摘要
采用数值模拟方法求解二维轴对称Navier-Stokes方程,对多个平移式和铰链花瓣式延伸出口锥喷管在级间热分离条件下的尾部流场进行了数值分析。结果显示,收拢状态下,受限于延伸锥壁面、发动机前封头及基础喷管外壁,尾部流场温度环境较恶劣;展开过程中,前封头边缘及延伸锥端部存在燃气回流,气动力变化剧烈,平移式喷管展开过程主要受气动力的阻碍影响,花瓣式喷管展开过程则受到燃气的正向推动作用。计算结果加深了2种延伸锥喷管的特性机理的认识,为发动机的热防护设计以及展开机构的驱动力设计提供了依据。
Two-dimensional axisymmetric equations are solved by numerical simulation,and the tail flow field of hinged petal extended conical nozzle and multiple translational extended conical nozzles under the condition of thermal separation between rocket stages was numerically analyzed.The results show that limited by the extended cone wall,the front head of the rocket engine and the external wall of the base nozzle in the closed state,the temperature environment of the tail flow field is harsh.During the process of expansion,gas backflow exists at the edge of the front head and the end of the extension nozzle,where the aerodynamic force changes dramatically.The expansion process of the translational nozzle is mainly affected by the obstruction of aerodynamic force,while the expansion process of the petal nozzle is positively promoted by the gas.The calculation results provide the basis for the thermal protection design of the rocket engine,the characteristics analysis of two kinds of extended nozzles and the driving force design of the developing mechanism.
作者
任孝宇
周子翔
许玉荣
徐节荣
祝珊
吕轩
REN Xiaoyu;ZHOU Zixiang;XU Yurong;XU Jierong;ZHU Shan;LU Xuan(Hubei Key Laboratory of Advanced Aerospace Propulsion Technology(System Design Institute of Hubei Aerospace Technology Academy),Wuhan 430040,China)
出处
《武汉大学学报（工学版）》
CAS
CSCD
北大核心
2021年第2期158-165,共8页
Engineering Journal of Wuhan University
关键词
固体火箭发动机
延伸喷管
尾部流场
数值模拟
solid rocket engine
extendible nozzle
tail flow filed
numerical simulation
