It has long been the dream of aircraft designers to produce civil airliners with no separate tail or fuselage, as with the B2 Spirit bomber (Fig. 4.19). The advantages would include much lower aerodynamic drag, and reduced weight. There are, however, several problems. Much of the structural load on a civil aircraft derives from the stresses due to pressurisation of the cabin, and by far the most efficient cross-sectional shape is a circle. Horizontally-arranged double or multiple bubble arrangements may be used, but passenger access between the bubbles then becomes an issue. Longitudinal stability considerations mean that the range of centre of gravity positions is relatively restricted, so passenger movements might need to be controlled. There are also difficulties involved in access and in the placing of passenger external view windows. None of these problems is insuperable, but the real constraint would be the very high costs of such a radical development.
Category Aircraft Flight
The stability criteria for a canard or tail-first configuration aircraft (Fig. 11.8) are essentially the same as for a conventional one. When the aircraft is trimmed, the forward wing (foreplane) should be arranged to generate a higher lift coefficient than the rearward wing (main-plane). The foreplane is therefore usually set at a higher geometric incidence than the main-plane, thus giving longitudinal dihedral. On a canard it is the larger rear wing surface that generates
Fig. 11.8 A stable canard arrangement
The aircraft has to be trimmed with the foreplane generating a higher lift coefficient than the main-plane. The foreplane is therefore normally set at a higher incidence.
most of the lift, so it follows that on a stable canard, both surfaces must be producing lift.
Since both surfaces on a canard produce positive lift, the overall wing area, total weight, and drag can all be lower than for the conventional arrangement. Also, as we have already mentioned, pitch control is achieved by lifting the nose by increasing the foreplane lift, rather than by pushing the tail down. This shortens the take-off run, and generally improves the pitch control characteristics. The manoeuvrability of the canard configuration is one of the features that makes it attractive for interceptor aircraft (see Figs 10.1 and 10.8).
Another claimed advantage of the canard is, that since the foreplane is at a higher angle of attack than the main-plane, the foreplane will stall before the main-plane, thus making such aircraft virtually unstallable. Unfortunately in violent manoeuvres, or highly turbulent conditions this may not be true, and once both planes stall, recovery may be impossible, because neither surface can be used to produce any control effect.
The main problems with the canard configuration stem from interference effects between the foreplane wake and the main wing. In particular, the down – wash from the foreplane tilts the main wing resultant force vector backwards, thus increasing the drag. By careful design, however, the advantages can be made to outweigh the disadvantages, and highly successful canard designs by Burt Rutan such as the Vari-Eze shown in Fig. 4.20 provoked renewed interest in the concept.
For forward-swept wings, as on the X-29 shown in Fig. 9.20, the foreplane interference can be a positive benefit, as the downwash suppresses the tendency of the inboard wing section to stall at high angles of attack.
For pressurised passenger aircraft, the canard arrangement has the added advantage that the main wing spar can pass behind the pressure cabin, as in the Beech Starship shown in Fig. 4.10. A problem remains in that, unless there is a rearward extension of the fuselage, the fin (vertical stabiliser) may have to be large to compensate for the fact that it is not very far aft of the centre of gravity.
Thus far we have considered the landing manoeuvre for aircraft operating from conventional runways. Within this group we include special short take-off and landing (STOL) aircraft such as the C-17 (Fig. 10.20), since the techniques employed are essentially similar.
Sometimes aircraft are required to have a shorter landing run than is obtainable by conventional means, as for example in carrier landing. Although the carrier can help by sailing into the wind as fast as possible, the deck is short, and additional deceleration has to be provided by an arrester hook which
Fig. 13.12 The A380 landing
Note the large number of wheels required because of the massive weight (Photo courtesy of R. Wilkinson)
engages with a wire across the deck. The ultimate in landing performance is of course provided by the vertical take-off and landing (VTOL) Harrier (Fig. 7.12) or Osprey (Fig. 1.30).
At the other extreme the Space Shuttle (Fig. 8.19) commenced its approach without power at hypersonic speed. We looked at the high speed part of the landing manoeuvre in Chapter 8. The final approach, however, was very similar to those we have already dealt with, except that there was no longer the option to fly down the glide path under power. The lack of this ability means that it was not possible to be nearly so precise in achieving a particular touchdown point, with the result that a long runway was needed. Since the whole of the re-entry and landing manoeuvre was unpowered accurate computer control was needed right from the point of re-entry if the Shuttle was to end up in the right continent, let alone the right airfield.
On older aircraft types, it is normally necessary to avoid flow separation and stalling, since it is very difficult to maintain proper control in the stalled
Fig. 1.19 Flow separation
At high angles of attack, as in the lower photograph, the flow no longer follows the contours of the upper surface, but ‘separates’, producing a highly turbulent recirculating region of flow (Photo courtesy of ENSAM, Paris)
condition. However, from Fig. 1.17, you will see that after an initial drop at stall, the lift starts to rise again at high angles of attack. For thin wings, the highest value of CL may indeed be obtained in the stalled condition. The overall aircraft lift is further increased by the fact that at these high angles of attack, the engine thrust begins to add a significant component to the lift. Such high lift can be a considerable advantage to combat aircraft performing violent manoeuvres, since it can be used to produce a large (centripetal) force for rapid pull-out from a dive. Alternatively, by rolling the aircraft on its side, the lift can be used to produce the cornering (centripetal) force for a rapid turn.
On missiles, where there is no loading on the pilot to consider, it is normal to make full use of this extended capability; indeed, missiles may spend short periods actually flying backwards after a sharp turn. In rapid manoeuvres, and with large amounts of available thrust, the high drag produced is unimportant.
The main difficulty of flight in separated flow is one of stability and control. The lift, drag, and most importantly, the position of the centre of lift, all vary rapidly. To overcome this problem, the aircraft may need artificial stability in the form of a quick-acting automatic control system. The development of reliable microelectronic systems has meant that it is now possible to fly in what would have previously been considered to be a highly unstable and dangerous condition. Recent combat aircraft have demonstrated controlled flight at angles of attack of more than 70°.
For military aircraft particularly, flight with separated flow provides considerable rewards in terms of improvements in both performance and manoeuvrability. However, even though it may be possible to control the aircraft in the stalled condition, the instability of the separated flow may still cause structural problems due to excessive buffeting. One solution is to control or stabilise the separated flow as described below.
On a swept wing, the pressure gradients are such that they cause the boundary layer to thicken towards the wing tips. Thus, unless corrective measures are taken, the flow is likely to separate near the tips before any other part of the wing. This is in addition to the inherent tip-stall tendency of swept wings due to upwash, described in Chapter 2. For moderately swept wings at high angles of attack, the outboard stalling is exacerbated by the formation of leading-edge conical vortices which curve inwards, away from the tips, as shown in Fig. 2.20.
One way to alleviate the problem, is to fit chordwise fences on the wing, as shown in Fig. 3.8(a) and Fig. 3.9. Wing fences effectively split the wing into separate sections and help to prevent spanwise thickening of the boundary layer. At the fence, a trailing vortex is shed, rotating in the opposite sense to the usual wing-tip trailing vortex. The vortex produced by the fence scours away the boundary layer locally.
Fig. 3.8 Devices for inhibiting flow separation on swept wings
(a) Wing fence (b) Vortilon (c) Saw-tooth leading edge
Fig. 3.9 A wing fence on an early jet transport
The fence helps to prevent the spanwise thickening of the boundary layer on a swept wing partly by inhibiting the spanwise flow, and partly by generating a vortex which draws in the slow-moving air of the boundary layer
It was found that this trailing vortex also had the useful effect of stabilising the position of the leading-edge conical vortices which form at high angles of attack, thereby tending to improve the stability and control near the onset of stall.
The fence need not extend over the whole chord, and the short leading-edge fence shown in Fig. 3.9 and Fig. 3.8(a) was a device used on many early swept wing aircraft.
The vortilon shown in Figs 3.8(b) and 3.10 is a small fence-like surface extending in front of the wing and attached to the under-surface close to the stagnation line. It is intended to generate a vortex over the upper surface, but only at high angles of attack, when it is most needed. Engine mounting pylons can conveniently be used for the same purpose.
In the saw-tooth leading-edge design shown in Fig. 3.11, the abrupt change of chord causes a strong trailing vortex to form at this point. A trailing vortex is formed wherever there is an abrupt change of wing geometry.
On forward-swept wings, the boundary layer tends to thicken towards the inboard end, encouraging the centre section to stall first. Although this is a safer characteristic than tip-stall, it still produces a diverging nose-up pitching
Fig. 3.10 The vortilon is intended to generate a vortex at high angles of attack. The vortex inhibits the spanwise thickening of the boundary layer, and helps to stabilise the position of the separated leading-edge vortex
Fig. 3.11 The saw-tooth leading edge also produces a vortex
Fig. 3.12 Inboard strakes on this model of a forward-swept-wing aircraft help prevent flow separation at the wing root
moment, and preventative measures are necessary. In the forward-swept model shown in Fig. 3.12, inboard strakes have been added so that the inboard section behaves like a slender delta, and does not stall in the conventional sense. The strong separated vortex also helps remove the thick boundary layer. On the forward-swept X-29 (Fig. 9.20) the downwash and trailing vortices produced by ‘canard’ foreplanes are used to inhibit inboard separation.
Let us look once more at the nose of our supersonic aircraft. We saw how the shock waves formed in front of it, slowing the air down almost instantaneously and providing a subsonic patch through which the pressure information could propagate a limited distance upstream at the speed of sound (Fig. 5.2). It should be noted that the shock wave itself is able to make headway against the oncoming stream above the speed of sound. Only weak pressure disturbances travel at the speed of sound. The stronger the shock wave is, the faster it can travel through the air.
Considering the problem from the point of view of a stream of air approaching a stationary aircraft, this means that the faster the oncoming stream, the stronger the shock wave at the nose becomes. Thus the changes in pressure, density, temperature and velocity which occur through the shock wave all increase with increasing air speed upstream of the shock wave. A mathematical analysis of the problem shows that the strength of the shock wave, expressed as the ratio of the pressure in front of the wave to that behind, depends solely on the Mach number of the approaching air stream.
If we now stand further back from the aircraft we see that the bow shock wave which forms over the nose is, in fact, curved (Fig. 5.3(a)). As we get further from the nose tip so the shock wave becomes inclined to the direction of the oncoming flow. In this region the shock wave is said to be oblique. At the nose, where it is at right angles to the oncoming flow, it is said to be a normal shock wave.
The oblique shock wave acts in the same way as the normal wave except that it only affects the component of velocity at right angles to itself. The component of velocity parallel to the wave is completely unaffected. This means that the direction of the flow is changed by an oblique shock (Fig. 5.6) whereas
Fig. 5.6 Flow deflection by oblique shock wave
Tangential component Vt remains unchanged but V„2 < Vn
Fig. 5.7 Flow deflection through bow shock wave
Deflection reaches a maximum and then reduces again
it is unaffected by a normal shock. In both cases, however, the magnitude of the velocity is reduced as the flow passes through the shock wave.
Looking more carefully at the effect of the bow shock wave (Fig. 5.7) we see that, in general, the same flow deflection can be obtained by two possible angles of oblique wave. The reason for this is given in Fig. 5.8. The wave of greater angle at A is stronger because the velocity component normal to the wave front is greater. It therefore changes the oncoming velocity component more than the weaker wave at point B.
Adding the resulting velocity components immediately downstream of the shock waves at two points (Fig. 5.8) shows how a particular point B (where the shock wave is weak) can be chosen with exactly the same flow deflection as at A (with a strong shock wave).
It should also be noted that for a normal shock wave the downstream flow is always subsonic, as it is for most strong oblique waves. The fact that the
Fig. 5.8 Weak and strong shock waves
Strong shock at A gives same deflection as weak shock at B, but greater pressure jump since V2 < V2
velocity component parallel to the wave is not changed means, however, that the flow downstream of the weak oblique wave is supersonic.
In the gas turbine, the burning process causes the air to be heated at virtually constant pressure, in constrast to the piston engine, where the air is heated in an almost constant volume with rapidly rising pressure. The (thermodynamic) efficiency of both types of engine can be shown to depend on the pressure ratio during the initial compression process. Increasing the pressure ratio increases the maximum temperature, and the efficiency is, therefore, limited by the maximum temperature that the materials of the hottest part of the engine can withstand.
The temperature limitation is rather more severe in the gas turbine, since the maximum temperature is sustained continuously, whereas in the piston engine, it is only reached for a fraction of a second during each cycle. For a long time, this factor led to a belief that the gas turbine was so inherently inefficient in comparison with a reciprocating engine, that it was not worth bothering with.
At high altitude, the atmospheric air temperature is reduced, so for a given compressor outlet temperature, a greater temperature and pressure ratio between inlet and outlet can be allowed. Thus, the thermodynamic efficiency tends to rise with increasing altitude. This factor, coupled with the advantages of high altitude flight, described in Chapter 7, makes the high speed turbo-jet – propelled aircraft a surprisingly efficient form of transport. In fact, as we show in Chapter 7, for long-range subsonic jet-propelled transport, there is no economic advantage in using an aircraft designed to fly slowly.
The thermodynamic efficiency of gas turbines improved dramatically during the first three decades of development mainly because of progress in producing materials capable of sustaining high temperatures, improvements in the cooling of critical components, and better aerodynamic design of compressors and turbines.
Figure 7.11 shows the forces acting on an aircraft in a steady climb. If the climb is steady then there can be no net force acting on the aircraft either along the flightpath, or at right angles to it. If we consider the forces acting along the flightpath we can see (Fig. 7.11) that the sine of the climb angle is given by the difference between thrust and drag divided by the aircraft weight. Thus to operate at the maximum angle of climb possible we need the biggest possible value of thrust minus drag.
If the thrust minus the drag is equal to the weight we have a vertical climb, e. g. the Harrier (Fig. 7.12). If thrust minus drag is greater than the weight then the aircraft will be in an accelerating, rather than a steady climb.
If, however the difference between thrust and drag is less than the aircraft weight, some lift must still be provided by the wings. To be able to climb at all the aircraft must be operating at a height at which the engine is capable of producing more thrust than the drag of the aircraft.
If, for instance, the aircraft is flying straight and level initially we can plot the now familiar variation of drag with flying speed. Let us suppose that the
Fig. 7.13 Climbing flight
Increased throttle setting gives excess of thrust over drag for climb Best climb angle is obtained when thrust minus drag is maximum
aircraft is operating at point A on this curve. An increase in throttle setting will give an available thrust-minus-drag difference for climb as shown (Fig. 7.13). If we know the engine characteristics at the new throttle setting we can optimise the airspeed to give the best possible thrust/drag difference.
Here we must turn our attention to the type of powerplant being used once again. If we are dealing with a turbo-jet and thrust will not vary very much with speed in the operating range we are considering. All we need to do therefore is to gratefully accept the maximum thrust that the engine will give and fly at the speed which produces the least amount of drag (point A in Fig. 7.14).
If we are using a piston engine/propeller combination, we have already seen that the thrust falls with increasing speed and so we must reach a compromise between the requirements of airframe and powerplant and operate at a speed somewhat lower than the minimum drag speed in order to achieve the maximum angle of climb (Fig. 7.15).
At this point a word of caution is necessary. We have estimated the best climbing angle using the drag curves derived for straight and level flight. When the aircraft is climbing examination of the forces normal to the flightpath (Fig. 7.11) shows that the lift developed by the wing will be reduced by a factor equal to the cosine of the climb angle and is thus no longer equal to the aircraft weight. Our drag curve will therefore need to be modified and this, in turn, may change the best speed for climb.
A large number of aircraft, such as civil airliners and military transport aircraft, are not required to indulge in particularly violent manouevres. Although the rate of climb might be quite high, because the forward speed is also high, the angle of climb is frequently not very great. In such cases our original approximation will not be too far from the truth.
When an aircraft is banked (turned about the roll axis), the resulting forces produce a tendency to sideslip, as illustrated in Fig. 10.11. In sideslip motion, the fin produces a sideforce and hence a yawing moment, as shown in Fig. 10.11. Thus, banking an aircraft will cause it to turn towards the direction
Fig. 10.11 Sideslip and yaw due to roll
When an aircraft rolls, one component of weight acts sideways relative to the aircraft axes. This causes the aircraft to slip sideways. Once the sideslip develops, the fin will generate a sideforce tending both to right the aircraft and to yaw it towards the direction of the sideslip of the lower wing, unless compensated for by applying opposite rudder. This is another example of the cross-coupling between motions.
Mounting the wings well above the centre of gravity aids roll stability, but not for the reasons often assumed. Figure 11.16(a) shows a high-winged aircraft which is rolling, but has not yet developed a sideslip. It will be seen that both the lift and weight forces pass through the centre of gravity, so there is no restoring moment. The fuselage does not swing like a pendulum under the wing, as is often incorrectly believed. Once the sideslip commences as in Fig. 11.6(b) the wing becomes yawed to the resultant flow direction and the lower wing tends to generate increased lift due to the onset of vortical lift at the
Fig. 11.17 The stabilising effect of sweep-back
If a swept-winged aircraft rolls, and tends to sideslip, the effective span of the leading wing will be greater than that of the other. This produces a righting moment
tip. Also, the cross flow on the fuselage, due to sideslip, produces an upwash on the lower wing and a downwash on the upper wing. There may also be a slight sideways drag component. As illustrated, the resulting force no longer passes through the centre of gravity, and a restoring moment is produced. The lower the centre of gravity is, the greater will be the moment arm. Thus, highwinged aircraft do not need so much dihedral as low-wing types, and may even need none at all.
The use of wing sweep also enhances roll stability, as may be seen from Fig. 11.17. When a sideslip occurs, the lower wing presents a larger span as seen from the direction of the approaching air, and as with dihedral, the effect is to roll the aircraft back towards the horizontal.
Excessive rolling stability can produce undesirable dynamic instabilities due to cross-coupling between roll and yaw modes, such as in the Dutch roll described in Chapter 12. Swept-wing aircraft, therefore, often have negative dihedral, which is known as anhedral. Anhedral is often found on swept-winged aircraft that are also high-winged, as on the Antonov shown in Fig. 12.13.