壁温对氢燃料超燃燃烧室流动特性的影响研究

In order to study the relationship between wall heat transfer and supersonic combustion in the scramjet combustor, using FLUENT software with RNG k-ε turbulence model, finite rate/eddy dissipation combustion model, density-based AUSM+implicit algorithm for the dual-mode combustion chamber model of China Aerodynamics Research and Development Center, the three-dimensional cold and hot flow fields are calculated. The simulated incoming Mach number is 2.05, the total temperature is 1870.9K. The flow fields and combustion fields of the scramjet combustor with wall temperatures of 500K, 700K, 900K, and 1300K, respectively, and adiabatic conditions were simulated. The numerical simulation results of the wall pressure of the combustor agree well with the experimental results. When the wall temperature is 500K, 700K, 900K, 1300K and adiabatic, respectively, the average errors are 8.89%, 5.78%, 14.41%, 13.97% and 16.53%. When the wall temperature is 700K, the numerical simulation results are the closest to the experimental results. Through the comparative analysis, it is found that with the decrease of the wall surface temperature, the wall pressure trend is almost unchanged, but the wall pressure decreases, and the starting point of the pressure rise of the wall pressure is greatly shifted back. The shock train generated by the combustion gradually moves downstream of the combustion chamber, and the shock train structure changes, but the front end of the shock train is X-shaped shock wave. The Mach number in the combustion chamber increases, while the area of the high temperature region of the combustion field decreases, then the combustion mode inside the combustor gradually shifts from the sub-combustion mode to the super-combustion mode.