Design studies of hypersonic lifting vehicles have generally indicated that aerodynamic heating may be reduced by using highly swept configurations with blunted leading edges. For laminar boundary layers the effect of sweep angle A on the heat transfer at the leading edge is usually taken as cos A as shown by the data of Feller (ref. 1) who measured the average heat transfer on the front half of a swept cylinder. More recent data (refs. 2 and 3) have indicated that the effect of sweep may be more nearly cos3/2 Lambda which, at a sweep angle of 75 deg, would result in a 50-percent reduction of the heat transfer predicted by the cos A variation. The data and theory of reference 4 also indicate a cos3/2 lambda variation but the theories of references 5 and 6 indicate a variation somewhere between cos A and cos3/2 lambda for large stream Mach numbers. The data of reference 7, in contrast to the investigations just cited, showed large increases in average heat transfer to a circular leading edge with increasing A up to a lambda of about 40 deg. These increases in heat transfer were probably caused by transition to turbulent flow which apparently resulted primarily from the inherent instability of the three-dimensional boundary layer flow on a yawed cylinder. The leading-edge Reynolds numbers of reference 7 were considerably larger than the values in references 1 to 4 and were also larger than typical values for full-scale leading edges of hypersonic vehicles; hence, the main application of the high Reynolds number tests will probably be to bodies at angle of attack.

The heat-transfer characteristics of the laminar compressible boundary layer on a hemisphere cylinder have been investigated at free-stream Mach numbers of 5, 6.5, and 8. The Reynolds number based on free-stream conditions and model diameter was varied from 70,000 to 700,000. Various conditions of steady-state heat trasnfer to the model were realized by circulating a coolant through the model, and by varing the tunnel supply air temperature. The wall to stagnation termperature ratio was varied from 0.43 to 0.75. Optical observations and Pitot pressure surveys of the boundary layer showed it to be laminar on both the hemisphere and the cylindrical afterbody. The heat transfer was evaluated from the temperature differences measured across the model wall under steady-state conditions. Over the hemisphere, the local non-dimensional heat-transfer parameters are, on the average, approximately twnty percent larger than predicted for an isothermal body by Korobkin's modified incompressible theory.