The possibility of using shock-induced combustion or an oblique detonation wave in hypersonic air-breathing engines has been investigated during the last few years. This presentation examines the interaction of a shock wave and a laminar boundary-layer in a combustible mixture. The Navier-Stokes' equations with a 7 species combustion model are cast into the form of a conservation law so that imbedded discontinuities are captured without oscillations. The algorithm developed employs the point-implicit TVD MacCormack method. In the preheated flow shock detonation cases, the scheme agrees well with published results and with theoretical analyses. An exothermic calculation was performed using a new diagonal method. The results obtained were identical to results obtained using the point implicit method, with a 20 percent reduction in CPU time. Solutions of flow with shock-deflagrations in oblique shock/boundary-layer cases indicate no evidence that the presence of the boundary-layer causes a catastrophic change, such as separation, in the overall flow field although there is a significant effect. Small regions of flow separation were observed in the ram accelerator oblique detonation cases where the detonation waves interact with the boundary-layer at the projectile surface. Under all the given conditions, there is no evidence that unsteady flow results from the interaction of the shock-deflagration or detonation wave system and the boundary-layer.
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