大扩张角涡轮过渡段性能试验和数值研究  

Experimental and Numerical Investigation on the Performance of an Aggressive Intermediate Turbine Duct

作　　者：施鎏鎏 罗华玲 张颜[2] 刘火星[2] 

机构地区：[1]中航商用航空发动机有限责任公司,上海201108 [2]北京航空航天大学航空发动机气动热力国家科技重点实验室,北京100191

出　　处：《航空发动机》2016年第1期75-78,共4页Aeroengine

摘　　要：为研究某型大扩张角涡轮过渡段气动性能,对过渡段内部流场进行了详细的试验测量,同时采用CFD数值模拟对过渡段内部流场进行仿真,并与试验结果进行对比分析。结果表明:过渡段机匣表面流动受强逆压梯度影响,容易发生流动分离;轮毂表面流场受支板前缘冲击绕流的影响,呈现周向不均匀性。来流气流角使得过渡段内部流场向支板一侧偏斜,随着气流角的增大,过渡段总压损失增大。CFD模拟结果与试验测量结果吻合较好,均能很好地捕捉流场的细节特征;过渡段进、出口总压恢复系数随着来流气流角的增大而减小,CFD模拟和试验测量值的偏差约为0.2%。To study the aerodynamic performance of an aggressive intermediate turbine duct, experimental measurements and CFD simulations of the duct flow field were carried out, and the CFD results were compared with experimental data for validation. The results show that the intermediate turbine duct casing is dominated by strong adverse pressure gradient, and flow separation is easy to occur. The circumferential difference of the flow filed is found near hub, which is caused by the disturbance of the impinging flow at the strut leading edge. The flow field inclines to one side of the strut as the incoming flow angle increases, the total pressure loss increases as well. The CFD results agree reasonably well with the experimental results, both of them can capture the main flow features. Total pressure recovery coefficient decreases as the incoming flow angle increases, discrepancy between the CFD prediction and experimental measurement is found to be merely 0.2%.

关 键 词：大扩张角 涡轮过渡段 气动性能 流动分离 气流角 试验 大涵道比发动机 

分 类 号：V235.1[航空宇航科学与技术—航空宇航推进理论与工程]