大流量、高压航天反力式涡轮最佳通道气动设计

Optimum Pneumatic Design of Reaction Turbine Channel for Large Flow Rate High Pressure Space Engine

基于闭式补燃循环液体火箭发动机流量大的特点,欲设计效率水平高的航天反力式涡轮。因涡轮进出口压力均很高,而膨胀比小、载荷系数大,为保证较高的涡轮效率水平,对涡轮气动设计方法进行了优化。在涡轮进口总温、总压、转速和功率一定条件下,以AMDC/KQ涡轮叶栅损失模型为基础,依据涡轮中径的一维气动计算,对涡轮子午通道、叶栅通道及叶栅造型几组参数组合分别进行了气动设计的优化,研究了涡轮中径、叶高、叶栅稠度、导动叶喉宽匹配及动叶进口构造角对涡轮效率的影响,实现了涡轮效率水平最高。

Attributing to the large flow rate, high efficiency level reaction turbine is adoptive in staged combustion cycle liquid rocket engine. As both turbine inlet and outlet pressures are extremely high, but the pressure ratio is small the turbine load coefficient is great. To ensure the high level of turbine efficiency, optimum pneumatic design of turbine channel was employed. Given the turbine inlet temperature and pressure, rotate speed and power, based on AMDC/KQ formulae for turbine cascade losses, the correlating structure parameters of turbine meridional channel, cascade channel and blade form were optimized by one dimensional average-diameter pneumatic performance calculation. By this means, the influences of turbine cascade average diameter, blade height, cascade blade denseness, match between the stationary and rotor blade throats and rotor inlet structure angle were investigated, and the highest level of turbine efficiency was obtained.