An investigation has been made in flight at free-stream Mach numbers up to O.76 to determine the behavior of laminar and turbulent boundarynlayers passing through shock waves at Reynolds numbers up to 26 x 10 . The free-stream Mach numbers obtained in the tests were sufficiently high to give extensive regions of local supersonic flow. Boundary-layer and pressure-distribution measurements were made on a short-span airfoil built around a wing of a fighter airplane. Laminar flow up to the position of the shock wave was obtained on the upper surface of the smooth airfoil. Turbulent flow ahead (upstream) of the shock wave was obtained by attaching a transition strip near the nose.nThe results of the investigation indicated that the abrupt increase in momentum thickness and displacement thickness in the region of the pressure rise associated with a shock wave was significantly less with laminar than with turbulent flow ahead of the shock wave. With laminar flow ahead of the shock wave, transition from laminar to turbulent flow was caused by the abrupt pressure rise of the shock wave and was accompanied by a large decrease in the value of the boundary-layer shape parameter H through the shock wave. The beneficial effect of having laminar flow ahead of the shock wave was indicated by 50-percent smaller values of momentum thickness and lower values of shape parameter (1.6 as compared with 2.9) downstream of the shock waves.nComparison of the present results with results obtained at low Reynolds numbers by other investigators indicated that with turbulent flow ahead of the shock wave the changes in boundary-layer characteristics through the shock wave were not much different for the full-scale than for the l/lO-scale data. With laminar flow ahead of the shock wave, the changes in momentum thickness and displacement thickness (expressed as a fraction of a characteristic length) were about 10 times as great for the 1/23- scale as for the full-scale data.
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