超声压气机叶型设计方法  

Supersonic Compressor Blade Profile Design Method

作　　者：邱名[1] 周正贵[1] 刘龙龙[1] 崔翠[1] 

机构地区：[1]南京航空航天大学能源与动力工程学院,江苏南京210016

出　　处：《航空学报》2014年第4期975-985,共11页Acta Aeronautica et Astronautica Sinica

摘　　要：在设计超声叶型时,为使得叶栅进口马赫数和气流角等于给定值,提出一种新的叶型参数化方法。该方法以经典唯一进气角计算方法为基础,将超声叶栅的唯一进气角和叶型几何形状关联起来,由进口马赫数和气流角确定吸力面进口段叶型;根据喉部面积、前后缘小圆半径、最大挠度和最大厚度等特征参数确定其他部分叶型。用此参数化方法设计了6个超声叶型,并用Fluent对设计结果进行了验证。结果表明,来流马赫数及进气角的设计值与Fluent求解结果基本一致,进气角最大误差仅为0.7°,进口马赫数最大误差仅为0.01;并且实现了多激波增压和减小激波损失等效果。In order to make inlet Mach number and flow angle equal to a given value in the design of supersonic compressor blade profile, this paper presents a new blade profile parameterization method. The method is based on solving unique inci- dence classic method, and gives the relationship between the blade profile shape with unique incidence by analysis and for- mula derivation. The front part of blade profile is determined by the inlet Mach number and flow angle, and the other part is determined by blade profile characteristic parameters such as the throat area ratio, the small circle radius at the leading edge and trailing edge, the maximum deflection, the maximum thickness, etc. The method is used to design 6 supersonic compressor blade profiles. And Fluent is used to verify the design results. The results show that the design value of inlet Mach number and flow angle are consistent with the Fluent solution result （the maximum error of the Mach number is only 0.01, and the maximum error of the flow angle is only 0.7°） ; the parameterization method is easy to achieve compression by much oblique shock and a normal shock.

关 键 词：超声叶栅 压气机 叶型设计 唯一进气角 叶型参数化 C形叶型 S形叶型 

分 类 号：V231.3[航空宇航科学与技术—航空宇航推进理论与工程]