Abstract:
The drag of a non-lifting swept wing of infinite span is investigated for supersonic flow when the Mach lines from the wing apex lie ahead of the wing leading-edge. The wing section is assumed to be arbitrary but identical over the entire wing-span. The drag is found according to the linear equations of supersonic flow by considering the flow due to a system of superposed source planes. The drag of such a wing is found to be finite and the effect of the speed of flight independent of the section shape assumed. The variation of drag with the section shape is shown to be proportional to the integral over the chord of the product of the local wing thickness and the value of the excess pressure existing in incompressible flow at the same position. The drag of wing-sections given by certain types of formula is evaluated in general terms, and some numerical results are given : the drag of sections with bluff noses and finite trailing-edge angles is generally between 0 and 15 per cent greater than the drag of a wing of the same sweep and thickness having a biconvex section. Finally, methods of reducing the drag by changing the section shape are considered.
