Abstract:
This report puts on record, as data, pressure distributions measured on a 5-in. chord aerofoil of EC 1250 section in the 20 x 8 in. Rectangular High-Speed Tunnel, at the National Physical Laboratory. The pressure distributions given here were obtained some years ago, and give detailed results on an aerofoil which has some interesting properties but differs in shape from those now used for aircraft wings. Some discussion of the results has been made elsewhere, for example in R. & M.'s 2560 and 2222. The curves show the now well-known phenomenon of the backward movement of shock waves and spread of the supersonic region ahead of them at a fairly constant limiting local Mach number along the surface, for a symmetrical aerofoil at moderate incidences. The changes of lift, pitching moment, etc., with Mach number can be estimated from the pressure distributions. The results resemble those obtained from German tests at higher Reynold's number, but smaller incidence range.
