Heat-transfer distributions were obtained on 700 swept slab delta wings with sharp and blunt noses and cylindrical leading edges. The free-stream Reynolds number based on model thickness was 9 x 10(4). Lees' theory for the heat-transfer distribution is shown to be in good agreement with the data obtained on the nose and leading edge. The point of peak heating on the leading edge is in close agreement with the Newtonian stagnation point at all angles of attack. At angles of attack from 40° to 60° cross-flow theory was in excellent agreement with experimental center-line data; however, at other model attitudes the theory differed significantly from experiment. Nose blunting resulted in reduced heating rates over the aft center-line section of the wing at angles of attack less than about 40°.

Transpiration and convective cooling concepts are examined for the fuselage and tail surface of a Mach 6 hypersonic transport aircraft. Hydrogen, helium, and water are considered as coolants. Heat shields and radiation barriers are examined to reduce heat flow to the cooled structures. The weight and insulation requirements for the cryogenic fuel tanks are examined so that realistic totals can be estimated for the complete fuselage and tail. Structural temperatures are varied to allow comparison of aluminum alloy, titanium alloy, and superalloy construction materials. The results of the study are combined with results obtained on the wing structure, obtained in a previous study, to estimate weights for the complete airframe. The concepts are compared among themselves, and with the uncooled concept on the basis of structural weight, cooling system weight, and coolant weight.