Copyright Julien Evans 2016
Published by Steemrok Publishing
steemrok.com

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HANDLING LIGHT AIRCRAFT

CONTENTS

Introduction

List of figures

1        BASIC THEORY OF FLIGHT

2        THE AIRCRAFT

2.1    THE AIRFRAME
2.1.1 The wing
2.1.2 The tailplane and elevator
2.1.3 The fin and rudder
2.1.4 The fuselage
2.1.5 The landing gear
2.1.6 Trim tabs
2.1.7 The cabin
2.1.8 Other components

2.2    THE ENGINE
2.2.1 Combustion
2.2.2 Behaviour of liquids and gases
2.2.3 The four-stroke cycle
2.2.4 The carburettor
2.2.5 The throttle
2.2.6 Mixture strength and mixture control
2.2.7 Carburettor heat control
2.2.8 Detonation
2.2.9 Fuel injection
2.2.10 The ignition system: the magneto
2.2.10.1 Mechanical generation of electricity
2.2.10.2 Conversion to high voltage
2.2.10.3 Timing and distribution
2.2.10.4 The spark plug
2.2.10.5 Dual ignition systems
2.2.10.6 The impulse magneto
2.2.10.7 Ignition control
2.2.10.8 Electronic ignition systems
2.2.11 The oil system
2.2.11.1 Oil quantity and consumption
2.2.11.2 Oil grade
2.2.11.3 Dry-sump systems
2.2.12 Engine cooling
2.2.13 The propeller
2.2.14 Engine mounting
2.2.15 The cowling
2.2.16 The tachometer
2.2.17 Engine-driven systems
2.2.18 FADEC systems

2.3    THE FUEL SYSTEM
2.3.1 Fuel quantity and consumption
2.3.2 Fuel grade

2.4    THE ELECTRICAL SYSTEM
2.4.1 The alternator
2.4.2 The electrical circuit
2.4.3 The electrical system
2.4.4 Fuses and circuit-breakers
2.4.4.1 The fuse
2.4.4.2 The circuit-breaker
2.4.5 Electric motors
2.4.6 The ammeter
2.4.7 Electrical services

2.5    THE HYDRAULIC SYSTEM
2.5.1 The brake system
2.5.2 Hydraulic fluid
2.5.3 Hydraulically-operated landing gear retraction

2.6    THE FLIGHT INSTRUMENTS
2.6.1 Visual flight
2.6.2 Instrument flight
2.6.3 The International Standard Atmosphere
2.6.4 The airspeed indicator
2.6.4.1 Pressure error and instrument error
2.6.4.2 Density error
2.6.4.3 Groundspeed
2.6.5 The altimeter
2.6.5.1 Pressure datum setting
2.6.5.2 Altimeter terminology
2.6.5.3 Barometric error
2.6.5.4 Terrain clearance
2.6.5.5 Temperature error
2.6.5.6 Pressure error and instrument error
2.6.5.7 GPS derived altimetry
2.6.6 The vertical speed indicator
2.6.6.1 Lag
2.6.6.2 Pressure error
2.6.7 Direction
2.6.8 The magnetic compass
2.6.8.1 Compass errors
2.6.9 The gyroscope
2.6.9.1 The suction-driven gyro
2.6.9.2 The electrically-driven gyro
2.6.9.3 Flight instrument power supplies
2.6.10 The attitude indicator
2.6.11 The direction indicator
2.6.11.1 DI errors
2.6.11.2 The self-synchronising DI
2.6.12 Toppling
2.6.13 The turn-and-balance indicator
2.6.13.1 The turn indicator
2.6.13.2 The balance indicator
2.6.14 Electronic flight instrument systems (EFIS)

2.7    THE PILOT'S CONTROLS
2.7.1 The flight controls
2.7.1.1 The ailerons
2.7.1.2 The elevators
2.7.1.3 The rudder
2.7.1.4 Control locks
2.7.2 The flaps
2.7.3 The elevator trim tab
2.7.4 The engine controls
2.7.5 Electrical services
2.7.6 Seats and harnesses

2.8     AVIONICS
2.8.1 The COM radio
2.8.2 The transponder
2.8.3 The GPS map display

2.9    THE CHECKLIST

3        DETAILED THEORY OF FLIGHT

3.1    WEIGHT AND CENTRE OF GRAVITY

3.2    LIFT
3.2.1 Angle of attack
3.2.2 Wing shape
3.2.3 Factors affecting lift
3.2.3.1 Angle of attack
3.2.3.2 Speed
3.2.3.3 Aerofoil shape
3.2.3.4 Wing area
3.2.3.5 Air density

3.3    TAIL DOWN-FORCE

3.4    DRAG
3.4.1 Factors affecting drag
3.4.1.1 Object shape
3.4.1.2 Speed
3.4.1.3 Object size
3.4.1.4 Air density
3.4.2 Wing drag
3.4.2.1 Speed
3.4.2.2 Angle of attack
3.4.2.3 Wing planform
3.4.2.4 Air density
3.4.3 Total aircraft drag

3.5    ANGLE OF INCIDENCE

3.6    WING EFFICIENCY

3.7    THE MOTION OF THE AIRCRAFT

3.8    STABILITY
3.8.1 Longitudinal stability
3.8.2 Directional stability
3.8.3 Lateral stability
3.8.4 Interaction of directional and lateral stabilities
3.8.5 Effect of position of centre of gravity on stability
3.8.5.1 Effect on longitudinal stability
3.8.5.2 Effect on directional stability
3.8.5.3 Effect on lateral stability

3.9    AIR DENSITY

3.10 THRUST: THE PROPELLER
3.10.1 Effect of varying engine power setting
3.10.2 Effect of speed on RPM
3.10.3 Windmilling
3.10.4 Propeller efficiency
3.10.4.1 Effect of speed on propeller efficiency

4        AIRCRAFT HANDLING

4.1    AIRFRAME LIMITATIONS

4.2    ENGINE HANDLING AND OPERATING LIMITATIONS
4.2.1 Control of power
4.2.2 Use of mixture control
4.2.3 Engine operating limitations

4.3    PICKETING AND USE OF CHOCKS

4.4    MANHANDLING AND POSITIONING THE AIRCRAFT FOR ENGINE-STARTING

4.5    PROPELLER HANDLING

4.6    AIRCRAFT INSPECTION

4.7    ENGINE STARTING

4.8    TAXYING
4.8.1 Brake failure

4.9    ENGINE TESTING

4.10  PRE-TAKE-OFF CHECKS
4.10.1 Airframe
4.10.2 Engine
4.10.3 Instruments
4.10.4 Electrical services

4.11  PROCEDURE AFTER LANDING

4.12  EFFECT OF FLIGHT CONTROLS IN FLIGHT
4.12.1 Effect of elevators
4.12.2 Effect of ailerons
4.12.2.1 Consequence of bank
4.12.3 Effect of rudder
4.12.3.1 Consequence of yaw
4.12.3.2 Propwash-induced yaw

4.13  EFFECT OF VARYING PROPWASH STRENGTH

4.14  EFFECT OF VARYING AIRSPEED ON FLIGHT CONTROLS

4.15  EFFECT OF VARYING CONTROL MOVEMENT

4.16  EFFECT OF FLIGHT CONTROLS IN DISPLACED ATTITUDE

4.17  FUNCTION OF FLIGHT CONTROLS
4.17.1 Function of elevators
4.17.2 Function of ailerons
4.17.3 Balance: function of rudder

4.18  FUNCTION OF TRIM TABS
4.18.1 Trimming technique
4.18.2 Trim changes

4.19  THE FLIGHT PATH: THE THIRD DIMENSION

4.20  STRAIGHT AND LEVEL FLIGHT
4.20.1 Level flight
4.20.2 Control of flight path with attitude
4.20.3 Control of speed with power
4.20.4 Straight flight
4.20.5 Balance
4.20.6 Drag
4.20.7 Power
4.20.8 Flying for maximum range
4.20.8.1 Effect of height on range
4.20.9 Flying for maximum endurance
4.20.9.1 Effect of height on endurance
4.20.10 Technique for straight and level flight
4.20.11 Correction of deviations

4.21  CLIMBING
4.21.1 Climbing at maximum rate
4.21.2 Climbing at maximum gradient
4.21.3 Summary of techniques for climbing
4.21.4 Cruise climb

4.22  DESCENDING
4.22.1 Gliding at minimum gradient of descent
4.22.2 Sideslipping
4.22.3 Powered descent
4.22.4 Summary of techniques for descending
4.22.5 Cruise descent

4.23  FLAPS
4.23.1 Effect of flaps on stalling speed
4.23.2 Effect of flaps on trim
4.23.3 Control of speed and vertical flight path
4.23.4 Effect of flaps on forward view from cabin
4.23.5 Use of flaps for take-off
4.23.6 Use of flaps for approach and landing

4.24  MEDIUM TURNS
4.24.1 Level turns
4.24.1.1 Balance: use of rudder
4.24.1.2 Rate of turn
4.24.2 Climbing turns
4.24.3 Descending turns
4.24.4 Increase of stalling speed
4.24.5 Summary of techniques for turning

4.25  STALLING AND SPINNING
4.25.1 Symptoms of impending stall
4.25.2 The stall
4.25.3 The spin
4.25.4 Recovery from the stall
4.25.5 Practising stalls
4.25.5.1 Entering the stall from straight and level flight
4.25.5.2 Effect of lowered flaps on stalling characteristics
4.25.5.3 Effect of power on stalling characteristics
4.25.5.4 Entering the stall from level turning flight
4.25.5.5 Stalling in climbing and descending flight
4.25.5.6 Stalling in pitch-up manoeuvres
4.25.6 Effect of loaded weight on stalling speed
4.25.7 Effect of CG position on stalling characteristics
4.25.8 Inadvertent stalls: recovery at the incipient stage
4.25.9 Recovery from the spin
4.25.10 Recovery from spiral dive
4.25.11 Practising spins
4.25.11.1 Entering the spin
4.25.12 Effect of CG position on spinning characteristics
4.25.13 Airframe stress during spin recovery
4.25.14 Inadvertent spins
4.25.15 Stall and spin avoidance

4.26  THE RUNWAY

4.27  THE CIRCUIT

4.28  THE TAKE-OFF
4.28.1 Normal take-off technique (flaps up)
4.28.1.1 Wind
4.28.2 Crosswind take-off
4.28.2.1 Crosswind component
4.28.3 Short take-off
4.28.4 Obstacle clearance after take-off
4.28.5 Rejected take-off

4.29  FLYING THE CIRCUIT

4.30  THE APPROACH
4.30.1 Base leg
4.30.2 Final approach
4.30.2.1 Approach path control
4.30.2.2 Centre-line tracking
4.30.2.3 Alternative technique for final approach

4.31  THE LANDING
4.31.1 Normal landing technique
4.31.1.1 Control wheel movement
4.31.1.2 Centre-line tracking
4.31.1.3 Wind
4.31.2 Crosswind landing
4.31.2.1 Crab method for crosswind landings
4.31.2.2 Crossed controls (sideslip) method for crosswind landings
4.31.3 Crosswind component
4.31.4 Short landing

4.32  GLIDE APPROACH AND LANDING
4.32.1 Glide approach technique
4.32.2 Mis-judged approaches
4.32.3 The flare

4.33  FLAPLESS APPROACH AND LANDING
4.33.1 The flare and landing

4.34  THE GO-AROUND

4.35  THE WINDSOCK

4.36  STEEP TURNS
4.36.1 Lift increase
4.36.2 Rate of turn
4.36.3 Increase of stalling speed
4.36.4 Steep level turns
4.36.4.1 Stall avoidance
4.36.4.2 Recovery from spiral dive
4.36.4.3 Maximum rate turns
4.36.5 Steep gliding turns

4.37  REVIEW OF CONTROL FUNCTIONS
4.37.1 Function of elevators
4.37.2 Function of ailerons
4.37.3 Function of rudder
4.37.4 Function of power
4.37.5 Flight near stalling regime

4.38  INSTRUMENT FLIGHT
4.38.1 Physiological aspects
4.38.2 Instrument scan
4.38.3 Control technique
4.38.4 Navigation
4.38.5 Terrain clearance
4.38.6 Collision avoidance
4.38.7 Meteorological aspects

4.39  CONSTANT-SPEED PROPELLERS
4.39.1 Theoretical considerations
4.39.2 Technical description
4.39.3 The manifold pressure indicator
4.39.4 Operating technique

4.40  RETRACTABLE LANDING GEAR

5       EMERGENCY HANDLING

5.1    ENGINE FAILURE
5.1.1 Forced landing procedure
5.1.1.1 Choice of most suitable field
5.1.1.2 Choice of landing direction
5.1.1.3 Planning descent path
5.1.1.4 Attempting to rectify failure
5.1.1.5 Impact checks
5.1.1.6 Flying the descent path
5.1.1.7 Overrunning
5.1.1.8 Mis-judged approaches
5.1.1.9 Summary of forced landing procedure
5.1.1.10 After landing
5.1.1.11 Engine failure at low height
5.1.1.12 Engine failure after take-off
5.1.2 Further considerations
5.1.3 Practising forced landing procedure

5.2    THE PRECAUTIONARY LANDING

5.3    ENGINE FIRE
5.3.1 Engine fire in the air
5.3.2 Engine fire on the ground

5.4    CABIN FIRE
5.4.1 Further considerations

5.5    DITCHING

5.6    LANDING GEAR EMERGENCIES

6       THE FLIGHT MANUAL

6.1    TECHNICAL DESCRIPTION

6.2    LIMITATIONS

6.3    PERFORMANCE
6.3.1 Length of take-off run required
6.3.1.1 Wind component
6.3.1.2 Airfield elevation
6.3.1.3 Air temperature
6.3.1.4 aircraft loaded weight
6.3.1.5 Flap position
6.3.1.6 Gradient of runway
6.3.1.7 Nature of runway surface
6.3.1.8 Practical considerations
6.3.2 Obstacle clearance after take-off

6.4    LOADING
6.4.1 Empty weight and loaded weight
6.4.2 CG position
6.4.3 Practical considerations
6.4.4 Dangers of incorrect loading

7    GROUND SCHOOL SUBJECTS

Introduction

This book is an updated partial rewrite of the book 'The Pilot's Manual', first published by T and A D Poyser (ISBN 978-0856610202), which dealt with all aspects of light aircraft operation, including the ground subjects forming the syllabus for the UK Private Pilot's Licence. In the intervening years there have been little or no changes in some aspects (handling these aircraft), more significant changes in others (aircraft construction materials and processes) and revolutionary changes in yet more (flight instrumentation, airspace complexity and regulation and navigational equipment and procedures). This book restricts itself to the technical description of conventional all-metal light aircraft, the theory of flight and aircraft handling in daylight visual weather conditions (although a brief dissertation on the fundamentals of instrument flight is also included).

On the subject of personal pronouns I have used the masculine 'he' form throughout to make the text more easily readable, asking the reader to assume the inclusion of the feminine 'she' form by inference. I hope this modus scribendi is acceptable to female readers.

List of figures

1   BASIC THEORY OF FLIGHT

1     The aircraft
2     Disposition of forces acting on an aircraft in flight at constant speed and height
3     High-winged aircraft

2   THE AIRCRAFT

4     Left wing
5     Flap movement
6     Aileron movement
7     Left tailplane and elevator
8     Elevator movement
9     Fin and rudder
10   Rudder movement
11   The fuselage
12   Landing gear unit
13   Oleo-pneumatic unit in various situations
14   Creep marks
15   Adjustable elevator trim tab
16   Fixed aileron trim tab
17   Flow of gas through a constriction
18   The cylinder and its associated components (induction stroke)
19   Compression stroke
20   Power stroke
21   Exhaust stroke
22   Horizontally-opposed four-cylinder aero-engine
23   Float-type carburettor
24   Throttle valve
25   Mixture control at intermediate position
26   Carburettor heat control 'on'
27   Mechanical generation of electricity
28   The magneto: the electrical set-up
29   The spark plug
30   Dual ignition system for four-cylinder aero-engine
31   The magneto switch
32   The oil system
33   Cooling fins
34   The propeller
35   The tachometer
36   The fuel system
37   Electrical circuits
38   Aircraft electrical system
39   The fuse
40   Ammeter location
41   Ammeter presentation
42   Ammeter located in battery circuit
43   Ammeter presentation (battery circuit)
44   The brake system
45   Landing gear retraction mechanism
46   The flight instruments
47   High and low attitudes
48   View from pilot's seat
49   Bank left and right (view from behind)
50   View from pilot's seat
51   Aircraft in cruising flight
52   ASI working principle
53   Airspeed indicator
54   The pressure head
55   Airspeed indicator presentation
56   TAS (knots) corresponding to IAS of 100 knots
57   TAS and GS
58   The altimeter
59   Altimeter presentation (showing 4650 feet)
60   Altimeter presentation (showing 8800 feet)
61   Altimeter presentation (showing 12000 feet)
62   Reference for height measurement
63   Barometric error
64   Terrain clearance
65   The vertical speed indicator
66   Vertical speed indicator presentation
67   Incorrect VSI indication during abrupt transition from climb to descent
68   Magnetic variation
69   Magnetic direction
70   Heading
71   The magnetic compass
72   Magnetic compass showing heading 270M
73   Heading indications
74   The gyro
75   Suction-driven gyro
76   Attitude indicator gyro erect
77   Attitude indicator presentation
78   Attitude indicator indications
79   Direction indicator gyro erect
80   DI indicator indicating 360
81   DI indicator indicating 255
82   Turn indicator gyro precession
83   Turn indicator presentation
84   Turn indications
85   Balance indicator
86   Unbalanced flight
87   Turn-and-balance indicator
88   EFIS display
89   Cabin layout
90   Aileron control
91   Elevator control
92   Rudder control and nosewheel steering
93   Flap control
94   Elevator trim tab control
95   Magneto switch control
96   Control of electrical services
97   COM transceiver
98   The headset
99   The transponder

3 DETAILED THEORY OF FLIGHT

100  Airflow past flat plate inclined at a shallow angle
101  Wing aerofoil shape
102  'Lift' and 'downwash'
103  Wing above stalling angle
104  Relationship between lift and angle of attack at constant speed
105  Generation of tail down-force
106  Disturbance of airflow
107  Effective lift force
108  Induced drag
109  Induced drag generated at low and high angles of attack
110  Low and high aspect ratios
111  Air spillage at wing tips
112  Angle of incidence and angle of attack
113  Relationship between lift/drag ratio and angle of attack
114  Pitching, yawing and rolling
115  Disturbance causing nose to yaw to left
116  Dihedral
117  Sideslip to the left
118  High-winged aircraft in sideslip to the right
119  Airflow pattern during sideslip to the left
120  Effect of load disposition on CG position of loaded aircraft
121  CG coincident with lift force
122  CG position behind lift force
123  Direction of airflow past propeller blades
124  Propeller blade at most efficient angle of attack
125  Thrust and torque
126  Difference in rotational speed at blade roots and tips
127  Airflow direction at roots and tips
128  Helical twist in propeller blades
129  Effect of increased engine power
130  Effect of increased speed
131  Windmilling propeller
132  Forces resulting from windmilling propeller
133  Propeller efficiency during take-off
134  Reduced efficiency resulting from lower blade angle in cruising flight

4   AIRCRAFT HANDLING

135  Picketing
136  Chocks
137  Propeller arc
138  Elevators displaced upwards
139  Effect of elevators
140  Ailerons displaced by moving the control wheel to the left
141  Ailerons used to select and maintain an angle of bank of 30 to the left
142  Pilot's view of 30 angle of bank to the left
143  Yaw caused by bank
144  Pilot's view of yaw caused by bank
145  Rudder displaced to the left
146  Effect of rudder
147  Skidding
148  Spiral propwash impinging on fin-rudder assembly
149  Pitching and yawing in banked attitudes
150  Unbalanced flight caused by propwash-induced yaw
151  Unbalanced and balanced turning flight
152  Airflow past tailplane-elevator-tab assembly
153  Airflow past fin-rudder-tab assembly
154  Angle of attack dependent on relationship between pitch attitude and actual flight path
155  Disposition of forces in climbing, level and descending flight

STRAIGHT AND LEVEL FLIGHT

156  Generation of required lift
157  Attitudes for level flight at various speeds
158  Sideslip nullifying action of fin opposed by application of rudder
159  Unbalanced straight flight
160  Induced drag during flight at high and low IAS
161  Variation of drag with IAS
162  Variation of power required to maintain level flight (at sea level) with IAS
163  Variation of power required to maintain level flight (at various heights) with TAS

CLIMBING

164  Power available and power required for varying IAS
165  Typical climbing attitude
166  Gradient of climb
167  Climb at maximum rate and at maximum gradient

DESCENDING

168  Gradient of descent
169  Disposition of forces in the glide
170  Gradient of descent dependent upon ratio of required lift to total drag
171  Gliding at low IAS
172  Sideslipping to the left
173  Pitching effects arising from lowered flaps
174  Forward view enhanced when flying with flaps down
175  Going around

MEDIUM TURNS

176  Restoring magnitude of vertical component of lift by increasing lift force
177  Effect of wing tip drag difference as bank applied
178  Unbalanced flight during application and removal of bank
179  Typical attitude for climbing turn with 15 bank
180  Typical attitude for descending turn with 30 bank

STALLING AND SPINNING

181  Pre-stall buffeting
182  Operation of stall-warning vane
183  Causes of downward pitching at stall
184  Spin to the left
185  Oblique airflow at pressure head during spin
186  Stall practice at safe height
187  Entering stall from straight and level flight
188  Stall occurrence not directly related to attitude
189  Stalling in abrupt pitch-up manoeuvre
190  Rudder screened by elevators when control wheel moved forwards during spin recovery
191  Spin recovery procedure
192  Turning with crossed controls

THE CIRCUIT

193  Runways
194  Left-hand circuit
195  Positioned for take-off
196  Airborne, accelerating to climbing speed
197  Take-off profile
198  Effect of wind strength on climb gradient
199  Crosswind take-off
200  Weathercocking
201  Crosswind blowing aircraft away from centreline
202  Drifting effect of crosswind
203  Tracking along extended centreline in crosswind
204  Wind components
205  Crosswind component
206  Obstacle clearance after take-off
207  Drift allowance
208  Drift allowance in crosswind conditions
209  Changing runway appearance as base leg is flown
210  Final approach
211  Ideal descent gradient
212  Approach too high
213  Approach too low
214  Correction of too-high and too-low approaches
215  Correction of centreline displacement
216  The flare
217  Landing attitude
218  Ballooning and its correction
219  Crosswind from the right on final approach
220  Crosswind landing
221  Premature elimination of drift
222  Crossed controls method for crosswind landing
223  Final stages of short landing approach
224  Downwind and upwind glide at minimum gradient IAS
225  Shortening downwind leg for practice glide approach
226  Reducing and conserving height surplus on glide approach base leg
227  Ideal flight path in practice glide approach
228  Undershooting
229  Extending downwind leg for practice flapless approach
230  Flapless approach
231  The windsock

STEEP TURNS

232  Lift increase necessary to give required vertical component in banked attitudes
233  Variation of stalling speed with angle of bank
234  Steep level turns using 45 angle of bank
235  Variation of maximum angle of bank sustainable in level flight with power
236  45 banked gliding turn to the left

INSTRUMENT FLIGHT

237  Instrument scan

VARIABLE-PITCH PROPELLERS

238  Variable-pitch propeller
239  Engine controls
240  Manifold pressure indicator

5   EMERGENCY HANDLING

241  Courses of action to be taken following engine failure
242  Choice of field for forced landing
243  Average field and runway dimensions
244  Choice of circuit direction dependent on aircraft height and orientation relative to field
245  Attempting to prevent overrunning
246  Forced landing procedure
247  Engine failure at low height
248  Danger of attempting to turn back to airfield after engine failure during climb-out

6   THE FLIGHT MANUAL

249  Taking off with headwind and tailwind components
250  Wind component
251  Lever arm
252  Moment of empty aircraft
253  Moment of pilot
254  CG positions for empty and loaded aircraft
255  Permitted range for CG position
256  Loading table
257  Completed loading table

   





1   BASIC THEORY OF FLIGHT






The design of a typical modern single-engined aircraft is as shown in Figure 1, with the main features annotated. The front end of the aircraft is called the nose, and the rear end is called the tail. The design and function of each component will be described later. In this Section the basic aerodynamic theory of flight will be considered. (Aerodynamics is the study of the movement of objects relative to the air.)


All the component parts of the aircraft have their own individual weight. However, it it easier to think of the separate weight contributions as combining to form a single equivalent force acting downwards from a point in the aircraft known as the centre of gravity, the single weight force (shown as W in Figure 2) having the same effect as the combination of the separate forces.



In order for the aircraft to fly, it is necessary to generate a force that opposes the weight. This force is generated by the aircraft's wings when they move through the air and is called lift. Obviously, each wing contributes half of the total lift generated. However, it is easier to appreciate aerodynamic relationships if the two separate contributions are thought of as combining to form a single equivalent force acting upwards from the centre of the aircraft above the wings, the single lift force (shown as L in Figure 2) having the same effect as the combination of the two separate forces.

Now, in order to confer aerodynamic stability to the flight of the aircraft, it is arranged that the tailplanes each generate a small downward-acting force. Again, it is easier to think of a single force, called the tail down-force, (TDF), having the same effect as the combination of the two separate forces. The explanation of aerodynamic stability will appear later.

In flight at constant speed and height the disposition of the weight force, the lift force and the tail down-force is as shown in Figure 2. Notice that the lift force is slightly greater than the weight force because it also has to oppose the tail down-force.

Whenever any object (such as an aircraft) moves through any fluid (such as air) the fluid tends to resist the movement of the object. This resistance is called drag and it can be considered to be a force acting backwards on the object relative to its direction of movement. As a practical example of drag, the backward force may be felt if a hand is held out of the window of a car moving at speed, the force being exerted by the airflow.

All the components of an aircraft in flight which are exposed to the airflow generate drag. Again, it is easier to think of the separate drag contributions as combining to form a single equivalent force acting backwards from the centre of the aircraft, the single drag force (D) having the same effect as the combination of the separate forces.

By careful design, using the principle of streamlining, an aircraft's drag can be minimised. Streamlining is the designing of components exposed to the airflow so that they are, as far as possible, smoothly shaped and elongated. As an example, refer back to Figure 1 and notice that the engine compartment is carefully shaped so that it joins the fuselage smoothly and that the fuselage itself tapers gradually towards the tail, in marked contrast to the design of a ground vehicle.

The aircraft's drag in flight is opposed by the forward-acting thrust (T) from the propeller, which is rotated by the engine. The importance of streamlining can now be seen - minimised drag requires only modest thrust to oppose it. Therefore sufficient power to provide this thrust can be generated by a small, light, economical engine.

In flight at constant speed and height, the thrust exactly balances the drag of the aircraft. In summary, the disposition of forces acting on an aircraft flying in this manner is as shown in Figure 2.

Note that the thrust and drag forces are considerably less than the lift and weight forces - in modern light aircraft the former two are about one tenth the magnitude of the latter two.

Figure 1 represented a low-winged aircraft. Some light aircraft are high-winged designs, an example of which is shown in Figure 3. A feature of many high-winged designs is the use of struts to improve the load-bearing characteristics of the wings.



The general aerodynamic principles outlined in this section apply to all conventional designs of aircraft.

   





2   THE AIRCRAFT






Figure 1 showed the layout of a typical modern single-engined light aircraft. The aircraft can be considered as consisting of the engine and the airframe, which is the remainder of the machine.

2.1 THE AIRFRAME

In many modern designs the entire structure of the airframe is of metal, usually a light, strong aluminium alloy. The following description refers to such a design.


2.1.1 The wing

The wings are attached to the fuselage, one on each side. Figure 4 is a cut-away diagram showing the main features of the left wing.

The lift generated by the wing in flight is transmitted from the skin to the ribs, and thence to the main spar, which is bolted to the fuselage at the wing root.

The skin is riveted to the ribs, which are bolted or riveted to the main spar. In this type of 'stressed skin' construction, the internal structure has its rigidity and strength enhanced by the attached skin. In other words, the skin contributes to the supporting of loads experienced in flight. For this reason, it is important to ensure before flight that the skin is undamaged, otherwise the structural integrity of the entire assembly is reduced.

Notice that the ribs have holes punched into them, to lighten the weight of the structure. The wing tip is usually of metal or glass-fibre and is attached to the end of the wing to round off its shape.

A section through A-A shows the aerofoil shape of the wing, as shown in Figure 5. Notice that the trailing edge is not fixed rigidly to the main body of the wing, but is mounted on hinged brackets attached to the auxiliary spar, and is therefore moveable. This moveable part is called the flap, and is usually arranged so that it can take up one of three positions, as shown.



The flap on the left wing and that on the right wing are linked, so that both move together. The purpose of the flaps is to alter the aerodynamic characteristics of the wings to suit the various phases of flight.

A section through B-B shows the aerofoil shape nearer the wing tip (Figure 6).



Again, the trailing edge is not fixed rigidly to the main body of the wing, but is mounted on hinges attached to the auxiliary spar. The moveable part is called the aileron and is arranged so that it can move freely up and down within certain limits, as shown.

The aileron on the left wing is linked to that on the right wing so that as one aileron moves up the other moves down, and vice versa.

Notice that, to gain access to the cabin door of a low-winged aircraft, it is necessary to walk on the top surface of the wing near the root, which is especially reinforced for this purpose, forming a walkway. It is important to realise that other areas of the wing must never be walked on, since damage will almost certainly be caused to the skin. This consideration is not, of course, relevant to high-winged designs.

Sometimes, fuel tanks are incorporated in the structure of the wing, as in the example in Figure 4.

The right wing is a mirror image of the left.

2.1.2 The tailplane and elevator

The construction of the tailplane and elevator is similar to that of the wing. Figure 7 shows the left tailplane and elevator.



The assembly is attached to the side of the fuselage near the tail. A section through C-C shows a typical aerofoil shape. The elevator is mounted on hinges and is therefore moveable up and down within certain limits, as shown in Figure 8.



Note that the left elevator is linked to the right, so that they both move up or down together.

Of course, the right tailplane and elevator assembly is a mirror image of the left.

2.1.3 The fin and rudder

The construction of the fin and rudder is usually of the same pattern as the tailplane and elevator, as shown in Figure 9.



The assembly is attached to the top of the fuselage near the tail. A section through D-D shows a typical aerofoil shape. The rudder is mounted on hinges and is therefore moveable to the left and right within certain limits, as shown in Figure 10.



The ailerons, elevators and rudder are collectively referred to as the flight controls, and they enable the pilot to control the motion of the aircraft in flight.

2.1.4 The fuselage

Many modern light aircraft feature stressed skin construction in their fuselage. A typical arrangement is shown in Figure 11.


The stringers, which are thin spars running from nose to tail, and the frames give the structure its basic rigidity and strength, which again are enhanced by the skin. The skin is usually attached by rivets.

At the nose end of the structure is the firewall, made of metal, to which is sometimes attached a sheet of fire-proof material. The firewall separates the engine compartment from the fuselage and its purpose, as its name suggests, is to form a barrier ahead of the cabin area so that, in the event of a fire in the engine compartment, flames are prevented from entering the cabin. Attached to the firewall are the engine mounting points. These are reinforced areas designed to bear the weight of the engine assembly, which is bolted to them.

2.1.5 The landing gear

Most modern aircraft feature a tricycle landing gear, which comprises three units - two mainwheel units and a nosewheel unit. On low-winged aircraft a mainwheel unit is attached to the main spar underneath each wing, near the root. On high-winged designs these units are attached to the lower fuselage beneath the wings. The nosewheel unit is usually mounted underneath the engine compartment. Each unit consists of a leg and a wheel, as shown in Figure 12.



The purpose of the landing gear is to support the aircraft whenever it is on the ground, and also to withstand the loads sustained during landing. It is noteworthy that most of the aircraft's weight is supported by the mainwheel units, and their construction is therefore usually stronger than that of the nosewheel unit.

Figure 12 shows a common type of landing gear unit - the oleo-pneumatic type. The leg features a telescopic construction, at the lower end of which is attached the wheel. Inside the leg are two compartments, one containing compressed air, and the other oil of a special kind. It is the compressed air which supports the aircraft on the ground and absorbs the landing loads. The oil acts as a damping agent - it smooths out the telescopic motion of the leg as the wheel moves up and down. Thus, between them, the compressed air and the oil ensure that the aircraft's structure is not jarred during landing or by manoeuvring on uneven ground.

The unit in Figure 12 also features a torque link, the purpose of which is to ensure that at all times the wheel is aligned correctly with respect to the airframe. Figure 13 shows an oleo-pneumatic unit in various situations.



Each wheel is equipped with a pneumatic tyre. Once the tyre has been fitted, it is usual to paint adjacent marks on the tyre sidewall and the wheel hub, as in Figure 14.



These marks make it easy to detect any movement, called creep, of the tyre around the wheel. Such movement is undesirable since damage might be caused to the tyre valve.

It is normal design practice to arrange that the nosewheel can be turned left and right within certain limits, to steer the aircraft when it is moving on the ground.

Most mainwheel landing gear units are fitted with brakes, usually of the disc type, activated hydraulically. Control of the brakes is often arranged so that the left and right mainwheels may be braked either together (symmetrical braking) or individually (differential braking). Symmetrical braking is used to slow down or stop the aircraft when it is moving on the ground. Differential braking is used to assist with steering the aircraft on the ground.

On higher performance aircraft, provision is sometimes made for the landing gear to be retractable. In other words, the separate units may be folded up into the airframe, thereby minimising drag during flight. With low-winged aircraft, the mainwheel units usually retract into recesses in each wing, whilst on high-winged designs, the recesses are usually incorporated into the fuselage structure. The nosewheel unit retracts either into the engine compartment or into the fuselage. Provision is usually made for folding doors to cover the retracted units, to smooth out the profile of the airframe and hence further minimise drag. In many aircraft the retraction mechanism is activated hydraulically, whilst others may have electrical retraction.

2.1.6 Trim tabs

On many light aircraft, the trailing edge of one of the elevators incorporates a trim tab, as shown in Figure 15.



The tab is attached to the elevator by hinges so that it can move up and down, within certain limits. A section through E-E is shown in Figure 15.

Some aircraft also have trim tabs incorporated in their rudders and ailerons. Figures 15 represents a trim tab which is adjustable during flight. Sometimes, fixed tabs are fitted instead, of simpler construction. They are usually of sheet metal and are attached to the trailing edge of the appropriate flight control surface. They are adjustable only when the aircraft is on the ground. Figure 16 shows a section through a wing which has a fixed trim tab attached to its aileron. In this example the tab has been adjusted slightly upwards.



The aircraft shown in Figure 1  features an elevator trim tab adjustable in flight, and fixed tabs on its rudder and left aileron. This is a common arrangement.

The function of the trim tabs is to assist the pilot in the use of the flight controls.

2.1.7 The cabin

The cabin is incorporated in the fuselage. It has several features - the pilot's controls, seats for the pilot and passengers and stowage for baggage. The cabin is usually enclosed with perspex windows, which are carefully manufactured to be optically correct so that the occupants of the cabin may have undistorted vision through them.

In most designs, access to the cabin is via two doors, one each side, which are fitted with latches to ensure that the doors remain firmly closed during flight.

The load intended to be carried in the cabin (the pilot, passengers and baggage) is governed by two requirements. Firstly, the aircraft's loaded weight must not be greater than the 'maximum total weight authorised' (MTWA) specified in its Flight Manual. Secondly, the disposition of the load must be such that the position of the centre of gravity of the loaded aircraft lies within the limits specified.

2.1.8 Other components

Incorporated in the airframe structure are various components needed to operate the aircraft. They include:

(a) operating linkages for activation of the flight controls, flaps and adjustable trim tabs;

(b) fuel pipelines;

(c) hydraulic fluid pipelines;

(d) electrical wiring.

Engine control linkages are routed through apertures in the firewall into the cabin.

2.2 THE ENGINE

The purpose of the engine is to drive the propeller, which in turn provides the thrust necessary for sustained flight. Nearly all light single-engined aircraft are powered by internal combustion piston engines, in which fuel (petrol) is mixed with air and burnt, the heat energy of combustion then being converted into mechanical energy.

In basic design, these aero-engines are similar to those of cars. In certain respects, however, the two types are quite different. For example, aero-engines are designed to revolve more slowly, thus reducing the internal stresses, and are of sturdier construction than car engines designed to give the same power output. These features minimise the chances of mechanical failure. In this respect the reliability of modern aero-engines is excellent. Other differences occur in the design of the ignition systems and the cooling systems.

In order to appreciate the workings of an internal combustion engine, it is first necessary to take note of certain relevant chemical and physical facts.

2.2.1 Combustion

One fifth of the earth's atmosphere consists of oxygen gas, most of the remainder being nitrogen gas. In the combustion process, the fuel combines chemically with the oxygen present in the air. This process is accompanied by the release of a considerable amount of heat energy. (The nitrogen in the air plays no active part in the combustion process.)

2.2.2 Behaviour of liquids and gases

The working of the internal combustion engine involves several behavioural properties of liquids and gases:

(a) when a gas is compressed, it becomes hotter;

(b) when a fixed volume of gas is heated, its pressure increases;

(c) when a gas flows through a tube which features a constriction, then, at the constriction, the speed of flow of the gas increases, and its pressure decreases (Figure 17). After the gas has passed the constriction, its speed of flow and pressure revert to their original values;

(d) the earth's atmosphere becomes thinner or less dense with increasing height. In other words, a given volume at height contains fewer molecules of air than the same volume at a lower height;

(e) a gas at given pressure becomes less dense as its temperature increases. In other words, a given volume at higher temperature contains fewer molecules of gas than the same volume at lower temperature;

(f) as a liquid evaporates, it undergoes cooling of itself and its surroundings.



2.2.3 The four-stroke cycle

The power that an internal combustion engine delivers is developed in its cylinders. Figure 18 shows the component parts associated with one of the cylinders.

Usually the components are of steel, except for the cylinder head, piston and crankcase, which are usually of aluminium alloy.



Consider now the sequence of events occurring at this cylinder when the engine is operating. As the engine turns, so the piston moves up and down in the cylinder. Notice that the piston is attached to one end of the connecting rod, the other end of which is attached to the crank of the crankshaft. It can be seen that, by this arrangement, reciprocating (up-and-down) motion of the piston is converted into rotational motion at the crankshaft. The crankshaft drives the propeller.

In Figure 18, the piston is moving downwards. Notice that the inlet valve at the cylinder head is open, allowing fuel-air mixture from the inlet manifold to be drawn into the cylinder. The exhaust valve is closed. A short time later, as shown in Figure 19, the piston is moving upwards. Since the inlet valve and exhaust valve are now both closed, it will be appreciated that the fuel-air mixture is being compressed. The compression has the effect of heating the mixture.



As the piston reaches its highest position in the cylinder, the spark plug causes a spark to occur at the cylinder head. The spark ignites the inflammable mixture, which, because it is hot and compressed, burns quickly and thoroughly. The heat released during combustion goes to increase considerably the pressure of the gases inside the cylinder. In consequence, the piston is forced down again, giving a rotational impulse to the crankshaft, as shown in Figure 20.



A short time later, the piston starts to move up once more, propelled by the still-rotating crankshaft, as shown in Figure 21. Notice that the inlet valve is closed and the exhaust valve is open, allowing the piston to drive away the spent gases into the exhaust manifold and thence to the exhaust pipe, from which they disperse into the atmosphere.



The cycle of events then repeats itself. This sequence is called the four-stroke cycle, and Figures 18, 19, 20 and 21 represent respectively:

(a) the induction stroke;

(b) the compression stroke;

(c) the power stroke;

(d) the exhaust stroke.

It should be appreciated that the piston is delivering power only during the power stroke. The impulse given to the crankshaft during this stroke is sufficient to ensure that it continues to rotate, propelling the piston accordingly, during the sequence leading up to the next power stroke.

In a four-cylinder engine, the arrangement is such that, during two complete rotations of the crankshaft, each of the four pistons delivers one power stroke; in other words, the crankshaft experiences a power impulse every half-revolution. This feature enhances the effect mentioned above, ensuring that, at any particular moment, the continuously-rotating crankshaft is propelling those pistons which are not involved in delivering a power impulse.
 
Each valve is held in the closed position by a valve spring. The valve is opened, against the pressure of the spring, by an operating mechanism at appropriate moments during the four-stroke cycle. This operating mechanism is driven mechanically by the crankshaft, thereby ensuring that the valve opens and closes at the correct moments in time in relation to the rotation of the crankshaft.

A typical layout for a four-cylinder aero-engine is the horizontally-opposed arrangement, shown schematically in Figure 22, in which the cylinders are arranged in two pairs, disposed on opposite sides of the crankcase.


The view from above shows the offset disposition of the cylinders. This is a common feature of horizontally-opposed engines and is to allow each piston to be connected to an individual crank on the crankshaft.

2.2.4 The carburettor

The carburettor is the means by which liquid fuel and air are mixed in a satisfactory manner prior to combustion. Figure 23 is a schematic representation of a float-type carburettor, used extensively in small aero-engines.



The carburettor mixes the fuel and air in the correct ratio for complete combustion of the fuel. The mixture in the inlet manifold consists of air, vaporised fuel and tiny fuel droplets. From the inlet manifold the mixture is drawn into whichever cylinder is undergoing its induction stroke at that moment. This cylinder will have its inlet valve open - all the others will have theirs closed.

As the mixture is taken from the inlet manifold, so air is drawn into the intake tube of the carburettor to replace it. In many designs the incoming air passes through a filter whose function is to remove any foreign matter such as dust or grit which might otherwise damage the engine if ingested. Notice that the intake tube features a streamlined constriction, sometimes called the choke. The air flowing past it undergoes a reduction in pressure (as was shown in Figure 17) which has the effect of drawing fuel from the fuel nozzle, located in the intake tube at the constriction. As it is drawn from the nozzle by the air speeding past, the fuel is broken up by the air into droplets. In other words, it is atomised. The smaller droplets evaporate in the air - the larger ones are carried along with it. It can be seen that, by this process, new mixture has been created to replace that taken into the cylinders.

As fuel is drawn from the nozzle, the level of fuel in the float chamber drops and so the float is no longer forcefully held up. Reference to Figure 23 shows that when this happens, the float valve on the other side of the pivot will rise under the pressure of the fuel from the pump. Accordingly, more fuel will enter the float chamber until the level has risen sufficiently, urging the float to rise with it, for the float valve to be lowered again, thereby cutting off the flow of fuel.

The cycle then repeats itself. In this way, the float chamber is constantly replenished with fuel to replace that drawn from the nozzle. In practice, when the engine is running, a steady state of affairs is reached, as in Figure 23, with the float slightly lowered from its highest position and the float valve slightly raised, permitting fuel to flow into the float chamber at the same rate as it is drawn from the nozzle.

2.2.5 The throttle

The throttle is the means by which the power output from the engine is controlled.

The amount of mixture passing from the carburettor into the inlet manifold is controlled by the throttle valve. The valve, shown in Figure 23, is in the form of a circular plate located in the intake tube just above the fuel nozzle. The plate can pivot about its centre, thereby effectively varying the area of the intake tube through which the mixture passes and hence controlling the amount of mixture supplied to the inlet manifold.

The position of the throttle valve is controlled by the throttle lever, to which it is linked. When the lever is used to make the valve pivot so as to cut off almost completely the mixture supply to the cylinders, as shown in Figure 24, the throttle is said to be closed.



In this situation, after combustion, the increase of pressure of the gases in the cylinders is not very great, and so the pistons deliver only weak power impulses. In other words, the power output from the engine is minimal.

As the lever is used to open the throttle progressively, it causes the valve to pivot so as to allow more mixture to enter the cylinders for combustion, and the pistons therefore deliver stronger power impulses - the power from the engine increases. Figure 23 shows an intermediate throttle setting.

When the throttle is fully open, therefore, the engine delivers its maximum power, with the valve as shown in Figure 24.

2.2.6 Mixture strength and mixture control

It has been stated that the carburettor mixes the fuel and air in the correct ratio for combustion. Obviously, it must fulfil this function regardless of the setting of the throttle valve. When the valve is set to give low power output it partially restricts the air flowing through the intake tube, as described above. Because the flow of air is restricted, the pressure reduction at the constriction is less marked, and so less fuel is drawn from the nozzle. Conversely, the throttle valve set to give high power output allows a strong flow of air through the intake tube, with the result that the pressure reduction at the constriction is more marked and so more fuel is drawn from the nozzle. Thus it can be seen that the very design of this type of carburettor automatically matches the fuel supply to the amount of air entering the intake tube.

Under certain circumstances, however, the fuel supply must be modified for reasons which are explained below, and, accordingly, the carburettor is designed to be able to bias the mixture either with extra fuel (giving a rich mixture strength) or with fuel (giving a lesslean mixture strength). Very lean mixtures are referred to as being weak.

When the engine is operating at high power output, for example during take-off, the carburettor is made to supply a rich mixture. The extra fuel, when it vaporises, helps to cool the mixture. This ensures that, after being heated during the compression stroke, the mixture is not above the right temperature for correct combustion. In addition, the cylinders and pistons are prevented from becoming overheated. Note that the purpose of the extra fuel is solely to cool the mixture by evaporation - it does not increase the engine power output, since all that happens during the combustion process is that the available oxygen in the air is shared by however much fuel there happens to be in the mixture.
 
With increasing height, the earth's atmosphere becomes less dense. If an aircraft is cruising at height, therefore, the carburettor will be taking in this less dense air. However, the air pressure reduction at the streamlined constriction is still considerable - enough to draw out fuel from the nozzle at a rate which, because of the reduced density of air, would yield a mixture over-rich in fuel. This over-rich mixture would cause poor fuel economy - that is, less distance flown through the air for each litre of fuel used - and possible rough-running of the engine.

To prevent this, the carburettor is fitted with a mixture control, represented schematically in Figure 23. The control takes the form of a variable restriction in the fuel supply from the float chamber to the nozzle. To correct over-richness at greater heights the control is operated to restrict the flow of fuel to the nozzle, thus bringing the mixture to the correct strength for the prevailing density of air. Appropriate use of the mixture control will therefore improve the fuel economy and ensure smooth running of the engine.

The mixture control is progressive in its operation - when positioned so that the fuel supply is unrestricted (the 'rich' position), as in Figure 23, the carburettor would give the rich mixture necessary for high engine power output, if needed, whilst for cruising flight, the control would be used as necessary to give the correct mixture strength at the chosen height - the greater the height, the more the control should be moved away from the 'rich' position to restrict the fuel supply further. This process of correcting the mixture strength for height is called 'leaning out' the mixture. Figure 25 shows the mixture control set to an intermediate position.



The position of the mixture control is set by a lever to which it is linked.

Note that, for any particular throttle setting, regardless of whether the mixture is correctly leaned out or not, an engine equipped with the type of carburettor described here delivers less power at greater heights than it does at lower heights. This is because a cylinder-full of mixture at greater height contains fewer molecules of oxygen available to combine with the fuel.

Usually, on small aero-engines, the mixture control is not automatic in operation - it is set as necessary by the pilot and is adjustable in flight. Because of this, care must be taken to ensure that the control is not used to make the mixture too weak when a richer mixture is required, for example at high engine power output or when cruising at low height.

If the throttle is opened very quickly, the throttle valve permits a sudden increase in the flow of air entering the intake tube. However, because of its inertia, the supply of fuel from the float chamber to the nozzle cannot immediately match the demand, with the result that temporarily fuel starvation occurs, and an over-weak mixture is supplied to the inlet manifold. This effect is only momentary because soon the fuel flow overcomes its inertia and is able to supply the high demand at the nozzle. However, with some carburettor designs, the weakness of mixture may be sufficiently marked to cause the engine to cease to deliver power temporarily, a dangerous consequence which is obviated by incorporation in the carburettor of an accelerator pump. This device is activated mechanically when the throttle is opened, and causes an extra squirt of fuel to be supplied to the intake tube for a sufficient length of time to prevent the weakening of the mixture that would otherwise occur. Figure 23 includes a schematic representation of the accelerator pump.

2.2.7 Carburettor heat control

At the fuel nozzle, the evaporation of fuel droplets has the effect of cooling the intake tube to the extent that, if the air is sufficiently humid, ice may form in the intake tube. This phenomenon is called 'carburettor icing' and can occur even in warm ambient temperatures if the air is sufficiently moist. If the build-up of ice is heavy, it can block the intake tube so much that the supply of air passing through it is impeded, causing the engine to lose power and run roughly. In severe cases, total loss of power is possible.

To guard against this occurrence, a carburettor heat control is fitted. When selected 'on', the control shuts off the normal supply of air to the intake tube whilst at the same time opening an alternative supply. The air from this alternative supply is pre-heated, usually by passing through a muff fitted round the engine's exhaust pipe. When it enters the carburettor, this pre-heated air melts any ice which may be present. The resulting water then passes through the engine to be ejected with the exhaust gases. Figure 26 shows the control selected 'on'.


 
When the control is 'off', the pre-heated air supply is shut off and the normal supply of unheated air comes back into use, as in Figure 23.

Note that whenever the heat control is 'on', the engine delivers slightly less power for a particular throttle setting than when the control is 'off'. This is because the pre-heated air is less dense than the unheated air and so a cylinder-full of mixture contains fewer molecules of oxygen available to combine with the fuel. For this reason, the engine is normally operated with the control 'off'.

The control is selected 'on':

(a) for periods of a few seconds every few minutes or so, to ensure that any ice in the carburettor is dispersed;

(b) if the engine develops rough-running or loss of power and it is suspected that carburettor icing may be causing the malfunction. In this case, the control is left 'on' until all the ice has been dispersed;

(c) on some designs of engine whenever low power settings are in use for protracted periods of time. These engines are susceptible to icing around the throttle valve when it is set for low engine power.

Note that, if the control is used as described in (a), then it is unlikely that carburettor icing would build up to the extent that loss of power or rough-running of the engine occurred - prevention is better than cure.

The position of the control is set by a lever to which it is linked.

2.2.8 Detonation

During the power stroke of the four-stroke cycle, the mixture burns steadily and evenly and the increase of gas pressure on the piston is therefore smoothly progressive. However, if the engine controls are not being used correctly, it is possible to arrive at a situation where the temperature of the mixture just prior to ignition is too high for normal combustion. Instead of smooth, progressive combustion, the over-heated mixture in the cylinder burns explosively, resulting in a sudden harsh increase in pressure that can strain the piston and possibly even damage it.

This explosive combustion is called 'detonation' and can occur for one of three reasons:

(a) incorrect mixture strength. If the throttle is set to give high engine power output, then the mixture must be rich, as already mentioned in 2.2.6, to help to keep it cool. If the mixture is too weak, its temperature may increase during the compression stroke to the extent that, after ignition, detonation occurs;

(b) incorrect use of carburettor heat control. With the control 'on', the mixture supplied to the cylinders has a higher temperature than with the control 'off'. At high engine power output, this increase in temperature, which is raised further during the compression stroke, may be sufficient to cause detonation;

(c) incorrect grade of fuel.

Notice that detonation is most likely to occur at high engine power output, and can be avoided in these circumstances by ensuring that the mixture control is set to 'rich' and that the carburettor heat control is 'off'.

2.2.9    Fuel injection

Some aero-engines, usually those of higher power output, feature a fuel injection system instead of a float-type carburettor. These injection systems are considerably more complicated and therefore more expensive, but they do confer several advantages, not the least of which is the elimination of the possibility of carburettor icing.

A detailed description is outside the scope of this book. The basic operating principle is that fuel is injected into the air entering each cylinder during its induction stroke. Otherwise the operation of the engine is similar to that already described.

2.2.10 The ignition system: the magneto

The ignition system is designed to supply electric current to the spark plugs in the cylinder head. The current is generated mechanically by the magneto, which is a modified form of electric dynamo.

2.2.10.1 Mechanical generation of electricity

Mechanical generation of electricity makes use of the interaction of a rotating magnet with a coil of metal wire. A schematic arrangement is shown in Figure 27. Alternatively, the coil can be designed to rotate inside a magnet array.



In either case, electric current is generated in the coil as long as rotation occurs. When rotation stops, no current is generated.

2.2.10.2 Conversion to high voltage

The device in Figure 27 can be considered as an 'electricity pump', and accordingly, the electricity delivered from it has a 'pressure' (in the same way as a water pump delivers water under pressure). This pressure is termed voltage. Not surprisingly, the voltage depends on the speed of rotation of the magnet - faster rotation gives greater voltage.

In an aero-engine magneto, the spindle to which the magnet is attached is rotated by the engine's crankshaft via a suitable arrangement of gear wheels.

The voltage produced by the coil when the engine is running normally is very considerably less than that required by the spark plugs, and so the electric current has to be modified. The modification is achieved by two devices - the contact-breaker and the transformer. The contact-breaker interrupts the electric current from the coil at suitable intervals. Each interruption causes an electrical interaction in the transformer (which is merely a particular arrangement of coils of metal wire), producing a pulse of high voltage electric current from it.

2.2.10.3 Timing and distribution

The high voltage pulses are supplied to the spark plugs in the cylinder heads. Obviously, the pulses must be made to occur at precise moments, so that, in each cylinder, the spark occurs as the piston reaches its highest point in the cylinder at the end of the compression stroke.

The correct timing is ensured by the contact-breaker, which is operated mechanically by the spindle, causing interruptions in synchronisation with the spindle's rotation. By this means, the pulses from the transformer are automatically made to occur at precisely the correct moments during the rotation of the crankshaft, regardless of the speed of rotation.

All that remains is to ensure that the pulses are delivered to the individual spark plugs in the correct sequence. This is done by the distributor. The distributor has a rotating arm, driven round by the spindle, which directs the pulses from the transformer to the spark plugs via their respective ignition leads, which are suitable lengths of heavily insulated electrical cable.

The term 'magneto' applies to the entire assembly of components described above. The components are arranged such that the spindle can drive all the rotating parts. When the magneto is bolted to the engine's crankcase, its spindle connects with gear wheels driven by the crankshaft.

A schematic representation of the electrical set-up is shown in Figure 28.



2.2.10.4 The spark plug

Figure 29 shows a section through a spark plug screwed into its cylinder head.



When the high voltage pulse is sent to the spark plug from the distributor it travels down the centre electrode and then, in the form of a spark, it jumps across the small gap to the 'earth' electrodes. (All the electrodes are of metal.) From the earth electrodes the electric current flows to the outside case of the plug and then away via the cylinder head into the main body of the engine, where it dissipates. The ceramic insulation prevents the pulse from leaking across to the plug case before it reaches the end of the centre electrode.

The electric pulses travelling from the distributor to the spark plugs would cause interference on the aircraft's radio equipment, and to prevent this, the ignition leads are 'screened' - in other words, they have an outer sheathing of finely-woven metal wire.

2.2.10.5 Dual ignition systems

Most aero-engines are equipped with two magnetos. In these dual ignition systems, each cylinder has two spark plugs, one supplied by each magneto, figure 30 shows a schematic arrangement for a four-cylinder engine.



There are two advantages of dual ignition systems:

(a) in the event of failure of one of the magnetos, the other will still supply one spark plug in each cylinder with high voltage pulses;

(b) with both magnetos operating, the mixture in the cylinders is ignited at two different locations, which makes for more efficient combustion, and therefore better power output from the engine.

Note that these aero-engine ignition systems are entirely self-contained and need no external supply of electricity. (In this respect they differ from car engine ignition systems, which require the battery to supply the electrical current to the transformer.)

2.2.10.6 The impulse magneto

When the engine is being started from rest, the crankshaft, and therefore the rotating components of the magnetos, are turning much more slowly than when the engine is running normally. Because of this, the high voltage pulses are much weaker than normal - so weak that they may not be able to jump from the centre electrodes of the spark plugs to the earth electrodes. If this is the case then no sparks occur and so no ignition takes place.

To overcome this problem, it is arranged that one of the magnetos of the dual ignition system, termed the impulse magneto, has its spindle rotated, not steadily, but in impulses. This impulse drive is achieved mechanically by an arrangement of springs and weights within the magneto, and occurs whenever the spindle is rotating very slowly, as during start-up. During each impulse, the spindle momentarily rotates faster - enough to ensure that sparks occur at the cylinder heads.

Once the engine is running normally, the magneto automatically reverts to direct drive

2.2.10.7 Ignition control

Each magneto is controlled remotely by its own magneto switch. When the switch is on, the contact-breaker causes interruptions in the electric current from the coil and the magneto functions normally.

When selected off, the switch allows the electric current from the coil to bypass the contact-breaker. Accordingly, no interruptions occur and no high voltage pulses are delivered to the spark plugs. If both magnetos are switched off, combustion of the mixture cannot be initiated and so the engine no longer delivers power.

Figure 31 shows a magneto switch diagrammatically.



2.2.10.8 Electronic ignition systems

A recent development in aero-engine technology is the replacement of conventional magnetos by electronic ignition systems. These have several advantages, including fewer mechanical components (thereby reducing maintenance requirements) and more precise timing of the high voltage pulses delivered to the spark plugs, which promotes more efficient fuel combustion and thereby enhances engine power output and fuel economy.

2.2.11 The oil system

It is apparent that the internal combustion engine has many moving parts. When the engine is operating, there would be friction between these parts as they moved against each other, causing excessive wear and overheating, were it not for the presence of a lubricant. The lubricant used is oil, an adequate supply of which is delivered to all moving parts in the engine.

Wherever the machinery has to withstand high stresses, as in the crankshaft bearings and connecting rod bearings, the oil is delivered under pressure. (The crankshaft bearings are the housings in the crankcase which support the crankshaft. The connecting rod bearings are the housings at the end of the connecting rods in which the cranks rotate.) The pressure feed ensures that a film of oil is forced in between adjacent metal surfaces to prevent their contact.

Other moving parts where the stresses are not so great are lubricated either with oil under pressure or by oil spray. On spray-lubricated parts, an adequate oil film forms without the need of pressure. The layout of a typical oil system is shown in Figure 32.



The sump, in which the oil is stored, is a deep metal tray forming the lower part of the crankcase. From the sump the oil, still hot from its circulation through the engine, is drawn by the pump through the suction screen, which filters foreign matter from the oil. From the pump the oil is piped, now under pressure, to the oil cooler. In the cooler the oil passes through a matrix of thin tubes which are exposed to the airflow experienced by the aircraft in flight, enabling the oil to lose much of its heat to the airflow.

The oil is now delivered to the pressure screen, where it is again filtered. On the way, automatic measurements of the temperature and pressure of the oil are made and transmitted to calibrated gauges in the cabin.

From the pressure screen the oil is directed through drillings in the crankcase to the crankshaft bearings and, through holes drilled in the bearings, into the hollow crankshaft, from which it is fed to the connecting rod bearings, again via holes drilled in the bearings. From these bearings the oil forces its way out in the form of a spray.

The resulting oil spray lubricates the cylinder walls and the piston bearings housing the other ends of the connecting rods. Part of the spray is collected in oil channels through which it flows to lubricate other moving parts in the engine.

The inlet and exhaust valve operating mechanism is usually lubricated by oil under pressure, taken from the pressure supply.

Having fulfilled its lubricating function, the oil drains back into the sump.

Note that, in order to ensure an adequate supply at all times when the engine is operating, the oil pump is designed to deliver more than is needed. The pump is driven by the crankshaft via a series of gear wheels. Hence faster rotation of the crankshaft causes the pressure of oil delivered from the pump to rise. When the maximum designed pressure is exceeded, the surplus oil is diverted by the pressure relief valve back into the sump.

When the oil is very cold, for example before the engine is started in wintry weather, it is so thick that it does not flow easily through the fine tubes of the cooler. Obviously, in this case, the cooler is acting as an obstruction to the flow of oil, and there is a danger that its tubes might be damaged by the oil under pressure. However, the pressure build-up caused by the obstruction is sensed by the by-pass valve, which opens to allow the oil to by-pass the cooler. Once the engine has warmed up, the now thinner oil can flow more easily through the cooler and the by-pass valve closes to make sure that it does so.

The oil system of some designs of engine may differ in certain respects from the arrangement described here, but the basic operating principle is the same in all designs.

The oil system also serves a secondary purpose - that of cooling the inner parts of the engine, which give up their heat to the oil as it flows through. In turn, the oil loses the heat as it passes through the cooler.

2.2.11.1 Oil quantity and consumption

Not all of the oil returns to the sump after it has been pumped round the engine - some escapes past the pistons and is burnt along with the fuel-air mixture. This loss is referred to as the engine's oil consumption. If the level of oil in the sump drops below a certain level as a result of the loss, then the pump may not be able to deliver an adequate supply to the engine. The consequent oil starvation would cause, initially, overheating of the engine because of the increased friction, and eventually, mechanical failure of the engine.

Because of these dangers, the manufacturer will specify the minimum permitted quantity of oil required in the sump and also the maximum permitted rate of consumption (in terms of amount consumed per hour of engine operation).

Before every flight, therefore, the pilot is responsible for ensuring that the quantity of oil in the sump is at least equal to the minimum permitted quantity plus the amount that would be consumed during the flight. An engine which is found to lose oil more quickly than the maximum permitted rate must be withdrawn from service until the fault is remedied.

In most designs, the quantity of oil in the sump is measured with a calibrated dip-stick, which is often integral with the filler cap. If the dip-stick indicates a quantity less than that needed for the flight, extra oil is added through the filler opening.

2.2.11.2 Oil grade

The oil used for engine lubrication is refined to a very high quality. It is also graded according to its thickness, or viscosity. Oil of high viscosity is thick and treacly; that of low viscosity is thin and runny.

Oil suffers from the drawback that its viscosity reduces as its temperature increases. At very high temperatures, the oil may become so thin that the lubricating films in the engine may break down, with consequent detriment to the machinery. Conversely, at very low temperatures, the oil becomes very thick, so much so that it may be difficult to turn the engine for starting. Additionally, once the engine is running, the oil pipes may be subjected to excessive strain.

To overcome these problems, oil grades are chosen to suit the ambient conditions. For example, during summer-time operations, oil of higher viscosity is used to that as the engine reaches its normal operating temperature, the oil is still thick enough to be able to perform its lubricating functions adequately.

In winter, use of a lower viscosity oil will ensure that engine starting is not hampered and that the oil system is not overstrained when the oil is cold. On some engine designs, it can be arranged that during cold weather, the flow of air through the oil cooler is restricted to assist the oil to warm up to a satisfactory working temperature.

The engine manufacturer will specify the grades of oil to be used according to ambient temperature, and most engines have this information placarded adjacent to the oil filler cap for ease of reference. From what has been stated above, it is apparent that use of an incorrect grade is potentially harmful to the engine.

2.2.11.3 Dry-sump systems

The oil system represented by Figure 32 is termed a wet-sump system, because its oil is stored in the engine sump. This layout is found in many training and touring light aircraft. However, some machines are designed to be aerobatic - that is, capable of performing extreme manoeuvres such as loops and rolls.

Obviously, for these aerobatic types, the wet-sump oil system would not be satisfactory, because the oil would be thrown around during manoeuvres, with a danger that the supply to the oil pump might be interrupted. This drawback is overcome by storing the oil in a separate tank designed to ensure an adequate supply to the pump regardless of the motion of the aircraft. After being pumped round the engine the oil drains to the sump and is removed by a scavenge pump which returns it to the tank. This arrangement is called a dry-sump system.

2.2.12 Engine cooling

Ideally, all the heat energy generated in the cylinders during combustion of the fuel would be converted by the engine into mechanical energy, but this is not achievable in practice. Much of the heat energy is wasted - some is taken away by the exhaust gases and some goes to raise the temperature of the cylinders and their associated components. If nothing were done about it, these parts would rapidly heat up to the extent that the lubricating oil films would break down. When this happens, of course, the eventual consequence is mechanical failure of the engine.

Obviously then the heat must be taken away before it can do any damage - the engine must be cooled.

Most aero-engines are air-cooled. The cylinders and cylinder heads are made with integral metal fins surrounding their outsides, as shown in Figure 33.



When the aircraft is in flight, part of the airflow that it experiences is made to flow through the fins, which give up the heat they have taken from the cylinders to the passing air. By this means the working temperature of the engine is kept at a reasonable level.

In a multi-cylinder engine the cooling airflow is directed by baffles such that each cylinder receives an adequate supply.

When the aircraft is on the ground with its engine running, it is apparent that there will be no strong airflow past the fins, and so it is not easy for the engine to get rid of its excess heat. To a certain extent the situation is alleviated by the effect of the air thrown back by the rotating propeller. Some of this air flows through the fins, helping to take away the heat. Nevertheless, it is essential that to prevent overheating in these circumstances the pilot exercises care in the use of his engine controls, avoiding high power settings whenever possible.

2.2.13 The propeller

The propeller is an assembly of specially-shaped blades which is rotated by the engine to provide thrust. In section, the blades have a similar aerofoil shape to the wings, and they generate thrust in the same manner as the wings do lift.

The most common design for light aircraft is the two-bladed fixed-pitch propeller, usually made of aluminium alloy. A fixed-pitch propeller is one in which the blades are attached to the hub at a fixed angle.

The propeller is bolted to the end of the crankshaft. To streamline the hub, and thus minimise drag in flight, the propeller is usually fitted with a spinner, which is a shaped metal cone bolted to the hub. Figure 34 shows a typical propeller design.



2.2.14 Engine mounting

The entire engine assembly with all its components and accessories (which will be described later) is housed in a rigid framework of strong steel tubing. This in turn is bolted to the engine mounting points on the firewall at the nose end of the fuselage structure.

2.2.15 The cowling

Figure 1 shows that the engine compartment is enclosed by a cowling, the main purpose of which is to streamline the shape of the nose end of the aircraft to minimise drag in flight. The cowling usually incorporates intakes from which air is ducted for engine cooling purposes. This air, having taken heat from the fins of the cylinders, flows out again from vents in the rear of the cowling. With some designs, the vents are partly covered by cowl flaps, which are opened to maximise the cooling airflow whenever engine overheating is most likely to occur, as during ground running and while flying at low speed with the engine at high power. In cruising flight, with power reduced and the aircraft at higher speed, the cowl flaps can be closed to restrict the airflow and so prevent the engine from being cooled below its designed operating temperature.

Other intakes in the cowling supply the oil cooler and the carburettor with air.

In modern machines the cowling is either of aluminium alloy or of glass-fibre and features hinged portions which can be opened for engine inspection before flight.

2.2.16 The tachometer

The tachometer enables the pilot to assess the power output from the engine. It is mounted on a panel in front of the pilot called the instrument panel, and indicates the speed of rotation of the crankshaft in terms of revolutions per minute (RPM). Of course, the propeller RPM are exactly the same, since the propeller is bolted to the crankshaft.

The tachometer is driven either mechanically or electrically. Its presentation on the instrument panel is as shown in Figure 35.



The tachometer gives a direct indication of engine power output, higher RPM signifying higher power output. To appreciate why this is so, remember that when the throttle is opened, it causes the pistons to deliver stronger power impulses to the crankshaft, forcing it to rotate more quickly.

The red-line marked on most tachometers indicates the maximum permitted engine RPM. In the example shown the red-line is at 2700 RPM.

2.2.17 Engine-driven systems

Besides its chief task of rotating the propeller, the crankshaft must also drive other components. Some of these are integral with the engine and have already been mentioned, namely:

(a) the inlet and exhaust valve operating mechanism;

(b) the oil pump and fuel pump;

(c) the magnetos.

Additionally, external accessories are powered by the crankshaft. They are:

(a) the alternator in the electrical system;

(b) the suction pump which activates the gyroscopic flight instruments;

(c) the hydraulic system pump.

Most light aircraft are equipped with electrical and suction systems. Engine-driven hydraulic systems are most likely to be found on machines incorporating retractable landing gear.

It is usual for the suction pump and hydraulic pump (where fitted) to be driven directly via a series of gear wheels. In contrast, the alternator is usually driven by a belt of reinforced rubber which passes round two pulley wheels, one on the alternator, the other on the crankshaft (as in car engines).

2.2.18 FADEC systems

We have seen that some aero-engines feature fuel injection systems for delivering fuel to the cylinders rather than carburettors and noted that a benefit of these systems is the elimination of the possibility of carburettor icing. We have also seen that electronic ignition systems offer several advantages over conventional mechanical magnetos. A further development which enhances engine efficiency is the Full Authority Digital Electronic Control (FADEC) system fitted to some new engines.

In a FADEC engine, sensors measure various parameters and send this data to a computer, which controls both the flow of fuel fed to the cylinders and the precise moment in the four-stroke cycle that the spark is triggered in the spark plugs, further improving engine efficiency and power output.

The parameters sensed include:

(a) the throttle lever position;

(b) actual engine RPM;

(c) ambient air pressure and temperature;

(d) air pressure in the inlet manifold;

(e) the temperature of the cylinder heads;

(f) the temperature of the exhaust gases.

In aircraft fitted with FADEC engines, then, there is no need for a mixture control lever or a carburettor heat control lever. Usual practice is for two independent computers to be incorporated in the system so that normal engine operation is retained even after failure of one of the computers.

2.3 THE FUEL SYSTEM

The fuel used by the aircraft's engine is stored in tanks, from which it is fed to the carburettor by the fuel pump. Figure 36 shows a typical arrangement for a fuel system having one tank in each wing. In other designs, tanks may be located in the fuselage.



Every tank has a vent to allow air to replace the fuel taken from it. In this way there is no danger of suction building up in the tank, which might prevent more fuel from being drawn from the tank by the fuel pump.

The fuel cock selects the tank(s) from which the fuel is taken at any particular time. With some designs it is only possible to feed from one tank at a time. Others allow fuel to be drawn simultaneously from more than one tank. The fuel cock also enables the fuel supply to the engine to be completely cut off when required.

From the fuel cock the pipeline leads to the engine-driven fuel pump, which pumps fuel to the carburettor whenever the engine is running. On the way, automatic measurement of pressure is made and transmitted to a gauge in the cabin.

To guard against the possibility of fuel starvation caused by failure of the engine-driven pump, many systems incorporate an electrically-driven pump as a back-up. This is switched on as necessary in accordance with the manufacturer's recommendations.

It is usual to include fuel drains in the system at appropriate points. Before flight, or as recommended, the drains are opened briefly to allow any water or sediment which may have accumulated to be expelled. As an extra safeguard, the fuel is filtered by a strainer before it reaches the pumps.

On some engine designs, a hand-operated primer is used for engine starting. When operated, the primer injects neat fuel into the inlet manifold adjacent to each cylinder inlet to ensure an adequate supply during the starting procedure.

2.3.1 Fuel quantity and consumption

The quantity of fuel in each tank is automatically measured by electrical apparatus and the measurements are transmitted to calibrated gauges in the cabin. However, the pilot should visually check the quantity in each tank before flight to satisfy himself that the gauges are indicating reliably. (This procedure is not possible with some designs of aircraft.)

The tanks are replenished as necessary through the filler openings. During this operation, precautions must be taken to minimise the attendant fire hazard. These include the prohibition of smoking and naked flames in the vicinity of the aircraft.

The rate at which the engine consumes fuel depends on two factors:

(a) the throttle setting;

(b) the mixture control setting.
 
Consumption is highest when the throttle is fully open and the mixture control is set to rich. The aircraft's Flight Manual will detail the rates of consumption for various settings of the engine controls. It is the responsibility of the pilot to ensure that the tanks contain sufficient fuel (including reserves) for the intended flight.

A point to bear in mind is that, during flight, the total quantity of fuel remaining in the tanks is always decreasing. Were this not the limiting factor, the aircraft could be made to fly indefinitely (or at least until the engine had used up its oil supply, which would take considerably longer than the exhaustion of the fuel supply).

2.3.2 Fuel grade

Aviation fuel is refined to a very high standard of purity. It is graded according to its ability to resist detonation, and the ability is expressed as an 'octane rating' (a chemical specification). Fuel with good resistance is said to have a high octane rating, and vice versa.

Pure petrol ('gasoline' in American parlance) has a relatively low octane rating (about 90). To improve its resistance to detonation various chemicals are added, the most common being tetra-ethyl-lead. Fuel containing this ingredient is said to be 'leaded'. Unfortunately, leaded fuels have the drawback that their combustion products tend to cause fouling of the cylinders, valves and pistons, and so the quantity of tetra-ethyl-lead is kept to the minimum necessary to improve anti-detonation resistance to an acceptable level. A commonly used grade of leaded aviation fuel for light aircraft is 100LL. (100 is the octane rating and 'LL' stands for 'low lead content'.)

Some aero-engines require the use of unleaded fuels - they are able to resist detonation by their very design. Many such engines use UL91 aviation fuel. (UL stands for 'unleaded' and 91 is the octane rating.)

The generic term 'avgas' (aviation gasoline) refers to various petrol-based types of fuel constituted for use in aero-engines. Some engines are cleared for operation using 'mogas' (motor gasoline), which has slightly different chemical and physical characteristics compared to avgas. Its chief advantages is that it is usually cheaper than avgas.

It is clear that only the correct grade of fuel must be used for any particular engine, in accordance with the manufacturer's instructions. If leaded fuel is used in engines for which unleaded fuel is specified, there is a danger that the components mentioned above will become badly fouled by the combustion products. Conversely, engines requiring leaded fuel are designed to resist this fouling. If unleaded fuel is used in them detonation will occur at high power settings (with consequent possible damage to the pistons). For convenience, it is the usual practice to placard an aircraft's required grade of fuel adjacent to the tank filler openings, so that mistakes can be avoided.

2.4 THE ELECTRICAL SYSTEM

Most modern light aircraft require a supply of electricity to power various ancillary services. This supply is generated mechanically by the alternator.

2.4.1 The alternator

In the alternator, a coil of wire rotates in a magnetic environment, or field. In contrast to the magneto, which uses magnets to produce the field, the alternator achieves the same effect with an array of stationary coils of wire. When electric current is made to flow through the stationary coils they produce a magnetic field. If now the moving coil is rotated inside the array of stationary coils, it generates current. Most alternators make use of the principle of self-excitation, in which some of the generated current is made to flow through the stationary coils to set up the magnetic field. This arrangement is preferred to a magnet assembly because it makes it easier to control the electrical pressure, or voltage, of the alternator's output.

The spindle on which the moving coil is mounted is driven by the engine via a reinforced rubber belt which passes round two pulley wheels - one attached to the spindle and the other to the crankshaft. Whenever the engine is operating, therefore, the alternator is able to generate electric current.

Because of its design the alternator produces what is called alternating current (AC) - the current flows from the coil alternatively in one direction and then the other. In this form the current cannot be used by the ancillary services - it must be rectified, or made to flow in one direction only, as direct current (DC). The conversion of AC to DC is achieved by an electronic device called the rectifier.

The DC is supplied to the ancillary services as needed.

2.4.2 The electrical circuit
 
Figure 37 shows a simple electrical circuit in which a battery is made to power a light bulb.



When the switch is on, the battery forces current through the bulb, making it light up - the arrows show the direction of the current flow. If the circuit is broken by putting the switch off, no current can flow and the light goes out.

Figure 37 also shows how the basic circuit can be modified to power two light bulbs, each with its own independent control switch.

2.4.3 The electrical system

The aircraft's electrical system represented in Figure 38 is merely a further modification of the set-up described above.



If the left-hand side of Figure 38 is covered up, it is easy to see the similarity with Figure 37. The modifications are:

(a) the inclusion of protectors to prevent possible overheating of the services which might be caused by excess electric current;

(b) the master switch. This enables the entire system to be switched off, regardless of the positions of the individual service switches;

(c) the 'earth'. The metal airframe is said to act as the 'earth' when it takes on the function of the bottom wire in the simple circuit. This wire can then be eliminated, with a consequent saving in cost and weight.

If the battery were used to power the aircraft's services it would rapidly become run down or discharged. Hence the role of powering the electrical system is taken on by the alternator, which will also charge the battery.

In normal operation, therefore, with the alternator functioning and the master switch on (the master switch actually consists of two switches, as shown) the current flow is as indicated by the solid arrows.

If the alternator is not functioning for any reason, for example when the aircraft is on the ground with the engine stopped, it isolates itself from the remainder of the electrical system and the battery alone powers the system, with the current flow as shown by the hollow arrows. In this situation, which would also arise after failure of the alternator, care must be taken to keep the electrical load to a minimum (by switching off unnecessary services) to prevent the battery from discharging ('draining') too rapidly.

2.4.4 Fuses and circuit-breakers

Each protector referred to above is either a fuse or a circuit-breaker. It forms part of the circuit and its function is to prevent excess electric current flowing through its associated service by breaking the circuit. This effect is achieved by making use of the fact that whenever excess current flows through a metal wire the wire experiences an increase in temperature, as is explained below.

Excess current flow would be most likely to occur after malfunction of the service - the fuse or circuit-breaker therefore prevents further damage and, more importantly, the overheating (and attendant fire risk) that would otherwise be likely to occur.

2.4.4.1 The fuse

The fuse is a thin strand of special metal enclosed in a protective insulating case, usually of glass (Figure 39).



If the current strength exceeds the designed value, it heats the strand to the extent that the metal melts, thereby breaking the circuit. When this occurs the fuse is said to be 'blown', and the associated service cannot receive current until the fuse is replaced.

Fuses are graded according to the maximum strength of current that they will allow to flow through themselves without blowing. The unit measure of current strength (the actual amount of electricity flowing in the circuit) is the ampere, usually abbreviated to amp. Fuses are therefore graded in amps. As an example, the circuit for the aircraft's navigation lights might include a 5-amp fuse.

It is noteworthy that if a service is functioning normally its current requirement is considerably less than that which would blow the fuse.

Most aircraft carry a bank of spare fuses. If a fuse blows during flight it can be removed and replaced with a spare of the correct grading. If the new fuse blows, then it must be assumed that the associated service has malfunctioned - no further attempt to use it should be made until it has been inspected by qualified engineers and, if necessary, repaired.

2.4.4.2 The circuit-breaker

The circuit-breaker can be considered as a 'resettable fuse'. It consists of two strips of metal which are normally in contact with each other, allowing current to pass. If the current strength exceeds the designed value, one of the metal strips, which responds to the consequent increase in its temperature, moves apart from the other, thus breaking the circuit - the circuit-breaker is said to be 'tripped'.

The design of the circuit-breaker is such that contact is not automatically re-made when the heat-sensitive strip cools down again - this can only be achieved by physically resetting the circuit-breaker.

If a circuit-breaker continues to trip after being reset, it must be assumed that the associated service has malfunctioned, in which case the same considerations apply as in 2.4.4.1.

2.4.5 Electric motors

So far in this book, we have seen two examples of how electricity can be generated mechanically - the magneto and the alternator.

It is possible to reverse this operation - if electric current is supplied to a coil of wire mounted on a spindle and surrounded by a magnetic field, then the coil will rotate. The spindle can be made to drive other mechanisms; the assembly is now an electric motor.

In light aircraft some of the electrical services may incorporate electric motors. Typical examples are the flap operating mechanism and landing gear retraction mechanism (when these are electrically powered) and the engine starter.

2.4.6 The ammeter

The ammeter measures electric current strength and is therefore calibrated in amps. Its indications enable the pilot to verify that the electrical system is functioning correctly. A common arrangement is to include the ammeter between the alternator and the services, as shown in Figure 40.



In this layout the ammeter indicates the total current strength being generated by the alternator. This total is the sum of the individual requirements of whichever electrical services are switched on at any time. It includes the current needed to charge the battery.

To prevent possible overheating of the alternator it is usual for the manufacturer to specify a maximum permitted total current strength (for example 30 amps). Note that this limit is usually greater than would be needed if all the aircraft's electrical services were in simultaneous use. This is not always the case however. If the ammeter shows a reading greater than the maximum permitted, then the pilot should switch off services (obviously in increasing order of priority) until the indication is within the limit specified.

A presentation of the type of ammeter described above is shown in Figure 41.



An indication of zero on the ammeter shows that the alternator is not functioning, either because the engine is not operating or because the alternator itself has failed. In this situation the battery alone can supply the services, and the considerations mentioned in 2.4.3 apply.

An alternative location for the ammeter is between the battery and the services, as in Figure 42.



With this arrangement the ammeter shows the total current strength in the battery circuit. It is designed so that it can also indicate the direction of flow of the current. Reference to Figure 38 shows that the direction of flow depends upon whether the alternator or the battery is powering the electrical system. In the former case the battery is being charged and in the latter, discharged. Accordingly, the ammeter presentation is as shown in Figure 43.



An indication of 'discharge' shows that the alternator is not functioning.

The discussions above assume that the master switch is on. If it is off, the indication on the ammeter would be zero in both arrangements. Of course, with the master switch off the entire electrical system is dead and no services can be operated.

2.4.7 Electrical services
 
 The electrical services likely to be found in most light aircraft are:

(a) the avionic (aviation electronic) equipment, such as the radio;

(b) the lighting equipment;

(c) the oil temperature gauge and fuel quantity gauges;

(d) the gyroscopic turn-and-balance indicator;

(e) the pressure head heater;

(f) the electrically-driven fuel pump;

(g) the stall-warning device;

(h) the engine starter.

To operate any service the master switch and, if applicable, the appropriate service switch must both be put on. All these electrical switches are located in the cabin, usually in a group on the instrument panel. Note that some services (such as the various engine gauges and the stall warning-device) do not incorporate switches and are electrically powered as soon as the master switch is switched on.

The battery is secured in a special compartment in the airframe.

2.5 THE HYDRAULIC SYSTEM

In hydraulically-operated systems a fluid replaces the mechanical linkages which would otherwise be needed to activate the associated components. The fluid is contained within an arrangement of steel pipes and cylinders.

Most light aircraft employ a simple hydraulic system to operate the brakes on the landing gear mainwheels. Additionally, the retraction mechanism of aircraft with retractable landing gear may be powered hydraulically, by a more a complex, engine-driven system.

2.5.1 The brake system

The brake controls are integral with the rudder pedals (which will be described later) and are operated by the pilot's feet. Each brake has its own independent hydraulic sub-system. Figure 44 represents the sub-system for the brake on the left wheel - the brake is applied by pivoting the left rudder pedal forward.



The linkage converts the pedal movement into a push force on the piston in the master cylinder. Since the fluid is incompressible, this push force is transmitted to the piston in the brake cylinder which in turn forces the friction pad against the brake disc fixed to the wheel.

When the pedal is allowed to pivot back the push force on the master cylinder piston is removed and so the friction pad is no longer forcefully held against the disc - the brake is released. A spring device returns the pedal to its 'brake off' position when foot pressure is relaxed.

If the landing gear leg is of telescopic construction flexible piping allows the leg to move telescopically without affecting the operation of the brake.

The right brake sub-system operates in a similar manner. The brakes may be applied either together (to slow down or stop the aircraft) or differentially (to assist with steering). A separate control enables the system to hold the brakes 'on' for parking.

To compensate for small leaks of fluid past the various pistons a reservoir ensures that the system is replete with fluid. The non-return valves permit fluid to flow from the reservoir into the master cylinders, if required, when the brakes are released but prevent reverse flow when they are applied. It should be noted that if the system is in satisfactory working condition, fluid loss should be very little. Evidence of greater loss (for example if the reservoir needs frequent topping up) indicates that a more serious leak has occurred. Rectification will need to be carried out by a qualified engineer.

2.5.2 Hydraulic fluid

The fluids used in hydraulic systems are produced specifically for this purpose. They can either be made synthetically or refined from petroleum oil. The aircraft manufacturer will specify which particular fluid should be used in the machine's system and this information is often placarded adjacent to the reservoir, for ease of reference when topping up. A common location for the reservoir is in the engine compartment.

2.5.3 Hydraulically-operated landing gear retraction

Hydraulically-operated landing gear retraction mechanisms are usually powered by an engine-driven hydraulic pump. A schematic representation of such a system is given in Figure 45.



The landing gear leg is raised and lowered by a hydraulic jack, which consists of a cylinder and piston. Fluid can be supplied to either side of the piston, as shown.

This layout differs from the foot-operated brake system because it includes a return line for the fluid - it is a complete hydraulic circuit.

When the selector is moved to the 'down' position, fluid is pumped from the reservoir into the jack at the upper inlet, forcing the piston down the cylinder and lowering the leg. Below the piston, the displaced fluid flows along the return line to the reservoir. Note that the selector directs fluid to all three landing gear legs simultaneously, each leg having its own jack.

The reader is left to determine the result if the selector is moved to the 'up' position.

When the jacks have raised or lowered the landing gear, locking devices secure the legs in position until an opposite selection is made. In modern aircraft the selector position is controlled electrically, by a landing gear selector switch located in the cabin.

It is emphasised that the arrangement in Figure 45 is purely schematic - its purpose is to show the operating principle. In practice the engine-driven system includes many components and modifications whose description is beyond the scope of this book.

To guard against the possibility of the retraction mechanism becoming inoperative as a result of failure of the hydraulic system when the landing gear is retracted (for example if a leak allowed all the fluid to drain away), it is arranged that an alternative means of lowering the landing gear is available. In light aircraft the alternative means is often mechanical, being hand-operated by the pilot if needed.

2.6 THE FLIGHT INSTRUMENTS

The flight instruments are indicators mounted on the instrument panel in front of the pilot. Together with the compass, they show the behaviour of the aircraft in specific terms. The standard layout of instruments is as represented in Figure 46.



Notice that the instruments are located on the left-hand side of the panel. This is because the pilot occupies the front left-hand seat in the cabin.

2.6.1 Visual flight

The method of reference to the instruments depends on the mode of flight. In fine weather, for example, the pilot can see the earth's horizon outside the cabin and he uses the horizon as the reference for assessing the attitude of the aircraft. The horizon is the line in the distance along which the ground appears to meet the sky, and in good visibility it is visible all around the aircraft. The concept of attitude is concerned with the way in which the aircraft lies in relation to the horizon and is defined in terms of high or low and bank left or bank right.

A high attitude is one in which the nose of the machine is higher than the tail, and vice versa (Figure 47).


 
Of course, seated in the cabin, the pilot is unable to see this view of the aircraft. Instead, the manner in which the top of the engine cowling lies in relation to the horizon serves as a reference. The views from the pilot's seat corresponding to the situations in Figure 47 are as shown in Figure 48.



The aircraft is said to be banked to the left when its left wing is lower than its right wing, and vice versa (Figure 49).


 
 

Again, the pilot's assessment of bank is by the view from the cabin, as shown in Figure 50.

The pilot uses the controls as necessary to adjust the attitude of the aircraft and hence to achieve the desired flight path. There are, of course, an infinite number of variations of attitude. Figure 51 represents the aircraft in cruising flight, at a level attitude (not high or low) with the wings level (without bank left or right), and shows what might be a typical view from the cabin.



In fine weather, with a good view of the horizon, the pilot will need only to glance occasionally at the instruments to verify that the aircraft is behaving as required. Most of the time he will be looking outside the cabin, both to assess the machine's attitude in relation to the horizon, and also to watch for the presence of other traffic. This operational technique is called visual flight.

2.6.2 Instrument flight

If the horizon cannot be seen, for example when flying in cloud, only the flight instruments can furnish information about the aircraft's attitude. In this situation the pilot relinquishes the view through the windscreen and instead concentrates on the readings of the instruments, using the aircraft's controls as necessary to achieve the desired indications. This technique is called instrument flight and requires special pilot training to attain a safe standard of operation. A hollow 'T' is marked around the four instruments referred to most often during instrument flight to emphasise their importance to the pilot (Figure 46).

Nowadays, instrument flight is a normal operational technique. However, careful flight planning is needed to ensure that the aircraft maintains separation from other traffic and adequate clearance above the terrain over which it is flying when the view from the windscreen is obscured by cloud.

Later on in this book a brief dissertation on the fundamentals of instrument flight will be given, but it is stressed that, taking all considerations into account, this information deals with only a few facets of what is a somewhat involved subject.

2.6.3 The International Standard Atmosphere

The most important properties of the earth's atmosphere so far as aviation is concerned are:

(a) density - the actual number of molecules of air in a given volume;

(b) pressure;

(c) temperature.

In many respects these properties are inter-related. For example, increased density at a fixed temperature has the effect of increasing pressure, and vice versa. Similarly, increased temperature at a fixed density has the effect of increasing pressure; the opposite is also true.

In the atmosphere the three properties are not uniform - with increasing height, density and temperature become less. Hence, pressure also becomes less. The properties are also affected by variation in geographical location, weather patterns, diurnal cycle (day and night), seasonal change and many other factors.

The airspeed indicator (ASI), vertical speed indicator (VSI) and altimeter all make use of the atmosphere's properties, and so it is necessary to adopt a set of standards for the calibration of these instruments. This set of standards is the International Standard Atmosphere (ISA). The ISA represents an averaging out of the properties' variations actually occurring day-to-day and place-to-place. It assumes that at sea level, the air has a density corresponding to a weight of air of 1225  grams per cubic metre, and a temperature of 15C. Fixing these values has the effect of fixing sea level air pressure at 1013 hectopascals (hPa). With increasing height, temperature is assumed to decrease at the rate of 2C every 1000 feet, and pressure at the rate of 1 hPa every 28 feet. In practice, in the lower levels of the atmosphere it is found that this latter relationship holds reasonably true even in conditions which otherwise differ from those assumed in the ISA.

2.6.4 The airspeed indicator

As its name suggests, the ASI registers the speed of the aircraft through the air.

To appreciate its working principle, imagine a cylinder, closed at one end by a flexible diaphragm, facing into an airflow (Figure 52).



Because the cylinder is closed, the moving air builds up a pressure inside it. This dynamic pressure, as it is termed, pushes against the diaphragm, moving it as shown. When there is no airflow the dynamic pressure is zero, and the diaphragm reverts to its rest position. Thus the amount of diaphragm movement is an indication of the speed of the airflow - faster airflow causes greater movement. All that is needed now is a way of registering the diaphragm movement on a calibrated scale, and we shall have an ASI. Figure 53 shows the instrument schematically.



The diaphragm has been modified into a capsule made of thin metal and enclosed in a case, part of which is of glass, so that the calibrated scale can be seen. Th pitot tube faces into the airflow and so dynamic pressure builds up inside it. The inside of the capsule therefore experiences a total pressure made up of the dynamic pressure together with the atmosphere's natural ambient pressure, or static pressure as it is called. The static tube is closed but has vents in it, allowing the inside of the instrument case, and hence the outside of the capsule, to experience purely static pressure. Thus the movement of the capsule is dependent only upon the magnitude of the dynamic pressure. The linkage converts capsule movement Into movement of the pointer against the calibrated scale.

The ends of the pitot and static tubes are positioned at some point in the airflow where it is, as much as is possible, undisturbed by the presence of the moving aircraft. Frequently, the two ends are combined into one assembly, called the pressure head, located under the lower surface of one of the wings, as shown in Figure 54.



In the type of pressure head shown in the diagram, the pitot tube is surrounded by the static tube.

During flight in humid weather conditions in sub-zero temperatures, it is possible for ice to form on the pressure head. If this occurs, the tubes may become obstructed, in which case the transmission of dynamic and static pressure to the ASI will be interfered with, causing the instrument to give erroneous indications. The danger is averted by the inclusion of a heater element in the pressure head. When the heater is switched on by the pilot, electric current is supplied to the element and the pressure head is warmed.

From the pressure head the pitot and static tubes run through the airframe structure and into the ASI, located on the instrument panel.

The ASI scale is usually calibrated in miles per hour (mph) or knots. (One knot is a speed of one nautical mile (6076 feet) per hour. The nautical mile is the standard unit of distance in air navigation.)
 
The reading on the ASI is called the indicated airspeed (IAS). Figure 55 shows the ASI presentation on the instrument panel.



In this example, the calibrated scale has markings on it showing the various limiting speeds that apply to the aircraft. The airframe structure is designed to bear the aerodynamic forces acting on it during flight, with an adequate margin of strength, so long as these limits are observed. If they are disregarded, the increased aerodynamic forces may overstress the airframe; severe overstressing may cause distortion or even total failure of the airframe structure.

At the higher end of the white segment on the scale is the flap limiting speed - the IAS above which the flaps must be up.

At the higher end of the green segment is the maximum normal operating speed ('Vno' in aerodynamic terminology). The aircraft should not be flown faster than this IAS during normal operation. If, as the result of an upset, Vno is exceeded, then the pilot should use the controls to reduce IAS to Vno, or less, as quickly as possible.

The speed which must never be exceeded under any circumstances (Vne) is marked at the end of the yellow segment. The Vne marking is often referred to as the red-line.

Of course, every aircraft design has its own particular limiting speeds. In the example shown in Figure 55, the flap limiting speed is 95 knots IAS, Vno is 135 knots IAS and Vne is 160 knots IAS.

The significance of the differing lower limits of the green and white segments will be explained later.

Because of various errors, which are described below, the IAS may not be a true indication of the speed of motion of the aircraft through the air.

2.6.4.1 Pressure error and instrument error

It was mentioned that, ideally, the pressure head should be positioned in undisturbed airflow. In practice this cannot be achieved, and the air in the vicinity of the pressure head is upset to a certain extent by the passage of the aircraft. Additionally, the vents in the static tube imperfectly relay static pressure to the ASI. Both of these effects may cause the instrument to indicate erroneously. Whether an under-reading or over-reading is the result depends upon the design of the pressure head and the exact whereabouts of its location. The discrepancy is called pressure error, and it may vary according to the aircraft's speed.

Minor imperfections in the construction of the instrument may reduce the accuracy of its indications. Today, advanced manufacturing techniques ensure that this instrument error, as it is called, is small. The aircraft's Flight Manual will state the values of the correction that should be applied to the IAS to obtain the rectified airspeed (RAS), that is, IAS corrected for pressure error and instrument error. In nearly all designs of light aircraft these errors are insignificant, the combination of the two usually being less than 5 knots, and so for all practical purposes RAS can be assumed to be the same as IAS.

2.6.4.2 Density error

The ASI is calibrated in accordance with the ISA sea level air density conditions. With increasing height, air density becomes less, as already stated - a given volume contains fewer molecules of air. Because of this the static pressure is less than at lower height. So too is the dynamic pressure in the pitot tube of the ASI at any particular airspeed - the instrument will under-read.

The overall result, therefore, is that with increasing height above sea level the ASI will progressively under-read, the discrepancy being called density error. To further complicate matters, the air density at any particular height depends upon ambient air temperature, warmer temperatures than ISA conditions reducing density further, and cooler temperatures lessening the reduction. In other words, the magnitude of the density error depends upon air temperature as well as height.

It is possible to calculate the value of the density error for any particular height and air temperature. If this error value is applied to the RAS, or to the IAS if pressure error and instrument error are disregarded, the result is then the true airspeed (TAS) - the actual speed at which the aircraft is moving through the air.

The density error is zero only in air density conditions equivalent to sea level in the ISA - in this case IAS and TAS are identical. Figure 56 is a table showing TAS corresponding to an IAS of 100 knots at selected heights. Also shown is the effect caused by temperature variations.



2.6.4.3 Groundspeed

As has just been explained, the actual speed at which the aircraft moves through the air is its TAS, which can be calculated by applying density error correction to the IAS. If the air is stationary with respect to the ground, then the aircraft's speed over the ground - its groundspeed (GS) - will equal its TAS. If, however, the air is moving over the ground - in other words, if there is a wind blowing - then the GS may differ from the TAS.

As an example, suppose that a 20-knot wind is blowing. If an aircraft whose TAS is 100 knots is flying into the wind then its GS will be 80 knots. Conversely, if it turns and flies in the same direction as the wind then its GS will be 120 knots. Figure 57 represents these two situations.

Aerodynamically, GS has no significance; its importance lies primarily in navigational planning.



2.6.5 The altimeter

The altimeter registers height. It makes use of the phenomenon already mentioned that as height increases, the atmosphere's static pressure becomes less.

Figure 58 is a schematic representation of an altimeter.



The capsule is made from thin metal and is evacuated of air. It is fixed to the case of the instrument. A tapping from the static tube (from the pressure head) leads into the case. Thus the outside of the capsule experiences local static pressure.

The static pressure acting on the evacuated capsule would cause it to collapse, were it not for the metal spring which exerts a pull force on the capsule, maintaining its shape - the effect of the spring balances the effect of the air pressure.

With increasing height, the static pressure in the case becomes less and so the spring pulls the capsule into an expanded position, as shown, greater height causing greater capsule movement. The movement is registered on a scale calibrated in feet, the foot being the standard unit of height.

In modern altimeters the capsule movement is geared to drive three pointers, one indicating hundreds of feet, another thousands of feet and the third tens of thousands of feet.

Figure 59 shows the presentation on the instrument panel.



The tens of thousands of feet pointer is often modified into the form of indicator shown. At zero height indicated, the area revealed by the window is blank. As height increases, a white-painted area comes more into view in the window, the edge of the white area indicating the tens of thousands of feet.

In Figure 59 the altimeter shows a height of 4650 feet. Figure 60 shows an indication of 8800 and Fig 61 shows 12000 feet.

The altimeter indications are subject to various errors, which will be described later.





2.6.5.1 Pressure datum setting

When measuring height, a datum is necessary for reference. In cruising flight, the altimeter is usually made to indicate height above sea level. But the pilot of an aircraft flying in the vicinity of an airfield may require the altimeter to register height above the level of the airfield.
 
As an example, consider an aircraft flying over an airfield, as represented in Figure 62. The airfield is 500 feet above sea level.

 

Depending upon his preference the pilot will want the altimeter to read either 1000 feet (above airfield level) or 1500 feet (above sea level). The altimeter must therefore have the capability of being adjustable to be able to achieve both requirements.

The adjustment mechanism is controlled by the pressure datum knob (Figure 59) and makes use of the decrease in pressure of the atmosphere as height increases. As has already been stated, the pressure decreases at the rate of 1 hectopascal (hPa) every 28 feet of height.

Suppose that, in the example above, the sea level air pressure is 1020 hPa. Using the relationship just mentioned, we can calculate that the air pressure at the airfield will be 1020 - (500/28) or 1002 hPa. These pressure values of 1020 and 1002 hPa can be used as reference datums. When the pressure datum in the altimeter is set to 1020 hPa with the knob, the altimeter will register height above this datum. If the setting is changed to 1002 hPa the altimeter will indicate height above this new datum.

In the example in Figure 62, therefore, the aircraft's altimeter will register 1500 feet when its pressure datum is set to 1020 and 1000 feet when set to 1002.

The altimeter in Figure 59 is shown set to 1007, and those in Figures 60 and 61 to 983 and 1034 respectively.

2.6.5.2 Altimeter terminology

Thus far in this book the word 'height' has been used exclusively to refer to vertical displacement from the earth. For air navigation purposes, however, more specific terminology is required. In this context:

(a) the word height is used to describe vertical displacement from a ground-based reference, usually an airfield;

(b) the word altitude is used to describe vertical displacement above sea level.

The phrases 'air pressure at airfield level' and 'air pressure at sea level' are respectively abbreviated into the international code forms QFE and QNH.

In Figure 62 the aircraft is said to have a height of 1000 feet on QFE 1002 hPa and an altitude of 1500 feet on QNH 1020 hPa.

It is apparent that care must be exercised by the pilot when adjusting his altimeter datum to ensure that the correct setting is used as appropriate.

In certain navigational circumstances the altimeter datum is set to what is called the standard setting, which is 1013 hPa. When this setting is in use, the vertical displacement of the aircraft is expressed in terms of flight level (FL), which is the altimeter indication divided by a factor of 100. For example, an indication of 4500 feet is referred to as FL 45.

2.6.5.3 Barometric error

At any particular location the atmospheric pressure is constantly changing, although usually not at a rapid rate. It also varies from one location to another. Suppose that the aircraft in Figure 62 has returned to the location shown after a period of several days, at an indicated altitude of 1500 feet, and that the QNH at this location had decreased from 1020 to 995 hPa.

If the pilot had neglected to reset his altimeter to this new QNH, then his indicated altitude of 1500 feet would be in error - it would be an over-reading of (1020-995) x 28 or 700 feet. The aircraft's true altitude would therefore be 800 feet.

Of course, the QFE at the airfield would now be 995 - (500/28) or 977 hPa. Use of the outdated value of 1002 hPa would again give a 700 feet over-read of height - an indicated height of 1000 feet would in fact be a true height of 300 feet. Figure 63 clarifies the situation and shows both the old and the new pressure datums.



Altimeter errors caused by incorrect setting of the pressure datum are called barometric errors, and they can be eliminated by appropriate use of the setting knob. Thus the pilot returning to the location in our example should set his altimeter either to 995 hPa to obtain correct indications of altitude or to 977 hPa for correct indications of height above airfield level.

From the foregoing remarks, it is clear that here is another reason, besides that mentioned in 2.6.5.2, why the pilot must take every care to ensure that his altimeter is correctly set. This is especially important when flying from a region of higher QNH to one where the QNH is lower, when failure to reset the altimeter will cause it to over-read - to show that the aircraft is higher than it really is.

2.6.5.4 Terrain clearance

Note that use of QNH does not enable the altimeter to register vertical distance above the terrain. Even when set to QFE, the altimeter will show vertical distance above only one selected area of terrain - the airfield. For cruising flight the pilot sets the QNH. If the flight is a long one, it is quite likely that this setting will need updating as the aircraft travels from one location to the next, where the QNH may be different.

It is clear, therefore, that the pilot should be aware of the vertical distance above sea level, or elevation, of the terrain over which the aircraft flies. With this knowledge, it is easy to assess the vertical distance of the aircraft above the terrain - in other words, its terrain clearance - if its altimeter is set to QNH. Information about terrain elevation is to be found on the maps, or charts, used for air navigation.



As an example, suppose that the aircraft in Figure 64 is in cruising flight. Its altimeter, set to QNH, indicates an altitude of, say, 4000 feet. The pilot will have determined, from consultation of the navigation chart, that the elevation of the terrain at this location is, say, 1600 feet. He will therefore know that the aircraft's terrain clearance is 2400 feet.

Of course, using the technique of visual flight, the pilot can see the terrain below the aircraft and can verify that the clearance is sufficient without having to consult his altimeter.

However, if the aircraft were flying in cloud, with the pilot using instrument flight technique, then he would be denied the sight of the earth's terrain. In this case it would be necessary to choose a cruising altitude safely above the terrain that would be encountered en route and to ensure that the altimeter indication did not fall lower than this altitude during flight.

2.6.5.5 Temperature error

Deviations of air temperature from ISA conditions cause corresponding deviations of pressure. In colder conditions, the air pressure at any particular height is less than at the same height in the ISA. The altimeter interprets the lowered pressure as an increased height - in other words, it over-reads, showing the aircraft to be higher than it really is. The converse is true in warmer conditions.

These discrepancies are called temperature errors and they are of no great significance except during instrument flight in very cold weather (terrain clearance being the relevant consideration in this case).

2.6.5.6 Pressure error and instrument error

The same considerations apply to the altimeter as to the ASI (refer back to 2.6.4.1). However, the errors are insignificant and for all practical purposes can be ignored.

2.6.5.7 GPS derived altimetry

Many modern aircraft incorporate Global Positioning System (GPS) receivers in their navigation equipment. Some of these GPS receivers can indicate altitude but it should be borne in mind that, although their indications may be more accurate in absolute terms than those of pressure altimeters, they are not used as primary reference since at the present time vertical traffic separation procedures are based on the indications of pressure altimeters.

2.6.6 The vertical speed indicator

The VSI shows the rate at which the aircraft is climbing or descending. Like the altimeter, it makes use of the decrease of atmospheric pressure as height increases. The instrument is represented schematically in Figure 65.



A tapping from the static tube leads into the VSI case and then branches into two tubes, one of which runs directly into the capsule. The other tube leads into the body of the instrument via the metering device, as shown. The metering device is merely a restrictive orifice which retards any movement of air past itself.

When the aircraft is in flight at constant height it is clear that the inside and outside of the capsule experience the same pressure. In this situation the capsule will be in its rest position and the pointer will indicate zero, showing no climb or descent.

If the aircraft is now made to descend, it will be flying into a region of denser air, where the static pressure is obviously greater. Whilst the inside of the capsule will experience this greater pressure directly, the metering device will retard the inflow of the denser air into the body and so the pressure experienced by the outside of the capsule will be less than that inside. Accordingly, the capsule will expand, its movement being converted by the linkage to a pointer indication of descent.

An opposite sequence of events occurs if the aircraft is made to climb. Note that the more rapid the climb or descent, the greater is the delaying effect of the metering device and so the difference in pressure inside and outside the capsule is more marked. This in turn causes greater capsule movement, which makes the pointer move further away from zero. In other words, the instrument indicates the rate of climb or descent. The scale is calibrated in feet per minute. Figure 66 shows the presentation on the instrument panel. A climb of 500 feet per minute is indicated.

The metering device is so constructed that the VSI indications are accurate regardless of height or air temperature. However, the indications are subject to the errors described below.



2.6.6.1 Lag

A period of a few seconds is required for the metering device to convert any change in static pressure into a pressure difference inside and outside the capsule. Because of this, there is a corresponding delay in the VSI indications - it takes a few seconds for the pointer to settle on a steady reading. The delay is termed lag and is most noticeable when abrupt changes occur in the rate of climb or descent.

Abrupt changes from climb to descent and vice versa may even cause the instrument to register in the reverse sense, albeit temporarily. Consider the aircraft in Figure 67 which has just changed its flight path from a rapid climb to a rapid descent.

Note that the VSI indication at the start of the descent incorrectly shows that the aircraft is still climbing.

It can be appreciated that the VSI gives reliable indications only when the indications are steady.



2.6.6.2 Pressure error

Any change in static pressure is interpreted by the VSI as a climb or descent. If the change is due to inconsistent sensing of static pressure at the pressure head (which could be caused by a change in the aircraft's speed, for example) then obviously the VSI indications may be false. This pressure error is negligible in most light aircraft because of the relatively gentle accelerations and decelerations that are experienced.

2.6.7 Direction

The magnetic compass shows the direction in which the nose of the aircraft is pointing, the reference for direction being magnetic north. At any point on the surface of the earth, magnetic north is the direction in which the north-seeking end of a freely-suspended magnetised rod aligns itself, acting under the influence of the earth's natural magnetic field. At most locations this direction differs from true north, which is the direction in which the earth's geographic north pole lies. The difference between the two directions is termed variation, and in European areas is not great (Figure 68).



Magnetic direction is measured as angular displacement from magnetic north. In air navigation, direction is expressed in terms of degrees ().

The three directions shown in Figure 69 are 008M, 063M and 251M. 'M' denotes magnetic direction.



In this nomenclature, the magnetic cardinal directions north, east, south and west are respectively expressed as 360M (rather than 000M), 090M, 180M and 270M. The magnetic inter-cardinal directions north-east, south-east, south-west and north-west are respectively expressed as 045M, 135M, 225M and 315M.

The direction in which the nose of the aircraft points is called its heading. The aircraft in Figure 70 have headings of 080M and 305M.



2.6.8 The magnetic compass

The aircraft's heading is indicated to the pilot by the magnetic compass, which is shown schematically in Figure 71.

The magnets are attached to the circular card and the whole assembly is suspended below a pivot which is fixed to the case. The assembly can thereby rotate independently of the case. On the card are marked the heading indications.



The case features a glass window so that the card can be seen from outside. A vertical lubber line in the middle of the window is the reference against which the heading indication is read. (For clarity of marking on the card, the final zero is omitted from the indications, as is the first zero on the indications of 030 and 060. The cardinal directions are marked as N, E, S and W.)
 
The magnets align themselves with their north-seeking ends pointing in the direction of magnetic north. Thus the card is held in one fixed orientation, regardless of the orientation of the case (which is secured to the airframe structure).

The case is filled with alcohol, the liquid serving two purposes:

(a) by helping to support the magnet assembly (in the same way that water supports ships), it minimises the friction at the pivot, making the compass more sensitive;

(b) it damps down oscillations of the magnet assembly, conferring steadiness to the compass indications.

In Figure 71 the aircraft has a heading of 360M, and in Figure 72, 270M. As an aid to understanding, the symbol shown on top of the case represents the aircraft within which the compass is mounted.

The compass is located in a suitable position in the cabin, where it is easily visible. Figure 73 shows various heading indications as the pilot sees them.





2.6.8.1 Compass errors

Objects made of ferrous metals (iron and steel) distort the earth's natural magnetic field in their vicinity. Aircraft components made of these metals, such as the engine, may therefore cause the compass to give erroneous indications (since the magnets will align themselves with the distorted field). A similar effect is caused by electrical equipment, which when switched on may set up weak magnetic fields, these fields locally distorting the earth's field. Discrepancy of heading indication resulting from such disturbances is called deviation, and can be minimised (but not completely eliminated) by:

(a) locating the compass as far as possible from electrical equipment and components made of ferrous metals;

(b) suitable adjustment of the correction mechanism built into the compass. (The correction mechanism contains tiny magnets which set up fields helping to cancel out those tending to disturb the compass.)

The residual deviation is usually less than 2 - it varies according to the aircraft's heading - and can be ignored for all practical purposes. Periodically the compass is checked, or swung, by qualified technicians, to verify that its deviation is not excessive.

It goes without saying that, during flight, objects made of ferrous metals (such as tools) must never be stowed near to the compass installation, since these might cause excessive deviation.

The various forces which act on the aircraft when it is flying in a banked attitude, and during acceleration and deceleration, upset the magnet assembly, with the result that the compass gives incorrect indications. Even gentle flight manoeuvres may cause considerable errors.

For this reason the compass can be relied upon to give correct heading information only when the aircraft is flying at a constant speed with its wings level.

Another problem arises if the aircraft encounters turbulent air, when the compass card will oscillate irregularly (despite the damping effect of the alcohol), making it difficult for the pilot to assess the aircraft's heading.

2.6.9 The gyroscope

A gyroscope (usually abbreviated to 'gyro') is an object designed to make use of the physical phenomena associated with rotational motion. Typically, the gyro comprises a metal wheel mounted in a suitable housing. The wheel is made to spin at high RPM (Figure 74).



Three of the flight instruments - the attitude indicator (AI), the direction indicator (DI) and the turn-and-balance indicator (TBI) - incorporate gyros, each instrument utilising one of the two most important gyroscopic properties, which are:

(a) rigidity - the natural tendency of the gyro to resist changes in its orientation;

(b) precession - the resulting motion of the gyro when a disturbance is forced upon it.

In the flight instruments mentioned above, the gyros are made to rotate by one of two means, as described below.

2.6.9.1 The suction-driven gyro

The gyro is mounted in a metal case, from which the air is sucked by a suction pump driven by the aircraft's engine. Replacement air is allowed to enter the case through a filter, whose purpose is to remove foreign matter. The air is then ducted to the gyro where it impinges upon special indentations on the periphery, making the gyro rotate (Figure 75).



The gyro will rotate steadily so long as suction is maintained. Failure of the engine or of the suction pump, or serious leaks in the suction tubes, will result in the gyro slowing down and stopping - the associated flight instrument will then become inoperative.

A gauge mounted on the instrument panel measures the strength of the suction. The aircraft manufacturer will specify the range within which the reading on the gauge should lie when the engine is running normally. A reading outside this range implies a fault in the suction system and is a warning to the pilot that the suction-powered flight instruments can no longer be depended upon to give reliable information.

2.6.9.2 The electrically-driven gyro

The electrically-driven gyro is merely a modified electric motor, in which the metal wheel incorporates the motor's moving coil. The assembly will rotate whenever electric current is supplied to the coil.

Failure of the current supply will cause the gyro to slow down and stop, rendering the associated flight instrument inoperative.

2.6.9.3 Flight instrument power supplies

In light aircraft the usual arrangement is for the AI and the DI to be suction-powered and the TBI to be electrically-powered, so that at least one of the three instruments remains operative after failure of one of the power supplies.

2.6.10 The attitude indicator

The AI indicates the attitude of the aircraft and is referred to by the pilot when the earth's horizon is obscured. Its gyro is mounted in housings, or gimbals, in such a way that its orientation is totally independent of the aircraft's attitude.

The gyro is so designed that, whenever it is spinning, it orientates itself with its axis vertical with respect to the earth's horizon. In this state the gyro is said to be erect. (Erection usually takes several minutes from the moment the gyro starts to rotate.) Figure 76 illustrates.



The gyro's rigidity maintains it in this orientation regardless of the attitude of the aircraft. Attached to the gimbal arrangement by pivots, the horizon display (see Figure 77) is held level with respect to the earth's horizon by the gyro - the horizon display is said to be gyro-stabilised. The diagram shows the presentation of the AI on the instrument panel.



The attitude datum (which symbolically represents the aircraft viewed from behind) is fixed inside the glass face of the instrument, and so change of aircraft attitude is indicated by change of position of the horizon display relative to the attitude datum. The AI indication in Figure 77 corresponds to the cruising flight attitude shown in Figure 51 . Indications corresponding to other attitudes are shown in Figure 78.



2.6.11 The direction indicator

It has been previously mentioned that the magnetic compass heading indications are unreliable during flight manoeuvres and unsteady in turbulence. These drawbacks are overcome in the gyro-stabilised DI.

When it is spinning, the gyro in the DI orientates itself with its axis pointing in a fixed direction. In this state it is considered erect (Figure 79).



The gyro is mounted in gimbals which allow its rigidity to maintain it in the fixed directional orientation regardless of the aircraft's heading (in the same manner that the compass magnets hold one fixed orientation). The gimbal arrangement is geared to drive the indicator card (Figure 80), so that change of heading causes the card to rotate behind the heading datum. The diagram shows the presentation on the instrument panel, with the DI indicating 360M. A heading of 255M is indicated in Figure 81.





Of course, although the DI gyro is held in a fixed directional orientation by its rigidity, this orientation is random - when erect, the gyro's axis might point in any direction. A means is therefore provided for orientating the axis correctly with respect to our chosen directional datum - magnetic north. When the synchronising knob is pushed in and turned, the gimbal assembly, and therefore the card, can be rotated so that the heading datum shows the same indication as the magnetic compass. Releasing the knob frees the gimbal assembly so that the card is again acting in its gyro-stabilised mode. The DI is now said to be synchronised with the compass and will indicate heading in terms of magnetic direction. Furthermore, the indications will be correct during manoeuvring flight and steady in turbulence.

2.6.11.1 DI errors

Ideally, once the DI has been synchronised with the compass and the knob released, it should remain synchronised indefinitely.

However, a complication arises from the fact that the gyro's fixed directional orientation makes no allowance for the effect of the rotation of the earth, which causes the DI indication to become increasingly incorrect with the passage of time. Fortunately this effect can be offset by an internal adjustment device within the DI. The adjustment device rotates the gyro axis at such a rate that the discrepancy is eliminated. It is set according to the geographical area within which the aircraft is expected to fly. If the aircraft is moved to a different area further north or south, the adjustment device will need to be reset to allow for the difference in the effect of the earth's rotation. Resetting is only required when the change of location is considerable. For example, it would not be worthwhile for an aircraft based in the UK to have its DI readjusted for flights in continental Europe.

A further problem is that the frictional forces in the gimbal assembly slowly change the directional orientation of the gyro axis. In other words, they make the gyro precess.

If the adjustment device described above has been correctly set and if the internal friction forces are not great, then the discrepancy in heading between the DI and the magnetic compass after synchronisation should not increase rapidly. (A rate of change of more than 20 per hour would be considered excessive.)
 
In practice, therefore, normal procedure is to synchronise the DI before take-off, and then to re-synchronise during flight at intervals of 15 minutes or so.

2.6.11.2 The self-synchronising DI

In aircraft fitted with more sophisticated equipment the conventional DI is sometimes replaced by a self-synchronising model. A description of the instrument is beyond the scope of this book. The operating principle is that an electrically-powered compass system keeps the DI synchronised with respect to magnetic north, relieving the pilot of the requirement to synchronise manually. The magnetic compass is usually retained as a back-up in case of failure of the electrically-powered system.

2.6.12 Toppling

With most designs of AI and DI the gimbals give their gyros independence of orientation only within certain physical limits of gimbal movement. These limits will not be reached during normal flight manoeuvres. However, extremes of attitude (such as may occur during aerobatics) may well bring the gimbals to their limiting positions. When this happens the gyros are made to precess by the disturbing forces, usually undergoing rapid oscillatory change of orientation. A gyro behaving in this manner is said to be toppled and its associated flight instrument will, of course, be inoperative. Random movement of the horizon display (in the AI) and rapid rotation of the indicator card (in the DI) are the usual symptoms of toppled gyros.

Once the aircraft has returned to a level attitude, a period of several minutes may be required for the AI to re-erect itself, during which time its indications should be disregarded. The DI can be re-erected immediately by pushing in and releasing the synchronising knob. (In practice the knob is pushed in and rotated to resynchronise the DI with the compass, and then released again.)

2.6.13 The turn-and-balance indicator

It is universal practice for the turn indicator and the balance indicator to be combined so far as instrument panel presentation is concerned - the TBI is therefore two separate instruments in one presentation.

2.6.13.1 The turn indicator

The turn indicator incorporates a gyro which utilises the phenomenon of precession rather than rigidity. As its name suggests, the instrument shows when the aircraft is turning - in other words, when the heading is changing.

The gyro is mounted with its axis running from left to right relative to the aircraft. In straight flight it is held in this position by springs attached to its gimbal. If the aircraft turns to the left, the ensuing forces acting on the gyro make it precess - it tilts against the restoring effect of the springs. A turn to the right makes the gyro tilt the opposite way (Figure 82).



As soon as the aircraft stops turning the springs restore the gyro to its central position. The movements of the gyro are geared to the turn indicator. Note that the more rapid the turn, the greater the precession and the more the gyro tilts against the effect of the springs. Thus the instrument indicates the rate of turning. Figure 83 shows the instrument panel presentation. The indicator is central, showing no turn.



In a rate 1 turn, the aircraft heading changes at the rate of 3 per second. To turn from 090 to 270 would take 60 seconds. In a rate 2 turn, the heading changes at the rate of 6 per second. To turn from 090 to 270 would take 30 seconds.



Figure 84 shows indications of a rate 1 turn left and a rate 2 turn right. Failure of the power supply to the instrument will result in the gyro stopping, in which case the indicator will remain central even when the aircraft is turning.

2.6.13.2 The balance indicator

The balance indicator is the simplest of the flight instruments. It consists merely of a sealed curved glass tube filled with liquid in which a solid ball is free to move (Figure 85).


 
The disposition of the various forces (including gravity) acting on the aircraft during flight determines the position of the ball in the tube.

When the disposition of these forces is such that the balance indicator ball lies within the two central markings, the aircraft is said to be 'in balance'. Unbalanced flight is indicated by displacement of the ball from the central position - the greater the unbalance, the more the ball is displaced. Figure 86 illustrates.



The pilot uses his controls to keep the ball centred since unbalanced flight is inefficient - it results in the aircraft generating more drag than is necessary. (Occasionally, when extra drag is desired for modification of the aircraft's flight path, the pilot may deliberately fly the machine out of balance. In this situation, of course, the ball in the balance indicator will be displaced from the central position.)

The purpose of the liquid in the tube is to help to damp out oscillations of the ball during manoeuvring flight. Figure 87 shows the instrument panel presentation of the complete TBI.



2.6.14 Electronic flight instrument systems

Many newer light aircraft are fitted with electronic flight instrument systems (EFIS), in which the mechanical sensors and indicators described above are replaced with electronic components.

The operating principle is that the inputs from the pitot pressure and static pressure lines are directed to electronic detectors. Aircraft attitude, magnetic heading and rate of turn are sensed by electronic components. After processing, the data generates a display on a liquid crystal display (LCD) screen on the instrument panel. Figure 88 shows a representation of an EFIS display.



In Figure 88 the instantaneous readouts are shown against scrolling background scales. In addition to a numerical readout, vertical speed is also indicated by a variable length arrow. The reversed colouring on the numerical readout is to prevent any confusion with altitude indication. Airspeed trend is also shown pictorially (red 'down' arrow for decreasing IAS and green 'up' arrow for increasing).

Compared to mechanical instruments EFIS installations have several advantages:

(a) greater reliability and reduced maintenance costs;

(b) lighter and more compact construction;

(c) automatic synchronisation of the direction indicator;

(d) elimination of requirement for engine-driven suction pump.

2.7 THE PILOT'S CONTROLS

Figure 89 shows what might be a typical layout for the forward end of the cabin of a modern single-engined light aircraft equipped with a fixed-pitch propeller and fixed (non-retractable) landing gear.



It is usual for interconnected dual controls to be fitted so that the machine can be controlled from either of the front seats, although the flight instruments (Figure 46) are positioned on the left of the instrument panel since the left seat is the position from which the aircraft is flown when only one pilot is on board.

2.7.1 The flight controls

The flight controls - the ailerons, elevators and rudder - are activated by the control wheel and the rudder pedals, as described below.

2.7.1.1 The ailerons

The control wheel is connected to the ailerons by cables or rods. Rotating the control wheel to the left raises the left aileron and lowers the right aileron, and vice versa. The control wheel is progressive in its operation - greater control wheel movement results in greater aileron displacements. When the control wheel is in its central position the ailerons are undisplaced (Figure 90 and Figure 6 ).



2.7.1.2 The elevators

The control wheel is also connected to the elevators by cables or rods. Pulling the control wheel backwards (towards the pilot) raises the elevators, and vice versa. The control wheel is progressive in its operation. When it is centralised the elevators return to their undisplaced position (Figure 91).



2.7.1.3 The rudder

The rudder pedals are connected to the rudder by cables or rods. Pushing the left pedal forwards deflects the rudder to the left; pushing the right pedal forwards deflects the rudder to the right. The pedals are progressive in their operation. When they are centralised the rudder returns to its undisplaced position.

In most tricycle landing gear aircraft, the rudder pedals are also linked either rigidly or via springs to the nosewheel steering mechanism, for steering the machine during ground manoeuvring. Pushing forwards the left pedal turns the nosewheel to the left, and vice versa. Figure 92 illustrates these two functions of the rudder pedals.


 
A third function of the pedals is to apply the mainwheel brakes. The pilot pushes the upper portion of the appropriate pedal with his toes to apply the associated brake. When these 'toe brakes' are not in use, the pilot drops his feet to the lower parts of the pedals.

2.7.1.4 Control locks

Whenever the aircraft is parked in strong winds, an internal locking device is used to hold the control wheel firmly in its central position. Any tendency for the wind to move the ailerons and elevators, which might lead to them being damaged if violent movement occurred, is thus prevented. A separate lock secures the rudder pedals, unless the pedals are linked to the nosewheel, in which case movement of the rudder is prevented by the fact that the nosewheel is in contact with the ground and will resist turning to the left or right.

Common sense dictates that, whenever the aircraft is to be left parked unattended for any length of time, the control lock(s) should be employed.

2.7.2 The flaps
 
In the layout shown in Figure 89, the flaps are lowered and raised mechanically by pulling up or lowering the flap lever, the linkage incorporating a ratchet so that intermediate selection(s) of flap position between fully up and fully down are possible (Figure 93 and Figure 5).



As an alternative, many aircraft feature electrically-operated flaps. When the control switch (usually located on the instrument panel) is pushed down, the electric flap motor lowers the flaps; when the switch is pushed up the motor raises the flaps. With the switch in its central position no current is supplied to the motor and the flaps are retained at whatever position they happen to be. Thus intermediate selections are possible.

2.7.3 The elevator trim tab

The control for the elevator trim tab is usually in the form of a wheel, to which the tab is connected by cables. When this 'trim wheel' is rotated upwards and forwards the tab is displaced upwards; rotation of the wheel backwards and downwards displaces the tab downwards. The control is progressive in its operation. When the wheel is released the tab is retained in whatever position it happens to be. Centralising the wheel brings the tab back to its undisplaced position (Figure 94).



If the aircraft features aileron and rudder tabs adjustable in flight, they are controlled by similar trim wheels.

2.7.4 The engine controls

The engine controls are usually grouped together in a low central position on the instrument panel, as in the example in Figure 89.

The throttle is opened by pushing its lever forwards, and closed by pulling the lever backwards.

If it is required that the throttle lever retains any set position, the friction control is tightened. Otherwise, it is loosened to permit easy movement of the lever.

The mixture control is set to 'rich' by pushing its lever forwards, and to 'weak' by pulling it backwards. On many designs the mixture control lever is used to stop the engine after flight by moving it to its rearmost position. The lever is then in the 'idle cut-off' position, and the fuel supply to the nozzle in the carburettor is completely cut off.

The carburettor heat control is set to 'off' by pushing its lever forwards and to 'on' by pulling it backwards. Note that the throttle and mixture control levers are progressive in their operation, whereas the carburettor heat control should be selected either 'on' or 'off'.

The magneto switches are usually amalgamated into one key-activated control which is also used to energise the electric starter motor for engine starting. Figure 95 shows a typical arrangement.



When the key is at the OFF position both magnetos are switched off. At the R position the right magneto is on, and the left off, and vice versa at the L position. At the BOTH position both magnetos are on. For engine starting, the key is turned fully clockwise to the START position. (In this function the control is similar to that of a car.)

2.7.5 Electrical services

It is usual design practice to group together the various switches for the electrical services (including the master switch) on the instrument panel. In Figure 96 the switches are of the 'rocker' type - each is selected on by pushing in its top portion and off by pushing in its lower portion.

Similarly, the fuses, spare fuses and circuit-breakers are usually grouped together.



2.7.6 Seats and harnesses

Both of the front two seats are adjustable forwards and backwards, and in some designs, up and down. Correct adjustment of his seat to suit the individual pilot is important - he should be able to reach comfortably all of the various controls. It is equally important to ensure that the seat is firmly locked in position once it has been adjusted.

Restraining harnesses are provided for every seat in the cabin. The front two seats are usually equipped with both lap straps and upper harnesses. These latter are normally of the inertial reel type, enabling freedom of movement when reaching for controls, but locking in the event of impact. For take-off and landing full harness should be worn, with the lap straps tight (but not uncomfortably so). During cruising flight most pilots prefer to retain at least the lap straps in position to guard against the possibility of being thrown around in turbulent air.

2.8 AVIONICS

We shall look at three examples of avionics (aviation electronic equipment) often found in modern light aircraft.

2.8.1 The COM radio

Although there is no mandatory requirement, most modern light aircraft are fitted with communications (COM) radio, since lack of this equipment limits the operational flexibility of the aircraft, and may preclude entry into certain regions of airspace.

The COM radio provides air-to-ground, ground-to-air and, if needed, air-to-air speech communication. It operates in the Very High Frequency (VHF) range. At the present time, the frequencies (or channels) available lie in the range 118 to 137 Megahertz (MHz). The COM radio is the means by which the pilot communicates with Air Traffic Control (ATC) when required.

The aircraft's COM equipment consists primarily of a transceiver (transmitter-receiver) located in the instrument panel. Figure 97 shows a typical installation.



The  frequency selector knobs allow any of the COM frequencies to be selected. Note that the frequency thus selected is not the active one, which is the frequency displayed on the left side of the transceiver. Pressing the transfer (XFER) button will swap the two frequencies so that the last selected frequency is now active. This arrangement confers two advantages:

(a) the next anticipated frequency can be preselected, reducing pilot workload;

(b) in the event contact can not be established on the new frequency, the previous frequency can be quickly reselected.

The transceiver will both transmit and receive on the active frequency.

A headset worn by the pilot is used in conjunction with the transceiver. The headset consists of two earphones to one of which is attached a boom-mounted microphone. A lead from the headset plugs into jack points located at a convenient position on the instrument panel (Figure 98).



To transmit speech, the pilot presses down the transmit button mounted on the control wheel (the button is connected to the transmitter) and then speaks into his microphone, held adjacent to his mouth by the boom. When his message is finished he releases the transmit button.

Messages from the receiver are heard in the earphones. Note that whenever the transmitter is in use, the receiver is automatically deactivated. If the pilot forgets to release his button after transmission, he will not be able to hear incoming signals.

The volume control (Figure 97) is used to control the volume of reception heard in the earphones. (It does not control the strength of transmission - this cannot be varied by the pilot.) The squelch control adjusts the sensitivity of the receiver. When reception is strong the control can be turned down, reducing sensitivity and hence eliminating background noise. If reception is weak the control is turned up, increasing sensitivity (although this will bring the disadvantage of more background noise).

If there are two pilots in the aircraft, conversation between them is facilitated by the use of the radio's built-in intercommunication ('intercom') system. When the headsets are worn and the intercom is switched on, conversation is relayed from both microphones to both pairs of earphones. Note that the transmit button is not used for intercom.

The aerial for the transceiver is usually located externally on the fuselage.

2.8.2 The transponder

The transponder is an electronic installation fitted to many light aircraft that can send coded radio signals to ATC. When required, ATC will ask the pilot to 'squawk' a specific 4-digit code, which is achieved by the pilot selecting these digits on the transponder (Figure 99). In this way ATC can identify the individual aircraft on their radar screens, which pictorially show the geographic locations of all the aircraft under their control.



The push-button controls allow the pilot to select the following modes:

OFF - transponder unpowered;

STBY - transponder in standby mode, electrically powered but not sending or receiving signals;

ON - transponder sending selected squawk;

ALT- transponder sending selected squawk and aircraft altitude data;

IDENT - transponder momentarily sends identification signal when requested by ATC to emphasise the aircraft position on the ATC radar screen.

2.8.3 The GPS map display

Many light aircraft are fitted with Global Positioning System (GPS) equipment, which operates in a similar manner to that found in cars, except that the display is usually set to show a plan view of the aircraft geographic location in the form of an electronic map. Besides the features usually to be found on conventional maps, the aircraft's GPS display will show airfields and the boundaries of regions of airspace where specific rules apply to aircraft flying within them.

2.9 THE CHECKLIST

The reader will appreciate the importance of checking that the aircraft and all its components and controls are serviceable before an intended flight - it is clearly better to discover any unserviceabilities while the machine is still on the ground.

Again, the pilot must be able to operate the various controls in the correct manner and sequence.

Both of these considerations explain the need for a checklist. The checklist usually takes the form of a flip-card booklet and contains instructions advising the pilot of the correct sequence of actions that must be performed at appropriate times before, during and after flight, in order to operate the aircraft safely and efficiently.

The checklist is usually divided into two sections, the first covering normal operating procedures. The second section, emergency procedures, lists the sequence of actions that must be carried out to deal with various emergency situations should they arise.

In some checklists brief technical details of the aircraft may be included, examples being loading limitations and fuel consumption rates.

Most pilots who fly regularly take the trouble to memorize certain parts of the checklist, such as the various actions that must be performed during flight, and the more important of the emergency procedures. In this latter case, use of memorized procedures rather than reading from the checklist will allow the pilot to deal with the situation more expeditiously.

Note

The descriptions in this Section are generally representative of light aircraft technology as it is at the present time. It must be emphasised that individual designs of aircraft that the reader may encounter may well differ from the examples given here to a lesser or greater extent.

The source of technical information for any aircraft is its Flight Manual.

   





3 DETAILED THEORY OF FLIGHT






In this Section the basic theory of flight will be expanded and the forces introduced in Section 1 will be analysed in greater detail. Additionally, some applications to practical flight will be mentioned.

3.1 WEIGHT AND CENTRE OF GRAVITY

It will be recalled that the centre of gravity (CG) of an aircraft was defined as the point from which the single equivalent total weight force acted (Figure 2) .

3.2 LIFT

Lift is generated by the aircraft's wings and is derived from two features of the wing, namely its angle of attack and its shape.

3.2.1 Angle of attack

Consider a flat plate moving through the air at a shallow angle. The effect of the motion is to cause an airflow past the plate, as shown in Figure 100.


 
The angle at which the plate meets the airflow is called the angle of attack. Because of the air pushing underneath, the plate experiences a force which attempts to move it upwards and backwards. This force is called the total reaction (R). It can be considered as split, or resolved, into two separate forces, or components, one acting upwards, the other acting backwards, opposing the direction of motion of the plate.

The upward-acting component is lift (L) and the backward-acting component is drag (D), as shown in Figure 100.

It will be appreciated that, when the angle of attack is zero, no lift will be generated (although, of course, there will be some drag).

3.2.2 Wing shape

The flat plate in Figure 100 is not an efficient producer of lift. Far better is the typical wing aerofoil shape, shown in Figure 101 with the airflow pattern that surrounds it when it is moving through the air and inclined at a small angle of attack.



Not only does this shape generate more lift than the flat plate under the same conditions, but its drag is less. Notice that, in Figure 101, a line has been drawn from the leading edge of the wing section to the trailing edge. This is called the chord line and is the reference for assessing the angle of attack of the wing. Notice, too, that the total reaction acts from a point on the wing about a third of the way back from the leading edge.

How does the aerofoil shape generate lift? As it moves through the air the curved shape speeds up the flow over the upper surface, which causes the air to lose some of its natural pressure, since the laws of physics demand that the sum total of its energy (that is, its energy of movement and its pressure energy) must remain constant. In other words, wing motion through the air creates a partial vacuum above it. The air at normal pressure underneath the wing therefore exerts a push force upwards and this is the lift force that keeps the aircraft flying. In contrast to the flat plate, the wing will generate lift even when its angle of attack is zero. If the wing is then inclined to the airflow at a shallow angle of attack (as in Figure 101) its lift will increase further, as would be expected.

Notice that the airflow, as it approaches the leading edge of the wing, starts to move upwards, attracted as it is towards the area of reduced pressure. Because of this, the air just ahead of the leading edge is termed upwash. After the airflow has passed the wing, it moves backwards and slightly downwards. Accordingly, the air behind the wing is termed downwash, and its significance will become apparent later on.

As a practical demonstration of lift and downwash, suspend a teaspoon adjacent to the flow of water from a tap and then move it closer until the back of the spoon just touches the flow. The spoon will then be pulled into the stream of water, deflecting it as shown in Figure 102. The pull is, of course, analogous to lift, whilst the deflected water corresponds to downwash. Compare the shape of the teaspoon bowl with the wing section in Figure 101.



3.2.3 Factors affecting lift

3.2.3.1 Angle of attack

As already mentioned, a wing will generate lift even when its angle of attack is zero. However, for a particular speed of airflow, an increase in angle of attack has the effect of increasing the lift, until an angle of about 15 is reached. At or above this angle, called the stalling angle, the airflow past the wing becomes turbulent and no longer follows the shape of the wing. This state of affairs is shown in Figure 103.



This break-up of the airflow destroys to a large extent the lift-generating ability of the wing, although a certain amount of lift is still generated by the 'inclined flat plate' effect. However, if the angle of attack is increased further, even this lift becomes reduced. The stalling angle differs according to the actual shape of the aerofoil section but, for most light aircraft wings, is about 15.

In summary, a graph illustrating the relationship between lift and angle of attack is shown in Figure 104. Remember that the graph assumes that the speed of the airflow is constant.


 
3.2.3.2 Speed

For a particular angle of attack, a wing will generate more lift the faster it moves through the air. Mathematically, the lift generated is proportional to the square of the speed - the wings of an aircraft flying at 200 knots will generate four times as much lift as they do when the aircraft flies at 100 knots, assuming that the angle of attack is the same at both speeds.

3.2.3.3 Aerofoil shape

It has already been seen that the wing shape contributes considerably to its lift-generating ability. The aerofoil shape shown in Figure 101 is typical of that utilised by many light aircraft with cruising speeds of about 100 knots.

If it were necessary to design an aircraft for low-speed flight, some means would have to be found to ensure that the wings were able to generate enough lift for the aircraft to fly. One solution would be to position the wings on the airframe so that they met the airflow at a high angle of attack. However, as will be explained later, this design would incur an unacceptably high drag penalty. A more efficient means of improving the lift-generating ability at low speeds would be the use of a thicker, more highly-curved aerofoil shape.

Conversely, an aircraft designed to cruise at speeds much higher than 100 knots would employ a thinner, less highly-curved aerofoil shape, this shape being more efficient at these speeds. Note that the curvature of the wing is sometimes referred to as its camber.

Consider again the average light aircraft, with its 100 knot cruising speed. For take-off and landing, it would be highly desirable to be able to fly at much lower speeds, so that long runway lengths would not be required. Would it not be advantageous if it were somehow possible to change the shape of the wings so that they were more highly curved? Now refer back to Figure 5. One reason for incorporating flaps into the wing structure is now evident - it is to improve the lift-generating ability of the wings at low speeds by conferring greater curvature to their shape.

3.2.3.4 Wing area

The relationship is as would be expected. Other things being equal, the greater the area of an aircraft's wings, the greater is the lift generated.

3.2.3.5 Air density

Not surprisingly, a wing moving through less dense air generates less lift than it does when moving at the same speed and angle of attack through more dense air.

3.3 TAIL DOWN-FORCE

As previously mentioned, the tailplanes together generate the tail down-force. To understand how this force is generated, it is merely necessary to visualise the airflow past the aircraft (Figure 105).


 
Notice that the downwash strikes the tailplanes at an angle. If the diagram is now turned upside-down, it will be appreciated that the tailplanes will generate a force in the same manner as do the wings. Figure 105 shows the force resolved into a downward-acting component, which is the tail down-force, and a backward-acting component, which is obviously drag.

3.4 DRAG

Drag is the force acting on an object in opposition to the direction of its motion as it moves through a fluid and arises from the resistance of the fluid to the disturbance caused by the object's passage. Every part of an aircraft in flight that is exposed to the airflow generates drag. We have already come across two examples - the wings and the tailplanes.

3.4.1 Factors affecting drag

3.4.1.1 Shape

In Figure 106, two objects are moving through the air at the same speed. One is a sphere, the other a tear-drop shaped streamlined object having the same cross-sectional area. The pattern of the airflow around each object is shown.



The sphere disturbs the airflow more than does the streamlined object, as evidenced by the greater turbulence of its wake. Because of this, the drag generated by the latter is considerably less than that from the former. In fact the streamlined object generates only about one tenth of the drag generated by the sphere - a considerable reduction.

Making use of the principle of streamlining, the designer therefore minimises the drag generated by his aircraft.

3.4.1.2 Speed

The faster an object moves through the air, the more drag it generates. Mathematically, the drag generated is proportional to the square of the relative speed of motion - an object moving at 200 knots will generate four times as much drag as it does when it moves at 100 knots.

3.4.1.3 Size

It is self-evident that a large object would generate more drag than a smaller, similarly-shaped one moving through the air at the same speed.

3.4.1.4 Air density

Not surprisingly, an object moving through less dense air generates less drag than it does when moving at the same speed through more dense air.

3.4.2 Wing drag

The considerations examined above apply to an aircraft's wings as much as they do to any object moving through the air. Of course, the wing shape is highly streamlined, which helps to minimise the drag caused by turbulent wake behind it.

Consider again the pattern of airflow around a wing shape as it moves through the air. Now, aerodynamically, the effective lift acts at right angles to the direction of motion of the airflow in the downwash, as shown in Figure 107. (This is an over-simplification of highly complex theoretical reasoning, but will suffice so far as the following discussion is concerned.) We will represent this effective lift force as L* in the diagram.


 
It will be appreciated that, relative to the direction of motion of the wing, the effective lift force is inclined slightly backwards. The force can be resolved into two components, one acting at right angles to the direction of motion of the wing, the other acting backwards, as in Figure 108.



The upward-acting component is truly lift.

Of course, the backward-acting component is not really lift at all - it is drag. Note that this drag is in addition to that generated by the wing by virtue of its turbulent wake. This additional force is called induced drag and it appears whenever the wing generates lift. The drag component in Figure 101 obviously represents the total wing drag, that is, induced drag together with turbulent wake drag. The factors affecting induced drag are discussed below.

3.4.2.1 Speed

Remembering that induced drag is directly concerned with the generation of lift, and that, for any particular angle of attack, lift increases with the square of the speed of the wing, it is not surprising to discover that induced drag increases according to the same relationship. A wing moving at 200 knots through the air generates four times as much induced drag as it does when moving at 100 knots at the same angle of attack. Since the wing drag caused by turbulent wake also follows the same relationship, it can be appreciated that the total wing drag increases with the square of the speed of the wing, assuming a constant angle of attack.

3.4.2.2 Angle of attack

Figure 109 shows a wing in two different situations. At the top of the diagram the wing is inclined at a shallow angle of attack. At the bottom the same wing is moving more slowly but is inclined at a higher angle of attack, such that it is generating the same lift force.



Now, compare the direction of motion of the downwash in each case. Notice that the downwash from the wing at high angle of attack is deflected further downwards than that from the wing at shallow angle of attack. Accordingly, the effective lift force, acting at right angles to the direction of motion of the downwash, is inclined further backwards, as Figure 109 shows.

Resolving both forces into components, as was demonstrated in Figure 108, it is easy to see that the wing at higher angle of attack will generate considerably more induced drag than at shallow angle of attack, because the effective lift force is inclined further backwards. At angles of attack near to the stalling angle, induced drag is very great indeed.

3.4.2.3 Wing planform

Figure 110 features two aircraft with identical total wing areas. As can be seen, the lower aircraft has long, slender wings compared to the other. In other words, its wings have what is termed a high aspect ratio. The wings of the top aircraft have a low aspect ratio.



It will be remembered that, when the aircraft is in flight, its wings experience a region of reduced air pressure above them and that it is the air at normal pressure beneath that gives the push upwards - the lift.

In the vicinity of the wing tips some of the air underneath, instead of pushing upwards on the wing, leaks round the wing tip. Because of the forward motion of the aircraft, this air spillage follows a spiral path, as shown in Figure 111.



If the view from the side is considered, it will be appreciated that this spiral flow of air will influence the downwash behind the wing, deflecting it further downwards. From what has been said before, it is clear that the induced drag effect is therefore greatest in the vicinity of the wing tips.

Now compare the wing planforms in Figure 110 and notice that the 'vicinity of the wing tips' represents a less sizeable portion of the entire high aspect ratio wing than of the low aspect ratio wing and so, other things being equal, the total induced drag of the former is less than that of the latter. With this in mind, the designer will employ long, slender wings on his aircraft unless structural or other considerations countermand this requirement.

3.4.2.4 Air density

As expected, a wing moving through less dense air generates less induced drag than it does when moving through more dense air at the same speed and angle of attack.

3.4.3 Total aircraft drag

It is clear that the entire aircraft will generate turbulent wake drag when it is flying. Additionally, the wings will generate induced drag. From the foregoing remarks, it can be appreciated that the total drag generated by any particular aircraft will depend not only on its speed, but also upon the angle of attack of its wings.

3.5 ANGLE OF INCIDENCE

The angle of incidence of a wing refers to its inclination relative to the fuselage of the aircraft and is totally unconnected with its angle of attack. The former is fixed when the aircraft is assembled, whereas the latter may be varied by the pilot during flight by appropriate use of the controls. Figure 112 demonstrates both angles.



3.6 WING EFFICIENCY

In 3.2.2 reference was made to wing efficiency. It was said that an aerofoil-shaped wing was more efficient than a flat plate because, at any particular speed and angle of attack, it generated more lift and less drag. A measure of wing efficiency is the ratio of lift generated to drag generated under any particular circumstances. Here, the relationship between lift/drag ratio and angle of attack at constant speed of motion through the air will be considered. Figure 113 demonstrates this relationship for a typical light aircraft wing.


 
An interesting conclusion can be drawn - although the wing generates more lift at higher angles of attack, it also generates so much drag that its efficiency is less. At angles of attack above the stalling angle, efficiency decreases markedly, as would be expected.

Figure 113 also shows that the most efficient angle of attack is about 4 for the typical aircraft wing. Furthermore, since the effect of varying speed is the same on both lift and drag, this particular angle of attack will be the most efficient for any speed at which the wing moves through the air. The same will be true for any prevailing air density.

On most aircraft designs, the angle of incidence of the wings is about 4, so that in cruising flight with the fuselage in a level attitude, the wings meet the airflow at or near the most efficient angle of attack.

3.7 THE MOTION OF THE AIRCRAFT

Any displacement of the aircraft (for example, rising and falling or swinging left or right of the nose, or rising and falling of one wing or the other) involves the aircraft pivoting about its CG. The motions described above are respectively termed pitching, yawing and rolling and they occur about three imaginary axes at right angles to each other. The axes intersect at the CG, as in Figure 114.



The aircraft pitches about the lateral axis. The tailplanes enable the aircraft to resist pitch disturbances - they confer what is termed longitudinal stability.

The aircraft yaws about the vertical axis. The fin enables the aircraft to resist yaw disturbances - it confers what is termed directional stability.

The aircraft rolls about the longitudinal axis. The wings are arranged in such a manner than they enable the aircraft to resist roll disturbances - the arrangement confers what is termed lateral stability.

3.8 STABILITY

3.8.1 Longitudinal stability

Remember that the tailplanes generate the tail down-force, and that this force, because of the manner in which it is generated, responds to the same factors which affect the lift force generated by the wings.

Suppose that while an aircraft is in level flight, a gust of air disturbs it, causing the nose to pitch upwards. The aircraft will now start to climb, decelerating as it does so (in the same manner as a car tends to slow down as it runs up a hill).

The reduced speed will decrease the tail down-force. From consultation of Figure 2, we can see that in this situation the lift force, acting behind the CG as it does, will now raise the tail, so restoring the aircraft to level flight. Thus although the gust upsets the aircraft, the stability conferred by the tail planes will soon restore it to its original flight path.

If a gust pitches the nose downwards, the aircraft will start to descend, accelerating as it does so. The faster speed will increase the tail down-force and the tailplanes will be pulled down, again bringing the aircraft back to level flight.

3.8.2 Directional stability

Figure 115 shows the view from above of an aircraft which has just been disturbed by a gust of air causing the nose to yaw to the left. The aircraft, because of its inertia, will tend to continue in its original direction of motion. The result is that the airflow is as shown in the diagram.



It can be seen that the fin, which is now at a shallow angle of attack to the airflow, will generate a force in the same manner as the wing generates lift, and that this force will have the effect of swinging the tail to the left, restoring the aircraft to its original flight path.

Of course, the fin will exert an opposite restoring force if a gust causes the nose to yaw to the right.

The aircraft designer can maximise the restoring effect of the fin by one of two means, or a combination of both:

(a) by increasing the area of the fin;

(b) by making the fuselage longer and thereby increasing the distance between the fin and the CG.

It is usual for the designer to limit the directional stability of his aircraft, for reasons which will be explained later.

3.8.3 Lateral stability

On most aircraft, lateral stability is conferred by canting the wings slightly upwards from root to tip. This arrangement is termed dihedral (Figure 116).



Now, imagine that this aircraft has been disturbed by a gust of air which has caused it to roll to the left. The effect is to adjust the disposition of the lift force, which is now tilted to the left) relative to the weight force (which, of course, still acts vertically downwards), as shown in Figure 117.



The effect is that these two forces between them produce a sideways-acting force (F) as shown. This force tends to pull the aircraft to the side, with the result that the airflow now strikes the left-hand side of the aircraft more than the right-hand side. This tendency of the aircraft to move to the side is termed sideslip (to the left in this case), and one consequence of the effect of the airflow is that the left wing has now a greater angle of attack than the right wing, because of its dihedral. (This can be more easily appreciated if a very marked wing dihedral is imagined.) So - the left wing generates more lift, which tends to level the aircraft, restoring it to its undisplaced state.

Of course, the dihedral will also help to level the aircraft after a gust has caused it to roll to the right.

High-winged aircraft tend to be laterally stable by their very design. Figure 118 shows such a machine in a sideslip to the right.



The airflow, striking the aircraft more on the right-hand side, exerts a drag effect on the wing, as shown in Figure 118, which induces the aircraft to roll to the left, restoring it to its original flight path. Because of this natural lateral stability, high-winged aircraft are usually designed with less dihedral than low-winged machines.

3.8.4 Interaction of directional and lateral stabilities

Refer back to Figure 117, which shows an aircraft in a sideslip to the left. Figure 119 represents the view from above, showing the airflow pattern at the fin.



Of course, in Figure 119 the fin will cause the tail to swing to the right, thus restoring the symmetrical airflow pattern past the aircraft. Effectively, the slideslip motion of the aircraft has been cancelled by the action of the fin.

When this occurs, the differential angle of attack effect imparted by the wing dihedral is nullified - there will be no tendency for the lower wing to level itself. Furthermore, while the tail is swinging to the right, the aircraft's right wing is travelling slightly faster through the air than the left. The result is that the right wing generates more lift, aggravating the situation. In other words, the aircraft is laterally unstable - a lowered wing starts off a sequence of events which finally causes it to drop still further.

Now the dilemma of the designer can be appreciated. If the directional stability is too strong, the aircraft will be laterally unstable, even though its wings may Incorporate dihedral. On the other hand, too little directional stability, although allowing the dihedral to carry out its function effectively, will permit the nose of the aircraft to yaw from side to side in gusty conditions, with consequent discomfort to the occupants of the machine.

In practice, a compromise solution is adopted, in which the aircraft is designed to have good directional stability without being too laterally unstable.

3.8.5 Effect of position of centre of gravity on stability

For this discussion, it will be imagined that a heavy box is to be stowed in the cabin of a typical light aircraft. If the box is loaded at the forward end of the cabin, the effect is to move the CG of the loaded aircraft further forward. Conversely, stowage of the box at the rear end of the cabin has the effect of moving the CG rearward. These situations are shown in Figure 120.



3.8.5.1 Effect on longitudinal stability

The tail down-force has the stabilising effect of bringing the nose of the aircraft back to its original position after disturbance. Refer back to Figure 2, which showed the usual disposition of the weight force, the lift force and the tail down-force during flight. It is possible for the pilot to vary the magnitude of the tail down-force, using the aircraft's controls.

Now, suppose that the aircraft is loaded so that its CG is further rearward, coinciding with the point from which the lift force acts, as in Figure 121. In this case, the original tail down-force would now pull the tail down, thus pitching the nose up.



To prevent this, the pilot would have to use his controls to reduce the tail down-force to zero, thus restoring the equilibrium. But now, if a gust upsets the aircraft, the tailplanes will have no stabilising effect, because, regardless of the aircraft's speed, the tail down-force would remain zero.

Taking the argument a stage further, suppose now that the CG was so far to the rear that it was behind the point from which the lift force acts, as in the top image in Figure 122.



It can be seen that the lift force, because it is now acting ahead of the CG, will have the effect of pitching the nose up, and to prevent this from happening, the pilot would have to adjust his controls so that the tail force acted upwards, as in the lower image in Figure 122.

Well, the equilibrium has been restored again, but imagine that a gust now pitches the nose down. The aircraft will, of course, start to accelerate - the tail force will increase, pulling the tail further up, and the situation would quickly develop into a steep dive. In other words, the aircraft is longitudinally unstable. (The reader is left to prove that this instability would aggravate any pitch up of the nose caused by a gust.)

The conclusion can be drawn that, as the CG of an aircraft is positioned further rearwards, so its longitudinal stability decreases, and, at extreme rear CG positions, the aircraft may become longitudinally unstable. Conversely, forward CG positions enhance longitudinal stability. (However, extreme forward positions of CG have a detrimental effect on aircraft controllability, for reasons which will be explained later.)

Now it can be understood why the manufacturer specifies limits for the CG positions in his aircraft. A CG position outside these limits renders the machine unsafe for flight. The importance of correct loading can be appreciated, and one of the pilot's responsibilities is to ensure that this requirement is complied with.

3.8.5.2 Effect on directional stability

In 3.8.2 it was mentioned that the stabilising effect of the fin was enhanced when it was positioned further from the CG. Using this argument, it is plain that rearward CG positions reduce the distance between the CG and the fin, with a consequent reduction in directional stability. Conversely, forward CG positions enhance directional stability.

3.8.5.3 Effect on lateral stability

Because of the interaction described earlier, rearward CG positions, decreasing the directional stability, have the effect of improving the lateral stability, and vice versa. The effect is too small to be of consequence.

3.9 AIR DENSITY

The effects of varying air density on the aerodynamic forces involved in flight have already been mentioned. Now, remembering that the earth's atmosphere becomes less dense with increasing height, the practical results of this phenomenon will be discussed.

As a preamble, the reader is recommended to consult 2.6.4.2 so that he may be familiar with the effect of varying air density on the relationship between IAS and TAS.

Now, suppose that an aircraft is cruising at sea level at 100 knots IAS, with its wings at the most efficient angle of attack. Of course, at sea level, assuming ISA temperature conditions, the aircraft's TAS will also be 100 knots.

However, at an altitude of, say, 5000 feet, it will be appreciated that TAS will be greater than IAS. So, if the aircraft is to cruise at 5000 feet with its wings at their most efficient angle of attack, will 100 knots IAS be the correct speed to fly at (giving a TAS greater than 100 knots), or should a lower IAS be aimed for (to give a TAS of 100 knots)?

To determine the answer, remember that the most efficient angle of attack at 5000 feet will be exactly the same as at sea level (as explained in 3.6). Now, as has been stated, a wing in less dense air generates less lift than in more dense air, assuming that its angle of attack and speed through the air remain constant. Of course, the 'speed through the air' is exactly equivalent to TAS.

So, let us suppose that an IAS is chosen which gives a TAS of 100 knots. But at 5000 feet, the wings set to their most efficient angle of attack will be moving through less dense air and so they will generate less lift than at sea level - the aircraft will therefore be unable to maintain level flight. It would be possible to restore the lift by one of two means:

(a) increase the angle of attack at this speed of 100 knots TAS - but then the wings are less efficient!

(b) increase the TAS until the most efficient angle of attack generates the required lift.

Obviously, (b) is the solution, the aircraft must be flown at a greater TAS. How much greater? Perhaps the reader will have guessed the answer - 100 knots IAS will give exactly the correct TAS for the most efficient angle of attack.

But what about drag? Will the total drag be greater at this higher TAS? No, because the air is less dense than at sea level. In fact, both the lift force and the drag force generated by this particular aircraft, flying at 100 knots IAS with its wings at this particular angle of attack, will always have the same respective values, regardless of height.

Although we have considered only the case of cruising at the most efficient angle of attack, it can be shown that the general aerodynamic characteristics of an aircraft are the same at any particular IAS, regardless of height, and this theoretical conclusion is, indeed, borne out in practice. Thus the limiting speeds referred to in 2.6.4 are in terms of IAS rather than TAS.

3.10 THRUST: THE PROPELLER

Thrust is necessary to oppose drag, so that an aircraft may maintain sustained flight. Thrust is generated by the propeller, which is rotated by the engine. The propeller produces thrust by the same aerodynamic action by which the wing produces lift, and the aerofoil shape of a propeller blade closely resembles that of a wing. For good propeller efficiency, therefore, the blade should meet the airflow at the most efficient angle of attack, which is about 4 for most designs.

The discussion which follows concerns propellers whose blades are attached to the propeller hub at a fixed angle. This angle, called the blade angle, is related to an aerodynamic property of the propeller, termed pitch. Accordingly, our discussion will concern fixed-pitch propellers.

In flight, the direction of the airflow which passes the propeller blades is governed by:

(a) the forward motion of the aircraft;

(b) the rotational motion of the propeller.

Consider a point on one of the blades. Its direction of motion through the air can be determined by reference to Figure 123.



Of course, the airflow direction at the point considered will be in exactly the opposite direction, as shown. So, it is merely necessary to arrange that the blade at that point meets the airflow at the most efficient angle of attack, as in Figure 124.



Note that the blade angle is the angle between the blade chord line and the direction of rotation.

Figure 124 shows the total reaction generated by the blade. In order to appreciate how the thrust is derived, the total reaction can be resolved, not into lift and drag, but into a forward-acting force (T) and a force (Q) acting in the opposite direction to the rotation, as in Figure 125.



Of course, T is the thrust. The force Q is called torque and it opposes the rotation of the propeller. The propulsive force developed by the engine is necessary to oppose the torque and thereby keep the propeller turning so that it may continue to generate thrust.

But now a problem arises - the tips of the blades are moving much faster in the direction of rotation than the roots (Figure 126).



And so the direction of airflow will differ at each location, as shown in Figure 127.



A means must be found to ensure that, at any point on the blade, it is inclined to the airflow at its most efficient angle of attack. This is achieved by designing the blades to incorporate a helical twist so that the blade angle is greatest at the root, and decreases towards the tip, as represented in Figure 128.



3.10.1 Effect of varying engine power output

Let it be assumed that the argument in 3.10 referred to a propeller driven by an engine set to cruising power. Now, suppose that the throttle is opened, to increase power. The sequence of events is that, firstly, the increased engine propulsive force exceeds the opposing torque and therefore makes the propeller rotate more quickly. In other words, there is an increase in RPM.

This has the effect of increasing the both the speed of airflow past the blades and their angle of attack, as shown in Figure 129. The result is that the propeller generates more thrust. Of course, it also generates more torque, which is exactly balanced by the increased engine propulsive force. The reader is left to determine the effect of closing the throttle from cruising power.



3.10.2 Effect of speed on RPM

Consider a point on one of the propeller blades when the engine is set to cruising power and the aircraft is flying at cruising speed at constant height (as was shown in Figure 124). Now suppose that the pilot uses his controls to make the aircraft's nose pitch down. This will cause the machine to descend, and it will therefore start to accelerate.

Reference to Figure 130 shows that the increased forward speed effectively reduces the angle of attack of the blades.



The result is that the total reaction is reduced in magnitude, and, accordingly, both the thrust and the torque. The propulsive force from the engine, exceeding this reduced torque, will make the RPM increase.

The effect of increased RPM, as previously noted, is to increase the blade speed and angle of attack, with consequent increase in thrust and torque. Equilibrium is re-established when the torque again matches the engine propulsive force. The engine will then have stabilised at higher RPM.

The reader is left to prove that, if the aircraft's nose is made to pitch up to make the machine fly more slowly, then the RPM will decrease.

Note that here it is purely the effect of varying speed which causes the RPM to change, and that this occurs even though the throttle setting is unaltered.

From the foregoing remarks, it may be appreciated that if the aircraft is flying at high speed with the throttle set for high power, it is possible that the RPM will exceed the limit imposed by the manufacturer. This overspeeding of the engine may subject it to forces it is not designed to withstand and can be prevented by judicious handling of the engine controls.
 
3.10.3 Windmilling

Suppose that while the aircraft was cruising, the pilot fully closed the throttle, thereby reducing the RPM to idling power. With the consequent large reduction in thrust, the aircraft's drag would make it decelerate. But suppose now that the pilot pitched the nose down so that speed was maintained by the assistance of gravity. The situation would be as shown in Figure 131.



It can be seen that the blade now meets the airflow at such an angle of attack that the total reaction acts in a completely different direction from before, as the diagram shows. Let us now resolve this total reaction into components (Figure 132).



The result is remarkable. Two conclusions can be drawn.

Firstly, the thrust is acting backwards, as drag. In other words, the propeller is attempting to slow down the aircraft. It can be appreciated that a force is necessary to oppose this drag and that of the rest of the aircraft, in order for speed to be maintained. In fact it is purely the effect of gravity, acting on the descending aircraft, which is providing the force to counter these drag effects.

Secondly, the torque is acting in the same direction as the rotation of the propeller - it is assisting the rotation rather than opposing it. Even with the engine delivering no power, the propeller will keep turning, driven by the airflow passing it. This windmilling of the propeller is a reversal of the normal situation - now it is turning the engine.

3.10.4 Propeller efficiency

A measure of propeller efficiency is the ratio of thrust generated to torque generated under any particular conditions. It has been said already that angles of attack for the propeller blades of about 4 were most efficient because of the aerodynamic similarity between the blades and the aircraft's wings, and this is so even though the total reaction is being resolved into thrust and torque rather than into lift and drag.

3.10.4.1 Effect of speed on propeller efficiency

Consider an aircraft making its take-off run. At the low forward speed and with the engine set to maximum power, the situation at the propeller is as shown in the top image in Figure 133.
 


Notice that, because of the aircraft's low forward speed, the propeller blades meet the airflow at a high angle of attack and so the propeller efficiency is lower than optimum - the ratio of thrust to torque is poor. This is a severe penalty and a waste of engine power at a time when the highest possible thrust is required to confer good acceleration to the aircraft so that it quickly becomes airborne.

During cruising flight, the propeller performs much better - with the higher forward speed of the aircraft and the engine set to cruising power, the blade will at or or near its most efficient angle of attack.

The situation during take-off would be improved if the blade angle were lower, to match the low forward speed of the aircraft, as represented in the lower image in Figure133.
 
The reduced angle of attack results in the propeller developing slightly less thrust, but considerably less torque. In the situation shown in the top image in Figure 133 the engine RPM were reduced by the retarding effect of the high torque. Now, with the lower blade angle, the engine will be able to attain higher RPM, with the result that the speed of motion of the propeller blades through the air will be greater. Accordingly, the thrust will now be greater, this effect more than offsetting the reduced angle of attack effect as would be expected with the blades working more efficiently.

But will this propeller with lower blade angle perform satisfactorily in cruising flight, when the aircraft's forward speed is higher? Figure 134 shows that the blade is now at a lower angle of attack than optimum, and so its efficiency is poor.



It seems that a fixed-pitch propeller will be efficient either at low speed, or at higher speed, depending on its blade angle, but not both. Most small aircraft, whose acceleration during take-off is not so critical, are equipped with fixed-pitch propellers designed for optimum efficiency in cruising flight.

Later on we shall look at propellers in which the blade angle can be varied by the pilot according to the phase of flight. These variable-pitch propellers confer efficiency at all aircraft speeds.

   





4   AIRCRAFT HANDLING






Every flight involves a certain amount of preparation beforehand so that, once the flight has begun, the pilot is able to operate his aircraft safely and efficiently. Not surprisingly, the depth of pre-flight planning is related to the complexity of the intended flight - a long cross-country operation will involve more detailed planning than a brief flight in the vicinity of the base airfield.

In this Section, we shall be concerned with the manner in which the controls are used to make the aircraft behave as required during the various phases of flight. It is emphasised that aircraft handling is merely one aspect of the whole concept of flight management.

Throughout any flight the aircraft should be operated in accordance with its checklist. The observations below are of a general nature, applicable to most light aircraft. Initially, we shall be concerned with machines having fixed-pitch propellers and fixed landing gear.

4.1 AIRFRAME LIMITATIONS

The reader will recall that certain limitations are applied to the airframe.

Firstly, the pilot must ensure that before flight the aircraft is loaded in compliance with the loading limitations set down in the Flight Manual.

Secondly, during flight, he must observe the various speed limitations. Besides the limitations discussed in 2.6.4 (flap limiting speed, Vno and Vne), the Flight Manual will specify the maximum speed permitted for flight in turbulent air (to protect the airframe from overstressing) and the maximum crosswind component acceptable for take-off and landing. This limitation is primarily to prevent the control difficulties which would arise from operation in stronger crosswinds.
 
4.2 ENGINE HANDLING AND OPERATING LIMITATIONS

4.2.1 Control of power

To increase engine power output the throttle is opened (by pushing the lever forwards) to attain higher RPM. Maximum power, or 'full power', implies throttle fully opened, or until the maximum permitted RPM are reached, whichever is the more restrictive. (Remember that for any particular throttle setting, RPM increase with the aircraft's speed. With many designs, the throttle may be opened fully at low speeds, but will require closing partially at higher speeds to prevent violation of the maximum RPM limitation.)

Minimum power (idling RPM) is delivered when the throttle is closed (lever fully rearwards).

Of course, increased power results in increased thrust from the propeller, and vice versa.

4.2.2 Use of mixture control

As explained in 2.2.6 the mixture control is used to lean out the mixture for cruising flight. The usual procedure is as follows:

(a) set the mixture control to 'rich' (if not already so);

(b) set chosen RPM with the throttle lever;

(c) move the mixture control lever backwards until the RPM start to decrease;

(d) move the mixture control lever forwards sufficiently to regain the original RPM setting.

The mixture strength is now correct for the chosen RPM setting and cruising altitude. A change in either of these factors will necessitate re-adjustment of the mixture control.

4.2.3 Engine operating limitations

The pilot must observe the engine operating limitations stated in the Flight Manual. The maximum permitted RPM will be specified, as will the restrictions on the use of the mixture control. These latter are often expressed in terms of the power setting above which, and the altitude below which, the control must be set to rich.

Periodically, the pilot should check that the oil temperature, oil pressure and fuel pressure gauge readings are within the limits specified in the Flight Manual.

Excessive oil temperature might occur in warm weather if the engine is operating for long periods at high power. Setting reduced power should rectify the situation.

An abnormally high oil temperature indication accompanied by abnormally low oil pressure signifies malfunction of the oil system and these may be the first symptoms of impending engine failure. If such indications are observed the flight should therefore be terminated as soon as possible and the fault investigated.

Abnormally low fuel pressure signifies malfunction in the fuel system. If the engine-driven pump is at fault, pressure might be restored by switching on the electrically-driven pump. If this remedy has no effect, it is possible that a leak has occurred in the fuel pipeline. In either case the flight should be terminated as quickly as possible, particularly if a suspected leak has occurred, since this is a potential fire hazard.

4.3 PICKETING AND USE OF CHOCKS

When circumstances require, a parked aircraft should be tied down, or picketed, and its wheels chocked.

Incorporated in the lower surface of each wing is a reinforced anchor (or picketing) point. Picketing involves connecting these to similar anchor points embedded in the ground, using suitable tie-down straps. This precaution will prevent any tendency for the machine to be overturned by strong winds. The straps should be neither too loose (which would allow a strong wind to lift the machine) nor too taut (to avoid straining the airframe structure). Figure 135 demonstrates.



The aircraft should be picketed whenever strong winds are blowing or if they are forecast to occur.

Chocks are blocks of metal or wood which are positioned in front of and behind each mainwheel. They prevent movement of the aircraft if the parking brake system fails (Figure 136).



Common sense dictates that whenever the aircraft is to be parked outside for a considerable period of time:

(a) it is picketed;

(b) the wheels are chocked;

(c) the control lock(s) are employed.

4.4 MANHANDLING AND POSITIONING THE AIRCRAFT FOR ENGINE-STARTING

If not already parked so, the aircraft should be manhandled into a suitable position for engine-starting.

Aircraft manhandling requires special care because the airframe structure is relatively delicate. Whilst being strong enough to bear the aerodynamic loads experienced during flight, the structure might easily be damaged by careless handling on the ground.

It is usual for each aircraft to be equipped with a tow bar which can be connected to the nosewheel landing gear unit, enabling the machine to be pushed, pulled or steered by hand for positioning. When this is the case, no part of the airframe structure should be pushed or pulled. If no tow bar is provided the machine should be manhandled in accordance with the manufacturer's instructions. Vulnerable areas, such as the flight controls, should never be pushed.

When the aircraft is in the required position the brakes should be applied with the park brake control and the tow bar disconnected and stowed. The machine should be positioned such that there are no buildings or other aircraft immediately behind the tail. This is because, whenever the engine is running, the propeller sends back a swirl of air - the 'propwash'. Any loose stones on the ground in the vicinity of the aircraft might be flung backwards by the propwash - the precaution mentioned above will ensure that such flying debris will not cause damage. Under no circumstances should the engine be started while the aircraft is inside a hangar.

4.5 PROPELLER HANDLING

Any person who comes into contact with a rotating propeller risks lethal injury. A propeller at rest should never be turned by hand unless absolutely unavoidable (for manhandling the aircraft clear of others in confined hangarage or for close inspection of the individual blades).

If the propeller must be turned for these or any other reasons, the following precautions should be taken:

(a) apply the brakes with the park brake control;

(b) ensure that both magnetos are switched off;

(c) assume that the propeller is 'live'. If the magneto control is faulty it is possible that turning the propeller might start the engine. By ensuring that one's body is kept clear of the propeller arc (Figure 137) injury will be prevented should this unlikely occurrence happen.



(A few designs of aircraft have no electric engine starter. Their engines are started by turning the propeller by hand after the magnetos have been switched on. This technique requires special training for the personnel involved.)

4.6 AIRCRAFT INSPECTION

Before starting the engine, the aircraft and its components should be inspected for serviceability in accordance with the checklist. The pilot should check that sufficient fuel and oil are carried for the intended flight, allowing adequate reserves for contingencies. He should ensure that:

(a) picketing straps are removed and stowed or placed well clear of the aircraft;

(b) the chocks are removed and stowed or placed well clear (once the brakes have been applied by the park brake control);

(c) the control lock(s) are removed and stowed;

(d) the ground immediately ahead of the machine's nose is clear of obstructions.

4.7 ENGINE STARTING

Once the pilot has entered the cabin he should close and latch the doors, adjust and lock his seat in position and fasten his harness. Internal serviceability checks should be actioned as detailed in the checklist.

Just prior to starting, the pilot should check visually that no personnel are in the vicinity outside the aircraft, especially near the propeller or behind the tail.

The starting procedure should be carried out in compliance with the checklist. Once the engine is running, check that the oil pressure is satisfactory. (If it is not, stop the engine again - it is obviously unserviceable; the aircraft should be grounded until the fault has been rectified by qualified engineers.) Adjust the throttle lever to achieve the recommended RPM for warming up.

Verify the correct function of each magneto by switching the key from BOTH to R and back to BOTH, and then to L and back to BOTH. With the key at R or L the RPM should have decreased slightly (since combustion in the cylinders is less efficient) and should have re-attained the warm-up setting when the key has been switched back to BOTH. If the engine stops when the key is at R or L then that magneto is evidently faulty. If no RPM drop is observed then the magneto control switch is faulty. In either case the engine is unserviceable.

After starting, electric services such as the radio and the anti-collision lights can be switched on if desired, and the DI synchronised with the magnetic compass.

4.8 TAXYING

'Taxying' is the word used to describe ground manoeuvring of the aircraft under the thrust from the propeller.

Before moving away from the parking position the pilot should first check that the vicinity of the aircraft is clear of obstructions. He should then close the throttle and release the park brake control. A gentle increase in power will now get the machine moving forward.

As soon as the aircraft starts to move the throttle should be closed and the toe brakes applied to check their operation. After releasing the brakes again the throttle can be used as before to start the machine moving forward.

Although modern aircraft are not difficult to manoeuvre on the ground, they are not as easily controlled as cars. For this reason they should not be taxied at high speed. A general rule when manoeuvring near buildings or other aircraft or on poor ground is not to taxi faster than a brisk walking pace.

Once the throttle has been used to start the machine moving forward, it should be partially closed again so that the engine delivers just enough power to maintain taxying speed. To slow down, the throttle should be closed. It is noteworthy that the propeller generates a certain amount of thrust even when the throttle is fully closed. Therefore to slow down further, or to stop, the brakes must be applied.

To turn the machine to the left or right the rudder pedals are moved as appropriate to activate the nosewheel steering mechanism. Tighter turns may be made by differential use of the brakes. (For example, pushing forward the left pedal will steer the aircraft to the left. Applying the left toe brake will tighten the turn.)

When taxying, the control wheel should normally be held in the central position. On some types of aircraft it is recommended that certain control wheel inputs should be made when taxying in conditions of strong wind in order to assist with control. The Flight Manual will specify the inputs to be made.

At all times the pilot should look out for the presence of obstructions or poor ground so that avoiding action can be taken in good time.

The following flight instrument checks should be made when convenient during a turn to the left and one to the right:

(a) the TBI indicates appropriate turns;

(b) the DI heading indications decrease when turning left and increase when turning right;

(c) the magnetic compass heading indications behave as in (b);

(d) the AI continues to show a level attitude with the wings level.

Whenever the engine is running, a little oil escapes past the pistons into the cylinders. At power settings well above idling RPM (throttle closed) this oil is burnt along with the fuel-air mixture. At lower power settings, however, some of the oil forms a film over the spark plug electrodes, inhibiting spark occurrence, possibly to the extent that the engine runs roughly and fails to deliver the power that it should when the throttle is opened again. To guard against this possibility, the throttle should be set to warm-up RPM whenever the aircraft is brought to a halt for more than a few seconds. This precaution should ensure that all escaping oil is burnt and the spark plug electrodes are therefore kept dry.

4.8.1 Brake failure

In the event of failure of the brakes the engine should be stopped immediately (unless its use will definitely assist the pilot in dealing with the situation) and the aircraft steered away from obstructions while it decelerates. It is worth remembering that grass will slow down the machine more quickly than hard ground. Steering onto grass (if practicable) will thus help to bring the aircraft to a halt.

4.9 ENGINE TESTING

With most designs of engine the manufacturer will recommend a testing procedure to be carried out before take-off as detailed in the checklist.

Since the test involves the use of a high power setting, the following precautions should first be taken:
 
(a) the aircraft should be halted on firm, non-slippery debris-free ground and the brakes applied firmly with the park brake control;

(b) the machine should be positioned, when practicable, facing nose into the wind. This will assist engine cooling and will help to prevent forward movement of the aircraft;

(c) the vicinity of the machine, especially behind the tail, should be checked as being clear of personnel and other aircraft;

(d) the oil temperature, oil pressure and fuel pressure should be verified as being within the specified limits.

Primarily, the purpose of the test is to check the efficiency of the ignition system, although other engine controls also require exercising. Typically, the engine test might proceed as follows:

(a) open the throttle to achieve a specified RPM indication (checking that the brakes are preventing the aircraft from moving forwards - if not, the throttle should be closed again and the brakes re-applied);

(b) switch off each magneto individually as described in 4.7. With the key at R or L the RPM will have decreased slightly from the original setting. In each case the RPM drop should not exceed the allowed maximum and the engine should continue to run smoothly. When the key is switched back to BOTH, the RPM should increase again to the original setting. If the RPM drop at R or L is more than the permitted maximum, or if the engine runs roughly, the cause might be oil on the spark plug electrodes, in which case the problem can usually be cured by running the engine (with the key at BOTH) at the test RPM setting for a minute or so to burn away the oil (making sure that the various gauge indications remain within the specified limits) and then repeating the ignition test. If the result is still unsatisfactory the engine must be considered unserviceable for flight. (Ensure that the key is set to BOTH after completion of the test);

(c) select carburettor heat 'on'. If the control is functioning correctly, the RPM will decrease slightly from the original setting (since with hot air entering the carburettor, the engine delivers less power). Select the control back to 'off' and verify that the RPM increase again to the original setting. If no RPM drop is observed when the control is 'on', or if the RPM fail to increase again after reselection to 'off', the engine is unserviceable. Besides testing for serviceability, the procedure just described will remove any ice which might have built up in the carburettor whilst taxying;

(d) check that the indications of the oil temperature, oil pressure, fuel pressure and suction gauges and the ammeter are within the permitted limits;

(e) close the throttle and check that the idling RPM and oil and fuel pressure indications are within the permitted limits;

(f) reset warm-up RPM with the throttle lever.

If any part of the test procedure gives unsatisfactory results, the flight should be abandoned and the aircraft taxied back to the parking area for the attention of engineers.

4.10 PRE-TAKE-OFF CHECKS

The pre-take-off checks are actioned by the pilot (using the checklist) so that the various controls are correctly set for the take-off and initial climb. Typically, they include the following:

4.10.1 Airframe

(a) trim wheel(s) set as recommended;

(b) flaps set as recommended (check visually that both flaps are at the selected position);

(c) check for full availability of unobstructed movement of the control wheel (and rudder pedals when spring-type nosewheel linkage permits such movement);

(d) doors closed and latched;

(e) harness tight;

4.10.2 Engine

(a) throttle friction tightened to personal preference;

(b) mixture control set to 'rich';

(c) carburettor heat control set to 'off;

(d) magneto control key set to BOTH;

(e) oil temperature, oil pressure and fuel pressure within the specified limits;

(f) fuel cock selected to draw fuel from a tank with sufficient quantity;

(g) all fuel tank quantity gauges checked for sufficiency;

(h) electrically-driven fuel pump switched on;

4.10.3 Instruments

(a) altimeter set to QFE or QNH;

(b) AI appears erect, that is, shows a level attitude with the wings level;

(c) DI synchronised with magnetic compass;

(d) suction gauge indication within the specified limits;

4.10.4 Electrical services

(a) anti-collision lights switched on;

(b) pressure head heater switched on if weather conditions warrant;

(c) radio (and other avionics when appropriate) controls correctly set;

(d) ammeter indication satisfactory.

4.11 PROCEDURE AFTER LANDING

After landing, the appropriate actions in the checklist are carried out and the aircraft is taxied to the parking position. The brakes are applied with the park brake control. The engine is shut down in accordance with the recommended procedure. All electric services, including the master switch, are switched off and the fuel cock is turned off.

If the aircraft is to be left unattended for a considerable length of time, the pilot should carry out the actions described in 4.3 and close and latch all doors after vacating the machine. If appropriate, covers can be fitted to protect the aircraft against adverse weather.

4.12 EFFECTS OF FLIGHT CONTROLS IN FLIGHT

During flight, an aircraft's stability helps it to resist pitching disturbances (nose rising and falling), yawing disturbances (nose swinging left and right) and rolling disturbances (rising and falling of one wing or the other). The flight controls are used to overcome this natural stability when it is desired to change the flight path.

4.12.1 Effect of elevators

Suppose that while the aircraft is in cruising flight, the pilot moves the control wheel rearwards. The result will be to displace the elevators upwards. Effectively, the camber of the tailplane-elevator assembly and its 'inverted angle of attack' will have been increased, and so the tail down-force will increase in magnitude (Figure 138).


 
It is clear that the tail of the machine will be pulled down by the increased force, correspondingly pitching the nose up. For aerodynamic reasons (which will be explained later) the aircraft will also lose speed.

As the machine decelerates, of course, the speed of the airflow over the tailplane-elevator assembly will decrease, thus reducing the magnitude of the tail down-force. Eventually, the nose will pitch up no further even though the elevators are still displaced, because the balance of the weight, lift and tail down-forces will have re-attained equilibrium. The IAS will then have stabilised at a lower figure.

For the pilot in the cabin, therefore, the effect of moving the control wheel rearwards will be to pitch the nose of the aircraft to a higher attitude and to cause a reduced IAS (Figure 139).



Further rearward movement of the control wheel will result in an even higher attitude and another reduction in IAS.

Conversely, forward movement of the control wheel pitches the nose down and will result in the aircraft eventually stabilising at a lower attitude and at an increased IAS, as shown.

4.12.2 Effect of ailerons

Suppose that, in cruising flight, the pilot moves the control wheel to the left. The result will be to displace the left aileron upwards and the right downwards. Effectively, the camber and angle of attack of the wing at the left tip are reduced and at the right tip they are increased (Figure 140).



Accordingly, the lift generated at the left wing tip is reduced, and that at the right wing tip is increased. The lift imbalance will make the aircraft roll to the left. Conversely, movement of the control wheel to the right will make the machine roll to the right. In either case, the rolling motion will continue as long as the ailerons are displaced. If the control wheel is centralised, the aircraft will remain in whatever bank attitude it happens to be at. Figure 141 shows the ailerons used to select and maintain an angle of bank of 30 to the left. The view from the pilot's seat is shown in Figure 142.





4.12.2.1 Consequence of bank

Whenever an aircraft is banked, the sequence of events described in 3.8.4 occurs as a result of the predominance of the machine's directional stability over its lateral stability. To recap:

(a) the machine sideslips in the direction of the bank;

(b) its nose yaws in the same direction under the stabilising influence of the fin.

The pilot can visually detect the yaw as a change in the aircraft's heading relative to a chosen directional reference point (such as a building) at some distance straight ahead of the nose (Figures 143 and 144).





Note that this yaw is not directly caused by the effect of the ailerons, but arises as a consequence of bank. By yawing, it will be appreciated that the aircraft turns; it will continue to turn as long as the bank is maintained, with the fin nullifying the sideslip resulting from the bank.

4.12.3 Effect of rudder

The result of pushing forward the left rudder pedal is to displace the rudder to the left, effectively cambering the fin-rudder assembly and inclining it to the airflow at an angle of attack (Figure 145).


 
Accordingly, the assembly will generate a force which will pull it to the right, as shown. Thus the tail of the aircraft will swing to the right and its nose will yaw to the left.

Conversely, pushing forward the right pedal will make the nose yaw to the right. In either case, the yawing motion will continue as long as the rudder is displaced. The pilot can visually detect the yaw as a change in the aircraft's heading relative to a chosen directional reference point (Figure 146).



4.12.3.1 Consequence of yaw

If the aircraft is yawing under the influence of its rudder, the airflow past the fuselage is not symmetrical. In this situation the machine is said to be skidding (from the motoring analogy) because even though it is turning, its path through the air is not in the same direction as its heading. Figure 147 illustrates.



If the aircraft features dihedral wing attachment, the asymmetrical airflow will effectively meet the two wings at different angles of attack. In Figure 147 the aircraft's right wing will meet the airflow at a greater angle of attack than the left wing. (This effect is more easily appreciated if marked wing dihedral is imagined.) The right wing will therefore generate more lift than the left wing and the machine will start to roll to the left.
 
Similarly, application of right rudder (by pushing forward the right pedal) will produce a yawing motion which will cause the machine to start to roll to the right.

Note that the roll is caused not directly by the application of rudder, but occurs as a result of the effect of the asymmetrical airflow on the physical features of the aircraft designed to impart lateral stability (wing dihedral or high-winged layout).

4.12.3.2 Propwash-induced yaw

In cruising flight the propeller is generating thrust to overcome the aircraft's drag. As it revolves, the propeller throws back a spiral propwash. If, as is the case with most designs, the propeller revolves clockwise (as viewed from the cabin) this spiral airflow will impinge on the left-hand side of the fin-rudder assembly, as shown in Figure 148.



Of course, the result will be to make the nose of the machine yaw to the left.

This propwash-induced yaw can be eradicated by sufficient application of right rudder, which will enable the fin-rudder assembly to generate a force opposing the effect of the propwash.

If the propeller revolves anticlockwise, application of left rudder will be needed to counteract propwash-induced yaw.

4.13 EFFECT OF VARYING PROPWASH STRENGTH

Let it be assumed that our aircraft is in cruising flight, with sufficient rudder applied to prevent propwash-induced yaw. Now, suppose that the pilot increases the engine power. Of course, the strength of the propwash will increase correspondingly. There will be three consequences.

Firstly, the speed of the airflow over the inner areas of the wings, and hence the downwash flowing past the tailplane-elevator assembly, will be greater. The assembly will therefore generate a greater tail down-force even though the elevator position is unchanged. Accordingly, the aircraft's nose will pitch upwards. Of course, the pitching motion can be prevented by sufficient forward movement of the control wheel.

Secondly, the propwash-induced effect described in 4.12.3.2 will be more marked, and the aircraft will start to yaw. The motion can be prevented by further application of rudder as appropriate.

Thirdly, the greater speed of airflow over the elevators and rudder will enhance their effectiveness. In other words, the aircraft will respond more rapidly to forward and rearward movement of the control wheel and to movement of the rudder pedals. The ailerons lie outside the propwash and so their effectiveness is not enhanced.

If the engine power is reduced the consequences described above will be reversed, as expected. The rudder applied for counteracting the effect of the propwash in cruising flight will make the aircraft yaw in the opposite direction to that caused by the propwash. The yawing motion can be prevented by reducing the rudder application.

4.14 EFFECT OF VARYING AIRSPEED ON FLIGHT CONTROLS

Not surprisingly, the effectiveness of all the flight controls becomes enhanced when the aircraft is made to fly at higher IAS, and reduced at lower IAS.

4.15 EFFECT OF VARYING CONTROL MOVEMENT

The greater the movement of the control wheel or rudder pedals, the greater is the displacement of the associated control and so the aircraft's response is enhanced.

4.16 EFFECT OF FLIGHT CONTROLS IN DISPLACED ATTITUDE

It was mentioned in 3.7 that the motions of pitching, rolling and yawing occur about three imaginary axes at right angles to each other intersecting at the aircraft's CG. The reader will appreciate that this remains the case when the machine is in other than a 'wings level' attitude. For example, forward and backward movement of the control wheel when the aircraft is banked will still result in pitching motion about the lateral axis. Figure 149 illustrates pitching motion (top image) and yawing motion (bottom image) in banked attitudes with the view from the cabin.



4.17 FUNCTION OF FLIGHT CONTROLS

Having discussed the effect of the flight controls, we can now consider their specific functions, in other words the way they are used to control the motion of the aircraft.

4.17.1 Function of elevators

We have seen that the elevators can be used to make the aircraft adopt and maintain any desired pitch attitude. In practice, the attitude is chosen to achieve either one of two purposes, depending on the phase of flight. They are:

(a) to control the vertical flight path of the aircraft;

(b) to control the IAS.

4.17.2 Function of ailerons

The ailerons are used to select and hold any chosen angle of bank when it is desired to make the aircraft turn and to remove bank when straight flight is desired. Remember that the turning effect results as a consequence of the directional stability imparted by the fin.

4.17.3 Balance: function of rudder

The reader will appreciate that the fuselage will generate minimum drag when the airflow past it is symmetrical (viewed from above). The aircraft's flight is then said to be balanced. Any asymmetry of airflow will result in increased total aircraft drag. In this case the flight is said to be unbalanced.

We have encountered the two causes of asymmetric airflow, namely sideslip and skidding, in both of which situations the flight is therefore unbalanced. Remember that sideslip occurs as the initial consequence of bank. Skidding results from the yaw arising from application of rudder or from the effect of propwash on the fin-rudder assembly.

As we have already seen, the aircraft's directional stability automatically cancels any tendency of the machine to sideslip. In the case of propwash-induced skidding, however, there is no automatic corrective action. The pilot must prevent the skidding by counteracting the yawing motion with suitable application of opposing rudder.

With the wings level, propwash-induced yaw can be detected as change of heading. The resultant skidding can therefore be eliminated by choosing a directional reference point and countering any tendency of the nose to yaw away from it by appropriate use of the rudder pedals. For example, yaw to the left can be prevented by pushing the right pedal to a more forward position, so that the aircraft's nose remains aligned with the reference point.

If the view ahead is obscured by mist or haze, or by the engine cowling in a high attitude, or if it is featureless as during flight over the sea, some other means must be found to detect the yaw. The balance indicator will serve this purpose.

The balance indicator makes use of the fact that whenever the aircraft yaws, it is turning. Consequently, the ball in the balance indicator is thrown outwards (in the same manner as the occupants of a car tend to be flung outwards when it turns a corner). As an example, in the situation shown in Figure 148 the ball would take up the displaced position shown in Figure 150.



It can now be appreciated that displacement of the ball from the central position is indicative of unbalanced flight. If, in the example above, the propwash-induced yaw is eradicated by application of right rudder, the aircraft will cease turning and the ball will revert to a central position, confirming balanced flight.

The purpose of the rudder is thus to enable the pilot to maintain balanced flight (and hence minimise drag) by preventing asymmetry of airflow. Displacement of the balance indicator ball to the right requires movement of the right rudder pedal to a more forward position to centralise it, and vice versa.

There are two methods available to the pilot for turning the aircraft. He can either apply rudder or select a banked attitude. In the former method, we have discovered that although the aircraft will turn, the flight will be unbalanced. For example, suppose that in the example above the propwash-induced yaw has been corrected as described, and that the pilot then applies more right rudder. The aircraft will yaw (and turn) to the right and the balance indicator will be as shown in the top image in Figure 151. Evidently, the flight is unbalanced.



If the alternative method of turning is used, that of adopting a banked attitude to the right, the ball in the balance indicator will again be thrown outwards. Because of the bank, however, the ball will lie centrally in the indicator.

The flight is shown to be balanced (Figure 151 bottom image), which is to be expected, since we know that the airflow past the fuselage is symmetrical in this situation. Application of bank is obviously a more efficient method of turning the aircraft.

To summarise: the rudder is used to balance the flight of the aircraft in accordance with the disposition of the ball in the balance indicator. In practice, this technique is used even when there is an adequate directional reference point ahead of the aircraft, since in cruising flight the propwash-induced yaw is often not marked enough to detect visually, whereas the balance indicator ball will show even slight unbalance. Note that the rudder is not used to turn the aircraft (except in circumstances which will be discussed later).

4.18 FUNCTION OF TRIM TABS

We have already seen that rearward movement of the control wheel will result in a higher attitude.

Of course, to hold the aircraft in this higher attitude, the elevators would have to be retained in the upward-displaced position, against the restoring effect of the airflow which would attempt to return them to their undisplaced position.

So far as the pilot is concerned, he would detect the restoring effect of the airflow as a pull force needed to hold the control wheel in its rearward position. It would be desirable if another means could be found to hold the elevators in the displaced position, and thus relieve the pilot of the need to maintain the pull force on the control wheel.

The effect is achieved by rotating the elevator trim wheel backwards and downwards, which in turn displaces the trim tab downwards. The airflow, striking the displaced tab, tends to push it, and therefore the elevators, upwards. Figure 152 shows the airflow following the contours of the tailplane-elevator-tab assembly after such adjustment of the tab position.


 
Sufficient movement of the trim wheel will enable the elevators to hold the aircraft in the desired attitude without the need for the pilot to apply a pull force on the control wheel. The aircraft is then said to be trimmed in the chosen attitude; it will retain this attitude even if the pilot relinquishes his hold on the control wheel.

Conversely, sufficient rotation of the trim wheel upwards and forwards will enable the aircraft to retain any chosen lower attitude set by forward movement ol the control wheel without the need of a continuous push force.

In a similar manner, the rudder trim tab is used to hold the rudder in the displaced position necessary to counteract propwash-induced yaw and so maintain balanced flight. If the tab is adjustable in flight, it can be made to nullify the foot force on the rudder pedals needed to displace the rudder as appropriate for any chosen power setting. When the foot force has been nullified, the aircraft is in trim; it will fly in balance even if the pilot removes his feet from the rudder pedals. Rotation of the trim wheel to the right will move the tab to the left, which in turn will hold the rudder in a displacement to the right, and vice versa (Figure 153).



If the tab is of the fixed type, it is usually set for cruising flight. At power settings other than the cruise setting, the pilot will have to apply a force on the rudder pedals to hold the appropriate rudder displacement needed for balanced flight. In other words, the aircraft will only be in trim in cruising flight.

Although it acts in precisely the same manner, the function of the aileron trim tab differs from that of the elevator and rudder tabs. When the aircraft is constructed, manufacturing imperfections make it virtually impossible to ensure that the wings are precisely formed and that each is attached to the fuselage at precisely the same angle of incidence.

Consequently, it may well be that during flight the machine tends to roll gently to the left or right as a result of the ensuing lift imbalance. Of course, the ailerons can be used to prevent this tendency, and the need for the pilot to physically hold the control wheel in the position necessary to maintain the appropriate aileron displacement can be obviated by adjustment of the aileron trim tab.

As an example, suppose that the aircraft tends to roll to the left in flight, and that the tab is attached to the left aileron. In this case, the tab should be adjusted upwards, so that the airflow tends to hold the left aileron in a downward-displaced position (and, by virtue of the control linkage, the right aileron in an upward-displaced position).

The adjustment is made by trial-and-error to cancel out the machine's rolling tendency. Once set, it is unlikely that the tab would need readjustment unless the aircraft were to be reassembled after dismantling.

If, in the example just described, the tab were attached to the right aileron, it would need to be adjusted downwards to counteract the tendency of the aircraft to roll to the left.

4.18.1 Trimming technique

To change the pitch attitude of the aircraft the control wheel is moved as appropriate. The pilot should then assess whether a pull force or a push force is necessary on the control wheel to hold the aircraft in the chosen attitude. If the former, he should turn the elevator trim wheel backwards and downwards to nullify the control wheel force; if the latter, the trim wheel should be turned upwards and forwards.

If the aircraft has been trimmed correctly it will retain the chosen attitude even when the pilot's hold on the control wheel is relinquished. Any tendency for the attitude to change when the control wheel is released shows that the machine is not in trim. To rectify the situation, the control wheel should be used to regain the desired attitude and then the trim wheel turned as appropriate. Note that the control wheel, and not the trim wheel, is used to make attitude corrections.

As already explained, the rudder is used to balance the flight of the aircraft. If a push force on the left pedal is needed to centralise the balance indicator ball, and the rudder features a tab adjustable in flight, the rudder trim wheel should be rotated to the left to eliminate the force. Conversely, if a push force on the right pedal is needed, the wheel should be rotated to the right.

4.18.2 Trim changes

We have seen that two consequences of increasing power are the tendencies of the aircraft to pitch up and to yaw, as was described in 4.13. These effects are called 'trim changes', because the machine will tend to depart from its original trimmed attitude and to fly out of balance.

The pitching effect is a change in longitudinal trim and the yawing effect a change in directional trim. These trim changes can be counteracted by the pilot by appropriate movement of the control wheel and rudder pedals. If the control forces are now eliminated with the trim wheels, the aircraft will again be in trim.

Decreasing power will result in opposite trim changes. Later on in this book we shall encounter another cause of longitudinal trim change.

4.19 THE FLIGHT PATH: THE THIRD DIMENSION

An aircraft differs from a surface vehicle in that it is capable of movement in three dimensions rather than the two available to the latter - it can move upwards and downwards. In other words, the aircraft can be made to climb, descend or fly level as desired by the pilot.
 
We will now consider the factors that control the flight path in this third, vertical, dimension.

Firstly, remember that there are five forces acting on the machine in flight. They are: the weight force (W), the lift force (L), the tail down-force (TDF), the drag force (D) and the thrust force (T). The magnitude and disposition of these forces will determine whether the aircraft climbs, descends or maintains level flight, and at what IAS it does so.

Of the five forces only W is unchanging (disregarding its decrease as fuel is consumed). The factors affecting the others are listed below:

(a) L is the dependent upon the angle of attack of the wings and upon the aircraft's IAS;

(b) TDF is dependent upon elevator displacement and upon IAS. In reality, TDF has approximately the same magnitude in climbing, descending and level flight paths, except when the control wheel is actually being moved from one position to another to change the aircraft's attitude;

(c) D is dependent upon angle of attack (because of induced drag) and upon IAS;

(d) T is directly related to engine power setting.

Additionally:

(e) angle of attack is dependent upon the relationship between the pitch attitude of the aircraft and its actual flight path. Examples are illustrated in Figure 154 in which both aircraft have the same attitude. Thus the pitch attitude alone is not a direct indication of angle of attack;



(f) IAS depends upon the magnitude of T and upon the actual flight path of the aircraft. As an example, a particular power setting will result in a higher IAS if the machine is made to descend than if it is made to climb (in the same manner as a car tends to accelerate running downhill and to slow down running uphill, assuming that the accelerator pedal position is unchanged).

Figure 155 shows the disposition of the five forces in a climb, in level flight and in a descent. In each case the IAS is constant, neither increasing nor decreasing. For simplicity, TDF has been combined with W and all five forces are shown as acting from one point, which represents the aircraft. The thick arrows show the actual flight path of the machine in each situation.



We shall now introduce the concept of 'required lift', shown as RL in Figure 155. RL is simply the lift required to maintain the aircraft in the particular flight path at constant IAS.

Some interesting observations can be made:

(a) in level flight, (W + TDF) is balanced by RL, and D by T. In magnitude, RL is equal to (W + TDF) and T is equal to D;

(b) in the climb, (W + TDF) is balanced by RL together with part of T. D is balanced by the remainder of T. It is clear that in magnitude, RL is less than (W + TDF), and T is greater than D;

(c) in the descent, (W + TDF) is balanced by RL together with part of D. The remainder of D is balanced by T. It can be seen that in magnitude, RL is less than (W + TDF), and T is less than D.

The correlation between T and D is as expected - in climbing flight gravity (in the form of W) is opposing the motion of the aircraft, whereas in descending flight it is assisting.

Figure 155 also shows that although RL is less than (W + TDF) during climbing and descending flight, it is not very much so, especially in the relatively gentle climbs and descents associated with light aircraft. In other words, we can take RL as being reasonably constant regardless of actual flight path.

Bearing this in mind, we can deduce that since the aircraft in the bottom image in Figure 154 is flying with its wings at greater angle of attack than the other, it must therefore be at a lower IAS, because RL will be approximately the same in both cases. We can extend this deduction to one of considerable practical importance, namely that IAS is the only reliable clue the pilot has as to the angle of attack of his aircraft's wings, low IAS indicating high angle of attack and high IAS indicating low angle of attack, regardless of whether the aircraft is climbing, descending or in level flight. Remember that the relationship does not apply when the flight path is actually changing, only when the aircraft has stabilised in a steady flight path.

If, as a result of the pilot resetting his various controls, the actual lift generated by the wings is greater or less than that required, the disparity between the two will cause the aircraft to change its flight path in such a way as to establish a disposition of all five forces in which the lift generated is again equal to that required.

Of course, the pilot has no way of assessing the disposition or magnitude of the various forces. So far as he is concerned, control of the vertical flight path is achieved by adjustments of the engine power setting and the pitch attitude of the aircraft. In effect, once the pilot has chosen a power setting and a pitch attitude the aircraft will stabilise at a particular IAS in a flight path which disposes the five forces in equilibrium. Whether the machine climbs, descends or flies level will depend upon the actual settings of power and attitude chosen, as will the IAS at which it stabilises. To make the aircraft take up any required flight path and IAS, it is merely necessary for the pilot to adopt the power and attitude settings that he knows will achieve the desired results.

The number of variations of power and attitude settings is infinite, within the bounds of the maximum and minimum power available from the engine and the limit imposed by the wings attaining the stalling angle of attack. It is evident that the pilot must avoid any combination of power and attitude that will cause the machine to exceed Vne or the engine to exceed the maximum permitted RPM.

4.20 STRAIGHT AND LEVEL FLIGHT

In straight and level flight the aircraft maintains a constant heading (straight flight), a constant height (level flight) and a constant IAS.

4.20.1 Level flight

To achieve level flight, the aircraft's wings must generate enough lift to balance exactly the combination of weight and tail down-force.

This required lift can be generated at low IAS by arranging a high angle of attack for the wings, or at high IAS by arranging a low angle of attack. Of course, there are an infinite number of combinations in between that will give the same result. Figure 156 illustrates.



4.20.2 Control of flight path with attitude

So far as the vertical flight path is concerned, the aircraft's behaviour is indicated to the pilot by the altimeter and VSI. In level flight the altimeter shows constant height and the VSI no climb or descent. Climbing flight is indicated directly by the VSI and as an increasing reading on the altimeter. Descending flight is indicated directly by the VSI and as a decreasing reading on the altimeter.

We have already seen that, after the pilot has chosen any particular power and altitude settings, the aircraft will take up a flight path which disposes the five forces in equilibrium. Because of its inertia, the aircraft tends to resist changes in its flight path. Any change in attitude that the pilot might make will therefore have the initial effect of altering the angle of attack of the wings. In turn, the lift generated will change in magnitude, so that it now differs from that required to maintain the original flight path. The resulting imbalance will modify the flight path in such a way as to match lift generated to that required to maintain the new flight path, so restoring equilibrium.

As an example, consider that the aircraft is in a descending flight path, with its engine set to cruising power. (The IAS will be relatively high, because of the effect of gravity assisting the thrust from the propeller.)

Suppose that the pilot pitches the nose up to a higher attitude. Initially, the angle of attack will be increased, and the generated lift will exceed that required. As a result the aircraft will adopt a shallower descent path. The new flight path will decrease the angle of attack again, lessening the lift generated to that required.

In the shallower descent path there is less assistance from gravity in opposing drag, with the consequence that the aircraft decelerates to a lower IAS. (Assuming that required lift always has approximately the same magnitude regardless of flight path, we can deduce that once the aircraft has stabilised in the shallower descent its angle of attack  is greater than in the steeper descent, because of the lower IAS.)

Adopting a higher attitude again, the pilot might find that his aircraft stabilises in a climb, in which situation gravity is opposing the thrust from the propeller, resulting in an even lower IAS (and hence higher angle of attack).

By trial-and-error, the pilot will be able to find a particular attitude somewhere between these latter two at which the aircraft maintains constant height. For future reference, he will know that this particular combination of power and attitude will always result in level flight.

4.20.3 Control of speed with power

In level flight an aircraft's speed is controlled by the power setting. In the example above, the combination of cruising power with the attitude for level flight will result in the machine stabilising at a particular IAS.

To fly level at any other speed will require an adjustment in the power setting. However, if the previous attitude is maintained, any change in power will also result in the aircraft taking up a new flight path in order to balance the forces again. In other words, this attitude will not result in level flight if the power setting has been altered.

Suppose, for example, that it is desired to fly at higher IAS. The pilot's first action will be to increase power. As the aircraft accelerates it will take up a climbing flight path in which the various forces are disposed in equilibrium at the higher IAS. To re-attain level flight, all that is needed is for the pilot to select a lower attitude. If the new attitude is chosen correctly, level flight will be achieved. (If the chosen attitude is not low enough, the aircraft will continue to climb, but in a shallower flight path. If too low, the machine will stabilise in a descent.)

The transition from climb to level flight will result in a further increase in IAS. Of course, the wings will now be at a lower angle of attack than they were during level flight before the speed was increased.

Conversely, level flight at lower IAS will entail the choice of lower power and higher attitude. Figure 157 shows what might be typical attitudes for level flight at various speeds.



In practice, the pilot will find that the attitude change corresponding to a particular speed alteration (say, 20 knots) is greater in the region of low IAS than in the region of higher IAS. This phenomenon is a result of aerodynamic factors and is illustrated in Figure 157.

With experience, the pilot will be able to fly level at any chosen IAS (within practical limits) by correct setting of power and attitude.

Note that the longitudinal trim changes resulting from alteration of power setting will be in opposition to the attitude adjustments needed to maintain level flight.

For example, increasing power will tend to cause pitch up at a time when a lower attitude is required to maintain level flight at the resulting higher IAS. The pilot must use the control wheel to hold the chosen attitude rigidly, against the opposing pitching effect. Of course, the trim wheel should then be used to eradicate the control wheel force, so that the aircraft is re-trimmed in its new attitude.

4.20.4 Straight flight

Any tendency of the aircraft to change its heading can only be caused by the yaw arising from banked flight or from the misuse of rudder, or by propwash induced yaw.

If, therefore, the pilot ensures that the wings are level (using the ailerons if necessary to make certain that this is so) and that the flight is balanced (using rudder as appropriate to maintain a central disposition of the ball in the balance indicator) then there can be no yawing influence acting on the aircraft, and so it will maintain a constant heading.

4.20.5 Balance

Suppose that the pilot has correctly eliminated propwash-induced yaw, but that he fails to keep the wings level. Perhaps he has inadvertently selected a slight bank to the right. The aircraft will start to turn to the right, of course. The pilot, instead of levelling the wings as he should, might prevent the turning effect by applying left rudder. (The turning motion is arrested because the action of the fin in nullifying the bank-induced sideslip is opposed by the effect of the airflow on the displaced rudder, as shown in Figure 158.)



It will be agreed that straight flight has indeed been achieved as the heading is not changing, even though it differs from the direction of flight. Since the aircraft is not turning, the ball in the balance indicator will take up its lowest position. Because of the bank, however, the disposition will not be central (Figure 159).



Evidently the flight is unbalanced. Note that, as a result of the continuing sideslip, the aircraft's lateral stability will attempt to roll it to the left. In other words, the bank can only be maintained by holding the ailerons in their displaced positions with the control wheel. The straight flight has been achieved by the ailerons and rudder acting in opposition; this situation is referred to as flight with 'crossed controls'.

We know that to rectify the unbalance the right rudder pedal must be pushed forwards, in accordance with the disposition of the ball in the balance indicator. If this is done, thereby restoring the rudder to its original position, and the wings are then levelled (using the ailerons), the aircraft will again take up a straight flight path. Now the flight will be balanced. (Not surprisingly, for a particular power setting and pitch attitude, balanced straight flight will result in a higher IAS than straight flight with crossed controls, since drag is minimised.)

Although marked unbalance is obvious to the pilot (because of the bank attitude), slight unbalance may not be so. The correct technique, therefore, is to use the rudder as dictated by the balance indicator, and to concentrate on holding the wings level by appropriate use of the ailerons.

4.20.6 Drag

It will be recalled that the total drag generated by an aircraft is made up of two contributions, namely turbulent wake drag (generated by the entire machine) and induced drag (generated by the wings).

We know already that the former increases with the square of the IAS. For example, doubling the IAS results in turbulent wake drag increasing to four times its magnitude.

At any particular angle of attack, induced drag follows the same relationship. Of course, induced drag is also affected by variation of angle of attack.

At low IAS, the high angle of attack needed to generate the required lift causes high induced drag; the opposite is true at high IAS. Obviously, the two factors (IAS and angle of attack) are in opposition so far as the magnitude of induced drag in concerned. To determine the overall effect, consider the 'effective lift force' generated in each of the extreme situations shown in Figure 156. (Remember that the effective lift force acts at right angles to the direction of the airflow in the downwash.) Figure 160 shows these forces resolved into lift and induced drag.



We can deduce that flight at low IAS incurs greater induced drag than at high IAS - the angle of attack factor predominates over the IAS factor.

In Figure 161 is shown the variation of turbulent wake drag, induced drag and total drag (the two contributions added together) with IAS.



The diagram shows that there is one IAS at which the total drag has its lowest magnitude. It is annotated Vmd (the IAS for minimum drag).

In level flight there will be only one angle of attack that will generate the required lift at Vmd. It is thus the angle of attack at which the ratio of lift to total drag, in other words the aerodynamic efficiency of the aircraft, is highest.

4.20.7 Power

It can be shown that the power required to propel the aircraft in level flight is the product of its total drag and its TAS. As an equation:

Power = Drag x TAS

We shall consider the case at sea level in the ISA (where TAS equals IAS), in which:

Power = Drag x IAS

Reference to Figure 161 shows that selection of an IAS slightly less than Vmd results in only a small increase in drag. In fact the increase in drag is less than the reduction in IAS, in relative terms. In this IAS range the product of drag and IAS is therefore less than at Vmd. However, if IAS is reduced further, drag increases at a greater rate than IAS decreases, and so the power required from the engine is increased.

We can deduce that there is an IAS at which minimum power is required (Vmp) and that this speed is less than Vmd. Figure 162 shows the variation in power required with IAS.



In level flight there will be only one angle of attack which will generate the required lift at Vmp. Of course, at this angle of attack the aircraft is not flying at maximum aerodynamic efficiency.

Note that flight at constant IAS below Vmp requires increased power. In practice, the aircraft is not usually flown in this speed range. The lowest IAS that can be maintained in level flight occurs when the wings are just below the stalling angle of attack; flight at this speed involves the use of a considerable amount of power.

The highest IAS that can be maintained in level flight occurs when the engine is delivering full power, with the aircraft in the appropriate pitch attitude.

4.20.8 Flying for maximum range

For economy of operation, it is highly desirable to be able to fly the aircraft at the speed at which it travels the maximum distance, or range, through the air for each litre of fuel consumed.

We can determine this maximum range speed from the equation introduced in 4.20.7. To recap:

Power required = Drag x TAS

Each litre of fuel enables the engine to deliver a certain amount of power for a certain length of time. The two quantities are related; for example, if the power output is doubled, it can be sustained for only half the length of time. In other words, the product of power and time is a fixed quantity, representing the work obtainable from the litre of fuel for propelling the aircraft. We can re-write the equation as:

Power x Time = Drag x TAS x Time

or

Work obtainable = Drag x TAS x Time

The product of TAS and time is equal to the actual distance travelled through the air. (As an example, an aircraft flying at 2 nautical miles per minute for 5 minutes travels 10 nautical miles.)

Our equation becomes:

Work obtainable = Drag x Distance

or

Distance = Work obtainable
                            Drag

Now, since it is known that the work obtainable is a fixed quantity, we can deduce that in order to maximise the distance flown, the aircraft's drag must be minimised.

We can conclude that the maximum range speed is identical with the minimum drag speed, Vmd.
 
In practical terms, for a flight from one particular location to another, total fuel consumed can be minimised by cruising at Vmd.

The IAS for maximum range for any aircraft will be specified in its Flight Manual.

4.20.8.1 Effect of height on range

Since Vmd is concerned with the aircraft's aerodynamic characteristics and is expressed as an IAS, it is not affected by height. At the same time, the TAS resulting from flight at this IAS increases with height. It would seem, therefore, that the range obtainable from each litre of fuel can be enhanced by cruising at the greatest height achievable by the aircraft.

However, the equation in 4.20.7 informs us that the power required to maintain any particular IAS increases with height, because TAS increases. To fly at Vmd at greater height therefore requires more power from the engine. If this is so, of course, our litre of fuel will be consumed more quickly. So the increase in TAS is offset by the reduced time for which it can be maintained.

In practice, the two effects tend to cancel out each other, and the range obtainable is not affected greatly by height. Nevertheless, an aircraft's Flight Manual may specify an altitude (height above sea level) at which range is maximised.

4.20.9 Flying for maximum endurance

Operationally, it may occasionally be necessary to use fuel at the lowest possible rate of consumption. Flight at the maximum endurance speed will minimise the rate of consumption and so maximise the length of time for which the aircraft can fly for each litre of fuel consumed.

The derivation of the maximum endurance speed is straightforward if we consider the litre of fuel as representing a fixed quantity of work obtainable. In 4.20.8 we used the relationship:

Work obtainable = Power x Time

or

Time = Work obtainable
                    Power

In order to maximise time, it is evident that the power output from the engine must be the lowest that will enable the aircraft to maintain level flight. In other words, the pilot must fly at the minimum power speed, Vmp. In practice, flight at this speed is impractical, since there is no reserve for manoeuvring. For this reason, it is usual for a factor of, perhaps, 5 knots or so to be added, giving the recommended maximum endurance speed that will be specified as an IAS in the Flight Manual.

4.20.9.1 Effect of height on endurance

In 4.20.7 we considered the case of power required with varying IAS at sea level in the ISA. If we were to examine the relationship at greater heights, the results would be as shown in Figure 163.



The diagram shows that the TAS at which minimum power may be used is increased at greater heights. In fact, at any chosen height, the correct TAS for minimum power is achieved by flying at Vmp, which is specified as an IAS.

However, since:

Power required = Drag x TAS

it is not surprising that Figure 163 shows an increase in minimum power required in greater heights. Since the work obtainable is a fixed quantity, therefore, the length of time for which flight may be maintained is lessened. In other words, endurance is reduced at greater heights.

The conclusion is that, to maximise endurance, the pilot should fly his aircraft at the recommended maximum endurance speed at as low a height as is compatible with adequate terrain clearance.

The reader will appreciate that the recommended maximum endurance speed is always less than the maximum range speed for any particular aircraft (since Vmp is less than Vmd).
 
4.20.10 Technique for straight and level flight

A summary follows in practical terms of the various observations concerning straight and level flight:

Level flight at any chosen IAS is achieved by setting the correct power and attitude.

IAS can be varied by appropriate resetting of power; attitude is adjusted as required to maintain level flight. When altering IAS, it should be borne in mind that the aircraft takes time to accelerate or decelerate, because of its inertia. The skilful pilot will be able to make his attitude changes at such a rate that, at any instant during the change in speed, the attitude matches the IAS, so that level flight is maintained throughout.

When the correct attitude for level flight is established the aircraft should be trimmed to remove the control wheel force.

Straight flight is achieved by holding the wings level and ensuring that the aircraft is in balance.

Greatest range is achieved by flight at the maximum range speed (and optimum altitude) specified in the Flight Manual.

Greatest endurance is achieved by flight at the recommended maximum endurance speed similarly specified, at as low a height as is practicable.

If time-saving is more important than maximising range or endurance, TAS can be increased by cruising at higher IAS or at greater height, or both.

4.20.11 Correction of deviations

Because of disturbances in the air or inadvertent movement of the controls by the pilot, it may well be that the aircraft deviates from the desired heading and height.

Heading deviation may be corrected by gently banking the wings in the appropriate direction so that the machine turns towards the chosen directional reference point (or the chosen heading indication on the DI). When the correct heading is re-attained, the wings should be levelled again to resume straight flight.

Excess height may be corrected by temporarily selecting a slightly lower attitude. (The ensuing increase in IAS may be prevented if desired by a small reduction in power.) As the aircraft descends to the chosen height, the attitude should be restored to its original setting (and the power reset as appropriate if it was reduced) to regain level flight.

Conversely, insufficient height may be corrected by temporarily setting a slightly higher attitude, preventing the resulting decrease in IAS with a small increase in power. As the aircraft climbs to the chosen height, the attitude and power should be readjusted to their original settings to regain level flight.

4.21 CLIMBING

In straight climbing flight the aircraft is made to climb at a constant IAS on a constant heading.

It has already been seen that certain combinations of power and attitude will result in climbing flight. In practice, specific combinations are chosen to optimise the aircraft's performance.

4.21.1 Climbing at maximum rate

When climbing, the pilot will often want to gain height as quickly as possible. In other words, he will want to climb at maximum rate.

Figure 155 showed that, in climbing flight, thrust is greater than drag. Alternatively, we can consider that the engine power output converted by the propeller into thrust is greater than that required for level flight at the same speed, with the excess of power responsible for the gain in height.

The power converted into thrust is that actually delivered by the engine reduced somewhat by propeller inefficiency. For a fixed-pitch propeller, efficiency varies with the aircraft's speed, as was explained in 3.10.4.1, and is usually highest in the cruising speed range, when about 80% of the engine's propulsive force is converted by the propeller into thrust.

The power available is that converted into thrust when the engine is delivering full power. Figure 164 illustrates the variation of power available with IAS at sea level. Also shown is the variation of power required to maintain level flight at sea level.



Some interesting observations can be made:

(a) IAS (1) is the lowest speed which can be maintained in level flight, with the wings just below the stalling angle. (The power required to maintain this speed is less than that available);

(b) at IAS (2) the excess of power available over that required is greatest. Thus flight at this speed with the engine at full power will result in the aircraft achieving its maximum rate of climb;

(c) at IAS (3) the propeller efficiency is greatest;

(d) IAS (4) is the maximum speed which the aircraft can maintain in level flight, with power available equal to that required. The machine cannot climb at all at this speed, since there is no excess of power.

As already stated, Figure 164 relates power available and power required to IAS at sea level. Obviously, in climbing flight the aircraft is increasing its height, and these relationships will be changed as height is gained. As a general rule of thumb for light aircraft, the IAS which gives the maximum rate of climb at sea level should be reduced by 5 knots for every 5000 feet of height gained, if maximum rate of climb is to be maintained.

The power required to maintain the appropriate IAS increases with height, since TAS increases at a greater rate than the IAS is decreased. At the same time, the power available is reduced, because the engine's maximum power output decreases, as was explained in 2.2.6. Hence the excess of power, and thus the maximum rate of climb, are reduced, until eventually a height is reached at which, even with the engine at full power, the aircraft can climb no higher. In other words, it has reached its ceiling.

Another factor must be considered, however, before deciding the best speed at which to climb, namely engine cooling. When the engine is set to full power it is difficult for the cooling fins to get rid of the heat from the cylinders, especially at the relatively low IAS at which the maximum excess of power occurs. (Remember that at low speed the airflow through the fins is not strong.) With some aircraft designs, it is necessary to prevent engine overheating by either increasing the IAS, or by reducing power slightly, or both. Of course, either of these actions will reduce the excess of power and hence the maximum rate of climb.

An aircraft's Flight Manual will therefore specify the recommended IAS and power setting which should be adopted for maximum rate of climb. For many smaller aircraft, it is permitted to climb with the engine at full power so long as the oil temperature limitation is not exceeded. To maintain maximum rate of climb as height is gained, the rule of thumb given above should be applied to the recommended IAS.

Having set the specified climbing power, the only means by which the IAS can be controlled is by attitude. In practice, an attitude is selected which the pilot believes will result in the correct IAS. If the aircraft stabilises at too high a speed, a higher attitude should be tried, and vice versa. (Because of the aircraft's inertia, several seconds elapse before the machine settles at a new speed after change of attitude.) After some experience has been gained this trial-and-error method will not be needed - the pilot will come to know the attitude setting that will achieve the correct IAS. The rate of climb will be indicated by the VSI.

Propwash-induced yaw will be more marked in the climb, because of the use of high power. Application of rudder will be needed to keep the balance indicator ball centred. If this is done, and the wings are held level, then the aircraft will climb on a constant heading. Figure 165 shows what might be a typical climbing attitude.



Note that the view ahead of the nose is restricted by the engine cowling. To verify constancy of heading a directional reference point can be chosen to the side of the cowling, as shown. Alternatively, consultation of the DI will serve the same purpose.

The restricted view ahead also denies to the pilot the ability to watch for the presence of other traffic. This problem can be overcome by banking the aircraft gently to one side (for example, to the right) for a few seconds to change heading, so that the area in question is brought into view to the side of the cowling. Having satisfied himself that there is no conflicting traffic, the pilot can then bank the machine back to the original heading. This procedure should be repeated from time to time.

Periodically, during a prolonged climb, the oil temperature should be checked. If the gauge reading approaches the maximum permitted, power should be reduced to prevent overheating. (In the climbing attitude, reduction of power will cause the aircraft to decelerate - the correct IAS can be maintained by setting a lower attitude.) Alternatively, or additionally, a lower attitude again can be set to give a higher IAS and so assist with cooling. Either of these actions will detract from the maximum achievable rate of climb.

Notice that the aircraft's controls are being used in a different manner compared to level flight. In the climb, the power setting is fixed and IAS is controlled by attitude.

The transition from climb to level flight when the aircraft has reached the desired height is termed 'levelling off'. The normal technique is firstly to lower the nose gradually to the cruising attitude. The aircraft will then accelerate. As the desired cruising speed is attained the power should be reduced to the correct setting so that this speed is maintained in level flight. If the first attempt does not result in exactly the required behaviour, further adjustments of power and attitude are made until the aircraft is flying level at the chosen speed. Since the flight path cannot be changed immediately, because of the machine's inertia, the levelling off process should be initiated just before the chosen height is reached. If the anticipation is judged correctly, level flight will be achieved at exactly this height.

4.21.2 Climbing at maximum gradient

The gradient of climb is the ratio of height gained to horizontal distance travelled in a given time. It can be considered as the angle at which the flight path is inclined (Figure 166).



Suppose that the aircraft is climbing at maximum rate, and that the pilot then raises the nose to a slightly higher attitude. The effect will be to reduce the IAS. In consequence, the horizontal distance travelled through the air in a given time will be reduced. Although at this lower IAS the power excess will be less, it will not be very much so, as reference to Figure 164 confirms. In other words, the reduction in rate of gain of height is less than the reduction in horizontal distance travelled, in relative terms, and so the gradient of climb will be steeper.

Of course, if the speed is reduced too much the rate of climb will lessen considerably (because of the smaller power excess) and so the gradient of climb will be reduced. We can conclude that there is one IAS at which maximum gradient of climb is achieved, and that this speed will be slightly lower than that for maximum rate of climb (and correspondingly, the attitude slightly higher).

Figure 167 illustrates the two cases, representing 1 minute's flight.



In practice, the aircraft is made to climb at maximum gradient when obstacles obstruct the flight path after take-off. For this reason, the technique is sometimes referred to as 'obstacle clearance climb'. The reader will appreciate that prolonged climb at maximum gradient IAS may cause engine overheating problems, and so it is normal practice to revert to maximum rate or cruise climb (described later) as soon as the obstacles have been safely overflown.

The aircraft's Flight Manual will specify the IAS at which maximum gradient of climb is achieved.

4.21.3 Summary of techniques for climbing

From level flight, climb at maximum rate is achieved by firstly setting climbing power (maximum if permitted) and then adopting the pitch attitude which results in the IAS recommended in the Flight Manual. The aircraft should be trimmed in this attitude.

Constancy of heading is achieved by balancing the flight with appropriate application of rudder and by holding the wings level. Banking the aircraft occasionally to change heading temporarily will enable the pilot to see any other traffic in the intended flight path.

During a prolonged climb close watch must be kept on the oil temperature gauge. Appropriate action should be taken if necessary to prevent the maximum permitted temperature being exceeded.

To level off, the nose is lowered to the cruising attitude and the power reduced to the corresponding setting once the IAS has increased to that desired. When level flight has been attained the aircraft should be re-trimmed.

Maximum gradient of climb is achieved by setting climbing power and adopting the attitude which results in the IAS specified in the Flight Manual. As soon as possible the pilot should revert to maximum rate climb or cruise climb.

4.21.4 Cruise climb

If it is necessary to increase height, and rate or gradient need not be maximised, the usual technique is the cruise climb. In the cruise climb the power setting is somewhat higher than during level cruising but less than maximum, and the attitude is set for an IAS between that for maximum rate climb and that for level cruising, according to the pilot's preference. Besides conferring the advantage of higher speed, the lower attitude also offers a better view ahead of the aircraft. The levelling off technique is as already described.

4.22 DESCENDING

In straight descending flight the aircraft is made to descend at a constant IAS on a constant heading.

As already stated, certain combinations of power and attitude will result in descending flight. In practice, specific combinations are used, depending on the required descent path. Descending with the engine at minimum power is referred to as 'gliding'.

4.22.1 Gliding at minimum gradient of descent

The gradient of descent is the ratio of height lost to horizontal distance travelled in a given time. It can be considered as the angle at which the flight path is depressed (Figure 168).



Occasionally, it may be necessary for the pilot to glide his aircraft at minimum gradient of descent.

Figure 155 showed that, in descending flight, thrust is less than drag. With the engine at minimum power, thrust is reduced almost to zero. In this situation gravity alone (in other words, the weight force) enables the aircraft to maintain constant IAS, with the forces disposed as in Figure 169. (The dotted arrow shows the component of the weight force which opposes drag.)



When the aircraft is gliding its gradient of descent is entirely dependent upon the ratio of required lift to total drag, as reference to Figure 170 will confirm.



Notice that, the greater the ratio of lift to total drag, the shallower is the gradient and so the greater is the horizontal distance travelled for any particular height reduction. Since the required lift is approximately constant regardless of the descent path, we can conclude that the minimum gradient occurs when the aircraft is gliding with its wings at the angle of attack at which minimum total drag occurs. The aircraft's Flight Manual will quote the gliding IAS which corresponds to this angle of attack. (In practical terms this speed will equal the speed for maximum range, which also implies flying at the speed which maximises the ratio of lift to total drag.)

With the engine at minimum power, the only means by which the IAS can be controlled is by attitude. Too high an attitude will result in too low an IAS, and vice versa. With experience, the pilot will know exactly which attitude achieves the correct IAS.

Since the IAS for minimum gradient gliding is in the lower speed range, the pilot will find that the corresponding attitude is not markedly low. In fact, with many aircraft designs, the correct attitude is not much lower than that for cruising in level flight at higher IAS. With this being the case, it will be appreciated that the view of the area into which the aircraft is descending is obscured by the engine cowling. The problem can be overcome by temporary changes of heading during the descent, so that the area in question is visible to the side of the cowling.

Note that gliding at any other than the specified IAS, either faster or slower, will result in a steeper descent. In the latter case, we have a clear example of a situation in which the vertical flight path of the aircraft is not in the same direction as the nose is pointing, because a relatively high angle of attack is involved. Figure 171 illustrates.


 
Propwash-induced yaw will be virtually non-existent during the glide. If the rudder trim tab has been set for cruising flight, therefore, the aircraft will tend to fly out of balance. Application of rudder in accordance with the balance indicator will rectify the situation. (Of course, the application will be the opposite to that needed during the climb, when the engine is at high power.)

Periodically during a prolonged glide, the engine should be opened up to cruise power for a few seconds. This action is called 'clearing the engine' and serves two purposes. Firstly, it ensures that any oil film which has formed on the spark plug electrodes is burnt away. Secondly, it prevents the engine and oil from cooling too far below the normal operating temperature.

Levelling off from descent to level flight is initiated as the aircraft approaches the desired height. From the glide, the usual technique is firstly to set cruising power, whilst maintaining the descent attitude. The aircraft will then accelerate. As the desired cruising speed is reached the nose is then raised to the corresponding attitude, so that level flight is established. If the first attempt does not result in exactly the desired behaviour, further adjustments of power and attitude are made to achieve level flight at the chosen speed.

The levelling off process should be initiated just before the chosen height is reached, so that level flight is achieved at exactly this height.
 
4.22.2 Sideslipping

Occasionally, the pilot may wish to glide more steeply than the minimum gradient. There are several methods of doing so available to him.

Firstly, by a suitable adjustment of attitude, he can fly at a faster IAS than that for minimum gradient. This technique will not be suitable if approaching to land, however, since the pilot will not want excessive speed at this stage.

Alternatively, he could fly at a slower IAS than that for minimum gradient, but this procedure reduces the effectiveness of the flying controls and involves the wings meeting the airflow at an angle of attack close to the stalling angle, and so is usually avoided.

Since the gradient of descent depends on the ratio of lift to total aircraft drag, it can be appreciated that, in order to steepen the descent path without departing from the normal (minimum gradient) gliding speed, extra drag must be created. If flaps are fitted to the aircraft, lowering them will have this effect and will therefore achieve the desired result. (We shall look at use of the flaps in more detail later.)

A few designs of aircraft do not incorporate flaps. If the pilot wishes to steepen the glide without increase or decrease of speed, he can resort to sideslipping. The technique involves deliberate unbalanced flight using crossed controls, to generate extra drag.

For example, suppose that left bank is applied, using the ailerons, and the consequent yaw prevented by application of right rudder, so that straight flight (constant heading) is maintained. Appropriate lowering of pitch attitude to prevent the extra drag from slowing the aircraft will enable the normal gliding speed to be maintained.

Note that, when sideslipping, the aircraft's flight path, although straight, is not in the same direction as its heading. Figure 172 illustrates the example above - sideslipping to the left.



The amount of bank that can be set (and hence the steepness of descent) is limited by the effectiveness of the rudder. When maximum deflection of the latter is required to maintain straight flight, any further application of bank will cause vaw in the direction of the bank.

When desired, balanced flight can be restored by centralising the rudder and levelling the wings. With the extra drag removed, a slightly higher pitch attitude will then be needed to prevent acceleration above normal gliding speed.

Sideslipping to the right involves left rudder opposing right bank.

4.22.3 Powered descent

In the glide at minimum gradient IAS the aircraft will stabilise at a steady rate of descent, indicated by the VSI. If it is desired to descend at a shallower gradient at this IAS, the rate of descent must be decreased. Of course, the only way to achieve this is with the assistance of engine power. If the glide attitude is maintained, any increase of power will tend to accelerate the aircraft. To maintain the original speed, therefore, a higher attitude will be required. The new combination of power and attitude will result in a lower rate of descent, and thus a shallower descent gradient. The aircraft is now in a powered descent.

Further increases of power (and corresponding raising of pitch attitude) will give even lower rates of descent and shallower gradients. (The reader will agree that if the process is repeated sufficiently, the aircraft will finally achieve a climbing flight path.)

Notice the manner in which the controls are being used, with power controlling the rate of descent and attitude controlling the IAS.

4.22.4 Summary of techniques for descending

From level flight at cruising speed, gliding at minimum gradient is achieved by firstly reducing power to minimum and maintaining attitude until the IAS has decayed to that specified, and then lowering the attitude as necessary to maintain this speed. The aircraft should be trimmed in this attitude.

Constancy of heading is achieved by balancing the flight and holding the wings level. Banking the aircraft occasionally to change heading temporarily will enable the pilot to see any other traffic in the intended flight path.

During a prolonged glide the engine should be cleared periodically.

To level off, the power is increased to the appropriate cruise setting and the nose is raised to the corresponding attitude when the IAS has increased to that desired. When level flight has been attained, the aircraft should be re-trimmed.

If required, the steepness of the descent path may be increased by either gliding at higher IAS or, if increase of speed is undesired, by lowering the flaps or by sideslipping.

In a powered descent adjustments of engine power are used to control the rate of descent and attitude is set as necessary to maintain the correct IAS.

4.22.5 Cruise descent

If during cruising flight it is desired to descend to a lower height without change of speed, the usual technique is to reduce power and adopt a lower attitude. The control technique is the same as during the powered descent described above, with power set according to the required rate of descent and attitude adjusted as necessary to maintain cruising speed. To level off, the pilot should reset cruising power and corresponding attitude.

4.23 FLAPS

The flaps are used to alter the lift- and drag-generating abilities of the wings. On many light aircraft, the flaps can be set either up or down, or to an intermediate, 'maximum lift' position.

Refer back to Figure 5. Let us consider the situation in which the wing is moving through the air at the same speed and angle of attack in each of the three flap configurations. (Note that the 'flaps up' chord line is the reference for assessing the angle of attack at other flap positions.) The overall camber of the wing is increased when the flaps are lowered, as the diagram shows.

Compared to the 'flaps up' case, the wing generates more lift in the 'maximum lift' flap configuration, because of the increased camber, and slightly more drag. (Nevertheless, the ratio of total lift to total wing drag, in other words wing efficiency, is decreased.)

In the 'flaps down' configuration, lift is further increased slightly, and drag considerably. (This being so, the reader will appreciate that the term 'maximum lift' at the intermediate flap position, although common parlance, in incorrect. The misnomer arises from the manner in which the flaps are used.)

The flaps serve two main functions. Firstly, in the 'maximum lift' position, they reduce the length of runway needed for take-off. The explanation is straightforward: at any particular angle of attack, the required lift can be generated with the flaps in this position at a lower speed than if they were up. If the aircraft's wings are at the take-off angle of attack, then, the machine will leave the runway at a lower speed, thereby using less runway distance during the take-off acceleration. After the aircraft has become airborne, the lowered flaps are a liability, because of the reduced wing efficiency. They are therefore raised to the 'up' position to get rid of the extra drag, so that the aircraft's subsequent climb performance is not impaired.

The second function of the flaps is to increase drag, so that steep descents may be made without the need to increase speed. Of course, for this purpose they are lowered to the 'down' position. (At any particular descent speed, the wings will be at a lower angle of attack than with the flaps up, since the required lift is the same in both cases. The lower angle of attack eradicates the increase of lift that the flaps would otherwise confer.)

4.23.1 Effect of flaps on stalling speed

When the aircraft is in flight, the required lift can be generated at high speed with the wings at low angle of attack, or vice versa. In the latter case the limit is reached when the angle of attack has increased to the stalling angle. The IAS at which this situation occurs is called the stalling speed.

The aircraft stalls when it slows to the stalling speed. To the pilot the main symptoms are partial or complete loss of roll and pitch control and loss of height, all these consequences arising from the disruption of the airflow over the wings. Thus flight near the stalling speed is generally avoided.

Consider the aircraft just at its stalling speed with the flaps up. Of course, the required lift can be generated at this same speed at a lower angle of attack if the flaps are at the 'maximum lift' position. If now the angle of attack is increased slightly, the required lift can be generated at a slightly lower IAS, without the occurrence of stalling. A further increase in angle of attack will, however, result in stalling, which will occur at a lower angle of attack than with the flaps up. Nevertheless, the stalling speed will be less than before.

A further small reduction in stalling speed results when the flaps are fully down.

So far as practical flight is concerned, the pilot may fly safely at slightly lower speeds with the flaps lowered than he could with them up, without erosion of the margin above the stalling speed.

4.23.2 Effect of flaps on trim

Figure 173 shows the flaps lowered to the 'maximum lift' position on both a low-winged and a high-winged aircraft. It will be assumed that the machines were trimmed in the attitudes illustrated before the flaps were lowered.



As already stated, the wings generate more drag in this configuration. In the case of the low-winged aircraft, reference to Figure 173 suggests that the extra drag will tend to make the nose pitch down. In the other, a pitch up is more likely.

The lowered flaps also have the effect of deflecting the downwash further downwards, which in both cases will increase the tail down-force, thus tending to cause pitch up.

We can conclude that lowering the flaps on a high-winged aircraft will cause the nose to pitch up, since the two effects just described are acting in conjunction. On a low-winged machine, they are acting in opposition. Whether pitch up or pitch down occurs depends on the relative magnitudes of the two effects. (On some designs they cancel out each other exactly.)

The pitching effect of lowered flaps is evidently a trim change, because the aircraft will tend to depart from the original trimmed attitude. It can be prevented by appropriate use of the control wheel, and the control wheel force then eliminated with the trim wheel. Moving the flaps fully down will incur a further trim change.

Once the aircraft has been trimmed with flaps down, raising them will cause a trim change opposite to that when they were lowered. Again, the effect can be prevented using the control wheel, and the machine then re-trimmed.

4.23.3 Control of speed and vertical flight path

If no change is made in power or attitude when the flaps are lowered, the extra drag will result in a reduced IAS. To achieve a desired vertical flight path and the IAS appropriate to it after lowering the flaps, then, adjustments of power and attitude may be needed. In fact, the aircraft's controls are used as already described. In descending flight, for example, an attitude is set which results in the appropriate IAS, and power to achieve the required rate of descent.

(Note that, depending on the design of the aircraft, the trim change arising from lowered flaps may either assist or oppose the attitude adjustment necessary to obtain the desired aircraft behaviour. The correct attitude should be held rigidly whilst the trim wheel is used to eradicate any control wheel force, so that the aircraft is re-trimmed in the chosen attitude.)

Similar power and attitude adjustments are made as necessary to achieve the desired vertical flight path and IAS when the flaps are raised, and the aircraft is again re-trimmed.

The flaps must never be lowered at speeds in excess of the flap limiting speed, otherwise they may be damaged by the airflow. Of course, once the flaps have been lowered, this speed must not be exceeded until they are raised again.

4.23.4 Effect of flaps on forward view from cabin

As we have already seen, the wings are at a lower angle of attack when the flaps are lowered than at the same IAS with them up. As a consequence, the aircraft's nose will be pointing more in the direction of its vertical flight path, especially in the lower speed (higher angle of attack) range in which climbing and descending flight usually take place. Figure 174 illustrates the effect in descending flight. Remember that both aircraft have the same IAS.



So far as the pilot is concerned, when flying at low speed he has a better view of the area into which his machine is flying when the flaps are in use, with his vision unobstructed by the engine cowling.
 
4.23.5 Use of flaps for take-off

As already stated, the aircraft's runway distance requirement can be reduced by taking off with the flaps at the 'maximum lift' position.

Once the machine is airborne, the flaps impair the climbing performance, because of the extra drag. However, if they are raised immediately after leaving the ground, the consequent loss of lift may result in the aircraft sinking back to the ground. Additionally, the pilot will possibly have to cope with a trim change during a critical phase of flight.

For these reasons, the flaps are not normally raised until the aircraft has attained a safe height. To minimise the performance loss before the flaps are raised, it is usual practice for the pilot to climb at a lower IAS than that for maximum rate of climb with flaps up, thereby gaining two advantages. Firstly, the gradient of climb is better than at the higher speed, because of the reduction in horizontal distance travelled. (However, the gradient will not be as steep as at maximum gradient IAS with the flaps up, because of the extra drag.) Secondly, since the aerofoil shape is more highly cambered, the reduction in wing efficiency (in other words, the magnitude of the extra drag) is less at the lower speed.

The aircraft's Flight Manual will usually specify the optimum climbing IAS to be used when the flaps are at the 'maximum lift' position.

Once a safe height has been reached the flaps can be raised (which may cause temporary sinking of the aircraft) and the attitude adjusted as necessary to achieve the recommended IAS for maximum rate of climb.

4.23.6 Use of flaps for approach and landing

The usual technique for flying the aircraft towards the runway for landing is the powered approach, in which the flaps are fully down and the speed and vertical flight path are controlled by variations in attitude and engine power setting. (We'll look at this procedure in more detail later.)

It seems illogical, at first sight, to have the power (which reduces the rate of descent at any particular approach speed) opposing the effect of lowered flaps (which increases it). In fact this technique confers three advantages over descent with flaps up.

Firstly, at the approach speed, the range of rates of descent is broadened. With minimum power set, the gradient is considerably steeper than with flaps up, if such a flight path is desired. With increasing power, the same IAS will give shallower descent if needed.

Secondly, the pilot will have a better view of the approach path.

Thirdly, the lowered stalling speed makes possible safe reduction of approach speed just prior to landing, so that, when the aircraft actually touches down, it will run a shorter distance along the runway while being braked to taxying speed.

The aircraft's Flight Manual will specify the IAS at which the powered approach should be flown. Usually, this speed is slightly less than the normal gliding speed.

It may be that the pilot has to break off his approach and climb away again. This procedure is called 'going around' (Figure 175).



With the flaps fully down, of course, the aircraft's climb performance will be considerably impaired. However, raising the flaps will cause the machine to sink, with possibly hazardous consequences if it is near the ground, and there may also be a trim change.

Normal technique, therefore, is firstly to establish a climbing flight path with the flaps still fully down. Full power should be set and an attitude chosen which gives an IAS at, or slightly below, that specified for climbing with the flaps at 'maximum lift', to optimise the gradient. As soon as a climb is established the flaps should then be raised to 'maximum lift' and the attitude adjusted as necessary to achieve the specified IAS. At a safe height the flaps can be fully raised and maximum rate climb established.

This technique of raising the flaps in two 'stages' (that is, to 'maximum lift' first and then fully up at a safe height) divides the attendant sink and trim change into more manageable quantities. Likewise, the resulting increase in stalling speed occurs as two smaller rather than one large increment.

4.24 MEDIUM TURNS

In medium turns, the aircraft is made to turn using angles of bank of 30 or less. Turns can be carried out in level flight, or when climbing or descending.

4.24.1 Level turns

A turn made in level flight is a level turn. In 4.17.3 it was explained that application of bank is a more efficient manner of turning than application of rudder, because the latter results in skidding, unbalanced flight.

Figure 117 showed that in a banked attitude the consequence of the interaction of the weight force and the tilted lift force is sideslip. The action of the fin in nullifying the sideslip then causes the turning effect. Referring again to Figure 117, it will be appreciated that, because the lift force is tilted, its vertical component will now be less in magnitude than it is with the wings level, as the top image in Figure 176 shows.



As a consequence, the aircraft will not be able to maintain level flight and will take up a descending flight path. To prevent this occurrence, it is necessary to Increase the lift force so that its vertical component equals the 'wings level' lift force, as shown in the bottom image in Figure 176.

Of course, the lift increase can be generated either by an increase in speed or by an increase in angle of attack. In practice, the latter method is chosen because it is more convenient for the pilot. Reference to Figure 143 shows that when the aircraft yaws as a consequence of bank, its pitch attitude becomes lower. The effect will be apparent to the pilot as a lowering of the engine cowling below the horizon. By moving the control wheel backwards to maintain the original pitch attitude, the pilot will be arranging that his wings meet the airflow at the higher angle of attack needed to generate the increased lift force.

Besides the lift increase, the wings will also generate more drag at this higher angle of attack. Consequently the aircraft will decelerate to a lower IAS. The resulting decrease in lift will need to be offset by a further small increase in angle of attack. So far as the pilot is concerned, it is merely necessary to adopt whichever pitch attitude achieves level flight.

In medium turns, the pitch attitude differs little from that in level straight flight at the same engine power setting, and the decrease in IAS is not great (less than 10 knots for most light aircraft designs). Figure 50 shows typical attitudes for level turns using 30 of bank. Note that the attitude in the turn to the right appears slightly lower than in that to the left, because the pilot's seating position is on the left-hand side of the cabin.

In practice, the pilot will find that a slight back pressure is needed to hold the control wheel in the position required to maintain the correct attitude for level flight during the turn. It is normal procedure not to use the trim wheel to eliminate this force, since the turn is only a temporary manoeuvre. When the desired heading has been attained, the wings are levelled to restore straight flight, in which the back pressure will no longer be needed.

4.24.1.1 Balance: use of rudder

In 4.12.2 it was explained that when the control wheel is moved to the left to apply bank, the ailerons are displaced in such a way as to increase the lift at the right wing tip and to decrease that at the left. The lift variation will entail a corresponding drag variation. In other words, the drag at the right wing tip will be greater than that at the left. As a result, the tendency of the aircraft to yaw to the left under the sideslip-nullifying action of the fin will be hindered by the opposite yawing effect arising from the difference in wing tip drag. Figure 177 illustrates the situation.



While the bank is being applied, then, the desired turning motion of the aircraft will be delayed. Meanwhile, the balance indicator ball will remain in its lowest position, showing unbalanced flight (Figure 178 top image).



Once the desired angle of bank has been attained, of course, the ailerons are centralised. This action will restore equality of lift and drag at the wing tips. Now the yawing action of the fin will be unhindered and the aircraft will turn to the left. The ball will be thrown outwards, adopting a central position in the indicator and so confirming balanced flight (centre image).
 
When the control wheel is moved to the right to level the wings after completion of the turn, an opposite sequence of events occurs, with the aircraft continuing to yaw under the influence of the wing tip drag difference. The result is a transient state of unbalanced flight, during which the balance indicator will be as in the bottom image.

The adverse yawing effects occurring when the control wheel is moved can be overcome by suitable use of the rudder. When the control wheel is moved to commence a turn to the left, simultaneous application of left rudder will overcome the adverse yaw. Similarly, movement of the control wheel to the right to level the wings after completion of the turn should be matched with simultaneous application of right rudder. Note that these rudder applications are as demanded by the balance indicator. Identical considerations apply to the case of initiating and completing a turn to the right. In conclusion, it can be stated that whenever the ailerons are being used to roll the aircraft, sufficient rudder should be applied in the same direction as the control wheel movement to maintain balanced flight.

Most light aircraft are designed to have good directional stability, so that in turning flight the airflow past the fuselage is symmetrical without the need for application of rudder. If the directional stability is weak, the machine will tend to sideslip slightly in the turn. The balance indicator will confirm the unbalance, showing that application of rudder in the direction of the turn is needed to rectify the situation.

4.24.1.2 Rate of turn

The concept of rate of turn was introduced in 2.6.13.1. At any particular IAS, the rate of turn depends on the angle of bank. The greater the angle of bank, the larger in magnitude is the sideways-acting force arising from the tilting of the lift force, and so the greater is the rate of turn. At a speed of 100 knots IAS, about 15 of bank is required for a rate 1 turn. A rate 2 turn requires about 30. At higher speeds, greater angles of bank are needed to achieve these rates of turn. (For example, about 20 is needed for a rate 1 turn at 150 knots IAS.)

Since the aircraft will turn whenever bank is applied (assuming that the flight is balanced), it will be understood that anticipation is necessary when completing a turn onto a desired heading. If the pilot maintains his bank until the chosen heading is attained, the aircraft will continue to turn while he levels the wings. Straight flight will only be restored after the wings have been levelled, by which time the machine will have turned past this heading. The problem is overcome by commencing the wing-levelling action just before the desired heading is reached.

4.24.2 Climbing turns

The climbing turn differs from the level turn in that the pitch attitude is adjusted as necessary to maintain the chosen climbing speed. Again, the wings are generating more lift, and drag, than in straight climbing flight. The increased drag will tend to decelerate the aircraft to a lower IAS. This effect can be prevented by adopting a slightly lower pitch attitude than for the straight climb. In practice, it will be found that, compared to level turns at the same angle of bank, less back pressure wil be needed to hold the control wheel in the correct position during climbing turns. This is because the natural tendency of the pitch attitude to become lower in turning flight will assist in the adoption of that required to maintain IAS.

Note that the greater the angle of bank, the greater is the increase in wing drag, and so the lower the pitch attitude must be to maintain the chosen climbing speed.

The extra drag detracts from the aircraft's climb performance. In specific terms, the rate of climb at any particular IAS will be reduced during turning flight. For this reason, it is usual practice to restrict the angle of bank to 15, so that the performance loss is minimised.

Besides preventing propwash-induced yaw, the pilot will need to use his rudder to counteract the wing tip drag effects previously described to ensure balanced flight while entering and completing the turn. Rudder application should be as dictated by the balance indicator.

Figure 179 shows a typical attitude for the climbing turn.



4.24.3 Descending turns

In the descending turn, the pitch attitude is again adjusted as necessary to maintain the chosen descent speed. Assuming no change in engine power setting, a slightly lower pitch attitude will be needed than that for straight descent at the same IAS. The pilot will find that less back pressure will be needed to hold the control wheel in the correct position than for level turns at the same angle of bank, because of the effect described in 4.24.2.

The greater the angle of bank, the greater is the increase in wing drag, and so the lower the pitch attitude must be to maintain the chosen descent speed. Because of the extra drag, the rate of descent at any particular IAS will be greater during turning flight than during straight flight, assuming no change in engine power setting.

While entering and completing the turn, balanced flight is ensured if the rudder is used in accordance with the balance indicator. Figure 180 shows a typical attitude for the descending turn.



4.24.4 Increase of stalling speed

At any particular IAS, the wings meet the airflow at a greater angle of attack in turning flight than during straight flight; in this way the lift force is increased sufficiently for its vertical component to equal that required for the particular vertical flight path (climb, level flight or descent).

This being the case, it is clear that if the aircraft is permitted to decelerate in turning flight, its wings will reach the stalling angle at a higher speed than in straight flight. In other words, its stalling speed will be higher. As the angle of bank increases, so does the aircraft's stalling speed. Up to angles of bank of 30 or so, the increase is not great. For example, an aircraft which has a stalling speed of 50 knots IAS in straight flight will stall at 54 knots IAS when its angle of bank is 30. It can be seen, then, that so long as the bank is restricted to medium angles, the speed range normally involved in climbing, descending and level flight will be safely above the stalling speed.

4.24.5 Summary of techniques for turning

To enter a turn in level flight, the ailerons are used to select the chosen angle of bank. At the same time, sufficient rudder should be applied to maintain balance during the entry. When the desired angle of bank is attained, the ailerons are centralised and the rudder application removed.

During the turn, the elevators are used as necessary to maintain level flight. (Slight backward movement of the control wheel will be needed.) Any tendency for the bank angle to change should be prevented by appropriate use of the ailerons. If the aircraft displays weak directional stability, rudder should be applied (to prevent sideslipping) as dictated by the balance indicator.

Anticipation is required when reverting to straight flight. Just before the chosen heading is attained, the pilot should use the ailerons to level the wings, simultaneously applying sufficient rudder to maintain balance. With good judgement, straight flight will be achieved on exactly the chosen heading. When the wings are level, the ailerons are centralised again and the rudder application removed. The control wheel should be moved forward as needed to adopt the correct attitude for level flight. (This action is merely a matter of releasing the back pressure required during the turn.)

 In climbing and descending turns, the elevators are used to maintain the chosen IAS by suitable adjustment of the pitch attitude. In the former, the bank angle should be restricted to 15 if the performance loss from greater angles is unacceptable.

4.25 STALLING AND SPINNING

We have already seen that an aircraft will stall whenever its wings reach or exceed the stalling angle. Remember that the stall is caused by the disruption of the airflow over the wings (refer back to Figure 103). It has been explained that some combinations of power and attitude result in the aircraft stabilising at a low IAS, with the wings at high angle of attack. In fact, as will be described later, it is possible for the pilot to use his controls in such a manner as to bring the wings to the stalling angle.

4.25.1 Symptoms of impending stall

As the wings approach the stalling angle, the pilot will be aware of several symptoms of the impending stall. Firstly, the increasing angle of attack is implied by the steadily reducing IAS.

Secondly, the flight controls will become progressively less effective (again because of the reducing speed).

Thirdly, as the airflow behind the wings becomes more turbulent, it will buffet the tailplanes. This buffeting will be felt throughout the airframe (Figure 181). Pre-stall buffeting is very marked on some designs of aircraft; on others it may be almost non-existent.



Fourthly, the stall-warning device (fitted to most aircraft) will be activated. The device consists primarily of a small metal vane attached by a hinge to the leading edge of one of the wings. At angles of attack well below the stalling angle, the vane is held down by the airflow, as shown in Figure 182 (top image).



As the angle of attack approaches the stalling angle, the airflow pattern at the leading edge gradually changes, until eventually it lifts the vane, as shown.

The vane is a switch in an electrical circuit. When it lifts, a warning light illuminates on the instrument panel. If now the angle of attack is made to decrease, the airflow will force the vane down again. The circuit will then be broken and the light will go out.

On some aircraft, the vane activates a warning horn instead of a light. A few machines feature both light and horn. Whichever device is fitted, it is designed to operate at an angle of attack a little below the stalling angle. The warning will therefore be triggered a few knots above the stalling speed.

4.25.2 The stall

If the angle of attack is increased to the stalling angle, the resulting sudden disruption of airflow over the wings will cause the immediate loss of most of the generated lift. The most serious consequence is that the aircraft will lose height rapidly. As it does so, the direction of airflow will be such that the tailplanes generate an upward-acting force, which will have the effect of pitching the nose of the aircraft downwards. This pitching motion will be assisted by the residual lift force, acting as it does to the rear of the CG. Figure 183 illustrates this.



Note that in this situation, any attempt to raise the nose by rearward movement of the control wheel will be unsuccessful, because of the direction of the airflow at the tailplanes.

To complicate the situation, it is possible that one wing will stall momentarily before the other, because of small manufacturing imperfections in the aerofoil shapes, or a slight difference in their angles of incidence. The result is a rolling motion towards the wing which stalls first. Depending on various factors, use of the ailerons to counteract this 'wing drop' may or may not be successful.

It will be appreciated that during the stall, the pilot will experience loss of control of his aircraft to a lesser or greater extent, because the machine will not respond in the normal manner to certain control wheel movements. Nevertheless, it is possible for full control to be recovered by reducing the angle of attack of the wings below the stalling angle, as we shall see later.

If recovery action is not taken when the aircraft stalls, any wing-dropping tendency may well lead to more serious consequences.

4.25.3 The spin

A wing dropping during the stall will effectively meet the airflow at a greater angle of attack than the other, rising wing. Reference to 3.4.2.2 tells us that the dropping wing will therefore generate more induced drag than the other. The result is that the aircraft will yaw in the same direction as the wing drop. (For example, if the left wing drops at the stall, the aircraft will yaw to the left under the influence of the increased drag generated by this wing.)

It can be seen that the yawing motion will now cause the dropping wing to move more slowly through the air than the other, and so its already reduced lift will be decreased further; this effect will reinforce the rolling motion of the original wing drop.

The machine is now in a situation in which it is rolling towards the lower wing and simultaneously yawing in the same direction. This motion is called autorotation, because it occurs automatically as a result of the initial wing drop rather than because of any application of the flight controls by the pilot.

If not arrested, the autorotative motion will stabilise itself as a spin, in which the aircraft is continuously rolling and yawing towards the lower wing in a steeply nose-down pitch attitude. The axis of rotation does not pass through the machine's CG, but through the inner region of the lower wing. Figure 184 illustrates a spin to the left.



In the spin, the airflow meets the wings at an angle of attack greater than the stalling angle; the aircraft is therefore in a stalled condition. Hence any attempt to raise the nose in this situation by rearward movement of the control wheel will be unsuccessful. Similarly, if the ailerons are used to try to raise the lower wing, the aileron displacement will increase the drag at this wing tip, while reducing that at the higher wing tip. Thus the yawing motion will be reinforced, so worsening the situation.

The reader will appreciate that during the spin the aircraft is out of control. Additionally, it rapidly loses height. Because of the oblique direction of the airflow at the pressure head, the dynamic pressure in the pitot tube is less than it would be at the same speed of motion through the air in normal flight (Figure 185).



A low IAS (at or near the stalling speed in straight and level flight) will therefore be shown on the ASI during the spin.

Note that, in order to enter a spin, the aircraft must first pass through the autorotative phase, and so a spin cannot occur unless the wings reach the stalling angle.

Later on we shall see how control of a spinning aircraft may be regained.

4.25.4 Recovery from the stall

The most serious consequences of the stall are the attendant loss of height and the danger of the aircraft entering a spin if wing drop occurs. In the event of stall occurrence, then, prompt and correct recovery action will both minimise the height loss and arrest the autorotation should wing drop be experienced.

The first requirement is to bring the wings below the stalling angle. This is achieved by moving the control wheel forward, so that the pitching-down effect of the tailplanes (shown in Figure 183) is assisted by the elevators. Simultaneously with the control wheel movement, full power is applied with the throttle lever, both to accelerate the aircraft above the stalling speed as quickly as possible and to confer greater effectiveness to the elevators.

The amount of control wheel movement required to unstall the wings depends (amongst other factors) on the individual aircraft design. If the control wheel is not moved forward enough, it will take longer for the wings to attain an angle of attack below the stalling angle, and so the stall situation, with its attendant hazards, will be prolonged. On the other hand, excess control wheel movement will result in a steeper descent than necessary and therefore greater height loss before a climb can be established to regain that lost during the stall. With practice and experience, the pilot will come to know the correct control wheel movement for his aircraft.

If a wing has dropped, any tendency of the machine to autorotate should be prevented by opposing the yaw with application of rudder. (This is merely a matter of keeping the aircraft's nose aligned with a suitable directional reference point during the recovery.) Note that use of the ailerons to raise a dropped wing may well have the opposite effect by inducing autorotation (because of the wing tip drag difference). Even aircraft designs which exhibit a degree of normal roll response to aileron displacement during the stall may not always do so if the stall occurs under certain circumstances. For this reason, the ailerons are not used to raise a dropped wing during the stall recovery procedure - the control wheel is held centrally, so that they are undisplaced.

Once the IAS is safely above the stalling speed, the pilot can deduce that his wings are below the stalling angle. Accordingly, the aircraft will now respond normally to movement of the control wheel. The next part of the recovery procedure, then, is to level the wings with the ailerons and centralise the rudder.

Finally, to convert the descent into a climb, so that the height lost during the stall is regained, the nose is gently raised to a suitable climb attitude. Rudder should now be used as necessary to balance the aircraft.

If it stalls, the typical 4-seat light aircraft may well lose about 300 feet before the climb can be established. The height loss may be greater if the recovery is carried out hesitantly or incorrectly. The reader will appreciate the seriousness of stall occurrence 200 feet above ground level.

In summary, the correct stall recovery procedure is:

(a) apply full power, simultaneously moving the control wheel forward, keeping the ailerons centralised, sufficiently to unstall the wings;

(b) maintain constancy of heading by appropriate application of rudder;

(c) once the IAS is safely above the stalling speed, level the wings (if they are not so) with the ailerons and centralise the rudder;

(d) raise the nose to establish a climb to regain lost height, using rudder to balance the aircraft.

4.25.5 Practising stalls

By now, the reader will have gathered that, should a stall occur, the consequences could be extremely serious. For this reason, pilots periodically practice stalling in circumstances where no harm can arise, to maintain familiarity with the recovery procedure. As well as being able to cope with a stall, the pilot must learn to recognise the pre-stall symptoms, so that inadvertent stalls may be prevented. The philosophy is that prevention is better than cure.

Intentional stalling may be prohibited in a few designs of aircraft. In most others, intentional stalling is permitted providing certain requirements are satisfied, as detailed in the Flight Manual. Typically, these requirements include:

(a) ensuring that the aircraft's loaded weight is not above the maximum permitted for stalling. (This may well be below the maximum total weight authorised (MTWA));

(b) ensuring that the CG of the aircraft as loaded is within the permitted range for stalling. (Again, the permitted range may be more restrictive than for normal operations.)

The reasons for these precautions will become clear later on.

Just prior to commencing his stalling practice, the pilot should action the pre-stalling checks detailed in the checklist. The following considerations will be covered by the pre-stalling checks:

(a) the stall and recovery should be conducted at a safe height above ground level. The pilot should aim to have completed his stall recovery by at least 2000 feet above ground level. This cushion will provide ample space for recovery should an inadvertent spin occur. Figure 186 illustrates the point;



(b) the aircraft's controls should be correctly set. Specifically, the pilot should check that the flaps are in the desired position, that the mixture control is set to 'rich' (because recovery entails use of full power), and that the carburettor is free of ice (by use of the carburettor heat control). The fuel cock should be feeding from a tank containing an adequate quantity of fuel, and the engine gauge indications checked as being within the permitted limits;

(c) the harnesses of all persons on board should be fastened and any loose articles (such as cases and coats) should be firmly secured or stowed;

(d) the pilot must ensure that the aircraft is not within airspace within which stalling is not permitted, nor over built-up areas;

(e) he must also verify that the aircraft is well clear of other traffic and cloud, both around the machine and also above and below it. Banking left or right into 'clearing turns' will enable the pilot to inspect the area in question.

4.25.5.1 Entering the stall from straight and level flight

From straight and level cruising flight the pilot closes the throttle. To prevent the height loss arising from the decaying IAS, a progressively higher attitude is adopted by appropriate rearward movement of the control wheel. In this way, the lift required for level flight is maintained by the increasing angle of attack. Figure 187 illustrates.



As the aircraft decelerates, the pre-stall symptoms will become apparent to the pilot. Eventually, a stage will be reached when, despite further rearward movement of the control wheel, the aircraft's nose will pitch down. The stall has occurred. Whether the nose drops abruptly or gently depends not only on the aircraft design but also upon other factors explained below. Simultaneous wing drop may also occur.

Recovery action should be taken as soon as the nose drops.

Note that the lower end of the green arc on the ASI (see Figure 55) shows the stalling speed in level flight with flaps up and the aircraft at MTWA. In the example shown the stalling speed under these conditions is 46 knots IAS.

4.25.5.2 Effect of lowered flaps on stalling characteristics

If the stall is entered with the flaps at the 'maximum lift' position, the deceleration will be more rapid, because of the increased drag. The stall will occur at a slightly lower IAS and, for aerodynamic reasons, is likely to be more abrupt than with the flaps up, with a greater tendency for wing drop to occur.

During the recovery, the pilot must ensure, by careful attitude control, that the flap limiting speed is not exceeded. As soon as the climb is established, the flaps should be raised.

Stall entry with the flaps fully down will accentuate the effects just described. When climbing after the recovery, the flaps should be raised in stages.

Note that the lower end of the white arc on the ASI (see Figure 55) shows the stalling speed in level flight with flaps fully down and the aircraft at MTWA. In the example shown the stalling speed under these conditions is 40 knots IAS.

4.25.5.3 Effect of power on stalling characteristics

If, instead of closing the throttle completely, the pilot sets the engine for low power, the deceleration will be slower, because of the thrust. The loss of effectiveness of the rudder and elevators during the stall entry will be less marked, because of the propwash. For aerodynamic reasons, the stall is likely to be more abrupt than with the throttle closed, with a greater tendency for wing drop to occur. The aircraft will stall at a slightly lower IAS. This is mainly because the strength of the propwash over the inner areas of the wings will delay the break-up of the airflow. It will also enhance the lift generated by these areas, so that the remainder of the lift required can be generated at the stalling angle at a lower speed.

4.25.5.4 Entering the stall from level turning flight

If bank is applied and maintained during the stall entry, the pilot will notice that the stall occurs at a higher IAS than in straight and level flight. The stalling speed increases with the angle of bank, as was explained in 4.24.4. The increase is not great for medium angles of bank.

At the stall, the wing drop effect may increase or decrease the angle of bank. Nevertheless, normal recovery action should be taken, with the ailerons held centrally until the IAS is safely above the stalling speed.

4.25.5.5 Stalling in climbing and descending flight

Remember that the aircraft will stall whenever its wings reach or exceed the stalling angle, regardless of the actual flight path. Remember, too, that attitude alone is not a direct indication of angle of attack. Refer back to Figure 154, which illustrates two aircraft with the same attitude, yet with different angles of attack, because of their differing flight paths. Of course, the pilot of the aircraft in the bottom image in Figure 154 will observe a lower IAS than the other, and this is his clue as to the greater angle of attack of his wings.

An aircraft can be made to stall by sufficient rearward movement of the control wheel, regardless of its flight path, which, besides adopted attitude, also depends on engine power setting. By itself, then, attitude is not related to stall occurrence. Figure 188 shows two aircraft with their wings at the stalling angle, but with considerably differing attitudes.



Only the pre-stall symptoms previously described reliably indicate an impending stall.

4.25.5.6 Stalling in pitch-up manoeuvres

If the pilot abruptly moves the control wheel rearwards, the aircraft will momentarily continue along its original flight path, because of its inertia, despite the pitch-up change in attitude. If the wings are near the stalling angle when the control wheel movement is made, it is clear that they may exceed the stalling angle and so precipitate a stall at an IAS above the stalling speed in straight and level flight. In other words, the stalling speed is higher during such manoeuvres. The pilot will be aware of stall occurrence as failure of the nose to rise further despite the rearward movement of the control wheel. Figure 189 illustrates the situation.



4.25.6 Effect of loaded weight on stalling speed

The greater the loaded weight of an aircraft, the higher the IAS must be at any particular angle of attack if the lift required to balance this weight is to be generated. Of course, this consideration will apply at the stalling angle.

We can conclude that the greater the loaded weight, the greater is the aircraft's stalling speed. For example, an aircraft which stalls at 45 knots IAS in straight and level flight when its loaded weight is 800 kilograms (kg) will stall at 50 knots IAS when loaded to 1000 kg.

An aircraft's Flight Manual will quote the power-off (throttle closed) stalling speed (as an IAS) in straight and level flight at MTWA with flaps up and with flaps down. Stalling speeds at lower weights may also be given.

4.25.7 Effect of CG position on stalling characteristics

In 3.8.5.1 it was explained that rearward CG positions reduced the magnitude of the tail-down force required to balance the relative displacement of the lift force from the CG. As a result, less rearward movement of the control wheel will be needed to bring the wings to the stalling angle. Consequently, careless handling of the yoke is more likely to precipitate a stall when the loaded aircraft has a rearward CG position than when it has a forward CG position.

Furthermore, the natural tendency of the nose to pitch down at the stall will be reduced, because of the reduced displacement of the residual lift force from the CG. Thus greater forward movement of the control wheel will be needed to unstall the wings during the recovery.

In summary, the aircraft will stall more easily, and show greater reluctance to recover, when loaded in such a way as to give a rearward CG position.

Conversely, forward CG positions will require greater rearward movement of the control wheel to bring the wings to the stalling angle. In fact, in extreme cases, it may be impossible to stall the aircraft, even with the control wheel fully rearwards. However, although extreme forward CG positions help to prevent inadvertent stalling, they also reduce the aircraft's pitch controllability, as will be explained later. A forward limit for CG position is therefore imposed for each aircraft design.

4.25.8 Inadvertent stalls: recovery at the incipient stage

Having thoroughly explored the stalling characteristics of his aircraft, the pilot will go on to practise recovery at the incipient stage, that is, when he first becomes aware of the pre-stall symptoms. Depending on circumstances, however, not all of these symptoms will make themselves apparent.
 
For example, loss of effectiveness of the elevators and rudder will not occur if the propwash is strong (engine delivering power). Additionally, some aircraft do not exhibit pre-stall buffeting.

So far as IAS is concerned, the pilot should bear in mind that the stalling speed will be reduced slightly when the flaps are lowered, and also when the engine is delivering power. More importantly, the stalling speed will be increased during turning flight and in pitch-up manoeuvres. Accordingly, the pilot must ensure, by judicious handling of attitude and power, that his IAS is always comfortably above the stalling speed in the particular flight path his aircraft is following.

Activation of the stall-warning device is the most reliable pre-stall symptom. The pilot's response should be immediate and correct stall recovery action. Note that when recovering at the incipient stage, less forward movement of the control wheel will be needed to reduce the angle of attack of the wings safely below the stalling angle. Height loss is considerably reduced when recovery action is taken at the incipient stage of the stall.

A point to bear in mind is that if power is not available, such as when gliding into land after engine failure, recovery from an inadvertent stall can be effected solely by forward movement of the control yoke. The recovery will therefore take longer and so will involve a greater loss of height before control of the aircraft is regained.

4.25.9 Recovery from the spin

We have seen that an important part of the stall recovery procedure is the prevention of yaw caused by wing drop, using the rudder. Failure to prevent the yaw may well result in autorotation, which in turn will develop into a spin.

Because of its self-perpetuating nature, the spin is a stable flight path, even though the aircraft is out of control. Even when the correct spin recovery action is initiated, the machine may not stop spinning for several seconds, because this stability has to be overcome. The importance of the height 'cushion' when practising stalls can now be fully appreciated.

Note that if the engine is delivering power during the spin, its only result will be to accelerate the descent. For this reason, the first action in the spin recovery procedure is to close the throttle (if not already so).

After analysis of the rather complex disposition of forces acting on the aircraft during the spin, it has been found that the only way that control can be regained is firstly to overcome the yawing motion. Accordingly, the next action in the recovery procedure is to apply and maintain full rudder in the opposite direction to the yaw. For example, if the aircraft is spinning to the left, full right rudder must be applied and maintained.

After a brief pause (1 second or so) the control wheel should then be moved progressively and steadily forward to try to unstall the wings. Note that the aircraft will still be spinning during this action, despite the rudder application. Because of the direction of the airflow during the spin, the elevators will to a certain extent screen the rudder when the control wheel is moved forward, and so reduce the latter's effectiveness. The brief pause beforehand will enable the rudder to have maximum effect in reducing the yawing motion. Figure 190 illustrates the point.



The progressive forward movement of the control wheel should be continued until the spin stops. At this point, the aircraft will be in a steeply nose-down, banked attitude. It will therefore accelerate rapidly. The final part of the recovery procedure, then, is to centralise the rudder, level the wings with the ailerons (to which the aircraft will respond normally, now that the wings are below the stalling angle) and gently raise the nose to a climbing attitude. (Abrupt rearward movement of the control wheel may well precipitate another stall.) The engine should now be set to climb power, so that a climbing flight path is established.

Three points to bear in mind are, firstly, that attempts to use the ailerons to overcome the rolling motion in the spin will most likely worsen the situation, because the wing tip drag effects will merely assist the yaw, and may therefore hinder or even prevent recovery. Unless the aircraft's Flight Manual states otherwise, the ailerons should be kept centralised as the control wheel is moved forwards.

Secondly, the motion of the aircraft during the spin may disorientate the pilot to the extent that he may not be able to identify the direction in which his machine is spinning. If such doubt exists, he should consult the turn indicator, which responds to the yawing motion. In the recovery procedure, the pilot should apply full rudder in opposition to the direction of turn shown by the indicator.

Thirdly, considerable height is lost in the spin, during which the rate of descent may well exceed 5000 feet per minute. The importance of prompt and correct recovery action will be appreciated.

In summary, the correct spin recovery procedure is:

(a) close throttle, if not already so;

(b) apply and maintain full rudder in opposition to the yaw;

(c) after a brief pause, move the control wheel progressively forward, keeping the ailerons centralised, until the spin stops;

(d) centralise the rudder;

(e) level the wings with the ailerons;

(f) gently raise the nose to a climbing attitude and apply climb power. Figure 191 illustrates the spin recovery procedure.



4.25.10 Recovery from spiral dive

It may well be that, following mis-handling of the controls, the pilot gains the impression that his aircraft has entered a spin. The ASI will confirm the diagnosis if it shows a low, steady IAS. An indication of increasing speed implies that the aircraft is in a spiral dive rather than a spin.

Recovery from a spiral dive is effected with normal control wheel movements. Specifically, the wings should be levelled by use of the ailerons and the nose raised using the elevators.

4.25.11 Practising spins

With spins, as with stalls, the philosophy is that prevention is better than cure. Nevertheless, conscientious pilots prefer to practise spinning periodically, to maintain familiarity with the recovery procedure should it ever be required.

In most light aircraft, intentional spinning is permitted providing certain requirements are fulfilled. These are to be found in the Flight Manual, and usually include those concerned with loaded weight and CG limitations mentioned in 4.25.5.

Prior to spinning, the pilot should action the pre-spinning checks as detailed in the checklist. Usually, the pre-stalling checks will also serve for spinning. In order to be able to recover by 2000 feet above ground level, the spin entry will need to be carried out well above this height. For the typical light aircraft, 5000 feet above ground level would be a safe entry height. In most machines, intentional spinning is prohibited unless the flaps are up.

4.25.11.1 Entering the spin

Of course, spin entry could be achieved simply by deliberately stalling the aircraft and neglecting to take recovery action. However, this method may involve considerable loss of height before autorotation sets in. Normal procedure, then, is to use application of rudder to yaw the aircraft just as it approaches the stall. The yaw will start a rolling motion (refer back to 4.12.3.1) and so induce autorotation.

In practice, full rudder is applied in the intended direction of spin as the stalling speed is approached, and the control wheel then moved fully rearwards to stall the wings. The aircraft will roll until almost inverted, with the nose dropping abruptly below the horizon. This is the autorotative phase. Very quickly, the machine will assume a steep nose-down attitude, rotating rapidly. The spin has occurred.

As a general rule, the spin tends to become more stable the longer it is allowed to continue, and the aircraft will show greater reluctance to recover. An aircraft's Flight Manual may therefore specify the maximum number of rotations, or 'turns', that are permitted during intentional spinning. If no such advice is given, recovery should be initiated after not more than three turns, since no practical value is to be gained from a prolonged spin, and valuable height is wasted.

4.25.12 Effect of CG position on spinning characteristics

It has already been demonstrated that an aircraft will stall more easily, and show greater reluctance to recover, when loaded in such a way as to give a rearward CG position.

Perhaps it is not surprising to discover that in these circumstances an aircraft will also enter a spin more easily, and that the spin itself will be more stable.

Again, the machine will show more reluctance to respond to the recovery procedure. In fact, extreme rearward CG positions (outside the Flight Manual limit for normal operations) may preclude spin recovery.

Since a practice stall may lead to an inadvertent spin, it is therefore necessary to take precaution (b) mentioned in 4.25.5, so that recovery action will quickly stop the spin.

4.25.13 Airframe stress during spin recovery

During the transition from descent to climb in the spin recovery procedure, the airframe is subjected to greater stresses than in normal flight manoeuvres, because the wings are momentarily generating considerably more lift, to overcome the machine's inertia. The greater the loaded weight of the aircraft, the more is its inertia, and so the greater are the stresses experienced by the airframe in this manoeuvre.

With some designs, it is therefore necessary to restrict the loaded weight for intentional stalling, so that if a spin occurs, the machine will not be overstressed during the recovery procedure; hence precaution (a) in 4.25.5.

4.25.14 Inadvertent spins

The reader will remember that the spin is caused by yawing motion, which leads to autorotation, when the wings are at the stalling angle. We have seen that the yaw may arise either as a result of wing drop, or by application of rudder (when the spin is intended). Without yaw, autorotation cannot occur, and so neither can the spin.

If a pilot inadvertently allows his aircraft to stall, he must take prompt and correct stall recovery action, part of which requires that any yawing motion is arrested by appropriate use of rudder. The stall recovery procedure will therefore prevent spin occurrence.

However, a stall with crossed controls will almost immediately lead to autorotation. Take the example of a pilot attempting to turn sharply at low speed using rudder rather than bank to effect the turn. Suppose that he wants to turn to the right and has applied right rudder to try to turn at a greater rate. Perhaps he is reluctant to bank his aircraft, and has moved the control wheel to the left to restrict the rolling motion to the right that occurs as a consequence of the rudder-induced yaw. Figure 192 illustrates the situation.



If the pilot now accidentally moves the control wheel far enough rearwards to stall the wings, then the yaw from the effect of the rudder, reinforced by that from the wing tip drag difference, will cause immediate autorotation, and the aircraft will enter a spin to the right. (In fact, stalling with crossed controls can be used to enter intentional spins.)

Note that the spin can always be averted if stall recovery action is taken during the autorotative phase.

4.25.15 Stall and spin avoidance

The reader may have gained the impression that only gross mishandling of the controls will lead to inadvertent stalls and spins. Nevertheless, it is worth remembering that most stall- and spin-related accidents are the result of several contributory factors. Take the case of an aircraft loaded to its MTWA, with a rearward (but within the Flight Manual limit) CG position. Possibly the weather is poor, with the horizon ill-defined, so that attitude judgement is difficult. Perhaps the pilot has just taken off, or is approaching to land (in both of which situations the IAS tends to be in the lower range) and becomes distracted for some reason or other, so that he fails to monitor his ASI. Careless handling of the controls in such circumstances might well have disastrous consequences.

Restating our dictum that prevention is better than cure, it cannot be overemphasised that an aircraft will never stall or spin unless the wings are brought to the stalling angle. Adherence to the correct handling techniques for climbing, descending, turning and level flight, together with frequent monitoring of IAS when flying in the lower speed range, will guarantee avoidance of stall situations.

Pitch-up manoeuvres, such as transition from descent to climb, should be made gently, not only to prevent stall occurrence, but also to guard against possible overstressing of the airframe.

Even if the correct techniques are accidentally disregarded, the pre-stall symptoms will warn of impending mishap. If such symptoms, particularly activation of the stall-warning device, become evident, the pilot must take immediate and correct stall recovery action.

4.26 THE RUNWAY

Every airfield has one or more runways for the use of aircraft taking off or landing. Each runway is a rectangular-shaped area whose surface may be hard (for example, concrete or tarmac) or grass-covered.

Hard runways usually have a lengthwise dotted white line painted along their middle - the centreline. Grass runways are normally delineated by dotted white edge markings, with their centrelines sometimes marked.

Additionally, most runways have identification figures marked at their ends, or thresholds. The identification refers to the magnetic direction in which the runway points, taken to the nearest 10, with the final zero omitted. Thus a runway pointing in the direction 076M will bear the identification numbers '08'. Considered from the opposite direction, the same runway would be designated '26' (because it would point in the direction 256M).

Figure 193 shows a tarmac runway 05/23 and a grass runway 17/35.



Runways vary considerably in their lengths and widths. The pilot will need to make sure that the runway intended for take-off or landing is long enough for his aircraft in the circumstances prevailing. This consideration will be dealt with in detail later on.

4.27 THE CIRCUIT

The circuit is a procedure in which the aircraft takes off, climbs to a certain height and then positions to approach and land on the same runway from which it departed. For reasons which will be explained later, the take-off and landing are made on a runway facing into the wind.

Each runway has an associated circuit pattern, which is the airspace within which circuits are flown. The circuit pattern has a rectangular form. Figure 194 shows the features of a left-hand circuit (so called because all turns are made to the left).



At most airfields the circuit height is 1000 feet above airfield level, as in the example shown. Other airfields may have a circuit height of, for example, 800 feet.

Since the take-off and landing are made facing into wind, two of the legs in the circuit pattern are named accordingly. The term 'crosswind' is self-explanatory; 'downwind' means opposite to the direction from which the wind is blowing.

4.28 THE TAKE-OFF

During the take-off, the aircraft accelerates on the runway from a rest position to a speed at which it can safely be made lo leave the ground and take up a climbing flight path. In other words, the machine must be accelerated to such a speed that an angle of attack safely below the stalling angle will generate the lift required to become airborne.

Of course, the IAS corresponding to this take-off angle of attack depends on the aircraft's loaded weight and on the flap position. The take-off speed quoted as an IAS in the Flight Manual refers to the case when the aircraft is loaded to its MTWA with the flaps up. The take-off speed with the flaps at 'maximum lift' may also be given.

In practice, the variation in take-off speed with weight is not significant, and take-offs at lighter weights are made at the speed quoted for the MTWA.

If the runway length required for take-off is well short of that available, as is usually the case with light aircraft, take-off is normally made with the flaps up, so that maximum rate of climb can be attained as soon as the machine is airborne.

4.28.1 Normal take-off technique (flaps up)

Having completed his pre-take off checks, the pilot taxies his aircraft onto the runway at the threshold and aligns the machine with the centreline. This procedure is called 'lining up'. The pilot's view of the runway will be as in Figure 195. (The lateral displacement of the machine due to his offset seating is negligible.)



The brakes are released and full power is applied. (The throttle should be held fully open during the take-off.)

Any tendency of the aircraft to deviate from the centreline is prevented with the rudder pedals, which have their effect through the nosewheel steering and also by virtue of the propwash flowing past the rudder. In other words, directional control during the take-off run is merely a matter of steering the aircraft along the centreline, as one would a car. If the centreline is not marked, the pilot maintains equal spacing from each runway edge.

Unless otherwise stated in the Flight Manual, the control wheel should be held centrally (with both ailerons and elevators undisplaced) while the aircraft accelerates to take-off speed.

The pilot will pay most attention to the view ahead during the take-off run to monitor the accuracy of his steering, with occasional glances at the ASI. As the aircraft accelerates, he will find that smaller rudder pedal movements will be needed to correct displacements, because of the increasing effectiveness of the rudder.
 
When the take-off speed has been reached, the control wheel is moved rearwards until the nose of the aircraft begins to rise. The machine will then become airborne, or lift off. Slight forward movement of the control wheel will prevent the nose rising further, so that the IAS will continue to increase (Figure 196).



When the aircraft has accelerated to the maximum rate climbing speed, the nose should be raised to the climbing attitude, so that this speed is maintained.

During the climb-out, the ailerons are used as necessary to keep the wings level and the rudder to balance the aircraft. Figure 197 shows the take-off profile.



At a safe height, the engine power can be reduced to climb power (if continuous use of full power is not permitted) and the attitude lowered correspondingly to maintain the correct IAS.

4.28.1.1 Wind

The main advantages to be gained from taking off from a runway facing into wind are:

(a) the aircraft reaches its take-off speed at a lower groundspeed (GS), so reducing both the stresses on the landing gear during the take-off run and also the length of runway required to become airborne;

(b) at any particular climbing IAS the subsequent gradient of climb is steepened, because of the reduced GS, and so any obstacles below the flight path are overflown with greater clearance (Figure 198).



The stronger the wind, the greater are these two effects.

4.28.2 Crosswind take-off

Frequently, the wind blows from a direction which does not coincide exactly with the available runway directions. When this is the case, the pilot is obliged to make a crosswind take-off. An example of such a situation is shown in Figure 199.



During the take-off run, the directional stability imparted by the fin will attempt to yaw the nose of the aircraft into the wind, to nullify the asymmetry of airflow. This weathercocking action will tend to steer the machine away from the centreline (Figure 200).



However, the crosswind will also tend to blow the entire machine away from the centreline in the opposite direction to the weathercocking action (Figure 201).



The two effects thus act in opposition to each other, although one may predominate, depending on the aircraft's design. So far as the pilot is concerned, however, the rudder pedals are merely used as necessary to steer the machine along the centreline.

The asymmetric airflow will also attempt to lift the 'into wind' wing prematurely, because of the aircraft's laterally stabilising design features (dihedral or high-wing arrangement). In the example shown, with a crosswind from the left, the tendency of the left wing to lift during the take-off run can be prevented by holding the control wheel partially to the left, so that the displaced ailerons oppose the wing-lifting effect. (Of course, the control wheel should be held partially to the right if the crosswind is from the right.)

As is explained below, the lift off should be delayed until the IAS is a few knots higher than the normal take-off speed. Immediately prior to moving the control wheel rearwards, the ailerons should be centralised again, so that the aircraft becomes airborne with its wings level.

Once airborne, the machine will be blown away from the centre-line by the crosswind (Figure 202). Note that this drifting effect is relative to the ground. The airflow past the aircraft is symmetrical, assuming that the flight is balanced.



The drifting should be allowed for by turning the aircraft onto a heading slightly into wind, so that its path over the ground, or track, is along the extended centre-line (Figure 203).



Note that if the aircraft momentarily touches the runway again after becoming airborne (perhaps because of unevenness of the surface), the drifting effect will impose a sideways-acting stress on the landing gear and tyres. By lifting off at a slightly higher speed, the machine will leave the ground more crisply, so avoiding any tendency to touch again.

4.28.2.1 Crosswind component

In Figure 199, the wind can be considered as being split into a headwind component (opposite to the runway direction) and a crosswind component (at right angles to the runway), as shown in Figure 204.



The magnitude of the crosswind component depends on the strength of the wind and the angle between the runway direction and the wind direction (that is, the direction from which the wind is blowing). An increase in either factor will increase the crosswind component.

To avoid difficulty in controlling the aircraft during take-off, the pilot must ensure that the crosswind component is not greater than the maximum permitted for his aircraft (as specified in its Flight Manual). Consultation of a crosswind table, such as that shown in Figure 205, will enable him to assess the crosswind component. Of course, the table is equally valid for a crosswind from the right. Note that wind strength is expressed in knots.



4.28.3 Short take-off

The length of take-off run required to become airborne can be reduced by use of the short take-off technique, which differs from the normal technique in two respects:

(a) the flaps are set to 'maximum lift' (during the pre-take-off checks);

(b) a lower take-off speed (usually specified in the Flight Manual) is used.

Once airborne, the aircraft is made to fly at the optimum IAS for climbing with the flaps at 'maximum lift'. When a safe height has been reached, the flaps may be raised and maximum rate climb established.

Most pilots use the short take-off technique when taking off from grass runways, because not only does the grass retard acceleration slightly, but also the surface is often more uneven than that of hard runways. The reduced take-off run and lower take-off speed therefore help to minimise the stresses imposed on the landing gear.

Note that a further small reduction in take-off run can be gained by adopting the following procedure at the start of the take-off:

(a) when the aircraft has been positioned at the threshold, firmly apply the brakes;

(b) set maximum power;

(c) release the brakes and continue with the take-off as already described.

4.28.4 Obstacle clearance after take-off

At some airfields, there are obstacles (such as trees) quite close to the ends of the runways. The pilot will want to make certain that his aircraft safely clears such obstacles after take-off. Maximum obstacle clearance would be assured by:

(a) becoming airborne in the shortest distance possible;

(b) thereafter achieving the maximum possible gradient of climb.

Straight away, we can see that a dilemma confronts us, for use of the flaps to shorten the take-off run will detract from the subsequent climb gradient. Conversely, take-off with the flaps up, whilst maximising the climb gradient, will lengthen the take-off run. Figure 206 illustrates the two cases.



The Flight Manual will usually state the recommended obstacle clearance technique for any particular design. The performance capabilities of most light aircraft are such that use of the short take-off technique, with the flaps at 'maximum lift', is more likely to give better clearance over obstacles close to the runway. Of course, such a choice is not available to the pilot of an aircraft not fitted with flaps.

4.28.5 Rejected take-off

If, during the take-off run, the pilot becomes unhappy about the proceedings, the take-off should be rejected, by:

(a) closing the throttle;

(b) braking the aircraft to taxying speed;

(c) steering clear of the runway as soon as practicable to permit its use by other traffic.

If the take-off is rejected at high speed, for example just before lift off, it may be necessary to apply maximum braking to avoid overrunning the runway. In fact, if the latter is not very long, overrunning may be unavoidable. The implications of this situation will be discussed later.

Reasons for rejecting take-off might include:

(a) loss of power or rough-running of the engine;

(b) poor acceleration;

(c) malfunction of the ASI;

(d) if it becomes evident to the pilot that insufficient runway remains available for take-off speed to be reached.

Note that situation (d) should not arise if the pilot has correctly assessed the runway length beforehand as being adequate for his aircraft in the circumstances prevailing.
 
4.29 FLYING THE CIRCUIT

The aircraft is flown around the circuit pattern as detailed in Figure 194. The following points are noteworthy:

(a) the climbing turn onto the crosswind leg is initiated as soon as a height of 500 feet has been attained;

(b) loss of climb performance during the turn is minimised by restricting the bank angle to 15;

(c) in order to maintain the rectangular track, the pilot will need to take up headings slightly into wind to allow for drift whilst flying the crosswind leg and the base leg (Figure 207);
 


If a crosswind prevails, similar compensation will be required during climb-out, on the downwind leg and on final approach. Figure 208 illustrates the case of a crosswind from the right;



(d) the pilot judges the accuracy of his track-keeping by reference to the aspect which the runway presents throughout the procedure. Accuracy will improve with practice and experience. As a general rule, the downwind leg is spaced about nautical mile from the runway;

(e) the pre-landing checks ensure that the aircraft's controls are correctly set for landing. Typically, they include the following:

(1) brakes off;

(2) harness tight;

(3) mixture rich;

(4) carburettor free of ice;

(5) fuel cock selected to draw fuel from a tank with sufficient quantity;

(6) electrically-driven fuel pump switched on.

The pre-landing checks will also prepare the aircraft for a go-around (refer back to 4.23.6) should one be made.

4.30 THE APPROACH

An aircraft is considered to be 'making an approach' after it has started its descent on base leg. The powered approach involves the rather complex transition from level cruising flight with flaps up to descending flight at approach speed with flaps lowered.

To divide the resulting trim change and drag increase into more manageable quantities, it is usual practice to lower the flaps in stages, setting them to 'maximum lift' on base leg and then fully down on final approach.

Once the approach is initiated, the pilot makes adjustments of power as necessary to control the aircraft's rate of descent, whilst maintaining the correct approach speed by appropriate changes of attitude. Ideally, the pilot should attempt to achieve a steady rate of descent, with the engine at low intermediate power setting, so that the aircraft finally reaches the runway threshold at ground level.

4.30.1 Base leg

From level cruising flight, the approach is initiated on base leg as follows:

(a) reduce power to a low intermediate setting, whilst maintaining the cruise attitude (which will assist the resulting deceleration);

(b) when the IAS has fallen below the flap limiting speed, lower the flaps to 'maximum lift', still maintaining attitude;

(c) as the IAS falls further to the intermediate approach speed, lower the nose as necessary to maintain this speed. The intermediate approach speed is usually a few knots greater than final approach speed;

(d) trim the aircraft.

A rate of descent should be chosen such that the aircraft is ready to be turned onto final approach at about 600 feet. On base leg, the oblique appearance of the runway will gradually decrease as the machine approaches the extended centre-line. Figure 209 shows the change in appearance of the runway as base leg is flown in a left-hand circuit.



The descending turn onto final approach should be anticipated, so that the aircraft is positioned on the extended centreline at the completion of the turn. Thus, in a left-hand circuit, the turn might be initiated when the runway appears as in the centre example of Figure 209.

4.30.2 Final approach

During the descending turn from base leg onto final approach, the pilot should carefully monitor his ASI, lowering the pitch attitude as necessary to maintain the correct intermediate approach speed. Here is a situation in which poorly handled controls, at too low an IAS, might well precipitate an inadvertent stall. To minimise the increase in stalling speed during the turn, the bank angle should be limited to 30.

If the base leg descent and final turn have been well-judged, the aircraft will have completed the turn at about 500 feet and will now be aligned with the runway's extended centreline. The flaps should be fully lowered and the power and attitude adjusted as necessary to achieve a steady descent path towards the threshold at final approach speed. This speed for most aircraft designs is slightly less than the normal (minimum gradient) gliding speed, and will be quoted as an IAS in the Flight Manual. Figure 210 shows a typical view from the windscreen on final approach.



If the circuit pattern has been flown correctly, a gradient of descent on final approach of about 5 will bring the aircraft to the threshold at ground level (Figure 211).



So far as the pilot is concerned, the gradient of descent actually achieved depends on two factors, namely the groundspeed (GS) and the rate of descent. Of course, GS itself will depend on the strength of the wind. Thus a lower rate of descent will be required when a strong wind results in a reduced GS, and a greater rate of descent when the wind is not so strong. In other words, the pilot will find that the stronger the wind, the higher will be the power setting needed to achieve the ideal descent gradient at the correct IAS.

4.30.2.1 Approach path control

Of course, the pilot cannot see a side view of his approach path. The only clue as to whether the flight path is satisfactory is the view ahead through the windscreen.

On any particular approach path, the position on the ground at which the aircraft will reach ground level will appear to take up a fixed point in the windscreen. This 'sighting point' serves as a reference for assessing the flight path. Figure 210 represents an ideal approach path, with the sighting point (shown as a white asterisk in the diagram) coincident with the threshold.

If the approach is too high, the threshold will appear to be moving below the sighting point, and the perspective of the runway and its surroundings will appear too 'deep'. The consequence of too high an approach is that the aircraft will reach ground level well past the threshold, as the sighting point indicates (Figure 212).



On the other hand, if the approach is too low, the threshold will appear to be moving above the sighting point and the perspective of the runway and its surroundings will appear 'flattened'. The consequence of too low an approach is that the aircraft will reach ground level well short of the threshold, again as the sighting point indicates (Figure 213).



A too-high approach is corrected with a temporary increase in rate of descent, by reducing power. During this corrective manoeuvre the lowered attitude required to maintain approach speed will result in the threshold appearing to move above the sighting point. The top image in Figure 214 represents a side view.



Conversely, a too-low approach is corrected by a temporary reduction in rate of descent (increase in power). During the corrective manoeuvre the threshold will appear to move below the sighting point. The bottom image in Figure 214 shows the side view.

In either case, the corrective action should be continued until the runway perspective changes to that associated with the ideal approach path. At this point the original rate of descent should be restored, so that the sighting point is now coincident with the threshold.

In conclusion, an ideal approach path is one in which the sighting point coincides with the threshold and in which the runway perspective is neither too 'deep' nor too 'flattened'. Deviations are corrected by suitable adjustments in power.

With practice and experience, the pilot will come to recognise such deviations in their early stages, and will therefore need only small power adjustments to rectify the situation.

4.30.2.2 Centre-line tracking

The reader will appreciate that the pilot's workload during final approach is considerable, because not only must he maintain the ideal approach path, whilst monitoring his ASI, but also the aircraft must be flown so that it tracks along the extended centreline.

Displacements from the centreline are indicated by the associated oblique appearance of the runway. They are rectified by turning onto a corrective heading until the centreline has been regained, at which point the aircraft is turned back onto the heading needed to track along the centreline. Figure 215 shows a plan view of corrective action needed to rectify a displacement to the right, with the pilot's forward view as the correction is initiated.



If displacements are detected in their early stages, then only small heading corrections will be needed to rectify the situation.

4.30.2.3 Alternative technique for final approach

Some pilots favour a different control technique for final approach. They adjust the vertical path using pitch attitude and speed using engine power setting. The rationale behind this technique is that the approach is a defined flight path (in the same way as constant altitude is a defined flight path in level flight, in which speed is also controlled by engine power setting).

Whichever technique is adopted, the aim is identical: to find the combination of attitude and power setting that brings the aircraft towards the runway threshold on the correct approach path at the correct final approach speed.

4.31 THE LANDING

The landing is a manoeuvre in which the aircraft is brought into contact with the runway and then decelerated to taxying speed. The primary considerations are, firstly, that the machine is landed in such a way that it will not become airborne again after touchdown, and secondly, that the landing gear is not subjected to excessive stresses.

4.31.1 Normal landing technique

For the typical light aircraft, the transition from powered approach to landing begins at about 200 metres from the threshold, and involves a change in the manner in which the controls are used.

The most important change is in the role of the elevators, which, rather than being used to control IAS (through attitude changes) as during the powered approach, now take on the function of controlling the aircraft's vertical flight path. This is because the landing manoeuvre involves a change from descending flight to level flight just above the runway. (The landing manoeuvre can therefore be considered as a defined flight path).

The first phase of the landing technique concerns the flight path from the transition point to the threshold, during which the power is gradually reduced to minimum. At the same time, the attitude is raised at such a rate, and to such an extent, that the aircraft arrives at the threshold in level flight just above the runway.

The reader will appreciate that the reduction of power during the transition, together with the raising of the nose, will result in the reduction of IAS. If the transition point has been chosen at the correct distance from the threshold, the aircraft will have decelerated to threshold speed (specified as an IAS in the Flight Manual) as it arrives at the threshold.

Figure 216 shows a side view of the first phase of the landing technique, called the 'flare', together with what might be a typical view from the windscreen just before the transition point, and as the threshold is reached.



The second phase, called the 'hold off', involves prolonging the level flight path, offsetting the decaying IAS with increasing angle of attack, until the latter is just below the stalling angle, at which point the aircraft is allowed to touch down. The technique of delaying ground contact until this landing angle of attack has been reached results in touchdown occurring at a low speed, so reducing the stresses imposed on the landing gear. The aircraft will contact the ground on its mainwheels because of the high attitude needed to achieve the high angle of attack in the final stages of the hold off. Figure 217 illustrates the situation immediately prior to touchdown, with the aircraft in the landing attitude.



After touchdown, the nose is lowered until the nosewheel contacts the runway. In this way the angle of attack of the wings, and hence the generated lift, is reduced, and so the aircraft will remain firmly on the ground. The brakes are now applied to decelerate the aircraft to taxying speed, so that the machine can be steered off the runway as soon as practicable.

It will be appreciated that use of the technique just described achieves both objectives stated in 4.31.

Throughout the landing manoeuvre the pilot's attention will be wholly concerned with the view ahead. During the flare and hold off he will be constantly assessing both the height of the aircraft above the ground and also its attitude.

4.31.1.1 Control wheel movement

Both the flare and hold off involve rearward movement of the control wheel. (Since the aircraft will have been trimmed in the approach attitude, back pressure will be needed on the control wheel.) More specifically, the rearward movement during the flare is interrupted as level flight is achieved just above the runway. Then, in the hold off, a more gentle movement is required, at a sufficient rate to prevent premature touchdown, but without the aircraft gaining height; a progressively higher attitude will be needed as the speed decays. When the landing attitude, shown in Figure 217, has been attained, the rearward movement of the control wheel is arrested, so that the aircraft then sinks to the runway, contacting the ground on its mainwheels. The control wheel is then gently moved forward to lower the nosewheel on to the runway.

If, at any stage during the flare or hold off, the control wheel is moved rearwards too quickly, the aircraft will start to gain height. This error is called 'ballooning' and is potentially dangerous because, if not corrected, it may precipitate a stall, with the aircraft dropping to the ground heavily. The error is rectified by a slight forward movement of the control wheel. Then, as the aircraft nears the ground again, the rearward control wheel movement is resumed. (In practice, this corrective action is merely a temporary relaxation of the back pressure needed on the control wheel to raise the nose.) Figure 218 illustrates ballooning and its correction.



Conversely, if the control wheel movement is too slow, the aircraft will strike the ground prematurely at high speed. Besides subjecting the landing gear, especially the nosewheel, to excessive strain, this error is likely to result in the aircraft bouncing back into the air again. Such a bounce, if it occurs, should be corrected as for ballooning.

With practice and experience, the pilot will be able to coordinate his control wheel movement so that the ideal landing technique is achieved.

4.31.1.2 Centreline tracking

As the ground is approached during the flare, it is no longer practicable to bank the aircraft, since the lower mainwheel will contact the ground prematurely. Indeed, excessive bank may result in a wing tip striking the ground.

If centreline displacements occur, then, the rudder must be used to make corrections, while the wings are held level with the ailerons. In fact, the pilot reverts to the technique used for take-off. In other words, the rudder is used to steer the aircraft along the centreline during the flare, the hold off and the subsequent landing run along the runway after touchdown. In the latter case, the nosewheel steering will assist the pilot in preventing displacements.

4.31.1.3 Wind

By landing on a runway facing into wind, the aircraft touches down at a lower GS, so reducing both the stresses imposed on the landing gear and also the length of runway required to land and decelerate to taxying speed.

4.31.2 Crosswind landing

If the wind direction does not coincide exactly with the available runway directions, the pilot will be obliged to make a crosswind landing. There are two methods available to accomplish this manoeuvre.

In the first method the pilot counteracts the sideways drifting effect of the crosswind on final approach by adopting a heading slightly 'into wind', so that the aircraft tracks along the extended runway centreline (as in Figure 208). This is called the 'crab' method, the name reflecting the apparent sideways movement of the aircraft relative to the ground, although it should be remembered that the airflow past the machine is symmetrical, assuming that the flight is balanced.

In the second method the pilot sets up a deliberate sideslip so that the aircraft's fuselage is aligned with the runway centreline. For obvious reasons this is called the 'crossed controls' method, and the airflow past the aircraft is not symmetrical.

4.31.2.1 Crab method for crosswind landings

Figure 208 illustrated the case of a crosswind from the right. Figure 219 shows a corresponding view through the windscreen on final approach. Deviations from the centreline are corrected as already described in 4.30.2.2.



The offset heading should be maintained throughout final approach, and during the flare and hold off. Again, the rudder is used to correct centreline deviations once the flare has been initiated.

It will be appreciated that if the aircraft is allowed to touch down on this offset heading, the mainwheel tyres and landing gear legs will be subjected to an undesirable sideways-acting force, because the wheels will not be aligned with the machine's actual direction of motion over the ground.

For this reason, the offset heading should be maintained only until the landing attitude has been reached, at which point the pilot should use the rudder to yaw the aircraft's nose into line with the runway direction, so that touchdown occurs with the machine actually heading in this direction (Figure 220).



This use of rudder is referred to as 'eliminating the drift'. Of course, there will be a simultaneous tendency of the aircraft to roll in the same direction as the application of rudder (refer back to 4.12.3.1), which must be prevented by appropriate use of the ailerons, so that the machine touches down with its wings level.

In the example discussed, with a crosswind from the right, elimination of the drift just prior to touchdown will require application of left rudder; simultaneously, the control wheel should be moved to the right sufficiently to hold the wings level. After touchdown, the pilot should hold the control wheel in this displaced position to prevent any tendency of the right wing to lift during the landing run, while using rudder inputs as required to steer the aircraft along the centreline.

The controls would of course be used in the opposite sense to cope with a crosswind from the left.

A more serious error than failing to eliminate the drift before touchdown is that of doing so prematurely during the hold off, because not only will the aircraft be drifting again when it does eventually contact the ground, but it will also be blown away from the centreline (Figure 221).



4.31.2.2 Crossed controls (sideslip) method for crosswind landings

In our example of a crosswind from the right the pilot applies right bank and left rudder such that the fuselage is aligned with the centreline throughout the approach. Effectively the aircraft's resulting sideslip is exactly nullifying the drift, as shown in Figure 222.



The sideslip is maintained throughout the flare and hold off so that the aircraft eventually touches down firstly on its 'upwind' mainwheel, in this example the right mainwheel. The machine is then allowed to settle onto its 'downwind' mainwheel, followed by its nosewheel, after which the controls are used as described for the 'crab' method during the landing run.

Some pilots combine the two methods just described, starting final approach using the 'crab' method but then switching to the 'crossed controls' method as the aircraft approaches the transition point for the flare.

4.31.3 Crosswind component

The pilot must not land his aircraft in circumstances giving a crosswind component greater than the maximum permitted for his aircraft.
 
4.31.4 Short landing

If it is necessary to minimise the length of runway required for landing and braking to taxying speed, the short landing technique is employed, in which the normal approach path is flown, but a lower threshold speed is aimed for. In this way, the aircraft arrives at the threshold with its wings at the landing angle of attack, so that touchdown occurs almost immediately, without a flare or hold off.

Typically, the short landing threshold speed is about 5 knots above the power-off stalling speed with flaps down. The reader will appreciate that flight in this speed regime demands careful handling of power and attitude if an inadvertent stall is to be avoided.

The short landing technique commences on final approach, with the aircraft initially established on the correct approach path at normal powered final approach speed. The IAS is now reduced, in 5-knot stages, so that the aircraft arrives at the threshold at the short landing threshold speed (specified in the Flight Manual as an IAS).

Each 5-knot reduction in IAS involves the following sequence of actions:

(a) reduce power slightly;

(b) raise the nose to the attitude needed to decay the IAS to the new target figure. (Action (a) is to prevent the aircraft momentarily deviating above the correct approach path as the nose is raised);

(c) increase power again to the setting necessary to continue along the correct approach path at the lower IAS;

(d) trim the aircraft in the new attitude.

Notice that the controls are being used in the normal manner, with IAS controlled by attitude, and rate of descent, and hence descent gradient, by power setting.

At the threshold, a brief increase in power will check the descent, after which the throttle is closed and the aircraft allowed to contact the runway without any further raising of the nose. The nosewheel is immediately lowered onto the runway and maximum braking applied (although the brake application should be reduced if the mainwheels 'lock').

A point to bear in mind is that the attitude during the final stages of the short landing approach is higher than for a normal approach (because of the higher angle of attack of the wings) and the sighting point will probably be below the windscreen. The pilot will have to assess the accuracy of his approach path by reference to the ground adjacent to the threshold. If the approach path is correct, the ground adjacent to the threshold will appear to be level with the sighting point (Figure 223).



When flying at these lower speeds, the pilot must avoid the use of very low power settings, because of the ensuing reduction of margin above the stalling speed and poorer effectiveness of the elevators.

4.32 GLIDE APPROACH AND LANDING

The glide approach and landing is a manoeuvre which the conscientious pilot practises regularly to sharpen his skill in bringing the aircraft safely into land should engine failure occur.

The aim of the practice glide approach is to arrive at the threshold at ground level after initiating a glide from circuit height on base leg.

Relative to the ground, the gradient of descent achieved during the glide depends upon both the aircraft's GS and also its rate of descent. Thus the wind strength will be an important factor to consider. Figure 224 shows an aircraft gliding at minimum gradient IAS both downwind and upwind. Of course, the rate of descent will be identical in both cases. However, since in the upwind glide the aircraft's GS is less than in the other, the gradient of descent is steeper.



Although the pilot will have a pretty good idea of the wind strength, he cannot guarantee that it will remain constant throughout his glide. For this reason, he will aim to arrange that the aircraft has a surplus of height during the glide approach, until the final stages. The height surplus will prevent failure to reach the threshold should the wind turn out to be stronger than anticipated. Once sure of being able to reach the threshold, the pilot can easily get rid of any residual surplus height, whereas if it becomes apparent that the threshold will not be reached, even when gliding at minimum gradient IAS, then there is nothing that the pilot can do to retrieve the situation.

4.32.1 Glide approach technique

The circuit preceding the glide approach is as already described, except that it is usual practice to shorten the downwind leg slightly, because the aircraft's gradient of descent in the latter stages of the glide approach will be steeper than for the normal powered approach (Figure 225).



The pilot maintains cruising speed at circuit height on base leg until the aircraft arrives at a point where he believes that the threshold can be reached, with a surplus of height to spare, without further use of power. He then fully closes the throttle and initiates minimum gradient glide.

If it appears that the surplus of height is becoming excessive, there are three ways in which height can be lost without departure from the minimum gradient gliding speed:

(a) by turning away slightly from the airfield (Figure 226 top image);

(b) by lowering the flaps;

(c) by sideslipping.



In practice, method (a) is chosen since it enables methods (b) and (c) to be held in reserve should more corrective action be needed at a later stage in the approach.

Conversely, if the height surplus seems to be disappearing, the pilot can turn slightly towards the airfield (Figure 226 bottom image).

Ideally, the aircraft should be established on final approach not below 500 feet above airfield level, to give the pilot time to assess and, if necessary, correct his flight path. If the base glide was well-judged, the pilot will find that when he has turned onto final approach, the sighting point will appear to be beyond the threshold, indicating that the threshold will be reached with a surplus of height to spare.

The flaps can now be set to 'maximum lift'. (The attitude should be lowered correspondingly to maintain speed.) The extra drag will steepen the descent path, bringing the sighting point nearer to the threshold.

When convinced that the new flight path will still enable the aircraft to reach the threshold with a surplus of height to spare, the pilot can set the flaps fully down. (Another lowering of attitude will be required to maintain speed.) The descent path will be steepened again. If the flaps have been used correctly, the sighting point will now coincide with the threshold.

Figure 227 shows a side view of the ideal final approach path, with the aircraft eventually arriving at the threshold at ground level.



4.32.2 Mis-judged approaches

Note that the flaps should be used only when it is clearly apparent that the threshold can be reached with a surplus of height to spare. For example, if the pilot turns onto final approach only to find that the sighting point coincides with the threshold, then it is obvious that use of flaps will bring the aircraft down to ground level short of the threshold. In this situation, it may be necessary to leave the flaps up until just before the flare is initiated.

Taking this consideration one stage further, if the sighting point is short of the threshold after the turn onto final approach, then it is impossible to glide to the threshold. This situation is called 'undershooting' (Figure 228).



It should be borne in mind that any attempt to 'stretch the glide' when undershooting by raising the aircraft's nose will only worsen the situation by steepening the descent gradient (because the IAS will drop below that for minimum gradient). The pilot should either revert to a normal powered approach or initiate a go-around.

If, having already lowered the flaps, it becomes clear that an undershoot is developing, there is nothing to stop the pilot raising them again to restore a shallower descent path.

At the other end of the scale, misjudgement may result in a situation where, even with the flaps down, the aircraft will still have a surplus of height as it crosses the threshold. Sideslipping (not permitted on some designs with flaps down) will help to rectify the error, with the pilot reverting to balanced flight just before the flare is initiated. On aircraft without flaps, of course, sideslipping is the only method available for getting rid of surplus height without change of gliding speed.

Assuming that the runway is long enough, it is permissible to continue a mis-judged practice glide approach through to the landing, even though the aircraft may still be too high as it passes the threshold. If any doubt exists about the adequacy of the runway length, however, a go-around should be made.

4.32.3 The flare

The flare should be initiated at such a point that the aircraft decelerates to the threshold speed specified for the glide approach as it arrives at ground level. (For most light aircraft, this speed is 5 knots or so higher than for the powered approach threshold speed to allow for the increased stalling speed during the flare with the engine at minimum power.) Greater control wheel movement will be required than for the powered approach because firstly, the approach path is steeper, necessitating a greater pitch change to achieve level flight, and secondly, the elevators will be less effective, because of the lack of propwash.

4.33 FLAPLESS APPROACH AND LANDING

If for some reason the pilot finds that he is unable to lower the flaps (perhaps because of failure of the activating mechanism) then he will have to make a 'flapless' (in other words, flaps up) approach and landing. The conscientious pilot will practise this manoeuvre occasionally so that he is competent to deal with such a failure should it occur.

Since the aircraft will generate less drag on final approach with flaps up, its descent gradient will be shallower than in the flaps-down powered approach. For this reason, it is usual practice to extend the downwind leg slightly in the preceding circuit (Figure 229).



The remainder of the circuit is flown as already shown in Figure 194. With the reduced drag, the pilot may find that, even having extended the circuit, less power will be needed to achieve a steady descent path to the threshold than for the flaps-down approach. To allow for the increased stalling speed, it is usual to fly the flapless approach 5 knots or so above the normal flaps-down powered approach speed. Again, power controls rate of descent and attitude controls IAS.

On final approach it will be found that a higher attitude is required than for the flaps-down approach, because the wings are at higher angle of attack, and the sighting point will appear lower in the windscreen. Because of the shallower descent gradient the runway aspect will also appear slightly 'flattened'. When the correct approach path is being followed, the sighting point will coincide with the threshold. Figure 230 illustrates.



Compare the view through the windscreen in Figure 230 with that in Figure 210, and refer back to Figure 174 to appreciate the reason for the different sighting points.

4.33.1 The flare and landing

The flapless threshold speed is usually 5 knots or so above that for the flaps-down powered approach, again to allow for the increased stalling speed. Otherwise the flare and hold off are as already described in 4.31.1. When the aircraft has attained the landing attitude, and hence the landing angle of attack, it will be flying faster than if its flaps were down, the higher speed matching the reduced lifting ability of the wings. Thus touchdown will occur at this higher speed.

Furthermore, the deceleration during the hold off will be slower, because of the reduced drag. The two effects combine to increase the length of the runway required to land and brake to taxying speed.

If a flapless landing is made from a glide approach, it is usual to add a further small increment to the threshold speed to allow for the increased stalling speed during the flare with the engine at minimum power.

4.34 THE GO-AROUND

The go-around manoeuvre serves the vital function of enabling the pilot to extricate himself from a poorly executed approach, when the procedure described in 4.23.6 would be carried out. Of course, if a flapless approach is being flown, the pilot can initiate maximum rate climb straight away.

A go-around should also be initiated if the pilot flares prematurely, or balloons or bounces excessively, or otherwise misjudges his landing. In this situation, the procedure is modified slightly:

(a) apply full power;

(b) adjust attitude as necessary to achieve level flight;

(c) when the IAS has increased sufficiently, initiate a climb and complete the procedure as already described.

Having carried out a go-around, the pilot should reposition his aircraft in the circuit pattern for a further attempt at approach and landing.

4.35 THE WINDSOCK

By now, the reader will understand the importance of knowledge of wind direction and strength when taking off or landing.

The windsock is a simple device enabling the pilot to assess the wind. It consists of a cylindrical sleeve of fabric, usually coloured yellow or orange, mounted on a pole adjacent to the runway. Besides indicating the wind direction, it is constructed in such a way that, the stronger the wind, the more the windsock streams away from the vertical. Figure 231 illustrates this.



4.36 STEEP TURNS

In steep turns, the aircraft is made to turn using angles of bank in excess of 30.

4.36.1 Lift increase

The aerodynamic factors involved in medium turns also apply to steep turns, differing only in degree. It has already been explained that it is necessary to increase the lift generated by the wings in turning flight relative to that required for straight flight. With increasing angle of bank, the lift increase necessary to give the required vertical component becomes considerable (Figure 232).



4.36.2 Rate of turn

Figure 232 shows that the greater the angle of bank, the larger is the sideways-acting force arising from the tilting of the lift force, and so the greater is the rate of turn at any particular IAS.

By banking steeply, high rates of turn are attainable. The practical value of steep turns, then, is to enable the aircraft to change direction more quickly than if medium angles of bank are employed.

4.36.3 Increase of stalling speed

As the angle of bank increases, so does the aircraft's stalling speed (refer back to 4.24.4). At steep bank angles the increase becomes considerable. For an aircraft which has a stalling speed of 50 knots IAS in straight flight, the stalling speed at various angles of bank is as stated in Figure 233.



4.36.4 Steep level turns

It has been seen that, in turning flight, the increased lift is generated by increasing the angle of attack of the wings. In medium level turns, the extra wing drag results in a slightly lower IAS, but the margin above the stalling speed is still comfortably adequate.

As the angle of bank is steepened, however, the IAS will decrease, and the stalling speed will increase, until finally the two coincide, with the wings at the stalling angle of attack. In other words, the aircraft will stall.

To generate the increased lift at these steeper bank angles without bringing the wings to the stalling angle of attack, it is therefore necessary to prevent the IAS decay. In practice, extra power is used to counteract the increased wing drag. Not surprisingly, it will be found that, the greater the chosen angle of bank, the greater will be the power setting needed to maintain speed.

So far as maintaining level flight is concerned, the pilot will find that greater back pressure is required on the control wheel to hold the correct pitch attitude, compared to turns with shallower angles of bank.

In summary, the steep level turn is entered as for the medium turn, except that power is simultaneously increased sufficiently to maintain IAS.

During the turn, elevators are used as necessary to maintain level flight, ailerons to prevent any tendency of the bank angle to change, rudder to balance the aircraft and power to control IAS.

Figure 234 shows typical attitudes for level turns using 45 bank angles. Notice that the attitude in the turn to the right appears lower than in that to the left, because the pilot's seating position is on the left-hand side of the cabin.



Compared to the medium turn, greater anticipation is required when reverting to straight flight on a chosen heading. The steep turn is completed as for the medium turn, except the power is simultaneously reduced to the setting in use before the turn was entered, so that the IAS remains constant.

4.36.4.1 Stall avoidance

The reader will appreciate that in steep turns the aircraft is being flown in an angle of attack regime close to the stalling angle, especially at higher angles of bank.

Suppose that momentarily the pilot accidentally relaxes the back pressure on the control wheel. The nose will drop and so the aircraft will being to descend. The pilot, observing the descent from consultation of his flight instruments (altimeter and VSI) will apply extra back pressure to try to raise the nose back to the correct attitude for level flight. This action might well precipitate a stall, with the nose failing to rise despite the rearward movement of the control wheel.

The problem is overcome by reducing the angle of bank, and thus reducing the stalling speed, prior to any attempt to raise the nose. As soon as level flight has been restored the bank can be increased again as desired.

If, during a steep turn, the pilot becomes aware of pre-stall symptoms, such as activation of the stall-warning device, forward movement of the control wheel (in practice, relaxation of the back pressure) will lower the angle of attack of the wings and so avert the stall. When satisfied that he has full control of the aircraft, the pilot can then make a further attempt to turn, either increasing power or reducing the angle of bank to prevent recurrence of the incipient stall.

If the aircraft actually stalls, as indicated by sudden dropping of the nose, recovery action should be taken without delay as described in 4.25.4. (Note that a wing drop at the stall may either increase or decrease the angle of bank.)

After taking the precautions described in 4.25.5, deliberate stalling in steep turns will enable the pilot to maintain proficiency with the recovery action.

4.36.4.2 Recovery from spiral dive

Failure to apply sufficient back pressure on the control wheel to maintain level flight in an attempted steep turn will probably result in a spiral dive, with the aircraft descending and its IAS increasing. This error is most likely to occur when high angles of bank are in use.

Recovery from the spiral dive is as described in 4.25.10.

4.36.4.3 Maximum rate turns

We have seen that in a medium turn, it is unnecessary to increase power to overcome the extra wing drag since only a small reduction in IAS is involved. It has also been shown that if the bank angle is increased, the IAS will decrease, and the stalling speed will increase, until the two eventually coincide and the aircraft stalls. The angle of bank at which the wings are just below the stalling angle of attack, then, is the maximum sustainable in level flight at that particular power setting.

If now the engine power setting is increased, this angle of bank can be sustained at higher IAS, in other words, with the wings at lower angle of attack. Taking the argument one step further, at this new power setting we can increase the angle of bank, without stall occurrence, until a situation is again reached in which the extra drag decelerates the aircraft to stalling speed, which will now be greater.

Thus the maximum angle of bank sustainable in level flight depends on power setting. To clarify the point, Figure 235 illustrates the case with the engine at cruise power and at full power.



We can conclude that angle of bank (1) is the maximum sustainable in level flight with the engine at cruise power, and angle of bank (2) is that sustainable at full power. In other words, sustained level turns are impossible at greater angles of bank, since no more power is available.

The diagram also shows that cruising speed can be maintained by increasing power up to angle of bank (3), when full power is needed. Greater angles of bank will entail reduction of speed.

Since we know that rate of turn depends upon angle of bank, it will be appreciated that the maximum achievable rate of turn in level flight occurs at the maximum sustainable angle of bank, with the engine at full power and the wings just below the stalling angle.

In practice, maximum rate turns are flown by applying full power and increasing the bank angle, simultaneously increasing the rearward movement of the control wheel to maintain level flight, until the stall-warning device activates, or until pre-stall buffeting is detected. This state of affairs is maintained until it is desired to complete the turn. Considerable back pressure will be required to hold the control wheel in the correct position during the turn. If the aircraft stalls, recovery action should be taken without delay. (The engine will already be at full power, of course.)

Most pilots practise maximum rate turns occasionally to sharpen their coordination in the use of aircraft's controls.

4.36.5 Steep gliding turns

If a steep turn were to be attempted whilst gliding at minimum gradient IAS, it is quite likely that at this speed the wings would reach the stalling angle. An increase in IAS is therefore required, and the only method available to the pilot of doing so when gliding is by the adoption of a lower pitch attitude.

As a rule of thumb the IAS should be increased by 5 knots for every 10 bank (or fraction of 10) over 30. Thus a 45 banked gliding turn should be flown at minimum gradient speed plus 10 knots.

To attain these higher speeds against the opposing effect of the increased wing drag, quite steep nose-down pitch attitudes will be necessary. With the nose so far below the horizon, accuracy of speed control will require considerable skill and practice. The increased drag will also result in steep decent gradients, and hence high rates of descent, which the pilot should bear in mind. Because of these factors, it is not usual practice to carry out gliding turns at angles of bank in excess of 45. Figure 236 demonstrates a typical attitude for a 45 banked gliding turn to the left.



Steep angles of bank in gliding flight might be required when bringing the aircraft into land after engine failure. Pilots therefore practise the manoeuvre occasionally to maintain competence.

4.37 REVIEW OF CONTROL FUNCTIONS

At this stage, it is worth reviewing the functions of the aircraft's most important controls in various phases of flight.

4.37.1 Function of elevators

The elevators are used to maintain or change the pitch attitude. The pitch attitude is chosen  to maintain either:

(a) a defined vertical flight path, such as level flight or the landing flare (and, optionally, final approach);

(b) a desired IAS, such as in climbing or descending flight.

4.37.2 Function of ailerons

The ailerons are used to select and hold any chosen angle of bank, and to remove the bank when so desired. However, they should not be used to attempt to rectify wing drop in a stall, nor during spin recovery.

4.37.3 Function of rudder

The primary function of the rudder is to balance the aircraft's flight. In specific phases of flight, it has different functions:

(a) to prevent bank-induced yaw when deliberately side-slipping;

(b) to maintain constancy of heading during stall recovery and so prevent autorotation;

(c) to oppose yaw during spin recovery;

(d) to steer the aircraft during take-off and landing (flare, hold off and landing run).

4.37.4 Function of power

When the pitch attitude is controlling the vertical flight path, power adjustments are used to control IAS. When climbing at maximum rate or maximum gradient, the engine is set to climb power (maximum if permitted). In descending flight, power is used to control rate, and hence gradient, of descent.

4.37.5 Flight near stalling regime

During flight the pilot should aim to keep his IAS comfortably above the stalling speed (and hence the angle of attack of the wings safely below the stalling angle). However, deliberate flight near the stalling regime is necessary:

(a) when stalling is intentional;

(b) during the hold off whilst landing;

(c) during the approach for a short landing;

(d) in steep turns.

4.38 INSTRUMENT FLIGHT

Instrument flight is a technique in which the pilot assesses the aircraft's behaviour entirely by reference to the flight instruments. This technique is used when poor visibility (such as when flying in cloud) obscures the view of the outside world to the extent that the attitude of the aircraft cannot be determined.

The discussion which follows is of a general nature, stressing the most important considerations which apply to instrument flight. However, it is beyond the scope of this book to make a detailed examination of all the factors involved.

4.38.1 Physiological aspects

On the ground, man's assessment of his orientation with regard to the vertical derives from the brain's interpretation of information from three sources:

(a) eyesight;

(b) the balance organs in the ears;

(c) various nerves throughout the body (for example, those in the feet when standing up).

The brain gives priority to visual information when this is available. If vision is denied (as when in total darkness) the two secondary sources supply sufficient information for the brain to assess the body's orientation.

In flight, the nature of the forces acting on the aircraft and its pilot differ from those experienced by the latter when he is groundborne. For this reason, the brain's interpretation of information from the balance organs and nerves may be erroneous.

Using the technique of visual flight, this shortcoming poses no problems, since sight of the outside world is available and the brain naturally gives priority to this information, which in any case is less likely to be misinterpreted.

If, however, the pilot is denied sight of the outside world, as when flying in cloud, then the secondary sources alone supply information, which the brain is likely to misinterpret. For example, the pilot may gain the impression that his wings are level when in reality the aircraft is in a banked turn. Conversely, the situation may arise where the wings are indeed level, but the pilot feels that the aircraft is in a banked attitude.

It will be appreciated, then, that it is impossible for the pilot to control the aircraft unless he has information about attitude which the brain will interpret correctly. We can see that the two pre-requisites for correct assessment of the aircraft's attitude are:

(a) reliable indications of attitude;

(b) correct interpretation of such indications by the brain.

When the outside world cannot be seen, the first requirement is fulfilled solely by the flight instruments. Interpretation of their indications would present no problems, except that the brain has difficulty in disregarding information from the balance organs and nerves. In other words, the pilot may come to disbelieve the flight instruments because he has the feeling that the aircraft's attitude is different from that indicated. The pilot cannot hope to retain control of his aircraft unless he accepts that the instrument indications are correct, and that any sensations he may have which are at variance with these indications must be disregarded. Perhaps the most difficult aspect of instrument flight so far as the pilot is concerned is learning to believe the instruments regardless of conflicting sensations that may be experienced.

It is worth pointing out here that for a pilot who has not been trained in instrument flight technique, the consequences of entering cloud could be disastrous. Control of the aircraft will be lost very quickly, because the pilot will instinctively react to his sensations rather than to the flight instrument indications of the aircraft's attitude.

4.38.2 Instrument scan

In view of the foregoing remarks, it will be appreciated that a pilot flying in cloud will need to devote nearly all of his attention to the flight instruments. Specifically, the AI should be consulted most frequently, perhaps every two seconds or so, so that the pilot has a continuous idea of the aircraft's attitude. In between such consultations, the remaining instruments should be read, both to cross check the AI's information and to assess the aircraft's performance. This technique is called 'scanning' the instruments (Figure 237).



Note that the aircraft's pitch attitude is indicated directly by the AI and indirectly by the ASI, altimeter and VSI. Of course, accurate knowledge of the engine power setting is also needed to infer pitch attitude from these last three instruments. For example, a particular IAS implies a higher attitude when the engine is at high power than when low power is set. Thus the tachometer will need to be referred to occasionally.

Bank is indicated directly by the AI and indirectly by the DI and TBI (both of which respond to the turning effect arising from a banked attitude).

4.38.3 Control technique

So far as use of the aircraft's controls is concerned, there is no difference in technique between visual flight and instrument flight. The AI replaces the earth's horizon for assessment of attitude.

If the flight instruments show that the aircraft's behaviour differs from that required, suitable alterations of power and attitude are made to rectify the error. When making a power correction, the pilot will refer to the tachometer, so that he can make an accurate assessment of the power setting in use. Similarly, he will refer to the AI when adjusting attitude. As in visual flight, power and attitude are set in a combination which the pilot believes will give the desired behaviour. If the results are not exactly as intended, further adjustments will be required. Having made corrective inputs, the pilot must wait for the instrument indications to stabilise before applying further corrections, allowing for the aircraft's inertia. He must avoid 'chasing the needles' - in other words, over-controlling. Care must also be taken not to become over-attentive to one instrument at the expense of the others. The pilot must discipline himself to scan all the instruments in a systematic manner, so that he has continuous knowledge of the aircraft's overall behaviour.

The pilot's workload during instrument flight is greater than when visual flight technique is in use. This is mainly because in the former case a conscious mental effort is needed to interpret the aircraft's attitude from the instrument indications, whereas in the latter the earth's natural horizon is to be seen all around the aircraft, and attitude is assessed subconsciously, so using less of the pilot's mental capacity. Of course, there are also other demands on the pilot's attention, such as monitoring the engine gauges and operating the carburettor heat control periodically to ensure that the carburettor is free of ice.

4.38.4 Navigation

In visual flight, navigation is mainly by reference to landmarks on the ground, correlating such features to those marked on a navigational chart. If the aircraft features GPS equipment, the GPS map display can offer confirmation of geographic location and track (actual path over the ground or sea).

When flying in cloud, sight of the ground is not available to the pilot, and GPS or alternative means must be used to determine the aircraft's geographic location. Alternatives include:
 
(a) by interpretation of information from ground-based radio facilities;

(b) with the assistance of ground-based personnel manning radio equipment designed for this purpose.

Method (a) requires the aircraft to be fitted with appropriate navigational radio equipment and the pilot to be knowledgeable in the use of such equipment and in interpretation of the information derived from it.

Method (b) is the only choice available to pilots of aircraft equipped solely with COM radio. Depending on their type of radio installation, the ground personnel will be able to advise the pilot either of the aircraft's direction, or bearing, relative to the installation, or of its actual position (in terms of bearing and distance from a chosen reference). The former type of radio installation is known as VDF (VHF direction-finding equipment), and the latter as 'radar' (radio detection and ranging).

4.38.5 Terrain clearance

Denied sight of the ground, the pilot will have to ensure at all times that his chosen altitude is safely above any terrain that the aircraft may overfly. (Of course, knowledge of the elevation of such terrain will be necessary.) Again, when descending in cloud, a flight path must be chosen in which the pilot is positive that terrain will be safely cleared. This requirement implies the necessity of knowing the aircraft's exact position throughout the descent. One of the navigation methods described in 4.38.4 will be needed.

Similar considerations apply when approaching to land. Each such approach has an associated decision altitude, down to which the pilot may safely fly in cloud. At decision altitude, if the ground cannot be seen, the pilot must make a go-around, following a previously specified flight path. He then has the choice of either making another approach attempt (if he believes that the weather conditions will be such that sight of the ground will be regained by the time decision altitude is reached) or flying to a location where better weather conditions are more likely to result in a successful approach. Having regained visual contact with the ground, the pilot can position his aircraft for landing.

4.38.6 Collision avoidance

Flying in cloud, a pilot will not be able to see conflicting traffic, because not only is the visibility from the cabin severely restricted (effectively the aircraft will be in fog), but the pilot's attention will be devoted mainly to the flight instruments.

To prevent collisions between aircraft flying in cloud, their pilots are obliged to conduct their flights in accordance with a special set of regulations, called Instrument Flight Rules (IFR), whose purpose is to ensure adequate separation between such traffic. A detailed description of IFR is beyond the scope of this book; suffice it to say that their purpose is achieved by imposition of certain restrictions on aircraft flying in weather conditions which are unsuitable for visual flight. In addition, traffic separation can be positively implemented with the assistance of ground-based personnel equipped with radar, which enables them to determine the positions of all aircraft in the vicinity of the installation.

A recent development in avionics is equipment showing nearby aerial traffic on aircraft GPS map displays. Besides relative position, this equipment shows relative altitude of the traffic so that pilots may assess the three-dimensional separation of their aircraft from this traffic. The equipment requires aircraft to be fitted with transponders (see 2.8.1) incorporating the appropriate avionics. This equipment will, of course, also be beneficial to pilots operating under visual conditions, supplementing their lookout.

4.38.7 Meteorological aspects

Cloud consists of minute particles of water, which have no detrimental effect on fight unless the air temperature is at or below freezing (0C). An aircraft flying through cloud in the latter conditions is likely to accumulate ice on the forward facing areas of its airframe as it meets these particles. The dangers of ice accumulation on aircraft are significant but their description is beyond the scope of this book.

4.39 CONSTANT-SPEED PROPELLERS

In 3.10.4.1 it was explained that a fixed-pitch propeller cannot function efficiently in all phases of flight, and that for small aircraft designs it is usual to fit a propeller whose blade angle gives optimum efficiency in cruising flight.

For larger aircraft, the impaired performance resulting from propeller inefficiency during particular phases of flight is unacceptable. Such machines require a propeller whose blades can be turned to a lower angle (fine pitch) for use at high engine power settings and low forward speed, for example during take-off and climb, and then turned to a higher angle (coarse pitch) for cruising flight. In other words, a variable-pitch propeller is required (Figure 238).



Ideally, the pilot should be able to adjust the pitch so that the propeller blades always meet the airflow at the most efficient angle of attack. However, there is no means by which he can assess the blades' angle of attack; instead, a constant-speed governor is fitted to control the pitch. The whole assembly is now termed a constant-speed propeller, and is remotely controlled by the pilot's propeller control lever (sometimes referred to as the pitch lever). The pilot uses the propeller control lever to set the desired speed of rotation of the propeller (and therefore of the engine) according to the phase of flight.

4.39.1 Theoretical considerations

The constant-speed governor compares the actual engine RPM with that demanded by the setting of the propeller control lever. The governor eradicates any discrepancy between the two by altering the pitch. Once the desired RPM requirement has been set with the lever, any alteration of circumstances which would tend to increase the RPM (for example, opening of the throttle or increase of forward speed) is countered by the governor, which coarsens the pitch (increases the blade angle). The effect is to increase the angle of attack of the blades, and therefore the torque, which retards the engine speed to the original RPM.

An opposite sequence of events occurs if any alteration of circumstances attempts to reduce the engine RPM.

Note that the governor can adjust the pitch only within the mechanical limits of the pitch-changing mechanism in the propeller hub, the limits being termed the fine pitch stop (at the lowest blade angle) and the coarse pitch stop (at the highest angle). For example, if the blades are at the fine pitch stop, any reduction of engine RPM cannot be countered - the behaviour will be the same as that of a fixed-pitch propeller.

4.39.2 Technical description

The pitch-changing mechanism forms part of the propeller hub assembly and is powered by a special supply of oil from the engine's main oil system. Oil is piped to the hub (from the hollow crankshaft) and its pressure is used to activate the mechanism as directed by the setting of the propeller control lever.

The propeller control lever is sited adjacently to the other engine controls in the cabin. Figure 239 shows a typical arrangement. Notice that the control knobs have different shapes (and usually different colours) to assist the pilot in their identification.



When the propeller is acting in the constant-speed range (that is, with its blades in between the fine pitch and coarse pitch stops), the pilot can increase RPM by setting the lever to a more forward position and vice versa. Thus maximum RPM are attained when the lever is fully forwards.

4.39.3 The manifold pressure indicator

Opening or closing the throttle will have no effect on engine RPM if the propeller is acting in the constant-speed range. Hence, in addition to the tachometer, aircraft featuring constant-speed propellers are equipped with a manifold pressure indicator (MPI) for assessment of engine power output. The MPI is mounted on the instrument panel (usually next to the tachometer) and shows the pressure of the fuel-air mixture in the inlet manifold. A typical presentation is illustrated in Figure 240.



The MPI is usually calibrated in inches of mercury (in Hg), these being units of pressure. As the throttle is opened it allows more mixture to enter the inlet manifold for combustion in the cylinders, and so the reader will appreciate that a higher indication on the MPI signifies a higher power output.

4.39.4 Operating technique

The basic operating technique for constant-speed propellers is to set the recommended manifold pressure setting (with the throttle lever) and the engine RPM setting (with the propeller control lever) appropriate to the phase of flight. These recommended settings are to be found in the Flight Manual. Correct use of the controls in this respect will ensure that the propeller blades are working at or near their most efficient angle of attack throughout the various phases of flight.

Departure from the recommended settings will result in the propeller blades working at other than the most efficient angle of attack. Usually, no harm will be done, except in the case where low engine RPM are set with the propeller control lever and high manifold pressure with the throttle lever. The consequence of such action is to make it difficult for the relatively slow-moving pistons to absorb the large pressure increases in the cylinders during combustion of the mixture. Most engines are not designed to withstand this misuse of the controls. (A motoring analogy is putting one's foot down on the accelerator pedal while the car is moving at low speed in high gear.)

Note that, for any particular manifold pressure setting, selection of higher RPM results in increased power output from the engine (and hence increased thrust from the propeller). Thus when maximum power is required from the engine (as during take-off), the pilot sets both throttle and propeller control levers fully forward.

For an aircraft fitted with a fixed-pitch propeller, use of the mixture control lever to lean out the mixture for cruising flight is as explained in 4.2.2. Of course, this procedure is unsuitable for an engine driving a constant-speed propeller - the governor would merely reduce the blade angle to maintain RPM until the blades were at the fine pitch stop, by which stage the mixture would be over-lean and the power output from the engine would be considerably reduced.

As the mixture is leaned out, the rate at which fuel is consumed by the engine is reduced. At the same time, the temperature of combustion in the cylinders (and hence the temperature of the exhaust gases) increases, because there is no surplus fuel to cool the mixture. A gauge which measures rate of fuel consumption can therefore be used for calibration of mixture strength, as can one which indicates exhaust gas temperature (EGT). The former is often referred to as a fuel flow meter.

An aircraft fitted with a constant-speed propeller will most likely feature either a fuel flow meter or an EGT gauge, or both, on its instrument panel. The procedure for leaning out the mixture is then as follows:

(a) use the throttle lever to set recommended manifold pressure;

(b) use the propeller control lever to set recommended RPM;

(c) move the mixture control lever backwards to achieve the recommended indication on the fuel flow meter or EGT gauge.

In some aero-engines fitted with FADEC systems (see 2.2.18) the FADEC computers also control propeller pitch to ensure optimum operational efficiency in all phases of flight. In aircraft so equipped there will be no need for the propeller control lever.

4.40 RETRACTABLE LANDING GEAR

A retractable landing gear is raised as soon as the aircraft is safely airborne on take-off (after a brief application of brakes to stop the mainwheels spinning). Lowering the landing gear prior to landing usually features as one of the items in the pre-landing checks.

Indicator lights on the instrument panel show the position of the landing gear legs, enabling the pilot to confirm that the three legs do indeed lock themselves in the position demanded by the selector.

The Flight Manual will probably specify the maximum permitted IAS for retraction and for lowering of the landing gear (the two speeds may differ). The pilot should not violate such limiting speeds, otherwise damage might be caused to the retraction mechanism or landing gear doors by the airflow.

When descending, lowering the landing gear also serves as a useful means of steepening the gradient (by increasing drag) if so desired.

   





5   EMERGENCY HANDLING






Serious emergency situations are uncommon in aviation, because of the careful regulation of flight procedures and the high standard of reliability of aircraft and aviation ground equipment.

Nevertheless, the pilot must be able to cope competently with such a situation should it arise, primarily to safeguard the lives of his passengers and himself. A satisfactory degree of competence in dealing with emergencies can only be assured by appreciation of the various factors involved and by frequent practice of the relevant procedures; in this way the pilot will not be caught unawares should unforeseen difficulties occur.

It is worth noting that there is usually sufficient time to take stock of the situation and plan a suitable course of action. Furthermore, the familiarity with emergency procedures gained by frequent practice will assist the pilot in making sensible decisions promptly, and will minimise the erosion of judgement that naturally occurs when anxiety besets the pilot.

An important consideration in dealing with emergency situations is sensible prioritisation of tasks. Retaining control of the aircraft and ensuring it is following a safe flight path at a safe speed is the first priority, followed by the appropriate actions or checklist(s).  A review of navigational factors may be required - can the aircraft continue along the original intended route or are changes necessary? In the following discussions, reference is occasionally made to transmission of distress calls on the radio to alert ground personnel to the emergency situation and thereby obtain their assistance or advice. Such radio calls should not be made before the more essential actions have been completed.

A useful mantra summarising prioritisation is : aviate, navigate, communicate.

5.1 ENGINE FAILURE

Modern aero-engines are extremely reliable, and malfunctions are more likely to occur as a result of pilot mismanagement than because of actual mechanical failure. Bearing this in mind, the pilot should take the following precautions before and during every flight:

(a) ensure that sufficient fuel and oil are carried for the intended flight, with adequate reserves;

(b) monitor the fuel quantity gauges frequently and ensure that no tank is allowed to run dry while it is supplying fuel to the engine;

(c) monitor the engine gauges frequently and terminate the flight as soon as possible if abnormal indications are shown;

(d) operate the carburettor heat control periodically to ensure that the carburettor remains free from ice;

(e) avoid mis-setting of the mixture control.

For reasons which will soon become clear, other factors which should be borne in mind throughout any flight are the wind direction (relative to the aircraft's heading) and the elevation of the terrain over which the aircraft is flying.

Should total engine failure occur, the aircraft will not be able to maintain height, and the pilot will be obliged to glide down and make a forced landing. If the aircraft were within comfortable gliding range of an airfield, the pilot would obviously choose to land there. However, it is more likely that the machine will not be able to reach the nearest airfield, in which case the pilot must select the best field within comfortable gliding range for the forced landing.

If the failure is only partial, it might be possible for the pilot to reach the nearest airfield, using whatever power is available from the engine and flying the aircraft at its minimum gradient gliding IAS to achieve the shallowest possible descent path. However, it should be borne in mind that a partial failure might develop into a total failure. It would be unwise, then, to attempt to reach an airfield if flight over terrain unsuitable for a forced landing were involved. In this case, the safest course of action would be to close the throttle and initiate the forced landing procedure as for total failure, remembering that the residual power available from the engine can be used should it be needed to reduce rate of height loss. The pilot should not rely on this residual power, however, since it will not be available if the failure becomes total.

If the failure is accompanied by severe vibration, signifying mechanical defect, the engine should be shut down immediately to prevent it from being wrenched from its mountings.

If the cause of the failure is not known, then after commencing the forced landing procedure the pilot should look for the cause and attempt to rectify the failure. Should he be successful in this, then the forced landing procedure can be discontinued and cruising flight resumed.

Figure 241 summarises the courses of action to be taken following engine failure.



5.1.1 Forced landing procedure

When in cruising flight, the pilot should always be aware of the wind direction relative to his aircraft's heading, to facilitate directional orientation when it comes to carrying out forced landing procedure after engine failure. Often, there are visible indications of wind direction, such as smoke from chimneys, or cloud shadows on the ground, to facilitate orientation. In the absence of such indications, it is worth remembering that if the aircraft is not greatly distant from the airfield of departure, the wind direction will probably not differ very much from the direction in which the take-off was made.

To illustrate the forced landing procedure, we shall consider the case where total engine failure occurs for reasons unknown.

The amount of time available to deal with the situation depends on the height of the aircraft above ground level when the failure occurs - the greater the height the more time there will be for dealing with the emergency. To avoid wasting whatever time is available, it is essential that a sensible order of priorities is given to the various actions that the pilot will take. For example, it would be unwise to concentrate on trying to ascertain and rectify the cause of failure, if meanwhile the aircraft were heading towards terrain unsuitable for a forced landing, for if the failure cannot be rectified, the pilot will then have to cope with landing on this terrain.

Following total engine failure, then, the pilot must carry out the actions listed below without delay, in the order given:

(a) establish a glide at minimum gradient IAS, to achieve the greatest gliding range should it be needed. Trim the aircraft in the appropriate attitude;

(b) choose the most suitable field within comfortable gliding range for the forced landing;

(c) choose the landing direction;

(d) plan a descent path which will enable a successful landing to be carried out in the chosen field in the chosen landing direction.

These four actions are essential to the safeguarding of the aircraft and its occupants. If time permits, the pilot can allocate some of his attention during the glide to secondary actions. In order of priority, these are:

(e) look for the cause of the failure and attempt rectification;

(f) if the failure cannot be rectified, transmit a distress call;

(g) prior to landing, carry out the impact checks (described later).

Let us now examine these actions in detail.

5.1.1.1 Choice of most suitable field

In our discussion of the glide approach in 4.32, it was explained that the pilot should aim to have a surplus of height until the final stages of the approach, to guard against undershooting resulting from misjudgement of the glide. It was pointed out that, whilst getting rid of surplus height presents no problems, undershooting cannot be countered.

The reasoning applies equally to the forced landing descent path. Thus it is better to choose a field around which a circuit (either left- or right-handed) can be flown, rather than one that demands a long, straight glide. Besides the advantage of giving the pilot more flexibility in his control of height loss, this procedure also enables the chosen field to be more closely inspected for suitability. At the cruising heights normally suited to light aircraft, these considerations may dictate choice of a field downwind of the aircraft's position, since the pilot will aim to land facing into wind, or approximately so.

Figure 242 illustrates these points and shows some of the alternatives available in the descent path, demonstrating that a smaller circuit can be flown if the height surplus appears to be diminishing, and vice versa.



The chosen field should fulfil as many as possible of the following criteria:

(a) it should be as large as possible, both in width and length;

(b) the surface should be free from buildings, livestock or similar obstructions;

(c) the surface should be level and even, preferably green in colour (signifying grass or young crops) rather than brown (ploughed land) or yellow (ripe crops);

(d) there should be no high obstructions (such as electricity cables and pylons) near the field;

(e) in sparsely populated regions, the chosen field should be near to a feature of civilisation, such as a road or settlement, so that assistance can be readily summoned after landing, if necessary.

A useful mnemonic to help remember the criteria described above is the 'five esses':

size
shape
surface
slope
surroundings

Of course, it is probable that the aircraft will not be able to reach a field satisfying all these criteria. Common sense will dictate the selection in the circumstances prevailing. Thus in most cases it would be better to aim to land in a large ploughed field than a small grass-covered one, since a large field gives the pilot greater chance of landing within its confines should he misjudge the descent path.

Again, it would be better to choose a collection of small fields separated from each other by low fences or walls (leaving the choice of individual field until a later stage in the descent), rather than a larger field completely surrounded by buildings or trees.

It is worth noting that nearly all fields in the open country are smaller in dimension than the runways at most airfields (Figure 243).


 
5.1.1.2 Choice of landing direction

Preferably, the landing direction should be into wind, as already stated, to minimise both the groundspeed on touchdown and the length of the landing run. However, if a considerably longer landing path is available by landing crosswind, then it would be logical to choose this alternative.

5.1.1.3 Planning descent path

As Figure 242 shows, the pilot should plan to fly a descent path such that the aircraft arrives at the start of base leg at about 1000 feet above ground level, with the base leg spaced at about the same distance from the landing path 'threshold' as for a practice glide approach at an airfield. (Knowledge of the terrain elevation will assist in assessment of height above ground level when the altimeter is set to QNH.)

Whether a left- or right-hand circuit should be flown depends primarily on the height of the aircraft above ground level and its orientation relative to the field, bearing in mind that the intention is to arrive at the start of base leg at about 1000 feet. In the situation shown in Figure 242 a right-hand circuit would be chosen if there were insufficient height for a left-hand circuit to be flown (Figure 244).



5.1.1.4 Attempting to rectify failure

Note that the propeller will probably continue to windmill, driven round by the airflow, even after total engine failure. If this is the case, no attempt should be made to use the engine starter, since serious damage would probably be caused to both engine and starter.

Instead, the pilot should carry out the specified cause-of-failure checks to try to restore power. As previously stated, engine failure is most likely to occur as a result of pilot mismanagement. Examples of mismanagement are allowing the fuel tank supplying the engine to run dry, failure to enrich the mixture after descent from cruising height and inadequate use of the carburettor heat control to prevent ice formation in the carburettor. The cause-of-failure checks are designed primarily to rectify such mistakes. Typically, they include the following actions:

(a) examine the fuel cock selection and fuel quantity gauges. If the aircraft has more than one tank, change the fuel cock selection to another tank containing fuel. This action will restore power if the cause of the failure was a blockage in the pipe line from the tank originally selected;

(b) switch on the electric fuel pump (if fitted), in case the engine-driven pump has failed;

(c) set the mixture control fully rich;

(d) set the carburettor heat control 'on';

(e) check that both magnetos are switched on.

If the propeller has stopped, and the engine failure is not because of obvious mechanical defect, it would be logical to use the starter to try to re-start the engine after carrying out the cause-of-failure checks.

5.1.1.5 Impact checks

If the aircraft is damaged on landing to the extent that fuel tanks or pipe lines rupture, there will be a risk of fire from the spilt fuel.

To minimise this hazard, the impact checks should be carried out prior to landing. The checks are designed primarily to render the aircraft and its components electrically 'dead', so that no sparks can occur to ignite spilt fuel.

Two further items included in the impact checks are, firstly, ensuring that the pilot and his passengers have their harnesses fastened tightly, and secondly, unlatching the cabin doors to facilitate evacuation after landing in the event of distortion damage to the fuselage.

A typical sequence for the impact checks is:

(a) select fuel cock off;

(b) switch off magnetos;

(c) switch off master switch;

(d) check harnesses tight;

(e) unlatch cabin doors.

Of course, if the engine is still capable of producing residual power, it would be logical to delay items (a) and (b) in the impact checks until the last moment, in case the pilot finds that he has to use this power to reduce the rate of height loss after misjudging the descent path.

5.1.1.6 Flying the descent path

As already stated, a descent path should be flown which positions the aircraft at the start of base leg at about 1000 feet above ground level.

If it appears that the height surplus will be excessive at the start of base leg (in other words, if the aircraft is going to be much above 1000 feet), the pilot should not correct the error by extending the downwind leg greatly, since it may subsequently prove impossible to reach the chosen field when the aircraft is turned back into wind on final approach. (This is because the gradient of descent is steeper when gliding upwind than when gliding downwind.) Instead, the aircraft should be positioned onto base leg and then turned slightly away from the field. If necessary, the height surplus can be further reduced by delaying the turn onto final approach, as shown in Figure 242.
 
Conversely, if the height surplus is disappearing, the aircraft can be turned slightly towards the chosen field and, if necessary, turned onto final approach earlier, again as shown in Figure 242.

On final approach, the height surplus will result in the sighting point appearing to be beyond the landing path 'threshold'. Once sure of being able to reach the 'threshold' with a surplus of height to spare, the flaps can be used (as was described in 4.32.1) to reduce the height surplus by steepening the descent path.

Note that so far as use of the flaps is concerned, it is better to lower them too late rather than too early. Far less harm will be done if the aircraft runs into the field's far boundary at low speed at the end of its landing run as a result of touching down well beyond the 'threshold' than if it runs into the near boundary at high speed as a result of undershooting.

The flare technique is as already described in 4.32.3. If the landing path is restricted in length, application of brakes may be necessary after touchdown to bring the aircraft to a halt within the confines of the field. The machine should be steered clear of any obstructions in its path during the landing run.

(In an aircraft equipped with retractable landing gear, it is advisable to delay lowering the landing gear until the pilot is sure that the chosen field can be reached, bearing in mind that the extra drag will steepen the descent path.)

5.1.1.7 Overrunning

Even with maximum application of brakes, it may well be that the chosen field is not long enough to bring the aircraft to a halt before it arrives at the far boundary. One way of trying to avoid overrunning is to increase the length of the available landing path by steering the aircraft to the left or right as the boundary is approached (Figure 245).



If not already done, the first three items of the impact checks listed in 5.1.1.5 should be carried out after the aircraft has come to a halt if it has collided with the boundary (fence, hedge or wall) or any other obstruction during its landing run.

5.1.1.8 Misjudged approaches

If it is clear during base leg or final approach that the aircraft is undershooting and will not reach the chosen field, no attempt should be made to 'stretch the glide' by raising the nose to a higher attitude, as this will have just the opposite effect by steepening the descent gradient. Instead, the best alternative field within gliding range should be chosen for the landing.

Conversely, the situation might arise where it is obvious to the pilot that the height surplus has become so great that, even resorting to the tactics of turning away from the field on base leg and delaying the turn onto final approach, the aircraft will eventually touch down too far beyond the landing path 'threshold', so reducing the length of landing path available and increasing the danger of overrunning. In this case, it would be logical to lower the flaps early or sideslip the aircraft (or both, if permitted for the particular design of aircraft) to try to minimise the error. In addition, the descent path can be steepened by increasing the gliding speed (by lowering the nose), although this advantage will be offset to a certain extent by the ensuing protracted hold off phase during the landing, as the aircraft decelerates to touchdown speed.

If the height surplus is so great that even these remedies would not prevent a touchdown too close to the far boundary of the field, then the best course of action is to choose an alternative field in which a touchdown can be made closer to the near boundary.

5.1.1.9 Summary of forced landing procedure

Figure 246 summarises the sequence of events in the forced landing procedure.



5.1.1.10 After landing

After landing, the pilot's first concern should be for his passengers and himself, summoning assistance if injuries have been incurred.

The aircraft should be protected from damage by livestock or sightseers. (Again, assistance will probably be needed.) Finally, the police (and if applicable the aircraft owner) should be consulted as to the removal of the machine from the field.

5.1.1.11 Engine failure at low height

We have seen that the lower the height at which engine failure occurs, the less time there is available to deal with the situation before the inevitable forced landing.

If the aircraft is not much above 1000 feet above ground level, there will clearly be insufficient time to fly a circuit, and the pilot will probably choose to land in a field upwind of the aircraft's position. If practicable, a base leg should be flown, to give greater flexibility in control of height loss (Figure 247).


 
Common sense dictates that greatest priority should be given to achieving a safe landing; the secondary actions described in 5.1.1 should only be carried out if time permits. If there is clearly insufficient time for all the secondary actions to be completed, it is probably best to omit the cause-of-failure checks so that the distress call can be transmitted and the impact checks carried out.

If the engine failure is experienced much below 1000 feet, the pilot should turn the aircraft into wind and land in the best area available, making further small turns as necessary to avoid obstructions. In the event of there being insufficient height for a turn into wind to be carried out, a landing should be made straight ahead. In either case, the flaps should be lowered prior to touchdown, if circumstances allow, to minimise the touchdown speed and the length of the landing run.

5.1.1.12 Engine failure after take-off

Perhaps the worst situation that a pilot might face is engine failure directly after take-off. The situation is aggravated by the fact that the aircraft will be in a nose-high attitude, and will very quickly decelerate to stalling speed unless the pilot immediately lowers the nose to maintain gliding speed.

If the aircraft is only just airborne, the obvious course of action is to close the throttle, lower the flaps, land on the remaining runway and attempt to stop before the airfield boundary is reached, if necessary resorting to the tactics described in 5.1.1.7 to lengthen the available landing path. These latter considerations also apply to the case of a take-off rejected at high speed (refer back to 4.28.5).
 
If the engine failure occurs during the climb-out, the pilot should lower the nose to maintain gliding speed and attempt to land in the best area available ahead of the aircraft. Although small turns might be necessary to reach a suitable area for landing or to avoid obstructions, or to turn into wind after a crosswind take-off, no attempt should be made to turn back to the airfield, since there may be insufficient height to complete the manoeuvre, and in any case the aircraft will contact the ground at high speed because of the tailwind. Figure 248 illustrates the point.



Should time permit during the descent, a distress call can be transmitted and as many as possible of the impact checks carried out. If circumstances allow, it is advisable to lower the flaps prior to touchdown, to minimise the touchdown speed and the length of the landing run.

5.1.2 Further considerations

The forced landing procedure described above has been evolved over many years and in most situations gives the pilot his best chance of successfully achieving a safe landing after engine failure. However, it is possible that circumstances demand modification of the procedure described. Even though common sense may dictate departure from this procedure, it is worth remembering that the first priority must always be to attempt to achieve a safe landing. All other considerations are of secondary importance.

If a forced landing in a heavily wooded area (such as a forest) is unavoidable, the pilot should try to land the aircraft onto the tree tops at the lowest speed possible. In other words, the aircraft should be brought to a stall just above the trees. The more supple branches in the upper regions of the trees will help to cushion the impact. This technique is also probably the best to adopt if forced to land on uneven or rocky ground.

Finally, it will be appreciated that flight over towns and cities should be at such a height as would allow a forced landing to be made clear of the built-up area in the event of total engine failure.

5.1.3 Practising forced landing procedure

The pilot can practise the forced landing procedure by closing the throttle to simulate engine failure. He should remember to clear the engine periodically during the glide. The impact checks should be simulated (rather than actually carried out) and the procedure discontinued at a safe height above ground level; the success or otherwise of the practice can be gauged without the need to continue the glide down to ground level.

5.2 THE PRECAUTIONARY LANDING

If for some reason a pilot in cruising flight must land before reaching an active airfield he should carry out a precautionary landing. Reasons might include:

(a) shortage of fuel;

(b) deteriorating weather;

(c) approach of night.

The reader will agree that these predicaments should not arise if the pilot has planned and managed his flight properly.

If it proves impossible to be able to land at an active airfield, the second best choice might be a disused airfield, although it should be remembered that many of these are in various degrees of dilapidation. As a last resort, the aircraft will have to be landed in a field.

When it becomes clear that a precautionary landing is unavoidable, the first action should be to decide how much time is available for searching for a field, remembering to allow perhaps 15 minutes for inspection of the field prior to landing. Flying in a downwind direction will allow a greater ground distance to be covered in the time available and so present a larger choice of fields.

If fuel remaining is the limiting factor, the aircraft should be flown at the IAS for maximum range. If deteriorating visibility is the problem, with the earth's horizon not clearly defined, it will probably be best to fly at a safe low speed, with the flaps at 'maximum lift', to give the pilot a better view ahead and more time to detect and avoid obstructions or higher terrain. Attitude control will be more difficult and will require greater concentration in these conditions.

The chosen field should fulfil as many as possible of the criteria listed in 5.1.1.1; it should be carefully inspected for suitability (this also applies to a disused airfield, of course) and the landing direction chosen (preferably into wind).

An approach should be made towards the intended landing path and the aircraft flown at a safe low speed, with the flaps at 'maximum lift', along the landing path on the right-hand side so that it can be inspected more closely. During this inspection run, the pilot should not neglect to monitor his ASI and to watch ahead periodically for the presence of obstructions. Satisfied that the landing path is acceptable, the pilot can go around and fly another circuit and approach.

A normal powered approach and landing should be carried out, unless the available length of the landing path necessitates a short landing. If the aircraft is damaged during the landing, the impact checks should be carried out after it has come to rest, and the machine evacuated. Subsequently, the considerations in 5.1.1.10 apply.

5.3 ENGINE FIRE

In modern aircraft, engine fire is even rarer than engine failure. Smoke or flames coming from the engine compartment are indicative of engine fire. It is most likely to be fuel or oil that is actually burning (possibly because of a broken pipe line), and so the logical course of action is to cut off the supply of fuel and shut down the engine, as directed by the specified engine fire drill, to try to extinguish the fire. (Most light aircraft are not equipped with engine fire extinguishers.)

5.3.1 Engine fire in the air

If the symptoms of engine fire are observed during flight, the pilot should immediately carry out the engine fire drill specified for his aircraft. The conscientious pilot will know the drill from memory. A typical engine fire drill might proceed as follows:

(a) close throttle;

(b) select fuel cock off;

(c) switch off electric fuel pump;

(d) set mixture control to 'idle cut-off';

(e) switch off magnetos.

With the engine shut down, the pilot will be obliged to carry out forced landing procedure (omitting the cause-of-failure checks, of course). Under no circumstances should the engine be restarted, as the fire is likely to recur.

After landing, the aircraft should be evacuated as quickly as possible.

5.3.2 Engine fire on the ground

If engine fire occurs whilst the aircraft is on the ground, the following actions should be carried out by the pilot:

(a) if moving, bring the aircraft to a halt and apply the parking brake;

(b) carry out the engine fire drill;

(c) order the evacuation of the passengers;

(d) transmit a distress call;

(e) switch off master switch;

(f) if near other aircraft or buildings, release the parking brake (to enable the machine to be pushed clear of them by the airfield fire crew);

(g) evacuate the aircraft.

After evacuation, the pilot should keep himself and his passengers well clear of the aircraft and let the fire crew deal with the situation.

5.4 CABIN FIRE

Even though flames may not be visible, smoke is indicative of fire. The most likely cause of fire in the cabin is overheating of a component in the electrical system as a result of malfunction. Most light aircraft are equipped with a fire extinguisher in the cabin for use in such situations.

Before taking any action, the pilot must decide whether or not the fire has an electrical source, assuming that it does so if any doubt exists. Then the specified cabin fire drill should be carried out. A typical drill might proceed as follows:

(a) if the source of fire is electrical, switch off master switch;

(b) use extinguisher at source of fire, if necessary;

If the fire is extinguished:

(c) open all available fresh air vents fully to ventilate cabin;

(d) arrange to land at the nearest suitable airfield.

If the fire cannot be extinguished:

(c) initiate a rapid descent;

(d) transmit a distress call;

(e) open all available fresh air vents fully to ventilate cabin. If this action increases the intensity of the smoke or flames, the vents should be reset as before;

(f) land in the best available field (preferably an airfield) in the vicinity. (Engine power will be available, of course, should it be needed.)

5.4.1 Further considerations

If an electrical fire has been successfully extinguished, further use of the electrical system should be kept to a minimum. In particular, no electrical service should be used if it is suspected that that service may have been responsible for the fire. When it is desired to make use of one of the services, the following procedure should be adopted:

(a) switch off all electrical services except that intended for use;

(b) switch on master switch;

(c) use service as desired;

(d) when finished with service, switch off master switch.

In the event of an electrical fire which cannot be extinguished, it will be necessary to switch on the master switch briefly to transmit a distress call.

The rapid descent technique is employed to bring the aircraft down to ground level as quickly as possible. One method is to close the throttle, lower the flaps and fly the aircraft at the flap limiting speed to maximise drag.

5.5 DITCHING

Whenever a single-engined aircraft is flown beyond gliding distance from land, there is always the possibility of having to ditch the machine in the water, for example after engine failure. During such flights, all occupants of the aircraft should therefore wear lifejackets (uninflated). As an extra precaution, many pilots carry an inflatable dinghy in their aircraft when flight over water is intended and some wear immersion suits and ask their passengers to do likewise. If ditching is inevitable, the following procedure should be adopted:

(a) fly towards any shipping in the area, or towards the nearest land if there is no shipping, flying at minimum gradient gliding IAS and using whatever engine power is available;

(b) transmit a distress call;

(c) carry out the cause-of-failure checks if engine failure has occurred for reasons unknown;

(d) carry out the impact checks prior to touchdown;

(e) attempt to touch down at the lowest possible speed (stalling the aircraft just above the surface), landing along the swell if the wind is light, or into wind if it is strong;

(f) leave the aircraft and inflate the lifejacket(s) (and dinghy).

5.6 LANDING GEAR EMERGENCIES

It is usual design practice for aircraft with hydraulically- or electrically-operated landing gear retraction mechanisms to have an alternative means of lowering the landing gear after failure of the normal mechanism. The alternative is often manual in operation.

If one or more of the wheels fails to lock in the 'down' position when the landing gear is lowered, recycling the landing gear (that is, retracting and then lowering it again) may cure the fault. Recycling should not be attempted, however, if it is suspected that the retraction mechanism is damaged.

Of course, it may well be that the wheels are in fact all locked down, but that the indicator lights are faulty. Flying the aircraft past the airfield control tower will allow ground-based personnel to observe any obvious landing gear defect. If any doubt exists, the pilot should assume that the wheel(s) in question are not locked down.

Whenever a landing must be carried out with one or more of the wheels not locked down, the impact checks (5.1.1.5) should be carried out before landing, leaving the switching off of the fuel cock and magnetos until immediately prior to touchdown. It is generally better to land on a hard runway rather than on grass as less damage is likely to be incurred by the aircraft, and therefore less harm caused to the occupants. If not already done, the impact checks should be completed after the aircraft has come to rest. The machine should then be evacuated.

If the landing has to be made with all three wheels in the 'up' position the hold off should not be prolonged. In other words, the aircraft should be allowed to touch the runway gently in a flatter attitude than normal, to prevent the nose dropping violently to the ground after the tail has touched.
 
If it is the nosewheel that has failed to lock down, a normal touchdown should be made on the mainwheels, and the nose then held clear of the ground as long as possible using the elevators.

In the case of one of the mainwheels failing to lock down, it may be advantageous to retract the landing gear again and carry out a 'wheels up' landing in order to minimise the problem of directional control after touchdown. On some aircraft designs, however, the landing should be carried out with as many wheels locked down as are available. The Flight Manual will explain the best procedure to adopt.

Whenever a landing has to be made with only one mainwheel locked down, the ailerons should be used during and after touchdown to prevent the other wing from touching the ground. When ground contact can no longer be avoided, the pilot should be prepared for the swing that will occur towards the grounded wing. Use of the nosewheel steering (if available) and rudder will help to minimise the swing, although it may prove impossible to prevent the aircraft leaving the runway. This consideration should be borne in mind when choosing the runway intended for landing - it is clearly desirable to have a minimum of obstructions to the side of the runway.

   





6   THE FLIGHT MANUAL






The Flight Manual is the main source of technical information for any particular design of aircraft. Familiarity with this information will enable the pilot to operate his aircraft safely and efficiently.

Broadly speaking, the Flight Manual can be divided into four parts, dealing with:

(a) a technical description of the aircraft and its components;

(b) limitations;

(c) performance;

(d) loading.

6.1 TECHNICAL DESCRIPTION

The general arrangement of the aircraft and its components will be described, including any special points of interests. The Flight Manual will also include checklists for use during normal operating procedures and during emergency procedures.

The Flight Manual will state the grade of fuel required and the rates of consumption for various settings of the engine controls. The grade of oil required and the maximum permitted rate of consumption will also be stated.

6.2 LIMITATIONS

The Flight Manual will state the various airframe limitations (refer back to 4.1) and the engine operating limitations (refer back to 4.2.3).

6.3 PERFORMANCE

We have seen that, to optimise the aircraft's performance, it is often necessary to fly the machine at specified indicated airspeeds.

The Flight Manual will specify the correct IAS for:

(a) maximum range (an optimum altitude may also be stated);

(b) maximum endurance;

(c) maximum rate of climb;

(d) maximum gradient of climb;

(e) optimum climb with flaps at 'maximum lift';

(f) minimum gradient of glide;

(g) take-off with flaps up, and with flaps at 'maximum lift';

(h) powered approach, and flapless approach;

(i) arrival at the runway threshold after powered approach, glide approach, flapless approach and for short landing.

Also stated will be the power-off stalling speed (IAS) in straight and level flight at MTWA with flaps up and with flaps down. As we have already noted, these speeds are frequently marked on the ASI. Additionally, the stalling speed at lower weights may be given.

The actual performance achieved when flying at the specified speeds depends upon other factors. The Flight Manual contains graphs which enable the pilot to assess his aircraft's performance (for example, gradient of climb) in the circumstances prevailing. Perhaps the most important performance consideration for the pilot of a light aircraft is the adequacy of the airfield for take-off and landing. In other words, the pilot will want to be sure that the runway intended for use is long enough. For most light aircraft designs it can be assumed that a runway long enough for take-off will also be sufficient in length for landing under the same conditions. Accordingly, we shall consider the factors affecting the length of take-off run required to become airborne.

6.3.1 Length of take-off run required

By reference to the appropriate graph in the Flight Manual, the pilot can calculate the length of runway required to become airborne. The length of the take-off run depends upon:

(a) wind component;

(b) air density, in other words, airfield elevation and air temperature;

(c) aircraft loaded weight;

(d) flap position;

(e) gradient of runway;

(f) nature of runway surface.

To illustrate these factors, let us take the example of a typical light aircraft, loaded to 800 kg, taking off from an airfield at sea level in ISA conditions (air temperature 15C) with no wind component. The performance graph for this particular design of aircraft shows that, in these circumstances, the machine requires a take-off run of 280 metres (m) to become airborne. Throughout the take-off, the aircraft's TAS will equal its IAS (because there will be no density error) and its GS will equal its TAS (because there is no wind component). The design in our example has a take-off speed of 60 knots IAS. At lift off, then, the aircraft's TAS and GS will both be 60 knots.

The reader will appreciate that any factor which increases the GS at lift off will increase the length of take-off run, and vice versa.

6.3.1.1 Wind component

If the aircraft is made to take off facing into wind, it is said to have a headwind component. Taking off in the opposite direction, it would have a tailwind component. In the former case, the aircraft's GS at lift off will be less than its TAS, and so the length of the take-off run will be reduced. In the latter, the GS will be greater than the TAS, and so the length of take-off run will be increased.

Consider our example above, with conditions as stated, except that the take-off is made with a 10 knot headwind component. The GS at lift off will be 50 knots and the length of the take-off run 250 m, compared to 280 m with no wind component. With a 10 knot tailwind component, the GS at lift off will be 70 knots and the length of take-off run 380 m - a considerable increase. Figure 249 illustrates the two cases.



When the wind direction is coincident with, or opposite to, the runway direction, it is easy to assess the headwind or tailwind component, because these quantities will be exactly equal to the wind strength. (Remember that the wind direction is the direction from which the wind blows.) In crosswind conditions, the pilot will need to refer to a wind component table to assess the headwind or tailwind component. Such a table is shown in Figure 250.



Common sense dictates that take-off should be made with a headwind component whenever circumstances allow, both to reduce the length of take-off run required to become airborne and to steepen the subsequent gradient of climb.

6.3.1.2 Airfield elevation

The higher an airfield is above sea level, the less the ambient air density will be. The TAS of an aircraft taking off from an airfield above sea level will therefore be greater than its IAS. Assuming no wind component, the aircraft's GS at lift off will equal its TAS and will therefore be greater than its IAS. This increased GS will result in a longer take-off run.

A further consideration is the fact that the aircraft's engine will deliver less power with the throttle fully open than at sea level (because the air density is less), so reducing the acceleration during take-off and further increasing the length of the take-off run.

The aircraft in our example, taking off from an airfield 2000 feet above sea level in ISA conditions (air temperature 11C) with no wind component, will require a take-off run of 360 m, compared to 280 m at the sea level airfield.

6.3.1.3 Air temperature

With increasing air temperature, the ambient air density at an airfield becomes less. So, for any particular IAS, the aircraft's TAS (and hence GS) will be greater. Additionally, the engine will deliver less power with the throttle fully open. These two effects will combine to increase the length of the aircraft's take-off run.

If we consider the case of our aircraft taking off from a sea level airfield in ISA+10C conditions (air temperature 25C) with no wind component, the take-off run will be 310 m, compared to 280 m in ISA conditions.

Of course, in conditions cooler than ISA, the length of the take-off run will be reduced.

6.3.1.4 Aircraft loaded weight

The take-off speed quoted as an IAS in the Flight Manual applies to the case when the aircraft is loaded to its MTWA. In practice, take-offs at lower loaded weights are also made at this speed.

However, the greater the loaded weight of the aircraft, the slower will be the acceleration during the take-off run and so the greater will be the length of the take-off run.

The aircraft design in our example has a MTWA of 1000 kg. When loaded to this weight and taking off from a sea level airfield in ISA conditions with no wind component, the length of the take-off run will be 340 m, compared to 280 m when loaded to 800 kg.

6.3.1.5 Flap position

The Flight Manual will specify the recommended flap position for take-off, to which the stated performance figures will relate.

For most aircraft designs, the take-off run will be shorter when the flaps are set to 'maximum lift' than with them up.

6.3.1.6 Gradient of runway

As a general rule of thumb, the length of the take-off run is increased (or decreased) by 10% for every 1 of runway slope uphill (or downhill) because of the slower (or faster) acceleration.

6.3.1.7 Nature of runway surface

The Flight Manual's take-off run performance figures usually relate to a tarmac or concrete runway. A dry, short-cut grass surface will slightly retard the aircraft's acceleration, so increasing the length of the take-off run. Long grass, wet grass or soft ground will further increase the length of the take-off run.
 
6.3.1.8 Practical considerations

It would clearly be tedious to have to calculate the length of take-off run required before every flight. Indeed, it is not usually necessary to do so. Most pilots calculate a 'standard' take-off run requirement for their aircraft assuming the following circumstances:

(a) no wind component;

(b) an average elevation for the airfields from which the aircraft is likely to be operated;

(c) air temperature warmer (for example 10) than ISA conditions;

(d) aircraft loaded to its MTWA;

(e) level tarmac or concrete runway.

If the available runway length at an airfield is comfortably in excess of this 'standard' figure then the pilot will not need to make a full calculation of take-off run required, since the runway is obviously long enough for take-off. The safety margin will be enhanced by ensuring that take-off is made with a headwind component if there is a wind blowing.

If the available runway length is only slightly in excess of, or is less than, the 'standard' figure, it will be necessary to make a full calculation for the circumstances prevailing. Should the calculated take-off run requirement exceed the available runway length it is clearly unsafe to attempt to take off, and it will be necessary for the pilot to reduce the loaded weight of his aircraft in order to be able to take off safely.

6.3.2 Obstacle clearance after take-off

The Flight Manual will contain graphs enabling the pilot to calculate the ground distance required from the start of the take-off run to the point where the aircraft has gained a certain height (usually 50 feet). When obstacles such as buildings or trees are sited close to the far threshold of the runway intended for use, these graphs will allow the pilot to verify that such obstacles will be safely cleared.

In practice, the pilot will check that the calculated distance to reach the specified height in the circumstances prevailing is less than the actual distance from the start of the take-off run to the obstacle in question, if necessary reducing the loaded weight of his aircraft to make sure that this is so.

Again, many pilots calculate a 'standard' figure for distance required to take off and climb to the specified height. This figure will ensure safe clearance over the airfield boundary (fence or hedge) if the boundary is comfortably beyond this distance.

6.4 LOADING

We have seen the importance of ensuring before flight that the aircraft is safely loaded. Two requirements must be complied with:

(a) the aircraft's loaded weight must not exceed the MTWA specified in the Flight Manual;

(b) the CG of the loaded aircraft must lie within the specified limits.

6.4.1 Empty weight and loaded weight

The empty weight of an aircraft is its weight when unladen, that is, with no fuel, oil, persons or baggage on board. The empty weight will be specified in the Flight Manual. Obviously, the loaded weight is the sum of the empty weight and the weights of whatever items are loaded on board.

6.4.2 CG position

Determination of the CG position is not quite so straightforward as calculation of the loaded weight, but is not a difficult procedure once practice has been gained.

The procedure involves the concept of lever arms and moments. The term 'lever arm' applies to a force and is the distance at which the force acts from a fixed datum point (Figure 251).



The moment of the force is the product of its magnitude and its lever arm:

Moment = Force x Lever arm

Consider an empty aircraft and take the extreme nose of the machine as the datum. It is possible to calculate a moment for the empty aircraft by multiplying its empty weight by the lever arm, that is, the distance of the CG from the datum (Figure 252).



This calculation is performed by the aircraft manufacturer and the calculated moment is stated in the Flight Manual.

Consider now loading the aircraft by placing the pilot in his seat in the cabin. The pilot's moment will be the product of his weight and his lever arm, as shown in Figure 253.



How can we now determine the position of the CG of the loaded aircraft? In fact it is merely necessary to divide the total moment, that is M(e) plus M(p), by the loaded weight, W(e) plus W(p). The result of this calculation is the distance of the CG of the loaded aircraft from the datum.

As equations:

Loaded weight = W(e) + W(p)

Total moment = M(e) + M(p)

CG position = Total moment
                        Loaded weight

Figure 254 illustrates the CG positions for the aircraft when empty and when loaded. In the diagram 'LA' is the lever arm for the loaded aircraft. Notice that, in this example, the loaded GC position is forward of the empty CG position because the pilot's seat is forward of the latter.



If we now add more load in the form of fuel, oil, passengers and baggage, the same process is repeated to determine the CG position of the loaded aircraft.

To illustrate the procedure let us consider an example. The Flight Manual of a typical aircraft states the following limitations:

(a) the MTWA of this aircraft is 1000 kg;

(b) the CG position of the loaded aircraft must lie within the range 225 to 244 centimetres (cm) to the rear of the datum.

The Flight Manual also states that the CG datum is the extreme nose of the aircraft. Figure 255 illustrates the permitted range for the CG position and shows the positions of the fuel and oil tanks, the cabin seats and the baggage compartment.



Let us suppose that it is desired to take off with the following load:

(a) pilot (76 kg);

(b) two passengers, one in the other front seat (69 kg) and one sitting behind (88 kg);

(c) three suitcases (total 43 kg);

(d) 8 litres of oil (the maximum capacity for this design);

(e) 112 litres of fuel.
 
To verify that this load distribution satisfies both the stated limitations, it is necessary to consult the loading section of the Flight Manual, where the table shown in Figure 256 is given.



Firstly, we can enter the weights of the various items of load in the appropriate column in the table. Note that there is no need to calculate the weight of oil since this has already been done by the aircraft manufacturer. 112 litres of fuel will weigh (112 x 0.72) or 81 kg. Note, too, that a maximum permitted baggage weight has been specified, which must not be exceeded.

Now we can complete the table by calculating the moment (weight multiplied by lever arm) for each item. It is interesting to note that the pilot has been combined with the front seat passenger, because the lever arm is the same for both. For convenience the moment values can be rounded to the nearest 100 units without significantly compromising the accuracy of the calculation. Finally, the individual weights can be added together to determine the loaded weight, and the individual moments to determine the total moment (Figure 257).



Immediately it can be seen that the first loading limitation has been met, since the loaded weight (954 kg) is less than the MTWA.

The CG position is determined by dividing the total moment by the loaded weight:

CG position = 219,800 = 230 cm to the rear of the datum.
                             954

This answer is within the permitted range. The intended load disposition is therefore acceptable. (The change in CG position during flight as fuel and oil are consumed is negligible and can be disregarded.)

6.4.3 Practical considerations

As with the take-off performance calculations we looked at earlier, it would clearly be tedious to have to carry out a full loading calculation before every flight. In fact it is not usually necessary to do so. Most pilots prefer to adopt 'standard' loading configurations for which they have performed full calculations and which they know will satisfy the Flight Manual limitations. These 'standard' configurations will assume an average weight for pilot and passengers. It is then necessary to make a full calculation only when the pilot intends to load his aircraft in a manner differing from the 'standard' configurations, or if the individual passenger weights differ considerably from the assumed average weight (for example, when small children are to be carried).

If such a calculation shows that the CG position of the loaded aircraft is ahead of the forward limit, then the pilot can move the CG rearwards by arranging a more rearward disposition of the intended load (for example, seating the heavier passengers in the rear seats). Conversely, a more forward disposition of the load will bring the CG position forwards if the calculated position is to the rear of the rearward limit.

6.4.4 Dangers of incorrect loading

Overloading an aircraft will reduce its performance (for example, increasing the length of take-off run and reducing the maximum achievable gradient of climb). Additionally, the landing gear may be overstressed when taking off and landing, and the airframe structure if turbulent air is encountered during flight. The stalling speed will also be increased.

As we have seen, a CG position to the rear of the rearward limit will reduce the longitudinal stability of the aircraft in flight and will make it easier for the machine to enter a stall or spin if the pilot handles the controls carelessly. Furthermore, the aircraft will show greater reluctance to recover from such a stall or spin.
 
A CG position ahead of the forward limit will require a greater tail down-force to hold the lift and weight forces in equilibrium. In other words, the elevators will have to be held in a more upward-displaced position during flight. This may leave insufficient upward elevator movement when it is desired to pitch the aircraft's nose up. Thus it may prove difficult to raise the nose for take-off and impossible to flare the machine for landing.

   





7   GROUND SCHOOL SUBJECTS






Besides aircraft handling skills, knowledge of the theory of flight and thorough familiarity with use of Flight Manuals, student pilots training for their licences will need to reach a high standard of knowledge in various ground school subjects. These include:

Navigation
Flight planning
Operational procedures
Radio communications
Aviation law
Meteorology
Human performance

Most flight training organisations offer courses in these subjects as part of the overall training syllabus for pilot licences.

   





Also by Julien Evans: Fiction

Madeleine's Quest
Chalk and Cheese
The Sommerville Case
The Damocles Plot
Flight 935 Do You Read

Non-fiction:

How Airliners Fly