Numerical investigations of the nozzle performance for a rocket-based rotating detonation engine with film cooling

Abstract

This study numerically investigates the nozzle performance of a rotating detonation engine with film cooling. The premixed stoichiometric hydrogen/air mixture serves as the reactant, and the pure air is pumped into the film cooling holes. The diverging section of the nozzle is designed based on the characteristic method and the maximum thrust theory. The spike truncation ratio is optimized to ΔLspike = 40% Lcowl, reducing the weight and maintaining a high axial thrust coefficient of 0.9456. In contrast with the steady condition, the flow loss of the practically transient condition mainly comes from three parts. The first is the pressure loss of the continuously sweeping shock waves, the second is the viscous effect of the flow passage, and the third is the energy dissipation of the recirculation zone in the base region. When the film cooling strategies are introduced on the inner and outer walls, the hole type a/b < 1.0/1.0 is preferred thanks to a more considerable coolant coverage induced by more violent interactions between the radial jet and the mainstream. The cooling effects and the aerodynamic performance can be further balanced within a moderate injection velocity. The final optimized nozzle configuration with ΔLspike = 40% Lcowl, a/b = 1.0/2.0, and Vsec = 100 m/s corresponds to an axial thrust coefficient of 0.9552, a discharge coefficient of 0.9344, a pressure gain of -16.48%, and a root mean square deviation of the flow deflection angle of 14.42°.