An investigation of the behavior of the three-dimensional hypersonic flow along intersecting planes has been conducted, considering both the viscous and inviscid flows and their mutual interaction. The experimental program involves measurements of surface pressures and heat-transfer rates in the vicinity of a 90° corner. The local forces and shear stresses can be obtained thereby, yielding some information on the nature of the three-dimensional boundary-layer behavior in this region. Pilot pressure contours are also obtained and utilized in the determination of the complex intersecting shock pattern in the corner. The test data was obtained at a freestream Mach number of 11.2 and a Reynolds number of 1.5 X 10 4 /in. A theoretical analysis of the corner flow was obtained by the development of a set of equations valid throughout the boundary layer, shock wave structure, and inviscid core as previously proposed by Rubin for two-dimensional and axisymmetric flows. This analysis has been shown to be valid in the continuum merged layer and in the viscous interaction regions downstream. The theory is extended here to the three-dimensional corner configuration. Where applicable, the theoretical solutions were compared to the experiments and very good agreement was found to exist over the entire spectrum of flow variables.
Nomenclatureheat-transfer coefficient H = stagnation enthalpy ft = coefficient of thermal conductivity L = yM m z /Re m = reference length M = Mach number p -pressure q -surface heat transfer rate Re x = pux/p, -Reynolds number based on x Re m = POO^OO/MOO = freestream unit Reynolds number St -Qw/p
The present paper deals with an experimental study of the viscousinviscid interaction occuring in a corner region under hypersonic, low density, free stream conditions. The tests were conducted in the Mach 11. 8 hypersonic tunnel at PIBAL over a range of free stream Reynolds numbers between 0. 15 x 10 and 0. 50 x 10 /ft. The model consists of two sharp edged plates mounted at an angle of 90 with respect to each other and with normal leading edges. Data obtained include surface measurements of pressure and heat transfer for values of the interaction parameter (x) between 1.0 and 10. The entire corner region is surveyed at a value of X = 2. 5 and measurements of total temperature, pitot, and static pressure aie obtained. The results indicate that the static pressure in the region of intersection of the shock layers is as much as twice that of the local two dimensional value. The local heat rates in this region are also considerably larger than their two dimensional counterparts.
Surface heat transfer and boundary layer transition easurements have been performed on a slender blunted cone at = 8.0. These measurements have been obtained on the conical surface with mass transfer occurring on the spherical nose cap. The free stream Reynolas number, based on nose radius, varied between .6,cl and i s-. Data was obtained both at zero angle of at-*tack and at angles as high as WX, which exceeds the semi-vertex. cone angle of "We. The technique utilized in the determination. of transition location was the observation of the RKS output of a thih film gauge, imbedded in the model surface, and maintained at a constant temperature by an anemoeter system. A sudden change in the RMS output of the film gauge r' s found to accurately locate the onset of boundary layer transition.
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