Abstract
Numerical and experimental approaches are conducted to investigate the flow features and secondary combustion performance induced by different air–fuel ratios in a boron-based solid-ducted rocket engine. The results indicated that the afterburning chamber flow features become more complicated owing to the multiple nozzles of the gas injector, and a number of recirculation zones are generated. Because of this, the mixing of the fuel gas and incoming air is enhanced. When the air–fuel ratio decreases, the heat release in the afterburning chamber increases continuously, which causes the pre-combustion shock train to continue to propagate upstream in the subsonic diffuser of the inlet isolator, along with the boundary layer separation zone also moving forward, and the stability margin of the direct-connect inlet decreasing gradually. Furthermore, the direct-connect inlet works at a critical state with an air–fuel ratio of 11.5. As the mass flow rate of the fuel-rich gas rises gradually, the engine thrust gradually increases, and the number of vortexes in the afterburning chamber and the corresponding region affected by the vortexes generally decrease. Meanwhile, the mixing and combustion of the fuel-rich gas and incoming flow were not substantially changed. Additionally, the combustion efficiency and specific impulse are proportional to the air fuel ratio.
Published Version
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