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首页> 外文会议>ASME turbo expo: turbomachinery technical conference and exposition >AERODYNAMIC MEASUREMENTS AND ANALYSIS IN A FIRST STAGE NOZZLE GUIDE VANE PASSAGE WITH COMBUSTOR LINER COOLING, SLOT FILM COOLING AND ENDWALL CONTOURING
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AERODYNAMIC MEASUREMENTS AND ANALYSIS IN A FIRST STAGE NOZZLE GUIDE VANE PASSAGE WITH COMBUSTOR LINER COOLING, SLOT FILM COOLING AND ENDWALL CONTOURING

机译：燃烧室衬管冷却，槽膜冷却和端壁修整的第一阶段喷嘴导流通道的气动测量和分析

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Endwalls impose a challenge to cool because of the complex system of secondary flows and separation lines disrupting surface film coolant coverage. The interaction of film cooling flows with secondary flow structures is coupled. The momentum exchange of the film coolant with the mainstream affect the formation the secondary flows, which in turn affect the coolant coverage. Therefore, to develop better endwall cooling schemes, a good understanding of passage aerodynamics as affected by interactions with coolant flows is required. This study presents experimental and computational results for cascade representing the first stage nozzle guide vane of a high-pressure gas turbine. The cascade is subsonic, linear, and stationary with an axisymmetrically-contoured endwall. Two cooling flows are simulated; upstream combustor liner coolantin the form of an aero-thermal profile simulated in the approach flow and endwall slot film cooling, which is injected immediately upstream of the passage inlet. The experiment is run with engine representative combustor exit flow turbulence intensity and integral length scales, with high turbine passage exit Reynolds number of 1.61×10~6. Measurements are performed with various slot film cooling mass flow rate to mainstream flow rate ratios (MFR). Aerodynamic effects are documented with five-hole probe measurements at the exit plane. Varying the slot film cooling MFR results in minimal effects on total pressure loss for the range tested. Vorticity distributions show a very thin, yet intense, cross-pitch flow on the contoured endwall side. Coolant distribution fields that were previously presented for the same cascade are discussed in context of the aerodynamic measurements. A coolant vorticity parameter presenting the advective mixing of the coolant due to secondary flow vorticity is introduced. This parameter gives developers a new prospective on aerodynamic-thermal performance associated with cooled turbine endwall. The numerical study is conducted for the same test section geometry and is run under the same conditions. The applicability of using RANS turbulence closure models for simulating this type of flow is discussed. The effects of including the combustor coolant in the approach flow is also briefly discussed in context of the numerical results
机译：端壁对冷却提出了挑战，因为复杂的二次流系统和分离线破坏了表面膜冷却剂的覆盖范围。膜冷却流与次级流结构的相互作用被耦合。薄膜冷却液与主流的动量交换会影响二次流的形成，进而影响冷却液的覆盖范围。因此，为了开发更好的端壁冷却方案，需要对与冷却剂流相互作用影响的通道空气动力学有很好的理解。这项研究给出了代表高压燃气轮机第一级喷嘴导流叶片的级联的实验和计算结果。级联是亚音速的，线性的，并带有轴对称轮廓的端壁。模拟了两个冷却流;上游燃烧室衬套冷却剂的形式为在进场气流和端壁缝隙薄膜冷却中模拟的空气热曲线，该冷却剂被直接注入通道入口的上游。实验以发动机代表燃烧器出口流的湍流强度和积分长度标度进行，涡轮通道出口的雷诺数高，为1.61×10〜6。使用各种缝隙薄膜冷却质量流量与主流流量之比（MFR）进行测量。空气动力学效应通过出口平面上的五孔探头测量得到记录。在测试范围内，改变缝膜冷却MFR对总压力损失的影响最小。涡度分布在轮廓化的端壁侧显示出非常薄而强烈的交叉螺距流量。在空气动力学测量的背景下讨论了先前针对同一级联提供的冷却剂分布场。引入冷却剂涡度参数，该参数表示由于二次流动涡度而引起的冷却剂的对流混合。该参数为开发人员提供了与冷却后的涡轮机端壁相关的空气动力-热力性能的新前景。数值研究是针对相同的测试截面几何形状并且在相同的条件下进行的。讨论了使用RANS湍流闭合模型来模拟此类流动的适用性。在数值结果的背景下，还简要讨论了将燃烧器冷却剂包括在进场流中的影响

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