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Skin Friction, Heat Transfer, and Pressure Measurements on Hypersonic Inlet Compression Surfaces in the Mach Number Range 7.5 to 16

Report Number: AFFDL TR 65-199
Author(s): M. O. Ryder, Jr.
Corporate Author: Cornell Aeronautical Laboratory, Inc.
Laboratory: Air Force Flight Dynamics Laboratory
Date of Publication: 1965-12
Pages: 194
Contract: AF 33(615)-1845
Project: 1366
Task: 136605
AD Number: AD0627798

An experimental study of boundary layer flow, under the influence of adverse pressure gradients typical of hypersonic inlets, was conducted on two two-dimensional and three axisymmetric compression surface models instrumented with skin friction, heat transfer and pressure gages. Tests were conducted over a Mach and Reynolds number range of 7.5 to 16 and 32,000 per ft. to 4,700,000 per ft., respectively. The boundary layers on the two-dimensional models were laminar and attached for all conditions tested. Local laminar separation occurred for some conditions of the axisymmetric tests. Boundary layer transition occurred for the high Reynolds number runs at a Mach number of 8 on the axisymmetric models but the adverse pressure gradients were not large enough to cause the turbulent boundary layers to separate. The skin friction gages gave a more accurate indication of localized boundary layer separation than either the heat transfer or static pressure distributions on the models tested.

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