The formation of shock waves as a limitation in aircraft design has been considered only, as far as the author knows, in the cases of wing sections at low operating lift coefficients and of the airscrew efficiency. There is, however, a further limitation for high flying aircraft in that the small air density causes the aircraft, even at the stall, to travel at a value of Mach number (V/a) which will cause a shock stall to form on the wing. In some cases the aircraft will be able to prevent this shock stall at high C L by flying faster, but in others, due to high wing loading, the wing sections will be shock stalled at all C L values and the only remedy will be to decrease the wing loading.

In most of the problems concerned with the buckling of plates it is assumed that the length of the plate is greater than its width, and that the edges of the plate, at which the compression loads are applied, are simply supported. The method of support of the side edges vary with the actual conditions.

In the case of short wide plates or plates with many lateral stiffeners, as shown in Figs, 1 and 2 respectively, the effect on the buckling load due to the conditions of support at the two side edges will become less, and that due to the conditions of support at the end edges will become important, so that they can no longer be assumed to be simply supported. For ease of calculation the side edges are assumed to be simply supported.

In this paper a simple approximation method is presented for rapidly computing the lift distributions of arbitrary aerofoils. The numerical results are compared with those obtained by an exact method and for many purposes show a satisfactory degree of accuracy. The latter, for all practically occurring cases, can be estimated at the start of the computation work with the aid of the comparison examples given.