 13.7 shows that drag is still increasing rapidly.

1 3


M 0.98 This shows the important results o the shock waves moving to the trailing edge and

no longer spoiling the suction or causing separation. The speed o the flow over the suraces is nearly all supersonic, the CP has moved af again and, owing to the goo d suction over nearly all the top surace, with rather less on the bottom, the lif coefficient has actually increased. The drag coefficient is just about at its maximum, as shown Figure 13.7 . M 1.4 The aerooil is through the transonic region. The bow wave has appeared. The lif

coefficient has allen again because the pressure on both suraces is nearly the same; and or the first time since the critical Mach number, the drag coefficient has allen considerably.

M 0 81

CL

M 0 75

M 0 98

CD

M 0 98

M 0 89

M1 4

M 0 81 M 0 89

0 5

M1 4

1 5

1 0 Mach number

M 0 75

1 0

0 5

Mach number

Figure 13.7 Changes in lif & drag in the transonic region

416

1 5

High Speed Flight

13

M 0 81

M 0 75

CP

CP

3 1

M 0 89


M 0 98

CP

CP

M1 4

CP

Figure 13.8 Pressure distribution in the transonic region

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13

High Speed Flight Properties of a Normal Shock Wave

NORMAL

SHOCK WAVE

SUPERSONIC

SUBSONIC

SUBSONIC

Figure 13.9 Normal shock wave ormation

C STRONG OBLIQUE SHOCK WAVE A NORMAL SHOCK WAVE

1 3


B

W EAK OBLIQUE SHOCK WAVE MACH LINE

Figure 13.10 Normal & oblique shock waves

When a shock wave is perpendicular (normal) to the upstream flow, streamlines pass through the shock wave with no change o direction. A supersonic airstream passing through a normal shock wave will also experience the ollowing changes: • The airstream is slowed to subsonic; the local Mach number behind the wave is approximately equal to the reciprocal o the Mach number ahead o the wave e.g. i the Mach number ahead o the wave is 1.25, the Mach number o the flow behind the wave will be approximately 0.80. (The greater the Mach number above M 1.0 ahead o the wave, the greater the reduction in velocity). • Static pressure increases. • Temperature increases. • Density increases. • The energy o the airstream [total pressure (dynamic plus static)] is greatly reduced. Minimum energy loss through a normal shock wave will occur when the Mach number o the airflow in ront o the shock wave is small but supersonic .

418

High Speed Flight

13

Oblique Shock Waves An oblique shock wave is a slightly different type o shock wave. Reerring to Figure 13.10, at ‘A’ the air is travelling at supersonic speed, completely unaware o the approaching object. The air at ‘B’ has piled up and is subsonic, trying to slip around the ront o the object and merge with the airflow. Through the shock wave supersonic air rom ‘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 o the shock wave, lying perpendicular or normal to the direction o 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 o 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 o airstream velocity normal to the shock wave will always be subsonic downstream, otherwise no shock wave).

3 1


The black dashed lines in Figure 13.10 outline the area o 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 rom this point the effects o the shock wave decrease until the air is able to pass the object without being affected. Thus the effects o the shock wave disappear, and the line cannot be properly called a shock wave at all; it is called a ‘Mach line’.

Shock Wave Summary • The change rom 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 ormed. A normal shock wave marks the change rom supersonic to subsonic flow. • I the shock wave 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 o the shock wave. However, the component o velocity normal to the shock wave will always be subsonic downstream o the shock.

419

13

High Speed Flight Effects of Shock Wave Formation The ormation and development o shock waves on the wing have effects on lif, drag, stability and control. Many o 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 ofen to separate. This increases the depth o the turbulent wake behind the wing.

Effect of Shock Waves on Lift At low subsonic speeds the lif coefficient C L is assumed to be constant at a given angle o attack. With increasing Mach number, however, it will vary as shown in Figure 13.11.

M 0 81

CL

M 0 75

M 0 98

M 0 89

1 3


0 4

M CRIT SHOCK STALL

1 2

1 4

MACH NUMBER

Figure 13.11 Variation o C L with Mach number at Constant α

At high subsonic speed C L increases. This is the result o the changing pattern o streamlines. At low speeds the streamlines begin to diverge well ahead o the aerooil, Figure 13.3 and Figure 13.12. 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 rom Chapter 7 that this phenomena causes the stall speed to increase at high altitudes.

420

High Speed Flight

13

HIGH SPEED

LOW SPEED

Figure 13.12 Streamlines at low & high subsonic speeds

At speeds above M CRIT a shock wave will have ormed on the upper surace. This may cause boundary layer separation af o the shock wave, causing loss o lif (above M 0.81, as shown in Figure 13.8 and Figure 13.11).

SHOCK WAV E 3 1


SEPA RAT ED A IRFLOW

Figure 13.13 Shock stall

This is known as the shock stall because it results rom a separated boundary layer just as the low speed stall does. Shock stall occurs when the lif coefficient, as a unction o Mach number, reaches its maximum value (or a given angle o attack). The severity o the loss o lif depends on the shape o the wing sections. Wings not designed or high speeds may have a severe loss o lif at speeds above MCRIT (Figure 13.11), but wings designed specifically or high speed flight, with sweepback, thinner sections and less camber will have much less variation o lif through the transonic region. Separated airflow caused by a shock stall can cause severe damage to the airrame, par ticularly the empennage. This will be ully discussed on page 427 . The lower end o the transonic region is where most modern high speed jet transport aircraf operate and a small shock wave will exist on the top surace o the wing in the cruise.

421

13

High Speed Flight Effect of Shock Waves on Lift Curve Slope and CLMAX

At a constant angle o attack, the increase o C L as speed increases rom about M 0.4 into the low end o the transonic region gives a steeper lif curve slope, i.e. the change o C L per degree angle o attack will increase. However, because o earlier separation resulting rom the ormation o the shock wave, CLMAX and the stalling angle will be reduced. Figure 13.14 and Figure 13.15 illustrate these changes.

HIGH SUBSONIC SPEEDS

CL

INCOMPRESSIBLE FLOW

ANGLE OF ATTACK

1 3


Figure 13.14 Effects o Mach number on lif curve

C LMAX

0 4

1 0

MFS

Figure 13.15 Effect o Mach number on C LMAX.

422

High Speed Flight

13

Effect of Shock Waves on Drag As speed increases above M CRIT shock waves begin to orm 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. The Mach number at which the aerodynamic drag begins to increase rapidly is called THE DRAG DIVERGENCE MACH NUMBER . The Drag Divergence Mach Number is usually close to, and always greater than, the Critical Mach Number, as shown in Figure 13.16 . Energy Drag Energy drag stems rom the irreversible nature o 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 aircraf. 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 o shock wave movement there is a considerable flow separation, as shown in Figure 13.8 and Figure 13.13. 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. 3 1


M 0 98

CD

M 0 89

M 0 75

M 0 81

M CRIT DRAG DIVERGENCE

1 2 MACH NUMBER

MACH NUMBER

Figure 13.16 Variation o C D with Mach number

The change in drag characteristics is shown by the C D curve or a basic section at a constant angle o attack in Figure 13.16 . The ‘hump’ in the curve rom M 0.89 to M 1.2 is caused by: • The drag directly associated with the trailing edge shock waves (energy loss). • Separation o the boundary layer. • The ormation o the bow shock wave above M 1.0.

423

13

High Speed Flight Effect of Shock Waves on the CL / CD Drag Polar Curve

Although the curve o C L / CD is unique at low speeds, at transonic speeds when compressibility becomes significant, the curve will change. Figure 13.16 shows the variation o C L / CD with Mach number. The point at which the tangent rom the origin touches the curve corresponds to the maximum CL / CD or maximum L / D. In the transonic region, the L/D ratio is reduced.

CL LOW MACH NUMBERS

HIGH MACH NUMBERS

L D MAX

1 3

CD


Figure 13.17 Effects o Mach number on C L / C D polar

Effect of Shock Waves on the Centre of Pressure The centre o pressure o an aerooil is determined by the pressure distribution around it. As the speed increases through the transonic region, the pressure distribution changes and the centre o pressure will move. It was shown in Figure 13.8 that above MCRIT the upper surace pressure continues to drop on the wing until the shock wave is reached. This means that a greater proportion o the ‘suction’ pressure will come rom the rear o the wing, and the centre o pressure is urther af. The rearward movement o the CP, however, is irregular as the pressure distribution on the lower surace also changes. The shock wave on the lower surace usually orms at a higher ree stream Mach number than the upper surace shock, but it reaches the trailing edge first. The overall effect on the CP is shown in Figure 13.18. As the aircraf accelerates to supersonic speed, the overall movement o the CP is af to the 50% chord position.

424

High Speed Flight

13

1 4

1 0

M CRIT

50%

100%

PERCENTAGE CHORD

Figure 13.18 CP movement in the transonic region

3 1


The wing root usually has a thicker section than the wing tip so will have a lower M CRIT and shock induced separation will occur at the root first. The CP will move towards the tip, and i the wing is swept, this CP movement will also be rearward. This effect will be discussed in detail later.

Figure 13.19 Low pressure area in ront o shock wave

425

13

High Speed Flight Effect of Shock Waves on CP Movement Rearward CP movement with increasing Mach number in the transonic region produces a nosedown pitching moment. This is known as ‘Mach Tuck’, ‘High Speed Tuck’ or ‘Tuck Under’. A urther actor contributing to the nose-down pitching moment is decreased downwash at the tail resulting rom reduced lif at the wing root. I the tailplane is situated in the downwash, its effective angle o attack is increased, giving an increase in the nose-down pitching moment. For a stable aircraf a push orce is required on the stick to produce an increase in speed, but as a result o Mach tuck, the push orce required may decrease with speed above M CRIT giving an unstable stick orce gradient, Figure 13.20.

PUSH UNSTABLE STICK FORCE GRADIENT

STICK FORCE 0

1 3

PULL

M CRIT


MACH NUMBER

Figure 13.20 Reduction in stick orce with increasing Mach number

The Effect of Shock Waves on Flying Controls A conventional trailing edge control surace works by changing the camber o the aerooil to increase or decrease its lif. Deflecting a control surace down will reduce M CRIT. I the control is moved down at high subsonic speed and a shock wave orms on the aerooil ahead o the control surace, shock induced separation could occur ahead o the control, reducing its effectiveness. At low speed, movement o a control surace modifies the pressure distribution over the whole aerooil. I there is a shock wave ahead o the control surace, movement o the control cannot affect any part o the aerooil ahead o the shock wave , and this will also reduce control effectiveness. Conventional trailing edge control suraces may suffer rom greatly reduced effectiveness in the transonic speed region and may not be adequate to control the changes o moment affecting the aircraf at these speeds. This can be overcome by incorporating some or all o the ollowing into the design: an all moving (slab) tailplane ( Figure 11.2), roll control spoilers, making the artificial eel unit in a powered flying control system sensitive to Mach number or by fitting vortex generators.

426

High Speed Flight

13

Control Buzz I 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 o hinge moment which can set up an oscillation o the control surace called control buzz.

Buffet In the same way that separated airflow prior to a low speed stall can cause airrame buffet, shock induced separation (shock stall) at high speed can also cause buffeting. Aerodynamic buffet is a valuable stall warning, but it can damage the aircraf structure. Because o the higher dynamic pressure when an aircraf is operating in the transonic speed region, any shock induced buffet will have a greater potential or severe airrame damage. High speed buffet must be completely avoided. The aircraf must thereore be operated in such a manner that a (saety) margin exists beore aerodynamic buffet will occur. I the variables which affect both high speed and low speed stall are considered it will be possible to identiy the conditions under which buffeting will occur and a chart can be drawn to show all the actors involved. This is called a ‘Buffet Onset’ chart (illustrated in Figure 13.26 ) which is used by flight crews to ensure their aircraf is operated at all times with a specified minimum buffet margin.

3 1


In Chapter 7 it was shown that stall speed is affected by several actors. In this study o low speed stall combined with high speed buffet, the actors to be considered are: • • • • • •

Load actor (bank angle). Mach number. Angle o attack. Pressure altitude. Weight. CG position.

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13

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 o sound to decrease. Mach number is proportional to TAS and inversely proportional to the local speed o sound (a): TAS M = a Thereore, i altitude is increased at a constant EAS, Mach number will increase. At low speed CLMAX is airly constant, but above M 0.4 C LMAX decreases as shown in Figure 13.21. Reer also to Figure 13.12 or the reason why C LMAX starts to decreases at speeds above M 0.4. C LMAX

1 3

1 0

0 4


MFS

Figure 13.21

From the 1g stall speed ormula: VS1g =

√

L ½ ρ CLMAX S

It can be seen that as C LMAX decreases with increasing altitude, the 1g stall speed will increase. A LT ALT 1 1g Stall Speed

E A S

Figure 13.22

428

High Speed Flight

13

Figure 13.22 shows the variation with altitude o stalling speed at constant load actor (n).

Such a curve is called the stall boundary or the given load actor, in which altitude is plotted against equivalent airspeed. At this load actor (1g), the aircraf cannot fly at speeds to the lef o this boundary. It is clear that over the lower range o altitude, stall speed does not vary with altitude. This is because at these low altitudes, V S is too low or 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 (Alt1 in Figure 13.22) is eventually reached when there is only one speed at which the aircraf can fly, since increasing or decreasing speed or banking the aircraf will result in a stall. In the case o a 1g manoeuvre, this altitude is called the ‘Aerodynamic Ceiling’. I the aircraf were allowed to ‘drif up’ to this altitude, the aircraf will stall. Not a pleasant prospect or a modern high speed jet transport aircraf. This state o difficulty is also called ‘coffin corner’. Reer also to Figure 13.25. Note: The recovery in C LMAX at supersonic speeds is such that it may still be possible to operate above this ceiling i enough thrust is available to accelerate the aircraf to supersonic speeds at this altitude. 3 1

FL


CONSTANT MACH NUMBER

E AS

Figure 13.23

Load Factor Because load actor increases the stall speed, curves like the one sketched in Figure 13.22 can be drawn or all values o load actor up to the maximum permissible ‘g’, and together they constitute the set o stalling boundaries or the given aircraf. Such a set o curves is shown in Figure 13.23. Superimposed on these curves are dashed lines representing lines o constant Mach number, showing how high Mach numbers can be achieved, even at relatively low EAS, at high altitudes.

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13

High Speed Flight Stall boundaries set a lower limit to the operating speed, according to the load actor. In the case o a high-speed aircraf, there is also an upper limit which is due to the approach o shock stall and the associated buffet which occurs i the aircraf enters the transonic speed range. The limits associated with these effects give the buffet boundaries.

FL

CONSTA NT MAC H BUFFET BOUNDARY

STALL BOUNDARY

E AS 1 3

Figure 13.24


Mach Number For a given aircraf there is a Mach number which, even at low angle o attack, cannot be exceeded because o the onset o shock stall. Figure 13.23 shows the EAS corresponding to this Mach number alling as altitude increases, so the range o op erating speeds is reduced at both ends.

Angle of Attack However, there is a urther effect which makes the buffet boundary a more severe limit than that suggested by a curve o constant Mach number. As the EAS associated with a given Mach number alls with increased altitude, so the required C L, and hence angle o attack, increases. This results in a reduction in the Mach number at which buffeting occurs, which results in a urther reduction in the permissible airspeed. This effect is made worse as the high angle o 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 o attack, as Figure 13.24 illustrates. Also, an increase in load actor (bank angle) requires an increase in lif at a given EAS, hence an increase in angle o attack and a urther reduction in limiting Mach number. Thus the greater the load actor (bank angle or gust), the more severe the limitation due to buffeting. There is a set o buffet boundaries or various load actors (bank angles), just as there is a set o stall boundaries. The restrictions on speed and ‘g’ can be summarized in the orm o a single diagram in which load actor is plotted against EAS, shown in Figure 13.25.

430

High Speed Flight

g

13

"COFFIN CORNER" MAXIMUM PERMISSIBLE

SEA LEVEL

g

STALLING BOUNDARY

ENVELOPE

AT ALTITUDE

BUFFET BOUNDARY

1 0

VS

E AS

Figure 13.25

Pressure Altitude At sea level there is a stall speed below which the aircraf cannot fly. As load actor increases, so does the stall speed (proportional to the square root o the load actor). The curve o ‘g’ against EAS modifies the low speed stall boundary. It will continue to rise until the ‘limit load actor’ is reached ( Chapt. 14). The ‘limit load actor’ must never be exceeded. At the high speed end, when g = 1, there is a limiting speed which must not be exceeded because o shock induced buffet. As the load actor increases, so does the CL at given speed, and the limiting Mach number alls, slowly at first and then more rapidly. This defines a buffet boundary, which eventually intersects the boundary o maximum permissible ‘g’ to constitute an overall envelope like the outer curve depicted in Figure 13.25. Thus the aircraf may operate at any combination o speed and load actor within this envelope, but not outside it.

3 1


At altitude the situation is similar. However, at altitude the equivalent stalling speed increases with ‘g’ rather more rapidly than at sea level, because o the Mach number effect on C LMAX. Also, the buffet boundary becomes much more severe. Above a certain altitude the buffet boundary may intersect the stall boundary at a value o ‘g’ lower than the structural limit, as shown in Figure 13.25. This ‘point’ is another representation o “coffin corner”.

Weight The weight o the aircraf 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 o attack at any given speed, so that the Mach number at which buffeting occurs is reduced, and the buffet boundary is moved to the lef. 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.

CG Position Forward movement o the CG increases stall speed so the buffet boundaries will be affected in a similar way to that due to weight increase.

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13

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 aircraf can fly. In the case o a 1g manoeuvre, this altitude is called the ‘Aerodynamic Ceiling’. Operating an aircraf at its aerodynamic ceiling would leave no saety margin. In 1g flight the aircraf would be constantly on the point o stall. It could not be manoeuvred nor experience the smallest gust without stalling. Regulations require an aircraf to be operated with a minimum buffet margin o 0.3g.

Use of the Buffet Onset Chart (Figure 13.26) 1.3g Altitude (1g + 0.3g = 1.3g): At this altitude a ‘g’ increment o 0.3 can be sustained

without buffet occurring. Using the data supplied: Follow the vertical solid red line upwards rom 1.3g to the 110 tonnes line, then horizontally to the 30% CG vertical line, then parallel to the CG reerence line, again horizontally to the M 0.8 vertical line. The altitude curve must now be ‘paralleled’ to read-off the Flight Level o 405. The 1.3g altitude is 40 500 f. I the aircraf is operated above FL405 at this mass and CG, a gust, or bank angle o less than 40°, could cause the aircraf to buffet. (40° o bank at high altitude is excessive, a normal operational maximum at high altitude would be 10° to 15°).

1 3


Buffet restricted speed limits : Using the data supplied:

Follow the vertical dashed red line upwards rom 1g to the 110 tonnes line, then horizontally to the 30% CG vertical line, then parallel to the CG reerence line. Observe the FL350 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 o the dashed red line, the FL350 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 M MO is the maximum operating Mach number which may not be exceeded under any circumstances. Aerodynamic ceiling : at 150 tonnes can be determined by:

Initially ollowing the red dashed line vertically upwards rom 1g, continue to the 150 tonnes plot, then move horizontally to the lef to M0.8 (via the CG correction). The interpolated altitude curve gives an aerodynamic ceiling o FL390. Load actor and bank angle at which buffet occurs : Using the data supplied:

From M 0.8, ollow the dashed blue line to obtain 54° bank angle or 1.7g.

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High Speed Flight

13

BUFFET ONSET

clean configuration REF

150

NOTE: FOR MACH NUMBERS ABOVE OR EQUAL TO .82 THERE IS NO C.G. VARIATION EFFECT FROM REFERENCE VALUE

FLIGHT LEVEL VMO LIMIT

140

130 60

120 80

100

110

120

100

140 90 160 80

180

3 1

200


220 W EIGHT ( t ) 240 250 270 290 310 330 350 370 390 410 .50

30 .55

.60

.65 .70 M AC H N U MB ER

.75

.80

.85 15 20 25 30 CG %

35 40

45 50 55

BANK ANGLE ( o) 60 65

1 1.2 1.4 1.6 1.8 2 2.2 2.4 LO AD F ACT OR

M MO

DATA : M = .80 FL = 350 WEIGHT = 110 tonnes CG = 30 %

RESULTS : BUFFET ONSET AT : O

M = 0.80 W ITH 54 BANK ANGLE, OR AT 1.7g LOW SPEED (1 g) : M = 0.555 HIGH SPEED : ABOVE M 0.84 ( M MO ) 1.3 g ALTITUDE = FL4 05

Figure 13.26 Example o a buffet onset chart

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13

High Speed Flight Delaying or Reducing the Effects of Compressibility To maximize revenue, airlines require their aircraf to fly as ast and as efficiently as possible. It has been shown that the ormation o shock waves on the wing results in many undesirable characteristics and a massive increase in drag. Up to speeds in the region o M CRIT the effects o compressibility are not too serious. It is thereore necessary to increase M CRIT as much as possible. Many methods have been adopted to delay or reduce the effects o compressibility to a higher Mach number, as detailed below.

Thin Wing Sections On a low t/c ratio wing, the flow acceleration is reduced, thus raising the value o M CRIT. For example i MCRIT or a 15% t/c wing is M 0.75, then M CRIT or a 5% t/c wing will be approximately M 0.85. The use o a low t/c ratio wing section has some disadvantages: • The lif produced by a thin wing will be less, giving higher take-off and landing speeds and increased distances. • A thin wing requires disproportionally wider main spars or the same strength and stiffness. This increases structural weight.

1 3

• Limited stowage space is available in a thin wing or:


• uel • high lif devices and their actuating mechanism and • the main undercarriage and its actuating mechanism.

Sweepback (see Page 449 for Sweepback Fact Sheet) One o the most commonly used methods o increasing M CRIT is to sweep the wing back. Forward sweep gives a similar effect but wing bending and twisting creates such a problem that sweepback is more practical or ordinary applications. A simplified method o visualizing the effect o sweepback is shown in Figure 13.27 . The swept wing shown has the ree stream velocity broken down to a component o velocity perpendicular to the leading edge and a component parallel to the leading edge. The component o velocity perpendicular to the leading edge is less than the ree stream velocity (by the cosine o the sweep angle) and it is this velocity component which determines the magnitude o the pressure distribution. M CRIT will increase since the velocity component affecting the pressure distribution is less than the ree stream velocity.

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High Speed Flight

13

V ELOCITY CO MPONENT PARALLEL TO LEA DING EDGE FREE STREAM V ELOCITY

SWEEP ANGLE V ELOCITY COMPON ENT PERPENDICULAR TO LEA DING EDGE

Figure 13.27 Effect o sweepback 3 1

Alternatively, it can be considered that compared to a straight wing, a swept-back wing o the same aerooil section has a smaller effective thickness chord ratio. Sweeping the wing back increases the effective aerodynamic chord or the same dimensional thickness, Figure 13.28.

SAME CHORD

FREE

The local velocity will be lower or a given ree stream velocity. In this way, the MCRIT o a swept wing will be higher than that o a straight wing.

STREAM FLOW

SAME CHORD

EFFECTIVE AERODYNAMIC CHORD INCREASED

Figure 13.28


Sweeping the wing back has nearly the same aerodynamic advantages as a thin wing, without suffering reduced strength and uel capacity. Unortunately, 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. Sweptback wings also increase the magnitude o high speed tuck.

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13

High Speed Flight Another advantage o sweepback is the reduced lif curve slope. This is illustrated by the lif curve comparison in Figure 13.29 or the straight and swept wing.

CL

STRAIGHT

SWEPT

ANGLE OF ATTACK

Figure 13.29 Effect o sweepback on sensitivity to gusts 1 3

Any reduction o lif curve slope makes the wing less sensitive to changes in angle o attack due to a gust or turbulence. Since the swept wing has the lower lif curve slope, a given vertical gust will increase the C L, and hence the load actor, by a smaller amount than would occur i the wing were straight.


Disadvantages of Sweep • Reduced CLMAX • gives a higher stall speed and increased take-off and landing distances. • Maximum lif angle o attack is increased, which complicates the problem o landing gear design (possibility o tail-strike) and reduced visibility rom the flight deck during takeoff and landing. The contribution to stability o a given tail surace area is also reduced. • A swept-back wing has an increased tendency to tip stall resulting in pitch-up at the stall and possible deep stall problems. • Reduced effectiveness o trailing edge control suraces and high lif devices because their hinge line is swept. To produce a reasonable CLMAX on a swept wing the hinge line o the inboard flaps may be made straight. Leading edge high lif devices are also used to improve the low speed characteristics.

436

High Speed Flight

13

Vortex Generators It has been shown that most o the unavourable 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 ormation increases the adverse pressure gradient so the loss o kinetic energy in the boundary layer will be greater. Increasing the kinetic energy o the boundary layer will reduce flow separation. Simple devices called vortex generators are used to re-energize the boundary layer. Vortex generators are small plates, vanes, blades or wedges mounted in spanwise rows along the wing surace, as illustrated in Figure 13.30.

VORTEX GENERATORS 3 1


Figure 13.30 Vortex generators (blade type)

Each vortex generator produces a vortex at its tip which will induce high energy air rom the ree 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. Vortex generators are usually located on the upper wing surace, particularly ahead o control suraces, 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 aircraf to prevent separation caused by high adverse pressure gradients due to the contours o the surace.

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13

High Speed Flight Area Rule In Chapter 6 it was stated that in addition to the drag o individual components there is an extra drag due to intererence between these components, principally between wing and uselage. This is especially important at high speed. Experiments have shown that a large part o the transonic drag rise or a complete aircraf is due to intererence. Intererence drag at transonic speeds may be minimized by ensuring that the cross-sectional area distribution along the aircraf’s longitudinal axis ollows a certain smooth pattern.

WING

TAIL FUSELAGE

NOSE

1 3

TAIL

Figure 13.31 Without area rule


With some early high speed aircraf designs this was not the case. The area increased rapidly in the region o the wing, again in the vicinity o the tail and decreased elsewhere, giving an area distribution like the one illustrated in Figure 13.31. On later aircraf, the uselage was waisted, i.e. the area was reduced in the region o 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 Figure 13.32. There is an optimum area distribution, and the minimization o transonic intererence drag requires that the aircraf should be designed to fit this distribution as closely as possible. This requirement is known as the ‘transonic area rule’. In practice, no aircraf has this optimum distribution, but any reasonably smooth area distribution helps to reduce the transonic drag rise.

WING

TAIL FUSELAGE

NOSE

Figure 13.32 Area rule

438

TAIL

High Speed Flight

13

Mach Trim It was stated on page 423 to page 426 that as speed increases beyond M CRIT, shock wave ormation at the root o a swept-back wing will generate a nose-down pitching moment because lif orward o the CG is reduced and downwash at the tailplane is reduced. At high Mach numbers an aircraf will tend to become speed unstable. Instead o an increasing

push orce being required as speed increases, a pull orce becomes necessary to prevent the aircraf accelerating urther. This is potentially very dangerous. A small increase in Mach number will give a nose-down pitching moment which will tend to urther increase the Mach number. This in turn leads to a urther increase in the nose-down pitching moment. This unavourable high speed characteristic, known as “ Mach Tuck”, ” High Speed Tuck” or “Tuck Under” would restrict the maximum operating speed o a modern high speed jet transport aircraf. Some improvement can be made by mounting the tailplane on top o the fin, wh ere it is clear o the downwash, but it has been shown that this can produce a deep stall problem. To maintain the required stick orce gradient at high Mach numbers, a Mach trim system must be fitted . This device, sensitive to Mach number, may:

• deflect the elevator up. • decrease the incidence o the variable incidence trimming tailplane. • move the CG rearwards by transerring uel rom the wings to a rear trim tank.

3 1


by an amount greater than that required merely to compensate or the trim change. This ensures the required stick orce gradient is maintained in the cruise at high Mach numbers. Whichever method o trim is used by a particular manuacturer, a Mach trim system will adjust longitudinal trim and operates only at high Mach numbers .

MACH TRIM

PUSH

INPUT RESULTANT STICK FORCE

STICK FORCE

0 BASIC STICK FORCE PULL M CRIT

MACH NUMBER

Figure 13.33 Effect o Mach trim

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13

High Speed Flight Supercritical Aerofoil A airly recent design development, used to increase efficiency when operating in the transonic speed region, is the ‘supercritical aerooil’.

FLAT UPPER SURFACE

LARGE THICKNESS

S - SHAPED CAMBER LINE THICK TRAILING

BLUNT NOSE

EDGE

12% CONVENTIONAL SECTION

17% SUPERCRITICAL SECTION

Figure 13.34 Supercritical aerooil shape

A supercritical aerooil shape, illustrated in Figure 13.34, differs rom a conventional section by having:

1 3


• • • • •

a blunt nose. large thickness. an S - shaped camber line. a relatively flat upper surace. a thick trailing edge.

Because the airflow does not achieve the same increase o speed over the flattened upper surace compared to a conventional section, the ormation o shock waves is delayed to a higher MFS and, the shock waves are much smaller and weaker when they do orm. Because the shock waves are smaller and weaker, there is not such a sharp pressure rise on the rear o the section and this gives a much more even ‘loading’ on the wing.

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High Speed Flight

13

The Advantages of a Supercritical Aerofoil • Because o the delayed ormation o shock waves and their weaker nature, less sweep angle is required or a given cruising Mach number, thus reducing some o the problems associated with sweepback. • The greater thickness gives increased stiffness and strength or a given structural weight. This also allows a higher aspect ratio to be used which reduces induced drag. • The increased section depth gives more storage space or uel. This type o wing section can be used to increase perormance in one o two ways: • Increased Payload By using existing cruise speeds, the uel consumption would be reduced, thus allowing an increase in payload with little or no drag increase over a conventional wing at the same speed. • 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

3 1

• The aerooil ront section has a negative camber to give optimum perormance at cruise Mach numbers, but this is less than ideal or low speed flight. C LMAX will be reduced, requiring extensive and complex high lif devices at the leading edge, which may include Krueger flaps, variable camber flaps, slats and slots.


• The trailing edge o the aerooil has large positive camber to produce the ‘af loading’ required, but which also gives large negative (nose-down) pitching moments. • This must be balanced by the tailplane, causing trim drag. • Shock induced buffet may cause severe oscillations.

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13

High Speed Flight Aerodynamic Heating Air is heated when it is compressed or when it is subjected to riction. An aircraf will have compression at the stagnation point, compression through a shock wave, and riction in the boundary layer.

500

400

300

200

100

0 1 3

40 100


0

1

2

3

4

MACH NUMBER

Figure 13.35 Surace temperature rise with Mach number

So when an aeroplane moves through the air its skin temperature will increase. This occurs at all speeds, but only becomes significant rom a skin temperature point o view at higher Mach numbers. It can be seen rom Figure 13.35 that the temperature rise at M 1.0 is approximately 40°C. Again rom a skin temperature point o view, this rise in temperature does not become significant until speeds in the region o M 2.0 are reached, which is the approximate limit speed or aircraf manuactured rom conventional aluminium alloys. Above this speed the heat treatment o the structure would be changed and the atigue lie shortened. For speeds above Mach 2.0, titanium or “stainless steel” must be used.

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High Speed Flight

13

Mach Angle Reerence to Figure 13.38 will show that as the Mach number increases, the shock waves become more acute. To illustrate why the angle o the shock waves changes, it is necessary to consider the meaning and significance o the Mach angle ‘µ’ (mu). I the TAS o the aircraf is greater than the local speed o sound, the source o pressure waves is moving aster than the disturbance it creates.

E

MACH LINE OR WAVE

a, C

LOCAL SPEED OF SOUND

B

D VELOCITY OF AIRCRAFT,

V

A

DIRECTION OF FLIGHT

3 1


Figure 13.36 Mach angle

Consider a point moving at velocity ‘V’ in the direction ‘A’ to ‘D’, as in Figure 13.36 . A pressure wave propagated when the point is at ‘A’ will travel spherically outwards at the local speed o sound; but the point is moving aster, and by the time it has reached ‘D’, the wave rom ‘A’ and other pressure waves sent out when the point was at ‘B’ and ‘C’ will have ormed circles as shown, and it will be possible to draw a common tangent ‘DE’ to these pressure waves. The tangent represents the limit which all the pressure waves have reached when the point has reached ‘D’. ‘AE’ represents the local speed o sound (a) and ‘AD’ represents the TAS (V) TAS As illustrated, M = 2.6 a The angle ‘ADE’, or µ, is called the Mach angle and by simple trigonometry: M =

a 1 = TAS M The greater the Mach number, the more acute the Mach angle µ. At M 1.0, µ is 90°. sin µ =

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13

High Speed Flight Mach Cone In three dimensions, the disturbances propagating rom a moving point source expand outward as spheres, not circles. I the speed o the source (V) is greater than the local speed o sound (a), these spheres are enclosed within a Mach cone, whose semi vertical angle is µ.

MACH CONE

a

V 1 3


Figure 13.37 Mach cone at approximately M 5.0

It can be seen rom Figure 13.37 that the Mach angle (µ) continues to decrease with increasing Mach number. The Mach angle is inversely proportional to the Mach number.

Area (Zone) of Influence When travelling at supersonic speeds the Mach cone represents the limit o travel o the pressure disturbances created by an aircraf: anything orward o the Mach cone cannot be influenced by the disturbances . The space inside the Mach cone is called the area or zone o influence. A finite body such as an aircraf will produce a similar pattern o waves but the ront will be an oblique shock wave and the wave angle will be greater than the Mach angle because the initial speed o propagation o the shock waves will be greater than the ree stream speed o sound.

Bow Wave Consider a supersonic stream approaching the leading edge o an aerooil. In order to flow around the leading edge, the air would suddenly have to turn through a right angle (see Figure 13.3). At supersonic speeds this is not possible in the distance available. The ree stream velocity will suddenly decelerate to below supersonic speed and a normal shock wave will orm ahead o the wing at the junction o 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 Figure 13.38.

444

High Speed Flight

13

BOW WAVE OBLIQUE SHOCK

M<1 M FS > 1 M FS > 1 NORMAL SHOCK

Figure 13.38 Bow wave

The shock wave ahead o the leading edge is called a bow wave and is normal only in the vicinity o the leading edge. Further away rom the leading edge (“above“ and ”below”) it becomes oblique. It can be seen in Figure 13.38 that the trailing edge shock waves are no longer normal because the ree stream mach number is greater than 1.0; they are also now oblique.

Expansion Waves 3 1

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 orms at the junction o the supersonic and subsonic flow, the generation o which is wasteul o 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 Figure 13.39. SUBSONIC FLOW

Figure 13.39 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”.

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13

High Speed Flight

EXPANSION WAVE

SUPERSONIC

VELOCITY

FLOW

UP PRESSURE, DENSITY AND TEMPERATURE

DOW N

Figure 13.40 Supersonic flow at a convex corner with expansion wave

Figure 13.40 shows that a supersonic airflow can ollow a convex corner because it expands

upon reaching the corner. 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.

1 3


OBLIQUE SHOCK

EXPANSION WAVES

Figure 13.41 Expansion waves in a supersonic flow

446

OBLIQUE SHOCK

High Speed Flight

13

Figure 13.41 shows a series o expansion waves in a supersonic airflow. Afer passing through

the bow shock wave, the compressed supersonic flow is ree to expand and ollow the surace 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 ollowing changes:• The airflow is accelerated; the velocity and Mach number behind the expansion wave are greater. • The flow direction is changed to ollow the surace. • The static pressure o the airflow behind the expansion wave is decreased. • The density o the airflow behind the expansion wave is decreased. • Since the flow change is gradual there is no “shock” and no loss o energy in the airflow. An expansion wave does not dissipate airflow energy.

Sonic Bang The intensity o shock waves reduces with distance rom the aircraf, but the pressure waves can be o sufficient magnitude to create a disturbance on the ground. Thus, “sonic bangs” are a consequence o supersonic flight. The pressure waves move with aircraf ground speed over the earth surace.

3 1


Methods of Improving Control at Transonic Speeds It has been seen that control effectiveness may decrease in the transonic region i conventional control suraces are used. Some improvement in control effectiveness may be obtained by placing vortex generators ahead o control suraces. However, alternative orms o control such as: • an all moving (slab) tailplane • roll control spoilers give better control in the transonic speed region. These types o control are explained in Flying Controls Chapter 11. Control surace buzz is sometimes remedied by fitting narrow strips along the trailing edge o the control surace, or it may be prevented by including dampers in the control system or by increasing the stiffness o the control circuit. Because o the high control loads involved at high speeds and the variation in loads through the transonic region, the controls will normally be ully power operated with artificial eel.

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13

High Speed Flight The table in Figure 13.42 is provided to summarize the characteristics o the three principal wave orms encountered with supersonic flow.

Supersonic Wave Characteristics

1 3


TYPE OF WAVE

OBLIQUE Shoc k wave

DEFINITION

A PLANE OF DISCONTINU ITY, INCLINED MORE THAN 90º FROM FLOW DIRECTION

NORMAL Shoc k wav e

A PLANE OF DISCONT INUITY, NORMAL TO FLOW DIRECTION

FLOW DIRECTION CHANGE

TURNED INTO A PRECEDING FLOW

NO CHANGE

EFFECT ON VELOC ITY and MACH NUMBER. BEHIND WAVE

DECREASE D BUT STILL SUPERSONIC

DECREASE D TO SUBSO NIC

INCREASED TO HIGHER SUPERSONIC

DECREAS E

EFFECT ON STATIC PRESSURE and DENSITY

INCREASE

GREAT INCREASE

EFFECT ON ENERGY OF AIRFLOW

DECREASE

GREAT DECREAS E

INCREASE

INCREASE

EFFECT ON TEMPERATURE

Figure 13.42 Characteristics o the three principle wave orms

448

EXPANSION wave

TURN ED AWAY FROM PRECEDING FLOW

NO CHANGE ( NO SHOC K )

DECREAS E

High Speed Flight

13

Sweepback - Fact Sheet Sweep Angle: The angle between the line o 25% chords and a perpendicular to the root

chord. Purpose o Sweepback: To increase MCRIT.

A Swept Wing Increases the Critical Mach Number (MCRIT).

All other effects rom a swept wing are by-products, most o them disadvantages. However, the benefits rom a higher MCRIT outweigh the associated disadvantages.

By-products of Sweepback 1.

Increased tendency to stall at the tip first - minimized by fitting wing ences, vortilons or saw tooth leading edges. • Tip stall can lead to pitch-up, a major disadvantage. • Pitch-up can give the tendency or a swept wing aircraf to Super Stall. • Aircraf that show a significant tendency to pitch-up at the stall MUST be fitted with a stall prevention device; a stick pusher. 3 1

Close to the stall, ailerons and coordinated use o rudder should be used to maintain wings level because the use o rudder alone would give excessive rolling moments. (V SR is adjusted so that adequate roll control exists rom the use o ailerons close to the stall). 2.


When compared to a straight wing o the same section, a swept wing is less aerodynamically efficient. • At a given angle o attack C L is less. • CLMAX is less and occurs at a higher angle o attack. • The lif curve has a smaller gradient (change in C L per degree change in alpha is less).

CL

HIGH ASPECT RATIO

LOW ASPECT RATIO (or sweepback)

Figure 13.43

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13

High Speed Flight • Swept wings must be fitted with complex high lif devices, both leading and trailing edge, to give a reasonable take-off and landing distance. ◦ The least efficient type o leading edge device is used on the inboard part o the

swept wing to help promote root stall. • Because o the higher stalling angle o attack, the fin or vertical stabilizer is swept to delay fin stall to a greater sideslip angle. • A swept wing must be flown at a higher angle o attack than a straight wing to give the required lif coefficient; this is most noticeable at low speeds. • One o the ew advantages o a swept wing is that it is less sensitive to changes in angle o attack due to gust or turbulence; a smaller change in Load Factor or 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 o M CRIT 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 aircraf’s longitudinal trim.

6.

The hinge line o trailing edge ‘flap’ type control suraces are not at right angles to the airflow, which reduces the efficiency o the controls.

1 3


450

Questions

13

Questions 1.

Identiy which o the ollowing is the correct ormula or Mach number:

a. b. c. d. 2.

c. d.

3 1

the ratio o the aircraf’s TAS to the speed o sound at sea level. the ratio o the aircraf’s TAS to the speed o sound at the same atmospheric conditions. the ratio o the aircraf’s IAS to the speed o sound at the same atmospheric conditions. the speed o sound.

s n o i t s e u Q

increase. decrease. remain constant. initially show an increase, then decrease.

The term ‘transonic speed’ or an aircraf means:

a. b. c. d. 6.

A severe nose-down pitching moment or “tuck under”. A high-speed stall and sudden pitch up. Severe porpoising. Pitch-up.

For an aircraf climbing at a constant IAS the Mach number will:

a. b. c. d. 5.

M = TAS × a

Mach number is:

a. b.

4.

constant

What is the result o a shock-induced separation o airflow occurring symmetrically near the wing root o a sweptwing aircraf?

a. b. c. d. 3.

TAS = M a M = IAS a M TAS = a

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

At M 0.8 a wing has supersonic flow between 20% chord and 60% chord. There will be a shock wave:

a. b. c. d.

at 20% chord only. at 20% chord and 60% chord. at 60% chord only. orward o 20% chord.

451

13

Questions 7.

As air flows through a shock wave:

a. b. c. d. 8.

For a wing section o given thickness, the critical Mach number:

a. b. c. d. 9.

1 3

11.

d.

the shock waves striking the tail. the high speed airflow striking the leading edge o the wing. wing flutter caused by the interaction o the bottom and top surace shock waves. the airflow being detached by the shock wave and the turbulent flow striking the tail.

The “area rule” applied to high speed aircraf requires:

a. b. c. d.

452

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).

High speed buffet is caused by:

a. b. c.

13.

its speed increases. its speed decreases. its speed remains the same. it changes direction to flow parallel with the Mach cone.

I an aeroplane accelerates above the critical Mach number, the first high Mach number characteristic it will usually experience is:

a. b. c. d. 12.

will start to increase. will start to decrease. will remain constant. is directly proportional to the Mach number.

As air flows through a shock wave:

a. b. c. d.

Q u e s t i o n s

will decrease i angle o attack is increased. will increase i angle o attack is increased. will not change with changes o angle o attack. is only influenced by changes in temperature.

At speeds above the critical Mach number, the lif 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.

that the cross-sectional area shall be as small as possible. that the variation o cross-sectional area along the length o the aircraf ollows a smooth pattern. that the maximum cross-sectional area o the uselage should occur at the wing root. that the uselage and the wing area be o a ratio o 3 : 1.

Questions 14.

An all moving tailplane is used in preerence to elevators on high speed aircraf:

a. b. c. d. 15.

b. c. d.

3 1

s n o i t s e u Q

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 rom the direction o the original airflow.

As airflow passes through a normal shock wave, which o the ollowing changes in static pressure (i), density (ii), and Mach number (iii) will occur?

a. b. c. d. 19.

Outward and orward. Inward and af. Outward and af. Inward and orward.

The airflow behind a normal shock wave will:

a. b. c. d. 18.

moves the centre o gravity to maintain stable lateral stick orces in the transonic region. automatically compensates or pitch changes while flying in the transonic speed region. prevents the aircraf rom exceeding its critical Mach number. switches out the trim control to prevent damage in the transonic region.

What is the movement o the centre o pressure when the wing tips o a sweptwing aeroplane are shock-stalled first?

a. b. c. d. 17.

because the effect o the elevator is reversed above the critical Mach number. because shock wave ormation on the elevator causes excessive stick orces. because shock wave ormation ahead o the elevator causes separation and loss o elevator effectiveness. because it would be physically impossible or a pilot to control the aircraf in pitch with a conventional tailplane and elevator configuration.

Mach Trim is a device which:

a.

16.

13

(i) decrease increase increase increase

(ii) increase decrease decrease increase

(iii) < 1.0 < 1.0 > 1.0 or < 1.0 < 1.0

An aerooil travelling at supersonic speed will:

a. b. c. d.

have its centre o pressure at 50 % chord. have its centre o pressure at 25% chord. give a larger proportion o lif rom the lower surace than rom the upper surace, and have its centre o pressure at 50 % chord. give approximately equal lif rom the upper and lower suraces, and have its aerodynamic centre at 50% chord.

453

13

Questions 20.

A bow wave is:

a. b. c. d. 21.

When an aircraf is flying at supersonic speed, where will the area o influence o any pressure disturbance due to the presence o the aircraf be located?

a. b. c. d. 22.

23.

Q u e s t i o n s

(ii) increase decrease decrease increase

(iii) increase decrease decrease decrease

(iv) increase decrease decrease increase

(v) decrease increase increase decrease

shock waves interering with the smooth airflow into the engine intakes. flying aster than MMO. the conversion o mechanical energy into thermal energy by the shock wave. flying aster than VMO.

What is the effect o a shock wave on control surace efficiency?

a. b. c. d.

454

(i) parallel to surace normal to wave parallel to wave parallel to chord

Wave drag is caused by:

a. b. c. d. 26.

a higher wing loading will increase MCRIT. low wing loading will give a higher MCRIT. wing loading does not influence M CRIT. wing loading and MCRIT are directly proportional.

What influence does an oblique shock wave have on the streamline pattern (i), variation o pressure (ii), temperature(iii), density (iv) and velocity (v)?

a. b. c. d. 25.

increases. decreases. is inversely proportional to the square root o the Mach number. remains the same.

The influence o weight (wing loading) on the ormation o shock waves is:

a. b. c. d. 24.

Within the Mach Cone. In ront o the Mach Cone. In ront o the bow wave. In ront o the Mach Cone only when the speed exceeds M 1.0

The temperature o the airflow as it passes through an expansion wave:

a. b. c. d.

1 3

a shock wave which orms on the nose o the aircraf at M CRIT. the shape ormed when the shock waves on the upper and lower wing surace meet at the trailing edge. a shock wave that orms immediately ahead o an aircraf which is travelling aster than the speed o sound. the shape o a shock wave when viewed vertically.

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 o efficiency, due to control deflection no longer modiying the total flow over the wing.

Questions 27.

13

At what speed does an oblique shock wave move over the earth surace?

a. b. c. d.

Aircraf ground speed. The TAS o the aircraf plus the wind speed. The TAS o the aircraf less the wind speed. The TAS relative to the speed o sound at sea level.

3 1

s n o i t s e u Q

455

13

Answers

Answers

1 3

A n s w e r s

456

1 a

2 a

3 b

4 a

5 c

6 c

7 b

8 a

9 b

10 b

11 d

12 d

13 b

14 c

15 b

16 d

17 a

18 d

19 a

20 c

21 a

22 b

23 b

24 a

25 c

26 d

27 a

Chapter

14 Limitations

Operating Limit Speeds. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 459 Loads and Saety Factors . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 459 Loads on the Structure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 459 Load Factor . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 460 The Manoeuvre Envelope (V - n Diagram) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 461 The CLMAX Boundary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 461 Design Manoeuvring Speed, VA . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 462 Effect o Altitude on V A . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 463 Effect o Aircraf Weight on V A . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . .

463

Design Cruising Speed V C . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 464 Design Dive Speed VD . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 464 Negative Load Factors . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 464 The Negative Stall. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 465 Manoeuvre Boundaries. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 465 Operational Speed Limits. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 466 Gust Loads. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 467 Effect o a Vertical Gust on the Load Factor . . . . . . . . . . . . . . . . . . . . . . . . . . . . 468 Effect o the Gust on Stalling. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 469 Operational Rough-air Speed (V RA / MRA) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 470 Landing Gear Speed Limitations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 472 Flap Speed Limit. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 473 Aeroelasticity (Aeroelastic Coupling) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 474 Flutter . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

477

Control Surace Flutter . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 478 Aileron Reversal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 480 Questions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

482

Answers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

486

457

14

1 4

L i m i t a t i o n s

458

Limitations

Limitations

14

Operating Limit Speeds In service an aircraf must observe certain speed limitations. These may be maximum speeds or minimum speeds, but in each case they are set to give sae operation in the prevailing conditions. The limits may be set by various considerations, the main ones being: • strength o the aircraf structure. • stiffness o the aircraf structure. • adequate control o the aircraf. Strength is the ability o the structure to withstand a load, and stiffness is the ability to withstand deormation.

Loads and Safety Factors • Limit load: • Ultimate load: • Factor o saety:

The maximum load to be expected in service. The ailing load o the structure. The ratio o ultimate load to limit load.

For aircraf structures the actor o saety is 1.5. The saety actor on aircraf structures is much lower than the saety actors used in other orms o engineering because o the extreme importance o minimum weight in aircraf structures. To keep the weight as low as possible, the saety actor must be kept to a minimum. Because o this it is extremely important not to exceed the limitations set on the operation o the aircraf, as the saety margin can easily be exceeded and structural damage may occur.

4 1

s n o i t a t i m i L

Loads on the Structure The airrame structure must obviously be strong enough to take the loads acting upon it in normal level flight, that is the orces due to lif, drag, thrust and weight. However, the aircraf is also required to manoeuvre and to fly in turbulent air. Under these conditions the loads on the aircraf will be increased, so it must also be strong enough to withstand whatever manoeuvres are specified or the aircraf and 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 o the aircraf.

459

14

Limitations Load Factor The loads which must be considered are given in the design requirements o an aircraf. They are given in terms o load actor (n), colloquially known as ‘g’. Load Factor (n) =

Lif Weight

In level flight, since lif equals weight, the load actor is 1.0 (1g). I the aircraf is perorming a manoeuvre such that, or example, the lif is twice the weight, the load actor is 2.0 (2g). The limit load is given in terms o load actor to make the requirement general to all aircraf. However, it should be appreciated that ailure o the structure will occur at some particular applied load. For example, i the structure ails at 10 000 lb load, an aircraf weighing 4000 lb will reach this load at a load actor o 2.5. However, i the aircraf weighs 5000 lb, the ailing load is reached at a load actor o 2.0, i.e. it takes less ‘g’ to overstress a heavy aircraf than a light one. Limit load actors are based on the maximum weight o the aircraf .

3 1 4

POSITIVE C LMAX A

C D

2


1

O 1

S

E VS F

H V A

NEGATIVE C LMAX

VC VD

Figure 14.1 The manoeuvre envelope

460

SPEED (EAS)

Limitations

14

The Manoeuvre Envelope (V - n Diagram) The maximum load actors which must be allowed or during manoeuvres are shown in an envelope o load actor against speed (EAS). Figure 14.1 shows a typical manoeuvre envelope or V - n diagram. The limit load actors will depend on the design category o the aircraf. The EASA regulations state that: a)

For normal category aircraf, the positive limit load actor 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 manuacturer will not design an aircraf to be any stronger than the minimum required by the regulations). The positive limit load actor or modern high speed jet transport aircraf is 2.5 .

b)

For utility category aircraf the positive limit load actor is 4.4

c)

For aerobatic category aircraf the positive limit load actor is 6.0

The negative limit load actor may not be less than: d)

-1.0 or normal category aircraf

e)

-1.76 or utility category aircraf

)

-3.0 or aerobatic category aircraf

4 1


The CLMAX Boundary The line OA in Figure 14.1 is determined by the C LMAX o the aircraf. In theory, the lif, and hence the load actor or a given weight, depends on the angle o attack o the wing and the airspeed. The maximum possible lif will occur at the angle o attack where C L is a maximum. At this angle o attack the lif will increase with speed as shown by the line OA. For level (1g) flight the speed at CLMAX will be the stalling speed (V S), represented by point S in Figure 14.1. At Point A, the load actor reaches its positive limit.

461

14

Limitations It can be seen rom Figure 14.2 that at speeds below point A the wing cannot produce a lif orce equal to the limit load actor, whereas at speeds above point A the limit load actor can be exceeded. Manoeuvres at speeds above point A thereore have the potential to cause permanent deormation to the structure or structural ailure i 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 perormed saely provided that the limit load actor is not exceeded.

PERMANE NT DEFORMATION OF STRUCTU RE POSSIBLE

ULTIMATE LOAD FACTOR

STRUCTURAL FAILURE

4

FACTOR OF SAFETY (1 5)

UNSAFE MANOEUVRE

3

A

POSITIVE LIMIT LOAD FACTOR

10

2 STALL ANGLE

1 4

1


5

O

SAFE MANOEUVRE

SPEED (EAS)

Figure 14.2 Loads imposed during manoeuvres

There is, o course, a saety actor on the airrame o 1.5 so complete ailure o the structure will not occur at the load actor o 2.5 but at 2.5 × 1.5 = 3.75. However, permanent deormation o the structure may occur at load actors between 2.5 and 3.75, so it is not sae to assume that the load actor may be increased above the limiting value just because there is a saety actor.

Design Manoeuvring Speed, VA The highest speed at which sudden, ull elevator deflection (nose-up) can be made without exceeding the design limit load actor.

When establishing V A the aeroplane is assumed to be flying in steady level flight, at point A1 in Figure 14.3, 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 o the C LMAX line and the positive limit load actor line (point A) to saeguard the tail structure because o the higher load on the tailplane during the pitch manoeuvre (Re. Page 274, Chapter 10, Manoeuvre Stability).

462

Limitations

CL

3

MAX

14

W ING FLA PS UP A

2

A 1

1

V 0

Vs

1

C

V

D

V A

-1

Figure 14.3 Design manoeuvring speed V A

Line OA in Figure 14.3 represents the variation o stalling speed with load actor. Stalling speed increases with the square root o the load actor, thereore; VA = VS1g √ n 4 1

For example an aircraf with a 1g stalling speed o 60 kt and limit load actor o 2.5 would have a VA o:


60 √ 2.5 = 95 kt

Effect of Altitude on VA At high altitude the equivalent stalling speed increases with ‘g’ rather more rapidly than at sea level because o the Mach number effect on C LMAX. Above a certain altitude the buffet boundary may intersect the stall boundary at a value o ‘g’ lower than the structural limit, thus VA will become more limiting at high altitude (Please reer to Figure 13.25 or a diagram).

Effect of Aircraft Weight on VA The 1g stalling speed depends on the weight o the aircraf. The line OA is drawn or the maximum design weight, so or lower weights the stalling speed will be less. For the same limit load actor VA will thereore decrease. For the example considered above, i VA is 95 kt at 2500 lb weight, then at 2000 lb weight it will be: 95

√

2000 = 85 kt 2500

Note: 20% decrease in weight has given approximately 10% decrease in V A.

463

14

Limitations

3

POSITIVE C LMAX A

C D

2 S

1

E

O

SPEED (EAS)

VS

1

F

H V A

NEGATIVE C LMAX

VC VD Figure 14.4 The manoeuvre envelope

1 4

Design Cruising Speed V C


Point ‘C’ in Figure 14.4 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 CS-25.335 and CS-23.335. It must give adequate spacing rom V B (see page 467 ) and VD to allow or speed upsets. For example CS-25 requires V C to be at least 43 kt above V B, and not greater than 0.8 V D. CS-23 has similar requirements. V C need not exceed the maximum speed in level flight at maximum continuous power (V H) or in CS-23, 0.9 V H at sea level

Design Dive Speed VD Point ‘D’ in Figure 14.4 is the design dive speed V D. This is the maximum speed which has to be considered when assessing the strength o the aircraf. It is based on the principle o an upset occurring when the aircraf is flying at V C, resulting in a shallow dive, during which the speed increases, until recovery is effected. I the resulting speed is not suitable because o buffet or other high speed effects, a demonstrated speed may be used. This is called V DF, the flight demonstrated design dive speed.

Negative Load Factors In normal flying and manoeuvres it is not likely that very large negative ‘g’ orces will be produced; however, some negative ‘g’ orces may occur during manoeuvres and the aircraf must be made strong enough to withstand them.

464

Limitations

14

The Negative Stall I the angle o attack o the wing is ‘increased’ in the negative direction, it will eventually reach an angle at which it will stall. (I the wing section is symmetrical this angle will be the same as the positive stall angle, but or a cambered wing, the angle and the negative C LMAX will usually be lower). The line OH in Figure 14.1 represents the negative C LMAX boundary. For large aircraf a limit load actor o -1 must be considered up to V C. From VC to VD the negative load actor varies linearly rom -1 to 0.

Manoeuvre Boundaries Taking into account the limiting values o positive and negative load actor, and the maximum speed to be considered, the aircraf is thereore sae to operate within the boundaries shown in Figure 14.5.

3

A

C D

2

1

S

SAFE

L

OPERATION E

O 1

H

SPEED (EAS)

4 1


F

Figure 14.5 Manoeuvre boundaries

Line SL represents level 1g flight. Line SA shows the load actors 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 airrame is required to withstand. Line OH shows the negative load actors that could be produced with the wing at its negative stalling angle, and line HFE is the negative ‘g’ limit. The design speeds V C and VD, already defined, are used or the purpose o assessing the strength requirements o the aircraf in various flight conditions. These speeds are not scheduled in the aircraf’s Flight Manual, but the operational speed limits which are scheduled, are related to them.

465

14

Limitations Operational Speed Limits The maximum airspeed at which an aircraf is permitted to fly is V MO or ‘large aircraf’ (CS-25) or VNE or other aircraf (CS-23) other than turbine engined aircraf. (For certification, a large aircraf is defined as one o more then 5700 kg Maximum Certificated Take-off Mass). Maximum Operating Speed (Large Aircraf) V MO / MMO : VMO is a speed that may not be

deliberately exceeded in any regime o flight (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 VMO is an Indicated Airspeed, as altitude increases the Mach number corresponding to VMO will increase. There will be additional limitations on the aircraf because o compressibility effects. In a climb VMO will be superceded by M MO (maximum operating Mach number) at about 24 000 to 29 000 f, depending on atmospheric conditions. Mach/Airspeed Warning System (Large Aircraf) : Two independent Mach/Airspeed warning

systems provide a distinct aural warning (clacker) any time the maximum speed o V MO /MMO is exceeded. The warning clackers can be silenced only by reducing air speed below V MO /MMO. When Climbing at Constant IAS It is Possible to Exceed M MO

1 4


When Descending at Constant Mach No. It is Possible to Exceed V MO Never Exceed Speed (Small Aircraf) V NE : VNE is set below V D to allow or speed upsets to be

recovered. (VNE = 0.9VD). VNE will be shown by a radial red line on the airspeed indicator at the high speed end o the yellow arc.

Maximum Structural Cruise Speed (Small Aircraf) V NO : VNO is the normal operating cruise

speed limit and must be not greater than the lesser o V C or 0.89VNE. On the airspeed indicator V NO is the upper limit o the green arc.

From VNO to VNE there will be a yellow arc, which is the caution range. You may fly at speed within the yellow arc only in smooth air, and then only with caution.

466

Limitations

14

Gust Loads The structural weight o an aircraf must be kept to a minimum while maintaining the required strength. The ollowing gust strengths were first ormulated in the late 1940s and their continued effectiveness has been verified by regular examination o actual flight data recorder traces.

+ 66 ft/sec

+ 50 ft/sec

GUST LOAD

+ 25 ft/sec

FACTOR

10 SPEED (EAS)

0

25 ft/sec 66 ft/sec VB

50 ft/sec VC

4 1


VD

Figure 14.6

Aircraf are designed to be strong enough to withstand a 66 f/sec vertical gust at V B (the design speed or maximum gust intensity). I an aircraf experienced a 66 f/sec vertical gust while flying at V B, it would stall beore exceeding the limit load actor. In turbulence an aircraf would receive maximum protection rom damage by flying at V B. VB is quite a low airspeed and it would take some time or an aircraf to slow rom V C (the design cruising speed) to V B i it flew into turbulence. Thereore, another design strength requirement is or the aircraf also to be strong enough to withstand a vertical gust o 50 f/ sec (EAS) at V C. Protection is also provided or the remote possibility o a vertical gust during a momentary upset to a speed o V D (the design diving speed). The aircraf must also be strong enough to withstand a vertical gust o 25 f/sec at V D. (VB, V C and VD are design speeds and are not quoted in an aircraf’s Flight Manual). In practice, a slightly higher speed than V B is used or turbulence penetration. This speed is V RA / MRA (the rough-air speed). VRA / MRA will give adequate protection rom over-stressing the aircraf plus give maximum protection rom an inadvertent stall .

467

14

Limitations Effect of a Vertical Gust on the Load Factor Vertical gusts will affect the load actor (n) by changing the angle o attack o the wing, Figure 14.7 . INCREASE IN LIFT

(C L ) AIRCRAFT TAS,

V

INCREASE IN ANGLE OF ATTACK VERTICAL GUST

EFFECTIVE AIRFLOW

VELOCITY

Figure 14.7 1 4

The ollowing example illustrates the effect o a vertical gust on the load actor (n).


An aircraf is flying straight and level at a C L o 0.42. A 1° change in angle o attack will change the C L by 0.1. I the aircraf is subject to a vertical gust which increases the angle o attack by 3°, what load actor will the aircraf experience? Load Factor =

LIFT WEIGHT

In straight and level flight: n = 1 or

0.42 0.42

A 3° increase in angle o attack will give: 3 × 0.1 = 0.3 the CL will increase by 0.3: 0.42 + 0.3 = 0.72 n =

0.72 = 1.7 0.42

A gust which increases the angle o attack by 3° will increase the load actor to 1.7

468

Limitations

14

For a given gust speed and aircraf TAS, the increment in the load actor depends on the increase in C L per change in angle o attack due to the gust (the slope o the lif curve). I the lif curve has a steep slope, the ‘g’ increment will be greater. Factors which affect the lif curve are aspect ratio and wing sweep.

CL

HIGH ASPECT RATIO

LOW ASPECT RATIO (or sweepback)

Figure 14.8

Wings having a low aspect ratio, or sweep, will have a lower lif 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 lif increment produced is a smaller proportion o the original lif orce or the more heavily loaded aircraf.

4 1


For a given TAS and gust speed, the increase o lif will be proportional to the wing area. Thereore, the increase in load actor is inversely proportional to the wi ng loading. Wing Loading =

Weight Wing Area

For a given aircraf the only variables or load actor increment in a gust are the aircraf TAS and the gust speed.

Effect of the Gust on Stalling I an aerooil encounters an upgust, it will experience an increase in angle o attack. For a given gust velocity the increment in angle increases as the aircraf TAS decreases. I the angle o 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 sae to fly i a gust is likely to be met so as not to stall in the gust.

469

14

Limitations Operational Rough-air Speed (VRA / MRA) For flight in turbulence an airspeed must be chosen to give protection against two possibilities: stalling and overstressing the aircraf structure. Turbulence is defined by a gust o a defined value. I this defined gust is encountered, the aircraf speed must be: • high enough to avoid stalling. • low enough to avoid damage to the structure. These requirements are ulfilled by calculating the stall speed in the gust and then building in sufficient strength or this speed. The key is the chosen value o the gust, as this will dictate the strength required and thereore the aircraf weight. The gust velocity is associated with the design speed, V B, and the vertical value o the gust is 66 f per second. Encountering a gust beore the pilot is able to slow the aircraf, plus the possibility o hitting a gust i the aircraf is ‘upset’ at high speed, must also be taken into consideration. Because these probabilities are lower however, progressively lower values o gust velocity are chosen at the higher speeds. These values are 50 f per second at the design cruise speed V C and 25 f per second at the design dive speed V D. The design gust values o 66, 50 and 25 f per second or gusts at the design speeds o V B, VC and VD have existed since the early 1940s. In the UK they were established as a result o the earliest “Flight Data Recorder” results. Modern flight recorder results and sophisticated design analyses continue to support the original boundaries o the design gust envelope.

1 4


Generally, design or strength is based on calculating the increase in load on the aircraf as a unction o an instantaneous increase in angle o attack on the wing page 469. On large aircraf, additional allowances have to be made or several reasons: • The greater dynamic response due to increased structural flexibility. • The possible implications o the smaller margin between actual cruise speed and design cruise speed. • The significance, in the more advanced designs, o the effects o build-up o gusts and unsteady flow generally. • The requency o storm penetrations. • The implications o the limited slow-down capabilities.

470

Limitations

14

All design speeds, and design gust values, are EAS. But, remember: the increase in angle o attack due to a gust is a unction o the TAS o the aircraf and the TAS o the gust. The choice o rough-air speed to be used operationally must be consistent with the strength o the aircraf. At the same time the aircraf must comply with both minimum stability and control criteria. There is also the important consideration o what maximum speed reduction can be achieved in a slow-down technique. A typical chart o the speeds to which the roughair speed is related, is shown below in Figure 14.9. The illustration is drawn or a single (mid) weight. Line AB is the 1g stall speed.

B

50

40

I

L

E

R

S P N H

30

K 20 4 1


10

0

A

100

C

O

200

M

G

300

J 400

SPEED - KNOTS EAS

Figure 14.9

Line CE is the stall speed in a 66 f per sec gust. (This assumes the 66 f per sec. gust up to maximum altitude. Note that point E would represent an extremely high true airspeed gust value). Line GHI is the V MO /MMO line. Line JKL is the V DF /MDF line. Line MN is an example o a maximum strength speed line or a 66 f per sec gust. Line RS is the 1.3g altitude.

471

14

Limitations At all speeds above the line CE the aeroplane will sustain a 66 ps gust without stalling and at all speeds below the line MN the aeroplane is strong enough to withstand a 66 ps gust. The rough-air speed thereore should lie somewhere between these two speeds, and the line OP gives equal protection between accidentally stalling and overstressing the aircraf. The line MN is a curious shape because different parts o the structure become critical at different altitudes. This line is actually the lowest speed boundary o a collection o curves at the higher speed end o the chart. Because o the obvious attraction o a single speed at all altitudes up to that at which the rough-air 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 halway speed would give equal pro tection against the 50-50 probability o being orced too ast or too slow. It has been stated that the diagram is drawn or a mid weight. The effect o weight change in terms o the lower and upper limits to rough-air speed is, o course, significant, but selcancelling. At low weights the stall line or a 66 f per sec gust alls to lower speeds and the maximum strength speed line increases to higher speeds. There is thereore no point in attempting a sophisticated variation o V RA with weight. The maximum altitude limit does, however, vary significantly with weight, and also varies or the level o manoeuvre capability chosen. A 0.3g increment to buffet is not too much protection in severe turbulence. A lower altitude will thereore be required or a higher level o protection, and, or a given level o protection, a lower altitude will be required or higher weights.

1 4


Landing Gear Speed Limitations The landing gear will normally be retracted as soon as possible afer take-off to reduce drag and increase the climb gradient. There is no normal requirement or the gear to be operated at high IAS so the retract and extend mechanism together with the attachment points to the structure are sized or the required task. To design the gear or operation at high IAS would unnecessarily increase structural weight. VLO: the landing gear operating speed is the speed at which it is sae both to extend and to

retract the landing gear. I the extension speed is not the same as the retraction speed, the two speeds must be designated as V LO (EXT) and VLO (RET).

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 IAS. Consequently V LO is usually lower than V LE. VLE: the landing gear extended speed. There may be occasions when it is necessary to erry the

aircraf with the gear down, and to do this a higher permissible speed would be convenient. VLE is the speed at which it is sae to fly the aircraf with the landing gear secured in the ully extended position. Because the undercarriage doors are closed, V LE is normally higher than V LO.

472

Limitations

14

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).

C LMAX W ING FLAPS DOWN

C LMAX W ING FLAPS UP

3

2

1 V 0

Vs

V

F

C

V

D

1

-1

4 1

Figure 14.10


Flaps increase C LMAX 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 rom the V-n diagram in Figure 14.10 that it is possible or a greater load to be applied to the structure at quite moderate airspeeds with flaps down. The limit load actor with flaps deployed is reduced rom 2.5 to 2 to give additional protection to the flaps and also the wing structure. I flaps are deployed in turbulence, a given vertical gust can generate a much larger lif orce which will subject the structure to a larger load, possibly exceeding the ability o the structure to withstand it, and the structure could ail. VFE : the Wing Flaps Extended Speed is the maximum airspeed at which the aircraf should be

flown with the flaps in a prescribed extended position. (Top o the white arc on the A SI).

Extending flaps or 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 aircraf Flight Manual.

473

14

Limitations Aeroelasticity (Aeroelastic Coupling) Aerodynamic orces acting on the aircraf produce distortion o the structure, and this distortion produces corresponding elastic orces 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 ailure occurs. This interaction between the aerodynamic loads and the elastic deormation o the air rame is known as aeroelasticity, or aeroelastic coupling. At low airspeeds, the aerodynamic orces are relatively small, and the resultant distortion o the structure produces only negligible effects. At higher speeds, aerodynamic loads and the consequent distortion are correspondingly greater. Aerodynamic orce is proportional to V , but structural torsional stiffness remains constant . This relationship implies that at some high speed, the aerodynamic orce build-up may overpower the resisting torsional stiffness and ‘divergence’ will occur. The aircraf must be designed so the speed at which divergence occurs is higher than the design speeds V D / MD. 2

Definitions: Elasticity 1 4

No structure is perectly rigid. The structure o an aircraf is designed to be as light as possible. This results in the aircraf being a airly flexible structure, the amount o flexibility depending on the design configuration o the aircraf. E.g. aspect ratio, sweepback, taper ratio etc.


Backlash

The possibility o movement o the control surace without any movement o the pilot’s controls. Mass distribution

The position o the CG o a surace in relation to its torsional axis. Mass balance

A mass located to change the position o the CG o a surace in relation to its torsional axis. Divergence

The structure will continue to distort until it breaks. Flutter

The rapid and uncontrolled oscillation o a surace resulting rom imbalance. Flutter normally leads to a catastrophic ailure o the structure.

474

Limitations

14

4

3

2

1 FLEXURAL AC

AXIS

Figure 14.11

Reer to Figure 14.11 which represents the view o a wing tip, and consider a vertical gust increasing the angle o attack o the wing. The additional lif orce will bend the wing tip upwards rom position 1 to 2 and the increase in lif acting through the AC, which is orward o the flexural axis, will twist the wing tip nose-up; this increases the angle o attack urther. The wing tip will rapidly progress to position 3 and 4. The wing is being wound up like a spring and can break i distorted too much.

4 1


How ar the structure is distorted depends on: • the flexibility o the structure. • the distance between the AC and the flexural axis. • the dynamic pressure (IAS). Methods o delaying divergence to a higher speed:

• The structure can be made stiffer, but this will increase weight. • A better solution is to move the flexural axis closer to the AC. This can easily be accomplished by mounting a mass orward o the AC. Instead o using a large piece o lead, as in control surace mass balance, the engines can be mounted orward o the leading edge and this will move the flexural axis closer to the AC. (Also see Flutter, page 477 ).

475

14

Limitations

W ING TIP

W ING ROOT

TRA ILING EDGE

LEADING EDGE

1 4


Figure 14.12 Typical flutter mode

476

Limitations

14

Flutter Flutter involves: • aerodynamic orces. • inertia orces. • the elastic properties o a surace. The distribution o mass and stiffness in a structure determine certain natural requencies and modes o vibration. I the structure is subject to a ‘orcing’ requency near these natural requencies, a resonant condition can result giving an unstable oscillation which can rapidly lead to destruction. An aircraf is subject to many aerodynamic excitations (gusts, control inputs, etc.) and the aerodynamic orces at various speeds have characteristic properties or rate o change o orce and moment. The aerodynamic orces may interact with the structure and may excite (or negatively damp) the natural modes o the structure and allow flutter. Flutter must not occur within the normal flight operating envelope and the natural modes must be damped i possible or designed to occur beyond V D / MD. A typical flutter mode is illustrated in Figure 14.12. Since the problem is one o high speed flight, it is generally desirable to have very high natural requencies and flutter speeds well above the normal operating speeds. Any change o stiffness or mass distribution will alter the modes and requencies and thus allow a change in the flutter speeds. I the aircraf 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 / MMO).

4 1


Wing flutter can be delayed to a higher speed, or a given structural stiffness (weight), by mounting the engines on pylons beneath the wing orward o the leading edge, Figure 14.13. The engines act as ‘mass balance’ or the wing by moving the flexural axis orward, closer to the AC.

AC

FLEXURAL AXIS MOV ED FORWA RD

Figure 14.13 Wing mass balanced by podded engines

477

14

Limitations Control Surface Flutter Control surace flutter can develop as a result o an oscillation o the control surace coupled with an oscillation in bending or twisting o the wing, tailplane or fin. A control surace oscillation can result rom backlash (ree play) in the control system or rom a disturbance (gust). Flutter can develop i the CG o the control surace is behind the hinge line, so that the inertia o the control surace causes a moment around the hinge. Torsional Aileron Flutter Figure 14.13 illustrates the sequence or a hal cycle, which is described below.

1.

The aileron is displaced downwards, exerting an upwards orce 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 surace lags behind due to the CG being af o the hinge line.

3.

The inherent stiffness o the wing has arrested the twisting motion (the spring is now wound up), but the air loads on the aileron, the stretch o the control circuit, and its upwards momentum, cause the aileron to ‘flick’ upwards, placing a down load on the trailing edge o the wing.

4.

The energy stored in the twisted wing and the reversed aerodynamic load o the aileron cause the wing to twist in the opposite direction. The cycle is then repeated.

1 4


Torsional aileron flutter can be prevented either by mass balancing the ailerons with attachment o a mass ahead o the hinge line to bring the CG onto, or slightly ahead o the hinge line, or by making the controls irreversible (ully powered controls with no manual reversion). Flexural Aileron Flutter This is generally similar, but is caused by the movement o the aileron lagging behind the rise and all o the outer portion o the wing as it flexes (wing tips up and down), thus tending to increase the oscillation. This type o flutter can also be prevented by mass balancing the ailerons. The positioning o the mass balance ‘weight’ is important the nearer the wing tip, the smaller the mass required. On many aircraf the mass is distributed along the whole length o the aileron in the orm o a leading edge ‘spar’, thus increasing the stiffness o the aileron and preventing a concentrated mass starting torsional vibrations in the aileron itsel.

Mass balancing must also be applied to elevators and rudders to prevent their inertia and the ‘springiness’ o the uselage starting similar motions. Mass balancing may even be applied to tabs. The danger o all orms o flutter is that the speed and amplitude o each cycle is greater than its predecessor, so that in a second or two the structure may be bent beyond its elastic limit and ail. Decreasing speed i flutter is detected is theoretically the only means o preventing structural ailure, but the rate o divergence is so rapid that slowing down is not really a practical solution.

478

Limitations

14

HINGE LINE TORSIONAL AXIS CG

1

2

4 1


3

4

Figure 14.14 Torsional aileron flutter

479

14

Limitations Aileron Reversal

15º

22º

Figure 14.15 Low speed aileron reversal

Low Speed It was described on page 147 that i an aileron is lowered when flying at high angles o attack, that wing could possibly stall, Figure 14.15. In that case the wing will drop instead o rising as intended. Hence the term low speed aileron reversal.

1 4


ELAST IC W ING

FLEXURAL AX IS

Figure 14.16

High Speed Aileron reversal can also occur at high speed when the wing twists as a result o the loads caused by operating the ailerons. In Figure 14.16 the aileron has been deflected downwards to increase lif and raise the wing. Aerodynamic orces act upwards on the aileron, and as this is behind the flexural axis o the wing, it will cause a nose-down twisting moment on the wing structure. This will reduce the angle o attack o the wing which will reduce its lif. I the twisting is sufficient, the loss o lif due to decreased angle o attack will exceed the gain o lif due to increased camber, and the wing will drop instead o lifing.

480

Limitations

14

SPOILER SURFACES OUTBOARD AILERONS (LOW SPEED ONLY)

INBOARD AILERONS (HIGH SPEED AND LOW SPEED)

Figure 14.17 Inboard & outboard ailerons & roll spoilers

High speed aileron reversal can be delayed to a speed higher than V D / M D by having inboard and outboard ailerons and/or roll control spoilers. The inboard ailerons, Figure 14.17 , 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. 4 1

On most high speed jet transport aircraf roll control spoilers assist the ailerons. Because they are mounted urther orward and on a stiffer part o the wing, roll control spoilers do not distort the wing structure to the same degree as ailerons.


481

14


I an aircraf is flown at its design manoeuvring speed V A:

a. b. c. d. 2.

The speed VNE is:

a. b. c. d. 3.

Q u e s t i o n s

4.

c. d.

c. d.

VLO is used when the aircraf is taking off and landing when the IAS 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 o the orward retracting nose gear. VLO is a lower IAS because the undercarriage doors are vulnerable to aerodynamic loads when the gear is in transit, up or down.

The phenomenon o flutter is described as:

a. b. c. d.

482

they are used only when preparing to land. the additional lif and drag created would overload the wing and flap structure at higher speeds. flaps will stall i they are deployed at too high an airspeed. too much drag is induced.

Why is VL E greater than V LO on the majority o large jet transport aircraf?

a. b.

6.

normal operations. abrupt manoeuvres. flight in smooth air. flight in rough air.

The maximum allowable airspeed with flaps extended (V FE) is lower than cruising speed because:

a. b.

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 aircraf 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.

1 4

it is possible to subject the aircraf to a load greater than its limit load during high ‘g’ manoeuvres. it is only possible to subject the aircraf to a load greater than its limit load during violent increases in incidence, i.e. when using excessive stick orce to pull-out o a dive. it is not possible to exceed the limit load. it is possible to subject the aircraf to a load greater than its limit load at high TAS.

rapid oscillatory motion involving only rotation o the control suraces, associated with the shock waves produced around the control suraces. oscillatory motion o part or parts o the aircraf relative to the remainder o the structure. rapid movement o the airrame caused by vibration rom the engines. reversal o the ailerons caused by wing torsional flexibility.

Questions 7.

What is the purpose o fitting the engines to an aircraf using wing mounted pylons?

a. b. c. d. 8.

b. c. d.

4 1

s n o i t s e u Q

not change. increase by 15%. increase by 7.5%. decrease by 7.5%.

VLO is defined as:

a. b. c. d. 12.

eliminate control flutter. aerodynamically assist the pilot in moving the controls. provide equal control orces on all three controls. return the control surace to neutral when the controls are released.

I an aircraf weight is reduced by 15%, V A will:

a. b. c. d. 11.

the down-going aileron increasing the semi-span angle o attack beyond the critical. flow separation ahead o the aileron leading edge. uneven shock wave ormation on the top and bottom surace o the aileron, with the attendant movement in control surace CP, causing the resultant orce to act in the opposite direction rom that intended. dynamic pressure acting on the aileron twisting the wing in the opposite direction, possibly causing the aircraf 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 or take-off and landing.

Aileron reversal at high dynamic pressures is caused by:

a.

9.

14

maximum landing gear operating speed. maximum landing gear extended speed. maximum leading edge flaps extended speed. maximum flap speed.

I flutter is experienced during flight, the preerable action would be:

a. b. c. d.

immediately increase speed beyond V MO / MMO, by sacrificing altitude i necessary. immediately close the throttles, deploy the speed brakes and bank the aircraf. rapidly pitch-up to slow the aircraf as quickly as possible. reduce speed immediately by closing the throttles, but avoid rapid changes in attitude and/or configuration.

483

14

Questions 13.

1 4

Q u e s t i o n s

484

Which o the ollowing statements are correct? 1.

It is a design requirement that control reversal speeds must be higher than any speed to be achieved in flight.

2.

The airrame must be made strong and stiff enough to ensure that the wing torsional divergence speed is higher, by a substantial saety margin, than any speed which will ever be achieved in any condition in flight.

3.

Flying control suraces are aerodynamically balanced to prevent flutter.

4.

An aircraf is not a rigid structure.

5.

Aeroelasticity effects are inversely proportional to IAS.

6.

Control reversal speed is higher i the aircraf is fitted with outboard ailerons which are locked-out as the aircraf accelerates; the inboard ailerons alone controlling the aircraf in roll at higher speeds.

a. b. c. d.

All the above statements are correct. 1, 2, 3 and 6. 1, 2, 4 and 6. 1, 3, 5 and 6.

Questions

14

4 1

s n o i t s e u Q

485

14

Answers

Answers 1 c 13 c

1 4

A n s w e r s

486

2 c

3 a

4 b

5 d

6 b

7 c

8 d

9 a

10 d

11 a

12 d

Chapter

15 Windshear

Introduction (Re: AIC 84/2008) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 489 Microburst. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 489 Windshear Encounter during Approach . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 491 Effects o Windshear . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 492 “Typical” Recovery rom Windshear . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 494 Windshear Reporting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 495 Visual Clues . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 495 Conclusions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

495

Questions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

496

Answers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

500

487

15

1 5

W i n d s h e a r

488

Windshear

Windshear

15

Introduction (Ref: AIC 84/2008) Windshear is a sudden drastic shif 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 aircraf to sudden updraughts, downdraughts or extreme horizontal wind components, causing sudden loss o lif or violent changes in vertical speeds or altitudes. Windshear will cause abrupt displacement rom 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 o experienced pilots flying modern and powerul aircraf. An encounter may cause alarm, a damaged undercarriage or a total catastrophe. The first and most vital deence is avoidance . The most powerul examples o windshear are associated with thunderstorms (cumulonimbus clouds), but windshear can also be experienced in association with other meteorological eatures such as the passage o a ront, or a marked low-level temperature inversion. The meteorological eatures o windshear will be dealt with ully elsewhere.

Microburst Microbursts are associated with thunderstorms and are one o the most dangerous sources o

windshear. Microbursts are small-scale intense downdraughts which, on reaching the surace, spread outward in all directions rom the downdraught centre. This causes the presence o both vertical and horizontal windshear that can be extremely hazardous to all types and sizes o aircraf, especially when within 1000 eet o the ground.

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A microburst downdraught is typically less than 1 mile in diameter as it descends rom the cloud base to about 1000 to 3000 eet above the ground. In the transition zone near the ground, the downdraught changes to a horizontal outflow that can extend to approximately 2.5 miles (4 km) in diameter. • Downdraughts can be as strong as 6000 eet per minute. • Horizontal winds near the surace can be as strong as 45 knots resulting in a 90 knot shear as the wind changes to or rom a headwind across the microburst.

• These strong horizontal winds occur within a ew hundred eet o the ground. An individual microburst seldom lasts longer than 15 minutes rom the time it strikes the ground until dissipation.

These are maximum values but they do indicate how it is possible or large and powerul aircraf to become uncontrollable when they meet such examples o the microburst.

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Windshear A microburst intensifies or about 5 minutes afer 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 or as long as an hour. Once microburst activity starts, multiple microbursts in the same general area are not uncommon and should be expected. STRONG DOW NDRA UGHT

Increasing Headwind

Increasing Tailwind

Outflow

Outflow 2

1 5

4

W i n d s h e a r

3

1

Figure 15.1 A microburst encounter during take-off

During take-off into a microburst, shown in Figure 15.1, an aircraf first experiences a headwind which increases perormance without a change in pitch and power (1). This is ollowed by a decreasing headwind and perormance, and a strong downdraf (2). Perormance continues to deteriorate as the wind shears to a tailwind in the downdraf (3). The most severe downdraf will be encountered between positions 2 and 3, which may result in an uncontrollable descent and impact with the ground (4).

490

Windshear

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Windshear Encounter during Approach The power setting and vertical velocity required to maintain the glide slope should be closely monitored. I any windshear is encountered, it may be difficult to stay on the glide path at normal power and descent rates. I there is ever any doubt that you can regain a reasonable rate o descent, and land without abnormal manoeuvres, you should apply ull 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 o windshear, it is best to assume ‘worse case’. It is impossible to predict at the first stages o a windshear encounter how severe it will be, and it is good advice to suggest that recovery action should anticipate the worst.

W INDSHEA R From

To

Headwind Calm or Tailwind

From

Tailwind

To

Calm or Headwind

INDICATIONS Indicated Airspeed

Decrease

Increase

Pitch Att itude

Decrease

Increase

Tends to Sink

Balloons

Increase

Decrease

Increase

Decrease

Up to Glideslope

Down to Glideslope

Reduce Power

Increase Power

Aircraft Ground speed

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ACTIONS Power Fly Be prepared to To Stay on Glide Path

Increase Rate of Descent (Due to faster ground speed)

Decrease Rate of Descent (Due to slower ground speed)

Figure 15.2 Indications & recovery actions or windshear encounter during approach

Reerring to Figure 15.2, this table gives guidance should you encounter windshear during a stabilized landing approach. Approaches should never be attempted into known windshear conditions.

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Windshear Effects of Windshear The relationship o an aeroplane in a moving air mass to its two reerence points must be ully understood. One reerence is the air mass itsel and the other is the ground. On passing through a shear line, the change o airspeed will be sudden, but the inertia o the aircraf will at first keep it at its original ground speed. The wind is a orm o energy and when it shears, an equivalent amount o energy is lost or gained. • A rapid increase in headwind (or loss o tailwind) are both ‘energy gains’, and will temporarily improve perormance, Figure 15.3. • Downdraughts or a sudden drop o headwind (or increase in tailwind) are the main danger at low altitude because they give an ‘energy loss’, Figure 15.4 and 15.5. 'ENERGY GAIN'

60 kt

Vertical Speed: Ground Speed: IAS :

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W i n d s h e a r

- Rapid increase in headwind.

10 kt

200 ft/min R.O.C. 130 kt 190 kt

Vertical Speed: Ground Speed : IAS:

GLIDE SLOPE

SHEAR LINE

Figure 15.3 “Energy gain” due to increase in headwind

492

700 ft/min R.O.D. 130 kt 140 kt

Windshear

'ENERGY LOSS'

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- Effect of downdraught.

10 kt

Vertic al Speed: Ground Speed: IAS:

1500 ft/ min R.O.D. 130 kt GLIDE 130 kt

Vertical Speed: Ground Speed: IAS:

SLOPE

700 ft/ min R.O.D. 130 kt 140 kt

SHEAR LINE

Figure 15.4 “Energy loss” due to downdraught 5 1

'ENERGY LOSS' - Loss of headwind.

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10 kt

Vertical Speed: Ground Speed: IAS:

1000 ft/min R.O.D. 130 kt 110 kt

Vertic al Speed: Ground Speed: IAS:

700 ft/ min R.O.D. 130 kt 140 kt

GLIDE SLOPE

20 kt SHEAR LINE

Figure 15.5 “Energy loss” due to loss o headwind

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Windshear “Typical” Recovery from Windshear The combination o increasing headwind, ollowed by downdraught, ollowed by increasing tailwind should be considered, as this is the sequence which might be encountered in a microburst on the approach, or ollowing take-off. • The presence o thunderstorms should be known and obvious, so the increase in speed caused by the rising headwind should be seen as the orerunner o a down-burst or microburst; any hope o a stabilized approach should be abandoned and a missed approach carried out as the only sae course o action .

• The initial rise in airspeed and rise above the approach path (balloon) should be seen as a bonus and capitalized on. Without hesitation, increase to go-around power, being prepared to go to maximum power i necessary , select a pitch angle consistent with a missed approach, typically about 15°, and hold it against turbulence and buffeting. • The next phase may well see the initial advantages o increased airspeed and rate o climb being rapidly eroded. The downdraught now strikes, airspeed may be lost and the aircraf may start to descend, despite the high power and pitch angle. It will be impossible to gauge the true angle o attack, so there is a possibility that the stick shaker (i fitted) may be triggered; only then should the attempt to hold the pitch angle normally be relaxed. • the point at which a downdraught begins to change to increasing tailwind may well be the most critical period. The rate o descent may lessen, but the airspeed may still continue to all; the height loss may have cut seriously into ground obstacle clearance margins. Given

1 5

that maximum thrust is already applied, as an extreme measure i the risk o striking the ground or an obstacle still exists, it may be necessary to increase the pitch angle ur ther and deliberately raise the nose until stick shaker is elt, then decrease back-pressure on the pitch control to try and hold this higher pitch angle , until the situation eases with the

W i n d s h e a r

aircraf beginning to escape rom the effects o the microburst. When there is an indefinite risk o shear, it may be possible to use a longer runway, or one that points away rom an area o 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 afer rotate already put the aircraf into a good configuration should a microburst then be encountered. The aircraf is, however, very low, there is little saety margin and the ride can be rough. I there is still extra power available, it should be used without hesitation. Ignore noise abatement procedures and maintain the high pitch angle, watching out or stick shaker indications as a signal to decrease backpressure on the pitch control. In both approach and take-off cases, vital actions are:

• Use the maximum power available as soon as possible . • Adopt a pitch angle o around 15° and try and hold that attitude. Do not chase airspeed. • 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.

494

Windshear

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Windshear Reporting I you encounter a windshear on an approach or departure, you are urged to promptly report it to the controller. An advanced warning o this inormation can assist other pilots in avoiding or coping with a windshear on approach or departure. The recommended method or windshear reporting is to state the loss or gain o airspeed and the altitudes at which it was encountered. I you are unable to report windshear in specific terms, you are encouraged to make reports in terms o the effect upon your aircraf.

Visual Clues You can see thunderstorms and hence receive a mental trigger to ‘think windshear’. Once alerted, look out or tell-tale signs such as: • Divergent wind sleeves or smoke. • Strong shafs o rain or hail, also ‘virga’ (intense precipitation which alls in shafs below a cumulonimbus cloud and evaporates in the dry air beneath). • Divergent wind patterns indicated by grass, crops or trees being beaten down or lashed. • Rising dust or sand. To observe and recognize any o the above will suggest that windshear danger is very close, i not imminent; nevertheless, a ew seconds o advance warning may make all the d ifference, i the warning is heeded and those seconds put to good use.

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Conclusions Most pilots will experience windshear in some orm or other; or most it may be no more than a very firm landing or a swing on take-off or landing requiring momentary use o, perhaps, ul l rudder or correction; they will probably put it down to ‘gusts’. Some ew pilots will experience more authentic examples o windshear which will stretch their skills to the limit. A very small number may find their skills inadequate. There is no sure way o knowing in advance the severity o 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 destabilize it to such an extent that a missed approach is the only sae course, and the sooner that decision is made, the saer it is likely to be . Other

encounters must be treated on their merits, but any hint o ‘energy loss’ should be met with a firm and positive response in line with the guidance put orward. Recognize -

that windshear is a hazard.

and Recognize Avoid Prepare Recover

-

the signs which may indicate its presence. windshear by delay or diversion. or the inadvertent encounter by a speed ‘margin’ i ‘energy loss’ windshear is suspected. know the techniques recommended or your aircraf and use them without hesitation i windshear is encountered.

495

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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 downdraught?

a. b. c. d. 2.

Maximum downdraughts in a microburst encounter may be as strong as:

a. b. c. d. 3.

Q u e s t i o n s

4.

c. d.

Decreasing headwind or tailwind. Increasing headwind and decreasing tailwind. Decreasing headwind and increasing tailwind. Increasing headwind or tailwind.

Which perormance characteristics should be recognized during take-off when encountering a tailwind shear that increases in intensity?

a. b. c. d.

496

Two minutes with maximum winds lasting approximately 1 minute. Seldom longer than 15 minutes rom the time the burst strikes the ground until dissipation. One microburst may continue or as long as 2 to 4 hours. For as long as 1 hour.

Which windshear condition results in a loss o airspeed?

a. b. c. d. 6.

80 kt. 40 kt. 90 kt. 45 kt.

What is the expected duration o an individual micro burst?

a. b.

5.

6000 f/min. 7000 f/min. 8000 f/min. 10 000 f/min.

An aircraf that encounters a headwind o 45 knots, within a microburst, may expect a total shear across the microburst o:

a. b. c. d.

1 5

Indicated Airspeed: constant. Vertical Speed: decreases. Pitch Attitude: decreases. Indicated Airspeed: increases. Vertical Speed: decreases. Pitch Attitude: constant. Indicated Airspeed: decreases. Vertical Speed: constant. Pitch Attitude: constant. Indicated Airspeed: decreases. Vertical Speed: decreases. Pitch Attitude: decreases.

Loss o, or diminished climb ability. Increased climb perormance immediately afer take-off. Decreased take-off distance. Improved ability to climb.

Questions 7.

Which condition would INITIALLY cause the indicated airspeed and pitch to increase and the sink rate to decrease?

a. b. c. d. 8.

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 o airspeed and resultant lif rom windshear?

a. b. c. d. 10.

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 o when a constant tailwind shears to a calm wind?

a. b. c. d. 9.

15

Maintain, or increase, pitch attitude and accept the lower-than-normal airspeed indications. Lower the pitch attitude and regain lost airspeed. Avoid overstressing the aircraf, pitch to stick shaker, and apply maximum power. Accelerate the aircraf to prevent a stall by sacrificing altitude.

Which o the ollowing would be acceptable techniques to minimize the effects o a windshear encounter? 1.

3. 4. 5. 6.

To prevent damage to the engines, avoid the use o maximum available thrust. Increase the pitch angle until the stick shaker activates, then decrease backpressure to maintain that angle o pitch. Maintain a constant airspeed. Use maximum power available as soon as possible. Keep to noise abatement procedures. Wait until the situation resolves itsel beore taking any action.

a. b. c. d.

1, 3, 5 and 6. 2, 3 and 5. 2, 3, 4, 5 and 6. 2 and 4.

2.

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s n o i t s e u Q

497

15

Questions 11.

Which o the ollowing statements about windshear is true? 1. 2. 3. 4. 5.

a. b. c. d. 12.

b. c. d.

Q u e s t i o n s

13.

b. c. d.

498

small-scale intense updraughts, which suck warm moist air into the cumulonimbus cloud. small-scale shafs o violent rain, which can cause severe problems to gas turbine engines. large-scale, violent air, associated with air descending rom the ‘anvil’ o a thunder cloud. small-scale (typically less than 1 mile in diameter) intense downdraughts which, on reaching the surace, spread outward in all directions rom the downdraught centre.

Thrust is being managed to maintain desired indicated airspeed and the glide slope is being flown. Which o the ollowing is the recommended procedure when you observe a 30 kt loss o airspeed and the descent rate increases rom 750 f/min to 2000 f/min?

a.

14.

1, 2, 3, 4 and 5. 1, 2 and 4. 1, 2, 4 and 5. 2, 3, 4 and 5.

A microburst is one o the most dangerous sources o windshear associated with thunderstorms. They are:

a.

1 5

Windshear can subject your aircraf to sudden updraughts, downdraughts, or extreme horizontal wind components. Windshear will cause abrupt displacement rom the flight path and require substantial control action to counteract it. Windshear only affects small single and twin engine aircraf. Large, modern, powerul, ast gas turbine engine powered aircraf will not suffer rom the worst effects o a microburst. Microbursts are associated with cumulonimbus clouds. Windshear can strike suddenly and with devastating effect which has been beyond the recovery powers o experienced pilots flying modern and powerul aircraf.

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 ull power and execute a go-around; report windshear to ATC as soon as practicable. Wait until the airspeed stabilizes and the rate o descent decreases, because microbursts are quite small and you will soon fly out o it.


A rapid increase in headwind is an ‘energy gain’. A rapid loss o tailwind is an ‘energy gain’. A shear rom a tailwind to calm is an ‘energy gain’. A shear rom calm to a headwind is an ‘energy gain’. A shear rom 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.

Questions 15.

16.


A downdraught is an ‘energy gain’. A rapid loss o tailwind is an ‘energy loss’. A shear rom a tailwind to calm is an ‘energy loss’. A shear rom calm to a headwind is an ‘energy gain’. A downdraught is an ‘energy loss’.

a. b. c. d.

1, 3 and 4. 1, 2, 3 and 5. 1, 4 and 5. 4 and 5 only.

Which o the ollowing sequences might be encountered when flying into a microburst?

a. b. c. d.

17.

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Increased headwind, ollowed by downdraught, ollowed by increased tailwind on the approach, or ollowing take-off. Increased headwind, ollowed by downdraught, ollowed by increased tailwind on the approach. Increased tailwind, ollowed by downdraught, ollowed by increased headwind ollowing take-off. Increased headwind, ollowed by downdraught, ollowed by increased tailwind on take-off. Increased tailwind, ollowed by downdraught, ollowed by increased headwind on the approach. Increased tailwind, ollowed by downdraught, ollowed by increased headwind on take-off. Increased headwind, ollowed by downdraught, ollowed by increased tailwind on the approach. 5 1

Which o the ollowing statements is correct when considering windshear? 1. 2. 3. 4. 5.

a. b. c. d.

s n o i t s e u Q

Recognize that windshear is a hazard to all sizes and types o aircraf. Recognize the signs which may indicate its presence. Avoid windshear by delaying departure or by diverting i airborne. Prepare or the inadvertent encounter by a speed ‘margin’ i ‘energy loss’ windshear is suspected. Know the techniques or recovery recommended or your aircraf and use them without any hesitation i windshear is encountered.

2, 4 and 5. 3, 4 and 5. 1, 2, and 5. 1, 2, 3, 4 and 5.

499

15

Answers

Answers

1 5

A n s w e r s

500

1 d

2 a

3 c

4 b

5 c

13 c

14 b

15 d

16 a

17 d

6 a

7 c

8 b

9 c

10 d

11 c

12 d

Chapter

16 Propellers

Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 503 Definitions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 503 Aerodynamic Forces on the Propeller . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 506 Thrust . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 506 Centriugal Twisting Moment (CTM) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 507 Propeller Efficiency . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 508 Variable Pitch Propellers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 509 Power Absorption. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 513 Moments and Forces Generated by a Propeller . . . . . . . . . . . . . . . . . . . . . . . . . . . 514 Effect o Atmospheric Conditions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 517 Questions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

518

Answers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

522

501

16

1 6

P r o p e l l e r s

502

Propellers

Propellers

16

Introduction A propeller converts shaf power rom the engine into thrust. It does this by accelerating a mass o air rearwards. Thrust rom the propeller is equal to the mass o air accelerated rearwards multiplied by the acceleration given to it. A mass is accelerated rearwards and the equal and opposite reaction drives the aircraf orwards.

Definitions The propeller blade is an aerooil and the definitions or chord, camber, thickness/chord ratio and aspect ratio are the same as those given previously or the wing. Additionally the ollowing must be considered.

Blade Angle or Pitch

BLADE ANGLE OR

PITCH

The angle between the blade chord and the plane o rotation. Blade angle decreases rom the root to the tip o the blade (twist) because rotational velocity o the blade increases rom root to tip. For reerence purposes, the blade angle is measured at a point 75% o the blade length rom the root.

PLANE OF ROTATION

Figure 16.1 Blade angle

6 1

Geometric Pitch

GEOMETRIC

PITCH

s r e l l e p o r P

The geometric pitch is the distance the propeller would travel orward in one complete revolution i 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 urther).

Figure 16.2 Geometric pitch

503

16

Propellers Blade Twist Sections near the tip o the propeller are at a greater distance rom the propeller shaf and travel through a greater distance. Tip speed is thereore greater. The blade angle must be decreased towards the tip to give a constant geometric pitch along the length o the blade. The blade angle determines the geometric pitch o the propeller. A small blade angle is called “fine pitch”, a large blade angle is called “coarse pitch”.

Effective Pitch

SLIP

GEOMETRIC

In flight the propeller does not move through the air at the geometric pitch; the distance it travels orward in each revolution depends on the aircraf’s orward speed. The distance which it actually moves orward in each revolution is called the “effective pitch” or “advance per revolution”.

PITCH

Propeller Slip The difference between the Geometric and the Effective Pitch is called the Slip.

EFFECTIVE PITCH

The Helix Angle HELIX ANGLE

1 6

The angle that the actual path o the propeller makes to the plane o rotation.

Figure 16.3 Effective pitch & slip

P r o p e l l e r s

Angle of Attack The path o the propeller through the air determines the direction o the relative airflow. The angle between the blade chord and the relative airflow is the angle o attack ( α), Figure 16.4. The angle o attack ( α) is the result o propeller rotational velocity (RPM) and aircraf orward velocity (TAS).

Fixed Pitch Propeller Figure 16.5 shows a “fixed pitch” propeller at constant RPM. Increasing TAS decreases the angle o attack o the propeller. Figure 16.6 shows a “fixed pitch” propeller at a constant TAS.

Increasing RPM increases the angle o attack o the propeller.

504

Propellers

16

RESULTANT PATH OF BLAD E ELEME NT ( RELATIVE AIRFLOW ) TAS OF AIRCRAFT + INDUCED FLOW

BLADE ANGLE OR PITCH HELIX ANGLE

PROPELL ER ( RPM )

PLANE OF ROTATION

Figure 16.4 Angle o attack

TAS INCREASED

CONSTANT PITCH

6 1

s r e l l e p o r P

CONSTANT ( RPM )

Figure 16.5 Angle o attack decreased by higher TAS

CONSTANT TAS

CONSTANT PITCH

INCREASED ( RPM )

Figure 16.6 Angle o attack increased by higher RPM

505

16

Propellers Aerodynamic Forces on the Propeller A propeller blade has an aerooil section, and when moving through the air at an angle o attack it will generate aerodynamic orces in the same way as a wing. The shape o the section will generate a pressure differential between the two suraces. The surace which has the greater pressure is called the “pressure ace” or “thrust ace”. When the propeller is giving orward thrust, the thrust ace is the rear, (flat) surace. The pressure differential will generate an aerodynamic orce, the total reaction, which may be resolved into two components, thrust and propeller torque.

Thrust A component at right angles to the plane o rotation. The thrust orce will vary along the length o each blade, reducing at the tip where the pressures equalize and towards the root where the rotational velocity is low. Thrust will cause a bending moment on each blade, tending to bend the tip orward. (Equal and opposite reaction to “throwing” air backwards).

Torque (Propeller) Torque is the equal and opposite reaction to the propeller being rotated, which generates a turning moment about the aircraf longitudinal axis. Propeller torque also gives a bending moment to the blades, but in the opposite direction to the plane o rotation. TOTAL REACTION

THRUST

1 6

AN GLE OF

P r o p e l l e r s

ATTACK

TORQUE

PLANE OF ROTATION

Figure 16.7 Thrust & torque

506

Propellers

16

Centrifugal Twisting Moment (CTM) Components ‘A’ and ‘B’, o the centriugal orce acting on the blade, produce a moment around the pitch change axis which tends to ‘fine’ the blade off.

A

B PITCH CHANGE AXIS

Figure 16.8 Centriugal turning moment (CTM)

Aerodynamic Twisting Moment (ATM) Because the blade CP is in ront o the pitch change axis, aerodynamic orce generates a moment around the pitch change axis acting in the direction o coarse pitch.

TOTAL REACTION 6 1

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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 Figure 16.15) and will reinorce it.

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Propellers Propeller Efficiency The efficiency o the propeller can be measured rom the ratio, Power out / Power in. The power extracted (out) rom a propeller, “Thrust Power”, is the product o Force (Thrust) × Velocity (TAS). The power into the propeller, “Shaf Power” is engine torque (Force) × Rotational Velocity (RPM). The efficiency o the propeller can be expressed as: Propeller Efficiency =

Thrust Power Shaf Power

Variation of Propeller Efficiency with Speed Figure 16.5 illustrated that or a fixed pitch propeller, increasing TAS at a constant RPM reduces

the blade angle o attack. This will decrease thrust. The effect o this on propeller efficiency is as ollows: • At some high orward speed the blade will be close to zero lif angle o attack and thrust, and thereore Thrust Power, will be zero. From the above ‘equation’ it can be seen that propeller efficiency will also be zero. • There will be only one speed at which a fixed pitch propeller is operating at its most efficient angle o attack and where the propeller efficiency will be maximum, Figure 16.10. • As TAS is decreased, thrust will increase because blade angle o attack is increased. Thrust is very large, but the TAS is low so propeller efficiency will be low. Thus no useul work is being done when the aircraf is, or instance, held against the brakes at ull power prior to take-off. The efficiency o a fixed pitch propeller varies with orward speed . 1 6

I blade angle can be varied as TAS and/or RPM is changed, the propeller will remain efficient over a much wider range o aircraf operating conditions, as illustrated in Figure 16.10.

P r o p e l l e r s

100 %

FINE PITCH

COARSE PITCH

AIRCRAFT FORWA RD SPEED

Figure 16.10 Efficiency improved by varying blade angle

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Propellers

16

Variable Pitch Propellers Adjustable pitch propellers These are propellers which can have their pitch adjusted on the ground by mechanically resetting 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 or take-off, climb and landing and coarse pitch or cruise. They will usually also have a eathered position. (Variable pitch) Constant speed propellers Modern aircraf 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 o TAS, giving improved take-off and climb perormance and cruising uel consumption.

Constant Speed Propeller OPEN

INCR

MIXTURE

THROTTLE

RPM

CLOSE

DECR

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Figure 16.11

Figure 16.11 illustrates a ‘typical’ set o engine and propeller controls or a small piston engine

aircraf with a variable pitch propeller. Throttle, prop’ and mixture are shown in the take-off (all orward) position. “Pulling back” on the prop’ control will decrease RPM. “Pushing orward” on the prop’ control will increase RPM. NB:

A reasonable analogy is to think o the prop’ control as an infinitely variable “gear change”. Forward (increase RPM) is first gear. Back (decrease RPM) is fifh gear.

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Propellers Figure 16.12 shows conditions during the

early stages o the take-off roll. The RPM is set to maximum and the TAS is low. The angle o attack is optimum and maximum available efficiency is obtained. As the aircraf continues to accelerate, the TAS will increase, which decreases the angle o attack o the blades. Less thrust will be generated and less propeller torque. This gives less resistance or 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 o attack constant.

FINE PITCH

( "small" blade angle )

AT THE START OF THE TAKE - OFF ROLL. LOW FORWARD SPEED, HIGH RPM

Figure 16.12 Low TAS, high RPM

Figure 16.13 shows the conditions at high

1 6

TAS

P r o p e l l e r s

COARSE PITCH

( "large" blade angle )

RPM

HIGH FORWARD SPEED, HIGH RPM

Figure 16.13 High TAS, high RPM

510

orward speed in level flight. As the TAS increased, the CSU continually increased the blade angle (coarsened the pitch) to maintain a constant blade angle o attack.

Propellers

16

Figure 16.14 shows conditions when the TAS

RPM

engine and prop’ have been set or cruise conditions. Optimum throttle and RPM settings are listed in the aircraf Flight Manual. The recommended procedure is to reduce the throttle first, then RPM. Whatever configuration into which the aircraf 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

Windmilling FINE PITCH

( "small" blade angle )

TAS

TORQUE

RPM

DRAG

TOTAL REACTION

I a loss o engine torque occurs (the throttle is closed or the engine ails), the prop’ will “fine off” in an attempt to maintain the set RPM. The relative airflow will impinge on the ront surace o the blade and generate drag and “negative propeller torque”. The propeller will now drive the engine, as shown in Figure 16.15.

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The drag generated by a windmilling propeller is very high.

STEADY GLIDE, THROTTLE CLOSED, NO SHAFT POWER, PROPELLER WINDMILLING.

Figure 16.15 Windmilling

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Propellers Feathering Following an engine ailure on a twin engine aeroplane the increased drag rom the windmilling propeller will seriously degrade climb perormance, limit range and add to the yawing moment caused by the ailed engine which will affect controllability. Also, by continuing to turn a badly damaged engine, eventual seizure o the engine or an engine fire might result.

ZERO LIFT ANGLE OF ATTACK

By turning the blades to their zero lif angle o attack, no propeller torque is generated and the propeller will stop, reducing drag to a minimum, as shown in Figure 16.16 . This will improve climb perormance because the ability to climb is dependent on excess thrust afer balancing aerodynamic drag. Windmilling drag is one o the “ingredients” o the yawing moment rom a ailed engine. Feathering the propeller o a ailed engine will also reduce the yawing moment and consequently, V MC.

1 6

Figure 16.16 Feathered

P r o p e l l e r s

COARSE PITCH

( "large" blade angle )

A single-engine aeroplane fitted with a constant speed propeller does not have a “eathering” capability, as such. However, ollowing engine ailure, drag can be reduced to a minimum by “pulling” the RPM (prop) control to the ully coarse position, as shown in Figure 16.17 . In a steady glide with no shaf power rom the engine (throttle closed), i the propeller pitch is increased by pulling back the prop’ lever, the aircraf Lif/Drag ratio will increase. This will decrease the rate o descent. The RPM would decrease because o the reduction in negative propeller torque.

STEADY GLIDE, THROTTLE CLOSED PROP' LEVER "PULLED BACK"

Figure 16.17

512

The opposite will be true i the propeller pitch is decreased.

Propellers

16

Power Absorption A propeller must be able to absorb all the shaf power developed by the engine and also operate with maximum efficiency throughout the required perormance envelope o the aircraf. The critical actor is tip velocity. I tip velocity is too high, the blade tips will approach the local speed o sound and compressibility effects will decrease thrust and increase rotational drag. Supersonic tip speed will considerably reduce the efficiency o a propeller and greatly increase the noise it generates . This imposes a limit on propeller diameter and RPM, and the TAS 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 o a multi-engine aircraf as close to the uselage as possible to minimize the thrust arm. Increasing the propeller diameter requires the engine to be mounted urther out on the wing to maintain adequate uselage clearance. To keep V MC within acceptable limits, the available rudder moment would have to be increased. Clearly, increasing the propeller diameter to increase power absorption i s not the preerred option.

Solidity To increase power absorption, several characteristics o the propeller can be adjusted. The usual method is to increase the ‘solidity’ o the propeller. Propeller solidity is the ratio o the total rontal area o the blades to the area o the propeller disc. It can be seen rom Figure 16.18 that an increase in solidity can be achieved by:

PROPELLER DISC

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• Increasing the chord o each blade . This increases the solidity, but blade aspect ratio is reduced, making the propeller less efficient. • Increasing the number o blades . Power absorption is increased without increasing tip speed or reducing the aspect ratio. Increasing the number o blades beyond a certain number (five or six) will reduce overall efficiency.

Figure 16.18 Solidity o a propeller

Thrust is generated by accelerating air rearwards. Making the disk too solid will reduce the mass o air that can be drawn through the propeller and accelerated. To increase the number o blades efficiently, two propellers rotating in opposite directions on the same shaf are used. These are called contra-rotating propellers.

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16

Propellers Moments and Forces Generated by a Propeller Due to its rotation a propeller generates yawing, rolling and pitching moments. These are due to several different causes: • • • •

Torque reaction. Gyroscopic precession. Spiral (asymmetric) slipstream effect. Asymmetric blade effect.

Note: The majority o modern engines are fitted with propellers which rotate clockwise when viewed rom the rear, so called “right-hand” propellers. The exceptions are small twin piston engine aircraf, which ofen have the propeller o the right engine rotating anti-clockwise to eliminate the disadvantage o having a “critical engine” (see Chapter 12), plus some older aircraf.

Torque Reaction Because the propeller rotates clockwise, the equal and opposite reaction (torque) will give the aircraf an anti-clockwise rolling moment about the longitudinal axis. During take-off this will apply a greater down load to the lef wheel, Figure 16.19, causing more rolling resistance on the lef wheel making the aircraf want to yaw to the lef. In flight, torque reaction will also make the aircraf want to roll to the lef. Torque reaction will be greatest during high power, low airspeed (IAS) flight conditions . Low IAS will reduce the power o the controls to counter the “turning” moment due to torque. 1 6

TORQUE

P r o p e l l e r s

PROPELLER ROTATION

Figure 16.19 Torque reaction giving lef turn during take-off

Torque reaction can be eliminated by fitting contra-rotating propellers . Torque rom the

two propellers, rotating in opposite directions on the same shaf, will cancel each other out. Co-rotating propellers on a small twin will not normally give a torque reaction until one engine ails. A lef “turning” tendency would then occur. Counter-rotating propellers on a small twin will reduce the torque reaction ollowing an engine ailure.

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Propellers

16

Gyroscopic Effect A rotating propeller has the properties o a gyroscope - rigidity in space and precession. The characteristic which produces “gyroscopic effect” is precession. Gyroscopic precession is the reaction that occurs when a orce is applied to the rim o a rotating disc. When a orce is applied to the rim o a propeller, the reaction occurs 90° ahead in the direction o rotation and in the same direction as the applied orce. As the aircraf is pitched up or down or yawed lef or right, a orce is applied to the rim o the spinning propeller disc. Note: Gyroscopic effect only occurs when the aircraf pitches and/or yaws. For example, i an aircraf with a clockwise rotating propeller is pitched nose-up, imagine that a orward orce has been applied to the bottom o the propeller disc. The orce will “emerge” at 90° in the direction o rotation, i.e. a right yawing moment. Gyroscopic effect can be easily determined when the point o application o the imagined orward orce on the propeller disc is considered. Pitch down - orward orce on the top, orce emerges 90° clockwise, lef yaw . Lef yaw - orward orce on the right, orce emerges 90° clockwise, pitch up . Right yaw - orward orce on the lef, orce emerges 90° clockwise, pitch down. Gyroscopic effect will be cancelled i the propellers are contra-rotating.

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Propellers Spiral Slipstream Effect As the propeller rotates it produces a backward flow o air, or slipstream, wh ich rotates around the aircraf, as illustrated in Figure 16.20. This spiral slipstream causes a change in airflow around the fin (vertical stabilizer). Due to the direction o propeller rotation ( clockwise) the spiral slipstream meets the fin at an angle rom the lef, producing a sideways orce on the fin to the right. Spiral slipstream effect gives the aircraf a yawing moment to the lef.

The amount o rotation given to the air will depend on the throttle and RPM setting. Spiral slipstream effect can be reduced by: • • • •

the use o contra or counter-rotating propellers. a small fixed tab on the rudder. the engine thrust line inclined slightly to the right. offsetting the fin slightly.

PROPELLER ROTATION

1 6

P r o p e l l e r s

SPIRAL SLIPSTREAM LEFT YAW

Figure 16.20 Spiral slipstream effect

516

Propellers

16

Asymmetric Blade Effect In general, the propeller shaf will be inclined upwards rom the direction o flight due to the angle o attack o the aircraf. This gives the down-going propeller blade a greater effective angle o attack than the up-going blade. The down-going (right) blade will generate more thrust. The difference in thrust on the two sides o the propeller disc will give a yawing moment to the lef with a clockwise rotating propeller in a nose-up attitude . Asymmetric blade effect will be greatest at ull power and low airspeed (high angle o attack).

Effect of Atmospheric Conditions Changes o atmospheric pressure or temperature will cause a change o air density. This will affect: • the power produced by the engine at a given throttle position. • the resistance to rotation o 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 .

Engine and Propeller Combined I the combined effect o 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: • i density increases, RPM will increase. • i density decreases, RPM will decrease. 6 1

Engine Alone

s r e l l e p o r P

I the shaf 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 o a fixed pitch propeller: • i density increases, power required will increase. • i density decreases, power required will decrease.

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As a result o gyroscopic precession, it can be said that:

a. b. c. d. 2.

A propeller rotating clockwise as seen rom the rear, creates a spiralling slipstream that tends to rotate the aeroplane to the:

a. b. c. d. 3.

b. c. d. 4.

5.

c. d.

twisting. effective pitch. geometric pitch. blade pitch.

Blade angle o a propeller is defined as the angle between the:

a. b. c. d.

518

the airstream in the wake o the propeller. the amount by which the distance covered in one revolution alls short o the geometric pitch. the increase in RPM which occurs during take-off. the change o blade angle rom root to tip.

The distance a propeller actually advances in one revolution is:

a. b. c. d. 7.

the distance it would move orward in one revolution i there were no slip. the angle the propeller shaf makes to the plane o rotation. the distance the propeller actually moves orward in one revolution. the angle the propeller chord makes to the relative airflow.

Propeller ‘slip’ is:

a. b.

6.

prevents the portion o the blade near the hub rom stalling during cruising flight. permits a relatively constant angle o attack along its length when in cruising flight. permits a relatively constant angle o incidence along its length when in cruising flight. minimizes the gyroscopic effect.

The geometric pitch o a propeller is:

a. b. c. d.

Q u e s t i o n s

right around the normal axis, and to the lef around the longitudinal axis. right around the normal axis, and to the right around the longitudinal axis. lef around the normal axis, and to the lef around the longitudinal axis. lef around the normal axis, and to the right around the longitudinal axis.

The reason or variations in geometric pitch (twisting) along a propeller blade is that it:

a.

1 6

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.

angle o attack and chord line. angle o attack and line o thrust. chord line and plane o rotation. thrust line and propeller torque.

Questions 8.

Propeller efficiency is the:

a. b. c. d. 9.

b. c. d.

As throttle setting is changed by the pilot, the prop governor causes pitch angle o the propeller blades to remain unchanged. The propeller control regulates the engine RPM and in turn the propeller RPM. A high blade angle, or increased pitch, reduces the propeller drag and allows more engine power or takeoffs. As the propeller control setting is changed by the pilot, the RPM o the engines remains constant as the pitch angle o the propeller changes.

6 1

s n o i t s e u Q

When does asymmetric blade effect cause the aeroplane to yaw to the lef?

a. b. c. d. 13.

difference between the geometric pitch o the propeller and its effective pitch. actual distance a propeller advances in one revolution. ratio o thrust horsepower to shaf horsepower. ratio between the RPM and number o blade elements.

Which statement best describes the operating principle o a constant speed propeller?

a.

12.

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 o thrust horsepower to shaf horsepower. ratio o geometric pitch to effective pitch. ratio o TAS to RPM.

A fixed-pitch propeller is designed or best efficiency only at a given combination o:

a. b. c. d. 10.

16

When at high angles o attack. When at high airspeeds. When at low angles o attack. In the cruise at low altitude.

The lef turning tendency o an aeroplane caused by asymmetric blade effect is the result o the:

a. b. c. d.

gyroscopic orces applied to the rotating propeller blades acting 90° in advance o the point the orce was applied. clockwise rotation o the engine and the propeller turning the aeroplane counter-clockwise. propeller blade descending on the right, producing more thrust than the ascending blade on the lef. the rotation o the slipstream striking the tail on the lef.

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Questions 14.

With regard to gyroscopic precession, when a orce is applied at a point on the rim o a spinning disc, the resultant orce acts in which direction and at what point?

a. b. c. d.

15.

The angle o attack o a fixed pitch propeller:

a. b. c. d. 16.

1 6

c. d.

520

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.

I engine RPM is to remain constant on an engine fitted with a variable pitch propeller, an increase in engine power requires:

a. b.

Q u e s t i o n s

depends on orward speed only. depends on orward speed and engine rotational speed. depends on engine rotational speed only. is constant or a fixed pitch propeller.

Counter-rotating propellers are:

a. b. c. d. 17.

In the same direction as the applied orce, 90° ahead in the plane o rotation. In the opposite direction o the applied orce, 90° ahead in the plane o rotation. In the opposite direction o the applied orce, at the point o the applied orce. In the same direction as the applied orce, 90° ahead o the plane o rotation when the propeller rotates clockwise, 90° retarded when the propeller rotates counter-clockwise.

a decrease in blade angle. a constant angle o attack to be maintained to stop the engine rom overspeeding. an increase in blade angle. the prop control lever to be advanced.

Questions

16

6 1

s n o i t s e u Q

521

16

Answers

Answers

1 6

A n s w e r s

522

1 b

2 d

3 b

4 a

5 b

13 c

14 a

15 b

16 c

17 c

6 b

7 c

8 b

9 a

10 c

11 b

12 a

Chapter

17 Revision Questions

Questions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

525

Answers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

552

Explanations to Specimen Questions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 553 Specimen Examination Paper . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .578 Answers to Specimen Exam Paper . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .586 Explanations to Specimen Exam Paper . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .587

523

17

1 7

Q u e s t i o n s

524

Questions

Questions

17

Questions 1.

A unit o measurement o pressure is:

a. b. c. d. 2.

Which o the ollowing are the correct SI units?

a. b. c. d. 3.

7 1

s n o i t s e u Q

A=F×M F=M×A M=F×A A=M/F

Which o the ollowing is the equation or power?

a. b. c. d. 7.

Newton. Psi. Joule. Watt.

Which o the ollowing expressions is correct?

a. b. c. d. 6.

m V squared. kg/square cm. kg - metres. kg/cubic metre.

What is the SI unit which results rom multiplying kg and m/s squared?

a. b. c. d. 5.

Density is kilograms per cubic metre, orce is newtons. Density is newtons per cubic metre, orce is kilograms. Density is kilograms per newton, orce is newton-metre squared. Density is kilograms per square metre, orce is kilograms.

What is the SI unit o density?

a. b. c. d. 4.

kg/square dm. kg/cubic metre. newtons. psi.

N/m. Nm/s. Pa/s squared. Kg/m/s squared.

At a constant CAS when flying below sea level an aircraf will have:

a. b. c. d.

a higher TAS than at sea level. a lower TAS than at sea level at ISA conditions. the same TAS as at sea level. the same TAS, but an increased IAS.

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Questions 8.

Static pressure acts:

a. b. c. d. 9.

TAS is:

a. b. c. d. 10.

12.

Q u e s t i o n s

dynamic pressure increases and static pressure increases. dynamic pressure increases and static pressure decreases. dynamic pressure is maximum at stagnation point. there is zero pressure at zero dynamic pressure.

Consider a uniorm flow o air at velocity V in a streamtube. I the temperature o the air in the tube is raised:

a. b. c. d.

526

(i) increases and (ii) decreases (i) increases and (ii) increases (i) decreases and (ii) decreases (i) decreases and (ii) increases

Bernoulli’s Theorem states:

a. b. c. d. 14.

(i) increases and (ii) decreases (i) increases and (ii) increases (i) decreases and (ii) decreases (i) decreases and (ii) increases

As subsonic air flows through a convergent duct: (i) static pressure (ii) velocity

a. b. c. d. 13.

increase with decreasing temperature. increase with increasing density. remain constant at all times. decrease with decreasing altitude.

As a smooth flow o subsonic air at a velocity less than M 0.4 flows through a divergent duct: (i) static pressure (ii) velocity

a. b. c. d. 1 7

higher than the speed o the undisturbed airstream around the aircraf. lower than the speed o the undisturbed airstream around the aircraf. lower than IAS at ISA altitudes below sea level. equal to IAS, multiplied by air density at sea level.

The difference between IAS and TAS will:

a. b. c. d. 11.

parallel to airflow. parallel to dynamic pressure. in all directions. downwards.

the mass flow remains constant and velocity V decreases. the mass flow will increase and velocity V remain constant. the mass flow will decrease and velocity V will remain constant. the mass flow remains constant and the velocity V will increase.

Questions 15.

17

In a subsonic flow venturi, the relationship between the total pressure, static pressure and dynamic pressure o undisturbed air and air in the throat will be: (i) Dynamic pressure will be constant, static pressure will decrease. (ii) Total pressure will be constant, dynamic pressure will increase.

a. b. c. d. 16.

In accordance with Bernoulli’s Theorem, where PT = Total Pressure, PS = Static pressure and q = Dynamic pressure:

a. b. c. d. 17.

remain the same. decrease. sonic. increase.

7 1

s n o i t s e u Q

What are the units or wing loading and dynamic pressure?

a. b. c. d. 20.

remain the same. decrease. increase. sonic.

The Principle o Continuity states that in a tube o increasing cross-sectional area, the speed o a subsonic and incompressible airflow will:

a. b. c. d. 19.

PT + PS = q PT = PS - q PT - PS = q PS + PT = q

The Principle o Continuity states that in a streamtube o decreasing cross-sectional area, the speed o a subsonic and incompressible airflow will:

a. b. c. d. 18.

both (i) and (ii) are correct. (i) is correct and (ii) is incorrect. (i) is incorrect and (ii) is correct. both (i) and (ii) are incorrect.

N/square metre and N/square metre. Nm and Nm. N and N/square metre. N/square metre and joules.

When considering the Principle o Continuity or incompressible subsonic flow, what happens in a streamtube with a change in cross-sectional area?

a. b. c. d.

The density at the throat will be the same as the density at the mouth. The density at the throat will be less than the density at the mouth. The density at the throat will be greater than the density at the mouth. Cannot say without knowing the change in cross-sectional area o the streamtube.

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17

Questions 21.

When considering the Principle o Continuity or subsonic flow, what happens in a streamtube or a change in cross-sectional area?

a. b. c. d. 22.

Which o the ollowing creates lif?

a. b. c. d. 23.

RHO 1 = RHO 2 RHO 1 > RHO 2 RHO 2 > RHO 1 Cannot say without knowing the change in cross-sectional area o the streamtube.

An accelerat accelerated ed air mass. A retarded air mass. A change in direction o mass flow. A symmetrical aerooil at zero angle o attack in a high speed flow.

Which o the ollowing statements about a venturi in a subsonic airflow is correct? (i) The dynamic pressure in the undisturbed flow and in the throat are equal. (ii) The total pressure in the undisturbed flow and in the throat are equal.

a. b. c. d. 24.

A line rom the centre o curvature o the leading edge to the trailing edge, equidistant rom the top and bottom wing surace is the:

a. b. c. d.

1 7

Q u e s t i o n s

25.

undisturbed airflow and chord line. undisturbed airflow and mean camber line. local airflow and chord line. local airflow and mean camber line.

How is the thickness o an aerooil section measured?

a. b. c. d.

528

A negative (nose-down) pitching moment. A positive (nose-up) pitching moment. Zero pitching moment. No aerodynamic orce.

Angle o attack is the angle between:

a. b. c. d. 27.. 27

camber line. upper camber line. mean chord. mean aerodynamic chord.

A symmetrical aerooil section at C L = 0 will produce?

a. b. c. d. 26.

(i) is correct and (ii) is incorrect. (i) is incorrect and (ii) is correct. (i) and (ii) are correct. (i) and (ii) are incorrect.

As the ratio o wing angle. Related to camber. As the percentage o chord. In metres.

Questions 28.

Lif and drag respectively are normal and parallel to:

a. b. c. d. 29.

b. c. d.

a symmetrical section at zero angle o attack will produce a small positive coefficient o lif. an asymmetrical section at zero angle o attack will produce zero coefficient o lif. a symmetrical section at zero angle o attack will produce zero coefficient o lif. aerooil section symmetry has no effect on lif coefficient.

7 1

s n o i t s e u Q

When considering the lif and drag orces on an aerooil section:

a. b. c. d. 34.

Angle o the chord line to the relative ree stream flow. Angle o the chord line to the uselage datum. Angle o the tailplane chord to the wing chord. Angle o the tailplane chord to the uselage datum.

When considering the coefficient o lif and angle o attack o aerooil sections:

a.

33.

the angle between the relative airflow and the horizont horizontal al axis. the angle between the wing chord line and the relative wind. the angle that determines the magnitude o the lif orce. the angle between the wing and tailplane incidence.

What is the angle o attack?

a. b. c. d. 32.

angle o incidence. glide path angle. angle o attack. climb path angle.

The term angle o attack is defined as:

a. b. c. d. 31.

the chord line. the longitudinal axis. the horizon. the relative airflow.

The angle between the aeroplane longitudinal axis and the chord line is:

a. b. c. d. 30.

17

they are only normal to each other at one angle o attack. they both depend on the pressure distribution on the aerooil section. they vary linearly. lif is proportional to drag.

Where does the lif act on the wing?

a. b. c. d.

Suction. Always orward o the CG. Centre o Gravity. Centre o Pressure.

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17

Questions 35.

Which o the ollowing creates lif?

a. b. c. d. 36.

Which o the ollowing is the greatest actor causing lif?

a. b. c. d. 37.. 37

b. c. d.

Q u e s t i o n s

39.

twice the wingspan above the ground. hal the wingspan above the ground. when the angle o attack is increased. upon elevat elevator or deflection.

The ormula or lif is:

a. b. c. d.

530

more thrust is required. less thrust is required. ground effect has no effect on thrust required. lif decreases.

On the approach to land, ground effect will begin to be elt at:

a. b. c. d. 41.. 41

0·25 0·5 2·0 4·0

On entering ground effect:

a. b. c. d. 40.

Lif acts perpendicular to the horizontal and drag parallel in a rearwards direction. Drag acts parallel to the chord and opposite to the direction o motion o the aircraf and lif acts perpendicular to the chord. Lif acts at right angles to the top surace o the wing and drag acts at right angles to lif. Drag acts in the same direction as the relative wind and lif perpendicular to the relative wind.

I IAS is doubled, by which o the ollowing actors should the original C L be multiplied to maintain level flight?

a. b. c. d.

1 7

Suction above the wing. Increased pressure below the wing. Increased airflow velocity below the wing. Decreased airflow velocity above the wing.

Which o the ollowing statements is correct?

a.

38.

A slightly cambered aerooil. An aerooil in a high speed flow. Air accelerat accelerated ed upwards. Air accelerat accelerated ed downwards.

L=W L = 2 rho V squared S CL L = 1/2 rho V squared S C L L = rho V S CL

Questions 42.

The influence o ground effect on landing distance will:

a. b. c. d. 43.

7 1

CL higher. CD higher. the same. CL much higher.

s n o i t s e u Q

What is the MAC o a wing?

a. b. c. d. 48.

Increase. Decrease. Remain the same. Induced drag will increase, but profile drag will decrease.

What is the CL and CD ratio at normal angles o attack:

a. b. c. d. 47.. 47

Static pressure and chord. Wingspan and dynamic pressure. Wing area and dynamic pressure. Wing area and static pressure.

What effect on induced drag does entering ground effect have?

a. b. c. d. 46.

The same. Greater in the higher aircraf. Greater Greaterr in the lower aircraf. Greate Less in the higher aircraf.

What do ‘S’ and ‘q’ represent in the lif equation?

a. b. c. d. 45.

increase landing distance. decrease landing distance. have no effect on landing distance. depend on flap position.

Two identical aircraf o the same weight fly at different altitudes. All other important actors remaining constant, assuming no compressibility and ISA conditions, what is the TAS o each aircraf?

a. b. c. d. 44.

17

Area o wing divided by the span. The same as the mean chord o a rectangular wing o the same span. The mean chord o the whole aeroplane. The 25% chord o a swept wing.

When an aircraf enters ground effect:

a. b. c. d.

the lif vector is inclined rearwards which increases the thrust required. the lif vector is inclined orwards which reduces the thrust required. the lif vector is unaffected, the cushion o air increases. the lif vector is inclined orward which increases the thrust required.

531

17

Questions 49.

When an aircraf enters ground effect:

a. b. c. d. 50.

When considering an angle o attack versus coefficient o lif graph or a cambered aerooil, where does the lif curve intersect the vertical C L axis?

a. b. c. d. 51.. 51

d. 53.

Q u e s t i o n s

b. c. d.

From the root to the tip on the top surace and rom the tip to the root on the bottom surace over the wing tip. From the root to the tip on the top surace and rom the tip to the root on the bottom surace over the trailing edge. From the tip to the root on the top surace and rom the root to the tip on the bottom surace over the trailing edge. From the tip to the root on the top surace and rom the root to the tip on the bottom surace over the wing tip.

Wing tip vortices are caused by unequal pressure distribution on the wing which results in airflow rom:

a. b. c. d.

532

Span divided by tip chord. Chord divided by span. Span divided by mean chord. Chord divided by span, measured at the 25% chord position.

Which o the ollowing most accurately describes the airflow which causes wing tip vortices?

a.

55.

Air spilling rom the top surace to the bottom surace at the wing tip. Air spilling rom the bottom surace to the top surace at the wing tip. Air spilling rom the bottom surace to the top surace at the lef wing tip and rom the top surace to the bottom surace at the right wing tip. Spanwise flow vector rom the tip to the root on the bottom surace o the wing.

Which o the ollowing is the correct definition o aspect ratio?

a. b. c. d. 54.

59%. 77%. 130%. 169%.

Which o the ollowing is the cause o wing tip vortices?

a. b. c.

1 7

above the origin. below the origin. at the point o origin. to the lef o the origin.

When in level flight at 1·3VS, what is the C L as a percentage o C LMAX?

a. b. c. d. 52.

the induced angle o attack increases. lif decreases and drag increases. lif increases and drag decreases. the aircraf will be partially supported on a cushion o air.

bottom to top round the trailing edge. top to bottom round the trailing edge. bottom to top round wing tip. top to bottom round wing tip.

Questions 56.

With flaps deployed, at a constant IAS in straight and level flight, the magnitude o tip vortices:

a. b. c. d. 57.. 57

7 1

CLMAX. CL squared. the square root o the C L. CL.

s n o i t s e u Q

Considering the lif to drag ratio, in straight and level flight which o the ollowing is correct?

a. b. c. d. 62.

aircraf weight. changes in thrust. angle between chord line and longitudinal axis. wing location.

CDi is proportional to which o the ollowing?

a. b. c. d. 61.. 61

increase by a actor o our our.. increase by a actor o two. decrease by a actor o two. decrease by a actor o our our..

At a constant IAS, induced drag is affected by:

a. b. c. d. 60.

increases induced drag. decreases induced drag. is structurally stiffer than a low aspect ratio. has a higher stall angle than a low aspect ratio.

An aircraf is flying straight and level; i density halves, aerodynamic drag will:

a. b. c. d. 59.

increases or decreases depending upon the initial angle o attack. increases. decreases. remains the same.

A high aspect ratio wing:

a. b. c. d. 58.

17

L/D is maximum at the speed or minimum total drag. L/D maximum decreases with increasing lif. L/D is maximum when lif equals weight. L/D is maximum when lif equals zero zero..

High aspect ratio:

a. b. c. d.

reduces parasite drag. reduces induced drag. increases stalling speed. increases manoeuvrability.

533

17

Questions 63.

How does aerodynamic drag vary when airspeed is doubled?

a. b. c. d. 64.

I dynamic (kinetic) pressure increases, what is the effect on total drag (i all important actors remain constant)?

a. b. c. d. 65.

1 7

68.

2 2 4 ¼

Increase. Decrease. Remain constant.

9 3 6 1·5

4 8 12 16

In a stream tube, i density is halved, drag will be reduced by a actor o:

a. b. c. d.

534

1

(ii) (ii) (ii) (ii)

I the IAS is increased by a actor o 4, by what actor would the drag increase?

a. b. c. d. 69.

2 4 ¼ /16

I the rontal area o an object in an airstream is increased by a actor o three, by what actor does drag increase?

a. b. c. d.

Q u e s t i o n s

(i) (i) (i) (i)

I pressure increases, with OAT and TAS constant, what happens to drag?

a. b. c. 67.. 67

Drag decreases. Drag increases. It has no effect on drag. Drag only changes with changing ground speed.

I IAS is increased rom 80 kt to 160 kt at a constant air density, TAS will double. What would be the effect on (i) C Di and (ii) Di?

a. b. c. d. 66.

4 16 1 2

3 4 6 2

Questions 70.

In straight and level flight, which o the ollowing would cause induced drag to vary linearly i weight is constant?

a. b. c. d. 71.

s n o i t s e u Q

Square o the speed. CLMAX. Speed. Surace area.

What effect does aspect ratio have on induced drag?

a. b. c. d. 77.. 77

7 1

elliptical pressure distribution increases. reduction in induced drag. decrease in stall speed. longitudinal static stability increases.

What does parasite drag vary with?

a. b. c. d. 76.

decreases energy. thinner. increased skin riction. less tendency to separate.

The effect o winglets is:

a. b. c. d. 75.

increased taper ratio. decreased aspect ratio. use o a wing tip with a thinner aerooil section. increased aspect ratio.

The advantage o a turbulent boundary layer over a laminar boundary layer is:

a. b. c. d. 74.

Parasite Parasit e drag greater than induced drag. CL and CD are minimum. Parasite Parasit e and induced drag are equal. Induced drag is greater than parasite drag.

Induced drag can be reduced by:

a. b. c. d. 73.

1/V. V. 1/V squared. V squared.

In subsonic flight, which is correct or V MD?

a. b. c. d. 72.

17

Increased aspect ratio increases induced drag. Increased aspect ratio reduces induced drag. Changing aspect ratio has no effect. Induced drag will equal 1·3 × aspect ratio/chord ratio.

What happens to total drag when accelerating rom C LMAX to maximum speed?

a. b. c. d.

Increases. Increases then decreases. Decreases. Decreases then increases.

535

17

Questions 78.

What is intererence drag?

a. b. c. d. 79.

What is the cause o induced angle o attack?

a. b. c. d. 80.

82.

83.

lif and drag increase linearly with an increase in angle o attack. lif and drag act normal to each other only at one angle o attack. lif and drag increase exponentially with an increase in angle o attack. lif increases linearly and drag increases exponentially with an increase in angle o attack.

When considering the properties o a laminar and turbulent boundary layer, which o the ollowing statements is correct?

a. b. c. d.

536

a turbulent boundary layer has more kinetic energy. a turbulent boundary layer is thinner thinner.. less skin riction is generat generated ed by a turbulent layer layer.. a turbulent boundary layer is more likely to separate separate..

When considering the aerodynamic orces acting on an aerooil section:

a. b. c. d. 84.

Wing tip vortices. Wing tanks. The increased pressure at the leading edge. The spanwise flow, inward below the wing and outward above.

When compared to a laminar boundary layer:

a. b. c. d.

Q u e s t i o n s

(CL) squared to S. (CL) squared to AR. ½ rho V squared. ½ rho V squared S.

What phenomena causes induced drag?

a. b. c. d. 1 7

Downwash rom trailing edge in the vicinity o the wing tips. Change in flow rom effective angle o attack. The upward inclination o the ree stream flow around the wing tips. Wing downwash altering the angle at which the airflow meets the tailplane.

CDi is the ratio o?

a. b. c. d. 81.

Airflow retardation over the aircraf structure due to surace irregularities. Drag caused by high total pressure at the leading edges when compared to the lower pressure pressure present at the trailing edge. Drag caused by the generation o lif. Drag due to the intera interaction ction o individual boundary layers at the junction o aircraf major components.

Friction drag is the same. Friction drag higher in laminar laminar.. Friction drag higher in turbulent. Separation point is most orward with a turbulent layer layer..

Questions 85.

When the undercarriage is lowered in flight:

a. b. c. d. 86.

V squared CL S V (CL) squared S V squared AR CD S V squared CD S

7 1

s n o i t s e u Q

Vortex generat generators ors diminish tip vortices. Flow on upper and lower wing suraces is towards the tip. They both decrease at high angle o attack. On the upper surace there is a component o flow towards the root, whilst on the lower surace it is i s towards the tip.

A jet aircraf flying at high altitude encounters severe turbulence without encountering high speed buffet. I the aircraf decelerates, what type o stall could occur first?

a. b. c. d. 91.

½ rho ½ rho ½ rho ½ rho

Which statement about induced drag and tip vortices is correct?

a. b. c. d. 90.

Constant velocity. Constant temperature. No flow normal to the surace. No vortices.

Which o the ollowing is the correct ormula or drag?

a. b. c. d. 89.

Wing ences. Anhedral. Winglets. Low aspect ratio planorm.

Which o the ollowing is a characteristic o laminar flow boundary layer?

a. b. c. d. 88.

orm drag will increase and the aircraf’s nose-down pitching moment will be unchanged. induced drag will increase and the aircraf’s nose-down pitching moment will increase. orm drag will increase and the aircraf’s nose-down pitching moment will increase. induced drag will decrease and the aircraf’s nose-down pitching moment will increase.

Which o the ollowing decreases induced drag?

a. b. c. d. 87.. 87

17

Low speed stall. Accelerated Acceler ated stall. Deep stall. Shock stall.

A swept wing aircraf stalls and the wake contacts the horizontal tail. What would be the stall behaviour?

a. b. c. d.

Nose down. Nose up and/or elevat elevator or ineffectiveness. Tendency to increase speed afer stall. Nose up.

537

17

Questions 92.

An aircraf aircraf at a weight weight o 237 402N stalls at 132 kt. At a weight weight o 356 103N it would stall at:

a. b. c. d. 93.

An aircraf at low subsonic speed will never stall:

a. b. c. d. 94.

1 7

97.. 97

b. c. d.

i rearward movement o the CG gives a reduced down-orce on the tail, V S will be higher h igher.. i orward movement o the CG gives a reduced down-orce on the tail, V S will be higher. i rearward movement o the CG gives a reduced down-orce on the tail, V S will be reduced. i rearward movement o the CG gives an increased down-orce on the tail, V S will be reduced.

How do vortex generators work?

a. b. c. d.

538

control stick pulled af. ailerons held neutral. control stick sideways against bank. control stick sideways towards bank.

Force on the tail and its effect on V S due to CG movement:

a.

98.

lif decreases, drag decreases. lif constant, drag increases. lif decreases, drag increases. lif decreases, drag constant.

During erect spin recovery the correct recovery actions are:

a. b. c. d.

Q u e s t i o n s

Upper surace, towards the leading edge. Lower surace, towards the trailing edge. Upper surace, towards the trailing edge. Lower surace, towards the leading edge.

At the point o stall:

a. b. c. d. 96.

as long as the CAS is kept above the power-on stall speed. as long as the IAS is kept above the power-on stall speed. as long as the maximum angle o attack is not exceeded. as long as the pitch angle is negative.

At high angle o attack, where does airflow separation begin?

a. b. c. d. 95.

88 kt. 162 kt. 108 kt. 172 kt.

Re-direct spanwise flow. Take energy rom ree stream and introduce it into the boundary layer layer.. Reduce kinetic energy to delay separation. Reduce the adverse pressure gradient.

Questions 99.

I a jet aircraf is at 60 degrees bank angle during a constant altitude turn, the stall speed will be:

a. b. c. d. 100.

7 1

s n o i t s e u Q

44 degrees. 30 degrees. 60 degrees. 32 degrees.

In recovery rom a spin:

a. b. c. d. 105.

100 kt. 140 kt. 80 kt. 119 kt.

In level flight at 1.4VS what is the approximate bank angle at which stall will occur?

a. b. c. d. 104.

122 kt. 150 kt. 81 kt. 100 kt.

I VS is 100 kt in straight and level flight, during a 45° bank turn V S will be:

a. b. c. d. 103.

83 kt. 70 kt. 85 kt. 60 kt.

I the straight and level stall speed is 100 100 kt, what will be the stall speed in a 1·5g turn?

a. b. c. d. 102.

1· 60 greater. 1· 19 greater. 1· 41 greater. 2· 00 greater.

I the stalling speed in a 15 degree bank bank turn is 60 kt, what would the stall speed be in a 45 degree bank?

a. b. c. d. 101.. 101

17

ailerons should be kept neutral. airspeed increases. ailerons are used to stop the spin. rudder and ailerons are used against the direction o spin rotation.

Stall speed in a turn is proportional to:

a. b. c. d.

lif. weight. the square root o the load actor actor.. TAS squared.

539

17

Questions 106.

Stalling speed increases when:

a. b. c. d. 107.

The angle o attack at the stall:

a. b. c. d. 108.

Q u e s t i o n s

c. d. 111.

200 kt. 119 kt. 141 kt. 100 kt.

What are the effects o tropical rain on: (i) CLMAX (ii) Drag

a. b. c. d.

540

take energy rom the laminar flow to induce boundary layer separation. use ree stream flow to induce laminar flow. prevent spanwise flow. use ree stream flow to increase energy in the turbulent boundary layer.

VS is 100 kt at n = 1; what will the stall speed be at n = 2?

a. b. c. d. 113.

increases with high altitude; more flaps; slats. may increase with increasing altitude, especially high altitude; orward CG and icing. decreases with orward CG and increasing altitude. altitude never affects stall speed IAS.

Vortex generators:

a. b. c. d. 112.

increase drag and increase stall speed. increase drag and decrease stall speed. decrease drag and increase stall speed. decrease drag and decrease stall speed.

The IAS o a stall:

a. b.

1 7

boundary layer ences and spanwise flow. tip stall o the wing. flow separation at the root due to spanwise flow. change in wing angle o incidence.

The effect o tropical rain on drag and stall speed would be to:

a. b. c. d. 110.

increases with orward CG. increases with af CG. decreases with decrease in weight. is not affected by changes in weight.

The CP on a swept wing aircraf will move orward due to:

a. b. c. d. 109.

recovering rom a steep dive. the aircraf is subjected to minor altitude changes, i.e. 0 to 10 000 f. the aircraf weight decreases. flaps are deployed.

(i) increase (ii) decrease (i) decrease (ii) increase (i) increase (ii) increase (i) decrease (ii) decrease

Questions 114.

What causes a swept wing aircraf to pitch-up at the stall?

a. b. c. d. 115.

7 1

s n o i t s e u Q

VS increases, stall angle remains constant. VS increases, stall angle increases. VS decreases, stall angle remains constant. VS decreases, stall angle decreases.

What is a high speed stall?

a. b. c. d. 121.

Increase. Decrease. Remain the same.

What influence does the CG being on the orward limit have on V S and the stall angle?

a. b. c. d. 120.

Increased anhedral increases stall speed. Fitting a ‘T’ tail will reduce stall speed. Increasing sweepback decreases stall speed. Decreasing sweep angle decreases stall speed.

What happens to the stall speed with flaps down, when compared to flaps up?

a. b. c. 119.

Activate at a certain angle o attack and pull the control column backwards. Activate at a certain angle o attack and push the stick orward. Activate at a certain IAS and vibrate the stick. Activate at a certain IAS and push the stick orward.

What effect on stall speed do the ollowing have?

a. b. c. d. 118.

CP moves af. CP moves orward. Root stall. Spanwise flow rom tip to root on wing upper surace.

What does a stick pusher do?

a. b. c. d. 117.

Negative camber at the root. Separated airflow at the root. Spanwise flow. Rearward movement o the CP.

What causes deep stall in a swept-back wing?

a. b. c. d. 116.

17

Separation o the airflow due to shock wave ormation. A stall caused by increasing the load actor (g) during a manoeuvre. A stall due to decreasing C LMAX at speeds above M 0.4. Excessive dynamic pressure causing airflow separation.

What is load actor?

a. b. c. d.

1 / Bank angle. Weight / Lif. Lif / Weight. Weight / Wing area.

541

17

Questions 122.

What is the percentage increase in stall speed in a 45° bank turn?

a. b. c. d. 123.

What is the standard stall recovery or a light aircraf?

a. b. c. d. 124.

c. d. 126.

Q u e s t i o n s

(ii) move orward. (ii) move af. (ii) not move. (ii) not move.

During the take-off run. The last part o rotation. Climb with all engines operating. All phases are equally important.

Which kind o stall occurs at the lowest angle o attack?

a. b. c. d.

542

(i) move af (i) move af (i) move af (i) not move

Which is the most critical phase regarding ice on a wing leading edge?

a. b. c. d. 128.

a swept-back wing will stall rom the root and the CP will move af. a non-swept rectangular wing will stall rom the root and the CP will move orwards. a non-swept rectangular wing will tend to stall rom the tip and the CP will move backwards. a swept-back wing will stall rom the tip and the CP will move orward.

When entering a stall, the CP o a straight rectangular wing (i) and a strongly swept wing (ii) will:

a. b. c. d. 127.

19%. 41%. 50%. 10%.

When an aircraf wing stalls:

a. b.

1 7

Pitch down, stick neutral roll, correct or bank with rudder. Pitch down, stick neutral roll, correct or bank with aileron. Pitch down, stick neutral, wait or neutral tendency. Pitch down, stick neutral roll, do not correct or bank.

What percentage increase in lif is required to maintain altitude while in a 45 degree bank turn?

a. b. c. d. 125.

45%. 41%. 19%. 10%.

Deep stall. Accelerated stall. Low speed stall. Shock stall.

Questions 129.

Which o the ollowing aircraf designs would be most prone to super stall?

a. b. c. d. 130.

7 1

s n o i t s e u Q

1.5 VS. 1.15 VS. 1.2 VS. Above VS.

Which o the ollowing is used to activate a stall warning device?

a. b. c. d. 135.

Increased drag. Increased weight. Blockage o the controls. Reduction in CLMAX.

Which o the ollowing is the speed that would activate the stick shaker?

a. b. c. d. 134.

VS1g VS1 VSO VSL

Which o the ollowing is the most important result/problem caused by ice ormation?

a. b. c. d. 133.

Swept wing and wing mounted engines. Swept wing and ‘T’ tail. Straight wing and wing mounted engines. Straight wing and ‘T’ tail.

Which o the ollowing is the correct designation o stall speed in the landing configuration?

a. b. c. d. 132.

‘T’ tail. Swept orward wing. Swept-back wing. Pod mounted engines beneath the wing.

Which o the ollowing combination o characteristics would be most likely make an aircraf susceptible to deep stall?

a. b. c. d. 131.

17

Movement o the CP. Movement o the CG. Movement o the stagnation point. A reduction in dynamic pressure.

Which o the ollowing would indicate an impending stall?

a. b. c. d.

Stall strip and stick shaker. Stall strip and angle o attack indicator. Airspeed indicator and stick shaker. Stick shaker and angle o attack indicator.

543

17

Questions 136.

Which stall has the greatest angle o attack?

a. b. c. d. 137.

With a swept wing the nose-up phenomena is caused by:

a. b. c. d. 138.

b. c.

Q u e s t i o n s

d. 140.

b. c. d.

Because VMCA with slats extended is more avourable compared to the flaps extended position. Because flaps extended gives a large decrease in stall speed with relatively less drag. Because slats extended provides a better view rom the cockpit than flaps extended. Because slats extended gives a large decrease in stall speed with relatively less drag.

An aircraf has trailing edge flap positions o 0, 15, 30 and 45 degrees plus slats can be deployed. What will have the greatest negative influence on C L / CD?

a. b. c. d.

544

increases the energy o the boundary layer and decreases the critical angle o attack. increases the wing leading edge radius by rotating orward and down rom its stowed position on the bottom side o the wing leading edge. deploys automatically under the influence o increased stagnation pressure at high angles o attack / low IAS. increases the energy o the boundary layer and increases the maximum angle o attack.

Afer take-off why are the slats (i installed) always retracted later than the trailing edge flaps?

a.

141.

increase because increasing weight increases the 1g stall speed. decrease because the 1g stall speed is an IAS. decrease because increasing weight increases the 1g stall speed. remain the same because increased weight increases the IAS that corresponds to a particular angle o attack.

A slat on an aerooil:

a.

1 7

deploying lif augmentation devices. wing ences. wing sweep prevents the nose-up phenomena. tip stall.

When flying straight and level in 1g flight, slightly below maximum all up weight, a basic stall warning system (flapper switch) activates at 75 kt IAS and the aircraf stalls at 68 kt IAS. Under the same conditions at maximum all up weight the margin between stall warning and stall will:

a. b. c. d. 139.

Low speed stall. High speed stall (shock stall). Deep stall. Accelerated stall.

Deploying slats. 0 - 15 flaps. 15 - 30 flaps. 30 - 45 flaps.

Questions 142.

Extending the flaps while maintaining a constant angle o attack (all other actors constant):

a. b. c. d. 143.

7 1

lif and drag increase. CLMAX increases. CL and drag increase. CL increases.

s n o i t s e u Q

I the angle o attack is maintained constant, what happens to the coefficient o lif when flaps are deployed?

a. b. c. d. 148.

Pitch up. Pitch down. Depends on CG position.

I flaps are extended in level flight:

a. b. c. d. 147.

Increases camber. Increases angle o attack. Changes position o CP. Decreases the Aspect Ratio.

How is the pitching moment affected i flaps are deployed in straight and level flight?

a. b. c. 146.

increases. decreases. increases or decreases depending on the aircraf. stays the same.

How does a plain flap increase C L?

a. b. c. d. 145.

the aircraf will sink suddenly. the aircraf will yaw. the aircraf will climb. the aircraf will roll.

For an aircraf flying straight and level at constant IAS, when flaps are deployed the induced drag:

a. b. c. d. 144.

17

Increased. Decreased. Changes with the square o IAS. Remains constant because angle o attack remains the same.

In order to maintain straight and level flight when trailing edge flaps are retracted, the angle o attack must:

a. b. c. d.

be increased or decreased depending on type o flap. be decreased. be increased. stay the same because the lif requirement will be the same.

545

17

Questions 149.

On a highly swept back wing with leading edge flaps and leading edge slats, which device would be fitted in the ollowing possible locations?

a. b. c. d. 150.

On a swept-back wing, in which o the ollowing locations would Krueger flaps be fitted?

a. b. c. d. 151.

b. c. d.

Q u e s t i o n s

153.

Increased minimum glide angle. Decreased minimum glide angle. Increased glide range. Decreased sink rate.

What is the purpose o a slat on the leading edge?

a. b. c. d.

546

Decrease CLMAX. Decrease the critical angle o attack. Not affect the critical angle o attack. Increase the critical angle o attack.

What is the effect o deploying trailing edge flaps?

a. b. c. d. 155.

greater. smaller. unchanged. smaller or greater, depending on CG position.

What is the effect o deploying leading edge flaps?

a. b. c. d. 154.

increase boundary layer energy, move suction peak on to slat and increase CLMAX angle o attack. increase camber, increase suction peak on main wing, increase effective angle o attack and move C LMAX to higher angle o attack. increase boundary layer energy, increase suction peak on main wing section, move CLMAX to a higher angle o attack. decrease boundary layer energy, move suction peak onto slat, move C LMAX to a lower angle o attack.

The maximum angle o attack or the flaps down configuration, compared to flaps up is:

a. b. c. d.

1 7

Inboard leading edge. Outboard leading edge. The leading edge. The trailing edge.

The effects o leading edge slats:

a.

152.

Slats inboard, leading edge flaps outboard. Slats outboard, leading edge flaps inboard. Alternate leading edge flaps and slats along the wing leading edge. There is no preerred position or these two devices.

Decelerate the air over the top surace. Thicken the laminar boundary layer over the top surace. Increase the camber o the wing. Allow greater angle o attack.

Questions 156.

What is true regarding deployment o slats / Krueger flaps?

a. b. c. d. 157.

d.

7 1

s n o i t s e u Q

Slats. Flaps. Spoilers. Ailerons.

A low wing jet aircraf is flaring to land. The greatest stick orce will be experienced with:

a. b. c. d. 162.

a higher angle o attack is required or maximum lif. glide distance is degraded. CLMAX decreases. VS1g increases.

Which o the ollowing increases the stall angle?

a. b. c. d. 161.

An aircraf nose-up pitching moment. An aircraf nose-down pitching moment. The nose-up pitching moment will be balanced by the nose-down pitching moment. The resultant aircraf pitching moment will depend upon the relative position o the CP and CG.

When trailing edge flaps are deployed:

a. b. c. d. 160.

Increase then decrease. Remain constant. Decrease. Increase.

What pitching moment will be generated when Fowler flaps are deployed on an aircraf with a high mounted (‘T’ tail) tailplane?

a. b. c.

159.

Slats increase the critical angle o attack, Krueger flaps do not. Krueger flaps increase the critical angle o attack, slats do not. Krueger flaps orm a slot, Slats do not. Slats orm a slot, Krueger flaps do not.

What must happen to the CL when flaps are deployed while maintaining a constant IAS in straight and level flight?

a. b. c. d. 158.

17

flaps up and CG at the af limit. flaps ully down and Cg at the af limit. flaps ully down and CG at the orward limit. flaps ully up and Cg at the orward limit.

Positive static lateral stability is the tendency o an aeroplane to:

a. b. c. d.

roll to the right in the case o a positive sideslip angle (aeroplane nose to the lef). roll to the lef in the case o a positive sideslip angle (aeroplane nose to the lef). roll to the lef in a right turn. roll to the right in a right turn.

547

17

Questions 163.

Positive static longitudinal stability means:

a. b. c. d. 164.

The CG o an aeroplane is in a fixed position orward o the neutral point. Speed changes cause a departure rom the trimmed position. Which o the ollowing statements about the stick orce stability is correct?

a. b. c. d. 165.

Q u e s t i o n s

167.

c. d.

The stick orce per ‘g’ increases when the CG is moved af. The stick orce per ‘g’ must have both upper and lower limits in order to assure acceptable control characteristics. I the slope o the e-n line becomes negative, generally speaking this is not a problem or control o an aeroplane. The stick orce per ‘g’ can only be corrected by means o electronic devices (stability augmentation) in the case o an unacceptable value.

What is pitch angle?

a. b. c. d.

548

(1) reduces (2) increases. (1) increases (2) increases. (1) increases (2) reduces. (1) reduces (2) reduces.

Which statement is correct?

a. b.

168.

too much aileron needed in a cross-wind landing. too much rudder needed in a cross-wind landing. constant aileron in cruise in a cross-wind.

What is the effect o an af shif o the CG on (1) static longitudinal stability and (2) the required control deflection or a given pitch change?

a. b. c. d.

1 7

An increase o 10 kt rom the trimmed position at low speed has more affect on the stick orce than an increase in 10 kt rom the trimmed position at high speed. Increase o speed generates pull orces. Aeroplane nose-up trim decreases the stick orce stability. Stick orce stability is not affected by trim.

Too much lateral static stability is undesirable because:

a. b. c. 166.

nose-down pitching moment when encountering an up gust. nose-up pitching moment with a speed change at a constant angle o attack. nose-down pitching moment with a speed change at a constant angle o attack. nose-up moment when encountering an up gust.

The angle between the chord line and the horizontal plane. The angle between the longitudinal axis and the horizontal plane. The angle between the chord line and the longitudinal axis. The angle between the relative airflow and the longitudinal axis.

Questions 169.

An aircraf o 50 tonnes mass, with two engines each o 60 000 N Thrust and with an L/D ratio o 12:1 is in a straight steady climb. Taking ‘g’ to be 10 m/s/s, what is the climb gradient?

a. b. c. d. 170.

7 1

s n o i t s e u Q

At constant TAS the Mach number decreases. At constant Mach number the IAS increases. At constant IAS the TAS decreases. At constant IAS the Mach number increases.

The regime o flight rom the critical Mach number (M CRIT) to approximately M 1.3 is called:

a. b. c. d. 175.

both (i) and (ii) are incorrect. (i) is incorrect and (ii) is correct. (i) is correct and (ii) is incorrect. both (i) and (ii) are correct.

Assuming ISA conditions, which statement with respect to the climb is correct?

a. b. c. d. 174.

the turn radius o ‘A’ will be greater than ‘B’. the coefficient o lif o ‘A’ will be less than ‘B’. the load actor o ‘A’ is greater than ‘B’. rate o turn o ‘A’ is greater than ‘B’.

VMCL can be limited by: (i) engine ailure during take-off, (ii) maximum rudder deflection.

a. b. c. d. 173.

Lif is less than weight, load actor is equal to one. Lif is less than weight, load actor is less than one. Lif is equal to weight, load actor is equal to one. Lif is equal to weight, load actor is less than one.

Two aircraf o the same weight and under identical atmospheric conditions are flying level 20 degree bank turns. Aircraf ‘A’ is at 130 kt, aircraf ‘B’ is at 200 kt:

a. b. c. d. 172.

12%. 24%. 15.7%. 3.7%.

In a straight steady descent:

a. b. c. d. 171.

17

transonic. hypersonic. subsonic. supersonic.

The speed range between high and low speed buffet:

a. b. c. d.

decreases during a descent at a constant Mach number. is always positive at Mach numbers below MMO. increases during a descent at a constant IAS. increases during climb.

549

17

Questions 176.

When does the bow wave first appear?

a. b. c. d. 177.

What can happen to the aeroplane structure flying at a speed just exceeding V A?

a. b. c. d. 178.

180.

Q u e s t i o n s

higher maximum thrust available. higher maximum efficiency. more blade surace area available. nearly maximum efficiency over wide speed range.

You are about to take off in an aircraf with a variable pitch propeller. At brake release: (i) Blade pitch and (ii) Propeller RPM lever:

a. b. c. d.

550

Increased L/DMAX, increased rate o descent. Decreased L/DMAX, increased rate o descent. Increased L/DMAX, decreased rate o descent. Decreased L/DMAX, decreased rate o descent.

The advantage o a constant speed propeller over a fixed pitch propeller is:

a. b. c. d. 182.

Neutral point. Aerodynamic Centre. Centre o Gravity. Centre o Thrust.

A single-engine aircraf with a constant speed propeller is in a gliding descent with the engine idling, what would be the effect o decreasing the propeller pitch?

a. b. c. d. 181.

Mass and pressure altitude. Mass only. Pressure altitude only. It remains a constant IAS.

With a vertical gust, what is the point called where the change in the vertical component o lif acts?

a. b. c. d. 1 7

It may suffer permanent deormation i the elevator is ully deflected upwards. It may break i the elevator is ully deflected upwards. It may suffer permanent deormation because the flight is perormed at to large a dynamic pressure. It will collapse i a turn is made.

Which o the ollowing can affect V A?

a. b. c. d. 179.

At MCRIT. At Mach 1. Just above Mach 1. Just below Mach 1.

(i) reduced, (i) reduced, (i) increased, (i) increased,

(ii) increase. (ii) decrease. (ii) decrease. (ii) increase.

Questions

17

7 1

s n o i t s e u Q

551

17

Answers

Answers

1 7

A n s w e r s

552

1 d

2 a

3 d

4 a

5 b

6 b

7 b

8 c

9 c

10 d

11 a

12 d

13 b

14 d

15 c

16 c

17 c

18 b

19 a

20 a

21 a

22 a

23 b

24 a

25 c

26 a

27 c

28 d

29 a

30 b

31 a

32 c

33 b

34 d

35 b

36 a

37 d

38 a

39 b

40 b

41 c

42 a

43 b

44 c

45 b

46 d

47 b

48 b

49 c

50 a

51 a

52 b

53 c

54 d

55 c

56 c

57 b

58 c

59 a

60 b

61 a

62 b

63 a

64 b

65 d

66 a

67 b

68 d

69 d

70 c

71 c

72 d

73 d

74 b

75 a

76 b

77 d

78 d

79 a

80 b

81 a

82 a

83 d

84 c

85 c

86 c

87 c

88 d

89 d

90 b

91 b

92 b

93 c

94 c

95 c

96 b

97 c

98 b

99 c

100 b

101 a

102 d

103 c

104 a

105 c

106 a

107 d

108 b

109 a

110 b

111 d

112 c

113 b

114 c

115 b

116 b

117 d

118 b

119 a

120 c

121 c

122 c

123 a

124 b

125 d

126 a

127 b

128 d

129 c

130 b

131 c

132 d

133 d

134 c

135 d

136 c

137 d

138 d

139 d

140 d

141 d

142 c

143 d

144 a

145 c

146 b

147 a

148 c

149 b

150 a

151 c

152 b

153 d

154 a

155 d

156 d

157 b

158 b

159 b

160 a

161 c

162 b

163 a

164 a

165 a

166 d

167 b

168 b

169 c

170 b

171 d

172 a

173 d

174 a

175 c

176 c

177 a

178 a

179 b

180 b

181 d

182 a

Answers

17

Explanations to Specimen Questions Q1

(d) psi is the “imperial” unit o pressure (pounds per square inch). Although not an SI unit it is still widely used in the aircraf industry or hydraulic pressure and air pressure – answer (d) is not only a unit o pressure, it is also the only possibly correct answer among those offered. The SI unit o pressure [orce per unit area] is newtons per square metre. This unit however is not on offer, but knowing this act helps eliminate some o the possible answers. Answer (a) is kilograms per square decimetre – the kilogram is the SI unit o mass, so straightaway this answer can be eliminated and answer (b) can be eliminated or the same reason. Answer (c) is incorrect because the newton is the SI unit o orce. Q2

(a) Density is mass per unit volume and the SI unit is kilograms per cubic metre. A orce is a push or a pull and the SI unit is the newton. Q3

(d) Density is mass (kg) per unit volume (cubic metre). Q4

(a) I a mass is accelerated a orce must have been applied. The kg is the SI unit or mass and m/s squared is the SI unit or acceleration. The applied orce can be determined by multiplying the mass by the acceleration and the answer must use the SI unit or orce - the Newton. Answer (b) is incorrect because psi is not an SI unit. Answer (c) is incorrect because the Joule is the SI unit or work. Answer (d) is incorrect because the Watt is the SI unit or power. Q5

(b) Acceleration (A) is proportional to orce (F) and inversely proportional to mass (M).

7 1

s r e w s n A

Q6

(b) Power is the rate o doing work - orce (N) × distance (m) divided by time (s). Q7

(b) The Airspeed Indicator is calibrated at Sea Level ISA density. To maintain a constant dynamic pressure (CAS) when below sea level (density higher) the TAS will have to be lower. [Q = density × TAS squared] Q8

(c) Static pressure is due to the weight o the atmosphere pressing down on the air beneath, so a body immersed in the atmosphere will experience an equal pressure in all directions due to Static pressure. Static pressure will exert the same orce per square metre on all suraces o an aeroplane. Q9

(c) True Airspeed (TAS) is the relative velocity between the aircraf and undisturbed air which is close to, but unaffected by the presence o the aircraf. Changing the TAS ( the speed o the aircraf through the air; the only speed there is) compensates or changes in air density and ensures a constant mass flow o air over the wing. I an altitude below ISA sea level is considered, the air density would be higher and thereore the TAS would have to be lower than IAS to compensate and keep Lif constant.

553

17

Answers

Q 10

(d) IAS is a measure o dynamic pressure, whereas TAS is the speed o the aircraf through the air. Changes in TAS are used to compensate or changes in air density to maintain a constant dynamic pressure. The lower the density, the higher the TAS must be to maintain a constant IAS. Answer (a) is incorrect because decreasing temperature increases air density, which decreases the difference between IAS and TAS. Answer (b) is incorrect because increasing air density decreases the difference between IAS and TAS. Answer (c) is incorrect because density changes with altitude. Q 11

(a) The Principle o Continuity states: “The product o the cross-sectional area, the density and velocity is constant” and Bernoulli’s theorem states: “Pressure plus kinetic energy is constant.” Subsonic airflow at speeds less than M 0.4 will not change the density significantly so density need not be considered. Through a divergent duct (increasing cross-sectional area in the direction o flow) velocity will decrease and static pressure will increase. Q 12

(d) The Principle o Continuity states: “The product o the cross-sectional area, the density and velocity is constant” and Bernoulli’s theorem states: “Pressure plus kinetic energy is constant. Subsonic airflow at speeds less than M 0.4 will not change the density significantly so density need not be considered. Through a convergent duct (decreasing cross-sectional area in the direction o flow) velocity will increase and static pressure will decrease. Q 13

(b) Bernoulli’s theorem states: “In the steady flow o an ideal fluid the sum o the pressure and kinetic energy per unit volume remains constant”. Put another way: Pressure + Kinetic energy = Constant.

1 7

Answer (a) is incorrect because it contradicts Bernoulli’s theorem. Answer (c) is not a true statement. Answer (d) is incorrect because there will be Static pressure.

A n s w e r s

Q 14

(d) Raising the temperature o the air in the streamtube will decrease its density. The Principle o Continuity states: “The product o the cross sectional-area, the density and velocity is constant”. Thereore, i the density o the air decreases, the mass flow must remain constant, and the velocity will increase. Q 15

(c) Bernoulli’s theorem states: “Pressure plus kinetic energy is constant”. Subsonic airflow at speeds less than M 0.4 will not change the density significantly so density need not be considered. Through a venturi, total pressure will remain constant. In the throat, dynamic pressure will increase and static pressure will decrease. Q 16

(c) Bernoulli’s theorem states: “In the steady flow o an ideal fluid the sum o the pressure and kinetic energy per unit volume remains constant”. Put another way: Pressure + Kinetic energy = Constant. ( page 44). Thereore, Total Pressure minus Static Pressure equals Dynamic Pressure.

554

Answers

17

Q 17

(c) The Principle o Continuity states: “The product o the cross-sectional area, the density and velocity is constant”. The question stipulates “subsonic and incompressible flow”, so the effects o density need not be considered. So i the cross-sectional area decreases the velocity will increase. Q 18

(b) The Principle o Continuity states: “The product o the cross-sectional area, the density and velocity is constant”. I the cross-sectional area increases the velocity o a subsonic and incompressible flow will decrease. Q 19

(a) Wing loading is the ratio o aircraf weight (a orce) to the wing area - newtons per square metre. Dynamic pressure is orce per unit area - also newtons per square metre. Q 20

(a) For incompressible subsonic flow it is assumed that the density o the air remains constant. Q 21

(a) See question 20. The answers merely use a different method o saying the same thing. The Greek letter RHO is the symbol or density. Q 22

(a) The Principle o Continuity and Bernoulli’s Theorem. In accordance with Bernoulli’s Theorem, it is the acceleration o the mass o air over the upper surace o the wing that creates lif. None o the other available answers are even remotely correct. Q 23

(b) Bernoulli’s Theorem states: In the steady flow o an “ideal” fluid the sum o the pressure and kinetic energy per unit volume remains constant. 7 1

Statement (i) is incorrect because the dynamic pressure in the throat o the venturi is higher than the ree stream flow.

s r e w s n A

Statement (ii) is correct. Q 24

(a) A line rom the centre o the leading edge to the centre o the trailing edge, equidistant rom the top and bottom surace is called the camber line, mean line or mean camber line. Q 25

(c) One o the advantages o a symmetrical aerooil section is that the pitching moment is zero. Answer (a) is incorrect because it is a positive camber aerooil section that gives a negative (nose-down) pitching moment. Answer (b) is incorrect because it is a negative camber aerooil section that will give a positive (nose-up) pitching moment. Answer (d) is incorrect because a symmetrical section at zero coefficient o lif will generate drag. Q 26

(a) The angle o attack is the angle between the relative airflow and the chord line. Undisturbed airflow is one o the conditions o relative airflow and is an acceptable alternative name or relative airflow. Answers (b) and (d) are obviously incorrect, but answer (c) is incorrect because the angle between local airflow and chord line is the Effective angle o attack.

555

17

Answers

Q 27

(c) The (maximum) thickness o an aerooil section is measured as a percentage o the chord. Q 28

(d) Lif is the aerodynamic orce that acts at right angles to the relative airflow and drag is the aerodynamic orce that acts parallel and in the same direction as the relative airflow. Q 29

(a) The angle between the chord line and longitudinal axis is called the angle o incidence which is fixed or a wing, but may be variable or the tailplane (horizontal stabilizer). Q 30

(b) Angle o attack is the angle between the chord line and the relative airflow. Relative wind is the American term or relative airflow and is an acceptable alternative. Q 31

(a) Angle o attack is the angle between the chord line and the relative airflow. (Relative ree stream flow is an acceptable alternative name or relative airflow). Aerodynamic angle o incidence is an out o date alternative name or angle o attack, which has gone out o general use to prevent conusion with the ANGLE OF INCIDENCE (the angle between the chord line and the longitudinal axis - fixed or a wing, but possibly variable or a tailplane). Q 32

(c) A symmetrical aerooil section at zero angle o attack will produce no lif, only drag. An “asymmetrical” section is what we reer to as cambered. Q 33

(b) Both Lif and Drag orces depend on the pressure distribution on the aerooil. 1. Total Reaction is split into two vectors: Lif, which acts at 90 degrees to the Relative Airflow and Drag, which is parallel to and in the same direction as the Relative Airflow - this is true or all “normal” angles o attack. 2. Lif varies linearly with angle o attack, but Drag varies exponentially. 3. The lif drag ratio at “normal” angles o attack is between approximately 10:1 and 20:1.

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Q 34

(d) Both top and bottom suraces o the aerooil contribute to lif, but the point along the chord where the distributed lif is effectively concentrated is termed the Centre o Pressure. Q 35

(b) This is a strange set o possible answers. The definition o an aerooil is: “A shape capable o producing lif with relatively high efficiency”. To work successully an aerooil does not need to be cambered, many symmetrical section aerooils are used on aircraf. To generate lif it is considered necessary to have an aerooil set at a suitable angle o attack and an air flow o reasonably high velocity, in the region between 65 kt and 180 kt, depending on the weight o the aircraf. Answer (a) is considered to be incorrect because a slightly cambered aerooil with no airflow will not create anything. Answer (c) is incorrect because accelerating air upwards will not create a lif orce. Answer (d) is considered to be incorrect because there is no NET acceleration o air downwards. Air upwashes in ront o an aerooil and downwashes behind - back to its original position. The Newton’s third law o motion explanation o lif generation is considered a allacy. There may be a very small amount o lif created in this way, but it is insignificant.

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Q 36

(a) The greatest contribution to overall lif comes rom the upper surace. Answer (b) is incorrect. Although a small amount o lif is generated by the increase in pressure beneath the wing, particularly at higher angles o attack, it always remains a small percentage o the total lif. Answer (c) is incorrect because the velocity below the wing is always lower. Answer (d) is incorrect because a decreased velocity on the top surace would increase the static pressure and lif would decrease. Q 37

(d) Drag acts parallel to and in the same direction as the relative wind (airflow). Lif acts at right angles (90 degrees or normal) to the relative wind. Q 38

(a) I IAS is doubled, dynamic pressure will be our times greater and lif will be our times greater. To maintain lif constant the angle o attack should be reduced to a quarter o its previous value. Q 39

(b) Thrust required is an alternative name or drag multiplied by TAS. On entering ground effect induced drag is decreased so thrust required is decreased. Answer (b) states less thrust is required. This means that to maintain the same IAS the throttle(s) should be closed urther. As the engine(s) are probably already at idle, the aircraf will accelerate. Q 40

(b) Ground effect becomes significant within hal the wingspan above the ground. Q 41

(c) L = 1/2 rho × V squared × CL × S Q 42

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(a) As an aircraf flies into ground effect (within hal the wingspan o the ground), its proximity to the ground will weaken the tip vortices. Downwash is decreased which decreases the induced angle o attack and increases effective angle o attack. Lif is increased and induced drag is decreased. It will take a greater distance rom the screen height or the aircraf to touch down. Because lif is increased there will be less weight on the wheels, making the brakes less effective. Because drag is decreased there is more work or the brakes to do. Thereore, landing distance will be increased by ground effect.

s r e w s n A

Q 43

(b) The only difference caused by the different altitudes will be the air density. To compensate or the decreased density at the higher altitude and maintain a constant Lif orce, the TAS o the higher aircraf will need to be greater. Q 44

(c) Lif = hal rho (Density) × Velocity (V) squared x Coefficient o Lif (CL) × The wing area (S). Hal rho (Density) × Velocity (V) squared = Dynamic pressure (Q). Q 45

(b) As an aircraf flies into ground effect (within hal the wingspan o the ground), its proximity to the ground will weaken the tip vortices. Downwash is decreased which decreases the induced angle o attack and increases effective angle o attack. Lif is increased and induced drag is decreased.

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Answers

Q 46

(d) For maximum aerodynamic efficiency it is necessary to generate enough lif to balance the weight, while at the same time generate as little drag as possible. The higher the Lif/Drag ratio the greater the aerodynamic efficiency. Thereore, at normal angles o attack CL is much higher than C D - between 10 and 20 times greater. Q 47

(b) A rectangular wing o this chord and the same span would have broadly similar pitching moment characteristics. The MAC is a primary reerence or longitudinal stability considerations. Q 48

(b) As an aircraf enters ground effect downwash is decreased which decreases the induced angle o attack and increases effective angle o attack. The lif vector, being at right angles to the effective airflow, will be inclined orwards and it is this orward inclination which reduces the induced drag (thrust required). Q 49

(c) Lif increases and induced drag decreases. When an aircraf enters ground effect the induced angle o attack decreases, which increases lif and reduces induced drag - making answers (a) and (b) incorrect. Answer (d) is obviously incorrect. Q 50

(a) The zero lif angle o attack or a positively cambered aerooil section is about minus 4 degrees. (Figure. 4.6 ). As the angle o attack is increased the C L will increase, at zero angle o attack the C L will be a small positive value - the lif curve will intersect the vertical C L axis above its point o origin. ( Figure 5.5).

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Q 51

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(a) (CLMAX is regarded as V S, making 1.3VS 30% aster than C LMAX). The lif ormula (L = hal rho × V squared × CL × S) can be transposed to give: (C L = L / hal rho × V squared × S). As density (hal rho), Lif (L) and wing area (S) are constant, this can be written: (C L is proportional to 1 / V squared). 1.3VS gives: 1 / 1.3 squared, which = 1 / 1.69, which = 0.59 = 59%. Q 52

(b) Air will flow rom areas o higher pressure towards areas o lower pressure. When a wing is generating lif, the air pressure on the top surace is lower than that outside the wing tip and, generally, air pressure on the bottom surace is slightly higher than that outside the wing tip. This causes air to flow inwards rom the tip towards the root on the top surace and outwards rom the root towards the tip on the bottom surace. The pressure difference at the wing tip will cause air to flow rom the bottom surace to the top surace around the wing tip and this rotating airflow generates the tip vortices. Q 53

(c) Aspect ratio is the ratio o the wingspan to the average or mean chord (AR = Span/Chord or Span squared/ Wing area). A high aspect ratio wing is one with a long span and a narrow chord, a low aspect ratio wing is one with a short span and a wide chord.

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Q 54

(d) Air will flow rom areas o higher pressure towards areas o lower pressure. When a wing is generating lif, the air pressure on the top surace is lower than that outside the wing tip and, generally, air pressure on the bottom surace is slightly higher than that outside the wing tip. This causes air to flow inwards rom the tip towards the root on the top surace and outwards rom the root towards the tip on the bottom surace. The pressure difference at the wing tip will cause air to flow rom the bottom surace to the top surace around the wing tip and this rotating airflow generates the tip vortices. Q 55

(c) Air will flow rom areas o higher pressure towards areas o lower pressure. When a wing is generating lif, the air pressure on the top sur ace is lower than that outside the wing tip and generally, air pressure on the bottom surace is slightly higher than that outside the wing tip. This causes air to flow inwards rom the tip towards the root on the top surace and outwards rom the root towards the tip on the bottom surace. The pressure difference at the wing tip will cause air to flow rom the bottom surace to the top surace around the wing tip and this rotating airflow generates the tip vortices. Q 56

(c) Wing tip vortices are strongest with the aircraf in the clean configuration. With flaps down, the flaps generate their own vortices which interere with and weaken the main, tip vortices. Q 57

(b) Increasing the aspect ratio, the ratio o the span to the mean chord, decreases the proportion o wing area affected by the tip vortices. (Only approximately one and a hal chord lengths in rom the tip are affected by tip vortices). This can best be remembered by reerring to the Induced Drag coefficient ormula: CDi is proportional to C L squared and inversely proportional to Aspect Ratio.

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Q 58

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(c) I density halves, drag will halve (decrease by a actor o 2). Q 59

(a) Generally speaking, induced drag is the result o lif production. I aircraf weight increases, more lif is required which will increase induced drag. Answer (b) is incorrect because thrust has no influence on induced drag. Answer (c) is incorrect because this is a definition o ‘angle o incidence’ which has no influence on induced drag. Answer (d) is incorrect because wing location has no influence on induced drag. Q 60

(b) CDi is proportional to C L squared and inversely proportional to the Aspect Ratio. Q 61

(a) Since flying at VMD incurs the least total drag or 1g flight, the aeroplane will also be at L/D MAX angle o attack (approximately 4 degrees). L/D max is a measure o aerodynamic efficiency and is a constant value or a given configuration. Answer (b) is incorrect because L/D MAX does not change with increasing lif, but the L/D ratio will change. Answer (c) is incorrect because lif can equal weight at any combination o angle o attack and IAS. Answer (d) is incorrect because with lif zero, there would still be drag which would decrease the L/D ratio.

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Answers

Q 62

(b) Increasing aspect ratio is the designer’s chie means o reducing induced drag. Answer (a) is incorrect because aspect ratio has no significant influence on parasite drag. Answer (c) is incorrect because increasing aspect ratio tends to increase C LMAX which reduces stall speed. Answer (d) is incorrect because increasing aspect ration decreases rate o roll. Q 63

(a) It has to be assumed that airspeed is increasing beyond V MD, when parasite drag will be dominant. Reer to the drag ormula. I airspeed is doubled, dynamic pressure will be our times greater due to the square unction o velocity. I dynamic pressure is our times greater, drag will be our times greater. Q 64

(b) With questions like this it is a good idea to reer to the appropriate ormula, in this case the drag ormula: /2 rho × V squared × C D × S. I dynamic pressure ( /2 rho × V squared) increases and the other actors remain the same, drag will increase. 1

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Q 65

(d) CDi is directly proportional to C L squared and inversely proportional to Aspect Ratio ( Page 121). Induced Drag (Di) = /2 rho × V squared × CDi × S (Page 121). I IAS is doubled, Dynamic Pressure will be our times greater and C L will need to be reduced to /4 o its previous value to maintain constant Lif. /4 squared = /16, making CDi /16 o its previous value. ‘Plugging’ all these new values into the Di ormula gives: Di = /2 rho × 4 × /16 × S, making Di /4 o its previous value. 1

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Q 66

(a) The Drag ormula shows that: D = /2 rho × V squared × CD × S. I pressure increases with outside air temperature (OAT) and True Airspeed (TAS) or (V) constant, density will increase. I density rho) increases, Drag will increase. 1

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Q 67

(b) Once again, using the Drag ormula is an easy way to remember the key acts. I the area ‘S’ is increased by a actor o 3, drag will also increase by a actor o 3. Q 68

(d) Due to the V squared unction, i IAS is increased by a actor o 4, our squared being sixteen times greater, drag will increase by a actor o 16. Q 69

(d) A streamtube is used to illustrate Bernoulli’s theorem. A streamtube is a streamlined flow o air with no losses. Drag is proportional to density, so i density is halved, drag will be halved. Q 70

(c) Induced Drag is inversely proportional to V squared (Induced drag is proportional to 1/V squared). Q 71

(c) Induced drag is dominant at low speed and Parasite drag is dominant at high speed. Because Induced drag decreases with an increase in speed and Parasite increases with an increase in speed - as speed is increased rom a low value, a speed will be reached at which Induced and Parasite drag are equal. This speed gives minimum Total drag and is known as V MD.

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Answer (a) is incorrect because at V MD Parasite drag and Induced drag are equal. Answer (b) is incorrect because C L is a minimum when no lif is produced and C D is a minimum when the aircraf is not moving, neither o which are practical propositions. Answer (d) is also incorrect because at V MD Parasite drag and Induced drag are equal. Q 72

(d) Increasing the aspect ratio, the ratio o the span to the mean chord, decreases the proportion o wing area affected by the tip vortices. (Only approximately one and a hal chord lengths in rom the tip are affected by tip vortices). This can best be remembered by reerring to the Induced Drag coefficient ormula: CDi is proportional to CL squared and inversely proportional to Aspect Ratio. Answer (a) is incorrect because with increasing speed rom the minimum level flight value, Total drag decreases, reaches a minimum value and then begins to increase. Answer (c) and (d) are incorrect because Parasite Drag and orm drag are both directly proportional to the square o the speed. Q 73

(d) A turbulent boundary layer is thicker and gives more skin riction, but has more resistance to separation due to its higher Kinetic Energy. A laminar boundary layer is thinner and gives less skin riction, but has a greater tendency to separate. Answer (a) is incorrect because a turbulent layer contains more kinetic energy. Answer (b) is incorrect because a turbulent layer is thicker. Answer (c) is incorrect because, though a true statement, increased skin riction is not an advantage. Q 74

(b) Winglets are small vertical aerooils which orm part o the wing tip; they reduce tip vortex intensity, thus reduce induced drag - this is their only unction. 7 1

Answer (a) is incorrect because the statement is not definitive. Answers (c) and (d) are incorrect because the only unction o winglets is to reduce tip vortex strength, thereby reducing induced drag.

s r e w s n A

Q 75

(a) Parasite Drag (Dp) varies with air density, velocity squared, coefficient o drag and area. [Dp = /2 rho × V squared × C D × S]. 1

Answer (b) and (d) are incorrect because any increase in Parasite area (drag) caused by increasing the angle o attack beyond the zero lif angle o attack is included in with induced drag. Answer (c) is less correct than (a) because although parasite drag does increase with speed it is the square o the speed to which parasite drag is proportional. Q 76

(b) Increasing aspect ratio decreases the proportion o the wing area effected by the tip vortices, thus reducing induced drag. Q 77

(d) Total Drag is made up o Parasite Drag and Induced Drag. Parasite Drag is directly proportional to the square o the IAS and Induced Drag is inversely proportional to the square o the IAS.

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At low speed, Total Drag is predominantly Induced Drag and at high speed, predominantly Parasite Drag. As speed is increased rom CLMAX, Induced Drag will be decreasing and Parasite Drag will be increasing. Eventually a speed will be reached when Parasite Drag will be the same as Induced Drag, this speed is the minimum (Total) drag speed (V MD). At speeds greater than V MD, Induced Drag will continue to decrease and Parasite Drag will continue to increase. From C LMAX Total Drag will decrease to a minimum at VMD the start to increase again Q 78

(d) Intererence drag is the result o boundary layer ‘intererence’ at wing/uselage, wing/ engine nacelle and other such junctions. Answer (a) is incorrect because this is a description o skin riction. Answer (b) is incorrect because this is a description o Form (pressure) drag. Answer (c) is incorrect because this is a description o Induced drag. Q 79

(a) Induced angle o attack is the angle between the relative airflow and the Effective airflow - a result o tip vortices increasing the downwash rom the wing trailing edge in the vicinity o the wing tips. Increasing vortex strength will increase downwash and increase the induced angle o attack Q 80

(b) CDi is the induced drag coefficient and is proportional to C L squared and inversely proportional to aspect ratio. Q 81

(a) Induced drag is due to the ormation o wing tip vortices. Answer (b) is incorrect because tip tanks reduce the strength o tip vortices, thereby reducing induced drag. Answer (c) is incorrect because increased pressure at the leading edge is the cause o Form (pressure) drag. Answer (d) is incorrect because this statement is wrong, in that spanwise flow is the opposite.

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Q 82

(a) A laminar boundary layer has less kinetic energy than a turbulent layer, is thinner and gives less skin riction but separates more easily, whereas a turbulent boundary layer contains more kinetic energy making it separate less easily, but will give more skin riction. Q 83

(d) With reerence to the lif curve on page 77 it can be seen that as angle o attack increases lif will increase linearly. With reerence to the drag ‘curve’ it can be seen that as angle o attack increases drag will increase exponentially (increase at an increasing rate). Q 84

(c) A laminar boundary layer has less kinetic energy than a turbulent layer, is thinner and gives less skin riction but separates more easily, whereas a turbulent boundary layer contains more kinetic energy making it separate less easily, but will give more skin riction. Answers (a) and (b) are incorrect because a turbulent boundary layer gives more skin riction than a laminar boundary layer. Answer (d) is incorrect: the turbulent boundary layer separates urther af because it contains more kinetic energy.

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Q 85

(c) Form drag is generated by the ore and af pressure differential on a body in an airflow, whereas Induced Drag is caused by the generation o lif (Wing tip vortices). I the undercarriage is lowered, the rontal area o the aircraf will increase which will increase orm drag. The undercarriage is below the aircraf CG, hopeully, giving a backwards orce below the aircraf CG, generating a nose-down pitching moment. Lowering the undercarriage will not affect induced drag. Q 86

(c) Winglets reduce the intensity o tip vortices, thus reducing induced drag. Answer (a) is incorrect because wing ences reduce spanwise flow and help minimize tip stalling. Answer (b) is incorrect because anhedral reduces lateral static stability. Answer (d) is incorrect because a low aspect ratio will increase the proportion o the wing area affected by the tip vortices and increase induced drag. Q 87

(c) The “key” characteristic o a laminar boundary layer is that there is no flow normal to the surace. Velocity is not constant because the relative velocity at the surace is zero and ull stream velocity at its outer limit. Neither is the temperature constant in a laminar flow boundary layer due to skin riction variations rom the surace to its outer limit. Just because a boundary layer contains no vortices does not make it laminar – a turbulent boundary layer may contain no vortices, but it has flow normal to the surace. Q 88

(d) Drag = /2 rho × V squared × CD × S 1

Q 89

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(d) Air will flow rom areas o higher pressure towards areas o lower pressure. When a wing is generating lif, the air pressure on the top surace is lower than that outside the wing tip and, generally, air pressure on the bottom surace is slightly higher than that outside the wing tip. This causes air to flow inwards rom the tip towards the root on the top surace and outwards rom the root towards the tip on the bottom surace. The pressure difference at the wing tip will cause air to flow rom the bottom surace to the top surace around the wing tip and this rotating airflow generates the tip vortices. ( page 86 ). Induced Drag is a result o the tip vortices inclining the effective airflow so as to decrease the effective angle o attack. To maintain the required lif orce, the whole wing must be flown at a higher angle o attack. This increase in angle o attack is called the induced angle o attack. Because Lif acts at right angles to the effective airflow the lif vector is Answer (a) is incorrect because vortex generators are used to re-energize the boundary layer in order to delay boundary layer separation. Vortex generators have no influence on either Induced Drag or tip vortices. ( page 159). Answer (b) is incorrect because flow on the upper surace o a wing is towards the root. Answer (c) is incorrect because both tip vortices and induced drag increase at high angles o attack (low IAS). This is due to the reduced chordwise vector at low IAS (high angles o attack) and the same pressure differential making the tip vortices stronger.

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Q 90

(b) An accelerated stall is a stall that occurs at a load actor greater than ‘1’ , in other words, at more than 1g. A stall can occur at any speed, but an accelerated stall, by definition, will occur at a speed higher than a 1g stall. The speed at which a deep stall occurs is difficult to define, but i the accepted indications o a stall are considered, a deep stall could be said to occur at a speed higher than the 1g stall speed, but at a speed less than that o an accelerated stall. As the question stipulates that the aircraf is decelerating a shock stall should not occur. Q 91

(b) By the time the separated airflow (wake) rom a stalled swept wing contacts the tailplane the aircraf would already be pitching-up due to tip stall. I the tailplane is immersed in separated airflow, the elevator will be ineffective. Q 92

(b) The ormula to calculate the effect o weight change on stall speed is: New stall speed equals the old stall speed multiplied by the square root o the new weight divided by the old weight. Q 93

(c) Stalling is due to airflow separation and airflow separation depends upon the relationship between the boundary layer kinetic energy and the adverse pressure gradient. Generally speaking, the adverse pressure gradient is a unction o angle o attack; the adverse pressure gradient will increase with an increase in angle o attack. Thereore, stalling is due to exceeding the critical angle o attack and has nothing to do with IAS, in and o itsel. Answers (a) and (b) are incorrect or the reasons given above. Answer (d) is incorrect because pitch angle is the angle between the aircraf’s longitudinal axis and the horizon, which has nothing whatsoever to do with angle o attack - the angle between the relative airflow and the chord line. The relative airflow direction is parallel to and in the opposite direction to the flight path, so an aircraf whose nose was below the horizon could still easily be at an angle o attack greater than the critical angle o attack.

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A n s w e r s

Q 94

(c) As angle o attack increases, the increasing adverse pressure gradient in the presence o a constant boundary layer kinetic energy causes the boundary layer to start to separate first at the trailing edge. Increasing angle o attack/adverse pressure gradient moves the separation point orward. Q 95

(c) At the point o stall lif decreases and drag continues to increase. Q 96

(b) The ailerons should either be held neutral or returned to neutral. Answer (a) is incorrect because the angle o attack must be decreased to unstall the wing. Answers (c) and (d) are incorrect or general spin recovery because rudder should be used against the direction o spin to equalize the angle o attack on both halves o the wing, thus preventing urther autorotation.

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Q 97

(c) For the vast majority o aircraf the CP is af o the CG and as the aircraf rotates around the CG this will generate a nose-down pitching moment. A tail downorce is required to generate the equal and opposite nose up pitching moment required or equilibrium. The tail downorce gives an ‘effective’ increase in weight which requires a slight increase in lif to maintain the balance o up and down orces. This increase in lif increases the stall speed. The urther orward the CG, the greater the increase in stall speed. Q 98

(b) Vortex generators are rows o small thin blades which project vertically about 2.5 cm into the airstream. They each generate a small vortex which causes the ree stream flow o high energy air to mix with and add kinetic energy to the boundary layer. This re-energizes the boundary layer and delays separation. Answer (a) is incorrect because vortex generators only have an effect immediately downstream o their location, so will not significantly influence spanwise flow. Answer (c) is incorrect because vortex generators re-energize the boundary layer. Answer (d) is incorrect because vortex generators to not directly influence air pressure. Q 99

(c) Increase in stall speed in a level banked turn is proportional to the 1g stall speed multiplied by the square root o the load actor or 1/cos phi. Q 100

(b) It is always necessary to use the 1g stall speed to determine the increased stall speed in a bank. In this case it is necessary to transpose the ormula in order to determine the 1g stall speed. Q 101

(a) ‘g’ is the colloquial symbol or load actor. Load actor is the relationship between Lif and Weight. When an aircraf is banked in level flight, Lif must be greater than Weight and the relationship can be calculated by using the ormula: L = 1/cos phi (where phi = bank angle). To calculated the stall speed in a 1.5g turn, multiply the 1g stall speed by the square root o 1.5, in this case 1.22. 100 × 1.22 = 122 kt. [It can be said that ‘g’ is the same as 1/cos phi].

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Q 102

(d) Stall speed in a turn equals the 1g stall speed multiplied by the square root o the Load Factor and the Load Factor in a turn equals 1 divided by the cosine o the bank angle. ( page 171). The cosine o 45 degrees is 0.707. 1 divided by 0.707 equals 1.41 (this is a 41 percent increase in lif). The square root o 1.41 equals 1.19. 100 kt multiplied by 1.19 equals 119 kt (a 19 percent increase in stall speed). Q 103

(c) The effect o bank angle on stall speed can be visualized by reerence to the geometry o a vector diagram o Weight, Lif and bank angle – Re: Figure 7.23. The trigonometry ormula VS1g multiplied by the square root o 1 divided by the cosine o the bank angle (phi) is used to calculate the actual change in stall speed at various bank angles. To answer this question it is necessary to transpose the trig’ ormula mentioned above to: cos phi = /1 .4 squared (1.4 being the margin above stall speed at which the aircraf is 1

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Q 104

(a) The recommended, but general, spin recovery technique is ull opposite rudder (against spin direction), reduce throttle(s) to idle, neutralize the ailerons and gently but progressively apply orward pitch control. All our actions can be accomplished simultaneously. Q 105

(c) Increase in stall speed in a level banked turn is proportional to the 1g stall speed multiplied by the square root o the load actor or 1/cos phi. Q 106

(a) Stall speed varies with wing contamination, configuration (flaps & gear), thrust and prop slipstream, weight, load actor, mach number and CG position. Load actor varies with manoeuvring and turbulence. When recovering rom a steep dive load actor will increase and stall speed will increase. Answer (b) is incorrect because lower level altitude changes will not effect stall speed; it is only at very high altitude that stall speed increases with altitude (approx. 29 000 f). Answer (c) is incorrect because a decrease in weight will decrease stall speed. Answer (d) is incorrect because deploying flaps decreases stall speed. Q 107

(d) Stall angle is affected by flaps and wing contamination. CG movement and changes in Weight will only affect stall speed. Q 108

(b) The increased tendency o a swept wing to stall rom the tips (over and above the tendency o a tapered wing to stall rom the tips) is due to the root to tip spanwise flow on the top surace. Because the tips stall beore the root, the CP moves orward, giving the tendency to ‘pitch-up’.

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A n s w e r s

Answer (a) is incorrect because boundary layer ences are fitted to reduce spanwise flow and thereore reduce the tendency o a swept wing to tip stall; their unction does not directly affect CP movement. Answer (c) is incorrect because the spanwise flow on the top surace o a swept wing is rom root to tip, which increases the tendency or airflow separation at the tip - the cause o CP orward movement. Answer (d) is incorrect because wing incidence is a fixed value - the angle between the chord line and the longitudinal axis. Q 109

(a) Tropical rain increases weight and distorts the aerodynamic shape, thus decreasing lif and increasing drag. The weight increase and the lif decrease will increase the stall speed. Q 110

(b) Stall speed varies: at very high altitude due to Mach number, with flap and slat position, with CG position, with wing contamination, with load actor, with engine thrust and propeller slipstream and with weight. Answer (a) is incorrect because though stall speed does increase at high altitude, more flaps or slats will decrease stall speed. Answer (c) is incorrect because stall speed increases with orward CG and unless high altitude is mentioned, moderate altitudes do not affect stall speed. Answer (d) is incorrect because very high altitude does affect stall speed because o increasing Mach number.

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Q 111

(d) Vortex generators are rows o small thin blades which project vertically about 2.5 cm into the airstream. They each generate a small vortex which causes the ree stream flow o high energy air to mix with and add kinetic energy to the boundary layer. This re-energizes the boundary layer and delays separation. Q 112

(c) ‘g’ is the colloquial symbol or load actor. Load actor is the relationship between Lif and Weight. When an aircraf is banked in level flight, Lif must be greater than Weight and the relationship can be calculated by using the ormula: L = 1/cos phi (where phi = bank angle). To calculated the stall speed in a 2g turn, multiply the 1g stall speed by the square root o 2, in this case 1.41. 100 × 1.41 = 141 kt. [It can be said that ‘g’ is the same as 1/cos phi]. Q 113

(b) Tropical rain increases Weight and distorts the aerodynamic shape, thus decreasing lif (CLMAX) and increasing drag. Q 114

(c) Pitch-up o an aircraf fitted with a swept wing is caused by tip stall. The increased tendency o a swept wing to tip stall is due to an induced spanwise flow o the boundary layer rom root to tip on the top surace. Q 115

(b) ‘Deep stall’ or ‘Super stall’ is the possible final result o ‘pitch-up’. ‘Pitch-up’ is caused by orward movement o the CP on a swept-back wing. Q 116

(b) A stick pusher is an automatic device which activates at a certain angle o attack and physically pushes the stick orward to prevent the angle o attack increasing beyond a certain maximum value. This device is fitted to aircraf that exhibit excessive pitch-up at the stall to prevent super stall.

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Q 117

(d) A swept wing has a smaller lif curve gradient and a reduced CLMAX. The reduced CLMAX gives an increased stall speed. Reducing sweep angle will thereore reduce stall speed. Answer (a) is incorrect because the unction o anhedral is to reduce static lateral stability and has no influence on stalling. Answer (b) is incorrect because a ‘T’ tail is generally incorporated in an aircraf design to remove the tailplane rom the influence o downwash rom the wing and its location on the top o the fin has no influence on stall speed. Answer (c) is incorrect because increasing the sweep angle decreases C LMAX which increases stall speed. Q 118

(b) The purpose o high lif devices is to reduce the take-off and landing run. This is generally achieved by increasing the wing camber, resulting in an increased C LMAX. Increasing CLMAX will decrease the stall speed and, hence, the minimum operational speed. Q 119

(a) The CG on the orward limit gives a large nose-down pitching moment. This must be balanced by a tail downorce. The tail downorce is an effective increase in weight which requires more lif. The increase in lif required increases the stall speed, but the angle at which the wing stalls remains constant at approximately 16 degrees.

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Answers

Q 120

(c) At speeds higher than M 0.4 the proximity o the aircraf to its leading pressure wave increases upwash. This decreases C LMAX which gives a higher stall speed. Answer (a) is incorrect because the separation o airflow due to shock wave ormation is known as “Shock Stall”. Answer (b) is incorrect because a stall that occurs due to load actors greater than 1 is known as an “Accelerated Stall”. Answer (d) is incorrect because dynamic pressure itsel has no influence on airflow separation. Q 121

(c) Load actor (‘n’ or ‘g’) is the ratio o Lif to Weight or Lif / Weight. Q 122

(c) To calculate the increase in Lif in a 45 degree bank: L = 1 / cos 45 or 1.41. The increase in stall speed is the square root o 1 / cos 45 or 1.19. This is a 19 percent increase in stall speed. Q 123

(a) Stalling is caused by airflow separation, which generally is due to exceeding the critical angle o attack. The standard procedure to prevent a ull stall or to recover rom a stall is to decrease the angle o attack. It is recommended that the nose be lowered to or slightly below the horizon to reduce the angle o attack and at the same time apply maximum power to minimize height lost during stall recovery. For small aircraf it is recommended that wings level be maintained by use o the rudder. I ailerons are used, the down-going aileron may ully stall the lower wing and make it drop aster. Answer (b) is incorrect because this is the recommended stall recovery or a swept wing aircraf - whose stall warning margin to C LMAX is sufficient or roll control still to be effective. Answer (c) is incorrect because all stall recovery techniques require positive pilot action to regain ull control o the aircraf. Answer (d) is incorrect because rudder is recommended to lif a dropping wing on small aircraf; not correcting or a dropped wing could leave the aircraf in a 90 degree bank!!

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Q 124

(b) The extra lif required when the aircraf is banked is dependant upon the bank angle. The lif required to maintain a constant vertical orce to oppose the weight is proportional to the length o the hypotenuse o a right angled triangle. In this case, Lif = 1 / cos 45 = 1.41 which is a 41% increase in Lif. Q 125

(d) A swept-back wing stalls rom the tip and the CP moves orward, giving the tendency to pitch-up. Answer (a) is incorrect because a swept wing stalls rom the tip and the CP moves orward. Answer (b) is incorrect because a rectangular wing will stall rom the root, but the CP moves rearwards. Answer (c) is incorrect because a rectangular wing will stall rom the root and the CP moves rearwards. Q 126

(a) When approaching the critical angle o attack with a rectangular wing the upper suction peak flattens and begins to collapse due to airflow separation. The lower surace pressure distribution is not immediately affected, resulting in the CP moving af. Because the swept wing stalls rom the tip and the tip is located behind the aircraf CG, at the stall the CP o a swept-back wing moves orwards. (Figure 7.14).

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Q 127

(b) Ice on a wing leading edge will produce large changes in the local contour (o the aerooil section), leading to severe local adverse pressure gradients. This will cause the wing to stall at a much smaller angle o attack than would occur with an aerodynamically clean wing. Angle o attack during the last part o rotation may well exceed the lower, icing induced critical angle o attack, thus preventing the aircraf rom becoming airborne. Speed will be greater than V 1, so it may not be possible to stop within the remaining take-off distance available. Answer (a) is incorrect because during the take-off run only the extra drag rom the ice would be a actor and while increasing the take-off run this would not be the most critical phase. Answer (c) is incorrect because during a steady climb with all engines operating the angle o attack should be that which gives L/D MAX (4 degrees), this should be ar enough below the lower icing induced critical angle o attack that it will not to be a actor. Answer (d) is incorrect because at the majority o phases o flight the angle o attack will be much below the icing induced critical angle o attack. Q 128

(d) Angle o Attack is the angle between the Relative Airflow and the chord line. A stall is caused by airflow separation and separation can occur when either the boundary layer has insufficient kinetic energy or the adverse pressure gradient is too great. Generally speaking: adverse pressure gradient is increased by increasing the angle o attack. Answer (d) is correct because shock stall is caused by the presence o the shock wave on the wing top surace. It is the shock wave that causes a marked increase in adverse pressure gradient; the angle o attack remains small. Answer (a) is incorrect because a deep stall is the “automatic” progression o tip stall on a swept wing leading to pitch-up, and would occur at a relatively large angle o attack. Answer (b) is incorrect because an accelerated stall is one that occurs at greater than 1g, but at the “normal” high angle o attack o 16 degrees. Answer (c) is incorrect because the low speed stall also occurs at 16 degrees angle o attack.

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Q 129

(c) A swept-back wing tends to stall first near the tips which moves the CP orward causing a phenomena known as ‘pitch-up’. This tends to urther increase the angle o attack, reducing lif and the aircraf can start to sink rapidly, urther increasing the angle o attack. Separated airflow rom the ully stalled swept wing can then immerse a ‘T’ tail and reduce elevator effectiveness and prevent recovery. It is the tip stalling o the swept-back wing which causes super stall. Answer (a) is incorrect because the contribution o the ‘T’ tail is to make super stall recovery more difficult, it does not cause deep stall. Answer (b) is incorrect because a swept orward wing would experience a rearward movement o the CP ollowing tip stall, which would give an aircraf nose-down pitching moment. Answer (d) is incorrect. The pylons o pod mounted engines below the wing act as vortilons and reduce spanwise flow which can lead to tip stall. Q 130

(b) The primary cause o ‘deep stall’ is ‘pitch-up’ which is the result o tip stalling o a swept wing. A contributory actor is a ‘T’ tail, which may place the tailplane and elevator in the path o the separated airflow rom the wing once ‘pitch-up’ has occurred - this reduces the effectiveness o the elevator and may prevent prompt recovery rom the ‘deep stall’. It must be emphasized that ‘pitch-up’ is the primary cause o deep stall and a ‘T’ tail is a contributory actor.

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Answers

Answer (a) is incorrect because wing mounted engines do not contribute to deep stall. Answer (c) is incorrect because neither an unswept wing nor wing mounted engines contribute to deep stall. Answer (d) is incorrect because an unswept wing does not ‘pitch-up’ at the stall and a ‘T’ tail is only a contributory actor i the aircraf suffers ‘pitch-up’. Q 131

(c) VS0 means “The stall speed or the minimum steady flight speed in the landing configuration”. Answer (a) is incorrect because V S1g means “The stall speed at which the aeroplane can develop a lif orce (normal to the flight path) equal to its weight”. Answer (b) is incorrect because V S1 means “The stall speed or the minimum steady flight speed obtained in a specified configuration”. Answer (d) is incorrect because there is no such designation. Q 132

(d) Reerence Figure 7.31. There will be a significant decrease in CLMAX due to ice ormation on the wing. This is due to a radical change in aerooil section contour. This effect could cause the aircraf to stall in the cruise and is o greater consequence than any o the other possible answers. Answer (a) is not the preerred answer; although drag will be increased by ice ormation and will be a contributory actor. Answer (b) is not the preerred answer, although increased weight will increase the stall speed. Answer (c) is a significant effect, but is still less significant than the reduction in C LMAX.

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Q 133

(d) CS-25 specifies that the stall warning must begin at 5 kt or 5% beore the stall, whichever is the greater. This equates to 1.05V S.

A n s w e r s

Answers (a), (b) and (c) are incorrect because they all exceed 1.05V S. Q 134

(c) Any stall warning device must be sensitive to changes in angle o attack. As angle o attack increases, the stagnation point will move downwards and backwards rom the leading edge to a position slightly below. A small sensitive electric switch attached to a vane can be positioned at the leading edge so that downwards and backwards movement o the stagnation point moves the vane up and closing a circuit, activates the stall warning device. Answer (a) is incorrect. Although the CP does move with changes in angle o attack, its movement cannot be detected. Answer (b) is incorrect because the CG does not move with changing angle o attack. Answer (d) is incorrect because stalling is due to increasing angle o attack causing airflow separation. The critical angle o attack can be exceeded at any dynamic pressure. Q 135

(d) Stalling is the result o airflow separation. Airflow separation is due to the combined effect o adverse pressure gradient and boundary layer kinetic energy. Adverse pressure gradient increases with increasing angle o attack. A stall warning device must thereore be sensitive to angle o attack. A stick shaker is signalled by a device sensitive to changes in angle o attack - a

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flapper switch (leading edge stall warning vane), an angle o attack vane or an angle o attack probe. An angle o attack indicator, i fitted, will indicate to the flight crew the angle o attack at any moment. Answer (a) & (b) are incorrect because a stall strip is a device used to encourage a stall to occur by locally decreasing the leading edge radius. Answer (c) is incorrect because Indicated Airspeed (IAS) is not a reliable indicator o an impending stall because the critical angle o attack can be exceeded at any IAS. Q 136

(c) As an aircraf enters a deep stall the ‘pitch-up’ tendency increases the angle o attack to a very high value and the aircraf also starts to sink, which urther increases the angle o attack. Answer (a) is incorrect because low speed (1g) stall occurs at approximately 16 degrees. Answer (d) is incorrect because an accelerated stall also occurs at about 16 degrees. Answer (b) is incorrect because a high speed (shock) stall occurs at a small angle o attack, being due to shock induced airflow separation above MCRIT. Q 137

(d) A swept wing stalls rom the tip which is behind the aircraf CG. As the portion o the swept wing in ront o the CG is still producing lif, an aircraf nose-up pitching moment is generated - this phenomena is known as ‘pitch-up’. Answer (a) is incorrect because, generally speaking, deploying flaps will give a modern high speed jet transport aircraf a nose-down pitching moment. Answer (b) is incorrect because wing ences are fitted to minimize spanwise airflow and will not themselves cause a reaction rom the aircraf. Answer (c) is incorrect because one o the major disadvantages o sweepback is the pitch-up phenomena. Q 138

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(d) Stalling is caused by airflow separation. The amount o airflow separation is due to the relationship between the adverse pressure gradient and boundary layer kinetic energy. The adverse pressure gradient will increase i angle o attack is increased. A 1g stall occurs at the critical angle o attack (CL MAX). A stall warning must begin with sufficient margin to prevent inadvertent stalling, so a stall warning device must also be sensitive to angle o attack. Thereore, a 1g stall will occur at the critical angle o attack and the stall warning will activate at an angle o attack which is slightly less. In 1g flight each angle o attack requires a particular IAS (dynamic pressure). An increase in weight will not alter the respective angles o attack, but will increase both the IAS at which the stall warning activates and the IAS at which the 1g stall occurs, but the margin between them will remain essentially the same.

s r e w s n A

Q 139

(d) Slats increase boundary layer kinetic energy which delays separation to a higher angle o attack (adverse pressure gradient). In and o themselves, deploying slats do nothing, but they enable a higher angle o attack to be used, thus decreasing the minimum operational speed. Slats also increase the critical angle o attack (Re: Figure 8.15). Answer (a) is incorrect because slats increase the critical angle o attack. Answer (b) is incorrect because this is a description o a Krueger flap. Answer (c) is incorrect on all counts. Q 140

(d) Extended slats do not change C L or CD significantly; they do, however, increase CLMAX and thereore give greater margin to the stall speed.

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Answers

Answering this question correctly requires a ull understanding o the properties o slats. Consider answer (a): VMCA is the minimum IAS at which directional control can be maintained ollowing ailure o the critical engine. Page 392 discusses the indirect effect o trailing edge flap position on V MCA, but the “intent” o the question is very precise, in that all types o aircraf (both turbo jet and propeller) are implied by the wording “..... why are the slats always retracted ....“. Also, propeller driven aircraf are unlikely to be fitted with slats - there is no need with their relatively unsophisticated wing design. Answer (b) can be shown to be incorrect by reerring to page 220. Answer (c) is incorrect because with trailing edge flaps extended there is a requirement or a greater aircraf nose-down pitch angle to maintain a given C L, this gives a better view over the nose rom the flight deck, with slats extended however, this is not the case. Q 141

(d) Reerence Figures 8.18 and 19. Deploying trailing edge flaps increases both C L and CD. At small flap angles there is a greater percentage increase in C L than CD. However, urther flap deployment gives a smaller percentage increase in C L and a larger percentage increase in C D. Any amount o flap deployment will decrease the maximum L / D ratio; the greater the flap angle used, the greater will be the decrease in L / D MAX. Q 142

(c) Reer to Figure 8.22. It can be seen that moving rom position ‘A’ to position ‘C’ ulfils the inormation provided in the question. CL would increase, Lif would be greater than Weight and the aircraf would gain altitude. Q 143

(d) Reerence Figure 8.22. For an aircraf to maintain level flight when flaps are deployed the angle o attack must be decreased to maintain a constant C L. CDi is proportional to C L squared and inversely proportional to Aspect Ratio. Thereore, induced drag will remain the same. 1 7

Q 144

(a) Trailing edge high lif devices unction by increasing the camber o the wing, thus increasing CLMAX and CL or a given angle o attack.

A n s w e r s

Answer (b) is incorrect because angle o attack is the angle between the chord line and the relative airflow. Lif curves or flaps are drawn with reerence to the original chord line. A plain flap has a decreased stall angle, so a smaller maximum angle o attack is possible. Answers (c) is incorrect because though the statement is correct changing the position o the CP does not alter the C L. Answer (d) is incorrect because a plain flap has no significant influence on either wing chord or span. Q 145

(c) Centre o Pressure movement will generate a nose-down pitching moment, whereas the change in downwash generates a nose-up pitching moment. The resultant aircraf pitching moment will depend upon which o these two moments is dominant. From the inormation given in the question it is not possible to say what aircraf pitching moment will result, so the only answer that can be given is: “It depends”. Q 146

(b) For level flight lif must remain the same as the weight, so as flaps are extended the angle o attack must be decreased. Flaps, generally speaking, increase the camber which increases CLMAX. Answers (a), (b) and (c) are incorrect because C L and lif remain constant and drag increases.

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Q 147

(a) Re to Figure 8.22. It can be seen that moving rom position ‘A’ to position ‘C’ ulfils the inormation provided in the question. CL would increase. Q 148

(c) Reerence Figure. 8.20. Moving rom Point ‘B’ to Point ‘C’ ulfils the requirements o the question and illustrates the need to increase the angle o attack as flaps are retracted in order to maintain CL constant. Q 149

(b) To preserve the tendency or root stall first on a swept wing, the least efficient leading edge high lif device is fitted inboard. The ollowing list shows leading edge devices in order o increasing efficiency: Krueger, Variable Camber, Slat. Q 150

(a) To preserve the tendency or root stall first on a swept wing, the least efficient leading edge high lif device is fitted inboard. The ollowing list shows leading edge devices in order o increasing efficiency: Krueger, Variable Camber, Slat. Answer (d) is incorrect because a Krueger Flap is a leading edge device. Q 151

(c) Reerence Figure 8.16 . Slats increase boundary layer kinetic energy and enable a higher angle o attack to be used. It can be seen that the “suction” peak does not move orward onto the slat and has no significant effect on the pitching moment. Q 152

(b) Reerence Figure 8.5. Angle o attack is the angle between the Relative Airflow and the ORIGINAL chord line - with no flap deployed. Trailing edge flaps decrease the critical angle o attack.

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Q 153

(d) Reerence Figure 8.13. Angle o attack is the angle between the Relative Airflow and the ORIGINAL chord line - with no flap deployed. Leading edge devices increase the critical angle o attack. Q 154

(a) ANY deployment o flap decreases the L/D ratio. Glide angle is a unction ONLY o the L/D ratio. Thereore, deploying flaps will decrease L/D ratio which will increase glide angle, decrease glide distance and increase sink rate. Q 155

(d) Reerence Figure 8.15. At a given angle o attack, deploying slats does nothing, but enables a higher angle o attack (adverse pressure gradient) to be used without airflow separation, thus decreasing the minimum operational IAS. Answer (a) is incorrect because slats increase boundary layer kinetic energy. Answer (b) is not the preerred answer, but is a true statement. Because slats increase boundary layer kinetic energy the boundary layer will have a slightly increased thickness, but this is not the purpose o slats. Answer (c) is not the correct answer because slats do not significantly increase wing camber.

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Q 156

(d) Reerence Figures 8.11 and 8.14. Answers (a) and (b) are incorrect because both slats and Krueger flaps increase the critical angle o attack ( Figures 8.13 and 8.15). Answer (c) is incorrect because Krueger flaps do not orm a slot, but slats do orm a slot ( Figures 8.11 and 8.14). Q 157

(b) Reerence Figure 8.22. Point ‘A’ to Point ‘B’ illustrates the wording o the question and it can be seen that CL must remain constant i IAS is constant and level flight is to be maintained as flaps are deployed, in order that the Lif remains the same as the Weight. Answers (a), (c) and (d) are incorrect because any change in C L at a constant IAS will change the Lif generated and will not allow the aircraf to maintain level flight. Q 158

(b) Q 159

(b) Lowering ANY amount o flap decreases L/D MAX and reerring to page 372 will remind you that glide angle is a unction ONLY o L/D ratio. Thereore, when trailing edge flaps are deployed glide distance is degraded, making answer (b) correct. With reerence to Figure 8.5 it can be seen that with deployment o trailing edge flaps a lower angle o attack is required or maximum lif (CL MAX), making answer (a) incorrect. Answer (c) is incorrect because flap deployment increases CL MAX, the whole object o deploying flaps – to reduce the take-off and landing distance. Answer (d) is incorrect because increasing CL MAX decreases stall speed. Q 160

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(a) Reer to Figure 8.15. It can be seen that slats increase the stall angle. They do this by increasing the boundary layer kinetic energy - enabling a higher adverse pressure gradient/angle o attack beore reaching C LMAX.

A n s w e r s

Answer (b) is incorrect because flaps decrease the stall angle, Figure. 8.5. Answer (c) is incorrect because spoilers would tend to decrease the stall angle slightly and answer (d) is incorrect because ailerons would not significantly affect stall angle. Q 161

(c) Reerence page 281: When landing, the most critical requirement or sufficient control power in pitch will exist when the CG is at the most orward position, flaps are ully extended, power is set to idle and the aircraf is being flared to land in ground effect. Q 162

(b) “When an aircraf is subject to a positive sideslip angle, lateral stability will be evident i a negative rolling moment coefficient results”. It can be seen rom Figure 10.57 that a positive sideslip angle is aeroplane nose lef – right sideslip. Answer (b) is correct because the tendency o an aeroplane to roll to the lef in a right sideslip is static lateral stability. Answer (a) the tendency o an aeroplane to – “roll to the right in the case o a positive sideslip angle (aeroplane nose to the right)” is incorrect on two counts; a roll to the right in the case o a positive sideslip angle is an example o negative static lateral stability, plus the act that a positive sideslip angle is nose lef. Answers (c) and (d) are both incorrect because static lateral stability is a unction o sideslipping (uncoordinated flight) and turns are coordinated, such that no sideslipping takes place.

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Q 163

(a) For positive static longitudinal stability any change in angle o attack must generate an opposing pitching moment so that the aircraf tends to return towards its trim angle o attack. Answer (a) is the right answer because a nose-down moment when encountering an up gust is an example o positive static longitudinal stability. Answer (b) and (c) are incorrect because no change in angle o attack takes place – which is a stated requirement or static stability. Answer (d) is incorrect because a nose-up pitching moment when encountering an up gust is an example o negative static longitudinal stability. Q 164

(a) Stick orce stability i affected by trim, making (d) an incorrect answer. It can be seen rom Figure 10.31 that the slope o the curve is much steeper at low speed than at high speed, indicating that answer (a) is correct. An increase in speed MUST generate a push orce, making answer (b) incorrect. Answer (c) is incorrect because aeroplane nose-up trim decreases stick orce stability. Q 165

(a) Q 166

(d) Moving the CG af reduces static longitudinal stability and increases manoeuvrability. Increased manoeuvrability gives a smaller control deflection requirement or a given pitch change. Q 167

(b) This question concerns stick orce per ‘g’ – reerence page 275. There must be both an acceptable upper and lower limit to stick orce. The illustrations o Figure 10.36 show the actors which affect the gradient o stick orce per ‘g’ and the text highlights the requirements or any transport aircraf.

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Answer (a) is incorrect – it can be seen rom Figure 10.36 that stick orce gradient decreases with rearward CG position. Answer (c) is incorrect because the stick orce gradient must always be positive. (The term “e-n line” is assumed to mean the stick orce gradient line). Answer (d) is incorrect because there are various methods o modiying stick orces which are not electronic in nature. Q 168

(b) In this context, pitch angle is defined as: “the angle between the longitudinal axis and the horizontal plane”. Pitch angle can also be reerred to as “Body Angle” or as “The Pitch Attitude”. Q 169

(c) Climb gradient is the ratio o vertical height gained to horizontal distance travelled, expressed as a percentage. Trigonometrically, the tangent o the climb angle (gamma) will give climb gradient (tan = opp/adj), where ‘opp’ is the vertical height gained and ‘adj’ is the horizontal distance covered. Unortunately these values are not provided in the question, or indeed in real lie - so other values must be substituted and certain assumptions made in order to “calculate” the answer. Climb angle is the same as the angle between the Weight vector and W cos gamma. The ‘adjacent’ is W cos gamma or Lif and the ‘opposite’ is the backward component o Weight or W sin gamma. From the question Weight (50 000 kg × 10 = 500 000 N), Thrust (60 000 N × 2 = 120 000 N) and Drag (1/12 o Lif) are known or can be estimated. The value o Lif is not given but we do know the Weight, so it has

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Answers

to be assumed that Lif and Weight are equal (at small climb angles [<20 degrees], although we know Lif is in act less than Weight, or practical purposes the difference is insignificant). Thereore, the value o Lif is assumed to be 500 000 N and the Drag to be 500 000 N / 12 = 41 667 N. The ormula or climb gradient is: Percentage Gradient = (T - D / W) × 100. i.e. Thrust minus Drag is the backward component o Weight or ‘opp’ and Weight is the ‘hyp’. For small angles [<20 degrees] o climb or descent the length o the hypotenuse and adjacent are, or all practical purposes, the same; so the sine ormula can be used and will give an answer which is accurate enough. We now have Thrust (120 000 N) minus Drag (41 667 N) divided by Weight (500 000 N) = 0.157 × 100 = 15.7% Climb Gradient. Q 170

(b) Lif is less than Weight in a steady descent. Load actor is Lif divided by Weight, but when the aircraf is in equilibrium in a steady descent the vertical orce opposing the Weight is the Total Reaction. However, CS-25.321 states: “Flight load actors represent the ratio o the aerodynamic orce component (acting normal to the assumed longitudinal axis o the aeroplane) to the weight o the aeroplane”. This clarifies the issue completely; such that in a steady descent Lif is less than Weight and the load actor is less than one. Load actor is useul when considering the loads applied to the aeroplane in flight. While the load actor will not be altered significantly in a steady descent, the concept holds true. Q 171

(d) When considering turning, remind yoursel first o the appropriate ormulae – these help consolidate the variables. 1. L = 1 / cos phi reminds us that the only variable or lif and hence load actor in a turn is bank angle. 2. The next two ormulae must be considered together: 1 7

(a) Radius = V squared / g tan phi

A n s w e r s

(b) Rate = V/ Radius Formula 1 shows that answers (b) and (c) are incorrect because the bank angle or both aircraf is the same. Aircraf ‘A’ is slower than aircraf ‘B’, so Formula 2a shows that answer (a) is incorrect because the turn radius o ‘A’ will in act be smaller than ‘B’. Q 172

(a) VMCL is the minimum IAS at which directional control can be maintained with the aircraf in the landing configuration, BUT with the added ability o being able to roll the aircraf rom an initial condition o steady flight, through an angle o 20 degrees in the direction necessary to initiate a turn away rom the inoperative engine(s), in not more than 5 seconds. V MCL is the “odd one out” among the VMC speeds or this reason. It can clearly be seen that neither statement is correct, making (a) the correct answer. Q 173

(d) Q 174

(a) The speed region between M CRIT and approximately M 1.3 is called “Transonic”. Q 175

(c) The speed range between high and low speed buffet decreases with increasing altitude.

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Q 176

(c) Q 177

(a) Q 178

(a) For SMALL aircraf VA is the speed at the intersection o C LMAX and the positive limit load actor and is dependent upon mass (which will affect the speed at which C LMAX is achieved). As this is the examination or ATPL, LARGE aeroplanes (CS-25) must be considered. V A is defined as: The highest speed at which sudden, ull elevator deflection (nose-up) can be made without exceeding the design limit load actor - making V A slower than the speed intersection o C LMAX and the positive limit load actor. This is due to the effect o the tailplane moving downward when the aircraf is being pitched nose-up increasing the effective angle o attack o the tailplane and increasing the load imposed on the whole. Q 179

(b) Changes in lif orce due to a gust are considered to act through the Aerodynamic Centre. Q 180

(b) The key to answering this question successully is an understanding o what is meant by “......decreasing the propeller pitch.” Decreasing the propeller pitch is reducing the blade angle. This would increase the aircraf’s Parasite area and Total Drag, which would decrease L/D MAX. Because o decreased L/D MAX the aircraf would have an increased rate o descent. Q 181

(d) IAS is a measure o dynamic pressure, whereas TAS is the speed o the aircraf through the air. Changes in TAS are used to compensate or changes in air density to maintain a constant dynamic pressure. The lower the density, the higher the TAS must be to maintain a constant IAS.

7 1

s r e w s n A

Answer (a) is incorrect because decreasing temperature increases air density, which decreases the difference between IAS and TAS. Answer (b) is incorrect because increasing air density decreases the difference between IAS and TAS. Answer (c) is incorrect because density changes with altitude. Q 182

(a)

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Questions Specimen Examination Paper 1.

What is the SI unit which results rom multiplying kg and m/s squared?

a. b. c. d. 2.

TAS is:

a. b. c. d. 3.

Newton. Psi. Joule. Watt.

higher than speed o the undisturbed airstream around the aircraf. lower than speed o the undisturbed airstream around the aircraf. lower than IAS at ISA altitudes below sea level. equal to IAS, multiplied by air density at sea level.

Which o the ollowing statements about a venturi in a subsonic airflow is correct? (i) The dynamic pressure in the undisturbed flow and in the throat are equal. (ii) The total pressure in the undisturbed flow and in the throat are equal.

a. b. c. d. 4.

The angle between the aeroplane longitudinal axis and the chord line is:

a. b. c. d.

1 7

Q u e s t i o n s

5.

increases or decreases depending upon the initial angle o attack. increases. decreases. remains the same.

Which o the ollowing is a characteristic o laminar flow boundary layer?

a. b. c. d.

578

Area o wing divided by the span. The same as the mean chord o a rectangular wing o the same span. The mean chord o the whole aeroplane. The 25% chord o a swept wing.

With flaps deployed, at a constant IAS in straight and level flight, the magnitude o tip vortices:

a. b. c. d. 7.

angle o incidence. glide path angle. angle o attack. climb path angle.

What is the MAC o a wing?

a. b. c. d. 6.

(i) is correct and (ii) is incorrect. (i) is incorrect and (ii) is correct. (i) and (ii) are correct. (i) and (ii) are incorrect.

Constant velocity. Constant temperature. No flow normal to the surace. No vortices.

Questions 8.

Which o the ollowing is the correct ormula or drag?

a. b. c. d. 9.

b. c. d.

increase because increasing weight increases the 1g stall speed. decrease because the 1g stall speed is an IAS. decrease because increasing weight increases the 1g stall speed. remain the same because increased weight increases the IAS that corresponds to a particular angle o attack.

Because VMCA with slats extended is more avourable compared to the flaps extended position. Because flaps extended gives a large decrease in stall speed with relatively less drag. Because slats extended provides a better view rom the cockpit than flaps extended. Because slats extended gives a large decrease in stall speed with relatively less drag.

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s n o i t s e u Q

What must happen to the CL when flaps are deployed while maintaining a constant IAS in straight and level flight?

a. b. c. d. 13.

200 kt. 119 kt. 141 kt. 100 kt.

Afer take-off why are the slats (i installed) always retracted later than the trailing edge flaps?

a.

12.

V squared CL S V (CL) squared S V squared AR CD S V squared CD S

When flying straight and level in 1g flight, slightly below maximum all up weight, a basic stall warning system (flapper switch) ac tivates at 75 kt IAS and the aircraf stalls at 68 kt IAS. Under the same conditions at maximum all up weight the margin between stall warning and stall will:

a. b. c. d. 11.

½ rho ½ rho ½ rho ½ rho

VS is 100 kt at n = 1, what will the stall speed be at n = 2?

a. b. c. d. 10.

17

Increase then decrease. Remain constant. Decrease. Increase.

I an aircraf is longitudinally statically unstable, at the same time it will be:

a. b. c. d.

dynamically unstable. dynamically neutral. dynamically stable. dynamically positively stable.

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Questions 14.

Positive static lateral stability is the tendency o an aeroplane to:

a. b. c. d. 15.

To provide the required manoeuvre stability, an aircraf in straight and level flight (n =1) requires a stick orce o 150 lb/g. I n = 2.5 what is the increase in stick orce required?

a. b. c. d. 16.

18.

Q u e s t i o n s

c. d.

dominant lateral static stability gives an increased tendency or Dutch roll. dominant lateral static stability gives an increased tendency or spiral instability. dominant directional static stability gives an increased tendency or Dutch roll. no effect because they are mutually independent.

Which statement is correct?

a. b. c. d.

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Anhedral. Dihedral. High wing. Large wingspan.

When considering the relationship between lateral static stability and directional stability:

a. b.

19.

Destabilizing dihedral effect. Stabilizing. Negative dihedral effect. No effect.

What type o wing arrangement decreases static lateral stability?

a. b. c. d.

1 7

225 lb. 375 lb. 150 lb. No increase.

What effect does a positive swept wing have on static directional stability?

a. b. c. d. 17.

roll to the right in the case o a positive sideslip angle (aeroplane nose to the lef). roll to the lef in the case o a positive sideslip angle (aeroplane nose to the lef). roll to the lef in a right turn. roll to the right in a right turn.

The stick orce per ‘g’ increases when the CG is moved af. The stick orce per ‘g’ must have both upper and lower limits in order to assure acceptable control characteristics. I the slope o the e-n line becomes negative, generally speaking this is not a problem or control o an aeroplane. The stick orce per ‘g’ can only be corrected by means o electronic devices (stability augmentation) in the case o an unacceptable value.

Questions 20.

At cruising speed an aircraf fitted with spoilers, inboard ailerons and outboard ailerons will use which o the ollowing combinations?

a. b. c. d. 21.

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s n o i t s e u Q

Lef aileron up 5 degrees, right aileron down 2 degrees. Right aileron up 5 degrees, lef aileron down 2 degrees. Lef aileron up 2 degrees, right aileron down 5 degrees. Right aileron up 2 degrees, lef aileron down 5 degrees.

Which statement in respect to trim settings o a stabilizer is correct?

a. b. c. d. 26.

Always on the hinge line, irrespective o the type o aerodynamic balance. On the hinge line i the control surace doe not have an inset hinge. On the hinge line i the control surace has an inset hinge. In ront o the hinge line.

Which o the ollowing is the correct example o differential aileron deflection to initiate a lef turn?

a. b. c. d. 25.

The angle between the chord line and the horizontal plane. The angle between the longitudinal axis and the horizontal plane. The angle between the chord line and the longitudinal axis. The angle between the relative airflow and the longitudinal axis.

What is the location o mass balance weights?

a. b. c. d. 24.

No change. Elevator up, trim tab down. Elevator down, trim tab up. Elevator changes due to horizontal stabilizer changing.

What is pitch angle?

a. b. c. d. 23.

Inboard ailerons and spoilers. Inboard and outboard ailerons. Outboard ailerons only. Spoilers and outboard ailerons.

How does the exterior view o an aircraf change when trim is adjusted to maintain straight and level flight with speed decrease?

a. b. c. d. 22.

17

With a nose heavy aeroplane, the stabilizer leading edge should be higher than or a tail heavy aeroplane. With a nose heavy aeroplane, the stabilizer leading edge should be lower than or a tail heavy aeroplane. With CG on the orward limit, the stabilizer should be ully adjusted nosedown to obtain maximum elevator deflection at rotation during take-off. Since typical take-off speeds are independent o CG position, stabilizer settings are dependent only on flap setting.

Why does a transport aircraf with powered controls use a horizontal stabilizer trim?

a. b. c. d.

Pilot input is not subject to aerodynamic control orces. Trim tabs are not effective enough. Overly complex mechanism. Trim tabs would increased MCRIT.

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Questions 27.

An aircraf o 50 tonnes mass, with two engines each o 60 000 N Thrust and with an L/D ratio o 12:1 is in a straight steady climb. Taking ‘g’ to be 10 m/s/s, wh at is the climb gradient?

a. b. c. d. 28.

I lif in straight and level flight is 50 000 N, the lif o an aircraf in a constant altitude 45 degree bank would increase to?

a. b. c. d. 29.

31.

Q u e s t i o n s

(i) af (i) af (i) wd (i) wd

(ii) increasing. (ii) decreasing. (ii) increasing. (ii) decreasing.

What is the regime o flight rom the critical Mach number (M CRIT) to approximately M 1.3 is called?

a. b. c. d.

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both (i) and (ii) are incorrect. (i) is incorrect and (ii) is correct. (i) is correct and (ii) is incorrect. both (i) and (ii) are correct.

As Mach number increases at transonic speed, tuck under is caused by the CP moving (i) and downwash at the tail (ii):

a. b. c. d. 33.

the turn radius o ‘A’ will be greater than ‘B’. the coefficient o lif o ‘A’ will be less than ‘B’. the load actor o ‘A’ is greater than ‘B’. rate o turn o ‘A’ is greater than ‘B’.

VMCL can be limited by: (i) engine ailure during take-off, (ii) maximum rudder deflection.

a. b. c. d. 32.

lif is less than weight, load actor is equal to one. lif is less than weight, load actor is less than one. lif is equal to weight, load actor is equal to one. lif is equal to weight, load actor is less than one.

Two aircraf o the same weight and under identical atmospheric conditions are flying level 20 degree bank turns. Aircraf ‘A’ is at 130 kt, aircraf ‘B’ is at 200 kt:

a. b. c. d.

1 7

50 000 N 60 000 N 70 000 N 80 000 N

In a straight steady descent:

a. b. c. d. 30.

12%. 24%. 15.7%. 3.7%.

Transonic. Hypersonic. Subsonic. Supersonic.

Questions 34.

The speed range between high and low speed buffet:

a. b. c. d. 35.

7 1

s n o i t s e u Q

VMO. VNE. MMO. VD.

An aircraf in straight and level flight has a C L o 0.42, and a 1 degree increase in angle o attack would increase the C L by 0.1. Following a gust which increases the angle o attack by 3 degrees, what load actor would the aircraf be subject to?

a. b. c. d. 40.

IAS. TAS. Mach number. EAS.

I an aircraf is descending at a constant Mach number, which o the ollowing operational speed limitations may be exceeded?

a. b. c. d. 39.

Increase. Constant. Decrease. Decreases then above a certain Mach number it will increase.

Which o the ollowing is required so the flight crew can determine the effects o compressibility?

a. b. c. d. 38.

Increase. Decrease. Remain the same. Decrease up to a certain Mach number and then increase.

What happens to the Mach number o the airflow as it passes through an expansion wave?

a. b. c. d. 37.

decreases during a descent at a constant Mach number. is always positive at Mach numbers below MMO. increases during a descent at a constant IAS. increases during climb.

What happens to the local speed o sound o air passing through an expansion wave?

a. b. c. d. 36.

17

1·7 0·7 1·4 1·0

Which o the ollowing can affect VA?

a. b. c. d.

Mass and pressure altitude. Mass only. Pressure altitude only. It remains a constant IAS.

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Questions 41.

A single-engine aircraf with a constant speed propeller is in a gliding descent with the engine idling, what would be the effect o increasing the propeller pitch?

a. b. c. d. 42.

A single-engine aircraf with a constant speed propeller is in a gliding descent with the engine idling. What would be the effect o decreasing the propeller pitch?

a. b. c. d. 43.

Q u e s t i o n s

584

higher maximum thrust available. higher maximum efficiency. more blade surace area available. nearly maximum efficiency over wide speed range.

With a clockwise rotating propeller (when viewed rom the rear) at low orward speed, the propeller asymmetric blade effect will cause:

a. b. c. d. 1 7

Increased L/D MAX, increased rate o descent. Decreased L/D MAX, increased rate o descent. Increased L/D MAX, decreased rate o descent. Decreased L/D MAX, decreased rate o descent.

The advantage o a constant speed propeller over a fixed pitch propeller is:

a. b. c. d. 44.

Increased L/D MAX, increased rate o descent. Decreased L/D MAX, increased rate o descent. Increased L/D MAX, decreased rate o descent. Decreased L/D MAX, decreased rate o descent.

roll to the lef. yaw to the lef. roll to the right. yaw to the right.

Questions

17

7 1

s n o i t s e u Q

585

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Answers

Answers to Specimen Exam Paper

1 7

A n s w e r s

586

1 a

2 c

3 b

4 a

5 b

6 c

7 c

8 d

9 c

10 d

11 d

12 b

13 a

14 b

15 a

16 b

17 a

18 a

19 b

20 a

21 b

22 b

23 d

24 a

25 b

26 b

27 c

28 c

29 b

30 d

31 a

32 b

33 a

34 c

35 b

36 a

37 c

38 a

39 a

40 a

41 c

42 b

43 d

44 b

Answers

17

Explanations to Specimen Exam Paper Q1

(a) I a mass is accelerated a orce must have been applied. The kg is the SI unit or mass and m/s squared is the SI unit or acceleration. The applied orce can be determined by multiplying the mass by the acceleration and the answer must use the SI unit or orce - the newton. Q2

(c) True Airspeed (TAS) is the relative velocity between the aircraf and undisturbed air which is close to, but unaffected by the presence o the aircraf. Changing the TAS (the speed o the aircraf through the air; the only speed there is) compensates or changes in air density and ensures a constant mass flow o air over the wing. I an altitude below ISA sea level is considered, the air density would be higher and thereore the TAS would have to be lower than IAS to compensate and keep lif constant. Q3

(b) Bernoulli’s Theorem states: In the steady flow o an “ideal” fluid the sum o the pressure and kinetic energy per unit volume remains constant. Statement (i) is incorrect because the dynamic pressure in the throat o the venturi is higher than the ree stream flow. Statement (ii) is correct. Q4

(a) The angle between the chord line and longitudinal axis is called the angle o incidence which is fixed or a wing, but may be variable or the tailplane (horizontal stabilizer). Q5

(b) A rectangular wing o this chord and the same span would have broadly similar pitching moment characteristics. The MAC is a primary reerence or longitudinal stability considerations.

7 1

Q6

s r e w s n A

(c) Wing tip vortices are strongest with the aircraf in the clean configuration. With flaps down, the flaps generate their own vortices which interere with and weaken the main, tip vortices. Q7

(c) The “key” characteristic o a laminar boundary layer is that there is no flow normal to the surace. Q8

(d) Drag = 1/2 rho × V squared × C D × S. Q9

(c) ‘g’ is the colloquial symbol or load actor. Load actor is the relationship between Lif and Weight. When an aircraf is banked in level flight, Lif must be greater than Weight and the relationship can be calculated by using the ormula: L = 1/cos phi (where phi = bank angle). To calculated the stall speed in a 2g turn, multiply the 1g stall speed by the square root o 2, in this case 1.41. 100 × 1.41 = 141 kt. [It can be said that ‘g’ is the same as 1/cos phi].

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Answers

Q 10

(d) Stalling is caused by airflow separation. The amount o airflow separation is due to the relationship between the adverse pressure gradient and boundary layer kinetic energy. The adverse pressure gradient will increase i angle o attack is increased. A 1g stall occurs at the critical angle o attack (C LMAX). A stall warning must begin with sufficient margin to prevent inadvertent stalling, so a stall warning device must also be sensitive to angle o attack. Thereore, a 1g stall will occur at the critical angle o attack and the stall warning will activate at an angle o attack which is slightly less. In 1g flight each angle o attack requires a particular IAS (dynamic pressure). An increase in weight will not alter the respective angles o attack, but will increase both the IAS at which the stall warning activates and the IAS at which the 1g stall occurs but the margin between them will remain essentially the same. Q 11

(d) Extended slats do not change C L or CD significantly, they do however increase C LMAX and thereore give greater margin to the stall speed. Q 12

(b) Reerence Figure 8.22. Point ‘A’ to Point ‘B’ illustrates the wording o the question and it can be seen that CL must remain constant i IAS is constant and level flight is to be maintained as flaps are deployed, in order that the Lif remains the same as the Weight. Q 13

(a) Negative longitudinal static stability means the aircraf will be urther displaced rom equilibrium ollowing removal o the original disturbing orce. Thereore, over a period o time (Dynamic Stability) it can NEVER be dynamically stable. Q 14

(b) Page 302 states: “When an aircraf is subject to a positive sideslip angle, lateral stability will be evident i a negative rolling moment coefficient results”. It can be seen rom Figure 10.57 that a positive sideslip angle is aeroplane nose lef – right sideslip. Answer (b) is correct because the tendency o an aeroplane to roll to the lef in a right sideslip is static lateral stability.

1 7

A n s w e r s

Q 15

(a) To calculate stick orce per ‘g’ it must be remembered that in straight and level flight the aircraf as at 1g. Thereore the increment is only 1.5g. 150 lb/g × 1.5 = 225 lb. Q 16

Sideslip angle decreases the effective sweep on the wing ‘into wind’ and increases the effective sweep on the wing ‘out o wind’. Decreasing effective sweep angle increases Lif and thereore Induced Drag. This will give a positive contribution to Directional Static Stability - making (b) the only correct answer. Q 17

Dihedral (Geometric) makes a powerul contribution to Lateral static stability. A wing mounted high on the uselage gives a positive contribution the Lateral static stability. Large wingspan makes no contribution to Lateral static stability. A reduction in Dihedral will reduce Lateral static stability - as the definition o Geometric Dihedral is “The upward inclination o the plane o the wing rom the horizontal. I the plane o the wing is angled below the horizontal, this will urther decrease Lateral static stability and is known as Anhedral - making answer (a) correct.

588

Answers

17

Q 18

(a) The relationship between Lateral Static and Directional Static Stability will determine which type o Dynamic instability the aircraf is most likely to exhibit. I Static Lateral Stability is dominant, the extreme increase in lif on the wing into wind will also give a significant increase in Induced Drag. Thus, as the wing ‘into wind’ is accelerating upwards it will also be accelerating rearwards. By the time the aircraf has reached ‘wings level’ the other wing tip will be moving orward about the CG, which will increase its lif and the aircraf will tend to roll back in the opposite direction and this process will continue and maybe diverge - this is Dutch Roll. Because the Lateral Static Stability is much ‘stronger’ than the Directional Static Stability, the fin is not able to prevent the yawing motion. It is the DOMINANCE o Lateral over Directional that determines the likelihood o Dutch Roll thereore, decreased Directional Static Stability AND increased Lateral Static Stability will make Lateral Static Stability dominant and the aircraf susceptible to Dutch Roll. I Static Lateral Stability is dominant, the aircraf will be susceptible to Spiral instability. This is because the fin will give a larger yawing and consequent rolling moment with the aircraf in a sideslip than the Lateral Static Stability is able to counter. Similarly to the case o Dutch Roll - the aircraf can be more susceptible to Spiral Instability due to a decrease in Lateral Static Stability AND and increase in Directional Static Stability - it is the dominance that should be considered. Q 19

(b) This question concerns stick orce per ‘g’ . There must be both an acceptable upper and lower limit to stick orce. The illustrations o Figure 10.36 show the actors which affect the gradient o stick orce per ‘g’ and the text highlights the requirements or any transport aircraf. 7 1

Q 20

At cruise speed the flaps will be up, which de-activates (locks-out) the outboard ailerons. Thereore, the inboard ailerons and the roll spoilers will operate, making (a) the only correct answer.

s r e w s n A

Q 21

There is only one tab that moves in the same direction as the control surace - the antibalance tab, so a good general rule is that all tabs (except one) move in the opposite direction to the control surace. The best approach to questions about controls and/or tabs is to first consider what you want the aeroplane to do. In this case, a speed decrease will generate a nose-down pitching moment. To oppose this, the pilot need to increase back pressure on the pitch control. This moves the elevator up. To hold the elevator in this new position, the trim tab is moved down. Thus (b) is the only correct answer. Q 22

(b) In this context, pitch angle is defined as: “the angle between the longitudinal axis and the horizontal plane”. Pitch angle can also be reerred to as “Body Angle” or as “The Pitch Attitude”.

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Answers

Q 23

Mass balance weights are used to prevent control surace flutter. Flutter is prevented by redistributing the mass o the control surace to move its CG orward onto its hinge line. To accomplish the orward movement o control surace CG a mass balance weight is attached in ront o the hinge line. This makes (d) the only possible answer. Q 24

(a) Differential ailerons are used to decrease adverse aileron yaw. Adverse aileron yaw is the result o increased Induced drag rom the down-going aileron. A mechanism makes the downgoing aileron move through a smaller angle than the up-going aileron. Q 25

A nose heavy aeroplane is one in which a backward orce on the pitch control is required to maintain level flight. To trim-out the backward stick orce a downorce on the tailplane is required. A trimming tailplane must have its incidence decreased (leading edge lowered) to generate the required tail downorce - making (b) the only correct answer. Answer (c) is incorrect because one o the advantages o a trimming tailplane is that the effective pitch control is not influence by the amount o pitch trim used. Answer (d) is incorrect because its statement is complete rubbish. Q 26

Compared to a trim tab, the advantages o using a Variable Incidence Trimming Tailplane are that it is very powerul and gives an increased ability to trim or a larger speed and CG range, it reduces trim drag, and it does not reduce the ‘effective’ range o the pitch control. Answer (a) is incorrect because pilot input moves the elevator, not the trimming tailplane. Answer (c) is incorrect because its relative complexity is a disadvantage. Answer (d) is incorrect because MCRIT is not affected by trim tabs. Answer (b) is the only possible correct answer.

1 7

Q 27

A n s w e r s

(c) Climb gradient is the ratio o vertical height gained to horizontal distance travelled, expressed as a percentage. Trigonometrically, the tangent o the climb angle (gamma) will give climb gradient (tan = opp/adj), where ‘opp’ is the vertical height gained and ‘adj’ is the horizontal distance covered. Unortunately these values are not provided in the question, or indeed in real lie - so other values must be substituted and certain assumptions made in order to “calculate” the answer. Climb angle is the same as the angle between the Weight vector and W cos gamma. The ‘adjacent’ is W cos gamma or Lif and the ‘opposite’ is the backward component o Weight or W sin gamma. From the question Weight (50 000 kg × 10 = 500 000 N), Thrust (60 000 N × 2 = 120 000 N) and Drag (1/12 o Lif) are known or can be estimated. The value o Lif is not given, but we do know the Weight, so it has to be assumed that Lif and Weight are equal (at small climb angles [<20 degrees], although we know Lif is in act less than Weight, or practical purposes the difference is insignificant). Thereore, the value o Lif is assumed to be 500 000 N and the Drag to be 500 000 N / 12 = 41 667 N. The ormula or climb gradient is: Percentage Gradient = (T - D / W) × 100. i.e. Thrust minus Drag is the backward component o Weight or ‘opp’ and Weight is the ‘hyp’. For small angles [<20 degrees] o climb or descent the length o the hypotenuse and adjacent are, or all practical purposes, the same; so the sine ormula can be used and will give an answer which is

590

Answers

17

accurate enough. We now have Thrust (120 000 N) minus Drag (41 667 N) divided by Weight (500 000 N) = 0.157 × 100 = 15.7% Climb Gradient. Q 28

In level flight, Lif is a unction o the bank angle. The ormula is L = 1 / cos phi In a 45 degree bank the lif is increased by 1.41 (41%) 50 000 N × 1.41 = 70 500 N, making (c) the correct answer. Q 29

(b) Lif is less than Weight in a steady descent. Load actor is Lif divided by Weight, but when the aircraf is in equilibrium in a steady descent the vertical orce opposing the Weight is the Total Reaction. However, CS-25.321 states: “Flight load actors represent the ratio o the aerodynamic orce component (acting normal to the assumed longitudinal axis o the aeroplane) to the weight o the aeroplane”. This clarifies the issue completely; such that in a steady descent Lif is less than Weight and the load actor is less than one. Load actor is useul when considering the loads applied to the aeroplane in flight. While the load actor will not be altered significantly in a steady descent, the concept holds true. Q 30

(d) When considering turning, remind yoursel first o the appropriate ormulae – these help consolidate the variables. 1. L = 1 / cos phi reminds us that the only variable or lif and hence load actor in a turn is bank angle. 2. The next two ormulae must be considered together: (a) Radius = V squared / g tan phi (b) Rate = V/ Radius

7 1

s r e w s n A

Q 31

(a) VMCL is the minimum IAS at which directional control can be maintained with the aircraf in the landing configuration, BUT with the added ability o being able to roll the aircraf rom an initial condition o steady flight, through an angle o 20 degrees in the direction necessary to initiate a turn away rom the inoperative engine(s), in not more than 5 seconds. V MCL is the “odd one out” among the V MC speeds or this reason. It can clearly be seen that neither statement is correct, making (a) the correct answer. Q 32

(b) The initial ormation o a shock wave is on the top surace o the wing at the point o maximum local velocity - this is usually the thickest part o the wing, at the wing root. Shock waves cause a localized reduction in C L. On a swept wing this gives a reduction in Lif orward o the CG and the CP will move af. In addition, shock wave ormation at the wing root reduces downwash at the tailplane. These actors together cause Mach Tuck, Tuck Under or High Speed Tuck (three names or the same phenomena). Q 33

(a) Reerence to Figure 13.2 will show that the speed region between M CRIT and approximately M 1.3 is called “Transonic”. Q 34

(c) It can be seen that the speed range between high and low speed buffet decreases with increasing altitude.

591

17

Answers

Q 35

(b) Speed o sound is proportional to temperature. The temperature decreases as it passes through an expansion wave, thereore the local speed o sound decreases. Q 36

(a) Mach number is proportional to TAS and inversely proportional to Local Speed o Sound. Through an expansion wave, velocity increases and temperature decreases. Thereore the Mach number o the airflow will increase. Q 37

(c) Compressibility, in this context, is the general term which reers to the effects on the aircraf when flying aster than approximately Mach 0.4. To determine the effects o compressibility the flight crew need to know the aircraf Mach number. Q 38

(a) Q 39

(a) Q 40

(a) For SMALL aircraf VA is the speed at the intersection o C LMAX and the positive limit load actor and is dependant upon mass (which will affect the speed at which C LMAX is achieved). As this is the examination or ATPL, LARGE aeroplanes (CS-25) must be considered. VA is defined as: The highest speed at which sudden, ull elevator deflection (nose-up) can be made without exceeding the design limit load actor - making V A slower than the speed intersection o C LMAX and the positive limit load actor. This is due to the effect o the tailplane moving downward when the aircraf is being pitched nose-up increasing the effective angle o attack o the tailplane and increasing the load imposed on the whole aircraf. This is aerodynamic damping, which is a unction o the tailplane vertical velocity and TAS. Thereore VA varies with both aircraf mass and pressure altitude.

1 7

A n s w e r s

Q 41

(c) Increasing the propeller pitch, by pulling the propeller RPM control lever backwards to “Decrease RPM” will drive the blades towards the coarse pitch stop. This decreases the Parasite drag o the aeroplane, thus increasing L/D MAX and allowing a decreased rate o descent. Q 42

(b) The key to answering this question successully is an understanding o what is meant by “...... decreasing the propeller pitch.” Decreasing the propeller pitch is reducing the blade angle. This would increase the aircraf’s Parasite area and Total Drag, which would decrease L/D MAX. Because o decreased L/D MAX the aircraf would have an increased rate o descent. Q 43

(d) IAS is a measure o dynamic pressure, whereas TAS is the speed o the aircraf through the air. Changes in TAS are used to compensate or changes in air density to maintain a constant dynamic pressure. The lower the density, the higher the TAS must be to maintain a constant IAS. Q 44

(b) Asymmetric blade effect gives more thrust on the side with the down-going blade with a clockwise rotating propeller, this gives a lef turning moment.

592

Chapter

18 Index

593

18

1 8

I n d e x

594

Index

Index Buffet Margin . . . . . . . . . . . . . . . . . . . . . . Buffet Onset Chart . . . . . . . . . . . . . . . . . .

Symbols 1g Stall Speed . . . . . . . . . . . . . . . . . . . . . .

168

A Accelerated Stall . . . . . . . . . . . . . . . . . . . . 187 Acceleration . . . . . . . . . . . . . . . . . . . . . . . . . 9 Adverse Aileron Yaw . . . . . . . . . . . . . . . . 342 Adverse Pressure Gradient . . . . . . . . . . . 113 Aerodynamic Balance . . . . . . . . . . . . . . . 335 Aerodynamic Centre (AC) . . . . . . . . . . . . . 60 Aerodynamic Force Coefficient . . . . . . . . . 58, 71

Aerodynamic Heating . . . . . . . . . . . . . . . 442 Aerodynamic Twisting Moment (ATM) . 507 Aeroelastic Coupling . . . . . . . . . . . . . . . . 474 Aeroelasticity . . . . . . . . . . . . . . . . . . . . . . 474 Aerooil . . . . . . . . . . . . . . . . . . . . . . . . . . . . 52 Aeroplane Total Drag . . . . . . . . . . . . . . . 124 Aileron Reversal . . . . . . . . . . . . . . . . . . . . 480 Aileron-rudder Coupling . . . . . . . . . . . . . 343 Air Density . . . . . . . . . . . . . . . . . . . . . . . . . 26 Airspeed . . . . . . . . . . . . . . . . . . . . . . . . . . . 32 Angle o Attack . . . . . . . . . . . . . . . . . . . . . 53 Angle o Attack Probe . . . . . . . . . . . . . . . 153 Angle o Attack Vane. . . . . . . . . . . . . . . . 153 Angle o Descent in the Glide . . . . . . . . . 372 Angle o Incidence . . . . . . . . . . . . . . . . . . . 52 Anti-balance Tab . . . . . . . . . . . . . . . . . . . 337 Area Rule . . . . . . . . . . . . . . . . . . . . . . . . . 438 Area (Zone) o Influence . . . . . . . . . . . . . 444 Artificial Feel (‘Q’ Feel) . . . . . . . . . . . . . . 339 Artificial Feel Trim . . . . . . . . . . . . . . . . . . 354 Aspect Ratio . . . . . . . . . . . . . . . . . . . . . . . . 85, 118

Asymmetric Blade Effect . . . . . . . . . . . . . Asymmetric Thrust . . . . . . . . . . . . . . . . . .

517 349, 384

Automatic Slots . . . . . . . . . . . . . . . . . . . . 220 Average Chord . . . . . . . . . . . . . . . . . . . . . . 85

B Backlash . . . . . . . . . . . . . . . . . . . . . . . . . . 474 Balance o Forces . . . . . . . . . . . . . . . . . . . 367 Balance Tab . . . . . . . . . . . . . . . . . . . . . . . 336 Basic Lif Equation . . . . . . . . . . . . . . . . . . . 72 Bernoulli’s Theorem . . . . . . . . . . . . . . . . . . 44 Blade Angle or Pitch . . . . . . . . . . . . . . . . 503 Blade Twist . . . . . . . . . . . . . . . . . . . . . . . . 504 Boundary Layer Fences . . . . . . . . . . . . . . 161 Boundary Layer Separation . . . . . . . . . . . 423 Bow Wave . . . . . . . . . . . . . . . . . . . . . . . . . 444 Buffet . . . . . . . . . . . . . . . . . . . . . . . . . . . . 427

18

432 432

C Calibrated Airspeed . . . . . . . . . . . . . . . . . . 31 Canards . . . . . . . . . . . . . . . . . . . . . . . . . . . 182 Centre o Gravity (CG) . . . . . . . . . . . . . . . . . 7 Centre o Pressure (CP) . . . . . . . . . . . . . . . 53 Centriugal Twisting Moment (CTM). . . 507 Chord . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 52 Chord Line . . . . . . . . . . . . . . . . . . . . . . . . . 52 Chord Ratio. . . . . . . . . . . . . . . . . . . . . . . . . 53 Climb Angle . . . . . . . . . . . . . . . . . . . . . . . 369 CLMAX Augmentation . . . . . . . . . . . . . . . . 209 CLMAX Boundary. . . . . . . . . . . . . . . . . . . . . 461 Compressibility Error . . . . . . . . . . . . . . . . . 32 Constant Speed Propeller . . . . . . . . . . . . 509 Control Balancing . . . . . . . . . . . . . . . . . . . 335 Control Buzz . . . . . . . . . . . . . . . . . . . . . . . 427 Critical Angle o Attack . . . . . . . . . . . . . . 145 Critical Engine. . . . . . . . . . . . . . . . . . . . . . 386 Crossed-control Stall . . . . . . . . . . . . . . . . . 187

D Deep Stall . . . . . . . . . . . . . . . . . . . . . . . . . 166 Density Altitude . . . . . . . . . . . . . . . . . . . . . 79 Design Cruising Speed V C. . . . . . . . . . . . . 464 Design Dive Speed VD . . . . . . . . . . . . . . . 464 Design Manoeuvring Speed, V A . . . . . . . 462 Differential Ailerons . . . . . . . . . . . . . . . . . 343 Dihedral Effect . . . . . . . . . . . . . . . . . . . . . 304 Divergence . . . . . . . . . . . . . . . . . . . . . . . . 474 Dorsal Fin . . . . . . . . . . . . . . . . . . . . . . . . . 294 Downwash . . . . . . . . . . . . . . . . . . . . . . . . 263 Drag . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 53 Dutch Roll . . . . . . . . . . . . . . . . . . . . . . . . . 309 Dynamic Pressure . . . . . . . . . . . . . . . . . . . . 27,

8 1

x e d n I

71

Dynamic Stability . . . . . . . . . . . . . . . . . . .

282

E Effective Airflow . . . . . . . . . . . . . . . . . . . . 95 Effective Angle o Attack. . . . . . . . . . . . . . 53, 95

Effective Pitch . . . . . . . . . . . . . . . . . . . . . . 504 effective range . . . . . . . . . . . . . . . . . . . . . 352 Effect o Weight . . . . . . . . . . . . . . . . . . . . 373 Effect o Wind . . . . . . . . . . . . . . . . . . . . . 374 Elasticity . . . . . . . . . . . . . . . . . . . . . . . . . . 474 Emergency Descent . . . . . . . . . . . . . . . . . 371 Energy . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8 Energy Drag . . . . . . . . . . . . . . . . . . . . . . . 423

595

18

Index Equilibrium . . . . . . . . . . . . . . . . . . . . . . . . Equivalent Airspeed: (EAS) . . . . . . . . . . . . Expansion Waves . . . . . . . . . . . . . . . . . . .

365 31

Force . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7 Form (Pressure) Drag . . . . . . . . . . . . . . . . 113 Fowler Flap . . . . . . . . . . . . . . . . . . . . . . . . 211 Free Stream Mach number . . . . . . . . . . . 411 Frise ailerons . . . . . . . . . . . . . . . . . . . . . . . 343 Frost . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 179

Laminar . . . . . . . . . . . . . . . . . . . . . . . . . . . 114 Landing Speeds . . . . . . . . . . . . . . . . . . . . 209 Large Aircraf . . . . . . . . . . . . . . . . . . . . . . 187 Leading Edge Flaps . . . . . . . . . . . . . . . . . 216 Leading Edge Radius . . . . . . . . . . . . . . . . . 53 Leading Edge Slat. . . . . . . . . . . . . . . . . . . 218 Leading Edge Slots . . . . . . . . . . . . . . . . . . 218 Lif . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 53 Lif Curve . . . . . . . . . . . . . . . . . . . . . . . . . . 76 Lif / Drag Ratio . . . . . . . . . . . . . . . . . . . . 214 Lif/Drag Ratio . . . . . . . . . . . . . . . . . . . . . . 80 Load Factor . . . . . . . . . . . . . . . . . . . . . . . . 460 Load Factor in the Turn . . . . . . . . . . . . . . 380 Local Mach Number . . . . . . . . . . . . . . . . . 411 Longitudinal Control . . . . . . . . . . . . . . . . 279 Longitudinal Dihedral . . . . . . . . . . . . . . . 262 Long Period Oscillation . . . . . . . . . . . . . . 287

G

M

Geometric Pitch . . . . . . . . . . . . . . . . . . . . 503 Geometric twist . . . . . . . . . . . . . . . . . . . . 157 Glide . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 372 ‘g’ Limit on Turning . . . . . . . . . . . . . . . . . 381 Ground Effect . . . . . . . . . . . . . . . . . . . . . . . 91 Gust Loads . . . . . . . . . . . . . . . . . . . . . . . . 467 Gyroscopic Effect . . . . . . . . . . . . . . . . . . . 515

Mach Angle . . . . . . . . . . . . . . . . . . . . . . . 443 Mach Cone . . . . . . . . . . . . . . . . . . . . . . . . 444 Mach Number . . . . . . . . . . . . . . . . . . . . . 408 Mach Number (M) . . . . . . . . . . . . . . . . . . . 31 Mach Trim . . . . . . . . . . . . . . . . . . . . . . . . . 311 Manoeuvre Boundaries . . . . . . . . . . . . . . 465 Manoeuvre Envelope . . . . . . . . . . . . . . . . 461 Manoeuvre Point . . . . . . . . . . . . . . . . . . . 275 Manoeuvre Stability . . . . . . . . . . . . . . . . . 274 Mass . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7 Mass Balance . . . . . . . . . . . . . . . . . . . . . . 340 Mass Distribution . . . . . . . . . . . . . . . . . . . 474 Maximum Camber . . . . . . . . . . . . . . . . . . . 53 Maximum Operating Speed . . . . . . . . . . 466 Maximum Structural Cruise Speed . . . . . 466 Mean Aerodynamic Chord (MAC) . . . . . . 85 Mean Line or Camber Line . . . . . . . . . . . . 53 Microburst . . . . . . . . . . . . . . . . . . . . . . . . 489 Minimum Control Airspeed . . . . . . . . . . . 391 Minimum Turn Radius . . . . . . . . . . . . . . . 381 MMO . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 466 Momentum . . . . . . . . . . . . . . . . . . . . . . . . . 9 Multiple Slotted Flaps . . . . . . . . . . . . . . . 210

445

F Fin Stall . . . . . . . . . . . . . . . . . . . . . . . . . . . Fixed Tabs . . . . . . . . . . . . . . . . . . . . . . . . . Flaperons . . . . . . . . . . . . . . . . . . . . . . . . . Flapper Switch . . . . . . . . . . . . . . . . . . . . . Flaps . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Flexural Aileron Flutter . . . . . . . . . . . . . . Flutter . . . . . . . . . . . . . . . . . . . . . . . . . . . .

348 352 344 152 209 478 474, 477

H Helix Angle . . . . . . . . . . . . . . . . . . . . . . . . 504 High Lif Devices . . . . . . . . . . . . . . . . . . . . . 82 High Mach Numbers . . . . . . . . . . . . . . . . 311 High Speed Buffet . . . . . . . . . . . . . . . . . . 189 Hinge Moments . . . . . . . . . . . . . . . . . . . . 334 Horn Balance . . . . . . . . . . . . . . . . . . . . . . 335

1 8

I n d e x

I Ice . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 179 Indicated Airspeed . . . . . . . . . . . . . . . . . . . 31 Induced Downwash . . . . . . . . . . . . . . . . . . 86, 116

Induced Drag . . . . . . . . . . . . . . . . . . . . . . 116 Induced Drag Coefficient (C Di ) . . . . . . . . 121 Inertia . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8 Instrument Error. . . . . . . . . . . . . . . . . . . . . 32 Intererence Drag . . . . . . . . . . . . . . . . . . . 115 Internal Balance . . . . . . . . . . . . . . . . . . . . 336 International Standard Atmosphere (ISA) 26

K Kinetic Energy . . . . . . . . . . . . . . . . . . . . . . . 8 Krueger Flap . . . . . . . . . . . . . . . . . . . . . . . 216

596

L

N Negative Load Factors . . . . . . . . . . . . . . . Negative Stall . . . . . . . . . . . . . . . . . . . . . . Negative Static Stability . . . . . . . . . . . . . . Neutral Dynamic Stability . . . . . . . . . . . . Neutral Point . . . . . . . . . . . . . . . . . . . . . .

464 465 241 283 249

Index Neutral Static Stability . . . . . . . . . . . . . . .

241, 283

Never Exceed Speed. . . . . . . . . . . . . . . . . 466 Newton’s First Law o Motion. . . . . . . . . . . 8 Newton’s Second Law o Motion . . . . . . . . 8 Newton’s Third Law . . . . . . . . . . . . . . . . . . . 9 Normal Shock Waves . . . . . . . . . . . . . . . . 414 Nose Wheel Steering . . . . . . . . . . . . . . . . 394

O Oblique Shock Waves. . . . . . . . . . . . . . . . Operational Rough-air Speed (VRA / MRA) Operational Speed Limits . . . . . . . . . . . .

419 470 466

P Parasite Drag . . . . . . . . . . . . . . . . . . . . . . 112 Parasite Drag Formula . . . . . . . . . . . . . . . 115 Phugoid. . . . . . . . . . . . . . . . . . . . . . . . . . . 287 Pilot Induced Oscillation . . . . . . . . . . . . . 285 Pilot Induced Oscillation (PIO) . . . . . . . . 310 Pitching Moment . . . . . . . . . . . . . . . . . . . 214 Plain Flap. . . . . . . . . . . . . . . . . . . . . . . . . . 209 Position Error (Pressure Error) . . . . . . . . . . 32 Positive Static Stability . . . . . . . . . . . . . . . 241 Power . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8 Power Absorption . . . . . . . . . . . . . . . . . . 513 Power Assisted Controls . . . . . . . . . . . . . 338 Power-off Stability . . . . . . . . . . . . . . . . . . 264 Power-on Descent . . . . . . . . . . . . . . . . . . 370 Power Required . . . . . . . . . . . . . . . . . . . . 130 Pressure Distribution . . . . . . . . . . . . . . . . . 71 Principle o Continuity . . . . . . . . . . . . . . . . 43 Profile Drag. . . . . . . . . . . . . . . . . . . . . . . . 114 Propeller Efficiency. . . . . . . . . . . . . . . . . . 508 Propeller Slip . . . . . . . . . . . . . . . . . . . . . . . 504

Q Q - Feel . . . . . . . . . . . . . . . . . . . . . . . . . . .

277

R Radius and Rate o Turn . . . . . . . . . . . . . 378 Rate Control . . . . . . . . . . . . . . . . . . . . . . . 342 Rate o Descent in the Glide . . . . . . . . . . 374 Rectangular Wing . . . . . . . . . . . . . . . . . . 156 Relative Airflow . . . . . . . . . . . . . . . . . . . . . 53 Roll Control Spoilers . . . . . . . . . . . . . . . . . 343, 345

Root Chord . . . . . . . . . . . . . . . . . . . . . . . . . 85 Rudder Arm . . . . . . . . . . . . . . . . . . . . . . . 394 Rudder Ratio Changer . . . . . . . . . . . . . . . 349

S Saw Tooth . . . . . . . . . . . . . . . . . . . . . . . . .

161

18

Secondary Effects o Controls . . . . . . . . . 350 Secondary Stall . . . . . . . . . . . . . . . . . . . . . 187 Servo Tab . . . . . . . . . . . . . . . . . . . . . . . . . 337 Shock Stall . . . . . . . . . . . . . . . . . . . . . . . . . 189 Short Period Oscillation . . . . . . . . . . . . . . 288 Single-engine Angle o Climb . . . . . . . . . 396 Single-engine Rate o Climb . . . . . . . . . . 396 Skin Friction Drag . . . . . . . . . . . . . . . . . . . 112 Slots . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 157 Slotted Flap . . . . . . . . . . . . . . . . . . . . . . . . 210 Small Aircraf . . . . . . . . . . . . . . . . . . . . . . 188 Snow . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 179 Solidity . . . . . . . . . . . . . . . . . . . . . . . . . . . 513 Sonic Bang . . . . . . . . . . . . . . . . . . . . . . . . 447 Speed Brakes . . . . . . . . . . . . . . . . . . . . . . 346 Speed o Sound . . . . . . . . . . . . . . . . . . . . . . 31, 407

Spinning . . . . . . . . . . . . . . . . . . . . . . . . . . 183 Spin Recovery . . . . . . . . . . . . . . . . . . . . . . 185 Spiral Divergence . . . . . . . . . . . . . . . . . . . 309 Spiral Slipstream Effect . . . . . . . . . . . . . . 516 Split Flap . . . . . . . . . . . . . . . . . . . . . . . . . . 210 Spring Bias . . . . . . . . . . . . . . . . . . . . . . . . 354 Stabilizer Stall . . . . . . . . . . . . . . . . . . . . . . 182 Stalling Angle . . . . . . . . . . . . . . . . . . . . . . 145 Stall Limit on Turning . . . . . . . . . . . . . . . . 381 Stall Recognition . . . . . . . . . . . . . . . . . . . 148 Stall Recovery . . . . . . . . . . . . . . . . . . . . . . 146 Stall Speed . . . . . . . . . . . . . . . . . . . . . . . . 148 Stall Strips . . . . . . . . . . . . . . . . . . . . . . . . . 158 Stall Warning . . . . . . . . . . . . . . . . . . . . . . 150 Stall Warning Devices. . . . . . . . . . . . . . . . 151 Static Margin . . . . . . . . . . . . . . . . . . . . . . 250 Static Pressure. . . . . . . . . . . . . . . . . . . . . . . 25 Static Stability . . . . . . . . . . . . . . . . . . . . . . 241 Stick Position Stability . . . . . . . . . . . . . . . 271 Stick Pusher. . . . . . . . . . . . . . . . . . . . . . . . 167 Stick Shaker. . . . . . . . . . . . . . . . . . . . . . . . 151 Straight Steady Climb . . . . . . . . . . . . . . . . 368 Streamlines . . . . . . . . . . . . . . . . . . . . . . . . . 45 Streamlining . . . . . . . . . . . . . . . . . . . . . . . 114 Streamtube . . . . . . . . . . . . . . . . . . . . . . . . . 45 Supercritical Aerooil . . . . . . . . . . . . . . . . 440 Super Stall . . . . . . . . . . . . . . . . . . . . . . . . . 166 Super Stall Prevention . . . . . . . . . . . . . . . 167 Surace Area . . . . . . . . . . . . . . . . . . . . . . . . 71 Sweep Angle. . . . . . . . . . . . . . . . . . . . . . . . 85 Sweepback . . . . . . . . . . . . . . . . . . . . . . . . 160

8 1

x e d n I

T Tapered Wing . . . . . . . . . . . . . . . . . . . . . . 157 Taper Ratio . . . . . . . . . . . . . . . . . . . . . . . . . 85

597

CAE-Oxford-Aviation-Academy-080-Principles-of-Flight-ATPL-Ground-Training-Series-2014.pdf

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