轴向喷注端面燃烧固液火箭发动机数值仿真  

Numerical simulation of axial-injection end-burning hybrid rocket motor

作　　者：李新田[1] 卢鑫[1] 李延成[1] 刘珺怡 张程[2] 

机构地区：[1]中国航天科技集团公司中国运载火箭技术研究院,北京100076 [2]国防科学技术大学航天科学与工程学院,长沙410073

出　　处：《航空动力学报》2016年第3期740-745,共6页Journal of Aerospace Power

摘　　要：针对一种轴向喷注端面燃烧固液火箭发动机开展了研究,分析了该类发动机的燃烧机理,建立了考虑气-固边界耦合及燃料热解的发动机燃烧流动仿真模型,对发动机内流场进行了数值仿真.分析认为:发动机内存在端面燃烧和侧面燃烧两种燃烧状态,燃烧状态主要受氧化剂流速的影响.当氧化剂流速跨越转变速度后,燃烧状态发生改变.发动机流场数值仿真结果同样表明:不同氧化剂流速下发动机内存在上述两种燃烧状态,仿真得到的转变速度区间与文献试验结果吻合较好.Numerical simulations of a axial-injection end-burning hybrid rocket motor were conducted.The combustion mechanism of this kind of hybrid rocket motor was studied.The simulation models of combustion and flow were established in consideration of both gas-solid interfacial coupling and solid fuel pyrolysis.Numerical simulations of motor combustion were carried out.Analysis shows that,two combustion models,end-burning and side-burning,exist in the motor,and the combustion models mainly depend on the oxidizer velocity.The combustion models change when the oxidizer velocity exceeds the transition velocity.Numerical simulation results also validate the existence of the two combustion models under different oxidizer velocities,and the range of transition velocity matches well with the experimental results in reference.

关 键 词：轴向喷注 端面燃烧 固液火箭发动机 转变速度 燃料燃速 

分 类 号：V436[航空宇航科学与技术—航空宇航推进理论与工程]