Transonic tunnel tests on a 6 percent thick, warped 55 degrees sweptback-wing model

Abstract

Tests have been made in the A.R.A. 9 ft x 8 ft Transonic Tunnel on a model having a 6% thick, 55° sweptback wing with a warp distribution designed to give a constant spanwise C m - distribution and a triangular chordwise load at C L = 0.15, M = 1.2. The wing-body junction was designed according to supersonic area rule for this Mach number. The results can be considered as encouraging. Subcritical-type flow is maintained over most of the wing under the design conditions. The margins in both Mach number at the design C L and in C L at the design Mach number before the start of any serious supercritical increase in drag or before the appearance of any significant shock-induced separations are of the order of 0.05. Major changes in the pitching-moment characteristics are even further delayed. At subsonic speeds and low C L, K is near 1.2 while at the design conditions, the approximate value from the experimental results is K = 1.55 as compared with a theoretical prediction of K = 1.33. The results suggest that it is unlikely that there was a sizeable sweep factor on the wing skin-friction drag although no firm conclusion can be drawn about this. Translating the wing-fuselage drag resuks to a typical full-scale Reynolds number, L/D for C m = 0.15 varied from about 12 at M = 1.2/1.25 to about 9 at M = 1.4. Some incremental drags are also given for a particular fin-tail unit but these are not necessarily representative of a full-scale aircraft since the fin-tail unit was designed specifically to suk the requirements of a subsequent free-flight test of the model. Finally, the report includes some data showing the dowmvash at the tail position both for this model and for a corresponding model having a symmetrical, untwisted wing.