采用大涡模拟技术,针对某分段式固体火箭发动机开展了发动机燃面退移0mm,160mm和280mm三个时刻下发动机燃烧室内部流动不稳定现象的数值分析,获得了三个典型时刻燃烧室内压强可能的振荡特性。计算结果表明,在发动机点火初期燃烧室内流动不稳定性主要由表面涡脱落导致；随着燃面的退移,端面限燃层暴露在燃气中,由于端面包覆结构残余的影响,燃烧室内流动不稳定性主要由障碍涡脱落决定,且与点火初期相比,压强振荡的频率逐步减小。 

In order to investigate the flow instability in a segmented SRM(solid rocket motor), numerical analyses based on LES(large eddy simulation) were carried out for a potential pressure oscillation characteristics in the chamber when the burning surface regression was at 0 mm, 160 mm and 280 mm,respectively. The numerical results indicate that initial stage flow instability is mostly caused by the parietal vortex-shedding in the chamber shortly after the motor firing. As the burning surface regresses continuously, the restrictors expose to the hot gases. The chamber flow instability is mostly induced by the obstacle vortex-shedding generated due to the effects of the residual annular thermal protection. Correspondingly, the frequency of pressure oscillation is decreased gradually as compared with that of initial stage. 