RBCC发动机纯火箭模态流场数值仿真研究  

Numerical simulation for flow field in pure rocket modality of RBCC engine

作　　者：张留欢[1] 南向军[1] 张蒙正[1] 

机构地区：[1]西安航天动力研究所,陕西西安710100

出　　处：《火箭推进》2016年第2期42-46,共5页Journal of Rocket Propulsion

基　　金：国家863项目(2010AA702308)

摘　　要：基于某火箭基组合循环(RBCC)发动机结构及气动参数开展了飞行高度30 km、飞行速度8 Ma时,发动机纯火箭模态三维流场数值仿真。对进气道、燃烧室、尾喷管、火箭发动机等组件流场结果进行分析,并计算了发动机总体推力。结果表明:纯火箭模态下,RBCC发动机进气道存在气流分离,喉部总压恢复系数约为0.34;燃烧室存在两股气流掺混,二级进出口总压损失约38.5%;二级燃烧室流场结构复杂,使得尾喷管入口截面气流参数分布不均,其总压畸变值为0.648;纯火箭模态下该RBCC发动机轴向推力约1 700 N。Based on the structure and gasdynamic parameters of a rocket based combined cycle （RBCC） engine, the 3D flow field numerical simulation for pure rocket modality of engine was performed under the flight conditions with altitude of 30 km and flight velocity of 8 Ma. The flow fields in air intake, combustor, nozzle and rocket engine are analyzed. The total thrust of the engine is calculated. The results show there is airflow separation in the air intake in the pure rocket modality, and the total pressure recovery coefficient at throat is about 0.34. The air from air intake and the gas from rocket have interaction （mixture loss, shear loss, shockwave loss）, which leads to the loss of kinetic energy. The loss of total pressure at entrance of secondary combustor is 38.5%. As the structure of flow field at secondary combustor is complicated, the distribution of air flow parameters at the entrance of nozzle are non-uniform. The value of total pressure distortion is 0.648. The axial thrust of RBCC engine in pure rocket modality is about 1 700 N.

关 键 词：火箭基组合循环发动机 纯火箭模态 数值仿真 

分 类 号：V434-34[航空宇航科学与技术—航空宇航推进理论与工程]