针对大长径比固体火箭发动机地面试验中出现的分段不稳定燃烧现象,通过声腔频率分析获取其燃烧室压强振荡频率分布范围,从声-涡耦合及推进剂燃烧响应角度分析了两段燃烧不稳定现象的主要因素,设置对比试验,对比同一装药构型、不同推进剂配方发动机出现的燃烧不稳定现象,结合起振初期、中期与结束时刻的燃烧室声腔模态、燃烧室流场涡结构计算结果以及T型燃烧器压强耦合响应函数测试结果判定:发动机工作初期产生的不稳定燃烧现象是由于燃烧室装药构型导致的旋涡脱落频率与燃烧室声场频率相近产生耦合增益,诱发燃烧室压强振荡;发动机工作至末期,其不稳定燃烧频率不同于初期的频率范围,存在多倍频等特征,该段不稳定燃烧是由于推进剂燃烧响应引起的。从地面试验中发生的两段不稳定燃烧结果开展分析,为固体火箭发动机的优化设计与工程设计中有效规避此类问题提供参考。

For the segmented combustion instability phenomenon of slender solid rocket motors in the ground test, it can obtained the distribution range of pressure oscillation frequency in the combustion chamber by mean of acoustic cavity frequency analysis, analyzing the main factors of two-stage combustion instability from a perspective of acoustic-vortex coupling and propellant combustion response. Then through comparison experiment, comparing the combustion instability of the motor with the same grain configuration and different propellant formulations and combiningewith the acoustic cavity mode of the combustion chamber at the initial, middle and end time and the flow field vortex structure computation of the combustion chamber and the pressure coupled response function test results of the T-burner, the conclusion is that the combustion instability phenomenon at the initial stage of the motor operation is due to the fact that the vortex shedding frequency caused by the grain configuration is close to the frequency of the acoustic field of the chamber to generate coupling gain and induce the pressure oscillation. At the end of the operation of the motor, the combustion instability frequency is different from the initial frequency range, and there are frequency multiplication characteristics, the instability is caused by the combustion response of the propellant. The analysis unfolds from the two-stage combustion instability in the ground test andprovids a reference to optimize solid rocket design and effective avoidance of such problems in engineering design.