以某型大涵道比涡扇发动机为对象,采用高压涡轮导向器喉道流量函数的方法,计算得到燃烧室的出口温度,并间接获得燃烧室出口燃气流量。计算结果表明,本文方法所得结果与设计方计算结果吻合较好。在此基础上,针对该计算方法进行了参数敏感性分析,其结果将有助于飞行试验工程师更加合理地选择测试方法及传感器类型,为后续的试验工作奠定技术基础。 Taking a large bypass ratio turbofan engine as an object, the outlet temperature of the combustion chamber was calculated by using the throat flow function of the high pressure turbine guide, and the gas flow at the outlet of the combustion chamber was indirectly obtained. The calculation results show that the results obtained by this method are in good agreement with the calculated ones. Based on this, parameter sensitivity analysis is carried out for the calculation method. The results will help flight test engineers to choose test methods and sensor types more rationally, and lay a technological foundation for subsequent test work.