摘要
超音压气机叶型面临着进口高马赫数带来的激波损失增加和激波与附面层之间强烈干涉等难题,这些因素严重限制了超音叶型的实际应用,而超音叶型的造型方法对叶型流动和损失影响显著。本文以唯一进气角理论为基础,将进口马赫数、气流角和叶型进口段设计参数关联,采用前尾缘精确控制方法,结合多段低阶Bezier曲线发展了超音压气机叶型参数化造型方法。利用该参数化方法对ARL-SL19叶型进行了拟合,并通过Fluent数值计算验证了造型方法的高精度特性。在此基础上,设计了进口马赫数适用范围1.6~1.7的超音叶型,并开展了喉道面积和喉道位置对叶栅流动和性能的影响机制研究。结果表明,喉道参数主要通过改变激波系的结构而影响超音叶栅的性能和流动;而喉道面积比是影响总压损失的关键因素;叶栅的喉道面积比越小、喉道位置越靠前,其总压损失系数越小,但做功的有效范围也越小。
The supersonic compressor blade profle is faced with some problems,such as the increase of shock wave loss and the intense interaction between shock wave and boundary layer.The modeling method of supersonic blade profile has a significant effect on the flow and loss of the blade.Therefore,the inlet Mach number,the inlet angle and the design parameters of the blade inlet are correlated based on the theory of the unique incidence.A parameterized modeling method for supersonic compressor blade profiles was developed based on multi-stage low-order Bezier curves,incorporating a precise control method at the leading and trailing edges.The ARL-SL19 blade profile is fitted using the parameterization method,and the high precision of the modeling method is verified by Fluent numerical calculation.Based on this,a supersonic blade profile with inlet Mach number range of 1.6~1.7 is designed,and the effects of throat area and throat position on cascade flow and performance are investigated.The results indicated that throat parameters impact the performance and flow of supersonic cascades mainly by altering the structure of the shock wave system,and the throat area ratio is the key factor affecting the total pressure loss.The smaller the throat area ratio and the higher the throat position,the smaller the total pressure loss coefficient,but the smaller the effective range of work.
作者
唐宇安
阳诚武
赵胜丰
卢新根
邓坤盈
TANG Yu'an;YANG Chengwu;ZHAO Shengfeng;LU Xingen;DENG Kunying(The institute of Engineering Thermophysics,Chinese Academy of Sciences,Beijing 100190,China;School of Aeronautics and Astronautics,University of Chinese Academy of Sciences,Beijing 100049,China;National Key Laboratory of Light Turbine Power,Beijing 100190,China)
出处
《工程热物理学报》
北大核心
2025年第8期2544-2556,共13页
Journal of Engineering Thermophysics
基金
航空发动机及燃气轮机重大专项项目(No.J2019-II-0003-0023)
航空发动机及燃气轮机基础科学中心项目(No.P2022-B-II-002-001)
中国科学院青年创新促进会(No.2020148)。
