Experimental study on measurement and calculation of heat flux in supersonic combustor of scramjet

An experimental measurement and calculation method which consist of thermal response model, convergence criteria and control algorithms, is proposed in this paper for the determination of heat flux in a scramjet combustor. Numerical simulations are done to evaluate the effectiveness of the proposed method, and experiments are made in the direct-connect hydrocarbon fueled scramjet combustor of Mach-6 flight for different equivalence ratios. The distribution of heat flux along the axial and circumferential directions can be obtained using the proposed method. The distribution of heat flux is uneven which is caused by the aerodynamic heating, combustion heat release and changes of section area, and the peak heat flux can be more than 2MW/m² during the experiments. Heat flux increases with the increase in equivalence ratio for the same Mach number. And axial distribution of heat flux is uniform for different equivalence ratios. In addition, the combustion heat release area of the combustion chamber can therefore be concluded which is useful for guiding the structural design of the thermal protection system.