Category Aircraft Flight

A fan is essentially a propeller with a large number of blades, and therefore provides a means of producing a large amount of thrust for a given disc area. When there are many blades, so that they are close together, each blade strongly affects the flow around its adjacent neighbours. This interference can

have a beneficial effect if the relative flow speed is supersonic. The flow can be compressed gradually through a series of reflected shock waves, creating a smaller loss of energy than when the flow is compressed through a single shock.

As the flight proceeds fuel is used and the aircraft weight changes. This change may be very significant; up to half the total weight on a long-range transport aircraft. Thus the lift will decrease. In order to operate at the best lift coefficient we therefore need to reduce the dynamic pressure. If the aircraft is powered by a gas turbine we do not wish to reduce speed, or the engine efficiency will suffer. The only alternative to reducing the speed is to reduce the air density by climbing as the flight proceeds. Fortunately this is possible even if we are operating near the limiting height described in the previous paragraph. As the weight of the aircraft reduces, so this limiting height increases (Fig. 7.8) and we can achieve the desired increase in altitude without being squeezed into ‘coffin corner’, the point where stall and high speed buffet occur at the same speed. This technique is known as ‘cruise climb’.

OTHER TYPES OF POWERPLANT 201

Some practical considerations

In practice, other considerations may influence the way in which the cruising height is selected and the way in which the cruise climb technique is operated.

In the first place, we must remember that the cruise is only part of the flight. The aircraft must both land and take off, and the flight plan must be optimised over all phases as a whole, not just during cruise. Thus for short journeys the cruising height is likely to be lower than for long-range flights.

As far as the cruise climb is concerned, factors such as Air Traffic Control requirements do not enable the technique to be followed as closely as the pilot would like. For example, in order to achieve safe separation of aircraft the height may be dictated by safety rather than economy, and airline pilots have to get permission from Air Traffic Control before changing height.

Since the very earliest days, the three primary control actions available to the pilot of a conventional aircraft have been those of pitch, roll, and yaw, as defined in Fig. 10.1. On most military interceptor aircraft, the controls are operated by the same type of control stick or joystick as used on early aircraft. On most other aircraft, some form of handlebars or spectacle grip is provided, either protruding from the instrument panel as in Fig. 10.2, or mounted on a movable control column.

With the introduction of completely electronically operated fly-by-wire sys­tems (described later), where the control column provides no direct mechanical operation of the control surfaces, a new form of control called a sidestick has

been introduced, as seen in Fig. 10.3. This is a miniature form of joystick designed for one-handed operation, and mounted at the side of the pilot’s seat. The use of a sidestick produces a less cluttered flight deck, as shown in the photograph.

In conventionally controlled aircraft, pulling back on the stick or handlebars produces nose-up pitching action. Note, however, that in a weight-control hang-glider or microlight (Fig. 11.10), the control action is reversed; the pilot pulls on the control bar to transfer his weight forwards, tending to produce a nose-down effect. Pilots of conventional aircraft need to be very careful when converting to microlights, and vice-versa.

Turning the handlebars on a conventional aircraft clockwise, or pushing the stick to the right, produces, as its primary effect, a clockwise roll (and a

consequential tendency to turn to the right). In this book, by left and right we mean pilot’s left and right.

Yaw control is provided by foot-operated pedals. Pushing on the pedal bar with the right foot causes the aircraft to yaw to the pilot’s right. Most people find this pedal action natural, which is curious, because unlike the other con­trols, the pedals work in the opposite sense to the turning direction required. On a bicycle, pushing the right handlebar would turn the bicycle left.

Note, that the amount of rotation of the handlebars affects the rate of roll, rather than the angle to which the aircraft rolls.

Delta wings are effectively swept, and tailless delta-winged types can be stabilised in the same way as other tailless aircraft. However, a major problem arises with supersonic tailless deltas, because the rearward shift of the centre of lift position has to be trimmed by a large up-elevator movement. This significantly reduces the lift, while increasing the drag. On Concorde, the prob­lem is mainly solved by rapidly pumping fuel from a front tank to a rear one, so as to move the centre of gravity back for high speed flight. The movements of fuel during any flight have to be carefully calculated before take-off. This method of stability control, though complicated, does result in efficient flight with little or no trim drag.

In addition, the shape of the camber line may be used to control movement of the centre of lift. In Chapter 1 we explained how the lift due to angle of attack, and lift due to camber, were almost independent. At low speeds, the lift coefficient and angle of attack are large, so the lift force is dominated by the angle of attack. The centre of lift will be about – of the way back from the leading edge (- chord position). At high speed, the lift coefficient and angle of attack are low, and the lift is dominated by the camber. By suitable shaping of the camber line, the centre of lift at high speed (low angle of attack) can be arranged to be at about the same position as at low speed (high angle of attack). The very pronounced droop of the leading edge of Concorde’s wing produced by the camber may be seen in Fig. 2.23.

Because of the problems of control and stability of tailless deltas, many delta-winged aircraft have a small tail, or a foreplane.

Structural flutter

Flutter is the name given to a form of structural vibration that normally involves a combination of motions; typically bending and twisting. It is most likely to occur on wings, but tailplane flutter is not uncommon. The Handley Page 0/400 bomber produced an early, well documented, case of tailplane flutter, during the First World War.

Figure 14.4 illustrates one flutter mode. In this example, the wing is oscil­lating in torsion (twisting) as well as bending. As it bends upwards as in position-1 in Fig. 14.4, it is twisting in a nose-down sense. When it passes the limit of upward travel and starts to spring back down, it presents a negative (nose-down) angle of attack as in position-4. As the lift force is now down­wards, it aids the motion.

Near the bottom limit of travel, the wing starts to twist nose-up, so that when it springs upwards (position-9), the angle of attack has become positive,

and the lift force again aids the motion. Because the lift force is assisting the movement, the motion does not damp out like a normal vibration, and it can continue, sometimes with the amplitude increasing until failure occurs.

It will be seen, that in the case described above, the torsional and bending oscillations are 90° out of phase; that is, the bending deflection reaches its max­imum as the twist approaches its mid-position.

The normal remedy is to increase the torsional stiffness of the wing, but flutter can also occur by a coupling of the wing flexure with a pitching oscillation of the whole aircraft. This problem was encountered on some tailless aircraft, for which the pitching inertia was low. In such cases, increasing the torsional stiff­ness would not help, and it was the flexural stiffness that had to be increased.

The mass distribution also critically affects the flutter behaviour of the wing, and the location of wing-mounted engines is an important factor.

As with divergence, the onset of flutter depends on the aircraft speed. The lowest speed at which flutter occurs, is known as the critical flutter speed, and again, it is necessary to ensure that the aircraft speed never reaches the critical value.

In the helicopter the rotor blades are, in effect, long rotating wings of small chord. The blades are mounted on an engine-driven shaft. As they move through the air, they generate lift in the same way as a fixed wing. The obvious advantage over a fixed-wing aircraft is that the rest of the aircraft does not need to move relative to the air, and it can therefore hover.

The torque reaction of the motor tends to rotate the fuselage in the opposite direction to that of the rotor, and on a conventional single rotor helicopter, a tail mounted propeller or fan is used to counteract this effect. The tail pro­peller, which is normally referred to as the tail rotor, wastes power, and is one cause of the poor efficiency of simple helicopters. A recent innovation is the so-called no-tail rotor (NOTAR) design in which the tail rotor is replaced by a jet of air which interacts with the main rotor downflow to produce the required torque. The NOTAR configuration has a number of operational advantages including reduced noise.

When two rotors are used, they can be arranged to rotate in opposite direc­tions, thus cancelling out the unwanted torque reaction, and removing the need for a tail rotor. The two rotors are normally arranged at opposite ends of the fuselage, but sometimes, particularly in Russian designs, they may be arranged so as to counter-rotate on concentric shafts as illustrated in Fig. 1.25. The use of two rotors considerably increases the cost and complexity of the aircraft. There are also problems due to interference between the two rotor wakes.

On a simple single rotor helicopter, direct control of the amount of lift gen­erated by the main rotor blades is provided by the collective pitch mechanism which changes the incidence or ‘pitch angle’ of all of the blades by the same amount simultaneously. In addition, a cyclic pitch mechanism is provided. This causes the incidence of the blades to increase and decrease once per cycle. The cyclic pitch control is used to control both the nose-up (or nose-down) attitude of the helicopter, and the roll motion about the longitudinal axis. The small tail rotor is also provided with a mechanism which can be used to vary the incidence of its blades, thus altering the amount of thrust produced. This allows it to be used to yaw the aircraft.

On conventional helicopters, particularly older designs, the rotor blades are hinged about two axes at the point where they are joined to the hub, in such a way that they are free both to swing back and forth, and to flap up and down (within limits). The swinging motion is primarily intended to damp out cyclic variations in drag, and this hinge is called the drag (or lag) hinge. A damper is fitted to prevent unwanted oscillations. The flapping motion is resisted by strong centripetal forces, which tend to keep the blades nearly at right angles to the axis of rotation. The lift forces on the blades are always very much smaller than the centripetal forces, but they nevertheless pull the blades up slightly, so that in flight, their rotational path describes a very shallow cone rather than a flat disc.

When the cyclic pitch control is used to vary the incidence cyclically, the blades will tend to flap up and down cyclically in response to the lift variation. If the blades are caused to flap up at the rear, as shown in Fig. 1.27, the effect is to tilt the axis of tip rotation forwards. This generates a horizontal thrust force component, as illustrated, and the helicopter is thus propelled forwards. The flapping hinges are normally offset from the centre of the hub axis, and as the axis of blade rotation is inclined relative to the shaft axis, the centripetal forces produce moments which tend to tilt the rotor axis (and hence the whole aircraft) nose-down. By suitably adjusting the cyclic pitch control, the hori­zontal propulsion force can be arranged to occur in any desired direction, not just forwards, and it is thus possible to fly a helicopter sideways or backwards.

Fig. 1.27 Helicopter in forward flight

The blades are caused to flap up at the rear by use of the cyclic pitch control. The resultant force provides both lift and thrust components

When a helicopter flies forwards, the relative air speed on the advancing blade side will be greater than on the retreating side. If no corrective action were taken, this would cause the blades to flap up towards the front and down towards the rear, thereby tilting the axis of the blade rotation backwards: a condition known as blowback. If this were allowed to happen unchecked, the helicopter would simply slow down due to the now rearward thrust com­ponent. This tendency therefore has to be overcome by making use of the cyclic pitch control to reduce the blade angle of attack when advancing, and increase it when retreating. By this means, the lift is equalised on the advancing and retreating sides. If simple rigidly fixed blades were used, as on a propeller, the blades would generate more lift when advancing than when retreating, and the aircraft would tend to roll.

Two-bladed rotors may use a teetering rotor instead of flapping hinges. In this arrangement, the two blades are rigidly connected together, but allowed to tilt about the hub to produce a flapping motion. A teetering rotor is used on the autogyro shown in Fig. 1.31.

Some more recent helicopter designs do not use hinged blades, but rely on carefully controlled flexure of the mounting points: an arrangement that is often misleadingly referred to as a rigid rotor. Figure 1.26 shows a ‘rigid rotor’ head mechanism. Its relative simplicity and compactness may be contrasted with the great complexity of the conventionally-hinged double unit of the Kamov shown in Fig. 1.25.

The rotor head contains not only the hinges or flexures but must also carry the mechanism for the cyclic and collective pitch control of the blades. This unit is thus a very complicated and heavily loaded item, and almost any mechanical failure is likely to be disastrous. The safety record of helicopters is generally inferior to that of fixed-wing aircraft.

It should be noted that the description above applies to conventional heli­copters. Various alternative rotor mechanisms have been tried, including ones with tilting shafts.

Since the rotor provides thrust, lift and the primary means of control, the helicopter can be seen as a good example of a radical departure from Cayley’s classical aeroplane, where each component serves only one specific purpose.

The helicopter also suffers from other problems stemming from the differ­ences in the relative airflow velocities on the advancing and retreating blades. In order for the retreating blade to generate any lift at all, it must be moving faster than the relative airflow. As you can see from Fig. 1.28, this means that the advancing blade must be moving through the air at more than twice the speed of the aircraft. The advancing blade will, therefore, approach the speed of sound, when the aircraft is still only travelling at well below half this speed. Having a blade that continually moves in and out of supersonic flow produces considerable structural and aerodynamic problems, not the least of which is the noise created.

Figure 1.28 also shows a common situation, where the inboard part of the retreating blade is not moving fast enough to overtake the air flow. The

Resultant

Resultant relative velocity

Fig. 1.28 Relative air flow velocities in forward flight air therefore actually flows backwards relative to the blade on this portion. These factors severely limit the maximum speed of conventional helicopters. Figure 1.29 shows the Westland Lynx, which achieved a record-breaking 249.10 mph in 1986. This is less than half the speed attained by the fastest propeller-driven conventional aircraft.

A complete description of helicopter aerodynamics is beyond the scope of this book, but the above outline gives some idea of why rotating wing aircraft have not displaced fixed wing types, despite the obvious attraction of vertical take-off and landing.

In Chapter 1, we described how the lift coefficient of a wing depends on its camber. The wing camber can be changed in flight by deflecting the trailing edge downwards as shown in Fig. 3.13. The hinged trailing edge is known as a flap. The simple hinged flap shown in Fig. 3.13(a) is often used on light aircraft. The split flap shown in Fig. 3.13(b) is an alternative arrangement that was commonly used during and just after the Second World War.

The stalling effect, caused by flow separation, however, limits the maximum value of CL that can be obtained in this way. The key to producing very high

lift coefficients is to be found in inhibiting or controlling the separation of the boundary layer.

Since separation is associated with the dissipation of energy in the bound­ary layer, it follows that we can prevent separation either by removing the boundary layer, or by adding energy to it. The slotted flap shown in Fig. 3.13(c) represents one simple method. The slot allows air from the undersurface to blow over the flap, so that a fresh new boundary layer is formed on the flap, helping to maintain attachment. The tired wake from the main wing element may also be re-energised by turbulent mixing with the air emerging from the slot, but this is a secondary effect.

On sophisticated aircraft, it is normal to use a flap element that slides out, thereby increasing the wing area. This type is known as a Fowler flap, and is illustrated in Fig. 3.13(d). For very high lift coefficients, the flaps may be split into two or more elements. A two-element slotted flap is shown deployed on the Tornado in Fig. 3.14, and illustrated in Fig. 3.13(e).

Because an oblique shock wave is able to impose a sudden change in the direc­tion of an oncoming air stream (Fig. 5.12), the necessary flow deflection around an aerofoil with a sharp leading edge can be achieved with an attached shock wave system (Fig. 5.13) in which the bow shock waves emanate from the lead­ing edge itself.

However, there is a limit to the angle through which a flow can be deflected. This depends on the Mach number of the flow. If this critical angle is exceeded, the shock wave becomes detached (Fig. 5.14) and looks very much like the bow shock wave of the blunt aerofoil described earlier (Fig. 5.1).

So far we have only considered sudden changes in flow direction. If the flow is turned gradually (Fig. 5.15), the picture looks slightly different. Near to the surface the flow compresses and turns without a shock wave, but one is observed further away from the surface. The reason for this is that as the flow compresses its temperature rises. The speed of sound therefore increases and if we draw ‘Mach lines’ to indicate the extent to which each point on the surface can influence the oncoming flow, we see that these get progressively steeper and eventually run together to form the shock wave.

The compression near the surface is known as a ‘shockless compression’ and we will see later how this type of compression can be exploited in practical design as it involves no wave drag.

As the air flows through a compressor, its pressure and temperature rise. The rise in temperature means that the speed of sound increases, so without raising the Mach number of the flow, we can afford to let the later (high pressure) stages run at a higher speed than the early (low pressure stages). On modern engines, it is therefore usual to use two or more concentric shafts or spools. Each spool is driven by a separate turbine stage and runs at a different speed. Figure 6.22 shows a two-spool layout based on the Rolls-Royce Olympus 593 fitted to Concorde.

In turbo-prop engines, it is normal to drive the propeller from a separate turbine stage and spool from that of the main or core section of the engine. The propeller and core engine speeds can, therefore, be partially independently controlled.

The Rolls-Royce Gem engine, shown in Fig. 6.21, is described as a turbo­shaft engine, as it is intended to drive a helicopter rotor shaft rather than a propeller. It combines many of the features described above. Three spools are used, one to drive a single stage high pressure centrifugal compressor, one to drive a multi-stage low pressure axial compressor, and one to drive the rotor shaft via a gearbox.

turbine

Hot and cold air

jets have nearly

the same velocity at the point where they

meet

Fig. 6.23 A two-spool low by-pass ratio jet engine (or low by-pass turbo-fan)

Only part of the air passes through the combustion chamber. The rest is by-passed around the core. This type of engine is quieter and more fuel-efficient than the simple type shown in Fig. 6.2. It is commonly used in high performance military aircraft

So far we have considered the aircraft to be flying in a straight line. We now turn our attention to the important manoeuvre of changing direction; the turn.

We look first at the case in which the turn is made at constant altitude. Any time a turn is made by any vehicle, whether it be a bicycle, car, train or aeroplane, a force must be provided towards the centre of the turn because there is an acceleration directed towards the centre. In the case of a car or bicycle this force is provided by the tyres; in the case of a train it is provided by the rails. For an aircraft some other means must be found and this is done by tilting, or banking the aircraft so that a component of the lift force produced by the wings acts in the required direction (Fig. 7.21). Thus the wings must produce a higher amount of lift than was required for normal straight and level flight.

This extra lift means that, for a given speed, the wing must be operated at a higher angle of attack in the turn and, in addition, the increase in the lift will be accompanied by an increase in drag. This drag will, in turn, mean that the power required to sustain a steady turn is greater than that required for flight in a straight line at the same speed. The angle of bank and increase in lift, drag and required power all increase as the turn is tightened, and it may be that the minimum radius of turn which can be achieved is limited by the amount of power that is available from the engine. Alternatively the demand for extra lift may cause the wing to stall before this point is reached, and stalling may therefore prove to be the limiting factor.

Vertical lift component balances weight

Lift component provides required centripetal force

Fig. 7.21 Forces in turning flight

Lift must increase to provide both the vertical component to balance weight and the required forces for the turn

Recommended further reading

Mair, W. A. and Birdsall, D. L., Aircraft performance, Cambridge University Press, Cambridge, 1996, ISBN 0521568366. A good general text on aircraft performance.