Abstract:
A two-dimensional aerofoil of NACA 0015 section was tested at zero incidence in the Royal Aircraft Establishment 10 ft x 7 ft High-speed Wind Tunnel and measurements were made of (a) Static pressure on the aerofoil surface at Reynolds numbers of 1.4 x 10power6 to 5.5 x 10power6 (b) Static pressure on the aerofoil surface, on the tunnel walls and in the stream between the aerofoil and the walls at R = 2.8 x 10power6. All the tests were made at Mach numbers of 0.7 upwards and were continued past the choking Mach number of 0.764 until either the maximum permissible fan speed was reached or the maximum available power was being used. The results showed that the choking Mach number was about 0.764 at Reynolds numbers from 1.4 x 10power6 to 2.8 x 10power6. Above M = 0.760 the development of the supersonic region towards the walls was extremely rapid in terms of tunnel Mach number. At M = 0.761 the sonic line was only about half-way out to the tunnel walls and at M = 0.764 it had reached them. Before and during choking quite large changes in the aerofoil pressure distributions were produced by varying the Reynolds number. At M = 0.73 and 0.75 the shape of the pressure distribution curves indicated the possibility of a λ-shock at the lower Reynolds numbers and a single shock at the higher Reynolds numbers.
