液体火箭发动机不同室压下冷却方案适用范围

为了精准评估不同冷却方案对高压液氧烃火箭发动机推力室传热特性的影响，建立了一套再生通道-液膜屏蔽-隔热镀层-辐射换热的整机模型，采用Ievlev半经验模型计算燃气侧壁面的对流换热过程，引入Shruvik安全裕度评估准则，计算推力室径向的分区温度和热流密度。基于某型大推力液氧煤油火箭发动机，研究了不同冷却结构组合的换热能力上限，分析了不同推力室压力对冷却设计方案的影响。结果表明:推力室压力在12 MPa及以下时，可主要依靠再生冷却技术满足冷却需求；在16 MPa及以下时需要配合内冷却环带满足冷却需求；在18 MPa及以下时需进一步设置隔热镀层提高热防护能力；室压在20 MPa甚至更高时，必须采用其他强化换热措施。 

Adaptability of cooling structure schemes of liquid propellant rocket engine thrust chamber under different pressures

In order to accurately evaluate the effect of different cooling schemes on the heat transfer characteristics of thrust chamber of high-pressure liquid oxygen and hydrocarbon rocket engine, a complete model combining regenerative channel, liquid film shield, thermal barrier coating and radiation heat transfer was established. The Ievlev semi-empirical model was adopted for calculation of convection on gas wall. The Shruvik safety margin evaluation criterion was used in analysis. The radial direction multi-layer temperature and heat flux of the thrust chamber were calculated. Based on a large thrust liquid oxygen and kerosene rocket engine, the upper limit of heat transfer capacity of different cooling schemes was studied, and the influence of different thrust chamber pressures on cooling design was analyzed. Results showed that, thrust chamber can mainly rely on regenerative cooling technology to meet the cooling requirements at 12 MPa or below; film cooling slots should be further used at 16 MPa or below; thermal barrier coating should be further installed at 18 MPa or below; several intensive cooling measures must be adopted at 20 MPa or higher.