• 主题
高级搜索  >
历史搜索 清空历史
首页> 外文会议>ASME turbo expo: turbine technical conference and exposition >DEVELOPMENT OF DESIGN METHOD FOR SUPERSONIC TURBINE AEROFOILS NEAR THE TIP OF LONG BLADES IN STEAM TURBINES PART 2: CONFIGURATION DETAILS AND VALIDATION
【24h】

DEVELOPMENT OF DESIGN METHOD FOR SUPERSONIC TURBINE AEROFOILS NEAR THE TIP OF LONG BLADES IN STEAM TURBINES PART 2: CONFIGURATION DETAILS AND VALIDATION

机译：汽轮机长叶片尖端附近的超音速涡轮机翼设计方法的发展，第2部分：配置细节和验证

获取原文
页面导航
  • 摘要
  • 著录项
  • 相似文献
  • 相关主题

摘要

Both inflow and outflow velocities near the blade tip become supersonic when the blade length exceeds a threshold limit. The aerofoil near the tip of such a long blade has four features that demand an original supersonic turbine aerofoil design: supersonic flow in the entire field, high reaction, large stagger angle, and large pitch-to-chord ratio. This paper describes design method development for the supersonic turbine aerofoil. First, the aerofoil shape is defined using a curve with continuity in the gradient of the curvature. Second, six loss generation mechanisms are clarified by turbulent flow analysis. Third, an allowable design space between the pitch-to-chord ratio, the stagger angle and the axial-chord-to-pitch ratio is clarified by formulating three geometrical constraints to accelerate supersonic flow smoothly. When there is no solution in the theoretically allowable design space because of the large pitch-to-chord ratio, methods to reduce shock wave losses are proposed. Increasing the outlet metal angle of the pressure surface by around 10 deg from the theoretical outlet flow angle reduces the loss caused by the trailing shock wave. The physical mechanism for this is as follows: the increased outlet metal angle increases the outlet flow passage area so that the overexpansion is suppressed downstream from the flow passage. Fourth, both a cusped leading edge and an upstream pressure surface which has both an angle corresponding to the inflow angle and near-zero curvature can reduce the loss caused by the leading shock wave and satisfy the unique incidence relation. Finally, the aerodynamic performance of the supersonic turbine cascade and the design method are validated by supersonic cascade wind tunnel tests.
机译：当叶片长度超过阈值极限时，叶片尖端附近的流入和流出速度都将变为超音速。如此长的叶片尖端附近的翼型具有四个特征，需要原始的超音速涡轮机翼型设计：整个领域的超音速流动，高反作用力，大的交错角和大的音高和弦比。本文介绍了超音速涡轮翼型的设计方法开发。首先，使用曲率梯度连续的曲线定义机翼形状。其次，通过湍流分析阐明了六种损耗产生机理。第三，通过制定三个几何约束条件来平稳地加速超声速，澄清了螺距与弦距比，交错角和轴向螺距与螺距之间的允许设计空间。当由于大的音高和弦比而在理论上允许的设计空间中没有解决方案时，提出了减少冲击波损耗的方法。将压力表面的出口金​​属角度从理论出口流动角度增加大约10度，可以减少后随冲击波引起的损耗。为此的物理机制如下：增大的出口金属角度增加了出口流道面积，从而抑制了流道下游的过度膨胀。第四，尖锐的前缘和具有与流入角相对应的角度和接近零曲率的上游压力表面都可以减少由前导冲击波引起的损失并满足唯一的入射关系。最后，通过超音速叶栅风洞试验验证了超音速涡轮叶栅的空气动力性能和设计方法。

著录项

相似文献

  • 外文文献
  • 中文文献
  • 专利
获取原文

客服邮箱：kefu@zhangqiaokeyan.com

京公网安备：11010802029741号 ICP备案号：京ICP备15016152号-6 六维联合信息科技 (北京) 有限公司©版权所有

  • 客服微信

  • 服务号