Abstract: This paper presents a model of numerical calculation for the tempreture distribution on nozzle throat inserter of solid rocket under moving boundery conditions related to the ablating recession,deposition of solid particles from the exhaust gases,and blowing process of deposit. The model is based on the principles of finite-volume energy coservation.The finite-difference method and the alternative direction implicit scheme which is of good stability are used to get numerical solution.To ensure the stability of the numerical solution when boundary recession is so much that near-boundary-cell will drop,the complete implicit interpolative-iterative method is introduced to calculate "the temprature on inner wall and at nodes of surface-cell. The feature of multilayer structure of nozzle throat, aniso-tropy of material, and the variable thermophysical properties with temprature of material are considered in the scheme so that the model is of certain generality and may be applied to the engineering design and research.Several experiments for the solid motor show that the model and the method proposed in this paper could mainly conform to the real working conditions of nozzle inserter and are of quite satisfactory accuracy.

Key words: Temperature field;Nozzle throat;Solid rocket engine;Nozzle abla-tion;Calculation

摘要： 本文探讨了固体火箭发动机喷管喉村在烧蚀退移、燃气中固体微粒的沉积以及沉积物吹除所引起的动边界条件下的温度场数值计算模型。模型的建立以有限体积能量平衡的基本原理为依据,采用有限差分的数值计算方法及稳定性较好的交替方向隐式格式求解。为确保数值计算结果在边界退移及网格脱落时的稳定性,采用了插值与迭代相结合的方法,以全隐式格式求解表层网格节点和受热壁面的温度。模型还考虑了多层背村材料结构、变物性材料特性等诸种因素,使之具有一定的普遍性,并可应用于工程设计和研究中。数次发动机实验证明,本文提出的温度场计算模型能够较真实地反映喉衬温度响应,具有较为满意的计算精度。

关键词: 温度场;喷管喉部;固体火箭发动机;喷管烧蚀计算