Numerical investigation of non-premixed oblique detonation operations in scramjet combustor

Abstract

The present study aims to develop strategies for hydrogen injection to operate scramjet combustor configuration in detonation mode at higher Mach number flights conditions. The reactive multi-species unsteady Navier-Stokes equations along with turbulence modelling are solved with detailed chemical kinetics for a two-dimensional computational domain of cavity based scramjet combustor. In order to establish detonation mode combustion a finite length wedge at angle is attached to the downstream of cavity in a scramjet combustor configuration. Initial simulations are performed at Mach 7 incoming air flow with freestream pressure of 40 kPa and temperature of 300 K for 2 ms time duration. The hydrogen fuel is injected at 30 mm upstream of cavity with angle of injection 15◦ using straight pipe of 2 mm width to provide mass flow rate equivalent to ϕ = 0.34 with respect to incoming air mass flow rate. It is found that the presence of cavity between fuel injector and wedge stabilizes the detonation mode combustion and suppress the intermittent transition between deflagration and detonation modes in comparison to without cavity case. Further the flow conditions at the starting of combustor based on hypersonic intake, operating at an altitude of 25 km with flight Mach number 9 are simulated for cavity based combustor with wegde. The outcome suggest that high temperature of incoming flow can have adverse effect to develop detonation mode combustion, but with cavity and distance between wedge and cavity stable detonation front can be established.