Numerical Investigation of Asymmetric Mach 2.5 Turbulent Shock Wave Boundary Layer Interaction

Supersonic shock wave boundary layer interactions are common to inlet flows of supersonic and hypersonic vehicles. This paper reports on wall-resolved implicit large-eddy simulations a canonical Mach 2.5 turbulent interaction experiment at the NASA Glenn Research Center. The upstream was nominally axisymmetric two-dimensional. A conical centerbody with 16 deg half-angle maximum radius 0.147D test section diameter employed generate wave, where D is diameter. Asymmetric (swept) were obtained by displacing generator away from centerline. present simulation for displacement D/6. Results asymmetric compared results an earlier corresponding interaction. experimental Reynolds number based ReD=4×106. For simulations, lowered ReD=4×105 keep computational expense within limits. Compared interaction, streamwise extent separation varies considerably in azimuthal direction strongest location that closest generator. separated flow regions noticeably reduced substantial crossflow observed between locations farthest Fourier analysis unsteady data indicates low-frequency content region Away this region, increasing sweep angle cross-flow, diminished. proper orthogonal decomposition captures spanwise coherent structures more two-dimensional parts