为准确预示典型固体火箭发动机工作过程中燃烧室喷管整体结构温度场及应力场分布规律,建立了包含绝热层和喉衬的燃烧室喷管一体化模型。通过Fluent流体分析软件和ANSYS结构分析软件进行了非稳态流热固耦合数值计算。仿真结果表明,固体火箭发动机工作过程中,由于绝热层的存在,燃烧室壳体温度上升缓慢。热流密度在喉道上游达到最大值,喉部区域对流换热最为严重。热应力在喷管收敛段达到峰值,且随时间增大而增大。

In order to accurately predict the typical burner nozzle in solid rocket engine working process of the whole structure of the temperature field and stress field distribution, established include thermal barrier and throat lining chamber nozzle integration model. Through Fluent fluid analysis software and ANSYS structural analysis software, the unsteady flow thermo-solid coupling numerical calculation was carried out. The simulation results show that solid rocket engine working process, as a result of the existence of thermal barrier, combustion chamber shell temperature rise is slow, upstream maximum heat flux density in the throat, throat area of convective heat transfer is most serious. Thermal stress peaks at the nozzle convergence and increases with time.