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首页> 外文期刊>Journal of turbomachinery >Experimental/Numerical Crossover Jet Impingement in an Airfoil Leading-Edge Cooling Channel
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Experimental/Numerical Crossover Jet Impingement in an Airfoil Leading-Edge Cooling Channel

机译：翼型前沿冷却通道中的实验/数值交叉射流冲击

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摘要

Technological advancement in the gas turbine field demands high temperature gases impacting on the turbine airfoils in order to increase the output power as well as thermal efficiency. The leading edge is one of the most critical and life limiting sections of the airfoil which requires intricate cooling schemes to maintain a robust design. In order to maintain coherence with a typical external aerodynamic blade profile, cooling processes usually take place in geometrically-complex internal paths where analytical approaches may not provide a proper solution. In this study, experimental and numerical models simulating the leading-edge and its adjacent cavity were created. Cooling flow entered the leading-edge cavity through the crossover ports on the partition wall between the two cavities and impinged on the internal surface of the leading edge. Three flow arrangements were tested: (1) and (2) being flow entering from one side (root or tip) of the adjacent cavity and emerging from either the same side or the opposite side of the leading-edge cavity, and (3) flow entering from one side of the adjacent cavity and emerging from both sides of the leading-edge cavity. These flow arrangements were tested for five crossover-hole settings with a focus on studying the heat transfer rate dependency on the axial flow produced by upstream crossover holes (spent air). Numerical results were obtained from a three-dimensional unstructured computational fluid dynamics model with 1.1 × l0~6 hexahedral elements. For turbulence modeling, the realizable k-e was employed in combination with the enhanced wall treatment approach for the near wall regions. Other available RANS turbulence models with similar computational cost did not produce any results in better agreement with the measured data. Nusselt numbers on the nose area and the pressure/suction sides are reported for jet Reynolds numbers ranging from 8000 to 55,000 and a constant crossover hole to the leading-edge nose distance ratio, Z/D_h, of 2.81. Comparisons with experimental results were made in order to validate the employed turbulence model and the numerically-obtained results. Results show a significant dependency of Nusselt number on the axial flow introduced by upstream jets as it drastically diminishes the impingement effects on the leading-edge channel walls. Flow arrangement has immense effects on the heat transfer results. Discrepancies between the experimental and numerical results averaged between +0.3% and-24.5%; however, correlation between the two can be clearly observed.
机译：燃气轮机领域的技术进步要求高温气体撞击到涡轮机翼型上，以增加输出功率和热效率。前缘是机翼的最关键和最限制使用寿命的部分之一，它需要复杂的冷却方案来保持坚固的设计。为了保持与典型的外部气动叶片轮廓的连贯性，冷却过程通常在几何复杂的内部路径中进行，其中分析方法可能无法提供适当的解决方案。在这项研究中，创建了模拟前沿及其邻近空腔的实验和数值模型。冷却流通过两个空腔之间的分隔壁上的交叉端口进入前缘空腔，并撞击在前缘的内表面上。测试了三种流动布置：（1）和（2）是从相邻空腔的一侧（根部或尖端）进入并从前缘空腔的同一侧或相反侧流出的流，以及（3）从相邻腔的一侧进入并从前沿腔的两侧流出的气流。测试了这些流动装置的五个交叉孔设置，重点是研究传热速率对上游交叉孔（废空气）产生的轴向流的依赖性。数值结果是从一个具有1.1×l0〜6个六面体单元的三维非结构化计算流体动力学模型获得的。对于湍流建模，将可实现的k-e与针对近壁区域的增强壁处理方法结合使用。其他可用的具有类似计算成本的RANS湍流模型未产生任何与测量数据更好地吻合的结果。据报道，在雷诺数为8000到55,000的情况下，鼻子区域和压力/吸力侧的努塞尔数为2.81，前缘鼻子距离比Z / D_h为恒定的交叉孔。与实验结果进行了比较，以验证所采用的湍流模型和数值获得的结果。结果显示，努塞尔数对上游射流引入的轴向流有很大的依赖性，因为它大大减小了对前缘通道壁的撞击作用。流动布置对传热结果有巨大影响。实验结果与数值结果之间的差异平均在+ 0.3％至-24.5％之间；但是，可以清楚地观察到两者之间的相关性。

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  • 来源
    《Journal of turbomachinery》 |2013年第1期|011037.1-011037.12|共12页
  • 作者

    K. Elebiary; M. E. Taslim;

  • 作者单位

    Mechanical and Industrial Engineering Department,Northeastern University,Boston, MA 02115;

    Mechanical and Industrial Engineering Department,Northeastern University,Boston, MA 02115;

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  • 正文语种 eng
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