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采用拓展中心线、不同的流通截面面积变化规律和倒圆半径变化规律对内并联型TBCC(turbine based combined cycle engine)进气道涡轮通道扩张段进行了设计.通过数值模拟的手段,对涡轮通道扩张段设计参数的影响规律和涡轮模态下涡轮通道扩张段的气动特性进行了研究,并利用高速风洞试验结果对数值模拟方法进行了验证.研究结果表明:中心线控制点纵坐标在1.50~2.25、涡轮通道扩张段出口等直段长度与出口直径比值在0.3~0.7的范围内取值时,涡轮通道扩张段可获得较高的出口总压恢复系数和较小的出口总压畸变指数;采用前急后缓的流通截面面积和倒圆半径变化规律能使涡轮通道扩张段获得较小的出口总压畸变指数;随着飞行马赫数的增加,进气道和涡轮通道扩张段的流量系数先不断减小,在飞行马赫数为0.9附近达到最小,之后又逐渐增加,涡轮通道扩张段出口总压恢复系数不断升高,在飞行马赫数为0.7附近达到最大,之后又逐渐降低;涡轮模态下,涡轮通道扩张段出口总压畸变指数均小于0.5,能很好地满足涡轮发动机对进口流场的要求.
The expanding section of the inlet of the turbine-based combined cycle engine (TBCC) was designed by using the expansion center line, different variation rules of the flow cross-sectional area and the changing rule of the radius of the radii of the turbine.Through numerical simulation, The influence law of the design parameters of the expansion section and the aerodynamic characteristics of the turbo channel expansion section under the turbo mode were studied and the numerical simulation method was validated by using the high speed wind tunnel test results.The results show that the ordinate of the centerline control point is 1.50 ~ 2.25, when the ratio of the length of the straight section to the exit diameter of the turbine passage expansion section is within the range of 0.3-0.7, the total outlet pressure recovery coefficient and the smaller total pressure distortion index of the outlet can be obtained in the expansion section of the turbine passage. The flow rate of inlet section and turbine section can be reduced with the increase of Mach number, The coefficient first decreased continuously, reached a minimum near the flight Mach number of 0.9 and then gradually increased, and the total pressure at the exit of the expansion of the turbine passage was restored The number increased continuously and reached a maximum at flight Mach number of 0.7 and then decreased gradually. Under turbo mode, the total pressure distortion index at exit of turbine passage was less than 0.5, which could well meet the requirements of turbine engine inlet flow field Claim.