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A Flight Investigation of Laminar and Turbulent Boundary Layers Passing Through Shock Waves at Full-scale Reynolds Numbers (1953)
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NACA TN 5056 15 Rc = (19.0 + 1.8) x 106. (a) Mq = 0.710 ± 0.002; cL = 0.10 ± 0.05; (b) Mq = 0.736 ± 0.002; CL = 0.11 + 0.02; Rc = (20.8 ± 2.5) x 106. … (c) Mq = O.76O ± 0.005; CL = 0.11 ± 0.02; Rc = (25.3 ± 2.5) X 106. Figure 3.- Distribution of Mach number through the boundary layer at various chordvise positions for several values of free-stream Mach number, as obtained in flight with laminar flow ahead of the shock wave.NASA Technical Reports Server (NTRS)
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