Abstract
Abstract : A thin-layer Navier-Stokes code, developed for projectile aerodynamics, has been used to compute the supersonic flow over a missile afterbody containing a centered exhaust jet. The thin-layer, compressible, Navier-Stokes equations are solved using a time dependent, implicit numerical algorithm. A unique flow field segmentation procedure is used which preserves the sharp base corner and facilitates the adaption of the grid to the free shear layer in the base region. Solutions have been obtained for an axisymmetric, boattailed afterbody where the free stream Mach number is 2.0 and the jet exit Mach number is 2.5. Computations were made at various jet static pressure to free stream static pressure ratios from 1 through 9. Qualitative features of the base region flow field seen experimentally are very well observed in the computed results. Quantitative comparisons of base pressure with experiment indicate good agreement at high pressure ratios and some disagreement at low pressure ratios.
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