高负荷跨声速涡轮转子叶顶激波系结构及其对热流分布影响的数值研究

Numerical Study on Shock Wave Structure and Its Effect on Heat Flux Distribution of Highly Loaded Transonic Turbine Rotor Tip

关键词:  航空发动机  跨声速涡轮  涡轮叶顶  激波结构  传热特性  变工况

Key words:  Aeroengine  Transonic turbine  Turbine blade tip  Shock wave structure  Heat transfer characteristics  Variable working condition

摘要:

叶顶泄漏流动会增强叶顶换热，为防止高负荷跨声速转子叶顶烧蚀，本文采用数值方法研究了某高负荷跨声速转子叶顶激波系结构及其对叶顶热流分布的影响，为跨声速转子叶顶的冷却设计提供参考。结果表明：在叶顶间隙压力侧出现分离泡，高马赫数气流流经分离泡后的折转产生了较强斜激波，该斜激波干涉导致机匣边界层出现分离，分离泡上下游的两个折转产生了两条反射激波。受两条反射激波影响，叶顶边界层增厚导致叶顶出现两道紧密的低热流条带；沿着流动方向，分离泡下游激波强度逐渐增大导致机匣边界层分离程度加剧，边界层分离产生的前后折转程度随之增大，导致两条反射激波强度增大，引起叶顶条带热流进一步降低。当涡轮级膨胀比大于2.0时，受超音堵塞影响，亚声速区域的马赫数基本不受膨胀比变化的影响，叶顶热流分布也基本不变；在超声速区域，随着膨胀比的减小，叶顶间隙内斜激波先出现强度增大，再转为正激波的现象，以实现逐渐升高的静压升，叶顶条带热流随之出现先减小后增大的现象。当涡轮级膨胀比降低至1.5时，叶顶间隙内激波结构完全消失，叶顶受分离再附着影响，呈现出较高热流分布。

Abstract:

The tip leakage flow enhances the tip heat transfer. In order to prevent highly loaded transonic rotor tip ablation， this paper used numerical methods to study the structure of a highly loaded transonic rotor tip shock wave system and its influence on the tip heat flow distribution， providing a guideline for transonic rotor tip cooling design. The results show that a separation bubble appears at the pressure side of the blade top clearance， and a strong oblique shock wave is generated by the turning of the high Mach number gas flow through the separation bubble. The oblique shock wave interference causes the separation of the boundary layer of the casing， and two reflected shock waves are generated by the two turns upstream and downstream of the separation bubble. Under the influence of two reflected shock waves， the thickening of the boundary layer at the tip of the blade leads to two tight low heat flux bands at the tip of the blade. Along the flow direction， the intensity of the shock wave at the downstream of the separation bubble gradually increases， resulting in the aggravation of the separation of the boundary layer of the casing， and the degree of the back and forth turning caused by the separation of the boundary layer increases， resulting in the increase of the intensity of the two reflected shock waves and the further reduction of the heat flux of the blade top strip. When the expansion ratio of the turbine stage is greater than 2.0， the Mach number in the subsonic region is basically not affected by the change of the expansion ratio and the heat flux distribution at the blade tip is basically unchanged due to the influence of the supersonic blockage. In the supersonic region， with the decrease of the expansion ratio， the oblique shock wave in the blade tip gap first increases in intensity and then turns into a positive shock wave to achieve a gradually increasing static pressure rise. The heat flux in the blade tip strip then decreases and then increases. When the expansion ratio of the turbine stage is reduced to 1.5， the shock wave structure in the tip gap disappears completely， and the tip is affected by separation and reattachment， showing a high heat flux distribution.