基于RBF曲线的涡轮叶栅激波控制与优化  

Shock Control and Optimization of Turbine Blade Cascades Based on RBF Curves

作　　者：张蔚然 陈建[1] 苏欣荣 袁新[2] Weiran Zhang;Jian Chen;Xinrong Su;Xin Yuan(School of Energy and Power Engineering,University of Shanghai for Science and Technology,Shanghai;Department of Energy and Power Engineering,Tsinghua University,Beijing)

机构地区：[1]上海理工大学能源与动力工程学院,上海 [2]清华大学能源与动力工程系,北京

出　　处：《建模与仿真》2025年第2期435-448,共14页Modeling and Simulation

基　　金：国家自然科学基金面上项目(No.52276031);国家科技重大专项(No.J2019-II-0008-0028)。

摘　　要：随着航空发动机压比不断提高,高压涡轮叶片的出口马赫数逐渐提高,流道中不可避免地出现强激波及其导致的激波/边界层干涉现象,导致了明显的气动损失。因此,发展激波/边界层干涉的有效控制方法对于设计高负荷、低损失的涡轮叶片具有很好的指导意义。本文提出了基于RBF曲线造型的激波控制方法并发展了相应的优化设计方法。在激波/边界层相互作用区域基于RBF曲线实现局部叶型的精细造型,利用RBF曲线灵活多变的凹凸结构在叶栅中产生对应的压缩波和膨胀波,从而实现对激波系的精细控制。结合参数化建模及优化算法进行迭代寻优,从而有效削弱激波强度、减弱激波/边界层干涉。数值结果表明:优化结果在较宽的出口Ma数范围内均能有效减小损失,在设计工况现有鼓包方法相对于基准叶型减小总压损失系数7.1%,而本文方法相对于基准叶型减小总压损失系数15.3%,明显优于现有方法。流场分析表明优化后吸力面侧激波和尾缘两侧燕尾激波的强度均显著降低,同时发现坡度较缓的背风坡能够使气流平稳流动以避免流动分离,从而进一步降低损失。本文创新性地提出基于RBF曲线造型的涡轮叶栅激波控制与优化方法,克服了传统鼓包方法中设计域有限、仅限于在基础鼓包结构上微调等局限性,利用RBF曲线的灵活性和多样性实现了较宽工况范围内的涡轮叶栅激波调控。As the compression ratio of aircraft engines continues to increase,the exit Mach number of highpressure turbine blades gradually rises,leading to the inevitable occurrence of strong shock waves and the associated shock/boundary layer interaction phenomena within the flow passage.This results in significant aerodynamic losses.Therefore,developing effective control methods for shock/boundary layer interaction is crucial for designing high-load,low-loss turbine blades.This paper proposes a shock control method based on Radial Basis Function(RBF)curve modeling and develops corresponding optimization design techniques.In the region of shock/boundary layer interaction,local blade shapes are finely modeled using RBF curves.The flexible and variable convexconcave structures generated by the RBF curves produce corresponding compression and expansion waves within the blade row,thereby achieving precise control of the shock wave system.By combining parametric modeling with optimization algorithms for iterative optimization,the method effectively reduces shock intensity and mitigates shock/boundary layer interaction.Numerical results indicate that the optimization results significantly reduce losses across a wide range of exit Mach numbers.At the design conditions,the existing bump method reduces the total pressure loss coefficient by 7.1%compared to the baseline blade,while the method proposed in this paper achieves a reduction of 15.3%,demonstrating a clear advantage over existing methods.Flow field analysis shows that the intensity of the shock waves on the suction side and the trailing edge’s wingtip is significantly reduced after optimization.Additionally,it is found that a milder leeward slope can facilitate smooth airflow to avoid flow separation,thereby further reducing losses.This paper innovatively presents a shock control and optimization method for turbine blade rows based on RBF curve modeling,overcoming the limitations of traditional bump methods,which are restricted by the design domain and only allow for mi

关 键 词：跨声速 激波 RBF曲线 总压损失系数 边界层 

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]