An investigation of the flow about a plane half-wing of cropped-delta planform and 6 per cent symmetrical section at stream mach numbers between 0.8 and 1.41

Abstract

Summary. A study has been made of the flow development over the wing as the incidence and stream Mach number vary and this is illustrated by surface pressure distributions and oil-flow patterns. The growth and movement of the two main surface shocks (the rear and forward shocks) is discussed, and conditions for flow separation through these shocks are considered. For the rear shock, which has little sweep, these conditions are similar to those for shock-induced separation on two-dimensional aerofoils. The forward shock is comparatively highly swept and separation seems to correspond to two rather different but simultaneously-attained conditions, one related to the component Mach number normal to the shock front and the other to the position of the reattachment line. The flow in the region between the leading edge and the forward shock is shown to have certain characteristics analogous to those found upstream of the shock on two-dimensional aerofoils. To the rear of the forward shock, but ahead of the rear shock, the flow at low supersonic speeds resembles in some respects that about a simple cone. The general flow development is related in the text to the wing lift and pitching moment, and the drag. The first two are most affected by the aft movement of the rear shock, which also stimulates the transonic drag rise. The lift-dependent drag is shown to be influenced by the appearance of leading-edge separation and possibly also by some stage in the development of the forward shock. The flow over the cropped-delta planform is noteworthy for the absence of the strong outboard shock and this is attributed partly to the cropped tip and partly to the unswept trailing edge. A comparison is made with results obtained during preliminary tests in which the wing planform closely resembled that of a true delta.