CST造型方法在涡轮叶片前缘修型中的应用研究

Application of CST Method on Modifying Leading Edgeof Axial-Flow Turbine Blade

关键词:  涡轮  前缘  压力尖峰  分离泡  边界层  叶型损失  能量耗散  熵产率

Key words:  Turbine  Leading edge  Pressure spike  Separation bubble  Boundary layer  Profile loss  Energy dissipation  Entropy production rate

摘要:

为了探究CST(形状函数变换技术)造型方法在涡轮叶片前缘修型中的应用效果，完善了CST方法在前缘型线重构中的实施细节，数值模拟了雷诺数对前缘修型前后叶型损失及边界层特性的影响，验证了CST前缘修型方法在新型高速飞行器低压涡轮中的实用性。结果显示：CST方法前缘修型可以消除HD叶型吸力侧前缘的压力峰和分离泡，从而使得高雷诺数条件下吸力侧分离诱导的边界层转捩现象延迟发生，叶型损失降低32%，拓展了低损失状态的雷诺数范围。吸力侧损失的降低在低雷诺数条件下主要来自于前缘附近的剪切层，而高雷诺数条件下主要来自于前缘剪切层和扩压段前的层流边界层。新型高速空天飞行器低压涡轮叶片采用CST前缘修型对提升效率是有效的，在设计点状态附近效率提高0.1%，而膨胀比较低的大负攻角状态下效率提升0.3%~0.5%，损失降低的位置主要集中在叶展中部压力侧边界层和根部的二次流区域。

Abstract:

In order to explore the effects of leading edge modified with the method of CST (Class Function/Shape Function Transformation Technique) on the aerodynamic performance of the blade profile， firstly, the implementation details of CST method in leading edge reconstruction are improved. Secondly, the effects of Reynolds number on profile losses and boundary layer characteristics are studied with numerical simulation. Finally, the practicability of the CST leading edge modification method is verified in the low-pressure turbine of a new high-speed aircraft. The results show that the CST method can eliminate the pressure spike and separation bubble near the suction leading edge of HD profile, which also delays the transition phenomenon induced by the suction leading edge separation bubble under high Reynolds number condition and expands the range of low-loss state Reynolds number. In addition, the profile loss is reduced by 32%. The reduction of suction profile loss under low Reynolds number condition is mainly from the shear layer near the leading edge, while the reduction of suction profile loss under high Reynolds number is mainly from the shear layer near the leading edge and the laminar boundary layer before diffuser section. The application of CST leading edge modification in high-speed aerospace vehicle low-pressure turbine rotor is verified to be effective on improving the aerodynamics efficiency, that making a contribution to increasing the efficiency by 0.1% around the design point and about 0.3%~0.5% at lower expansion state with large negative incidence. The position where the loss is reduced is mainly concentrated in the boundary layer of pressure side and the secondary flow region near the root.