一种小型全流量补燃循环火箭发动机实验装置

介绍一种新的液体火箭发动机动力循环型式—全流量补燃循环的概念及其相对于其它动力循环的优点。为研究这一先进的循环系统,设计了一套小型全流量补燃循环氢/氧火箭发动机实验装置。结合该装置的系统方案,对其进行一维管路计算;通过对2个预燃室进行热力计算,确定了其燃烧温度和预燃气体的热物理性质;在燃烧室压强和混合比大范围变化的情况下,对氢氧推进剂的比冲特性进行探讨,以此确定燃烧室压强为4.0MPa,推进剂余氧系数为0.75。最后估算出该实验装置所能产生的推力为4018.77N。

An Experimental Equipment for Small ScaleRocket Engine with Full Flow Staged Combustion Cycle

A new kind of power cycle of the propellant feed system for liquid propellant rocket engine is presented.The full flow staged combustion cycle and its superior performance compared to the traditional cycle are discussed.For further study of this advanced cycle system,experimental equipment for small scale rocket engine with full flow staged combustion cycle was designed and established.One-dimensional propellant feeding pipe system calculation has been carried out to confirm the pressure and volume needed for the propellant store tank.Through the thermodynamic calculation of two pilot burners,the burning temperature and the thermodynamic properties of the burning gas were obtained for further design of the pilot burner.A large number of specific impulse for the hydrogen-oxygen propellant was investigated with the chamber pressure and mixture ratio varying in a wide range.According to the performance analysis done,the chamber pressure and mixture ratio were determined to be 4.0 MPa and 0.75 respectively with certain consideration of the practical experimental conditions.Finally,the thrust generated was predicted to be 4 018.77 N.