Application of Detonation Waves to Rocket Engine Chamber

Abstract

We present the results of experiments performed with a rotating detonation engine using continuous detonation in an annular combustor to create thrust. Detonation waves propagate in a supersonic and very small region, allowing shortening of the combustor. The combustor of RDE causes high-pressure loss when the propellant is injected, and cooling is necessary due to high heat flux. However, the combustion efficiency of detonation combustion in an annular combustor is the most important, but have not been fully elucidated. In addition, the influence of the injector shape and direct cooling of a rotating detonation combustor require clarification. This paper reports the measurement results of combustor stagnation pressure and thrust, the influence of injector shape on c * efficiency, and the estimate of heat flux. The c * efficiency was 88-100% when we used the convergent or convergent-divergent nozzle and the equivalence ratio was less than 1.0. The shape of the injec-tor influenced wave propagation mode, but the mode did not change the c * efficiency. We estimated time-spatial average heat flux from the terminal temperature, and the heat flux was 8.1 ± 1.8 MW/m 2 in no water injection condition. The rocket RDE sled test was successfully performed. The total mass of the rocket RDE system was 58.3 kg, total time averaged thrust was 201 N, the time averaged mass flow rate was 143 g/s, and the specific impulse was 144 s.