Abstract:
Tests have been made in the N.P.L. 18 in. by 14 in. High-Speed Tunnel at stream Mach numbers between 0.6 and 1.2 on two finite wings of identical planform and having single-wedge and double-wedge sections of 14% and 7% thickness/chord ratio respectively. The wings were untapered and the sweepback could be set at five values: 0°, 15°, 30°, 45° and 60°. The test Reynolds number based on the streamwise chord varied with stream Mach number and wing sweepback; and was between 1.3 x 10power6 and 3.4 x 10power6. The experiment was intended to assist in assessing the validity of the simple sweepback concepts currently in use, particularly for predicting the flow about infinite sweptback wings from two-dimensional data. The three-dimensional effects present are reduced to some extent by the type of section used, especially at transonic speeds, and the measured pressure distributions a): a particular spanwise station, and the associated pressure forces (normal and chordwise) correlate quite well on the basis of the simple theory. In addition it is shown that the general flow development, including the initial growth of the leading-edge separation, and the transonic flow-attachment about the leading edge also correlate satisfactorily, as do the pressures measured on the base of the single-wedge wing. There is some discussion on the effect of shock sweep on the conditions required for shock-induced boundary-layer separation and a modification is suggested to the curve put forward tentatively in Ref. 8.
