Abstract: Numerical simulation of supersonic combustion under the equivalence ratio of 0.97, 0.78, 0.65, 0.46 and 0.32 was performed, and the computational results were validated by the experimental data of the kerosene combustion on a direct connected supersonic combustion experiment facility of the School of Astronautics, Beijing University of Aeronautics and Astronautics. The computational results of the flow fields were obtained by the finite-volume method with implicit lower-upper decomposition （LU） approach and k-g turbulence model, and the chemical source terms were treated by a point implicit method. The results indicate that combustion mainly takes place near the lower wall where a kerosene atomizer is mounted in a cavity. The flow is divided into subsonic and supersonic layers. When the combustion generates high back pressure under high equivalence ratio, the boundary layer separates from the combustor wall, and shock train flow structure can be captured in the combustor segment corresponding to the cavity. The calculated wall pressure distribution agrees well with the available experimental data.

Key words: scramjet, supersonic combustion, kerosene fuel, numerical simulation, shock train, turbulence models