通过建立固体火箭发动机点火瞬态数学模型,对某大型分段式固体火箭发动机工作初期小火箭式点火装置的火焰喷射方式、分段对接部位火焰传播过程以及前后翼燃面的传播过程等进行数值计算研究。计算结果表明,发动机点火过程中,燃烧室内的流动顺畅,没有出现压强异常振荡现象,点火初期的火焰冲击对分段对接部位的绝热结构影响很小,但整个后翼槽药面全部点燃用时在整个火焰传播期用时占比过大。数值计算结果与全尺寸发动机地面热试车结果对比表明,数值计算点火平衡压强、压强爬升时间以及升压速率与地面热试车结果吻合性好。 By establishing the transient mathematical model of solid rocket motor ignition, the numerical simulation of the flame spraying method, the flame propagation process and the spreading process of the front and rear flames of a small rocket ignition device at the initial stage of a large-scale solid rocket motor are carried out. Computational research. The calculation results show that the flow in the combustion chamber is smooth and there is no abnormal pressure oscillation during the ignition process of the engine. The flame impact at the initial stage of ignition has little effect on the thermal insulation structure of the butt joints. However, The entire flame propagation time is too large. Comparing the result of numerical calculation with that of the full-scale engine, it shows that the ignition equilibrium pressure, the pressure climbing time and the pressure increasing rate are in good agreement with the ground test results.