跨声速涡轮静叶尾缘激波对动叶前缘气膜冷却效果影响的研究

Effects of Vane Trailing Edge Shockwave on Rotor Blade Leading Edge Film Cooling Effectivenessin Transonic Turbine Stage

关键词:  跨声速涡轮  气膜冷却  尾缘激波  非定常数值模拟  冷却效果

Key words:  Transonic turbine  Film cooling  Trailing edge shockwave  Unsteady simulation  Cooling effectiveness

摘要:

为进一步探究跨声速涡轮中非定常激波对气膜冷却效果的影响，以跨声速涡轮级叶型作为研究对象，采用非定常数值模拟方法，通过在动叶前缘不同弦向位置进行冷气喷射，探讨了静叶尾缘外伸激波的扫掠对动叶前缘气膜冷却效率的影响。研究结果表明，静叶尾缘外伸波接触到动叶前缘时会导致接触点下游气膜冷却效率降低；冷气喷射孔距离前缘越近，每周期内受上游静叶尾缘外伸波影响时间越长，受影响时长在10%～20%周期之间；不同方案中气膜孔下游相同距离位置上，时均冷却效率相差最大可达89.5%。

Abstract:

In order to further clarify the effects of unsteady shockwave on film cooling performance in a transonic turbine， unsteady numerical simulations were conducted on the profiles of a transonic turbine stage. Film holes were located at different positions near leading edge of rotor blade to study the effects of sweeping vane trailing edge outer-extending shock wave on film cooling effectiveness near rotor blade leading edge. The results showed that the sweeping shockwave can cause a decrease in film cooling efficiency downstream of the shockwave on blade surface when it reaches rotor blade. The closer the film hole to the leading edge， the longer the time that coolant jet flow suffers from shockwave， which is within a range of 10%～20% of a period. At positions which have the same distance downstream of film holes， the maximum deviation of film cooling effectiveness can be 89.5%.