一种低空满流的大面积比液体火箭发动机喷管

A Big Area Ratio Liquid Rocket Engine Nozzle with Full-Flow in Low Altitude

关键词:  液体火箭发动机  特征线法  喷管  型面设计  流动分离

Key words:  Liquid rocket engine  Method of characteristics  Nozzle  Contour design  Flow separation

摘要:

针对传统大面积比液体火箭发动机喷管在低空过膨胀状态下易产生流动分离的问题，采用特征线法，基于最大推力喷管，对其扩张段后半部分型面进行了控制压力设计，以保证新生成的大面积比喷管（低空满流喷管）壁面压力不小于分离临界压力。而后通过仿真手段对设计方法进行了校验，并对低空满流喷管的性能进行了评估。结果表明：基于最大推力喷管型面的控制压力设计方法能够实现预定的设计目标，生成的型面不仅保证了喷管在海平面条件下处于满流状态，还使得喷管对燃烧室压力脉动具备了一定的抵抗能力。当燃烧室压力为8.5MPa、燃气比热比为1.144时，相较于将要产生分离的面积比为40的最大推力喷管，低空满流喷管能够将面积比增加至60，从而提高真空比冲约5.24s。而相比于面积比为60的最大推力喷管，等面积比的低空满流喷管真空比冲损失约为1.57s。

Abstract:

For the problem that conventional big area ratio liquid rocket engine nozzles easily generate flow separation in low altitude over-expanded condition， based on thrust optimized nozzles， method of characteristics were used to design divergent latter section contours by pressure controlled， which ensures generated big area ratio nozzle （named low altitude full-flow nozzle） wall pressure no less than the critical separation pressure. Then the design method was verified by simulation， and performances of low altitude full-flow nozzles were assessed. Results show that， expected design goals can be achieved by method of contour pressure controlled based on thrust optimized nozzle. Not only full-flow states in sea-level condition are ensured by generated nozzle contours， but also some ability nozzles had to resist combustion chamber pressure fluctuation. When combustion chamber pressure is 8.5MPa and hot gas specific heat ratio is 1.144， comparing to thrust optimized nozzle with area ratio 40 which is going to separate， area ratio can increase to 60 by low altitude full-flow nozzle， thereby vacuum specific impulse can increase by 5.24s. However， comparing to thrust optimized nozzle which area ratio is 60， vacuum specific impulse is dropped by 1.57s on low altitude full-flow nozzle with same area ratio.