Category Aircraft Flight

Dynamic pressure

The quantity – pV2 is usually referred to as the dynamic pressure. There is a more precise definition of dynamic pressure, but this need not concern us now. Although it has the same units as pressure, dynamic pressure actually represents the kinetic energy of a unit volume (e. g. 1 cubic metre) of air.

Aerodynamic forces such as lift and drag are directly dependent on the dynamic pressure. It is, therefore, a factor that crops up frequently, and for simplicity, it is often denoted by the letter q. Pilots sometimes talk of flying at ‘high q’, meaning high dynamic pressure.

The other term in the expression, the pressure p, is often referred to as the static pressure.

Biplanes and multiplanes

It is tempting to ignore biplanes and multiplanes as being of purely historical interest, but old ideas have a habit of returning, and small biplanes have once again become popular for aerobatic and sport flying.

The wings of early aircraft had little or no bending stiffness, and had to be supported by external wires and struts. The biplane configuration provided a simple convenient and light structural arrangement, which was originally its main attraction.

A biplane produces virtually the same amount of lift as a monoplane with the same total wing area and aspect ratio. The biplane, however, has the smaller overall span, which makes it more manoeuvrable. The highly aerobatic Pitts Special, shown in Fig. 2.26, is an example of a modern biplane which is not merely a gimmick. The manoeuvrability of biplanes was one factor that led to their retention even when improvements in structural design had removed the necessity for external bracing.

Wing-tip shape

Reductions in drag can also be obtained by careful attention to the shape of the wing tip. This is particularly true in the case of aircraft with untapered wings. Although untapered wings are not the best shape in terms of minimising drag, they are often used on light aircraft because of their relative simplicity of con­struction, and their docile handling characteristics. (The inboard section tends to stall first.)

Two simple approaches; the bent and the straight-cut tip are illustrated in Fig. 4.13. Both of these tip designs are said to reduce drag by producing

Wing-tip shape

Fig. 4.11 Influence of lifting fuselage on lift distribution and drag

(a) Fuselages of cylindrical cross-section produce little or no lift, so there is a gap in the lift distribution at the centre (b) By using a lifting fuselage shape, the lift distribution can be brought closer to the optimum for low induced drag

separation of the spanwise flow at the tip, resulting in a beneficial modification of the tip flow-field. It should be noted, however, that unusual tip shapes are often intended primarily to inhibit tip stall, rather than reduce drag. Upward bent tips are evident on the Aerospatiale Robin shown in Fig. 4.14.

Matching propeller to engine

For aerodynamic efficiency large slowly rotating propellers are preferable, but unfortunately, small piston engines develop their best power-to-weight ratio at relatively high rotational speeds. For light aircraft, therefore, the added cost, complexity, weight and mechanical losses of gearing sometimes make it pre­ferable to use direct drive, and accept a slight degradation in engine efficiency due to running at low rotational speed. When small automotive engines are adapted for home-built aircraft, some form of gearing is often used. In the case of turbo-prop propulsion, the rotational speed of the primary engine shaft is so high that gearing is almost essential.

Once gearing is accepted, then the propeller diameter is limited only by prac­tical considerations such as ground clearance, so highly efficient propellers can be used. The propellers of the Lockheed Super Hercules shown in Fig. 6.6 are driven by geared gas-turbines.

Effect of altitude on the drag curve

The effect of altitude on the drag curve is very similar. As the altitude increases the density is reduced and this can be compensated by an increase in cruising speed to keep the dynamic pressure constant. If the aircraft attitude is kept con­stant both lift and drag coefficients will remain constant as before, and the drag curve will be shifted to the right in exactly the same way as before (Fig. 7.5).

Maximum speed

The maximum speed in level flight that can be attained by the aircraft can be deduced very simply from Fig. 7.4. In order to achieve the maximum speed we need the intersection between the drag and engine thrust curves to be as far to the right as possible. This is clearly obtained when the engine is at the maximum throttle setting.

This seems to be a very simple situation, but a word of caution is necessary. We have assumed a comparatively simple form for the drag curve in our dis­cussion. Compressibility effects may have an important influence on this for a particular aircraft. Such factors as the buffet boundary (Chapter 9) may then limit the maximum speed. High speed aircraft may also be limited by the maximum permissible structural temperature, which may be approached due to kinetic heating effects (Chapter 8). These factors may restrict the permitted maximum to a value below that which would be suggested by the simple ‘avail­able thrust’ criterion. Additional limitations may be imposed by the constraints on the engine operating conditions (Chapter 6).

Increasing the wing loading has the primary effect of shifting the whole of the drag curve to a higher speed (Fig. 7.5) without increasing the drag itself. Therefore a high wing loading, and consequently small wings, is desirable from the point of view of obtaining high speed.

A similar argument might lead the reader to suppose that high altitude is also desirable for high speed. To some extent this is true but it must be remem­bered that increase in altitude implies a reduction in temperature, and thus a lower speed of sound. This means that the flight Mach number will be increased for a given air speed at high altitude. Compressibility effects will therefore be apparent at a lower air speed and this will impose an important restriction, par­ticularly for aircraft designed for operation at subsonic or transonic speeds.

Wing sections in transonic flow

The conventional aerofoil revisited

In Chapter 5 we saw how the flow characteristics over a conventional aerofoil changed with increasing free-stream Mach number from a shock-free low speed flow (Fig. 5.18(a)) through the developing shock wave system at tran­sonic speeds (Fig. 5.18(b)) until the fully developed shock system is obtained at higher Mach numbers (Fig. 5.18(c)). In transonic aircraft we are particularly concerned with the intermediate type of flow shown in Fig. 5.18(b) in which the oncoming flow is still subsonic.

First let us take another look at the pressure distribution on a conventional aerofoil section (this is shown again in Fig. 9.4) and how this relates to the flow is shown in Fig. 5.18(b). We see at once that there are two potential problems. First there is a very high suction peak which occurs locally near the leading edge of the aerofoil. This means very high velocities in this region, and con­sequently high Mach numbers. The second problem occurs because of the very high adverse pressure gradient on the downstream side of this suction peak. This is liable to coalesce into a relatively strong shock wave (the shock wave which terminates the supersonic patch in Fig. 5.18(b)) and this may also induce boundary layer separation, with all the problems that entails!

Fig. 9.4 Low speed aerofoil pressure distribution

Mach number below 1.0 over surface

Note leading edge suction peak and adverse pressure gradient on top surface

Thin sections

The increase in the surface velocity over the aerofoil section is caused by two factors – the thickness of the section and its angle of attack. Thus one way in which the local Mach number over the top can be limited is to use a thin sec­tion. This has certain aerodynamic penalties associated with it, however, as we have already seen in Chapter 2. Firstly the range of angle of attack over which the wing will operate without stalling will be reduced, and secondly it is obvi­ous that the problems of fitting in a satisfactory wing structure get more and more severe as the section thickness is reduced (Chapter 14).

Supercritical sections

So far we have attacked the problem of developing a wing section suitable for transonic flight simply by using as thin a section as we can in order to limit the velocity increase due to thickness. However, as we get near to the speed of sound, the achievable wing loading is limited unless the flow becomes locally supersonic. We therefore have to design supercritical aerofoils in which this supersonic flow is adequately catered for.

Control harmonisation

For an aircraft to be comfortable and easy to fly, all of the primary control actions should require roughly the same amount of effort to operate them. The correct harmonisation of controls is often difficult to achieve with manual controls, but with the powered systems they can normally be tuned with precision.

Engine control

The power output of an aircraft piston engine is controlled in much the same way as a road vehicle engine, by means of a throttle, which varies the amount of air/fuel mixture admitted to the engine. A mixture control lever is used to give a rich fuel/air mixture for an extra, but inefficient boost of power for take­off and to adjust for air density changes.

In addition, most propeller-driven aircraft are fitted with an rpm control lever, which is used to set the propeller, and hence, engine speed.

On turbocharged engines, a means of varying the boost pressure may be provided, although in some installations, the process is automatically controlled.

The correct setting of the various controls depends on the chosen flight plan, and a good pilot will work out the best settings for each stage before take-off.

It is important to note that on a reciprocating engine, the rate of fuel con­sumption depends mainly on the power output. The pilot’s primary control of the power is by means of the throttle lever.

In gas-turbine systems, the primary engine control operated by the pilot is the fuel flow control. This serves a similar purpose to the throttle on a piston engine, except that in the gas turbine, it controls the thrust produced. The thrust and engine speed of a gas-turbine system cannot be separately varied to any significant extent, and any movable vanes, nozzles or surfaces are primarily used to fine-tune the operation of various components.

On more complex gas-turbine systems, with adjustable nozzles and multiple spools, there are many variables to control and monitor, and some form of automatic engine management control system is necessary. The pilot’s primary input is still via a single lever (or set of levers in a multi-engined installation). Further controls are required for reversed thrust, and reheat, where fitted. A host of minor controls can also be found, depending on the particular aircraft type. In the turbo-prop, there is also a means of selecting the propeller rpm.

For the pilot, the most obvious difference between a piston engine and a turbo-jet installation is the lack of torque reaction, and the relatively slow throttle response of the turbine. Changes in thrust and speed have to be anticipated much more carefully in the latter case. The lack of propeller drag braking effect can also make jet-propelled aircraft more difficult to handle.

Artificial stability – Mach trimmers and yaw dampers

In principle the pilot can control an unstable motion, by operating the controls directly to provide suitable forces and moments to oppose the motion. For example, in the case of the Dutch roll, the rudder is extremely effective in suppressing the yaw and hence controlling the motion. If the motion is of high frequency and poorly damped, however, this makes the aircraft very tiring to fly, and at some frequencies the pilot’s reactions will be such that he will not be able to ‘follow’ the motion correctly. In this event his efforts may well make a bad situation worse.

One way of overcoming the problem is to relieve the pilot of this part of his task altogether by the use of an automatic control system. In the case of the Dutch roll, the yawing motion can be sensed, both in terms of the degree of yaw and the rate at which it is developing, by the use of gyroscopically based instruments. In this case a position gyro can be used to sense the degree of yaw and a rate gyro to sense its rate. Once the information concerning the aircraft motion is available the rudder can be moved automatically to provide the required correction. Such a device is present on all large modern jet trans­port aircraft and is known as a yaw damper.

Details concerning the design of either the gyros or the damper control system are outside the scope of this book, however it is perhaps interest­ing to mention a few features which must be considered before leaving the subject.

One obvious feature is that the control system employed in the yaw damper must be able to distinguish between a conscious control input on the part of the pilot, and the control movement generated as a result of the unwanted motion. Thus the total movement must be determined as a combination of both inputs. Another point which must be carefully considered is the integrity of the control system. Should failure occur, the safety of the aircraft must not be comprom­ised. This means that either suitable back-up must be provided, or the system must revert to full manual control on failure. In the latter event the character­istics of the aircraft must be such that manual flight is reasonably possible, even if not very pleasant.

Further damping can be provided by the use of a similar system to control the ailerons in such a way as to oppose the rolling component of the motion. This system is known as a ‘roll damper’.

As mentioned above, the longitudinal characteristics deteriorate due to the rapid centre of pressure movement which results from comparatively small changes in the aircraft operating condition in transonic flight. This again can be ‘fixed’ by the use of a suitable automatic control system. This system uses elevator movement to compensate for the change in centre of pressure and is known as a ‘Mach trimmer’.

Flight with separated flow

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

Flight with separated flow

Flight with separated flow

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.

Boundary layer and stalling problems on swept wings

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.

Boundary layer and stalling problems on swept wings

Boundary layer and stalling problems on swept wings

Fig. 3.8 Devices for inhibiting flow separation on swept wings

(a) Wing fence (b) Vortilon (c) Saw-tooth leading edge

Boundary layer and stalling problems on swept wings

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

Boundary layer and stalling problems on swept wings

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

Boundary layer and stalling problems on swept wings

Fig. 3.11 The saw-tooth leading edge also produces a vortex

Boundary layer and stalling problems on swept wings

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.