某固体火箭发动机工作末期不稳定燃烧  

Combustion instability at end of burning in a solid rocket motor

作　　者：苏万兴[1] 李世鹏[1] 张峤[2] 赵艳栋 叶青青[1] 王宁飞[1] 

机构地区：[1]北京理工大学宇航学院,北京100081 [2]中国空间技术研究院载人航天总体部,北京100094

出　　处：《航空动力学报》2013年第10期2376-2383,共8页Journal of Aerospace Power

基　　金：国家自然科学基金(51076015)

摘　　要：针对某固体火箭发动机工作末期出现的压力振荡现象开展了数值研究与线性预估.通过有限元方法得到了燃烧室空腔的声模态及固有声振频率,轴向1阶与2阶声振频率随燃面退移先减小后增大;利用大涡模拟方法分析了燃烧室内的流场特性及压力振荡特性,振荡频率与试验结果一致,判定该发动机出现了以轴向1阶声振频率为主导的不稳定燃烧;其次分析了发动机内阻尼特性,其阻尼随燃面退移不断减小;最后通过不稳定燃烧线性理论解释了该发动机工作末期出现压力振荡的机理,表明燃面退移过程中喉通比下降是导致发动机由线性稳定转向线性不稳定状态的关键因素.Based on a solid rocket motor （SRM）, numerical simulation with linear pre diction was carried out to study the pressure oscillation at the end of burning. Acoustic modes and natural acoustic frequencies of combustor chamber were obtained by finite element analysis （FEA） method. The results indicate that the first and second axial acoustic fre quencies first decrease and then increase with the regression of the burning surface. The flowfield and pressure oscillation characteristics of the combustor were analyzed via large ed dy simulation （LES） method. The oscillation frequency was well consistent with the experi mental value, confirming that the SRM presented fundamental acoustic combustion instabili ty. Then the damping effect of the motor was analyzed. It shows that the total damping continuously decrease with the regression of the burning surface. Finally, the pressure os cillation mechanism at the end of burning was explained via linear combustion instability the ory. The decrease of the throattoport area ratio is a key factor that makes the SRM turn from linear stable state to linear unstable state.

关 键 词：固体火箭发动机 不稳定燃烧 压力振荡 声模态分析 阻尼特性 

分 类 号：V435.12[航空宇航科学与技术—航空宇航推进理论与工程]