提高超声速压气机级喘振裕度方法研究

doi:10.13675/j.cnki. tjjs. 180550

推进技术 ›› 2019, Vol. 40 ›› Issue (8): 1780-1791.DOI: 10.13675/j.cnki. tjjs. 180550

提高超声速压气机级喘振裕度方法研究

南京航空航天大学能源与动力学院

发布日期:2021-08-15

Study of Improving Surge Margin for aSupersonic Compressor Stage

College of Energy and Power Engineering,Nanjing University of Aeronautics and Astronautics,Nanjing 210016,China

Published:2021-08-15

摘要/Abstract

摘要： 在超声速压气机气动设计时，为实现设计点高性能和宽喘振裕度，提出采用优化方法以设计点性能为目标进行叶片设计，通过转/静子叶片几何手动修改提高压气机喘振裕度。以NASA Rotor 37为原型，应用此方法进行更高性能超声速压气机转子气动设计，并匹配静子，构成压气机级。结果表明：超声速压气机转子通道激波推出和静子大攻角分离是失速发生的主要原因，因此分别进行转子叶片前掠设计、改变叶尖稠度，以控制激波位置，单转子喘振裕度可从约7%提高到18%以上；静子上采用前掠、切向弯、修改叶片数及几何进口角等措施，最终将此压气机级的喘振裕度由约18%提高到30%以上。

关键词: 压气机;流场计算;喘振裕度;叶片前掠;叶尖稠度;几何进口角

Abstract: In order to achieve high performance at design point and wide surge margin in the aerodynamic design of a supersonic compressor, optimization method was adopted for the blade design at design point, and then the geometry of rotor/stator blade was modified manually to improve the compressor surge margin. This method was applied to the aerodynamic design of the higher performance supersonic compressor rotor and the matching stator to form the compressor stage based on NASA Rotor 37 design parameters. The results show that the supersonic compressor stall occurs because rotor shock is pushed out passage at the tip, or stator boundary layer separates as a result of large incidence. The surge margin of the single rotor improved from 7% to above 18% if measures， such as forward sweep and increasing blade tip solidity， are taken to control the shock position in the passage. Beneficial methods such as forward sweep tangential lean, modification of blade number and geometric inlet angle are also adopted for the stator. The predicted surge margin of the compressor stage finally improved from 18% to above 30%.

Key words: Compressor;Flow field calculation;Surge margin;Forward sweep;Blade tip solidity;Geometric inlet angle

. 提高超声速压气机级喘振裕度方法研究[J]. 推进技术, 2019, 40(8): 1780-1791.

参考文献

[1] Smith L H. Axial Compressor Aerodesign Evolution at General Electric[J]. ASME Journal of Turbomachinery, 2002, 124(3): 321-330.
[2] Hah C， Puterbaugh S L， Wadia A R. Control of Shock Structure and Secondary Flow Field Inside Transonic Compressor Rotors Through Aerodynamic Sweep[R]. ASME 98-GT-561.
[3] Denton J D， Xu L. The Effects of Lean and Sweep on Transonic Fan Performance[R]. ASME GT2002-30327.
[4] Denton J D， Xu L. The Exploitation of Three-Dimensional Flow in Turbomachinery Design[J]. Proceedings of the Institution of Mechanical Engineers， 1999， 213(2): 125-137.
[5] Yamaguchi N， Tominaga T， Hattori S， et al. Secondary-Loss Reduction by Forward-Skewing of Axial Compressor Rotor Blading[C]. Japan： The International Gas Turbine Congress， 1991.
[6] Passrucker H， Engber M， Kablitz S， et al. Effect of Forward Sweep in a Transonic Compressor Rotor[J]. Proceedings of the Institution of Mechanical Engineers， 2003， 217(4): 357-365.
[7] Gümmer V， Wenger U， Kau H P. Using Sweep and Dihedral to Control Three-Dimensional Flow in Transonic Stators of Axial Compressors[J]. ASME Journal of Turbomachinery， 2001， 123(1): 40-48.
[8] Braun M， Seume J R. Forward Sweep in a Four-Stage High-Speed Axial Compressor[R]. ASME GT2006-90218.
[9] 冯秀莲，金东海，桂幸民. 叶片弯掠对压气机静子叶片气动性能影响的三维数值模拟[J]. 航空动力学报， 2009， 24(10): 2338-2343.
[10] 刘宝杰，邹正平，严明，等. 叶轮机计算流体动力学技术现状与发展趋势[J]. 航空学报， 2002， 23(5): 394-404.
[11] Reid L， Moore R D. Design and Overall Performance of Four Highly Loaded， High Speed Inlet Stages for an Advanced High-Pressure-Ratio Core Compressor[R]. NASA-TP-1337， 1978.
[12] Suder K L. Experimental Investigation of the Flow Field in a Transonic， an Axial Flow Compressor with Respect to the Development of Blockage and Loss[D]. Cleveland: Case Western Reserve University， 1996.
[13] Spalart P R， Allmaras S R. A One-Equation Turbulence Model for Aerodynamic Flows[J]. Recherche Aerospatiale， 1994， 1(1): 5-21.
[14] 宁方飞，徐力平. Spalart-Allmaras湍流模型在内流流场数值模拟中的应用[J]. 工程热物理学报， 2001， 22(3): 304-306.
[15] Jameson A， Schmidt W， Turkel E. Numerical Solution of the Euler Equations by Finite Volume Methods Using Runge-Kutta Time Stepping Schemes[C]. Palo Alto: AIAA 14th Fluid and Plasma Dynamics Conference， 1981.
[16] Subramanian S， Bozzola R. Application of Runge-Kutta Time Marching Scheme for the Computation of Transonic Flows in Turbomachines[C]. California: 21st Joint Propulsion Conference， 1987.
[17] 邱名，周正贵，刘龙龙，等. 超声压气机叶型设计方法[J]. 航空学报， 2014， 35(4): 975-985.
[18] 周正贵，邱名，徐夏，等. 压气机/风扇二维叶型自动优化设计[J]. 航空学报， 2011， 32(11): 1987-1997.
[19] 徐国华. 遗传算法的改进及风扇/压气机通用优化平台的研制[D]. 南京:南京航空航天大学， 2013.
[20] 周正贵. 压气机叶片自动优化设计[J]. 航空动力学报， 2002， 17(3): 305-308.

提高超声速压气机级喘振裕度方法研究

Study of Improving Surge Margin for aSupersonic Compressor Stage

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