Numerical investigation of non-premixed oblique detonation operations in scramjet combustor
Date
2024-09-23Author
Vashishtha, A.
Dias, S.M.
Palateerdham, S.K.
Nolan, C.
Ingenito, A.
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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.
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