By H. Schuh, R. L. Bisplinghoff and W. S. Hemp (Auth.)

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If the leading edge is sharp, the shock on the upper side may be replaced by an expansion fan. If the angle of attack is small, its influence is of the same order of magnitude as the influence of thickness and may therefore be neglected. For comparatively large angles of attack, a slender wing may be replaced by a flat plate at the same angle of incidence and hence Fig. 13b may be used. The inviscid flow conditions on each side of this plate are equal to the conditions at the outer edges of the respective boundary layers.

Similar solutions for the compressible boundary layer on a yawed cylinder with transpiration cooling, NACA TN 4345 (1958). RESHOTKO, E. and BECK WITH, I. , Compressible laminar boundary layer over a yawed infinite cylinder with heat transfer and arbitrary Prandtl number, NACA Rep. 1379 (1958). VAN DRIEST, E. , Investigation of the laminar boundary layer in compressible fluids using the Crocco method, NACA TN 2597 (1952). SOMMER, C. S. and SHORT, B. , Free-flight measurements of turbulent-boundary-layer skin friction in the presence of severe aerodynamic heating at Mach numbers from 2-8 to 7-0, NACA TN 3391 (1955).

The rate of heat transfer is largest close to the narrowest cross section. A method for calculating turbulent heat transfer in a nozzle is due to Bartz, Refs. 14. 5 The influence of a non-uniform wall temperature on the heat transfer coefficient Suppose that a boundary layer develops from a leading edge and that the wall temperature Tw is equal to the adiabatic wall temperature Taw from the leading edge to a distance xx from it. Hence the rate of heat flow would be zero over this distance. At the point x1 the wall temperature FIG.

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