Cpl Atg

CPL AIRCRAFT TECHNICAL & GENERAL

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INDEX CPL ATG 1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 11. 12. 13. 14. 15. 16. 17. 18. 19. 20. 21.

Distribution of the 4 forces Straight & Level Stability Climbing, Descending & Turning Stalling Lift augmentation Flying Controls Multi engine flight High Speed Flight Wing & Fuselage construction Hydraulics Landing gear Ice removal systems Pressurisation & Oxygen systems Air conditioning systems Fuel system Electrics Fire warning systems The piston engine Propellors Basic Gas Turbines

Annex A Sample Exams Annex B Answers to questions

01 05 35 51 69 87 93 111 117 125 131 145 163 171 179 185 195 275 281 319 329 353 365

Copyright  2001 Flight Training College of Africa All Rights Reserved. No part of this manual may be reproduced in any manner whatsoever including electronic, photographic, photocopying, facsimile, or stored in a retrieval system, without the prior permission of Flight Training College of Africa.

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CHAPTER 1 DISTRIBUTION OF THE FORCES THE FORCES ACTING ON AN AIRCRAFT IN LEVEL FLIGHT With reference to the diagram, the following forces act on an aircraft in level flight.

WEIGHT Weight is the effect of gravity drawing the aircraft vertically downwards towards the centre of the earth. Weight acts through the Centre of Gravity (C of G) of the aircraft.

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LIFT Lift is the upward force produced by the wings of the aircraft. Lift acts through the Centre of Pressure (C of P) of the aircraft and should be equal in magnitude, but opposite in direction to weight. The two factors which are most commonly variable are angle of attack and speed. If the speed of the aircraft is increased then the angle of attack is decreased and if the speed of the aircraft is decreased, the angle of attack will have to be increased. This will achieve a constant value of lift to balance weight. THRUST Thrust is the force produced by the engine. It acts along the thrust line and propels the aircraft through the air.

DRAG With reference to the Drag chapter, the drag experienced by an aircraft is the sum of: Induced Drag Induced drag is inversely proportional to speed and directly proportional to the angle of attack. The higher the speed, the lower the angle of attack and the less the downwash and thus induced drag. Profile Drag Profile drag is the result of the motion of the aircraft through the air. The higher the speed of the aircraft, the higher the profile drag. Total Drag Total drag is the sum of induced drag and profile drag.

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THE BALANCE OF FORCES IN LEVEL FLIGHT It can be said that in Straight and Level flight (unaccelerated) : Lift = Weight & Thrust = Drag Lift acts through the centre of pressure and is equal in magnitude, but opposite in direction to weight, which acts through the centre of gravity. Two opposite forces acting about a common arm create a couple. The lift/weight couple imparts a nose down pitching moment to the aircraft. Thrust acts through the thrust line and is equal in magnitude, but opposite in direction, to drag which acts through the drag line. The thrust/drag couple imparts a nose up pitching moment to the aircraft.

THE EFFECT OF THE TAILPLANE For level flight to be achieved, not only must lift equal weight and thrust equal drag, but the sum of the pitching moments caused by the lift/weight couple and the thrust/drag couple must be zero. On most aircraft, the magnitude of the lift/weight couple is far greater than that of the thrust/drag couple, leaving the aircraft with a residual nose down pitching moment. To solve this, the a tailplane is inclined at an angle of incidence to carry a down load at cruise speed. This download on the tailplane causes a nose up pitching moment of the thrust/drag couple and together balance the nose down pitching moment of the lift/weight couple.

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CHAPTER 2 STRAIGHT & LEVEL FLIGHT Lets look at the major forces affecting straight and level flight: LIFT When considering how the local airflow is affected by the passage of a body through the air, the following factors are of consequence: i) ii) iii) iv) v) vi)

The The The The The The

shape of the body. attitude (profile) of the body. size of the body. density of the air. viscosity of the air. speed of the air/body.

Factors (iii) to (vi) are collectively known as Reynolds Number. AIRFLOW In the study of the properties of the local airflow as it passes over a body, the classic example used is the simple Venturi. The Venturi is a tube which constricts, or narrows, at a point and then opens up to its original width again. The Equation of Continuity and Bernoulli's Theory accurately describe the changing properties of the air as it flows through the Venturi. THE EQUATION OF CONTINUITY The Equation of Continuity states that air mass flow is constant. This statement is represented by the following formula: A x V x 

=

Constant

Where: A V 

= = =

Area Velocity Density

In subsonic aerodynamics, air is considered to be incompressible, therefore density may be removed from the formula, which now reads: Ax V

=

Constant

The significance of this equation is as follows: As air flows through a Venturi, the cross-sectional area of the tube decrease and in order to maintain a constant, the velocity of the airflow must increase. Once through the throat of the Venturi, the velocity of the airflow must decrease as it encounters a greater area, in order to maintain a constant. BERNOULLI'S THEORY CPL ATG CPL DOC 01 Revision : 1/1/2001

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Bernoulli's Theory states that in the streamline flow of an ideal fluid (one which is not viscous), the sum of all energies will remain constant. A particle of air in a streamline flow through a Venturi possesses the following energies: Kinetic, or Dynamic pressure energy = ½pV²

Static pressure energy = P

Because the particle of air undergoes no appreciable change of height as it passes through a Venturi, Potential energy is not included in the formula. Bernoulli's Theory is thus represented by the following formula: ½pV² + P = constant The significance of Bernoulli's Theory is as follows: As air passes through a Venturi, the area decreases and the velocity of the airflow increases (Equation of continuity). If the velocity of the airflow increases, the Dynamic pressure energy increases, and in order to maintain a constant energy level, the Static pressure energy must increase (Bernoulli's Theory). The value of the Venturi in these principles becomes apparent when related to an aircraft in flight. The bottom half of the Venturi corresponds to the top of an aircraft's wing, horizontal stabiliser or vertical stabiliser. The top half of the Venturi corresponds to the free stream flow of air (air unaffected by the passage of the aircraft). Relating an aircraft wing to the Equation of Continuity and Bernoulli's Theory, as air flows over the top surface of the wing, the velocity of the air increases and the static pressure decreases. If the bottom half of the wing is considered to be flat, there will be no appreciable change in either the velocity or static pressure of the air. There now exists a pressure differential, with normal pressure below the wing and relatively lower pressure above the wing. A body will always move from an area of high pressure to an area of low pressure and this force is what keeps an aircraft airborne. It is known as Lift and will now be discussed in detail . CPL ATG CPL DOC 01 Revision : 1/1/2001

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AEROFOIL TERMINOLOGY AND DEFINITIONS

CHORD LINE The straight line joining the centres of curvature of the leading edge and the trailing edge of an aero-foil. CHORD LENGTH The distance between the leading edge of an aerofoil, measured along the chord line. THE POINT OF MAXIMUM THICKNESS That point where the upper and lower surfaces of an aerofoil section are furthest apart. THE THICKNESS/CHORD RATIO The ratio of the maximum thickness to the chord length of an aerofoil, usually expressed as a percentage. FINENESS RATIO

=

LENGTH MAXIMUM THICKNESS

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RELATIVE AIRFLOW (ALSO KNOWN AS FREE STREAM FLOW) The airflow that is parallel and opposite to the flight path of the aircraft. This airflow must be unaffected by the passage of the aircraft. ANGLE OF ATTACK The angle between the Chord Line and the Relative Airflow. ATTITUDE Body angle as viewed from inside the cockpit i.e. where the neutral horizon cuts the canopy window or glare shield. ANGLE OF INCIDENCE The angle formed between the Chord line and the longitudinal axis of the aircraft. An aircraft has positive incidence when the leading edge is above the longitudinal axis. THE MEAN CAMBER LINE The curved line joining the leading and trailing edges of an aerofoil, which is equidistant from the upper and lower surfaces. CAMBER A term used to describe the measure of curvature of an aerofoil. An aerofoil has positive camber when the Mean Camber Line lies above the Chord Line.

ASPECT RATIO The ratio of the span of a wing to its chord, expressed with the formula: Aspect ratio

=

Span ² Chord

WHERE : SPAN = CHORD =

distance from wing tip to wing tip

Distance between leading edge and trailing Edge for a rectangular wing. If wing is Tapered, then an average chord length must be obtained.

TOTAL REACTION That net force representing the sum of all the forces of lift and drag acting on an aircraft in flight. THE CENTRE OF PRESSURE CPL ATG CPL DOC 01 Revision : 1/1/2001

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That point on the aerofoil through which the total reaction acts. DRAG That component of the Total Reaction which acts parallel to the Relative Airflow. LIFT That component of the Total Reaction which acts at 90 º to the Relative Airflow. STREAMLINE FLOW If succeeding molecules follow the same steady path in a flow, then this path can be represented by a streamline. There will be no flow across the streamlines, only along them.

At any fixed point on the streamline, each air molecule will experience the same velocity and static pressure as the proceeding molecules when they passed that particular point. These values of velocity and pressure may change from point to point along the streamline. A reduction in the velocity of streamline flow is indicated by wider spacing on the streamlines whilst increased velocity is indicated by decreasing spacing of the streamlines. Any molecule following a streamline will experience the same velocity and pressures as the proceeding molecules. Streamline flow around an aircraft is very desirable.

TURBULENT FLOW In turbulent flow, succeeding molecules do not follow a streamlined flow pattern. Succeeding molecules may travel a path quite different to the proceeding molecules. This turbulent flow is also known as unsteady flow or eddying and is an undesirable feature in most phases of flight.

Steady streamline flow is desirable in most phases of flight, and turbulent flow is best avoided.

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HOW LIFT IS GENERATED If the aerofoil is seen to represent the lower half of a Venturi tube, then the following deduction can be made:

As air flows over the aerofoil, it encounters a reduced area and the velocity of the airflow is increased (the Equation of Continuity).

When the velocity of the airflow is increased, the static pressure reduces (Bernoulli's Theory). As air flows under the aerofoil, the velocity and therefore the static pressure remains unchanged, provided the lower half of the aerofoil is flat and not inclined in the airflow.

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This differential in the static pressures: normal static pressure below the aerofoil and lower static pressure above the aerofoil, gives rise to a lifting force called the total reaction. Where the undersurface of the aerofoil is concave in shape, the airflow under the aerofoil encounters an enlarged area, thus the velocity of the airflow is reduced, (the Equation of Continuity). When the velocity of the airflow is increased, the static pressure is decreased, (Bernoulli's Theory). The pressure differential between the upper and lower surfaces of the aerofoil is now greater and the magnitude of the total reaction is increased.

Where the aerofoil is inclined in the airflow, (greater angle of attack), the area through which the airflow must pass above the wing is effectively reduced even further, resulting in a greater acceleration of airflow and a further drop in static pressure. Below the aerofoil, the airflow strikes the underside of the inclined aerofoil given rise to a greater increase in static pressure due to ram effect. This increase in lift at higher angles of attack is known as FLAT PLATE LIFT.

The pressure differential between the upper and lower surfaces of the aerofoil is now greater and the magnitude of the total reaction is increased. CPL ATG CPL DOC 01 Revision : 1/1/2001

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By tracing the airflow as it encounters the aerofoil, another factor can be found. The is light upwash of the airflow before reaching the aerofoil and a downwash after passing the aerofoil. In this process of forcing the airflow downwards, there is an equal and opposite force which forces the aerofoil upwards and, in so doing, adds to the total reaction.

PRESSURE DISTRIBUTION AROUND AN AEROFOIL From the aforementioned regarding the generation of lift, a simple experiment was developed to quantify the pressure distribution around an aerofoil. In this experiment, a series of small holes in the aerofoil surface are connected to glass manometer tubes which are immersed in a liquid. Where there is a decrease of pressure over the upper surface of the aerofoil, a "suction" effect results and the liquid in the corresponding tube is sucked up. Where there is an increased pressure below the aerofoil surface, the liquid is compressed in the tube.

From this experiment, the following points are noteworthy: a)

There is a pressure decrease on the upper surface of the aerofoil and a pressure increase on the lower surface of the aerofoil.

b)

The pressure is not evenly distributed. The pressure decrease above the aerofoil is greater than the pressure increase below the aerofoil, therefore although both upper and lower surfaces contribute to lift, it is the upper surface which contributes most to lift, (almost 80 %).

On highly cambered aerofoils, the airflow is forced to increase its velocity more than on a symmetrical aerofoil. By increasing the angle of attack, the aerofoil becomes effectively thicker, thus giving the airflow greater acceleration over the upper surface of the aerofoil, resulting in a larger pressure decrease with each increased angle of attack. If all the distributed pressures acting on the aerofoil were replaced by a single resultant force, the position on the chord at which this resultant force acts is called the centre of pressure. Its location is a function of both camber and angle of attack CPL ATG CPL DOC 01 Revision : 1/1/2001

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MOVEMENT OF THE CENTRE OF PRESSURE WITH CHANGING ANGLE OF ATTACK The following series of diagrams describe lift and the movement of the centre of pressure at various angles of attack on a general purpose aerofoil.

At approximately -4º angle of attack, no lift is being developed. At 0º angle of attack, the aerofoil is developing lift mainly due to the camber of the upper surface of the aerofoil. As the angle of attack is progressively increased, the lift becomes greater firstly because of the greater acceleration of airflow over the upper surface of the aerofoil, and sedondly because of the increased high pressure below the aerofoil due to ram effect. It is important to note that as the angle of attack is progressively increased, the centre of pressure moves progressively further forward. The movement of the centre of pressure is explained simply with the following philosophy: The centre of pressure is that point through which the total reaction may be considered to act. The total reaction representing the net result of the aerodynamic forces acting on the aerofoil will act at a point where the Venturi is the narrowest, (aerofoil is effectively the thickest), and the airflow over the aerofoil is subject to the greatest acceleration. The increasing lift, with successive increases in angle of attack, cannot carry on indefinitely and at about 15º angle of attack, for general purpose aerofoils, there is a dramatic reduction in lift. The aircraft pitches down and the centre of pressure moves backward. At this point, the aerofoil has been stalled, a subject which will be dealt with fully later.

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THE LIFT FORMULA From the previous experiment, it was found that there are several factors which influence the pressure distribution and, therefore, the lift of an aerofoil. In isolating these factors, the lift formula was derived: LIFT = CL½V²S where: CL

=

coefficient of lift. This represents:

i) ii)

Angle of attack of the aerofoil. The shape of the aerofoil.

This described in terms of thickness/chord ratio and a camber and can be altered in flight, for example, by lowering the flaps. 

=

The density of the air.

For any given altitude during subsonic flight, the density may be considered to be constant. V²

=

The velocity of the aircraft, squared.

S

=

The wing area.

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LIFT/DRAG RATIO It is apparent that for a given amount of lift it is desirable to have the least possible drag from the aerofoil. Typically the greatest lifting effort is obtained at an angle of about 16  and least drag occurs at an angle of attack of about -2. Neither of these angle is satisfactory, as the ratio of lift to drag at these extreme figures is low. What is required is the maximum lifting effort compared with the drag at the same angle, the highest lift/drag ratio. The L/D ratio for an aerofoil at any selected angle of attack can be calculated by dividing the CL at that angle of attack by the corresponding C D. In practice the same result is obtained irrespective of whether the lift and drag or their co-efficients are used for the calculation.

From the above diagram you can see the L/D ratio increases rapidly up to an angle of 4 at which point the lift may be between 12 –25 times the drag, the exact figure is dependant on the aerofoil fitted to the plane. At larger angles the L/D ratio steadily, because even though the lift itself is still increasing the proportion of drag is rising at a faster rate. The most important feature of this graph is the indication of the Angle of Attack for the highest L/D ratio; this angle is the one at which the aerofoil gives its best all round performance.

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DRAG When a body is set in motion through the air, a resistance force opposing its motion is produced. In aerodynamics, this resistance force is known as drag. Drag is that component of the total reaction which acts parallel and opposite to the flight path of the aircraft. THE SOURCE OF DRAG Total drag as expressed in the definition, is made up of two types of drag:  Profile Drag.(parasite)  Induced Drag. PROFILE DRAG Profile drag is also known as zero lift drag or parasite drag. The sources of profile drag are: Form Drag Form drag results whenever a body is in the path of an airflow. The airflow strikes the body causing a resistance called drag. the greater the area being struck by the airflow, the greater the drag. Besides the area, the shape of a body also has a decisive effect on the amount of form drag. In the case of an unstreamlined body, the airflow behind the body is of a turbulent nature, which adds to the amount of drag being produced. In aerofoil terminology, this streamlining is expressed as Fineness Ratio. Fitness Ratio

=

Length Breadth

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The above diagram illustrates how drag can be reduced by changing the shape of the object in the freestream. Skin Friction Drag

As air flows over a body, the roughness of the surface of the body (even if only microscopically rough), causes the airflow to be retarded. This retarding action of the air against the surface causes the airflow above it to be retarded to a lesser degree the further we move away from the surface. This retarding action causes the airflow to change from laminar to turbulent in nature, and it is this turbulence which causes skin friction drag. The rougher the surface, the greater the drag and the larger the surface area, the greater the drag. Furthermore, the greater the viscosity (stickiness) of the air, the greater the skin friction drag. Cold air is more viscous than warm air. Interference Drag Where various components of the aircraft are attached, for example the wings and the fuselage, conflicting airflows cause turbulence which, in turn, increases the drag. Interference drag is effectively reduced by the use of fairing. Parasite drag and all of its components, vary in proportion to the square of the speed of the aircraft.

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INDUCED DRAG Induced drag is also known as lift dependent drag. It is drag which is produced as a direct result of producing lift. Bearing in mind that air always flows from an area of high pressure to an area of low pressure, the air flows around the wing describes the following pattern: On the upper surface of the wing, the air flows from the tip to the root. This is because the ambient air pressure outside of the wingtip is high and the air pressure on the top surface of the wing is low. On the lower surface of the wing, the air flows from root to tip. This is because the air pressure below the wing is high and the ambient air pressure outside of the wingtip is comparatively lower. As these conflicting airflows meet at the trailing edge op the wing, they form small vortices, which move from root to tip. At the wingtip, these vortices combine with the strong under-spillage of air from below the wingtip to above the wingtip and form what is called a Shed Tip Vortex. When viewing an aircraft from behind, the Ram's horn vortex rotates clockwise on the left wing and anti clockwise on the right wing. In order to set up a strong circular motion of air requires energy, and this energy absorption is a part of induced drag. Bearing Newton's Third Law in mind, it becomes apparent that in order to keep a wing suspended in the air, the air below the wing must be forced down. In order for the vortices to force the air down behind the wing requires energy, and this absorption of energy is another part of induced drag. Now, although induced drag is always present whenever lift is being produced, its value is the greatest when the aircraft is at a low speed and high angle of attack. In order to keep the wing suspended in the air requires a certain volume of air to be forced down over a given distance. At a low speed, the distance being covered is far shorter and, therefore, the magnitude of the air being forced down must be far greater in order to keep the volume the same. This results in a greater induced drag. A simple analogy would be to consider a water-skier at low speed. Digging his ski deep into the water and exerting a lot of drag on the motor boat.

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AERODYNAMIC EXPLANATION Consider a section of a wing of infinite span, which is producing lift but has no trailing edge vortices. The diagram shows a total aerodynamic reaction which is divided into lift and drag. The lift component, being equal and opposite to weight, is at right angles to the direction of flight, and the drag component is parallel and opposite to the direction of flight. The angle at which the total reaction lies to the relative airflow is determined only by the angle of attack of the aerofoil. The effect of the downwash – due to the vortices – is to tilt downwards the effective relative airflow thereby reducing the effective angle of attack. To regain the consequent loss of lift the aerofoil must be raised until the original angle of attack is restored. The total reaction now lies at the original angle, but relative to the effective airflow, and the component parallel to the direction of flight is longer. The additional value of the drag – resulting from the presence of wing vortices – is known as INDUCED DRAG.

Summary. Wingtip vortices and the resultant downwash, represent an acceleration, i.e.a change of speed and direction of the airflow. The power absorbed in doing this work may be expressed in terms of an additional drag force known as INDUCED DRAG. CPL ATG CPL DOC 01 Revision : 1/1/2001

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Factors Affecting the Induced Drag The following factors affect the Induced drag: Aspect Ratio Aspect ratio is the ratio of an aerofoil's span to its chord. ASPECT RATIO

=

Span² Chord

A high aspect ratio wing has small wingtips in relation to its wing span. There is thus a smaller area involved in the under-spillage of airflow at the wingtips, resulting in a lower Induced drag. A high aspect ratio aerofoil thus has low Induced drag and a low aspect ratio aerofoil has high Induced drag. Plan Form Planform describes the shape of an aerofoil as seen from above. An aerofoil which is tapered to have small wingtips will have smaller wingtip vortices and therefore less Induced drag.

Weight The greater the weight of an aircraft, the greater the amount of lift the aerofoil will have to produce. This means that the pressure differential between the upper and lower surfaces of the aerofoil will be greater, resulting in stronger vortices with more downwash and greater Induced drag.

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Angle of Attack and Speed In order to maintain a constant value of lift, angle of attack and True Airspeed are inversely proportional. Thus, at low speed the aerofoil must have a large angle of attack, resulting in a high Induced drag. Factors Which Reduce the Induced Drag The following are design features which can be incorporated to reduce Induced drag. Tip Barriers By incorporating tip barriers or winglets on the wingtips, the under-spillage of airflow at the wingtips is reduced, thereby reducing the magnitude of the tip vortices and the Induced drag.

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Washout Washout describes a design feature where the wingtips have a lower angle of incidence than the wing root. This means that the wingtips will always have a lower angle of attack than the wing roots, resulting in a smaller pressure differential at the tips and smaller wingtip vortices with less Induced drag.

Change of Aerofoil Section This describes an aerofoil where the camber is reduced near the wingtips. In this way, the pressure differential and vortices are reduced at the wingtips, resulting in a lower Induced drag.

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The Induced Drag Curve

TOTAL DRAG The total drag on an aircraft during flight is the sum of the Profile drag and the Induced drag. Graphically, the Total drag curve is the combination of the Profile drag and the Induced drag curves.

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Taking into account all of the previously mentioned factors which account for Total drag, the Total drag formula can be derived. TOTAL DRAG

=

CD½pV²S

Where: CD p V² S

= = = =

Coefficient of drag, incorporating angle of attack and aerofoil shape. Density of the air. The True airspeed squared. The surface area of the aerofoil

TOTAL DRAG PROFILE DRAG FORM DRAG

SKIN FRICTION

INDUCED DRAG INTERFERENCE DRAG

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LEVEL FLIGHT THE EFFECT OF POWER CHANGES IN LEVEL FLIGHT When an aircraft is in straight and level flight, lift is equal and opposite to weight and thrust is equal and opposite to drag. Furthermore, to satisfy Newton 1, the nose up pitching moment from the thrust/drag couple and the tailplane is equal to the nose down pitching moment from the lift/weight couple. Visualise the diagram of the 4 forces in straight and level, if power is increased, the state of equilibrium is broken and the thrust/drag couple becomes greater than the lift/weight couple, causing the aircraft to pitch nose up. THE EFFECT OF FLAPS IN LEVEL FLIGHT Referring to the balance of forces in level flight, when flaps are lowered, the following occurs: A nose up pitching moment is caused by the fact that the centre of pressure is moved further forward due to the increased camber. This means that the arm between the lift and weight forces is smaller, resulting in a weaker nose down moment. A nose down pitching moment is caused by the fact that the drag line is effectively lowered when flaps are down. This means that the arm about which the thrust and drag vectors work is shorter, resulting in a weaker nose up pitching moment. A further effect worth noting, is that because the airflow follows the camber of the wing, the increased camber from the lowering of flaps means that the downwash from behind the wing is deflected further downwards, often rendering the elevators less effective.

THE EFFECT OF UNDERCARRIAGE ON LEVEL FLIGHT Unlike the lowering of flaps, there is no lift benefit from lowering undercarriage. The only effect is this - the drag line is lowered, thus shortening the arm about which the thrust/drag vectors work. This has the effect of weakening the nose up thrust/drag moment, resulting in a nose down pitching moment. FLYING FOR RANGE The definition of flying for range is to cover the greatest distance possible for a given amount of fuel. To achieve this requires a compromise between aerodynamic considerations, engine considerations and weather considerations. AERODYNAMIC CONSIDERATIONS To achieve the greatest range for a given amount of fuel, the engine must be overcoming as little drag as possible. This point is found at the best lift/drag ratio, or the velocity for minimum drag, which is the point formed by a tangent on the Preq curve. CPL ATG CPL DOC 01 Revision : 1/1/2001

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ENGINE CONSIDERATIONS For maximum range, the engine must be achieving the best possible fuel consumption. This is achieved at full throttle height for a specific speed. At this height, the throttle has just reached the fully open point to achieve the best range speed. If height is decreased, the engine must be throttled back to maintain the speed. If the height is increased, the speed drops below the optimum and there is no more power to regain the speed. The mixture must be as lean as possible for minimum fuel consumption. As altitude is increased, the air density decreases and the mixture must be leaned off to maintain the correct burning ratio. The RPM must be low. The lower the RPM, the coarser the pitch. The coarser the pitch, the less the drag. Also the coarser the pitch, the better the efficiency of the propeller because energy is being expended to accelerate the airflow, rather than accelerate the propeller. WEATHER CONSIDERATIONS Flying at full throttle height at the correct range speed is ideal for still air conditions. In the event of a headwind, the range will be reduced. Depending on the severity of the headwind, it may be advantageous to sacrifice full throttle height and correct speed by descending to a lower altitude in order to increase the speed and spend less time under the effect of headwind. The net effect however will still be a reduction in range. In the event of a tailwind, the range will be increased. Under these circumstances, full throttle height and best range speed should be maintained, thus enjoying the full benefit of aerodynamic, engine and weather considerations.

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FLYING FOR ENDURANCE The definition of flying for endurance is to remain in the air for as long a time as possible for a given amount of fuel. To achieve this requires a compromise between aerodynamic considerations and engine considerations. AERODYNAMIC CONSIDERATIONS Flying for endurance means that the engine must produce the lowest possible fuel consumption. This is found at the velocity for minimum power (Vmp), which is the lowest point on the Preq curve. ENGINE CONSIDERATIONS Because Preq = Drag x TAS, and TAS is lowest at sea level, the aircraft should be flown at the lowest safe altitude. For the same reasons as mentioned under Flying for Range, the aircraft should still be flown with low RPM and the mixture correctly leaned. NOTE: Because the object is to stay airborne for as long a time as possible and distance covered is not at question, the optimum speed for endurance is unaffected by wind conditions.

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Revision Questions on Lift 1.

The principles on which the production of lift are based on are (a) Boyle's Law (b) Bernoulli's Theorem (c) Reynolds Number

2.

When considering aerodynamic forces, the effect of size is related to (a) Joule's Law (b) Bernoulli's Theorem (c) Reynolds Number

3.

Newton's first Law of Motion, generally termed the Law of Inertia states (a) to every action there is an equal and opposite reaction (b) force is proportional to the product of mass and acceleration (c) every body persists in a state of rest, or of motion in a straight line, unless acted upon by an external unbalanced force.

4.

The angle between the chord line of the wing and the longitudinal axis of the aeroplane is known as the angle of (a) incidence (b) dihedral (c) attack

5.

In order to maintain altitude while decreasing airspeed: a) b) c)

6.

increase angle of attack to compensate for decreasing lift, increase angle of attack to produce more lift than drag, decrease angle of attack to compensate for increasing drag.

If the angle of attack and other factors remain constant and the airspeed is doubled, the lift produced at a higher speed will be: a) twice than that at a lower speed, b) three times more than that at the lower speed, c) four times more than that at the lower speed

7.

A wing designed to produce lift resulting from relatively : a) b) c)

high air pressure below and above the wing surface, low pressure below and high pressure above the wing surface, high pressure below and low pressure above the wing surface.

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8. The changes in aircraft control which must be made to maintain altitude while airspeed is decreasing are to (a) increase angle of attack to produce more lift than weight (b) maintain a constant angle of attack until the desired airspeed is reached, then increase angle of attack (c) to increase angle of attack to compensate for decreasing lift 9.

The point of an aerofoil through which lift acts is the (a) centre of pressure (b) centre of gravity (c) centre of rotation

10.

Lift on a wing is most properly defined as the: (a) differential pressure acting perpendicular to the chord of the wing, (b) force produced perpendicular to the relative flow, (c) reduced pressure resulting from a smooth flow of air over a curved surface.

11.

On a wing, the lift force acts perpendicular to and the drag force acts parallel to the: (a) chordline, (b) longitudinal axis, (c) flight path

12.

During flight at zero angle of attack, the pressure along the upper surface of the wing would be: (a) less than atmospheric pressure, (b) equal to atmospheric pressure, (c) greater than atmospheric pressure.

13.

The angle of attack of a wing directly controls the: (a) amount of airflow above and below the wing, (b) point at which the CG is located, (c) distribution of high and low pressure acting on the wing.

14.

One of the main functions of flaps during the approach and landing is to: (a) (b) (c)

15.

It is true to say concerning the use of flaps during approach and landing that (a) (b) (c)

16.

permit a touchdown at a higher indicated airspeed; increase the angle of descent without increasing airspeed; decrease lift, thus enabling a steeper than normal approach to be made.

flaps decrease lift which increases the stall speed, flaps provide an increase in lift, a steeper than normal approach is necessary because of increase in stall speed.

It is true to say regarding the use of flaps during level turns that:

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(a) (b) (c) 17.

using a constant flap setting and varying bank has no effect on stall speed, the addition of flaps increases the stall speed, the addition of flaps decreases the stall speed.

The maximum allowable airspeed with flaps extended is lower than cruising speed because: (a) the additional lift and drag created would overload the wing structure at higher speeds, (b) the flaps will retract automatically at higher speeds, (c) too much drag is induced.

18.

When gliding into a head wind the best glide angle will be achieved at (a) an IAS which produces the best lift/drag ratio (b) an IAS which is higher than that for best lift/drag ratio (c) an IAS which is lower than that for best lift/drag ratio, but which is higher than that required for best endurance

Revision Questions on Drag 1. If airspeed doubles while the angle of attack remains the same, the drag will: (a) remain the same, (b) double, (c)be four times greater. 2. As airspeed increases in level flight, total drag of an aircraft becomes greater than the total drag produced at the maximum L/D speed because of the: (a)increase in induced drag (b)increase in profile drag (c) decrease in profile drag 3. Changing the angle of attack of a wing, enables control of the: (a)lift, gross weight and drag, (b)lift airspeed and drag, (c) airspeed, weight and drag. 4. As airspeed decreases in level flight, total drag of the aircraft becomes greater than the total drag produced at the maximum L/D speed because of the (a) increase in induced drag, (b) increase in parasite drag, (c) decrease in induced drag.

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(a) a decrease in angle of attack will increase impact pressure below the wing and decrease drag, (b) an increase in angle of attack will decrease impact pressure below the wing and decrease drag, (c) an increase in angle of attack will increase impact pressure below the wing and increase drag, 6. An aircraft at 100 mph produces 1,000 lb. of drag. If angle of attack remains the same but airspeed is doubled, total drag would increase to: (a) 2 000 lbs., (b) 3 000 lbs., (c)4 000 lbs. 7. In comparison with a low aspect ratio wing, a high aspect ratio wing in a constant airflow velocity will have: (a) decreased drag, especially at high angles of attack, (b) increased drag, especially at high angles of attack, (c)increased drag, especially at low angles of attack. 8. In comparison with a high aspect ratio wing, a low aspect ratio wing in a constant airflow velocity will have, (a) decreased drag, especially at low angles of attack. (b) decreased drag, especially at high angles of attack, (c)increased drag, especially at high angles of attack.

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CHAPTER 3 STABILITY Stability of an aircraft means its ability to return to the original condition of flight without any action on the part of the pilot.

An aircraft is in a state of equilibrium if the sum totals of the external forces and of the moments of those forces about any point is zero. An aircraft is stable if on disturbance from its state of equilibrium it tends to return to it. This is usually called "latent stability". STATIC STABILITY The term static stability is concerned with the tendency of resulting motion of a body in attempting to return to a condition of equilibrium. When a body that has been displaced returns to equilibrium, a state of positive static stability exists. When a body that has been displaced has the tendency to continue in the direction of the disturbance, a state of static instability exists. When a body that has been displaced neither tend to return to equilibrium or continue in the displaced direction, a state of neutral static stability exists. The following points concerning the concept of static stability are important: 1.

The concept only has real significance in a system which is initially in equilibrium.

2.

It is only concerned with the tendency of the system to return to equilibrium and not what happens to the system after displacement.

3.

It is only concerned with small displacements within the system and not with relatively large displacement;

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DYNAMIC STABILITY This involves an element of time with that of motion. Consider a system that is statically stable. Once having been displaced from equilibrium, the system will, due to the forces created, return or gently subside back to its original condition, without any tendency to overshoot the point of equilibrium. Such motion is called subsidence or dead-beat return. Such a system is both statically and dynamically stable.

However, while still possessing the tendency to return to equilibrium, the system may, when it reaches the point of equilibrium, do so with finite velocity so that it overshoots. Because the system is statically stable, the new forces retard the system so that it eventually moves back again towards equilibrium with a series of decreasing oscillations. Such motion is called damped phugoid.

The diagram below depicts a statically stable system (i.e. the tendency is always to return or attempt to return to a state of equilibrium) but over a period of time the oscillations may either remain constant or get worse.

A highly undesirable system is one possessing both negative static stability (static instability) and negative stability, as there is no tendency to return to equilibrium, even with an element of time involved.

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LONGITUDINAL STABILITY THE POSITION OF THE CENTRE OF GRAVITY The position of the centre of gravity must remain within certain limits laid down in the Flight Manual. This is necessary because the position of the centre of gravity is a very important factor in determining the stability of the aircraft in the pitching plane. The centre of gravity is located ahead of the centre of pressure so that a nose down moment is provided by the two forces. The further forward the centre of gravity, the more stable the aircraft is in the pitching plane. As we shall see when discussing the effect of the centre of pressure movements, these in normal flight are stable and so the stability provided by the forward location of the centre of gravity is necessary to overcome this tendency.

With the centre of gravity in the middle of its normal range, the aircraft will be normally stable, but with the centre of gravity towards the aft limit the level of stability will be reduced because the arm between the centre of gravity and the centre of pressure is reduced, allowing the effect of movement of the centre of pressure to have greater significance . If the centre of gravity is behind the aft limit the aircraft will be dangerously unstable. As we will see when studying spinning this aft position of the centre of gravity also has dangerous results in relation to the Inertia Moment during a spin.

The nose down moment produced by Lift being behind Weight is usually counteracted by thrust being below drag, but as these are two small forces a down load on the tailplane is usually required to balance the moments.

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POSITION OF THE CENTRE OF PRESSURE We shall soon see that any increase in angle of attack prior to the stall causes the centre of pressure to move forward; this is an unstable movement because by decreasing the nose down moment of Lift and Weight it tends to make the angle of attack increase further which is an unstable action. The position of the Centre of Pressure depends upon the angle of attack which depends upon the following:  Speed: At high speeds, the centre of pressure lies well aft of the centre of gravity but moves forward as the speed is decreased.  Weight: Any increase in weight requires a larger angle of attack resulting in a forward movement of the centre of pressure, or an increase in speed which would keep the centre of pressure in the same position.  Manoeuvres: Any demand for increased lift requires a larger angle of attack with forward movement of the centre of pressure.  Turbulence: Turbulence will result in a change of the relative airflow, which consequently alters the angle of attack, causing the centre of pressure to fluctuate.  Configuration: During a flap or gear change, the centre of pressure is affected. DESIGN OF THE TAILPLANE A wing on its own is not stable and, therefore, need a smaller wing placed at some distance behind it to provide the stabilizing effect required. This small wing is called the tailplane and its effectiveness depends on the following: 1) 2) 3) 4)

Its size. The distance of its centre of pressure behind the aircraft centre of gravity. Its shape. Its aspect ratio.

On most occasions, the exact position and incidence of the tailplane is determined by the downwash of the mainplanes. It is so positioned that this downwash meets the tailplane at an angle different to the angle at which the airflow meets the mainplanes. This method of obtaining longitudinal stability is called longitudinal dihedral. Consider an aircraft in level flight. Assume the angle of attack on the mainplanes to be 4 degrees and that of the tailplane 2º. Should a disturbance be encountered giving a 2 degree nose up pitching moment, then the angle of attack of the mainplanes would be 6 degrees, representing a 50 % increase in lift. However, the tailplane is also subjected to the disturbance, thus resulting in an angle of attack of 4 degrees. This represents a 100 % lift increase, thus raising the tail resulting in a nose down correcting tendency.

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LATERAL STABILITY Lateral stability is stability in the rolling plane. Factors that affect lateral stability are: i) ii) iii) iv)

Dihedral. Keel effect. Sweepback. High lift and low centre of gravity.

DIHEDRAL When the span of an aircraft wing is tilted upwards, this inclination is called dihedral. Downward inclination is called Anhedral and provides a means of instability, especially in the case of fighter type aircraft. When an aircraft is displaced from its lateral position, it begins to slip sideways. This is due to the inclination of the tilted lift vector and the weight. During the slip, the direction of the relative airflow is altered, as shown. However, due to dihedral, the lower wing is now at a greater angle of attack than the higher wing, so producing greater lift which tends to roll the aircraft back to the laterally stable position.

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KEEL EFFECT When an appreciable area of the fuselage and vertical tailplane lie above the centre of gravity of an aircraft, the "keel effect" helps return the aircraft to a laterally stable position. Below depicts an aircraft in a side slip. Although such an attitude is unlikely, it serves as an illustration to explain keel effect which is exactly what happens in a sideslip of only a few degrees.

SWEEPBACK Consider an aircraft with swept wings in a sideslip. With regard to the lower starboard wing, the following is noted: 1)

The effective span is larger than that of the higher wing, thus producing more lift on the lower wing causing the aircraft to roll out of the banked attitude. 2) It has an effective increase in camber i.e. more lift produced. 3) It has a greater aspect ratio. 4) It is not shielded as with higher wings. These factors contribute to an increase in lift and thereby rolling the aircraft to a level attitude. This provides lateral stability.

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HIGH WING AND LOW CENTRE OF GRAVITY This method acts similarly to a pendulum, thus producing a righting moment due to the "pendulum effect".

DIRECTIONAL STABILITY Directional stability is stability in the yawing plane. methods: CPL ATG CPL DOC 01 Revision : 1/1/2001

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HIGH/LARGE VERTICAL FIN Weathercock action is obtained from the vertical fin. The distance between the vertical axis and the fin depends on the amount of stability required. Too long a moment arm could result in an "over corrected" situation. The same weathercocking action is directly provided by the fuselage area behind the C of G.

SWEEPBACK As discussed, a larger frontal area provides an increase in drag on that wing which is yawed forwards. This increase in drag, therefore, tends to oppose the disturbing yaw and maintain a satisfactory condition of directional stability.

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THE RELATIONSHIP BETWEEN LATERAL AND DIRECTIONAL STABILITY In the study of stability, lateral and directional stability are so closely interconnected, that it is impossible to consider one without the other: SPIRAL INSTABILITY Consider an aircraft that is very stable directionally, but not laterally stable e.g. large fin and little or no dihedral. The aircraft would have a marked tendency to sideslip during a banked attitude. Due to the large tail fin, a weathercocking action is set up and this only increases the tendency of the nose to yaw towards the lower wing. So, roll plus yaw, the nose yaws below the horizon and the aircraft enters a spiral dive. Bank becomes steeper and steeper and the aircraft becomes uncontrollable. This is called spiral instability, caused by too much stability (directional). This is in fact similar to the further effect of the ailerons, primary effect being to roll the aircraft, further effect being yaw (weathercocking). ROLL WITH YAW When yaw is induced, either by deliberate application of rudder or the failure of one engine, in a multi-engined aircraft, the outer wing travels further than the inner wing. This results in an increased speed in which case the outer wing develops more lift and begins to rise (i.e. roll). So we have yaw plus roll, the nose of the aircraft is still yawing and it yaws below the horizon and the aircraft enters a spiral dive; the primary effect of rudder is yaw and the further effect is roll.

OSCILLATORY INSTABILITY When an aircraft has a combination of roll and yaw, it is said to have oscillatory instability. The motion is called Dutch Rolling when rolling predominates and Snaking when yawing predominates. Smaller dihedral angles and tail fins will minimise this type of instability.

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YAW DAMPERS Dutch roll can cause big handling problems in large aircraft so yaw dampers are often fitted to prevent yaw from developing by automatically applying a cancelling rudder input when yaw is sensed. The 2 main types of yaw dampers found in aircraft today are PARALLEL type and the SERIES type. The parallel type moves the pilots rudder pedals as it operates and so must be switched off for take off and landing to avoid problems with crosswinds and asymmetric performance. The series type is the more modern type and does not move the rudder pedals as it operates and so may be switched on during all phases of flight.

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Revision Questions 1. If an aircraft is stable, this means that: (a) it is in a state of balance; (b) if it is displaced it will return to its original position without any correction by the pilot; (c) if it is displaced it must be returned to its original position by the pilot operating the controls. 2. For an aircraft which is neutrally stable in roll, following a wing drop: (a) the wing would tend to return to the level position; (b) the wing would continue to drop; (c) the wing would remain in its displaced position. 3. After a disturbance in pitch an aircraft oscillates in pitch with increasing amplitude. It is: (a) statically and dynamically unstable; (b) statically stable but dynamically unstable; (c) statically unstable but dynamically stable. 4. Longitudinal stability is given by: (a) the fin; (b) the wing dihedral; (c) the horizontal tailplane. 5. An aircraft is constructed with dihedral to provide: (a) lateral stability about the longitudinal axis; (b) longitudinal stability about the lateral axis; (c) lateral stability about the normal axis. 6. Lateral stability is given by: (a) the ailerons; (b) the wing dihedral; (c) the horizontal tailplane. 7. Stability about the lateral axis is given by: (a) wing dihedral; (b) the horizontal tailplane; (c) the ailerons. 8. Stability about the longitudinal axis is given by: (a) elevators; (b) ailerons; (c) wing dihedral

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9. Moving the centre of gravity aft will: (a) reduce longitudinal stability; (b) increase longitudinal stability; (c) have no effect on longitudinal stability. 10. If the aircraft has a nose-up pitch displacement, the effective angle of attack of the tailplane: (a) remains the same; (b) changes and causes the tailplane to apply a restoring moment; (c) will not change if the pitch up was due to elevator selection. 11. The longitudinal static stability of an aircraft: (a) is reduced by the effects of wing downwash; (b) is increased by the effects of wing downwash; (c) is not affected by wing downwash. 12. To ensure longitudinal stability in flight, the position of the C of G: (a) must always coincide with the C of P; (b) must be forward of the Neutral Point; (c) must be aft of the Neutral Point. 13. Wing dihedral gives a stabilising rolling moment by causing an increase in lift: (a) on the down-going wing when the aircraft rolls; (b) on the lower wing when the aircraft sideslips; (c) on the lower wing whenever the aircraft is in a banked altitude. 14. A high wing configuration gives: (a) more lateral stability than a low wing; (b) less lateral stability than a low wing; (c) the same lateral stability as a low wing. 15. After a disturbance in pitch an aircraft oscillates for a long time with only small reductions of amplitude on each oscillation. It would be said to have: (a) low damping; (b) high damping; (c) negative damping. 16. The presence of the fuselage in an aircraft with a high wing during a sideslip: (a) increases the lift on the lower wing and decreases the lift on the upper wing thus creating a stabilising moment; (b) increases the lift on both wings thus creating a stabilising moment; (c) decreases the lift on the lower wing and, increases the lift on the upper wing thus creating a destabilising moment. 17. Sweepback of the wings will: CPL ATG CPL DOC 01 Revision : 1/1/2001

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(a) increase lateral stability; (b) decrease lateral stability; (c) not affect the lateral stability. 18. If an aircraft is yawed to a large angle of sideslip: (a) directional stability will be lost; (b) if the sideslip angle is too large the fin may stall and directional stability will be decreased; (c) the rudder will always have to be used to return the aircraft to its original position. 19. The fin gives: (a) directional stability about the longitudinal axis (b) directional stability about the normal axis; (c) longitudinal stability about the lateral axis. 20. Increasing the size of the fin: (a) reduces lateral stability; (b) increases longitudinal stability and directional control; (c) increases the size of the keel surface giving increased directional stability. 21. Pendulum stability is a property possessed by: (a) aircraft with swept back wings; (b) aircraft with high wing configuration; (c) aircraft with low wing configuration. 22. An aircraft with a `Dutch roll' instability will: (a) go into a spiral dive following a lateral disturbance; (b) experience simultaneous oscillations in roll and yaw; (c) experience oscillations in pitch. 23. Dutch roll may be prevented by: (a) having the wings swept back; (b) reducing the size of the fin; (c) fitting yaw dampers. 24. An aircraft is yawed to starboard and the rudder is then centralized. if it then yaws to port it is: (a) directionally neutrally stable; (b) directionally statically stable; (c) directionally dynamically stable.

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25. Increasing the size of the fin will: (a) increase lateral stability (b) decrease lateral stability (c) not affect lateral stability. 26. After a disturbance in pitch an aircraft oscillates with increasing amplitude. It is: (a) dynamically neutral (b) dynamically stable but statically unstable (c) dynamically unstable longitudinally. 27. To ensure longitudinal stability in flight, the position of the C of G: (a) should not be forward of the neutral point (b) should not be aft of the neutral point (c) should coincide with the neutral point. 28. Moving the centre of gravity aft will: (a) increase longitudinal stability (b) reduce longitudinal stability (c) have no effect on longitudinal stability. 29. After a disturbance in pitch, an aircraft continues to oscillate at a constant amplitude. It is: (a) longitudinally neutrally stable (b) laterally unstable (c) longitudinally unstable. 30. Dutch roll is: (a) A type of slow roll (b) primarily a pitching instability (c) a combined rolling and yawing motion. 31. When the C of G is close to the forward limit: (a) very small forces are required on the control column to produce pitch (b) longitudinal stability is reduced (c) larger stick forces are required to pitch because the aircraft is very stable. 32. Stability of an aircraft about its lateral axis is normally provided by the: (a) tailplane (b) ailerons (c) elevators.

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33.

The aircraft nose initially tends to return to the original position after the elevator is pressed forward and released, the aircraft displays:

a) b) c)

negative stability, positive static stability negative dynamic stability.

34.

Rotation about the lateral axis is known as:

a) b) c)

pitching and is controlled with the elevator, rolling and is controlled with the ailerons, yawing and is controlled with the ailerons.

35.

If an aircraft is loaded to the rear of the CG range, it will tend to be:

(a) (b) (c)

unstable about the lateral axis, sluggish in aileron control, unstable about the longitudinal axis.

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CHAPTER 4 CLIMBING, DESCENDING & TURNING CLIMBING THE BALANCE OF FORCES IN THE CLIMB

WEIGHT Although weight always acts towards the centre of the earth, the weight vector can be resolved into a horizontal component and a vertical component. The horizontal component is parallel to and in the same direction as drag. It therefore supplements the drag and is known as weight apparent drag. The effect of increased weight in the climb is as follows: If weight is increased, the total lift must be increased. The increase in lift can come from one of two sources - angle of attack or speed. To increase the angle of attack is not practical because of the danger of reaching the critical angle of attack. To increase the speed is the only alternative, but to do this, the angle of attack must be reduced and in so doing, the overall climb performance of the aircraft will be reduced. LIFT In the climb, the lift is perpendicular to the flight path of the aircraft . Generally one would think that lift should be more than weight in the climb. Refer to the diagram above, it can be seen that lift is actually less than weight during the climb because thrust is inclined upward and therefore has a vertical component aiding lift.

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THRUST Thrust acts parallel and opposite to drag. In terms of magnitude, thrust must be equal to the sum of drag and weight apparent drag. When speaking of thrust, it is correct to speak in terms of power available. Power available

=

Thrust x TAS

At zero speed, there is zero Pav. At high speed, Pav approaches zero because propeller efficiency reduces beyond a certain speed because the high speed airflow is passing through the propeller faster than the propeller is forcing the airflow back. There is, therefore, a reducing amount of true thrust from the propeller. As the aircraft gains altitude in the climb, the Pav reduces. Although the TAS increases, this is overcome by the fact that the thrust reduces as a result of reduced air density. DRAG In flight, an aircraft requires power to overcome drag; this is the source of the power required curve. Power required

=

Drag x TAS

At low speed, the aircraft requires a lot of power to overcome high induced drag . At high speed, the aircraft requires a lot of power to overcome high profile drag. As altitude is gained during the climb, the TAS increases and thus the power required curve moves upward.

THE BEST RATE OF CLIMB The best rate of climb is defined as climbing at a speed which will ensure the most height per unit of time ( Vy). Referring below, it can be seen that the best rate of climb occurs at that speed at which there is the greatest difference between Pav and Preq (most excess power).

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THE BEST ANGLE OF CLIMB ( VX) By definition, the best angle of climb is a climb at a speed which will ensure the greatest amount of height per distance flown.

Referring to the diagram and formula, it can be seen that the best angle of climb occurs at a speed for which there is the greatest difference between thrust and drag. Given that the aircraft is at full thrust during the climb, the object must then be to fly at the speed at which there is the minimum drag (best Lift/Drag ratio). This speed is traditionally lower than the best rate of climb speed. THE EFFECT OF WIND ON THE CLIMB When considering the effect of wind on the climb, it helps to conceptualise that, firstly, the aircraft climbs in a block of air and, secondly, the block of air moves over the ground because of wind. Referring to the diagrams: The effect of headwind on the rate of climb - Nil. Angle of climb - increased.

Effect of a tailwind on the  

Rate of climb - Nil. Angle of climb - decreased.

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THE CRUISE CLIMB By definition, a cruise climb is one which will ensure a good rate of climb as well as a high forward speed. The cruise climb is typically used for cross-country flights. A good example for a light aircraft is this: An increase in speed of + 20kt above the best rate of climb speed, will usually lead to a reduction in the best rate of climb of + 7 %, with an increase in speed of + 25 %. DESCENDING THE BALANCE OF FORCE IN THE GLIDE/DESCENT

With reference to the balance of forces in level flight, when all or some of the thrust is removed, the aircraft will experience a nose down pitch. Once the correct attitude has been selected, the aircraft will commence a glide/descent at a specified speed and constant rate of descent. The forces in the descent reach a state of equilibrium, diagrammatically represented above. WEIGHT Weight continues to act towards the centre of the earth. With the removal of the thrust vector, the Lift/Weight couple overcomes the Thrust/Drag couple, causing a nose down pitching moment. With the angular inclination between the flight path and weight vector, the weight can be resolved into a vertical and horizontal component. As can be seen, the horizontal component acts parallel and opposite to drag and is therefore, in effect, thrust. This force is commonly referred to as weight apparent thrust.

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LIFT In a descent, the aircraft still has forward speed and a positive angle of attack and thus develops sufficient lift to balance weight. Although the lift vector has been tilted forward and it is the vertical component of lift which must balance weight, the trigonometric relationship between the two is a cosine function and even at unusually high descent angles, such as 15º, the angular component still remains very close in magnitude to the vertical component (97 %). In order to produce the correct value of lift, speed and angle of attack (attitude) must be adjusted. The two have an inverse proportion, the higher the attitude, the lower the speed, and vice-versa. THRUST Referring to the balance of forces when an aircraft is descending, the origin of the thrust is: Weight apparent thrust in the case of a power off glide, or The sum of weight apparent thrust and engine thrust in the case of a power assisted descent. The glide path or angle of an aircraft can be determined trigonometrically with the equation: The greater the power in the descent, the less the weight apparent thrust must be for the sum of the two to equal drag. The less the weight apparent thrust, the smaller the gliding angle or the shallower the glide. Provided the speed is kept the same, the smaller the gliding angle, the lower the rate of descent and the further the glide distance. DRAG In all phases of flight, including the descent, total drag is equal to the sum of the induced drag and profile drag. In the descent, total drag continues to act parallel and opposite to the flight path of the aircraft. With reference to the Total Drag curve, there is a specific speed at which drag is the least (best L/D ratio). To achieve the smallest gliding angle, the aircraft should be flown at this speed. If the aircraft is flown any slower or any faster than this speed, the drag will increase. If the drag is increased, the weight apparent thrust must be increased by increasing the gliding angle. To achieve the least drag and consequently the smallest glide angle, the aircraft should be flown in the clean configuration and with a coarse propeller pitch. THE MINIMUM ANGLE / MAXIMUM RANGE GLIDE By obtaining the minimum glide angle, the aircraft will achieve the greatest gliding range. This means that the horizontal distance covered per amount of height lost will be the greatest. Referring to the diagrams and formulae below, it can be seen that by flying the aircraft at the speed for the best L/D ratio (lowest drag, the weight apparent thrust needs to be the least and, therefore, the gliding angle will be the least.

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ACHIEVE SHALLOW GLIDE ANGLE GLIDE ANGLE SIN Ø =

T-D (W SIN ) W

FLY AT SPEED FOR MIN DRAG

THE MINIMUM RATE OF DESCENT/MAXIMUM ENDURANCE GLIDE By flying at a specific speed which ensures the minimum rate of descent, the aircraft will achieve the maximum endurance or remain airborne for the longest possible time. Referring to the diagrams and formulae below, it can be seen that by flying at the velocity for minimum power, (to overcome drag), the rate of descent will be the lowest, thus ensuring the maximum endurance in the glide.

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THE EFFECT OF WIND ON THE GLIDE When considering the effect of wind on the glide, it helps to conceptualise the fact that the aircraft flies in a mass of air and then this mass of air moves over the ground due to the wind. Referring below it can be seen that: In a headwind:  The glide angle from the cockpit remains the same.  The glide angle from a ground observer appears steeper.  The gliding distance is reduced. In tailwind:  The glide angle from the cockpit remains the same.  The glide angle from a ground observer appears shallower.  The gliding distance is increased.

THE EFFECT OF WEIGHT ON THE GLIDE Referring to the diagram below: If the weight of the aircraft is increased, the weight apparent thrust increases. If the weight apparent thrust increases, the speed of the aircraft will increase. If the speed increases, the drag increases - now drag again balances thrust. If the speed increases, the lift increases - now lift again balances weight. The net result is that the balance of forces is restored. There is no change to the gliding angle, thus no change to the gliding range. The only difference is that the aircraft flies at a higher speed and therefore has a higher rate of descent, i.e. it reaches the ground sooner, but at the same place.

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THE CRUISE DESCENT The cruise descent is a power assisted descent. The aircraft has a lower rate of descent and a higher forward speed. Although the engine consumes more fuel than it would in a glide, the earlier arrival at destination due to the higher forward speed make the descent more efficient. TURNING THE BALANCE OF FORCES IN A TURN In a turn, the lift vector is tilted in the direction of the turn by an amount equal to the angle of bank. The vertical component of Total Lift is called effective lift and it must still balance weight. In order for the effective lift to balance weight so that the aircraft can achieve a level turn, the total lift must be increased. To achieve this added total lift, the angle of attack must be increased in the turn. An increase in the angle of attack will lead to an increase in the induced drag, which in turn reduces the speed of the aircraft slightly if the thrust setting is maintained constant. The horizontal component of Total Lift is called Centripetal Force. Centripetal force works towards the centre of the turn and it is the force which turns the aircraft . In order to satisfy Newton's Laws, there must be a force equal and opposite to Centripetal Force. This force is called Centrifugal Force, which is often verified with the popular analogy of releasing a stone swinging at the end of a piece of string.

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LOAD FACTOR Load factor is best described as an increase in apparent weight. Mathematically

LF

=

L W

In straight and level flight, L = W, therefore LF = 1.

During a level turn, the total lift is greater than the weight, yet level flight is maintained. There has been an apparent increase in weight and the load factor is greater than 1.

If load factor is described by the relationship between lift and weight, then its value can be determined trigonometrically with the function Secant.

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Total Lift Weight L W

Sec Ø

=

Sec Ø

=

Sec Ø

=

n

=

N

1 COS 

1 COS 300

E.g. An aircraft at 30º bank

=

1.15

Therefore, the load factor is equal to 1.15 and the aircraft and pilot experience and apparent weight of 1.15 times their actual weight.

In order to sustain a level turn, the angle of attack of the aircraft must be increased. When the angle of attack is increased, the induced drag and, by association, the Total Drag is increased. Preq = TAS x Drag, therefore the Preq is increased in a turn. Because of the increase in drag, the speed of the aircraft will reduce. Should the pilot wish to maintain a constant speed in the turn, it would require the expenditure of excess power, (difference between Pav and Preq). With each successive increase in the angle of bank, the angle of attack must increase, the drag increases, the Preq increases and the excess power to maintain the speed decreases.

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TURN RADIUS The term turn radius describes the distance of the aircraft from the centre of the turn. The mathematical formula for determining turn radius (r) is: r

=

V2 g tan 

Ø

=

angle of bank.

where:

This formula is described fully under Minimum Radius and Maximum Rate Turns. From the formula, the following deductions can be made:  If TAS and bank angle are maintained constant, the turn radius will be constant and the size and weight of the aircraft will have no affect on the turn radius. Therefore if a B747 and a C172 have the same TAS and angle of bank, they will have the same turn radius.  If TAS is increased, but bank angle is kept the same, the radius will increase.

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 If the TAS is kept the same, but bank angle is increased, the radius will decrease. 

 If TAS is increased and the same radius is required then the AoB must be increased.

TURN RATE The term turn rate describes the amount of degrees of turn that the aircraft covers per unit of time, usually assessed in degrees per second. The mathematical formula for determining turn rate n is: n

=

v r

This formula is described fully under Minimum Radius and Maximum Rate Turns.

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For the formula, the following deductions can be made: The greater the TAS, the greater the rate of turn.  The smaller the radius, the greater the rate of turn.

THE EFFECT OF WIND ON THE TURN When considering the effect of wind on the turn, it helps to conceptualise that firstly, the aircraft turns in a mass of air and, secondly, this mass of air moves over the ground. This means that turn rate and turn radius are unaffected by the wind conditions. What does change however, is the aircraft's apparent change in turn radius when its position is viewed in relationship to the ground. THE EFFECT OF AIRCRAFT WEIGHT IN RELATION TO TURNING An increase in weight requires an increase in lift. Lift is increased by increasing the angle of attack. When the angle of attack is increased, induced drag increases. When induced drag is increased, the aircraft speed is reduced. Thus the following factors are affected:

 The aircraft attitude (angle of attack) is higher during the turn.  Unless power is applied, the speed is lower during the turn.  Because the speed is reduced, the radius of turn is reduced: 

r

=

V2 g tan 

 Although both the TAS and radius of turn are decreased, the radius decreases proportionally more, thus the rate of turn is increased. n

=

v r

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THE EFFECT OF BALANCE ON THE TURN Rudder applications must be made in the direction of turn to keep the turn balance.

If too much rudder is used in the direction of turn, the excessive lifting force from the tail will cause the aircraft to skid out of the turn, indicated on the Turn Co-ordinator by the ball being displaced out of the direction to turn.

If too little rudder is used in the direction of turn, the lifting force from the tail will be too small, causing the aircraft to slip into the turn, indicated by the ball being displaced in the direction of turn.

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AIRCRAFT ATTITUDE IN THE TURN In a turn, the angle of attack must be greater then for the straight and level condition of flight. From the cockpit, the direct indication of angle of attack is attitude which, clearly, will also be higher than for the straight and level conditions. THE EFFECT OF DENSITY ALTITUDE ON THE TURN Up to a certain point, TAS increases with increased density altitude . This has the following effect on the turn: An increase in TAS will cause the turn radius to increase: r

=

V2 g tan 

Although the TAS and radius both increase, the radius increases proportionally greater, thus the turn rate will decrease: n

=

v r

CLIMBING TURNS Referring to the diagram, it can be seen that the movement of the inner and outer wings can be resolved into an upward component and a forward component. Over a given time, the upward component for both wings are the same, but over a given time the outer wing has travelled a greater distance. By resolving the upward and forward velocities into a relative airflow, it can be seen that the angle of attack of the outer wing is greater than the angle of attack of the inner wing. From the lift formula then, the outer wing develops more lift than the inner wing and to maintain a constant angle of bank in the climbing turn, bank must be held off.

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DESCENDING TURNS Referring to the diagram, it can be seen that the movement of the inner and outer wings can be resolved into a downward component and a forward component. Over a given time, the downward component for both wings is the same, but over a given time, the outer wing has travelled a greater distance than the inner wing. By resolving the downward and forward velocities into a relative airflow, it can be seen that the angle of attack of the inner wing is greatest. From the lift formula then, the inner wing develops more lift than the outer wing and to maintain a constant angle of bank in the descending turn, bank must be held on. In some aircraft. Other aircraft may be more adversely effected by the fact that the outer wing is travelling faster and that the outer wing is producing more lift, hence some aircraft may overbank in a descending turn and require bank to be held off.

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Revision Questions 1. The angle of climb is proportional to: (a) the amount by which the lift exceeds the weight (b) the amount by which the thrust exceeds the drag (c) the angle of attack of the wing. 2. As altitude increases the excess thrust at a given IAS: (a) decreases because drag increases and thrust decreases (b) increases because drag decreases and thrust is constant (c) decreases because thrust decreases and drag is constant. 3. As altitude increases the excess power available: (a) decreases because the power available decreases and power required is constant (b) increases because the power required decreases and power available is constant (c) decreases because the power available decreases and power required increases. 4. To cover the greatest distance when gliding, the gliding speed must be: (a) near to the stalling speed (b) as high as possible within VNE limits (c) the one that gives the highest L/D ratio. 5. If weight is increased the maximum gliding range of an aircraft: (a) decreases (b) increases (c) remains the same. 6. As bank angle is increased in a level turn at a constant IAS, the load factor will: (a) remain the same (b) increase (c) decrease. 7. In a level turn at a constant IAS: (a) the drag will be greater than in level flight because of the increased induced drag (b) the drag will be the same as in level flight because the IAS is the same (c) the drag will be less than in level flight because the lift is less. 8. For a level turn at a constant IAS if the radius of turn is decreased the bank angle and load factor will: (a) increase (b) decrease (c) remain the same.

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CHAPTER 5 STALLING When any object moves through a gas or liquid, molecules of the fluid tend to adhere to the surface of that object. This is due to the viscosity of the fluid. If a layer of molecules clings to the surface what of the molecules adjacent to the layer? Viscosity also means that there is a degree of attraction between the individual molecules of that fluid. Thus the molecules dragged along at the surface will tend to drag their neighbours along, but not with the same velocity. The layer of air adjacent a body in which this dragging effect is felt is termed the boundary layer, and as we shall see it is this layer that actually determines the maximum amount of lift that can possibly be generated by any particular wing. VELOCITY PROFILES Velocity profiles are drawn to help visualize the local velocity of an airstream in the boundary layer. Below shows one of these. The relative velocity at the surface is zero, and the relative velocities at various levels away from the surface increase until the outer edge of the boundary layer is reached.

LAMINAR AND TURBULENT FLOW The beginning of airflow at the leading edge of a smooth airfoil surface produces a very thin layer of smooth airflow. This type of airflow is called laminar flow and is characterized by smooth regular streamlines. As the airflow moves back from the leading edge, the boundary layer thickens and becomes unstable. Small pressure disturbances cause the unstable airflow to tumble, and intermixing of the air particles take place. This type of airflow is called turbulent flow. Above shows the boundary layer flowing on a flat plate and shows the increasing thickness of the boundary layer and transition from laminar to turbulent flow. The thickness is greatly exaggerated in the drawing. Velocity profiles are different in laminar and turbulent flow. Laminar profiles show a gradual decrease in the relative velocity from the outer edge of the boundary layer to the surface . Turbulent flow involves rapid intermixing of the air levels, and fast removing air speeds up the particles near the surfaces, result in profiles as shown

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TRANSITION POINT The point at which laminar flow converts to turbulent flow i.e. transition to turbulence, is termed Transition Point . SUMMARY Air is viscous and clings to the surface of an object over which it flows . This is the basic mechanism for the creation of the boundary layer. The boundary layer may be laminar or turbulent in flow. When turbulent it is thicker and possesses more kinetic energy than a laminar flow boundary layer. INVERSE AND ADVERSE PRESSURE GRADIENTS As air flows over a wing it comes into contact with different pressures. Initially air flows from the leading edge (High pressure) to the low pressure at about 25 % cord during this stage the air wants to flow from the high pressure to the low pressure. This is called the INVERSE PRESSURE GRADIENT. After 25 % chord the air enters an area of ADVERSE pressure gradient from the low pressure area to a high pressure area at the trailing edge of the wing. It is important that you understand that a turbulent boundary layer by virtue of its higher Kinetic energy, can progress further into an adverse pressure gradient than can a laminar boundary layer. AIRFLOW SEPARATION AND SEPARATION POINT The airflow in the boundary layer is being acted on by two forces, friction forces and adverse pressure gradient forces. The friction forces between the surface of the object in the airstream and the air particles, and the friction forces between the particles themselves, tend to reduce the relative velocity to zero. This results in an aerodynamically 'dead' air layer close to the surface. Second, the air is being slowed by the adverse pressure gradient (discussed above). The result of the air slowing creates a stagnated region close to the surface of the object. Airflow from outside the boundary layer will overrun the point of stagnation and cause the boundary layer to separate from the surface. The point at which this occurs is termed the separation point, and in effect very little lift is produced from the airfoil surface aft of the separation point.

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In the presence of an adverse pressure gradient, such as exists at high angles of attack, the boundary layer is slowed down, stopped, then reversed in its direction of flow. Since the boundary layer no longer follows the form or shape of the object over which it flows neither does the airflow riding on the boundary layer. The separation of the boundary layer causes a breakdown of the orderly flow of air about the wing, and very little lift is produced aft of the separation point. If a large portion of the wings has separated flow the wings will no longer support the aircraft in level flight and the aircraft is said to be stalled. STALLING, AUTOROTATION AND SPINNING The stall may be described as that condition of flight wherein airflow separates from lifting surfaces at high angles of attack. Lift is drastically reduced, drag increased and controlled flight and manoeuvre is not possible whilst the airfoils remain stalled. THE FORCES ACTING ON AN AIRCRAFT DURING THE STALL On a normal subsonic section no flow separation occurs at low incidences, the now being attached over the rear part of the surface in the form of a turbulent boundary layer. As the incidence increases, so the adverse pressure gradient is increased and the boundary layer will begin to separate from the surface near the trailing edge of the wing. As the angle of attack is further increased the separation point will move forwards along the surface of the wing towards the leading edge. As the separation point moves forward the slope of the lift incidence curve decreases and eventually an incidence is reached at which the wing is said to stall, i.e. the separation point moves rapidly forward. The flow over the upper surface of the wing is then broken down and the lift produced by the wing decreases

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CRITICAL ANGLE The marked reduction in the lift coefficient, which accompanies the breakdown of airflow over the wing, occurs at the critical angel of attack for a particular wing. In subsonic flight an aircraft will always stall at the same critical angle of attack. A typical lift curve showing the critical angle is shown, it should be noted that not all lift is lost at the critical angle, in fact the aerofoil will give a certain amount of lift up to 90° ('flat - plate' lift is still available).

An aircraft will stall if the critical angle is exceeded. Nothing else can make an airfoil stall. The speed at which an airfoil stalls is determined by weight and load factor, but the stall angle remains the same, regardless of the speed at which the critical angle is obtained

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Traditionally, the angle of incidence of the horizontal stabiliser is less than that of the wings. Thus, when the wings are stalled, the horizontal stabiliser still produces lift and this will lead to a nose down pitching moment, or reduction in angle of attack. As it was an excessive angle of attack which leads to the stall, the reduction in angle of attack will cause the CP to move rearward and the aerofoil to unstall, regardless of the speed of the aircraft. As airflow separates from the wing, downwash is reduced and is eventually destroyed. This causes an increase in angle of attack on the horizontal stabilizer which reduces the down load on the tail to pitch the nose of the aircraft down. This action assists recovery and forms part of the longitudinal stability of the aircraft.

THE BASIC STALLING SPEED It must be remembered that an aircraft may be stalled at any speed; it is the angle of attack which remains constant for the stall. Because most smaller aircraft do not indicate the angle of attack on a gauge, it is customary for the aircraft manual to quote a stalling speed which corresponds to the staIling angle of attack under given conditions. These are: Aircraft at a given weight, with flaps and undercarriage retracted and the engine throttled back - basic stalling speed (Vs). If any of these factors are changed, the stall speed will be different; for example the stalling speed in the landing configuration, with flap and undercarriage extended (Vso).

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The basic stall speed is derived as follows: During level flight, L = W. Therefore L = W = CL½pV²S Before the stall, the CL will be at a maximum value (CL max) and the speed will be at a minimum (Vs)- basic stall speed. The formula will now read: W =

CL max ½ pV²S

The symptoms of the approaching stall are:  High nose attitude - which corresponds to the high angle of attack.  Low speed.  Ineffective controls - due to the reduced speed of the airflow over the controls.  Stall warning - buzzer or light.  Buffet - due to the turbulent airflow from behind the separation point striking the tailplane. Aircraft which have a T-tail configuration will experience less buffet than aircraft with a conventional tailplane due to the fact that the tailplane is elevated above the turbulent airflow from behind the wings. On aircraft with a conventional tailplane, the application of flaps will tend to deflect the turbulent airflow downward and below the tailplane, thus reducing the amount of turbulence felt by the pilot on the controls.

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THE RECOVERY FROM THE STALL The recovery procedure is designed first and foremost to unstall the mainplane.This can only be achieved by reducing the angle of attack At this point, the aircraft will still have a low airspeed and before any attempt can be made to recover, the aircraft to a climb, the speed will have to be increased. To do this, one of two options may be used: (ii.) The transfer of potential energy (height), into kinetic energy (speed), by selecting a nose low attitude. (iii.) the transfer of mechanical energy (power), into kinetic energy (speed), by selecting a level flight attitude and immediately opening full power. This method will obviously result in a smaller height loss. Once the aircraft has again safely reach flying speed, a climb away may be initiated. THE EFFECT OF WEIGHT AT THE STALL The effect of greater weight at the stall will make no difference to the stalling angle of attack. However, by referring to the formula for the basic stalling speed –

Vs

=

W CL max ½ pS

it can be seen that the greater the weight, the higher the basic stall speed

THE EFFECT OF WEIGHT DISTRIBUTION AT THE STALL If the centre of gravity is too far forward (weight in the nose), the nose down pitch at the stall will be very strong due to the length of the arm between the weight vector and the lift vector from the tailplane. This strong nose-down pitch could result in an excessively noselow attitude and consequent height loss before recovery. If the C of G is too far aft (weight in the tail), the C of P may move ahead of the C of G as the aircraft approaches the stall. The lift/weight couple now reverses its trend to a nose-up pitching moment, thus accelerating the stall. A forward C of G will require greater down load on the tail during S & L flight. This will increase the amount of weight that the lift must support. An increase in weight will cause an increase in Stalling Speed. So aircraft which are nose heavy will have higher stalling speeds than aircraft which are tail heavy. This effect is very small on small aircraft.

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EFFECT OF MANOEUVRES, "G" LOADING, ON THE STALL SPEED In any condition of flight where lift exceeds weight, for example during a turn, the aircraft will experience an apparent increase in weight or "G" loading and will stall at a higher speed.

The ratio of manoeuvre stall speed to basic stall speed can be calculated as follows: Vs (manoeuvre stall speed) =

L CL max ½ pS

Vs (basic stall speed)

=

W CL max ½ pS

PILOT SIMPLE FORMULA: VsNEW = VsOLD x √Load Factor

An aircraft in a level turn, if the total lift is two times the weight, the square root of two is 1.4, therefore the manoeuvre stall speed will be 1.4 times the basic stall speed. The aircraft will, however, still stall at the same angle of attack.

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THE EFFECT OF AIRCRAFT CONFIGURATION ON THE STALL SPEED

Vs

=

WEIGHT CL max ½ pS

WING AREA (S) An increase in the wing area, such as by using Fowler type flaps will reduce the stall speed. CHANGE IN CL MAX The use of all types of flaps will increase the CL max An increase in the CL max will reduce the stalling speed. THE EFFECT OF THRUST AND SLIPSTREAM AT THE STALL The effect of thrust and slipstream at the stall is two fold: As the aircraft approaches the stall, it has a high angle of attack. The thrust vector can be resolved into a vertical and horizontal component. The vertical component of thrust acts parallel and opposite to weight and thereby assists lift. In this process, the stalling speed is reduced.

The effect of high speed airflow from behind the propeller is to re-energise the boundary layer over the wing, thus delaying separation. In the process, the stalling angle of attack is increased. It must be remembered that this effect is local and restricted to the inboard sections of the wing where the slipstream passes. The outboard section of the wing (wing tips) now stand the risk of stalling first, (normal stalling angle of attack), which could lead to a wing drop and incipient spin if one wing stalls before the other.

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WING TIP STALLING Good aircraft design dictates that an aerofoil be designed to stall from root to tip for the following reasons: (a) When the root section stalls first, the turbulent airflow behind the root will strike the tailplane and warn the pilot of the impending stall. (b) The ailerons, situated on the outboard ends of the wings, will remain effective up to higher angles of attack. (c) The danger of a large rolling moment when one tip stalls before the other is reduced FACTORS PROMOTING TIP STALLING Power/Slipstream Slipstream from behind the propeller re-energises the boundary layer over the inboard sections of the wing, thereby delaying separation. The effect is local; the inboard section of the wing will stall at a higher angle of attack, but the tips will stall at the same angle of attack, i.e. sooner. Flaps Flaps increase the camber of the wing at the inboard sections. In doing so , the larger deflection of the downwash will incline the relative airflow vector, effectively reducing the angle of attack and delaying the stall. The effect is local to the inboard sections of the wing where the flaps are, thus the wing tips will be inclined to stall first.

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FACTORS REDUCING WING TIP STALLING Washout The term washout describes an aerofoil design where the angle of incidence of the aerofoil reduces from root to tip. For any given flight conditions, the root will always be at a higher angle of attack than the tip and thus will stall first. Change of Aerofoil Section This describes a change in camber of the aerofoil from root to tip. The more highly cambered an aerofoil the more the downwash behind the aerofoil. This means that the relative airflow will be deflected to a greater degree and the effective angle of attack will be reduced. By having greater camber wing tip, the wing tip stall can effectively be delayed until after the root stall. Root Spoilers Root spoilers are small triangular metal strips attached to the inboard leading edges of the wing. They disturb the streamline flow of air over the wing root causing an earlier transition from laminar to turbulent airflow, thus causing the wing roots to stall sooner than the wing tips. Slats and Slots The effect of slats is to decrease the angle of attack at the outboard section of the wing, (similar to leading edge flaps), thereby delaying the stall at the tips. The effect of slots is to re-energise the boundary layer behind the slots at the wing tip, thereby delaying the stall

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THE EFFECT OF ASPECT RATIO ON THE STALL Aspect ratio is the ratio of the span of a wing to its chord, described by the formula ASPECT RATIO

=

SPAN² CHORD

The lower the aspect ratio of a wing (short, but wide), the greater the amount of downwash. behind the wing. The effect of this downwash is to deflect the relative airflow and reduce the effective angle of attack. The attitude of the aircraft wiIl therefore be far higher before the stalling angle of attack is reached, (higher stalling angle - same stalling angle of attack). Swept back and delta wing aircraft have low aspect ratios. THE EFFECT OF ICING AND WING DAMAGE ON THE STALL Icing and damage to an aerofoil have the effect of spoiling the laminar airflow and causing it to become turbulent. This means that separation will occur sooner and the aerofoil will stall at a lower angle of attack. DEEP STALL A swept wing aircraft suffering pitch up may take the angle of attack well past the stalling angle due to inertia effect of the pitching moment. Unlike a straight wing, which will experience a pitch down when stalled, a stalled swept wing is likely to have a steady nose up pitching moment from the only remaining attached flow near the wing root. If a conventional low wing tailplane is fitted, the tailplane/elevator can be used to pitch the nose down and regain stable flight. However , if a high tail is fitted, the separated flow from the wing may envelope the tailplane thus making it ineffective. In this case no restoring control is available and it may not be possible to recover. The ensuring rate of descent will further increase the angle of attack thus perpetuating the deep stall condition. Swept wing aircraft in the configuration below are all fitted with stick pushers which provide an undemanded nose down control at a very high angle of attack. This prevents the aircraft from achieving very high angles of attacking in the 1 st place.

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SPINNING With reference to the stalled aircraft if the aircraft starts to roll there will be a component of flow induced tending to increase the angle of attack of the down-going wing and decrease the angle of attack of the up-going wing. The cause of the roll may be either accidental (wing-drop) deliberate (further effect of applied rudder) or use of aileron at the stall.

The effect of this change in angle of attack on the CL and CD curves is that the "damping in roll" effect normally produced at low incidence is now reversed. The increase in angle of attack of the down-going wing decreases the CL and increases the CD. Conversely, the decrease in angle of attack of the up-going wing increases the CL and decreases the CD. The difference in lift produces a rolling moment towards the down-going wing tending to increase the angular velocity. This angular acceleration is further increased by the roll induced by the yawing motion due to the large difference in drag.

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The cycle is automatic in the sense that the increasing rolling velocity sustains or even increases the difference in angle of attack. It is emphasized that autorotation like the stall, is an aerodynamic event which is dependent on angle of attack. It is therefore possible to auto-rotate the aircraft in any attitude and at speeds higher than the basic stalling speed. This principle is the basis of many of the more advanced acrobatic manoeuvres AUTOROTATION Autorotation, sometimes called the incipient spin, is an aerodynamic event before the full spin occurs and stabilizes. During the spin, the aircraft is in a stalled condition of flight, simultaneously rolling, pitching and yawing. In the erect spin to the right illustrated, the aircraft is rolling right, pitching up and yawing right. For convenience the direction of the spin is defined by the direction of yaw.

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SPIN LIMITATIONS It is most important to observe any limitations imposed on spinning of a particular aircraft type. These limitations are usually imposed to eliminate any tendency of the aircraft to get into a flat spin. This can result from spinning an aircraft with the centre of gravity in an excessively aft location. The tendency is exaggerated when the spin is allowed to continue for too many turns. For these reasons, the limitations imposed usually apply to centre of gravity aft limits; in some cases it is stipulated that no baggage may be carried in aft baggage compartments when carrying out spins. In some aircraft, the maximum number of turns permitted before recovery is stipulated. INSTRUMENT INDICATIONS DURING THE SPIN 1. Low airspeed, fluctuating around the stalling speed. 2. Maximum rate of turn on the turn co-ordinator in the direction of yaw (spin). 3. Inclinometer shows a skid. (Bank opposite direction to that of the spin.) 4. High rate of descent confirmed on the VSI and altimeter

SPIN RECOVERY  Throttle closed  Full opposite rudder to the direction of the spin (anti-spin moment).  Progressive forward moment of the control column to unstall the wings (anti-spin moment).  Once the spin stops centralize the rudders to prevent a spin in the opposite direction.  Ease gently out of' the dive. Harsh use of elevators at high speed may induce structural aircraft damage or cause a high speed stall to develop.

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Revision Questions 1. A stall warning must be set to operate: (a) at a speed just below stalling speed (b) at a speed above stalling speed (c) at the stalling speed. 2. In a steady turn an aircraft experiences 3g, the stalling speed will be: (a) above the normal stalling speed (b) below the normal stalling speed (c) the same as the normal stalling speed. 3. At altitudes above sea level the IAS stalling speed will be: (a) the same as at sea level (b) less than at sea level (c) greater than at sea level 4. An aircraft wing stalls at: (a) a constant true airspeed (b) a constant angle of attack (c) a constant indicated airspeed 5. A typical stalling angle of attack is: (a) 30 (b) 15 (c) 5 6. With engine power on, an aircraft will stall: (a) at the same speed as with power off (b) at a lower speed than with power off (c) at a higher speed than with power off 7. If the aircraft weight changes by 6% the stalling speed will change by approximately: (a) 3% (b) 12% (c) 6% 8. A fixed spoiler on the leading edge of the wing at the root will: (a) prevent a root stall (b) induce a root stall (c) give a shorter landing run.

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9. At angles of attack above the stalling angle: (a) the lift decreases and the drag decreases (b) the lift decreases and the drag increases (c) the lift increases and the drag increases. 10. A leading edge slat is a device for: (a) increasing the stalling angle of the wing (b) decreasing the drag of the wing (c) decreasing the stalling angle of the wing. 11. Wing tip staIling may be prevented: (a) wash out on the wing (b) wash in on the wing (c) giving the tip a sharp leading edge. 12. Increasing aircraft weight will: (a) decrease the stalling speed (b) not effect the stalling speed (c) increase the stalling speed. 13. A stick shaker is: (a) a high Mach number warning device (b) an artificial stability device (c) a device to vibrate the control column to give a stal1 warning. 14. The stalling speed is determined by: (a) the maximum value of CL (b) the CL for zero lift (c) the CL for maximum L/D ratio. 15. A stick pusher is a device for: (a) assisting the pilot to move the controls against high air loads (b) preventing the aircraft from getting into a stall (c) automatically compensating pitch changes at high speeds. 16. Wing tip stalling may be prevented by: (a) giving the tip a sharp leading edge (b) wash-in on the wing (c) wash-out on the wing.

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17. if weight is increased the stalling angle of attack is: (a) increased (b) decreased (c) the same. 18. With the C of G on the forward limit, the stalling speed would be: (a) lower than with the C of G on the aft limit (b) higher than the C of G on the aft limit (c) the same as with the C of G on the aft limit. 19. If an aircraft is flying close to the stall, and ailerons are operated: (a) a stall could occur on the wing with the down aileron (b) a stall could occur on the wing with the up aileron (c) there would be no effect on stalling. 20. On a highly tapered wing without wing twist the stall will commence: (a) at the tip (b) at the centre of the span (c) At the root. 21. The indicated stalling speed of an aircraft at a given weight: (a) increase with altitude increase (b) is constant at all altitudes (c) decreases with altitude increase. 22. The separation point of the boundary layer: (a) moves forward with increased angle of attack (b) moves aft with increased angle of attack (c) remains constant up to the stalling angle.

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CHAPTER 6 LIFT AUGMENTATION

FLAPS The function of flaps is as follows: In order to ensure that an aircraft can have a reasonably short landing distance, the approach and landing speed must be reasonably low. From the lift formula, it can be seen that the lower the speed, the less the lift. Under these circumstances, the pilot is compelled to raise the angle of attack in order to restore the lift. The situation is undesirable, firstly because raising the angle of attack places the aircraft close to the stall while close to the ground and secondly, raising the angle of attack reduces the forward visibility at a time when it is important to be able to see the runway. The use of flaps solves the problem. Firstly, flaps produce drag which helps to slow the aircraft down and secondly, by increasing the camber of the wing, using flaps increases the lift without having to increase the angle of attack. Flaps are also useful during the take-off. When it is necessary to take-off from a short runway, the aircraft can be rotated at a lower speed, thus using less runway for speed acceleration. The loss of lift due to the lower speed is compensated for by the increased lift from the flaps. When it is necessary to clear a high obstacle after take-off, the aircraft can be climbed at a lower speed and steeper angle, with flaps once again restoring the lift lost due to the low speed. The stalling angle of attack refers to the angle between the chord line and the relative airflow when the aircraft stalls he diagram below illustrates the effect flap has on stalling attitude and angle of attack in level flight.

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The plain or cambered flap, illustrated above, will have the effect of increasing the maximum lift by + 50 %. The stalling angle of attack will be reduced.

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TYPES OF FLAP As mentioned, flaps may be situated on the trailing edge of the leading edge. Both leading edge and trailing edge each have many variations. For this reason a table summarizing the various types is presented

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LEADING EDGE FLAPS Swept wing aircraft operating at high angles of attack are prone to significant increases in induced drag due to wing vortices and flow separation along the leading edge. To prevent this, and to generate more lift at high angles of attack, all swept wing transport aircraft are fitted with some form of leading edge lift augmentation device. Leading edge flaps are one of the options. One common design is the Krueger flap which has a folding nose to vary the effective camber of the leading edge. Alternatively, a drooped leading edge flap may be used. Both these designs are shown. The B747 employs Krueger flaps on the inboard wing section and a slightly different design on the outboard sections which allows the airflow to be re-energized by means of a slot in the flap design. In all cases, leading edge flaps are hydraulically actuated.

LEADING EDGE SLATS For some types of swept wing aircraft, an alternative use of leading edge slats is used to reduce spanwise flow of the boundary layer and increase lift at high angles. The MD DC-10 uses this method. The slats are cambered surfaces which when retracted, lie flush with the upper surface of the leading edge of the wing. They are arranged in sections along the leading edge and operated at low speeds by hydraulic actuators in the same way as flaps. In the extended position, as shown, a slot is formed between the slat and the wing surface which acts as a venturi accelerating the flow and energizing the boundary layer. In this way slats prevent spanwise flow which in turn reduces the tendency to tip-stall.

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CHAPTER 7 FLYING CONTROLS An aircraft has three axes about which it moves. Movement about these axes is achieved with the use of the appropriate controls. All three axes pass through the centre of gravity.

THE AXES OF MOVEMENT THE LONGITUDINAL AXIS The longitudinal axis runs from the nose to the tail of the aircraft. When the control column is moved to the left, the left aileron goes up and the right aileron goes down. When the control column is moved to the right, the right aileron goes up and the left aileron goes down. The wing with the up-going aileron experiences a decrease in camber and a decrease in angle of attack and thus a decrease in lift. The wing with the down-going aileron experiences an increase in camber and an increase in angle of attack and thus an increase in lift. This differential lift causes the aircraft to roll about the longitudinal axis. Damping in Roll When aileron is applied, an unbalanced force is created about the longitudinal axis, and this force causes the mass of the aircraft to accelerate about this axis. The rate of roll is initially zero as the ailerons are applied but the radial velocity of the wingtips will increase proportionally to the length of time that the rolling force is applied. This implies that the response, or number of degrees rotation per second for a fixed deflection, is changing continually. But does this mean that for a fixed deflection an aircraft will roll faster and faster until it is spinning like a bullet? Theory would imply so, but it does not happen due to a phenomenon known as 'damping in roll'. This simply means that as the wings start rotating an additional component of free stream flow is created that changes the relative angle of attack of the wings. The down going’s wing's angle of attack is increased and the up going’s decreased. This reduces the lift differential between the wings. There will come a time when the alteration of the angle of attack of the wings cancels out the lift differential from the ailerons and the forces on the wing will be in equilibrium once more. When this happens a steady rate of roll will exist.

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As rate of roll increases the damping force increases until a steady rate of roll results. In most low performance, conventional aircraft the damping in roll takes full effect well within the control application time (usually within one second). Thus for most of the time of application there is a steady response from the ailerons, and they are then termed rate controls. THE LATERAL AXIS The lateral axis runs from the wingtip to wingtip of the aircraft. When the control column is moved forwards, the elevator deflects downwards. The camber and angle of attack and thus the lift is increased on the horizontal stabiliser. The tail of the aircraft moves up and the nose of the aircraft pitches down about the lateral axis. When the control column is moved aft, the elevator is deflected upwards. The camber and angle of attack decrease and thus lift is decreased on the horizontal stabiliser. The tail of the aircraft moves down the nose of the aircraft pitches up about the lateral axis. Damping in Pitch As illustrated it can be seen that an elevator deflection will cause a rotation that will increase the restoring moment of both tail and main plane. It is for this very reason that the moment produced by the elevator is rapidly damped or equalized as this restoring moment is very large and increases rapidly. Thus an elevator deflection will rotate the aircraft through to a fixed and constant angle of attack. For example, if an elevator produces 500 N of force it will rotate the aircraft until the restoring moment is also 500 N. The angle of attack will remain constant with the forces in equilibrium i.e. there is no further rotation about the axis as there was in the case of the ailerons and roll. The aircraft, because of its increased lift, will then follow a curved flight path, such as a loop.

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Elevator Deflection Moment Balanced by restoring Moments of Wing and Tail, giving rotation to a fixed angle of attack. THE NORMAL AXIS The normal axis passes vertically through the centre of gravity of the aircraft. When the left rudder pedal is depressed, the rudder is deflected to the left. Lift is increased on the right hand side of the vertical stabiliser. The tail of the aircraft moves to the right and the nose of the aircraft yaws to the left about the normal axis. When the right rudder pedal is depressed, the rudder is deflected to the right. Lift is increased on the left hand side of the vertical stabiliser. The tail of the aircraft moves to the left and the nose of the aircraft yaws to the right about the normal axis. WEATHERCOCKING Weathercocking refers to an occurrence which describes an aircraft's movement about it's normal axis, i.e. Yaw. If the nose of an aircraft is yawed to the left, the airflow will strike the right- hand side of the fuselage and vertical stabiliser. This unbalanced force on the right-hand side of the aircraft, behind the C of G will cause the nose of the aircraft to yaw to the right and thus restore the aircraft to flight parallel with the airflow. The restoring force is called the weathercock effect.

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THE EFFECTIVENESS OF CONTROLS The effectiveness of the aircraft's controls depends upon the following factors: THE SIZE AND SHAPE OF THE CONTROL The greater the size or area of a control surface, the greater is the potential in developing a lifting force and therefore, the greater its effectiveness, (Lift = CL½pV²S). THE SQUARE OF THE AIRCRAFT'S SPEED (EAS²) The greater the speed of the aircraft, the greater the speed of the airflow over the control surfaces and the greater the potential for developing a lifting force, (Lift = CL½pV²S), and therefore, the greater the control effectiveness. THE ANGLE OF DEFLECTION The greater the angle through which a control surface is deflected, the greater its change in camber and angle of attack and therefore, the greater its lifting force and effectiveness, (Lift = CL½pV²S). THE LENGTH OF THE MOMENT ARM The distance between where a force acts, (at the control surface), and where the aircraft moves, (at the C of G about the appropriate axis), is called a moment arm. The longer the length of the moment arm, the greater the control effectiveness for any given control deflection. ADVERSE AILERON YAW By definition, adverse aileron yaw is an initial yaw in the opposite direction to the intended roll. With reference to Chapter 3, it is clear that the greater the angle of attack and lift of an aerofoil, the greater the induced drag. When an aileron application is made to roll the aircraft, the down going aileron/ up going wing, experiences an increase in angle of attack and lift and therefore experiences more induced drag. Conversely, the up going aileron/down going wing experience a decrease in angle of attack and lift and therefore experiences a reduction in induced drag.

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This differential drag results in an initial yaw in the opposite direction to roll and is called adverse aileron yaw. Adverse aileron yaw is described as an initial yaw in the opposite direction to roll. The directional stability of the aircraft due to the weathercock effect will oppose the yaw as it occurs and restore the longitudinal axis of the aircraft to flight parallel with the airflow. FACTORS TO REDUCE ADVERSE AILERON YAW DIFFERENTIAL AILERONS For any given control deflection, the up going aileron moves through a greater arc than the down going aileron. This helps to balance out the differential drag caused when the down going aileron experiences more induced drag than the up going aileron.

FRISE TYPE AILERONS In the construction of Frise type ailerons, the ailerons are hinged at a point slightly aft of their leading edges. When deflected, the nose of the up going aileron deflects into the airflow, and the nose of the down going aileron does not. The up going aileron thus experiences more profile drag than the down going aileron. This helps to balance out the differential drag caused when the down going aileron experiences more induced drag than the up going aileron.

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SPOILERS Spoilers are flat plates on the upper wing surface which are activated with aileron application. When ailerons are applied, the spoiler on the side of the up going aileron is extended. The down going wing thus experiences more profile drag than the up going wing. This helps to balance out the differential drag caused when the down going aileron experiences more induced drag than the up going aileron.

AIRSPEED EFFECTS AILERON REVERSAL Wings are not totally rigid and at high speeds the force produced by an aileron deflection can be so strong that the wing twists about its torsional axis. As illustrated this twisting moment causes a reduction in angle of attack thus reducing the effectiveness of the ailerons for a fixed deflection. At a certain speed, known as aileron reversal speed, an application of aileron will cause such a twisting moment that there is no change in lift on the wing in total thus there will be no rolling moment. If the aircraft is flown beyond aileron reversal speed the twisting moment can cause a reduction in angle of attack to a negative value and the aircraft will roll in a direction opposite to that intended.

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LOW SPEED EFFECTS At low speeds it is usually necessary to make larger deflections in order to achieve a desired roll rate. As the wing itself is operating at higher angles of attack at these lower speeds, an application of aileron can cause the aileron deflected downwards to operate very close to the critical angle. At these high angles of attack the following can occur: AILERON SNATCH At high angles of attack the airflow over the aileron may begin to oscillate and separate. This can. result in rapid movements of the centre of pressure on the control surface such that at time the centre of pressure moves past the hinge line of the control, towards the leading edge. This anti-restoring force causes a rapid increase in angle of attack of the aileron, and the C of P moves further forwards, aggravating the problem. The resultant force can be such that the control stick is literally snatched out of one's hand in the intended direction of movement.

AILERON OVERBALANCE On some aircraft one may find that the larger the deflection angle the smaller is the stick force needed to hold the control in that position. This is because the larger the deflection the more effective is the aerodynamic balance e.g. more and more section ahead of the hinge line protrudes into the flow, giving greater balancing. At some stage the control force will lessen to zero and beyond that deflection the controls may begin moving of their own accord, and overbalance may be confused with snatch. With snatch, however, there is no progressive decreasing of control forces before the control stick begins moving of its own accord.

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FLUTTER An important problem associated with aircraft controls is that of "flutter". This can be prevented by mass-balancing the system. On a few low performance aircraft the additional weight necessary to mass-balance the control is added externally. Usually, however, the weight is stored internally, often inside the horn balance, in the leading edge of the control surface on an arm which is inside the appropriate main surface.

It is also of the utmost importance that a spring or servo tab is individually mass-balanced about its own hinge-axis. Control surface flutter arises as the result of coupling between the rotation of a control surface and the bending or torsional motion of the wing, tailplane, or fin. As an example, the next diagram illustrates the case of a wing flexural-aileron flutter in which the motions involved are wing bending and aileron rotation. If the wing is oscillating in bending and the CG of the control surface is behind the hinge line of the control the inertia effects are such that as the wing moves upwards the control tends to rotate downwards. Conversely as the wing moves downwards the control moves upwards and the aerodynamic force produced by rotation of the aileron may assist the motion of the wing. The motion is opposed by the rigidity of the wing and the aerodynamic damping effects, and at low speeds flutter does not occur. As the forward speed is increased, however, the aileron excitation forces increases as the square of the speed, whereas the aerodynamic damping effect only increases linearly with speed. Thus above a certain critical speed flutter will occur. TYPES OF FLUTTER Torsional Flexural Flutter

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Twisting of the wing under load

Torsional Aileron Flutter

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Wing twists due to aileron input (shown below)

Flexural Aileron Flutter

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Aileron lags behind up and down movement of the wing.

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CROSS-COUPLING OF CONTROLS When a control produces a response in a plane other than that desired it is termed a crosscoupling response, and one method of solving the problem is to cross-couple the controls of the plane affected. In the case of adverse yaw the linkage system of the ailerons and rudder can be geared so that each time a movement is made in the rolling plane the rudder automatically moves to cancel out the adverse yaw. USE OF RUDDER DURING AILERON APPLICATION FOR BALANCE Although the above mentioned devices certainly reduce the amount of adverse aileron yaw, most aircraft still require a rudder input in the direction of the intended roll to overcome the effects of adverse aileron yaw completely and keep the aircraft balanced.

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CONTROL BALANCING On larger and faster aircraft, the forces required by the pilot to deflect the flight controls of the aircraft would be enormous. Clearly, some form of assistance is required and this comes in the form of control balancing. Some of the more common forms of control balancing are listed below: THE SETBACK HINGE OR INSET HINGE The net lifting force on a control surface acts through the centre of pressure. The product of this force and the arm between the CP and the hinge line is the moment which the pilot must oppose in order to deflect the control surface. By moving the hinge line aft, the arm and therefore the moment are reduced, requiring less force by the pilot to deflect the control surface. Certainly, the hinge line should not be positioned too far aft. As the angle of deflection of a control surface is increased, its CP moves forward. Should the CP move ahead of the hinge line, the moment would be reversed, causing the control surface to remain in the deflect position with no action on the part of the pilot and the "feel" would be lost.

THE HORN BALANCE When a control surface is deflected, a portion of the leading edge of the control surface protrudes into the airflow on the opposite side of the aerofoil. Airflow striking this position of the control surface assist the pilot to deflect the control surface.

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TAB BALANCES A tab is a small fixed or hinged surface forming part of the trailing edge of a control surface. The general principle of operation of the tab is as follows: When a control is deflected, the tab is deflected in the opposite direction. The large lifting force on the control surface and the short arm between the CP and hinge line form a moment which is balanced by the smaller lifting force on the tab surface and the longer arm between its CP and the hinge line of the control surface. The force required by the pilot to keep control surface in the deflected position is thus reduced or removed altogether. Some of the many types of tab balances are listed below. Fixed Tabs or Adjustable Tabs These tabs cannot be adjusted in flight. After a series of test flights, they are bent to a position which will ensure that no force is required by the pilot to keep the control surfaces in the deflected state which will keep the aircraft in straight and level flight at cruise power and cruise speed.

Trim Tabs Movement of the trim tab is achieved by the pilot in flight, either by means of manually turning a trim wheel in the cockpit which is connected to the tab by cables, or by activating an electrical switch in the cockpit connected to the tab via an electric motor. The trim tab can thus be deflected by the pilot to remove the forces required to move a control surface in any number of aircraft configurations and speeds. CPL ATG CPL DOC 01 Revision : 1/1/2001

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Servo Tabs The controls in the cockpit are directly linked to the tab instead of the control surface. When a control movement is made, the tab is moved in the incorrect sense and the lifting force from the tab causes the control surface to be moved in the correct sense. The principle of the servo tab is that it is far easier for the pilot to move the small tab surface than the large control surface.

Balance Tabs A balance tab is not under the control of the pilot. It is linked to the aerofoil in such a way that when the pilot deflects the control surface, the tab automatically moves in the opposite sense, thereby reducing the force input required by the pilot.

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Spring Tabs With balance tabs, the tab is linked to the aerofoil and its movement is coupled to be proportional to the movement of the control surface. With spring tabs, the tab is linked to the control surface with a spring between the two. In this way, at low speeds, the tab does not become operational until the control surface has been deflected by a certain number of degrees and at high speeds, the degree of tab deflection is small to prevent overbalancing, (the effectiveness of a control is dependent upon its angle of deflection and the speed of the airflow).

SPRING BIAS TRIMMERS With spring bias trimmers, no tab is involved at all. When the trim wheel is moved in the cockpit, it exerts a spring pressure on the cables of the flight controls or the flight controls themselves, to hold them and thus the control surface in the deflected position. VARIABLE INCIDENCE TAILPLANES Where very large trim inputs are required to control an aircraft in the pitching plane, the forces produced by a small tab surface are often insufficient. In this case, the action of the tab is augmented with a variable incidence tailplane. The tailplane's movement is achieved electrically or hydraulically. Its movement is coupled to the movement of the tab and the two will be activated by one control.

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Variable-incidence tailplanes of different design are used on all large transport aircraft with slight differences. The diagram below shows a variable-incidence tailplane which is used mainly as a trimming system. An elevator is also attached which is used for primary flight control in manual flight.

The variable-incidence function is operated by trim controls in the cockpit, by autopilot servo motors as the primary control in automatic flight , and also as a Mach trimmer through the autopilot servomotors. The diagram below shows a more advanced system where the variable-incidence function is used as the primary pitch control for automatic and manual flight as well as all trimming operations. This system is generally referred to as a ‘flying tail’. Elevators are fitted to this system but direct movement of the stabilizer itself, thereby supplementing its control functions rather than serving as an independent pitch maneuvering control surface can only operate them.

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POWERED CONTROLS A power operated system is shown Below. This system the pilot moves the cockpit control (control wheel) which is connected to a control valve. Depending on the amount of valve movement, hydraulic fluid under pressure moves to the appropriate side of the operating jack which is connected to the control surface.

CONTROL ACTUATOR Given the limited control input needed by the pilot, the power control actuator must be capable of achieving the necessary control deflections throughout the aircraft’s flight envelope. This is mainly a problem of hydraulic jack design, considering such factors as piston area and the pressure of the hydraulic supply. If the aerodynamic loads become too great for the jack to oppose, a condition known as ‘jack stalling’ occurs. In such a situation, aircraft maneuverability is reduced. The control system must remain stable and must not be influenced by signals that do not originate from the pilot. to maintain this stability it is essential that the linkage is free from backlash and that the hydraulic system is free from varying pressures. ARTIFICIAL FEEL SYSTEMS The assessment of the handling characteristics of an aircraft is very dependent on the amount of control movement and the control forces or feedback that the pilot experiences through the control linkage. In an aircraft with conventional non-powered controls the forces increase in direct proportion to the square of the aircraft speed. This relationship forms an important part of a pilot’s ‘feel’ for the aircraft and consequently can affect flying performance. In a fully powered control system however, such feedback is not available, the only natural forces being the friction generated when moving the control and control valve mechanism. To prevent aircraft overstress and to provide a neutral datum control position for pilot reference, all powered control system have an artificial feel system. The most common form of artificial feel system is the ‘q’ system where resistance to control movement is directly proportional to dynamic pressure (q). Below it shows a typical hydraulic ‘q’ feel system. Pitot and static pressures are fed to either side of a pressure capsule in much the same way as an airspeed indicator. Movement of the diaphragm is thus a function of dynamic pressure which is in turn connected to a hydraulic servo-valve. The servo-valve then provides metered hydraulic pressure which is equivalent to an amplified value of dynamic pressure. This pressure is fed to a jack which opposes the movement of the pilot’s control and thus provides the feel for the system. CPL ATG CPL DOC 01 Revision : 1/1/2001

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Although the q’ feel system provides a natural feedback to the pilot in proportion to airspeed, the system can be adapted to provide maneuver protection for the aircraft. A common technique which is sometimes used in large aircraft rudder systems is the ‘Vcubed’ method. Instead of providing control stick resistance based precisely on ‘q’ (which of course is based on V-squared), at higher speeds, the resistance increases dramatically as a function of V-cubed. When applied to a rudder system it prevents the application of large control deflections with corresponding large side-slip angles at high speed. Equally, a Vcubed system could help prevent an aircraft overstress at high speed if fitted to the longitudinal control system. Another method to prevent overstress is to increase the control resistance as a function of ’g’ or load factor. This is normally called a G feel system. POWERED CONTROL FAILURE In the event of a failure of the powered control system some form of back-up must be provided. Modern large transport aircraft normally provide some form of redundancy to the system by having parallel hydraulic jacks driven from separate hydraulic systems. Some aircraft even have three distinct hydraulic systems for the primary flight controls. If possible, an aircraft manufacturer will also provide the option of a manual reversion where the pilot can make some movement of the controls mechanically. With very large aircraft this may not be possible under any circumstances due to the very large components involved, and on medium sized aircraft may only be possible once the aircraft has slowed to low airspeed where aerodynamic loads are less. In this situation, one of the powered control systems may he fitted with an emergency hydraulic accumulator which provides sufficient fluid for control operation whilst the aircraft is slowed to a speed where manual control is possible.

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Revision Questions 1.To produce the desired effect, trim tabs must be adjusted: (a)in the same direction as the primary control surfaces they affect, (b)in the opposite direction to the primary control surfaces they affect, (c)depending upon the design of the trim tab controls. 2.Flight manoeuvres are generally divided into four flight fundamentals: (a)aircraft power, pitch, bank and trim, (b)straight and level flight, turns, climbs and descents, (c)take-offs, slow flight, fast flight and stalls. 3. If the aircraft is in an unusual flight attitude and the attitude indicator has exceeded limits, the instruments to rely on first to determine pitch attitude before starting recovery are (a) turn indicator and VSI (b) ASI and altimeter (c) Turn indicator and ASI 4.The most important function of a rudder during coordinated flight is that: (a)it prevents skid, (b)properly applied, it helps to overcome adverse yaw, (c)applying rudder overcomes the asymmetrical thrust of the propeller as the turn is initiated. 5.Application of aileron alone when rolling into a turn will result in unbalanced flight for the duration of the aileron input and will result in (a) sideslip (b) skid (c) either of the above may be correct depending on the direction of the turn

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CHAPTER 8 MULTI ENGINE FLIGHT TWIN ENGINE PERFORMANCE As the title implies, this section deals with aircraft performance that results from the failure of one of its engines. One of the most critical situations in the event of engine failure is during the take-off roll where safety margins are very slim. The problem of engine-out takeoff performance is not covered here since it merges with the study of flight planning performance, particularly the problem of obstacle clearance planning. ASYMMETRIC BALANCE OF FORCES Any multi-engine aircraft that suffers the failure of one of its power plants will need to stabilize in an asymmetric situation, unless all engines share the same thrust line. Clearly, all modern transport aircraft have a potential asymmetric problem because of the differing engine thrust lines. Once one engine has failed, apart from the obvious reduction in total thrust, the aircraft is subjected to a rearrangement of forces along the flight path as shown below.

Notice that the thrust line is now to one side of the main aircraft axis, acting through the thrust line of the one remaining engine in this case. Also, the drag line has moved to the other side because of the increase in drag on the failed engine side. In this example the drag is the result of a wind milling propeller, which even when feathered offers some drag. In the case of a jet engine drag is also experienced as a result of flow through the wind milling fan. Whatever the type of power plant, the thrust line will move toward the remaining engine(s) and the drag on the side of the failed engine will increase. These forces, together with the moment arm that is evident from the diagram will create a yawing moment toward the failed engine. The yawing moment would of course be less on an aircraft with fuselage-mounted engines such as an MD-80, than that experienced on a wing-mounted configuration such as a B737, for a given engine thrust. This is because the moment arm is very different in the two cases. Left uncorrected, this moment will yaw the aircraft at a rate that will be resisted by the aircraft’s directional stability. This yaw will also create a roll toward the failed engine because of the increased lift on the ‘faster’ wing. It is important to under-stand that although the yawing moment is the root cause of the control problem, on modern aircraft it is imperative to control the roll with aileron as well as controlling the yaw with rudder. CPL ATG CPL DOC 01 Revision : 1/1/2001

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Once corrective inputs are made and provided adequate speed is maintained, the aircraft will maintain a constant heading under asymmetric thrust with an infinite number of bank and sideslip combinations, depending on the size of the following resultant forces: • The sideslip force of the fuselage/fin as a result of its natural directional stability (Ssf). • The rudder sideforce generated by the use of rudder (Rsf). • Lateral component of weight due to angle of bank. • Thrust from the live engine(s). • Total drag. And for propeller driven aircraft the following additional forces: • Propeller torque, which tends to roll the aircraft in the opposite direction to propeller rotation. • Asymmetric blade effect, which displaces the thrust line of the engine to the left if the propeller rotates anti-clockwise when viewed from behind, particularly if the aircraft is operating at a high angle of attack. This is due to the effective increase in angle of attack of the down-going blades when the propeller axis is inclined to the aircraft flight path.

Above the diagram shows the arrangement of the main forces (ignoring the forces peculiar to propeller aircraft), for an aircraft maintaining heading with wings level. In this situation the yawing moment of the combined thrust force and sideslip force must be balanced by a large rudder side force. Note that the aircraft is sideslipping at a small angle so that heading and direction of flight are not coincident. Also the skid ball is centered since both lift and weight are aligned on the same plane. The diagram below shows the arrangement where a small angle of bank is used toward the live engine. This reduces the sideslip angle which has the benefit of reducing the sideslip force (ideally to zero) and consequently the amount of rudder side force that must be generated. The rudder side force that must be created by rudder deflection to oppose the yawing moment of thrust is balanced along the lateral plane by a side component of weight (Wsf), which will vary with bank. For this reason, bank must be contained to about 5-10  if optimum performance is to be achieved. In this situation the yawing moment generated by the rudder balances the thrust yawing moment. The skid ball is not centered since lift and weight are not exactly aligned. CPL ATG CPL DOC 01 Revision : 1/1/2001

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MINIMUM ASYMMETRIC CONTROL SPEEDS VMCG minimum control speed ground VMCG is the minimum speed under take-off power and configuration conditions at which it is possible to recover and maintain directional control on the ground following the loss of the critical engine using rudder alone. Critical engine considerations are not applicable for jet aircraft since a failure on either side produces the same magnitude of yawing moment. Propeller driven aircraft however, will suffer slightly more yaw from a failure on one side due to the combined effects of torque and asymmetric blade effect (if applicable). Although the aircraft nose wheel is assumed to be in contact with the ground, the operation of nose wheel steering is not permitted when VMCG is calculated for a given aircraft. In the real case however, nose wheel steering would be used. Because the setting of VMCG depends on the yawing moment of the live engine(s), the speed will vary with airfield altitude and temperature. Flight manual figures of VMCG refer to sea level standard rated take-off thrust situations. VMCG is below the take-off decision speeds (V1).

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VMCA minimum control speed air. As the name implies, this is the minimum speed that directional control can be maintained following the loss of the critical engine in flight. The specification of VMCA assumes the take-off configuration and allows up to 5ْ of bank toward the live engine for the reasons explained earlier. For propeller driven aircraft the matter of the critical engine is important since both torque effects and asymmetric blade effects at the high angle of attack on takeoff need to be controlled. Like VMCG, VMCA is affected by thrust on the live engine and will vary with altitude and temperature. Flight manual figures are specified for a standard day at sea level. EFFECT OF ENGINE FAILURE IN CRUISE AT HIGH ALTITUDE In straight-and-level unaccelerated flight, thrust is equal to drag. At high level, quite often the engines are operating close to maximum thrust for cruising flight. Consequently, in the event of an engine failure, a speed loss is inevitable unless a descent is made to an altitude where the thrust available from the remaining engine(s) is sufficient to sustain engine-out cruising flight. Inevitably the maximum possible cruising level for engine out flight will depend on ambient temperature and aircraft weight. Under certain conditions at high level, e.g. turbulence or lack of attention on the part of the crew, the speed reduction following an engine failure could be rapid, resulting in the aircraft stalling. This is particularly the case when operating at and above FL400, i.e. at levels where the speed range between buffet boundaries is much smaller.

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Revision Questions 1. What procedure is recommended for an engine out approach and landing? a. The flightpath and procedures should be almost identical to a normal approach and landing. b. The altitude and airspeed should be considerably higher than normal throughout the approach. c. A normal approach, except do not extend the landing gear or flaps until over the runway threshold. 2. What criteria determines which engine is the ‘critical’ engine of a twin-engine airplane? a. The one with the centre of thrust closest to the centreline of the fuselage. b. The one designated by the manufacturer which develops most usable thrust. c. The one with centre of thrust farthest from the centreline of the fuselage. 3. What effect, if any, does altitude have on Vmc for an airplane with un-supercharged engines? a. None b. Increases with altitude c. Decreases with altitude. 4. Under what condition should stalls never be practiced in a twin-engine airplane? a. With one engine inoperative. b. With climb power on. c. With full flaps and gear extended. 5. In a light, twin-engine airplane with one engine inoperative, when is it acceptable to allow the ball of a slip-skid indicator to be deflected outside the reference lines? a. While manoeuvering at minimum controllable airspeed to avoid overbanking b. When operating at any airspeed greater than Vmc. c. When practicing imminent stalls in a banked attitude 6. What is the safest and most efficient takeoff and initial climb procedure in a light, twin-engine airplane? ACCELERATE TOa. Best engine-out rate-of-climb airspeed while on the ground, then lift off and climb at that speed. b. Vmc, then lift off at that speed and climb at maximum angle-of-climb airspeed. c. An airspeed slightly above Vmc, then lift off and climb at the best rate-ofclimb airspeed. 7. What performance should a pilot of a light , twin-engine airplane be able to maintain at Vmc? a. Heading. b. Heading and altitude. c. Heading, altitude and ability to climb 50 ft/min. 8. What does the blue radial line on the airspeed indicator of a light, twin-engine airplane represent? a. Maximum single-engine rate of climb b. Maximum single-engine angle of climb. c. Minimum controllable airspeed for single-engine operation. CPL ATG CPL DOC 01 Revision : 1/1/2001

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CHAPTER 9 HIGH SPEED FLIGHT Slow flying aircraft can use highly cambered wing designs without any concern for the effects of air compressibility. Aircraft designed to fly faster then about 300kts need wings designed to cope with the changes in air density and the flow disturbances caused by compressibility effects.

As an aircraft moves through the air, velocity and pressure changes occur which create pressure disturbances in the airflow surrounding the aircraft. At speeds below the speed of sound, air ahead of the aircraft is moved aside by pressure change, which is transmitted ahead of the aircraft at the speed of sound. When an aircraft flies faster than the speed of sound, the air ahead of the aircraft has no advance warning of the aircrafts approach because the aircraft is flying faster then the pressure waves it is generating. These pressure disturbances are actually transmitted in all directions, but at speeds in excess of the speed of sound, the pressure waves stack up in front of the aircraft to form shock wave. The same thing happens at the leading edge of an aerofoil in supersonic flight. It is important to note that compressibility effects are not confined to aircraft speeds at and above the speed of sound. Most aircraft have aerodynamic surfaces, which produce lift by acceleration of airflow, so these surfaces will have local flow velocities faster then the aircraft flight speed. In this way an aircraft can experience compressibility effects and shockwave problems at flight speeds well below the speed of sound. Since the speed of sound is not constant and varies with temperature, it is very useful to define high aircraft speed as a percentage of the current speed of sound. This ratio is known as the aircraft Mach number. All high speed, highflying aircraft are equipped with Mach meters to indicate to the pilot the current Mach number of the aircraft. The mach number of an aircraft can be calculated from the inputs of pressure altitude and indicated airspeed. To obtain the aircraft True Airspeeds (TAS) for a given Mach number, the temperature must be measured since the speed of sound is dependant on air temperature in accordance with the following formula: MACH No. = 38.945 x  K (where K = degrees Kelvin (O c = +273 K))

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FLIGHT AT INCREASING MACH NUMBERS

Flow velocity less than the speed of sound is termed subsonic flow. Flow velocity at the speed of sound is termed sonic flow. Flow velocity greater than the speed of sound is termed supersonic flow. Since it is possible to have both subsonic and supersonic flow existing around an aircraft in flight, it is helpful to define some significant speeds and some distinct flight speeds and some distinct flight speed ranges on the basis of Mach numbers.

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Mfs Free stream Mach No. The Mach no. of the aircraft through the air (not influenced by local flow effects) Local Mach No. The Mach No. of the airflow at a particular location on and aircraft or aerofoil. MCRIT The critical mach number …..The mach number, which produces the first evidence of local sonic flow. MCRIT marks the start of the transonic speed range. Above M CRIT shockwaves form on the wing. MCDR The critical drag rise mach number. The Mach number slightly above M CRIT, which results in significant shockwave, related drag problems as the shockwaves grow in size and move rearwards. Forced Divergence Mach No. The Mach number above MCRIT, which results in large CP changes and pitching tendencies. MDET Detachment Mach Number…The aircraft mach number at which the bow wave attaches to the leading edge of the aerofoil or aircraft nose. All flow is supersonic above MDET. It marks the end of the transonic range. MMO Maximum Operating Mach No…..The highest mach number at which an aircraft may intentionally be flown. Sometimes MMO will lead to shockwave formation, control problems and high airframe stresses. Jet transport category aircraft have over speed warnings, which sound, at speeds in excess of MMO. Normal shockwave a vertical shockwave. Airflow undergoes a reduction in velocity and Mach number to subsonic and a rise in pressure density and temperatures as it passes through the shockwave. Oblique shockwave An inclined supersonic flow shockwave, Airflow undergoes a reduction in velocity and mach number but the flow is still supersonic. There is also a rise in density and pressure of the flow as it passes through the shockwave. Subsonic Speeds (0-M0.75) In this speed range all mach numbers are less than one. All flow is subsonic. The subsonic flight range extends from zero up to aircraft mach numbers of approx. M0.75, depending on aircraft design. Transonic Speeds (M0.75-M1.2) in the transonic range the flow is mixed, part subsonic and part supersonic. The transonic speed range begins at an aircraft Mach number less than 1, since some local flow velocities reach or exceed the speed of sound due to aerodynamic acceleration even though the aircraft as a whole may still be flying slower than the speed of sound. Similarly after the aircraft reaches Mach 1 the transonic speed range extends through to low supersonic aircraft speeds since compressibility effects raise the temperature of the air at stagnation points along the wing leading edges and nose. This results in lowering the local mach numbers at these points compared to the aircraft as a whole. The transonic range includes aircraft mach numbers from approx 0.75 through to 1.20, depending on aircraft design. A bow wave forms ahead of the aircraft but it not yet attached to the leading edge. Supersonic Speeds (M1.2-M5.0) in the supersonic speed range all flow is supersonic, even at the stagnation points. The shockwave is able to attach itself to the leading edge of the aerofoil or aircraft nose. Above M5.0 speeds are termed hypersonic. CPL ATG CPL DOC 01 Revision : 1/1/2001

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VARIATIONS IN LIFT AND DRAG AT HIGH MACH NUMBERS When designing aircraft and aerofoils for high-speed flight, a major aim is to minimize or delay the adverse effects of shockwaves. These effects are a large drag increase with a large loss of lift, airframe buffeting and control problems with large variations in the Centre of pressure, and in the lifting characteristics of the wing. As the shockwaves form and grows the CP moves rearwards generating a nose down pitch tendency (MACH TUCK) which may require the installation of a Mach trimmer which automatically applies back elevator trim to compensate as an aircraft accelerates to simulate a natural sense. The pilot applies forward stick against the mach trimmer as speed increases. Wing taper, aspect ratio, sweepback and fuselage area ruling and can produce major effects on the aerodynamic characteristics of an aircraft in high-speed flight. Of these options, sweepback is the most significant feature for managing compressibility effects. With a swept wing the component of speed perpendicular to the leading edge, which determines the amount of acceleration, imparted to the flow and therefore the pressure distribution and lift characteristics of the wing. In this way sweepback delays the onset of compressibility effects, since a swept wing aircraft is able to fly faster before the top surface acceleration produces shockwave problems. The main advantages of sweepback are: 1. 2. 3. 4. 5.

Delayed compressibility effects increase MCRIT and increase MCDR increased force divergence Mach number reduced magnitude of CL and CD changed with both speed and AoA changes improved lateral stability and a minor improvement in directional stability

The disadvantages of sweepback are: 1. 2. 3. 4. 5.

tendency to stall tip first due to strong spanwise flow at high angles of attack. This can cause a pitch up at the stall as the CP moves forwards and in. Artificial stall warning devices may be required. reduced CL max and higher stalling AoA reduced trailing edge device effectiveness excessive lateral stability which gives rise to undesirable Dutch roll tendencies structural complexity.

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WING SECTIONS

The supercritical wing section features a flattened top surface to reduce top surface acceleration and therefore delay the onset of compressibility effects. The reflex camber or cusp on the under-surface compensates for the loss of upper camber and helps provide lift at low speeds and high angles of attack. DIFFICULITES ASSOCIATED WITH HIGH SPEED FLIGHT AND HIGH SPEED DESIGNS. As an aircraft passes its critical mach number, the formation of shockwaves results in a very large and sudden increase in drag, accompanied by a loss of lift and often nose down trim changes due to the rearwards movement of the CP and possibly reduced tail plane download as flow separation causes turbulent air to blanket the tail plane and elevator if they are in the line of the flow. This turbulent airflow behind the shockwave also may cause severe vibration and buffeting as it strikes other aircraft surfaces. In early high-speed designs, the very high air loads at high speed led to a problem called control reversal. The disturbance of airflow at high mach numbers could cause a wing drop, which the pilot attempted to correct with aileron. The down going aileron on the low wing caused an increase in lift over the rear part of the wing near the trailing edge and the force was so great that the wing would twist around the spar causing the overall reduction in wing angle of attack instead of the anticipated increase. The wing would then drop further leading the pilot to believe that the control sense was reversed. Shockwaves can also lead to loss of control effectiveness by causing the flow separation ahead of the control surface hinge. Large high speed aircraft now have hydraulically powered controls to cope with the high air loads at high speed but even powered controls can be overcome at high speeds if the rate of deflection is too great or the speed too high. When this occurs the control is said to be jack stalling-meaning that air loads are preventing full extension of the hydraulic jack. Some aircraft use spoilers to assist roll control for these reasons.

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DUTCH ROLL Dutch roll is a result of the increasing dominance of the lateral stability over the directional stability of an aircraft at altitude. Highflying fast aircraft are likely to exhibit Dutch roll tendencies at high altitude since directional stability is reduced at altitude due to low IAS Dutch Roll values with high TAS. The lateral stability remains relatively unchanged with altitude so the aircraft has excess lateral stability leading to Dutch Roll, or spiral stability – the tendency to recover from a spiral dive…

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REVISION QUESTIONS 1. a. b. c.

At what Mach range does the subsonic flight range normally occur? Below .75 Mach. From .75 to 1.20 Mach. From 1.20 to 250 Mach.

2. a. b. c.

What is the highest speed possible without supersonic flow over the wing? Initial buffet speed. Critical Mach number. Transonic index.

3. What is the free stream Mach number which produces first evidence of local sonic flow? a. Supersonic Mach number. b. Transonic Mach number. c. Critical Mach number 4. What is the result of a shock-induced separation of airflow occurring symmetrically near the wing root of a swept wing aircraft? a. A high-speed stall and sudden pitch up. b. A severe moment or ‘tuck under’. c. Severe porpoising. 5. What is the movement of the centre of pressure when the wingtips of a swept wing airplane are shock-stalled first?. a. Inward and aft. b. Inward and forward. c. Outward and forward. 6. What is the principal advantage of a sweepback design wing over a straight wing design? a. The critical Mach number will increase significantly. b. Sweepback will increase changes in the magnitude of force coefficients due to compressibility. c. Sweepback will accelerate the onset of compressibility effect. 7. a. b. c.

What is one disadvantage of a swept wing design? The wing root stalls prior to the wingtip section. The wingtip section stalls prior to the wing root. Severe pitch down movement when the centre of pressure shift forward.

8. What is the condition known as when gusts cause a swept wing-type airplane to roll in one direction while yawing in the other? a. Porpoise. b. Wingover. c. Dutch roll. 9. A wing with marked sweep-back a. b. c.

has a low induced drag coefficient throughout the speed range has a high induce drag coefficient throughout the speed range has a reduced induced drag coefficient at the high speed range

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CHAPTER 10 WING AND FUSELAGE CONSTRUCTION STRESS & LOADS ON AN AIRFRAME Loads considered to act on an aircraft come from a number of sources. Most aircraft operating at higher altitudes are pressurised internally giving a rather high load on the aircraft structure , likened to a pressure tank. Structure around doors, windows and other cut-outs become particularly critical in these designs. High Frequency vibration loads have become common place in modern aircraft equipped with turbine engines. Loads peculiar to transonic and supersonic aircraft are the result of shockwave formation. Loads due to buffeting at low speeds are fairly well known and are in the inertia or dynamic load class. Vibratory loads in the frequency range of those produces by piston engines are still present on many aircraft and must be accounted for in a successful design. Stress is defined as any load applied to a unit area of material, and producing a deflection or deformation in the material. This movement of the material is termed a strain. There are tension, compression shear and bending stresses.

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WING CONSTRUCTION

SPARS In order to resist the bending forces imposed on it, an ideal spar is given a certain depth. An example of this is an ordinary ruler, which will flex easily when loaded on the upper or lower surfaces, but is very stiff when a load is applied to the edge. Unfortunately, the modern wing is thin in cross-section, precluding the use of a deep spar. Two, three or more spars are used in the wing to give the necessary strength. The spar usually consists of solid booms at the top and bottom, connected by a thin plate web. Normally these are manufactured as separate items and riveted together, but some spars are made in one piece from heavy forgings, machined to perfect shape. Figure 1 illustrates three typical spar sections.

a) b) c)

Simple plate web and extruded booms. "Fail-safe" spar in which no crack can propagate across the structure Spar machined from heavy forging.

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STRESSED-SKIN

Although some light aircraft still have parts of the airframe covered in fabric, most aircraft are metal clad. In subsonic aircraft, the wing skeleton of spars and ribs are covered with a light alloy skin. This is riveted to the framework and is designed to stiffen the wing by taking some of the load. This type of construction is known as "stressed-skin" and produces a relatively strong wing without too large a weight penalty. The wing can withstand twisting or torsion loads and is usually strengthened by the addition of span-wise stringers to withstand the bending of flexure loadings.

MACHINED SKIN

The faster the aircraft flies, the greater the rigidity of the structure is required. To achieve this, the stressed-skin of the slower aircraft is replaced by a machined skin. This consists of a skin that is manufactured from a solid billet of metal. The metal is milled away by machines so that in its final form, the contour of the wing is very accurately reproduced, together with the necessary strengthening buttresses and ribs. Altogether up to 90% of the original metal will be cut away, leaving a structure that is extremely strong but light in weight. The panels so produced are joined together to form a rigid, strong wing.

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In this form of construction, the skin of the upper and lower surfaces of the wing joins the front and rear spars rigidly together in the form of a box. To the front is attached the leading edge, and to the rear, the aileron and flaps. To increase the load carrying capacity of the skin between the spars, it is common to corrugate it, and then cover the corrugations with a thin sheet. This form of construction is much used and a variation of it, which has a number of spars one behind the other forming a series of boxes, appears particularly suited to aircraft with low aspect ratios.

D-SPAR CONSTRUCTION The front spar, which takes most of the bending load, is placed as near as possible to the point of maximum thickness of the wing, and the skin of the leading edge is rigidly attached to it, to form a D-shaped tube, which takes nearly all the torsional stresses of the wing. CONTROL SURFACES For speeds up to 300-350 kts, fabric-covered ailerons, built up on a spa and ribs are usually satisfactory. Higher speeds demand a rigidity that can only be obtained by a stressed-skin covering built up in much the same way as a D-spar wing. Additional stiffness can be obtained by employing longitudinal fluting of the skin (i.e. spaced corrugations) in this design most of the ribs can be eliminated. BRACED WINGS This design, feature is used almost exclusively in small high wing aircraft. The bracing struts, running from the fuselage to a point about half-way along the wing, relieve the spars of much of their vertical load, and anchor them in tension. The designer, can therefore, save weight in the wing, but because of the additional drag, this form of construction is limited to aircraft with a low top speed.

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Fuselages present basically a simpler structural problem than wings. A fuselage is usually built up from a framework of frames, or transverse members, joined by longitudinal girder members or "stringers". The whole framework being covered by stressed skin. The shape of the cross-section of the fuselage will vary with the job that the aircraft has to perform. Pressurized transport aircraft have circular cross-sections; this has been found to be the most suitable shape to resist the differential pressures. Fighter aircraft have a cross-section of minimum area, the shape is determined by the engine it encloses and/or the radar, armament, or other equipment carried. Light aircraft usually have a square fuselage; this being an easy and strong shape to construct.

"SAFE-LIFE" AND "FAIL-SAFE" In concluding, mention must be made of the structural concept of "safe- life" and "fail-safe". Structure designed for a given safe life, is one in which actual testing of similar structures has enabled the designer to calculate the minimum flying hours before structure failure will occur. This figure is then the "safe-life" for that particular structure. A "fail-safe" structure is one in which, by duplicating primary structures, an alternative path is available for a load. Therefore, if one member fails, the remaining structure can carry the load for a limited time. In some cases this will involve an extra weight penalty, but often the standby part can justify its existence by performing some separate task. An example of this is the window of a pressure cabin, which consists of two layers of glass with a sandwich of dry air between. Normally, the pressure differential is supported by the inner layer, but should this fail then the outer layer can be made to take the load.

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CHAPTER 11 HYDRAULICS Hydraulic systems are a means of transmitting power through liquid filled pipes, due to the fact that liquids, being incompressible, are able to transmit a force applied at one end of a pipe to the other end. Hydraulics can move large load s by applying pressure over a large area. ADVANTAGES Hydraulic systems have numerous advantages over their mechanical counterparts in that hydraulic systems can be designed to: i)

Provide smooth and steady movement.

ii)

Provide hydraulic power which is confined to pipelines and components, and therefore much of the strengthening of the aircraft structure is not required as is the case with mechanical systems.

iii)

Provide greater transmission of power for weight of equipment.

iv)

Simplify installation in that pipelines can be easily bent to go around obstructions.

v)

Provide variation in speed of operation without the use of gearing as would be required in a mechanical system.

vi)

Easily obtain its power to drive pumps etc., from the aircraft engines.

vii)

Relieve the pilot of effort to operate the cockpit controls by requiring minimal force to move the switches and levers to operate hydraulic controlled systems.

HYDRAULIC FLUID Hydraulic fluids generally are developed to provide the following requirements: i)

Be free flowing at all operating temperatures.

ii)

Have low freezing and high boiling points.

iii)

Not affect or be affected by the materials in the components.

iv)

Have good lubrication qualities.

v)

Not deteriorate or form sludge.

vi)

Have a high flash point.

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Hydraulic Fluids DTD 585 This is a widely used fluid that is coloured red, its mineral based and therefore requires the system to be fitted with synthetic rubber seals. CASTOR or VEGETABLE These as the name suggest are made from natural compounds, thus require only rubber seals in the system and are generally coloured yellow. SKYDROL Another popular fluid that is coloured blue or purple, and due to its ingredients requires the system to have butyl rubber seals. The idea of a hydraulic system The word hydro is based on the Greek word for water. A study of hydromechanics is the study of how the characteristics of water and other liquids can be used to do useful work, particularly mechanical work such as the movement of a mechanical component. The principles of hydromechanics are employed in every modern aircraft in the form of hydraulic systems which are used to power mechanical systems such as wing flaps, spoilers and flight control systems. A sound knowledge of the basic operating principles of hydraulic systems is essential for pilots.

In such a system, power is transmitted as fluid pressure within a given volume of liquid trapped in the pipes and components of the system. The pressure is then converted to a mechanical force to move a component. If liquid is trapped under pressure within any container, the pressure exerted by the liquid on the surfaces of the container is the same in all directions. In a closed container there is no resultant mechanical movement, but if one side of the container is movable, such as a piston within a cylinder as shown, then movement will occur.

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The movement of the piston will depend on the force that is applied on the piston as a result of fluid pressure. The force will be equal to the fluid pressure (P) multiplied by the area over which the pressure is applied it’s the area of the piston (A). Thus the resultant mechanical force F = P x A. This concept is easy to understand but needs to be reversed to fully appreciate the benefits of hydraulic power. Since liquids are incompressible, it follows that applying a force (F) from the piston of area (A) on the liquid will result in a pressure (P) being applied through the liquid. The greater the force, the greater the pressure experienced through the liquid. Below shows the basis of a closed hydraulic system where a small piston applies a force over area. This results in a pressure P being transmitted through the liquid. At the other end of the system is a larger piston of area. Since the pressure experienced on the face of this piston is the same as that generated by the smaller one, it follows that the resultant mechanical force on the larger piston is equal to the pressure P multiplied by area. Thus the force F2 is larger than the original force F1. In effect, the force has been magnified by the ability of the system to transmit the pressure P. Clearly, the success of the system depends on the ability of the liquid to transmit the pressure, and this in turn depends on the liquid being incompressible. It is this feature of liquids that allows the system to work in the first place. Compressible fluids such as gases (air) are not suitable for this type of system since they cannot transmit the pressure to the same degree.

FLUID FLOW THROUGH AN ORIFICE This information is of importance when understanding how a restrictor, or similar valve, performs. As can be seen below, if you consider a horizontal pipe through which fluid is flowing, where at some section there is an orifice, then experiments have shown that a steady motion is set up in the fluid. These streamlines converge toward the orifice, where the velocity of the fluid increases. After passing through the orifice, the streamlines continue to converge until the point B is reached. This point is known as the 'Vena contracta', and at this point the fluid will expand outwards into the full diameter pipe. The orifice constitutes a sudden contraction of a pipe and will cause a loss of pressure due to the sudden contraction itself. The total reduction in pressure is not only due to the contraction, but is also due to the sudden expansion after the vena contracta is reached.

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THE HEATING EFFECT OF LEAKAGE With a hydraulic shock absorber, the energy to be absorbed is first given to the fluid in the shock absorber, and this fluid is forced through small orifices where the energy is converted into heat energy. In this case the heat effect is deliberate. Again, in an aircraft hydraulic system, when the pump has cut out it takes fluid from the reservoir and circulates it round a closed circuit. The fluid, after leaving the pump, is under pressure, and hence has some energy. This energy is dissipated when the fluid reaches the reservoir, and causes the reservoir temperature to rise. If there were no heat losses the temperature would rise indefinitely, but in practice the reservoir temperature settles down to a steady value. A leakage in the system will cause the fluid temperature to rise if it is an internal leak. HYDRAULIC SYSTEM TYPES In general terms, there are two main types of hydraulic system which are sometimes known by some authorities as: i)

Self Idling.

ii)

Non-self Idling.

Modern hydraulic systems used on aircraft operate at a constant pressure, and the above type names reflect on the ability of the pump to off load or idle when the system has reached maximum normal working pressure. The Self Idling Type of system has a pump fitted which is capable of controlling the system pressure within the pump itself, hence the name Self Idling. The Non-Self Idling system uses a pump which delivers a constant flow and requires other components to be fitted to the system to control the system pressure. In the following paragraphs a supply system with each of the types of pump fitted will be examined.

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SELF IDLING TYPE SUPPLY SYSTEM (VARIABLE VOLUME) This type of hydraulic supply system generally comprises the following components, some aircraft however, may require variations in the layout or components used. The following displays all the basic requirements that may be found in a supply system of this type. RESERVOIR The reservoir is normally situated in the aircraft where ease of access may be gained for servicing purposes, such as checking the fluid level, or topping it up. Wherever possible it is also positioned so it is above the level of the pump to ensure it has a head of pressure, in order that it may provide a positive fluid supply to the pump. In situations where the reservoir cannot be located above the pump, then the reservoir is pressurised, in order to provide a head of pressure. The reservoir must also be large enough to provide a reserve of fluid, to compensate for minor system leakages, allow for jack, or actuator, ram displacement, and the return fluid from the jacks or relief valves and other such components. In some systems the reservoir must also be large enough to ensure an adequate volume of fluid is maintained in the total system, for temperature control purposes. Under normal operating conditions, fluid is drawn from the reservoir by the engine driven pump via a filter.

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LOW PRESSURE FILTER This filter, situated between the reservoir and engine driven pump, is sometimes known as a suction filter. It is situated at this point in the system to ensure that any foreign matter which has collected in the reservoir will be removed from the fluid before it enters the pump.

ENGINE DRIVEN PUMP All normal operational supply functions are provided by the engine driven pump which is driven by the aircraft engine, normally via a gearbox, and in a self idling system, or variable volume system, when a selection of a service is made the pump moves to the 'on load', or 'on stroke', position and starts to pump a flow of fluid to the selected service. In the example that flaps are selected 'down', when the flap actuators reach the end of their travel, pressure will build up in the system, when the pressure reaches the system maximum normal working pressure a relief valve assembly within the pump, or similar device, will cause the pump to move to the idle mode and stop delivering further fluid that passes through it, for this reason, when the pump is in the idle position, a small amount of fluid is still passed through the engine driven pump for such lubrication purposes. As can be seen this type of pump has three connections, that is, the inlet from the reservoir, to the reservoir. On some older type pumps when the pump is not supplying a selected service and maximum normal working pressure has been reached, the pump continues to supply a full flow of fluid, but that flow is directed back to the reservoir. On modern pumps however, the return line to the reservoir allows small quantities of fluid to be returned to the reservoir, to maintain a constant temperature and viscosity within the pump.

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GROUND SERVICING CONNECTIONS As can be seen ground servicing connections are provided in each of the pipelines, to and from, the engine driven pump. These are normally of a 'quick release' type and allow servicing functions to be carried out by connecting a ground, or external, source of hydraulic fluid flow from a type of servicing trolley. NON-RETURN VALVES As can be seen below, a non-return valve, or check valve, is fitted after the engine driven pump. This is fitted to prevent fluid, being pumped by the hand pump circuit, flowing back through the engine driven pump instead of to the circuit.

PRESSURE SWITCH It is essential that a warning device is fitted to the supply system. To indicate to the pilot that a fault exists in the system. This normally takes the form of a warning light, usually coloured red, which illuminates when there is a reduction in hydraulic pressure below a certain value. The warning light is activated by a pressure switch, which is usually located in the position. On some aircraft an audible warning system is also provided, and is also activated by the pressure switch. HAND PUMP On some aircraft a hand operated pump is provided for use in an emergency, in the event the engine driven pump fails. In some cases the hand pump may also be provided with a separate supply of reserve fluid, in order to operate some services in such circumstances. On most modern larger aircraft, this type of emergency arrangement would be inadequate, and so a second, or standby, engine driven pump is provided. Such emergency systems will be explained later. On most larger modern aircraft a hand pump is provided to assist with servicing operations, and is normally a double acting type, that is to say, it provides a flow of fluid continuously as fluid is displaced by the action of the pump. HAND PUMP PRESSURE RELIEF VALVE As with the engine driven pump, the hand pump must not be allowed to over pressurise the system, and so, when maximum normal working pressure is reached, the pump must be 'off loaded'. In the engine driven pump, the variable volume type, the pump is moved to an idle mode. In the case of the hand pump, a pressure relief valve is provided to allow the pump to idle when maximum pressure is reached. The pressure relief valve is usually fitted within the hand pump assembly itself, it is shown as a separate component for simplicity.

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HAND PUMP NON-RETURN VALVE A non-return valve, or check valve, is also fitted after the hand pump to prevent fluid from the engine driven pump passing back through the hand pump.

HAND PUMP FILTER A simple low pressure suction filter is fitted between the reservoir and the hand pump to provide clean fluid to the system. NON SELF IDLING TYPE SUPPLY SYSTEM (CONSTANT VOLUME)

The diagram above shows an example of a non self idling supply system which is essentially the same as the self idling type, with the exception of the following features.

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ENGINE DRIVEN PUMP This type of pump provides a continuous flow of hydraulic fluid to the system, and has no direct control over the system pressure. Note in the system diagram there are only two pipes connected to this type of pump, the inlet from the reservoir, and the outlet to the system. The pumping capacity of this type of pump for a given rpm is fixed, hence the terms constant volume, or fixed volume pump. To control the system pressure this type of pump relies on an automatic cut out valve. AUTOMATIC CUT OUT VALVE In simple terms, the automatic cut out valve is a sensitive pressure relief valve which is fitted after the pump, and when system pressure is reached, the automatic cut out will close the pipeline to the system and return, or redirect, the fluid from the pump back to the reservoir. ACCUMULATOR Stores -

fluid under pressure: dampens pressure surges supplements pump output during high volume demand allows for expansion and contraction of fluid with temperature permits faster operation of services provides emergency source of pressure pump fails maintains system pressure with small leaks absorbs shock waves

The accumulator consists of a chamber with a flexible diaphragm, on one side of which is fluid at system pressure, the other side has compressed air at a predetermined pressure. As hydraulic pressure builds up the gas is compressed until fluid and gas pressures equalise at normal system pressure. At this point the pressure regulator r valve unloads the pressure pump and it commences to idle. System pressure is then maintained by the accumulator.

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EMERGENCY OPERATION OF HYDRAULIC SERVICES Emergency operation of a hydraulic system, in the event of normal supply failure, may be provided in a number of ways. The following are some of the more common methods of emergency operation that are employed in aircraft hydraulic systems at the present time. It should be noted that any one aircraft may adopt a number of emergency operating methods within its hydraulic systems. HAND PUMP In the event that the engine driven hydraulic pump fails, on some aircraft with relatively small hydraulic systems, a hand operated pump may be provided to operate certain services. On aircraft with very large hydraulic systems, and large volume hydraulic jacks, or actuators, a hand pump would be totally inadequate. When hand pumps are used for emergency use, the hydraulic reservoir has a stack-, or stand-pipe fitted in the bottom. The normal, or engine driven pump obtains its supply of hydraulic fluid from the base of the stack pipe. This is illustrated. As can be seen, should a leak occur in the system which causes the fluid level to fall in the reservoir, it will only fall as far as the top of the stack-pipe. The fluid that now remains in the reservoir can only be provided by the hydraulic hand pump, in other words, the hand pump may be provided with its own reserve of emergency fluid.

On most large aircraft the hand pump is connected to the hydraulic system for servicing purposes only, and is often located in such a way that it cannot be operated in flight.

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DUPLICATED SYSTEMS Hydraulic supply systems on modern aircraft are normally duplicated, and on some aircraft may be triplicated. In the case of a duplicated system, in the event of one pump failing, the remaining pump will continue to supply the hydraulic circuits with fluid under pressure. In such a system a shuttle valve may be provided, which in the event one pump failing will close off the supply line from the unserviceable pump, thus preventing any loss of fluid, or fluid pressure. In this arrangement Number Two Pump is being used as a standby pump. In this case only the pumps are really being duplicated. In other systems the entire supply system is duplicated.

.

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RAM AIR TURBINE Ram Air Turbine Units are used on many modern aircraft as a means to drive certain components in an emergency. In this case the Ram Air Turbine (RAT) is used to drive an emergency hydraulic pump. When the hydraulic supply system fails, the RAT automatically lowers into the airflow under the wing, or fuselage, and a small propeller, or turbine driven by the airflow, drives an emergency hydraulic pump.

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Revision Questions 1.

DTD 585, hydraulic fluid is: a) b) c)

2.

A mineral based hydraulic fluid will require all components to be fitted with: a) b) c)

3.

maintains a constant viscosity at all working temperatures maintains a constant viscosity at a specific working temperature will not form sludge or thicken when it is stationary in the system.

A variable volume pump: a) b) c)

8.

a flap circuit to reduce the rate of movement in both directions an undercarriage circuit to reduce the speed of operation on lowering a speed brake to ensure it moves out quickly and moves in slowly.

A visco-static fluid is one which: a) b) c)

7.

prevent cavitation at the pump provide, in an emergency, a supply of fluid for the pump assist in damping out system pressure fluctuations.

A one-way restrictor valve may be fitted to: a) b) c)

6.

accumulator charge pressure is too low the engine RPM is too low a pressure relief valve is stuck slightly open.

The purpose of an accumulator is to: a) b) c)

5.

natural rubber seals butyl rubber seals synthetic rubber seals.

In a hydraulic system, when the engine is at idle RPM, the system pressure fails to rise above two thirds maximum normal working pressure, the probable fault is: a) b) c)

4.

blue light straw red.

maintains a constant volume of fluid to the circuits at all times maintain a constant pressure in the system. maintain a constant temperature throughout the system.

Excessive pressure due to thermal expansion in a closed circuit may be relieved by a: (a) Flow control valve (b) Pressure reducing valve (c) Pressure relief valve.

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9.

Excessive system pressure fluctuations may be due to: a) b) c)

10.

Most modern hydraulic reservoirs are pressurised to: a) b) c)

11.

allow fluid pressures to become excessive reduce the emergency source of hydraulic power reduce the amount of fluid stored in the accumulator

Whilst operating a hydraulic system, the hydraulic pressure dropped abnormally low, but slowly returned to normal after activation of the system was completed. The most likely cause of this would be a) b) c)

15.

pressure relief valve accumulator non-return valve

A lower than specified accumulator gas pressure may a) b) c)

14.

pump hydraulic fluid from the accumulator into the reservoir maintain system pressure in flight provide an alternative, in an emergency to the engine driven pump

The component in a hydraulic system that will be set to the highest pressure is the a) b) c)

13.

eliminate cavitation in the pump supply to provide pressure for emergency use to eliminate hydraulic hammering.

The hand pump in a hydraulic system is used to a) b) c)

12.

high accumulator charge pressure low accumulator system pressure low accumulator charge pressure.

low accumulator gas pressure a leak in the component actuator low fluid level

If fully extended hydraulic flaps start to re tract when the aircraft is below Vfe, a likely cause is a) b) c)

a leak in the flap actuating cylinder low hydraulic fluid level expansion of hydraulic fluid due to high temperature

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CHAPTER 12 LANDING GEAR When landing, the undercarriage is required to absorb the impact of the landing at touchdown by means of the following: a)

A shock absorber to absorb compression loads.

b)

A drag strut to absorb drag loads.

c)

A side load strut to absorb side loads.

“ You know you’ve forgotten to lower the undercarriage when it takes full power to taxi back to the terminal”

RETACTABLE UNDERCARRIAGE UNIT In order to reduce drag when the aircraft is in normal flight, most undercarriages are retractable, which is normally achieved by hydraulic actuation. Most main undercarriages are retracted sideways into the wing, or the bottom of the fuselage adjacent to the wing root. Some high winged aircraft retract their undercarriages into special fairings at the base of the fuselage. The nose undercarriage unit is normally retracted forwards or backwards into a compartment near the nose of the aircraft. Where space in the nose, or front fuselage area of the fuselage is a problem, the undercarriage may be retracted sideways. Below shows a number of examples of retraction.

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FIXED UNDERCARRIAGE UNIT Some undercarriage units, particular on light aircraft, may be non-retractable or fixed. The undercarriage must be attached to strong points of the aircraft which are capable of withstanding the considerable impact loads experienced when landing; to this end, the main undercarriage units are usually attached directly or indirectly to the main spar.

The nose undercarriage is normally attached to strengthened frames adjacent to the nose of the aircraft. A similar method is employed on tail wheeled aircraft in that a strengthened area of the rear fuselage is used.

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SHOCK ABSORBER The shock absorber has two primary functions, the first to absorb the compression loads imposed by the impact of landing, and the second to dampen the tendency to recoil, that is to say, the shock absorbers tendency to rapidly extend again immediately after absorbing the landing impact. The shock absorber must also be designed to withstand considerable bending loads, which it will also experience when landing, taxying, and in particular when the aircraft is being turned when on the ground. The shock absorber consists primarily of two cylinders which slide one inside the other much like a telescope. To prevent one cylinder rotating inside the other, torque links or toggles are used. Below shows an example of a set of Torque Links or Toggles as used on a shock absorber unit.

One some lighter types of aircraft the inner cylinder of the shock absorber is prevented from rotating within the outer cylinder by splines on the cylinders.

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WHEEL ASSEMBLIES The wheel assembly consists of a wheel and tyre, and usually housed within the main undercarriage wheels in the brake unit. The number of wheels employed on a particular aircraft is dictated by the physical wheel loading and the type of runway to be used. It is the function of the wheel and tyre assemblies to spread the load of the aircraft over a given area, to produce an individual wheel loading that is acceptable within the strength limits of the runway. Hence large heavy aircraft tend to have a greater number of wheel assemblies. The tyre size and shape is also important in this function, in that a large balloon type of tyre will again spread the load over a greater area. At the same time, consideration must be made to the space available for stowage of the undercarriage units, and as a result often a compromise must be reached in design. This in some cases has led to complex undercarriage trim systems in which, in simple terms, the assembly is physically folded prior to retraction.

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UNDERCARRIAGE POSITION INDICATION

The most common form of undercarriage position indicator is a system of lights. Undercarriage down and locked down is shown by a green light, or lights, normally one light for each undercarriage unit being used. On a three undercarriage arrangement three green lights will be illuminated, one for the nose or tail undercarriage, and one for each main undercarriage. When the undercarriage is in motion and/or unlocked, the green lights are replaced with the illumination of red lights. When the undercarriage is fully up and locked up all lights are extinguished. BASIC UNDERCARRIAGE OPERATION DURING RETRACTION Most modern undercarriages are retracted by hydraulic actuation, a small number of older types use pneumatics (compressed air) to achieve this function. In most cases when the undercarriage is selected up, the sequence followed is as follows: a)

On selection of undercarriage up, the undercarriage door opens.

b)

Undercarriage down locks are withdrawn.

c)

Undercarriage position lights change from green to red.

d)

Undercarriage retracts.

e)

Undercarriage uplocks engage, door closes, door locks engage and lights change from red to extinguished.

The reverse takes place when the undercarriage is selected down. On most undercarriage systems the undercarriage doors close again after the undercarriage has extended, to reduce aerodynamic drag.

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UNDERCARRIAGE EMERGENCY OPERATION In the event of hydraulic system failure, the undercarriage may be lowered in an emergency by either a back up, or standby, hydraulic system or more commonly, by the use of compressed air which is stored in cylinders in the aircraft for such a purpose. Selection for emergency lowering of the undercarriage is normally by a separate emergency selector.

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WHEEL BRAKE OPERATING SYSTEM BASIC WHEEL BRAKE HYDRAULIC SYSTEM

Shown is a simple wheel brake system as used on many light aircraft. The system comprises a hydraulic supply, normally provided by an engine-driven pump, which, on entering the brake system, passes through a non-return valve to the brake control valve. The non-return valve creates a safety situation, in that fluid under pressure, which has entered the system cannot flow back into the hydraulic supply, or be influenced by the operation of other hydraulic sub-systems.

After passing through the non-return valve, the fluid is directed through the pipelines to the brake control valve. The brake control valve is operated by the pilot, by foot pedals situated on the rudder bar, or by a brake lever which may be located on the control column. Most modern wheel breaks are operated by foot pedals. In the system shown in Foot Pedal Operation is employed. Operation of the foot pedals, or brake pedals, controls the brake control valve.

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The brake control valve performs a number of functions: i) ii)

iii)

To reduce the hydraulic supply system pressure to a lower value, by a pressure reducing valve located within the brake control valve. Generally, supply system pressure would be too high, which would cause the braking to be too harsh. The second function the brake control valve provides, is the ability to employ progressive braking on the landing run. Progressive braking is the progressive selection of increased brake pressure as speed reduces on the landing run. By having the ability to vary the brake pressure as required, better control can be achieved. The third primary function the brake control valve provides, is the ability to use differential braking when required. Differential braking is the application of separate, or individual brakes in order to achieve steering through the wheel brakes. This practice is widely employed on lighter types of aircraft. Larger aircraft tend to be equipped with nose wheel steering.

When brakes are selected fluid is directed to the wheel brake units. The degree of braking required is achieved by the amount of pressure applied to the brake pedals when they are depressed. The further the pedals are depressed the greater the degree of braking achieved. Progressive braking can be obtained by progressive depression of the pedals. Differential braking can be obtained by application of one brake or the other as required. THE BRAKE ACCUMULATOR During normal brake operation when brakes are selected, an instant supply of fluid under pressure is required; this is initially provided by the brake accumulator. If the supply system provided the sole supply of hydraulic fluid under pressure, there may be a delay while the main supply system pump builds up sufficient pressure to operate the brakes. Such an important system as the wheel brakes cannot rely solely on pump supply. The main supply pump will continuously recharge the accumulator with system fluid pressure, but in each case of brake application, the accumulator will supply the initial application of the brakes, in other words, provide the initial impetus. In the event of main system supply failure, the accumulator will provide a source of hydraulic fluid under pressure for emergency operation of the wheel brakes. The accumulator is so designed to provide sufficient pressure in an emergency, to operate the wheel brakes for a complete landing run plus a reserve. OPERATION OF THE BRAKE CONTROL VALVE The brake control valve is operated on most modern aircraft by servo pressure. The servo pressure is generated by the force being applied to the brake pedals, which is turn, operates a master cylinder, converts the pressure energy back into mechanical energy, and operates the brake control valve. It must be noted the fluid within the servo system is totally independent of the main supply fluid, and to this end a filler point is usually provided at the master cylinder to top up the servo system when required.

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BRAKE OPERATION Brakes are applied by depression of the brake pedals, by which, via the servo system, the brake control valve is operated, initially directing fluid from the brake accumulator to the wheel brake units. As the fluid passes through the control valve, the system pressure is reduced to a lower value. Fluid is directed to the brake unit inlet, and then to the brake unit operating cylinders. It should be noted that unlike some other systems, the amount of fluid flow to actually apply the brakes is quite small. When brakes are released, the fluid normally flows from the brake unit back to the brake control valve, and then to return, i.e. back to the reservoir. The accumulator in the meantime is re-pressurised with system fluid, ready for further brake applications. HYPDROPLANING Hydroplaning is a phenomenon that occurs when a tyre loses contact with the runway surface due to a build-up of water in the tyre-ground contact area. NASA has researched this problem since the late 1950's and has identified three forms of hydroplaning: a.

Dynamic.

b.

Viscous.

c.

Reverted Rubber.

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DYNAMIC HYDROPLANING Dynamic hydroplaning is caused by the build-up of hydrodynamic pressure at the tyrepavement contact area. The pressure creates an upward force that effectively lifts the tyre off the surface. When complete separation of the tyre and pavement occurs, the condition is called total dynamic hydroplaning, and complete wheel stoppage and loss of control can occur. Below shows the forces on a tyre without the presence of water on the runway. The ground friction causes a spin-up moment and wheel rotation results. When the tyre is rolling freely at a fixed speed an a dry runway, the vertical ground reaction shifts forward of the axle and a spin-down moment that offers resistance to the wheel rotation is developed. When these two moments are equal the wheel is turning at a constant RPM. The introduction of water on the runway leads to dynamic hydroplaning which is illustrated. Deep fluid on the runway creates additional drag on the tyre when it is displaced from the tyre path, and a high spray pattern is produced as shown . As the forward speed of the aircraft is increased the spray pattern thrown up the tyre lowers and the wedge of water penetrates the tyre-ground contact area and produces a hydrodynamic lift force on the tyre. This is partial hydroplaning. As the speed increases the spray pattern becomes flatter, and the wedges of fluid penetrates farther into the ground contact area until at some high forward speed complete separation of the tyre and runway takes place and total hydroplaning occurs. The ground friction force is progressively reduced as the wedge of water penetrates beneath the tyre. It approaches zero at total hydroplaning, and the spindown moment causes the tyre to stop the wheel rotation Obviously no braking action is available when the wheel is not making contact with the runway and has stopped rotating. Total dynamic hydroplaning is more of a landing than a takeoff problem. However, crosswind takeoffs are dangerous under the conditions. Total dynamic hydroplaning usually does not occur unless there is a minimum water depth present on the runway to support the tyre. The exact depth required cannot be predicted since other factors influence dynamic hydroplaning. These factors include runway smoothness and tyre tread. Smooth runway texture induces hydroplaning with lower water depths than coarse textured surfaces. Smooth tread tyres will hydroplane more easily than ribbed tyres. While the exact depth of water required for hydroplaning has not been accurately determined, a conservative estimate for an "average" runway is that water depth in excess of 0. 1 may induce full hydroplaning of an aircraft. VISCOUS HYDROPLANING Viscous hydroplaning is more common than dynamic hydroplaning. Viscous hydroplaning may occur at lower speeds and at lower water depths than dynamic hydroplaning. Viscous hydroplaning occurs when the pavement surface is lubricated by a thin film of water. The tyre is unable to penetrate this film and contact with the pavement is partially lost. Viscous hydroplaning often occurs on smooth runway pavements of where tyre rubber deposits are present, usually in the touchdown area when a thin water film can significantly reduce the coefficient of fiction. Air Force measurements have shown that the coefficient of friction is nearly 25 % lower at the ends of a runway than in the middle.

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REVERTED RUBBER This third type of hydroplaning is known as reverted rubber hydroplaning. White streaks on the runway are an indication that this type of hydroplaning has occurred. Examination of the aircraft tyre will show an elliptically shaped tacky or melted rubber condition. This condition occurs when the heat that is generated during a locked wheel skid reverts the rubber to a molten or unvulcanized state. This is called reverted rubber. METHOD OF PREVENTING HYDROPLANING Method to prevent hydroplaning include transverse runway grooving, frequent removal of rubber deposits from the touchdown areas, maximum use of aerodynamic braking and firm, but not hard touchdowns. BRAKE UNIT ANTI-SKID SYSTEMS The Anti-skid Unit is fitted to achieve maximum retardation of the wheel brakes without the wheels skidding, the general result being: a) b) c) d)

Reduced tyre wear. Shorter landing run for a given set of conditions. Better aircraft control during braking. Reduced tendency for tyre blow-out.

ANTI SKID BRAKING The purpose of the anti-skid system as fitted to aircraft wheel brakes, is to prevent the wheels from skidding on wet, or icy surfaces, and to ensure that optimum braking effect can be obtained under all conditions, by modulating the hydraulic pressure to the brakes. Antiskid units sense the rate of change of wheel deceleration, decreasing the hydraulic pressure applied to the brakes when a high rate of increase in deceleration exists, consistent with an impending skid, and restoring it as the wheel accelerates again. A modulator valve, a form of hydraulic restrictor, is often fitted in conjunction with the anti-skid unit, to restrict the flow of fluid to the brake unit after initial brake application, and to conserve main system pressure. This action tends to smooth out the brake operation. There are two basic types of anti-skid systems in use, they are:

 

Mechanical System. Electronic System.

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MECHANICAL ANTI-SKID SYSTEM In the mechanical anti-skid system, the anti-skid unit is mounted either on the brake unit torque plate, or within the axle bore. The anti-skid device is located within the hydraulic wheel brake system, between the brake control valve and the brake unit, usually just prior to the brake unit. The anti-skid unit consists of a valve assembly connected to a flywheel, which is driven by the associated aircraft wheel. MECHANICAL ANTI-SKID SYSTEM OPERATION During normal braking action, when no skid is present, the flywheel rotates at the same speed as the drive, and the valve is closed, allowing hydraulic fluid through the unit to the brake unit. This fluid is supplied from the brake control valve, and is at maximum pressure for the control valve selection made. On leaving the anti-skid unit, the fluid is directed to the inlet of the brake unit, and then to the brake unit operating cylinders. When the rotational speed of the aircraft wheel decreases rapidly, as when the aircraft wheel is about to skid, the inertia of the flywheel causes the anti-skid unit valve mechanism to operate, opening the valve, and reducing the hydraulic pressure in the brake unit. The anti-skid unit directs the fluid back to return. The reduced pressure, which has also reduced the braking effect, allows the aircraft wheel to accelerate, the flywheel returns to its normal position, causing the valve to close and the brake to be applied again, until the wheel is about to skid and the action is repeated. If the wheel bounces clear of the ground after brakes have been applied, the adjustment of the anti-skid unit allows the brake to be completely released for a period of time to prevent the wheel locking, prior to making contact with the ground. Without this action a blow-out may occur. ELECTRONIC ANTI-SKID UNIT The system comprises a wheel speed transducer, a control unit, and an anti-skid valve in the brake pressure line, together with associated switches and check-out and warning lamps. The wheel speed unit may supply either d.c. or a.c. depending on the type of system used. Operation is basically similar to the mechanical system, but the use of sophisticated logic circuits in the later types of electronic control units enables much finer control to be exercised. Further refinements such as, strut oscillation, damping circuits, touch-down protection, and locked wheel protection, may also be incorporated, and some systems automatically de-activate at low speed to prevent interference with normal taxying manoeuvres. The method by which the wheel speed signal is processed in the control unit varies from type to type, but all operate on the basis that, if any brake produces more torque than can be supported by the friction between the tyre and the ground for the existing wheel load, the resulting impending skid will produce a smaller rotational velocity signal from the affected wheel. This reduced signal is detected by the anti-skid control circuits, which send a signal to the anti-skid control valve, causing brake pressure to be reduced sufficiently to correct the skid condition. Brake pressure will be re-applied, to a level just below that which caused the skid, and will then increase at a controlled rate. Control units normally contain circuits which provide warning of failure in the system, and a self-test facility which enables the serviceability of the various components to be checked. Controls for the operation and testing of the anti-skid system are contained in the control unit and in the flight compartment.

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FUSIBLE ALLOY PLUGS The majority of tubeless tyre main wheels are fitted with small pressure relief valves which are known as fusible alloy plugs. In the event a tubeless tyre is subjected to excessive heat due to heavy wheel braking there is a danger the tyre may blow out. If in the event of a blow, loss of control of the aircraft during the landing run may follow. To avoid this situation, it is better to slowly deflate the tyre, relieving the excess pressure whilst maintaining control of the aircraft. This is the main objective of the fusible alloy plug. The fusible alloy plug is retained in the wheel by a screw thread and is located between the flanges or wheel rims. In the event of the temperature of the air exceeding a certain value, the fusible alloy in the centre of the plug melts and slowly allows the air to escape, giving a controlled deflation of the tyre thereby preventing the tyre from bursting or blowing out. Some wheels may be fitted with a series of fusible alloy plugs located at various points around the wheel and normally set to melt at different temperatures. Generally there may be one or a combination of three temperature settings employed on a wheel. The common temperature at which fusible alloy plugs melt are as follows: a) b) c)

155 degrees centigrade 177 degrees centigrade 199 degrees centigrade

-

plug body coloured red. plug body coloured green. plug body coloured yellow.

In the event of any tyre being deflated during the landing run, and any tyre on the same axle, then the tyres are automatically scrapped. This means that regardless of whether the tyres on the same axle are deflated or not the tyres are all scrapped.

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AUTO BRAKES

Automatic braking at a pre-selected deceleration rate following touchdown. Various deceleration rates can be selected, however, the maximum autobrakes deceleration rate is less than that produced by full manual braking. Antiskid protection is provided during auto brakes operation. The auto brakes system is armed by selecting a deceleration rate. Latching of the Auto Brakes Selector indicates the system is armed and operative. During auto brakes operation the deceleration selection can be changed by rotating the selector without disarming the system. With the auto brakes system armed and both Thrust Levers at idle, automatic braking is initiated at touchdown when main Landing gear truck un-tilt and wheel spin-up occurs. In order to maintain the selected deceleration rate, auto brakes pressure is adjusted as other deceleration devices such as speedbrakes and thrust reversers contribute to total deceleration. The system provides braking to a complete stop or until It is disarmed. The system disarms immediately when an auto brakes or normal antiskid system fault occurs. Disarming also occurs if the following pilot actions are taken during auto brakes operation:  manual braking  advancing either Thrust Lever after landing  moving the Speedbrake Lever to the DOWN detent after speedbrakes have been extended on the ground  selecting the DISARM or OFF position When the system disarms, the selector moves to the DISARM position, terminating auto brakes operation. Illumination of the Auto Brakes Inoperative Light indicates that the system has disarmed. Rotating the Auto Brakes Selector to OFF removes power from the system and extinguishes the light. The RTO (rejected takeoff) mode can only be selected on the ground. Latching of the selector indicates the RTO mode Is armed. With the RTO mode armed, the auto brakes system applies maximum brake pressure if both Thrust Levers are retarded to idle above 85 knots. The brake pressure applied is equivalent to that provided by full manual braking. When the RTO mode disarms, the Auto Brakes Inoperative Light illuminates. During a rejected takeoff, after 85 knots, the same conditions which disarm the Landing mode also disarm the RTO mode. The Auto Brakes Selector remains in the RTO position after system disarming. At liftoff, the selector trips OFF. CPL ATG CPL DOC 01 Revision : 1/1/2001

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Revision Questions 1.

In a wheelbrake circuit, the hydraulic accumulator will provide an increase in the number of brake applications in an emergency: a. b. c.

2.

Excessive brake travel could be due to a. b. c.

3.

the brake pedal will experience no resistance as it travels to the floor the brakes will feel spongy application of light foot pressure will cause the brakes to grab

The purpose of a shimmey damper is to a. b. c.

5.

air in the system a weak return spring a fluid leak in the master cylinder

If air has leaked into the lines of a foot operated hydraulic brake system a. b. c.

4.

if the initial charge pressure is increased if the initial charge pressure is reduced if the initial charge pressure is exhausted

eliminate vibration of the tailwheel eliminate vibration of the main wheels reduce nosewheel vibration during ground operations

With reference to aeroplane operating manuals, the term Vle refers to a. b. c.

the maximum speed at which the undercarriage may be lowered the maximum speed at which the aeroplane may be flown with the undercarriage extended the maximum speed of the aeroplane in the landing configuration

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CHAPTER 13 ICE REMOVAL SYSTEMS DEFINITIONS ANTI-ICING Used to prevent the formation of ice on the critical areas of an aircraft. DE-ICING Used to remove ice from less critical areas after it has formed. TYPES OF ICE PROTECTION There are four main types of ice protections: FLUID In order to prevent the adhesion of ice, this type uses a fluid which is pumped over the area to be protected. The fluid is carried rearwards by the slipstream and forms a film over the surface. The older type of aircraft uses this system to protect aerofoils, propellers and windscreens. The main disadvantages lies in the limited amount of fluid which can be carried. MECHANICAL This is a de-icing aid where flexible rubber boots are fitted to the leading edges of the airframe. These boots expand and contract, thus cracking the ice off and allowing it to be removed by the airflow. THERMAL Hot air is employed to give either anti-icing or de-icing protection. The supply of hot air is taken from the aircraft engines, through ducting to the required areas, then exhausted overboard. The aerofoils, air intakes, windscreens and system components of many modern aircraft are protected in this manner. ELECTRICAL Heating elements are attached or sprayed on to the areas requiring protection. Electrical power from the aircraft system is supplied to the elements, and heating of the area takes place. Continuous heating provides anti-icing whilst de-icing can be obtained by the intermittent application of heat. Aerofoils, propellers, spinners, intakes, windscreens and system components can all be protected by this method.

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Some of the systems described can be used either for de-icing or anti-icing. Though designed for one particular mode, operational reasons may justify its use for the other. The areas sensitive to ice formation are: i)

Aerofoils.

ii)

Propellers.

iii)

Engine intakes.

iv)

Windscreens.

v)

Pitot and static heads.

FLUID SYSTEMS This system prevents the adhesion of ice on surfaces by pumping de-icing fluid over them.

The fluid is supplied from the storage tank to the pump through an integral filter. The pump has a single inlet and a number of delivery outlets to feed the distributors on the aerofoil leading edges. A diagrammatic layout is shown.

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PROPELLER ANTI-ICING In this system, de-icing fluid is sprayed along the leading edge of each propeller blade. The fluid is held in a small tank, usually in one of the engine nacelles. This supply is fed to a small rotor-type pump driven by an electric motor. The fluid is fed through a valve to a slinger ring attached to the rear of the propeller hub. As the propeller rotates, centrifugal force sends the fluid from the slinger ring out, via a slinger tube, to the leading edges of each blade root as shown. The fluid continues along the blade under centrifugal force, remaining in the blade boundary layer and spreading over the blade chord. MECHANICAL AIRFRAME DE-ICING

With mechanical de-icing, ice is allowed to form and accumulate for a period of time, and is then broken up and swept away by the airflow over the aircraft. The system is applicable to the low subsonic aircraft only, and is very common. It is used on the leading edge of the mainplane, tailplane and fin. Rubber de-icer boots, bonded to the leading edges, are pulsated pneumatically, the successive inflation and deflation breaking up the ice allowing it to be swept away.

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THERMAL ICE PROTECTION In this type of system, hot air is used to heat wing, tailplane, fin leading edges and also the area of engine air intakes this source of hot air can be tapped from: i) The engine exhaust, and ducted to atmosphere through a head exchanger. This component raises the temperature of the ram air passing through it, which is then used in the anti-icing/de-icing system. ii) The rear of the engine compressor, where it will be at a high temperature and pressure.

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AIRFRAME SYSTEM The temperature of the air supply is controlled to prevent local overheating. The hot air is then ducted along the inside of the leading edges, and allowed to exhaust to atmosphere at the tips. Electrically operated shut-off valves control the flow of hot air to the leading edge sections, and these may be operated by thermo-switches to provide temperature control. Temperature indication and overheat warning is normally provided. One disadvantage of the thermal system is that a loss of engine power occurs when air is drawn from the compressor. Because of this, it is usually operated as a de-icing rather than anti-icing system. However, the system can be linked to an ice detector so that it is automatically switched on when icing conditions exists. ENGINE HOT AIR ANTI-ICING The hot air system provides surface heating of the engine and-or power plant where ice is likely to form. The affected parts are the engine intake, the intake guide vanes, the nose cone, the leading edges of the nose cowl and, sometimes the front stage of the compressor stator blades. The protection of rotor blades is rarely necessary, because any ice accretions are dispersed by centrifugal action. The hot air for the anti-icing system is usually taken from the last stage of the compressor, and externally ducted through pressure regulation valves, to the parts requiring protection When the nose cowl requires protection, hot air exhausting from the air intake manifold may be collected and ducted to the nose cowl. Exhaust outlets are provided to allow the air to pass into the compressor intake or vent to atmosphere, thus maintaining a flow of air through the system. On some engines, however, the engine and nose cowl anti-icing systems are independent. The engine nose cone is protected by a continuous unregulated supply of hot air tapped off the compressor and internally ducted to the nose cone. The nose cowl receives its supply of hot air from the high pressure compressor, via external ducting and a pressure regulating valve.

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WINDSCREEN DE-ICING To prevent icing and misting, hot air is sometimes blown over the outside and inside of the windscreen and other transparent panels. The principle of the outside air blast system is also extended to rain clearance. Aircraft using this latter system include the Lightning, Jet Provost and DC8. Great care must be taken when ground running aircraft with this system to prevent overheating the windscreen.

ELECTRICAL SYSTEMS There are two main methods, spaymat and heater mats, of utilizing electricity for the prevention of removal of ice; they are both used for the protection of aerofoils, engine intakes and propellers.

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Each mat is designed for a specific application, the heat output being obtained from whatever electrical source is available. Mats are available both for anti-icing or de-icing. Anti-icing mats on intakes are supplied continuously with electricity, while the de-icing mat is intermittently heated. The total area to be heated is often divided into several smaller areas with independent mats for each. The electrical power is then arranged to be switched to each small area in turn. Thus, on any particular area, there is no heating for a given period during which the ice builds up and then, when power is switched to that area, adhesion is broken by heat and the ice removed by the airflow. WINDSCREEN PROTECTION Electrically heated windscreens are used on the majority of modern military and civil aircraft, both for ice and mist prevention, and to increase the resistance of the panels to bird strike at very low temperature. The older type used very thin wire heater elements sandwiched between the glass laminations, the wire being connected to the electrical circuit. However, the introduction of transparent electrically conductive heater films has almost eliminated these types. . A transparent film element in the panel is connected to the electrical system and used as a heater panel. Also incorporated in the panel is a temperature sensitive element, which automatically regulates the heater temperature and an overheat sensing unit to guard against failure of the temperature regulator.

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CHAPTER 14 PRESSURIZATION DEFINITIONS Absolute pressure: pressure measured along a scale which has a zero value at a complete vacuum. Aircraft altitude: the actual height of the aircraft above sea level. Ambient pressure: the pressure in the area immediately surrounding the object under discussion. Standard barometric pressure: the weight of gases in the atmosphere sufficient to hold a column of mercury 760 mm high (approx. 30 in) at sea level, i.e. 17.7 psi at sea level. This pressure decreases with altitude. Cabin altitude: used to express cabin pressure in terms of equivalent altitude above sea level. Differential pressure: the difference between cabin pressure and atmospheric pressure. Positive differential: cabin is said to be having positive differential when cabin pressure is more than ambient pressure. Negative differential: cabin is said to be having negative differential when cabin pressure is less than ambient pressure. Gauge pressure: a measure of pressure in a vessel, container or line as compared to ambient pressure.

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PRESSURIZATION SYSTEM REQUIREMENTS PHYSIOLOGICAL CONSIDERATION - OXYGEN CONTENT Lack of oxygen is the most important factor in the adverse physiological effects encountered at altitude. At ground level and at rest, a person normally breathes 30 cubic inches or approximately 500 cubic centimetres of air at each breath, and at a normal respiration rate of 12 to 15 cycles per minute. Which means that the oxygen requirements for a human body are six to eight litres of air per minute. The primary effect of exposure to low air pressure is that the oxygen content of the blood is decreased, the immediate effect being on the brain. Oxygen absorption into the blood across the surface of the lungs decreases with reduction in atmospheric pressure. The effects may be felt at altitudes greater than 10 000 feet, although they are not generally serious even when they aggravate feelings of fatigue. At an altitude of about 20 000 feet, reduction in oxygen content leads to a phenomenon called HYPOXIA with consequent dulling of all the senses. It is more serious in that the victim is generally unaware of its onset (hypoxia) and believes that he is acting in a competent and cheerful manner. Meanwhile one's judgement is seriously impaired, and disastrous mistakes can ensue. Further decrease in atmospheric pressure leads to another phenomenon, called ANOXIA, which is due to serious lack of oxygen in the human body. Anoxia causes lack of clarity, loss of sight (defective vision), deterioration in hearing, uncontrolled trembling of muscles, and the victim becomes weak and sick. Total unconsciousness eventually occurs, and the time taken to fall unconscious decreases rapidly with increase in altitude. For example, altitudes of about 25 000 feet produce unconsciousness in six minutes and at about 40 000 feet in 20 seconds followed by death. It is also necessary for a certain minimum oxygen pressure to exist in the lungs, for the oxygen to diffuse through the lung tissue and pass to the red corpuscles in the blood stream. The minimum ambient pressure to maintain life, in this respect is 2.7 psi absolute, the equivalent altitude being 40 000 feet, even if 100 per cent oxygen is being breathed. From this aspect of passenger comfort, the ideal cabin conditions to maintain would be those corresponding to sea level. However, considering cabin structural requirements and consequent power and weight penalty, the difference between cabin and ambient pressure should be kept to the minimum. Experiences have shown that the vast majority of unselected persons suffer no discomfort, due to hypoxia at altitudes of up to 8 000 feet. For all healthy persons, in all conditions, cabin altitude of 8 000 feet is generally accepted as a safe figure to maintain passenger comfort, and from aircraft structural considerations.

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RATE OF CHANGE OF CABIN PRESSURE Apart from altitude, the question of the rate of change of cabin altitude within the aircraft must also be considered. It is common practice for an aircraft to climb at a rate of 1 000 feet per minute or more and engines are designed to maintain their performance with the changing atmospheric conditions during a climb. But the effect of rapid altitude changes on the human body causes physical pain and discomfort. If the cabin pressure decreases violently, the nitrogen and other gases in solution in the blood stream expand rapidly in the form of bubbles. This causes acute pain and injury, but would not normally occur except in the event of explosive decompression. If however, the rates of change are large but not violent, the most common effects are: i)

Sickness.

ii)

Expansion of gases in the abdomen (uncomfortable).

iii)

Mild effects of bends in exceptional cases.

iv)

Expansion of gases in the ear.

Cabin pressure can be decreased at slightly higher rates than increased without undue discomfort. Here again it has been established that a rate of change of cabin altitude of 500 feet/min for climb, and 300 feet/min for descent are acceptable limits for pressurized passenger aircraft. PRINCIPLE OF PRESSURIZATION REQUIREMENTS OF A SEALED CONTAINER. Firstly we require a sealed container, strong enough to withstand all stresses, and which could be taken to any altitude and still maintain the same pressure internally. Every endeavour must be made on the part of the manufacturer to ensure that the shell is perfectly sealed. This is essential if the air that is entering and leaving the cabin is to be effectively controlled. Unwanted leaks can occur at rivet points, windows, doors, escape hatches and each hole in the skin through which a service must pass to reach the unpressurized portion of the aircraft. Through these various holes must pass the main flying and trimmer controls, hydraulic, pneumatic, electrical and other service lines and cables. Each hole must be effectively sealed, but main flying control rods, trimmer cables and throttle controls must allow freedom of movement without too much static friction. AIR INPUT INTO THE SEALED CONTAINER To cater for the ventilation requirements, fresh air must be supplied to the sealed container through the inlet port, and be allowed to escape to the ambient through controlled discharge valves which will offer a restriction the outflow from the container. If the pressure in the cabin is to be controlled accurately by the position of the outflow valve, and for simplicity of system design, the input air into the cabin should be kept at a constant mass flow. That is, the control will automatically allow for changes in the air density and changes in cabin back pressure on the inlet, due to changes in the position of the outflow valve, maintaining a constant mass flow (lbs/min) into the cabin irrespective of cabin pressure.

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BASIC METHOD OF PRESSURE CONTROL The principle of pressurization is that whilst the aircraft is climbing at high rate to its cruise altitude, the cabin is climbing at a slower rate, to a lower altitude. Consider a sealed container (fuselage) where a constant mass air flow is coming in, if the same mass air flow is going out through a control valve, then the internal pressure will be held constant. Furthermore if the control valve stays in one position, the internal pressure must finally stabilize. Closing the control valve (outflow valve) will increase the pressure, and opening the valve will decrease the pressures and when the pressure is stabilized, there will be as much air going out as there is coming in. CONTROL OF RATE OF PRESSURE CHANGE The cabin rate of climb should be selected to ensure that the set cabin altitude is reached at approximately the same time as the aircraft reaches its cruise altitude. For example, if the aircraft was flight planned to fly at 20 000 feet altitude, and when pressurized to its maximum differential of 5 psi, it would have a cabin altitude of 6 000 feet. Then the time taken for the cabin to reach 6 000 feet from sea level would be the same as the time taken for the aircraft to reach 20 000 feet. So that comfortable conditions are made available for passengers and crew during climb and descent, the rate of cabin pressure change should be approximately 500 feet/min during climb and 300 feet/min during descent. This is achieved by controlling of the outflow valve.

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RATE CONTROL

A rate control can be incorporated in order to control the rate of change of cabin altitude, and also to limit the minimum and maximum rates of change. High rates of change are uncomfortable for the occupants and may cause severe pain in the ears. TYPICAL PRESSURIZATION SYSTEM CHARACTERISTICS When the aircraft is on the ground and with the cabin altitude set for field elevations and barometric corrections made, the cabin pressure will be the same as the ambient pressure. After take-off the desired cabin altitude for cruise flight and the desired rate, are selected. If the aircraft climbs at a rate higher than selected rate the cabin will climb at the selected rate until the selected cabin altitude is reached. If the aircraft climbs at the same rate as the selected rate, or at a lower rate than the selected rate, then the cabin will climb at the same rate as the aircraft until the selected cabin altitude is reached. If the aircraft levels off below the selected cabin altitude, the cabin will climb to the aircraft altitude and then will also level off until the aircraft climbs again. The cabin altitude will be maintained at the selected level during cruise, provided the maximum pressure differential is not exceeded. If the aircraft climbs further, the cabin altitude will be maintained at the selected level until the differential pressure limit is reached, and then the cabin will also climb. The rate of cabin climb will be such that the maximum pressure differential is maintained. In effect, the actual cabin climb rate will be lower than the aircraft climb rate. Before descent, the cabin altitude is set to the landing field elevation, and barometric correction is done. The cabin will descend at the selected rate, provided the maximum pressure differential is not exceeded, and will automatically go to the landing field elevation on touch down.

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Revision Questions 1

To measure the rate of climb of the cabin in a pressurised aircraft, the cabin VSI uses: a. static pressure measured at the aircraft altitude b. cabin differential pressure c. static pressure measured inside the cockpit

2.

After a flight in a pressurised aircraft, none of the doors are able to be opened. This could be caused by a. b. c.

3.

During a climb, the pilot notices that the cabin altimeter and VSI are reading the same as the aircraft altimeter and VSI. A possible cause is A. b. c.

4.

remain the same as outside pressure up to 8000 ft decrease at twice the rate of the outside pressure decrease at half the rate of the out side pressure

Oils and greases should not be used on joints in oxygen lines because a. b. c.

6.

the negative relief valve is stuck closed the aircraft is not above 10,000 ft yet the safety valve has opened

On a given aircraft’s pressurisation panel the cabin altitude has been set to 8000 ft and cabin rate of change is set to maximum (2000 fpm). If after take-off from sea level the aircraft climbs at 1000 fpm, the cabin pressure would initially a. b. c.

5.

the safety valve failing to operate due to faulty weight switches the negative relief valve opening one landing the use of incorrect door opening procedures

they can clog the lines they may start a fire they will cause the connections to oxidise (rust)

The use of supplemental oxygen by a pilot in a pressurized aircraft is mandatory a. b. c.

at all times for a flight planned above 10,000 ft at all times when the aircraft is above 10,000 ft for longer than 60 mins at all times when the cabin altitude is above 10,000 ft for longer than 60 mins

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7.

For cruise in an unpressurised aircraft at FL270 a diluter demand oxygen system a. b. c.

8.

is not suitable for prolonged operations would deliver oxygen under pressure need not supply 100% oxygen

Continuous flow oxygen systems a. b. c.

are suitable for prolonged use up to 25,000 ft are suitable for prolonged use up to 33,000 ft are suitable for prolonged use up to 40,000 ft

9.

Which a. b. c.

statement is incorrect regarding autopilot operation? Many autopilots should not be engaged above certain limiting IAS values Autopilots should always be off for take-off and landing Autopilots are designed to maintain attitude and so should be used during flight through turbulence

10.

When disengaging an autopilot a pilot’s main concern should be to a. b. c.

11.

guard against sudden changes in aircraft attitude maintain straight and level flight maintain a rate one turn

Which of the following ice protection systems is designed to be used only for de-icing? a. b. c.

Alcohol spray Hot air Pneumatic system

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