Category Aircraft Flight

An example of a hybrid rocket engine is shown in Fig. 6.41. In the design shown, a rocket is used to drive a turbine and to produce a hot jet. At low altitude, part of the thrust comes from heating and expanding the air as in a turbo-jet. At very high altitude, where the air density is too low for an air-breathing engine, it becomes a pure rocket. A wide variety of alternative air-breathing or hybrid arrangements has been investigated.

Engines involving a combination of rocket and air-breathing propulsion have so far only been used for missiles, but applications to hypersonic and orbital aircraft have been proposed. Further discussion of propulsion for hypersonic aircraft is given in Chapter 8.

In Chapter 5 we discussed the sudden drag rise which takes place in the transonic speed range. Different aircraft are designed to operate efficiently over different speed ranges. All are required at least to be able to take off and land subsonically and to accelerate or decelerate safely through the transonic range. Some, however, such as the Tornado we discussed above, may need to operate for prolonged periods at a variety of transonic and supersonic Mach numbers. The selection of a suitable planform must thus be a compromise.

Figure 8.17 shows how the drag of different planforms varies with Mach number and from it is possible to see why the various planforms shown in Figs. 8.1—8.4, 8.8, 8.16 and 8.18 were selected to satisfy particular perform­ance requirements.

In low speed flight, where the wing may be close to the stall angle, the down­ward deflection of an aileron will produce an increase in drag associated with the increased camber and onset of stalling, while the upgoing aileron will pro­duce a reduction. This causes the aircraft to turn towards the lowered aileron, which is the reverse of the normal response described earlier. The problem can be partially overcome by the design of the control surfaces, as in the Frize aileron design shown in Fig. 10.14. Alternatively, the ailerons may be geared, so that the upgoing surface moves more than the downgoing one.

Spoilers present another, and sometimes preferable solution. The operation of a spoiler causes a loss of lift, and the required rise in drag on the downgoing wing. Spoilers are also used for control in high subsonic and supersonic flight, where conventional ailerons may either lose their effectiveness, or become too effective.

In Chapter 11 we examined the stability of an aircraft from the simple view­point of whether a disturbance from the steady flight condition produced forces and moments that tended to restore the aircraft to this equilibrium state. This is known as static stability. Static stability is, however, not the end of the story. We can have a statically stable aeroplane which is still not satisfactory in practice because it oscillates about the equilibrium position (Fig. 12.1). If the amplitude of the oscillation grows with time then it is said to be negatively damped, and the aircraft is dynamically unstable.

Figure 12.1 illustrates the concepts of static and oscillatory dynamic stabil­ity for an aircraft flying with the wings level, without roll or bank. In this case it is the longitudinal motion of the aircraft that is of interest to us. Let us con­sider a statically stable aircraft which is slightly disturbed by increasing the angle of attack. The subsequent motion may take a wide variety of forms as shown in Fig. 12.1(a) to (c).

The first of these (Fig. 12.1(a)) shows a motion in which the aircraft simply returns to the state it was in before the disturbance was applied. The motion is thus stable and the motion is not oscillatory in nature. Figure 12.1(b) shows an oscillatory motion, but since the oscillations die out with time the motion is again dynamically as well as statically stable. Finally Fig. 12.1(c) shows the oscillations becoming greater with time rather than dying away. This motion is the dynamically unstable negatively damped motion referred to in the first paragraph. It is statically stable because whenever the aircraft is at a pitch angle that differs from its initial value, the moment acting on the aircraft is in the direction which tends to restore it to its original position.

By the principle outlined above, it follows that any object rotated so as to pro­duce a vortex or circulation, will generate lift when placed in a stream of air. This is known as the Magnus effect. Figure 1.12 shows streamline patterns for air flow past a rotating cylinder.

It is possible to generate a very large amount of lift by using a rotating cylinder or paddle, but the mechanical complexities of such a system normally outweigh any potential advantages. Despite considerable interest, and many patents, the effect has rarely been exploited for commercial advantage, except by professional sportsmen; most noticably, tennis players, who use the prin­ciple to swerve a ball by imparting a large initial spin.

When air flows past any part of an aircraft it appears to try to stick to the surface. Right next to the surface, there is no measurable relative motion. The relative velocity of the air flow increases rapidly with distance away from the surface, as illustrated in Fig. 3.1, so that only a thin ‘boundary’ layer is slowed down by the presence of the surface. Note, that individual air molecules do not actually physically stick to the surface, but fly around randomly, at a speed that is related to the temperature.

In reality, there is no precise edge to the boundary layer, the influence just fades. For the purposes of calculations, however, it is necessary to arbitrarily define an edge. In the simple case of the flow over a flat plate with no stream – wise variation in pressure, shown in Fig. 3.1, it is customary to define the edge of the layer as being the position where the flow speed reaches 99 per cent of the free-stream value.

From an aeronautical point of view, it is the wing boundary layer that is of greatest importance, and in Fig. 3.2 we show a typical example of how the boundary layer develops on an aerofoil. It will be seen that the thickness of this layer grows with distance from the front or leading edge.

There are two distinct types of boundary layer flow. Near the leading edge, the air flows smoothly in a streamlined manner, and appears to behave rather like a stack of flat sheets or laminae sliding over each other with friction. This type of flow is, therefore, called laminar flow. Further along, as indicated in

Fig. 3.2, there is a change or transition to a turbulent type in which a random motion is superimposed on the average flow velocity.

The two types of flow have important differences in properties that we can exploit. In simple terms, the main practical effects are that the laminar layer produces less drag, but the turbulent one is less liable to separate from the sur­face, as described later. To understand why these differences occur we need to look at the two types in a little more detail.

At the time of writing, the most popular wing-tip device appears to be winglet, which may be seen on the Airbus A340 shown in Fig. 4.17.

As illustrated in Fig. 4.18, winglets take advantage of the strong side – wash that occurs at the wing tip. Due to the sidewash, the air flow meets the vertical winglet at an angle of attack, and thus a sideways force is generated. The winglet therefore has its own horseshoe vortex system, as shown in Fig. 4.18(a). At the wing-tip/winglet junction, the winglet vortex system partly cancels the wing-tip vortex, so that effectively, the main ‘tip’ vortex forms at

the tip of the winglet. This vortex is above the plane of the main wing, and so its downwash effect is reduced. In fact, the winglet modifies the whole of the spanwise distribution of trailing vorticity in a way that reduces the downwash and induced drag. In addition, the sideforce on the winglet can have a forward thrust component, as shown in Fig. 4.18(b). This also contributes to the reduc­tion in drag.

An inward sidewash occurs on the upper surface of the wing, as air is drawn in towards the low pressure. Conversely, an outward sidewash occurs on the lower surface, where air flows away from the high pressure. Thus, winglets can be fitted both above and below the wing tip. Because of the requirements of ground clearance, however, they are often only fitted above the tip.

There was some initial scepticism concerning the claimed advantages of such devices, because it seems to be a little like picking oneself up by one’s boot­straps. However, theoretical study (Yates et al. 1986) shows that they do not contravene any of the laws of nature, and that significant reductions in trailing vortex drag are possible using such devices. These theoretical predictions are well supported by experimental evidence.

Devices such as winglets are described as non-planar, because the wing is not in a single plane. A full analysis of non-planar lifting surfaces is given in Yates et al. (1986). In general, theoretical analysis indicates that for a given span, there is a wide range of non-planar wing shapes that should give less

trailing vortex drag than a simple elliptical planform wing. The monoplane with winglets, and the biplane are two examples. However, in many cases, including that of the biplane, unless there is some good reason for limiting the wing span, it is cheaper and easier to use a simple monoplane, and to reduce the drag by increasing the aspect ratio.

It should be noted that winglets will not significantly reduce the drag if added to a wing that has already been optimised for low drag. If winglets are to be used to full effect, the wing has to be designed to take account of their presence from the outset. They can also be used to reduce the drag of a wing which was not optimised for low drag.

Both sails and winglets modify the distribution of vorticity downstream of the wing, and generally inhibit the formation of a well defined vortex at the tip. This has been shown to be a useful side-effect for crop-spraying aircraft, as it prevents the spray from being lifted above the wing and blown off target by a cross-wind.

Although these devices modify the trailing vortex field in a way that has a beneficial influence on the trailing vortex drag, they do not destroy the trailing vorticity as is popularly believed. Spillman (1988) reports that in flight trials with wing-tip sails, the disturbance effects far downstream were, if anything, slightly increased.

Winglets and other devices can produce a low-drag wing, but they add to the cost and complexity of construction. They also modify the handling and stability characteristics. In one case tested, the cross-wind stability of the air­craft in landing was severely affected, and in another, interference with the flow over the ailerons produced a control reversal effect in some circumstances. Even though the influence on handling and stability may not be detrimental in all applications, the effects must be fully evaluated for certification purposes, and this can also be a costly process.

An ingenious use of winglets was made in the design of the Beech Starship (Fig. 4.10). Here, the winglet also served as the vertical fin, and is thus a necessary, rather than additional feature.

The modern use of composite construction makes it possible to design much more complex out-of-plane wing geometries, as on the Airbus A350XWB shown in Fig. 14.6.

In addition to the use of these fixed devices, drag reductions have also been obtained by using spanwise blowing (Tavella et al. 1985).

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’.

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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.

We have already had something to say about the swept wing in Chapters 2 and 8. In this section we will look at the advantages to be gained from using swept wings in the transonic speed range, emphasising how conflicting design requirements are resolved and a suitable compromise solution is reached for a particular aircraft.

We saw in Chapter 2 that sweeping the wing works because only the velocity component at right angles to the leading edge of the wing contributes to the aerodynamic performance, and so the free stream Mach number is effect­ively reduced (Fig. 2.17). For the supersonic wing this had two advantages. Firstly the strength of the bow shock wave was reduced, and secondly the char­acteristics of the wing could be made to be similar at both low and high speeds.

In transonic flow sweep works because of the same basic principle. In the case of a typical aircraft designed only for transonic cruise, however, the oncoming flow will be just below the speed of sound and sweep is used to maintain a high cruise Mach number while reducing the effective Mach num­ber seen by the wing section to a value just below the transonic drag rise.

Sweep is therefore an important weapon in the armoury of the transonic aerodynamicist, but it has its limitations. There is still a need to use relatively thin sections in order to delay the transonic drag rise as much as possible, and consequently wings tend to be quite flexible with resulting problems which are described later. Furthermore sweep introduces problems of stability (Chap­ters 11 and 12). The lift-to-drag ratio is also reduced for the reasons given in Chapter 2.

Because of such problems sweep angle is kept as low as possible and the transonic wing section is generally still a lot thinner than its subsonic counter­part. Thus the basic low speed performance of such wings is not very good and, as would be expected with a thin section, stall occurs at a comparatively low angle of attack; an effect which is made worse by the tendency for the local load at the tip of the wing to be high, as described in Chapter 2.

It is for these reasons that such aircraft, when flying at low speed, usually require to vary the section geometry by the use of leading and trailing-edge slats and flaps. Although these are expensive in terms of weight and mechanical complexity, they do permit a thin section configuration to be adapted to give reasonable performance at low speeds. The design of such devices is a complete story in itself, since there are different performance requirements at cruise, landing, take-off and during any low speed waiting (or stand off) which may be required by air traffic control.

It must be reiterated that the use of sweep simply allows us to use a higher cruise Mach number than the drag rise Mach number for the particular wing section employed. This is merely one technique which can be employed to give acceptable cruise performance and its use is coupled with ever improving detailed section design.

In our discussion above we considered the swept wing simply from the point of view of a wing of infinite span yawed to the main flow direction. In reality, as with the supersonic swept wing, there will be both a tip section and a centre section to complicate the issue.

Furthermore, the basic planform will modify the way in which the trailing vortex sheet forms (Chapter 2). The load distribution is consequently affected so that the load becomes concentrated near the tip, as we mentioned earlier. This is just what we would wish to avoid. Firstly the concentration towards the tip means that the bending moment at the root of the wing will be more severe, and secondly the increased loading peak in the tip region will make the stalling problems there even more severe.

To compound the problem, although we may have solved some of the exist­ing problems by the use of sweep, the flexural behaviour of a conventional wing structure causes the tip angle of attack to be reduced relative to the rest

of the wing (Fig. 9.13). In some ways this is a good thing since load alleviation at the tip is what is required. However this has the effect of moving the centre of pressure of the whole wing forward and consequently altering the longitud­inal trim of the aircraft.

Looking at the above catalogue of woe, the reader might be forgiven for thinking that sweep should be avoided at all costs. This is not so. It is a vital technique in the design of aircraft of this type. However the problems discussed above will, at least, indicate that it needs to be used with due caution and that it is not such a complete answer to the proverbial maiden’s prayer as it might seem at first sight.