Numerical study of combustion and convective heat transfer of a Mach 2.5 supersonic combustor

• Convective heat transfer of supersonic combustor was numerically studied.
• Peaks of wall heat flux at varied fuel/air ratios are identified.
• Shock structures and vortices are related to flow separation caused by combustion.

In this paper, characteristics of combustion and convective heat transfer of a supersonic combustor at two fuel/air equivalence ratios of 0.9 and 0.46 were numerically studied. The numerical method of Favre averaged Navier–Stokes simulation with SST k-ω turbulence model and a multiple-step reaction mechanism of ethylene is introduced. The inlet Mach number of the combustor is 2.5 and inlet total temperature is 1650K, corresponding to Mach 6 flight conditions. Ethylene is injected at two locations upstream of a flame-holding cavity. The numerical method was validated by comparing the present results of wall pressures and heat fluxes to experiments and theoretical analysis. It is found that, due to injection of fuel at the bottom wall, fuel/air mixing and combustion occurs mainly in the vicinity of the bottom wall. High non-symmetry in distributions of the bottom and the top wall heat fluxes is observed. Peaks of wall heat flux at different locations and at varied fuel/air equivalence ratios are identified, which are caused respectively by effect of cavity and by shock structure formed upstream of the injection points. It is also found that heat flux peaks are strongly related to the reaction step of CO→CO2, contributing to major heat releasing.