Abstract:
PART I. The development of the flow pattern on a swept wing with incidence and stream Mach number is described. The wing, of aspect ratio 2.828, taper ratio 0.333 and leading-edge sweep 53.5 deg, was tested at Mach numbers between 0.6 and 1.6 at incidences up to about 12 deg. The test Reynolds number varied with Mach number, being typically 2.3 x 10power6 at M 0 = 1.0. Boundary-layer transition was fixed by a roughness band at the leading edge. It is shown that the flow pattern at moderate incidences develops smoothly from a subsonic type involving leading-edge separation to a supersonic type where the flow is attached near the leading edge and with shock-induced separation further aft. The formation and movement of the shock-wave system and the vortices near the wing surface are briefly discussed. PART II. The development of the flow pattern on a wing of aspect ratio 2.828, taper ratio 0.333, leading-edge sweepback 53.5 deg and 6 per cent thickness/chord ratio in the streamwise direction has been described in Part I, which discussed oil-flow patterns obtained on the surface of the wing. The complete programme of tests also included pressure plotting at four spanwise stations and force measurements. These are discussed in relation to the flow development in this part of the Report. The wing was tested at Mach numbers between 0.6 and 1.6 for incidences up to about 14 deg. The tunnel stagnation pressure was held constant at a value near atmospheric pressure during the tests, so that the Reynolds number varied with Mach number: at M 0 = 1.0 it was 2.3 x 10power6 based on the mean aerodynamic chord. Boundary-layer transition was fixed by a roughness band at the leading edge. A detailed analysis has been made of the pressure distributions on the surface of the wing and the chordwise distributions integrated to determine the spanwise loading. The overall lift and pitching moment of the wing were also obtained from these data, as well as from direct measurements using a strain-gauge balance, by means of which the wing drag was also determined. These results are considered in some detail to illustrate the effects of Mach number and incidence on the flow about the model. A preliminary analysis is also made of the conditions for boundary-layer separation due to shock waves on the wing surface. The principal factor appears to be the component of Mach number normal to the shock front.
