Note on the Flow Near the Tail of a Two-dimensional Aerofoil Moving at a Free-stream Mach Number close to Unity

This note discusses the flow near the trailing edge of a two-dimensional aerofoil moving at a free-stream Mach number close to unity, for cases where the effects of viscosity are slrall. A qualitative argument, which is supported by experimental evidence, suggests that the local Mach number downstream of the trailing-edge shock waves is approximately independent of tie-stream Mach number, a&rofoil geometry, and incidence. It follws from this result that there is a unique relationship between the flow deflection angle at the trailing edge and the local Mach number just upstream of the trailing-edge shock waves. This relationship is determined by using results obtained during wind-tunnel experiments on aerofoils df the R.X.E. series, and rray sometimes be used to give rapid estimates of the local Nach number at the trailing edge of an aerofoil in terms of the trailing-edge angle, incidence, and control angle. When the Mach number, ixmeediately ahead of the trailing-edge shock has been detezdned, the local Mach numbers over the surface ahead of the trailing edgo can be estimated by using simple-wave theory. The characteristics of straight-sided controls are considered as en exa&e.