Abstract: The direct-connection test of solid rocket scramjet with cavity flameholder using boron-containing fuel-rich propellant was carried out. The simulated flight conditions was 23km and 5.5Ma. The supersonic combustor performance was calculated by measured pressure， thrust and flow rate. The experimental results show that due to the deposit of the combustion product， the fuel mass flow rate continuously increased during the test， and the equivalence ratio increased from 0.44 to 0.54. The cavity flameholder studied in the current paper indeed prompt the combustion efficiency of gas phase， however， no obvious improvements on the combustion efficiency of condensed phase has been observed. In this experimental condition， the total combustion efficiency was 48%， and the specific impulse was 423s， higher than that of the central strut injection configuration mentioned in reference.

Key words: Solid rocket scramjet;Cavity;Gas generator;Boron-containing fuel-rich solid propellant;Combustor performance

摘要： 以含硼贫氧固体推进剂为燃料，对带凹腔火焰稳定器的固体火箭超燃冲压发动机燃烧室构型首次开展了地面直连试验研究。试验模拟了23km，5.5Ma的飞行工况，通过测量压强、推力和流量等参数，得出了燃烧室性能。试验结果表明：一次富燃燃气在燃气发生器喉部沉积导致燃气流量持续提高，试验过程中当量比由0.44逐渐增加至0.54；本文所研究的凹腔稳焰结构提高了富燃燃气中气相可燃组分的燃烧效率，但对于凝相颗粒燃烧的促进效果不明显；试验工况下发动机总燃烧效率约48%，高空比冲约为423s，高于文献中所报道的中心支板喷射稳焰的固体火箭超燃冲压发动机试验比冲性能。

关键词: 固体火箭超燃冲压发动机;凹腔;燃气发生器;含硼贫氧固体推进剂;燃烧室性能