摘要
为了探究航空发动机过渡段内部流场的流动机理和损失机制,促进航空发动机的整机性能提升,采用数值模拟方法对某航空发动机增压级过渡段进行了流场分析,并与实验结果进行对比验证。结果表明:过渡段出口流场在20%~90%径向位置处分布较为均匀,气流的总压损失主要集中在支板尾缘和上下壁面附近;主要损失机理包括下壁面曲率引起的附面层分离和支板表面附面层分离;角区的二次流扰动沿流向始于下壁面曲率最大处;总压损失会随流速的增加而增大,流速越大的总压损失增量也会越大,流速为140~170 m/s时总压损失系数增量是流速为80~110 m/s时的7倍。
In order to explore the flow mechanism and loss mechanism of the internal flow field in the transition stage of aircraft engine and promote the overall performance improvement of aircraft engines,this article conducts numerical simulation and flow field analysis on the transition stage of a certain aircraft engine,and compares it with experimental results.The results show that the flow field at the outlet of transition stage is evenly distributed in the radial position of the 20%-90%flow channel,and the total pressure loss of air flow at the outlet is mainly concentrated near the trailing edge of the support plate and the upper and lower walls;the loss in the flow channel is caused by the combined curvature of the lower wall and the boundary layer seperation on the support plate surface;the secondary flow disturbance in the corner region starts along the flow direction at the point where the curvature of the lower wall is maximum;and the total pressure loss will increase with the increase of flow rate.The larger the flow rate,the greater the increment of total pressure loss.The increment of total pressure loss at flow rate of 140 to 170 m/s is 7 times that at flow rate of 80 to 110 m/s.
作者
张劭钦
刘臻梁
何中海
吴亚东
ZHANG Shaoqin;LIU Zhenliang;HE Zhonghai;WU Yadong(School of Mechanical Engineering,Shanghai Jiao Tong University,Shanghai,China,200240)
出处
《热能动力工程》
CAS
CSCD
北大核心
2024年第6期40-48,共9页
Journal of Engineering for Thermal Energy and Power
基金
国家重点专项(2019-Ⅱ-0004-0024)。
关键词
轴流压气机
过渡段
数值模拟
角区分离
1/3倍频程
axial compressor
transition stage
numerical simulation
corner separation
one-third octave
