Supercritical aerofoils for transonic flow

The conventional section, as we have seen, relies heavily on the leading-edge suction peak to develop lift. This means that most of the lift is developed at the front of the section and relatively little at the rear. One way of improving the situation, without incurring the penalty of high local Mach numbers, is to ‘spread’ the load peak and load up the rear of the aerofoil thus producing the type of distribution shown in Fig. 9.5, the so-called ‘roof top’ distribution.

The way in which this is done is to reduce the camber at the front of the aerofoil (the camber may even be negative here) and to increase it towards the rear. This gives the typical section shown in Fig. 9.5. In this way the locally high Mach numbers and strong shock waves associated with a conventional section can be avoided.

In Chapter 5 we saw that a region of shock-free compression can exist in a flow provided Mach lines drawn within the compressive region do not con­verge. An example of this was given in Fig. 5.15 where a shock-free (so-called isentropic) compression region is shown near a smooth corner on a surface. The same technique can be used to avoid the formation of shock waves in the recompression of the flow over a supercritical aerofoil.

In this case the process is complicated by the existence of the subsonic flow outside the local supersonic ‘patch’. Complex wave reflections will occur both at the sonic boundary between the two areas as well as from the aerofoil

Pressure lower than surrounding atmosphere

Pressure greater than surrounding atmosphere

Fig. 9.5 ‘Roof top’ pressure distribution

Local surface Mach number is close to 1.0 between A and B

surface itself (Fig. 9.6). The local surface slope in the compression region must therefore be carefully designed to suppress any tendency for the compressive wave to coalesce into a shock wave within the supersonic ‘patch’.

In Fig. 5.15 the formation of the shock wave away from the surface is inevitable because the flow is supersonic everywhere. The waves generated in the supersonic region over our aerofoil can, however, reach a region of sub­sonic flow before they run together, if we get the design right. In this case no shock wave will be formed.

The process of design is much more complicated than may appear from the above account. While a satisfactory solution may be obtainable for a single design point it will be necessary to ensure that the off design flow is both stable and not subject to large drag rises. Careful design will also be required to prevent adverse shock/boundary layer interactions in the off design condition. Because of this such aerofoils do not usually run with entirely shock-free recompression but the supersonic region of flow is terminated by a near normal

atmosphere

Fig. 9.7 Peaky pressure distribution

Flow on top surface is supersonic up to weak shock wave shock wave of low strength. This feature may well improve buffet behaviour which will be discussed shortly.

With improved computational methods the design of supercritical aerofoils is advancing rapidly. Figure 9.7 shows the pressure distribution over a modern supercritical aerofoil similar to that used on the A320 Airbus. A large area of supersonic flow is employed over the top surface ending with an almost shock – free compression so that losses are kept low. This means that the local loading in this area can be higher, leading to a somewhat more ‘peaky’ distribution than the ‘roof top’ distribution shown in Fig. 9.5.

The aerodynamic problems involved in designing supercritical sections with suitable ‘off-design’ performance are severe. They are an example of one area in which the use of computers in the solution of the basic equations of the air flow has produced dramatic results.