高压氢氧火箭发动机推力室燃烧稳定性分析
刘倩1,李敬轩2,孙纪国1,梁炫烨2,向小林2,潘亮1,郑孟伟1

(1.北京航天动力研究所,北京100076;2.北京航空航天大学宇航学院,北京100191)

高压氢氧火箭发动机; 推力室; 燃烧稳定性; 数值仿真; 火焰传递函数

Analysis on combustion stability of thrust chamber in high pressure hydrogen-oxygen rocket engine
LIU Qian1,LI Jingxuan2,SUN Jiguo1,LIANG Xuanye2,XIANG Xiaolin2,PAN Liang1,ZHENG Mengwei1

(1.Beijing Aerospace Propulsion Institute, Beijing 100076, China; 2.School of Astronautics, Beihang University, Beijing 100191, China)

high pressure hydrogen-oxygen rocket engine; thrust chamber; combustion stability; numerical simulation; flame transfer function

备注

针对高压氢氧火箭发动机推力室不设置隔板喷嘴和声腔的结构方案,利用火焰传递函数+低阶声学模型的解耦预测仿真方法,分析了不同喷注参数和结构参数下燃烧室的燃烧稳定性裕度。采用非定常雷诺平均NS方程(URANS)计算同轴直流喷嘴非稳态燃烧过程以获取火焰传递函数,其中采用Soave Redlich Kwong(SRK)状态方程计算密度等物性参数; 考虑到同轴直流喷嘴的火焰长度与声波量级相当,采用分布式火焰结构进行火焰传递函数建模。采用商业软件COMSOL计算加载了火焰传递函数的燃烧室声学模态,使用模态增长率为评定标准,预测燃烧不稳定性。结果 表明,给定不同燃气/氧喷注速度比、混合比、相对喷嘴压降、缩进深度比、富氢燃气喷前温度等各工况下,预测得到的燃烧室均未出现燃烧不稳定现象。在推力室设计中通过增加燃气/氧喷注速度比或降低燃烧室混合比,有利于提升燃烧稳定性裕度。所做工作为高压氢氧火箭发动机喷注器设计及燃烧稳定性裕度评估提供参考。

Aiming at the structural scheme of the thrust chamber without baffle nozzle and acoustic cavity in high-pressure hydrogen-oxygen rocket engine, the combustion stability of the combustion chamber under different injection parameters and structural parameters was analyzed by using the method of decoupling of unsteady combustion and acoustic system.The unsteady Reynolds-averaged NS equation(URANS)was used to calculate the unsteady combustion process of the shear coaxial injector and obtain the flame transfer function, and the Soave Redlich Kwong(SRK)state equation was used to obtain density and so on. Considering that the flame length of the shear coaxial injector was equivalent to the magnitude of the sound wave, the distributed flame structure was used for modeling the flame transfer function.The combustion chamber acoustic modes loaded with the nozzle flame transfer function were calculated by COMSOL, and the combustion instability was predicted by combining the calculated mode frequencies and their growth rates.The results showed that the predicted combustion stability of the chamber was stable under different gas/oxygen injection velocity ratio, mixing ratio, relative nozzle pressure drop, retraction depth ratio and per-injection temperature under given operating conditions.In the thrust chamber design, it is beneficial to improve the combustion stability by increasing the gas/oxygen injection velocity ratio or reducing the mixing ratio in the combustion chamber.The work provides a reference for the design of injector in high-pressure hydrogen-oxygen rocket engine and the combustion stability evaluation.