考虑转捩影响的地面重型燃气轮机压气机叶型优化设计
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摘要
随着数值模拟技术的不断成熟,基于优化理论的气动设计方法逐渐兴起并吸引了众多学者进行研究与应用。本文的工作就是从这一方面着手,将优化理论引入到地面重型燃气轮机压气机叶型设计中,建立可用于地面重型燃气轮机压气机的二维气动优化设计体系。在叶轮机械中,转捩过程受到来流湍流度、来流马赫数、压力梯度及其它诸多因素的影响。本文中,优化设计考虑到地面重型燃气轮机压气机入口气体在较高雷诺数及较高湍流度下流动转捩带来的影响。
本文首先对某地面重型燃气轮机压气机的末级静叶中径叶型进行参数化。叶片参数化造型方法有很多种,本文采用的“中弧线贴厚度分布法”的参数化造型方法。是在已有方法的基础上改进得到的一种新的叶片参数化造型方法,适用于亚音速压气机叶片的参数化优化设计。
建立具有自动优化设计能力的平台。该平台由自主开发的二维叶型生成程序、网格生成软件GAMBIT、流场计算分析软件CFX以及自主开发的优化模块组成。整个过程可以靠计算机自动完成,省去了手动调节浪费时间的弊端,优化的叶型范围更加宽广。
以某地面重型燃气轮机压气机末级静叶中径叶型为优化对象,以总压损失系数为目标函数,流场分析中考虑转捩情况,利用iSIGHT优化器进行优化计算。优化后,叶型气动性能有较明显改善。设计工况下叶型总压损失系数较原型下降了3.63%,气流折转角增大了1.44%,静压比升高了0.11%,扩压因子增大了1.04%。比较分析发现优化叶型通过减小吸力面的逆压梯度抑制附面层的发展来减小叶型尾迹,降低尾迹损失,总压损失系数下降。同时因为气流转折角增大,静压比升高,扩压因子增大,从而使优化叶型性能高于原型。在非设计工况条件下,较低马赫数时,优化叶型的低损失工作范围较大,较高马赫数时,优化叶型的低损失工作范围较小。
将优化出的叶型在保证弦长、最大厚度、最大厚度相对位置、前缘半径、安装节距、安装角、前缘角、后缘角、控制叶型中弧线形状的无量纲参数不变的前提下,应用到原始叶片各截面叶型上生成新的优化叶型。将各截面上新的优化叶型按照原始叶片的积叠方式生成优化叶片。对原始叶片与优化叶片进行数值模拟,优化叶片在主流区内通过减小吸力面的逆压梯度抑制附面层的发展来减小叶型尾迹,降低尾迹损失,总压损失系数下降。在端壁角区范围内,通过控制角区分离来减小损失,总压损失系数下降。同时因为在不同叶高处,优化叶片的气流转折角增大,静压比升高,从而使优化叶片性能高于原始叶片。
With the mature development of numerical simulation technology, aerodynamic design method based on optimization theory rises gradually. And it has become focus of many scholars to study. From this point the work establishes aerodynamic optimization design system that is applicable to heavy-duty gas turbine compressor by introducing optimization theory to heavy-duty gas turbine compressor blade design. In turbomachine there are so many influencing factors of transition process such as inflow turbulence intensity, inlet mach number, pressure gradient and so on. In this paper the factor that air flow transition at high reynolds number or high turbulence intensity in ground-based heavy-duty gas turbine compressor is also discussed.
Firstly the paper parameterizes middle-radius profile of stator blade in heavy-duty gas turbine compressor which is studied. There is many blade parameterization modeling method. The work adopts with“camber line add thick”. It is a new blade parameterization modeling method improved on basis of existing methods, and it is suitable for subsonic compressor blade.
Then a design system has been established with the ability of automatic optimization. The system consists of in-house program for 2D compressor blade design, GAMBIT for grid generating, CFX for flow field computation and analysis and in-house codes for optimization. The entire process is auto-completed by computer; in addition, the range of blade profile is much larger.
In this paper, a compressor blade of a ground-based heavy-duty gas turbine has been optimized by the design system referred above. Total pressure loss coefficient is considered as object function of the optimization, and transition is especially considered in flow field analysis, and the optimization calculation is performed with ISIGHT Optimizer. After optimization, aerodynamic performance of the blade is improved, in the design point, the total pressure loss coefficient is decreased by 3.63%, the flow angle is increased by 1.44%, static pressure ratio is increased by 0.11% and diffusion factor is increased by 1.04%. Compared to the original blade profile, it can be seen that, the optimized profile gets better performance by decreasing the adverse pressure gradient of surface and inhibiting the separation of boundary layer. The trailing edge loss and total pressure loss ratio are both decreased under this profile. On off-design condition, the work range of low loss is larger under low mach number and the range is small under high mach number.
The optimized blade profile was applied to each section of the original blade to form a new optimized blade profile with chord length, maximum thickness, relative position of maximum thickness, radius of leading edge, radius of trailing edge, installation pitch, installation angle, leading edge angle and trailing edge angle unchanged. The new blade was stacked by each section of optimized blade profile in the same way as the original blade. The three-dimensional numerical simulation is carried on for the original blade profile and the optimized profile with NUMECA. It shows that the optimized profile gets better performance by decreasing the adverse pressure gradient of surface and inhibiting the separation of boundary layer. The trailing edge loss and total pressure loss coefficient are both decreased under this profile, total pressure ratios of different blade height are increased; in the endwall corner, total pressure loss coefficient is decreased and the separation is restrained, total pressure ratios of different blade height are increased. Meanwhile, the flow angles and static pressure ratios of different blade height are increased, so the performance of optimized blade is much better than the original blade.
本文首先对某地面重型燃气轮机压气机的末级静叶中径叶型进行参数化。叶片参数化造型方法有很多种,本文采用的“中弧线贴厚度分布法”的参数化造型方法。是在已有方法的基础上改进得到的一种新的叶片参数化造型方法,适用于亚音速压气机叶片的参数化优化设计。
建立具有自动优化设计能力的平台。该平台由自主开发的二维叶型生成程序、网格生成软件GAMBIT、流场计算分析软件CFX以及自主开发的优化模块组成。整个过程可以靠计算机自动完成,省去了手动调节浪费时间的弊端,优化的叶型范围更加宽广。
以某地面重型燃气轮机压气机末级静叶中径叶型为优化对象,以总压损失系数为目标函数,流场分析中考虑转捩情况,利用iSIGHT优化器进行优化计算。优化后,叶型气动性能有较明显改善。设计工况下叶型总压损失系数较原型下降了3.63%,气流折转角增大了1.44%,静压比升高了0.11%,扩压因子增大了1.04%。比较分析发现优化叶型通过减小吸力面的逆压梯度抑制附面层的发展来减小叶型尾迹,降低尾迹损失,总压损失系数下降。同时因为气流转折角增大,静压比升高,扩压因子增大,从而使优化叶型性能高于原型。在非设计工况条件下,较低马赫数时,优化叶型的低损失工作范围较大,较高马赫数时,优化叶型的低损失工作范围较小。
将优化出的叶型在保证弦长、最大厚度、最大厚度相对位置、前缘半径、安装节距、安装角、前缘角、后缘角、控制叶型中弧线形状的无量纲参数不变的前提下,应用到原始叶片各截面叶型上生成新的优化叶型。将各截面上新的优化叶型按照原始叶片的积叠方式生成优化叶片。对原始叶片与优化叶片进行数值模拟,优化叶片在主流区内通过减小吸力面的逆压梯度抑制附面层的发展来减小叶型尾迹,降低尾迹损失,总压损失系数下降。在端壁角区范围内,通过控制角区分离来减小损失,总压损失系数下降。同时因为在不同叶高处,优化叶片的气流转折角增大,静压比升高,从而使优化叶片性能高于原始叶片。
With the mature development of numerical simulation technology, aerodynamic design method based on optimization theory rises gradually. And it has become focus of many scholars to study. From this point the work establishes aerodynamic optimization design system that is applicable to heavy-duty gas turbine compressor by introducing optimization theory to heavy-duty gas turbine compressor blade design. In turbomachine there are so many influencing factors of transition process such as inflow turbulence intensity, inlet mach number, pressure gradient and so on. In this paper the factor that air flow transition at high reynolds number or high turbulence intensity in ground-based heavy-duty gas turbine compressor is also discussed.
Firstly the paper parameterizes middle-radius profile of stator blade in heavy-duty gas turbine compressor which is studied. There is many blade parameterization modeling method. The work adopts with“camber line add thick”. It is a new blade parameterization modeling method improved on basis of existing methods, and it is suitable for subsonic compressor blade.
Then a design system has been established with the ability of automatic optimization. The system consists of in-house program for 2D compressor blade design, GAMBIT for grid generating, CFX for flow field computation and analysis and in-house codes for optimization. The entire process is auto-completed by computer; in addition, the range of blade profile is much larger.
In this paper, a compressor blade of a ground-based heavy-duty gas turbine has been optimized by the design system referred above. Total pressure loss coefficient is considered as object function of the optimization, and transition is especially considered in flow field analysis, and the optimization calculation is performed with ISIGHT Optimizer. After optimization, aerodynamic performance of the blade is improved, in the design point, the total pressure loss coefficient is decreased by 3.63%, the flow angle is increased by 1.44%, static pressure ratio is increased by 0.11% and diffusion factor is increased by 1.04%. Compared to the original blade profile, it can be seen that, the optimized profile gets better performance by decreasing the adverse pressure gradient of surface and inhibiting the separation of boundary layer. The trailing edge loss and total pressure loss ratio are both decreased under this profile. On off-design condition, the work range of low loss is larger under low mach number and the range is small under high mach number.
The optimized blade profile was applied to each section of the original blade to form a new optimized blade profile with chord length, maximum thickness, relative position of maximum thickness, radius of leading edge, radius of trailing edge, installation pitch, installation angle, leading edge angle and trailing edge angle unchanged. The new blade was stacked by each section of optimized blade profile in the same way as the original blade. The three-dimensional numerical simulation is carried on for the original blade profile and the optimized profile with NUMECA. It shows that the optimized profile gets better performance by decreasing the adverse pressure gradient of surface and inhibiting the separation of boundary layer. The trailing edge loss and total pressure loss coefficient are both decreased under this profile, total pressure ratios of different blade height are increased; in the endwall corner, total pressure loss coefficient is decreased and the separation is restrained, total pressure ratios of different blade height are increased. Meanwhile, the flow angles and static pressure ratios of different blade height are increased, so the performance of optimized blade is much better than the original blade.
引文
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2陈云永,刘波,马聪慧,高丽敏,石建成,温泉.全三维、多叶排内外涵风扇压气机叶型优化研究.航空动力学报. 2008年3月第23卷第3期
3 Ku¨sters, B., Schreiber, H. A., Ko¨ller, U., and Mo¨nig, R., 1999,‘‘Development of Advanced Compressor Airfoils for Heavy-Duty Gas Turbines: Part II—Experimental and Theoretical Analysis,’’ASME J. Turbomach., 122, this issue, pp. 406–415
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6 Dawes.W.N. A Numerical Analysis of the Three-Dimensional Viscous Flow in a Transonic Compressor Rotor and Comparison With Experiment. 1986,ASME 86-GT-16
7 Hah,C., J.D. A Numerical Modeling of Endwall and Tip Clearance Flow of an Isolated Compressor Rotor. ASME Paper 85-GT-116, 1985
8 Arnone.Viscous Analysis of Three Dimensional Rotor Flow Using a Multigrid Method. Journal of Turbomahcinery. 1994, 116(2): 435-445
9 L.He, W.Ning. Efficient Approach for Analysis of Unsteady Viscous Flows in Turbomachines. AIAA Journal. 1998, 36(11):2005-2012
10 Denton J D, Dawes, W N. Computational fluid dynamics for turbomachinery design[c]//proceeding of instin Mech Engrs,1999,213(part c):107-124
11李志广.发展我国重型燃气轮机产业的可行技术途径.航空发动机. 2000年第3期
12钟兢军,王会社,王仲奇.多极压气机中可控扩散叶型研究的进展与展望第一部分可控扩散叶型的设计与发展.航空动力学报,Vol.16,No.3,2001,16(3):205-211
13 Sasha M. Savic, Marco A. Micheli, Andreas C. Bauer, Redesign of a multistage axial compressor for a heavy duty industrial gas turbine (GT11NMC) , ASME Paper, GT2005-68315, 2005
14 Rechter, H., Steinert, W., and Lehmann. Comparison of Controlled DiffusionAirfoils with Conventional NACA 65 Airfoils Developed for Stator Blade Application in a Multistage Axial Compressor, ASME Paper No. 84-GT-246, June 1984
15 NerurkarA C et al .Design study of turbomachine blades by optimization and inverse techniques AIAA-96-2555, Lake BuenaVista ,FL: AIAA,1996
16 Shahpar, S., Lapworth, L., Depablos, T., and Taylor, M.. A linear approach to the Multiparameter Design of Three-Dimensional Turbomachinery Blades.AIAA 99-0363, 37th AIAA Aerospace Sciences Meeting and Exhibit, Reno,NV
17 June Chung, Ki D.Lee. Shape Optimization of Transonic Compressor Blades Using Quasi-3D Flow Physics. ASME 2000-GT-489
18周正贵.压气机/风扇叶片自动优化设计的研究现状和关键技术.航空学报. 2008年3月第29卷第2期.Vol129 No12
19汪光文,周正贵,胡骏.基于优化算法的压气机叶片气动设计.航空动力学报. 2008年7月第23卷第7期
20 S. Goel, J.I. Cofer IV and H. Singh(1996). Turbine airfoil design optimization. ASME Paper 96-GT-158
21 S. Pierret and R.A. Van den Braembussche. Turbomachinery Blade Design Using a Navier-Stokes Solver and Aritificial Neural Network. ASME Paper 98-GT-4
22 JUNE CHUNG,LEE KI D. Shape optimization of transonic compressor blades using quasi - 3D flow physics[R].ASME Paper ,2000 - GT- 489
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24 SHAHROKH SHAHPAR DR. A comparative study of optimization methods for aerodynamic of turbomachienry blades[ R] . ASME Paper ,2000 - GT - 523
25 AHN CHAN SOL ,KIM KWANG YONG. Aerodynamic design optimization of an axial flow compressor rotor [ R] . ASME Paper ,2002 - GT -30445
26 Lee, S. Y. and Kim, K.Y. Design Optimization of Axial Flow Compressor Blades with Three Dimensional Navier Stokes Solver. ASME Paper 2000-GT-0488
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2陈云永,刘波,马聪慧,高丽敏,石建成,温泉.全三维、多叶排内外涵风扇压气机叶型优化研究.航空动力学报. 2008年3月第23卷第3期
3 Ku¨sters, B., Schreiber, H. A., Ko¨ller, U., and Mo¨nig, R., 1999,‘‘Development of Advanced Compressor Airfoils for Heavy-Duty Gas Turbines: Part II—Experimental and Theoretical Analysis,’’ASME J. Turbomach., 122, this issue, pp. 406–415
4 Wu.C.H. A General Theory of ThreeDimensional Flow in Subsonic and Supersonic Turbomachine in Radial, Axial and Mixed Flow Types. 1952,NASA TN-2604
5 Denton, J.D. A Time Marching Method for Two-and Three-dimensional Blade to Blade Row. 1974, ARC R&M 3775
6 Dawes.W.N. A Numerical Analysis of the Three-Dimensional Viscous Flow in a Transonic Compressor Rotor and Comparison With Experiment. 1986,ASME 86-GT-16
7 Hah,C., J.D. A Numerical Modeling of Endwall and Tip Clearance Flow of an Isolated Compressor Rotor. ASME Paper 85-GT-116, 1985
8 Arnone.Viscous Analysis of Three Dimensional Rotor Flow Using a Multigrid Method. Journal of Turbomahcinery. 1994, 116(2): 435-445
9 L.He, W.Ning. Efficient Approach for Analysis of Unsteady Viscous Flows in Turbomachines. AIAA Journal. 1998, 36(11):2005-2012
10 Denton J D, Dawes, W N. Computational fluid dynamics for turbomachinery design[c]//proceeding of instin Mech Engrs,1999,213(part c):107-124
11李志广.发展我国重型燃气轮机产业的可行技术途径.航空发动机. 2000年第3期
12钟兢军,王会社,王仲奇.多极压气机中可控扩散叶型研究的进展与展望第一部分可控扩散叶型的设计与发展.航空动力学报,Vol.16,No.3,2001,16(3):205-211
13 Sasha M. Savic, Marco A. Micheli, Andreas C. Bauer, Redesign of a multistage axial compressor for a heavy duty industrial gas turbine (GT11NMC) , ASME Paper, GT2005-68315, 2005
14 Rechter, H., Steinert, W., and Lehmann. Comparison of Controlled DiffusionAirfoils with Conventional NACA 65 Airfoils Developed for Stator Blade Application in a Multistage Axial Compressor, ASME Paper No. 84-GT-246, June 1984
15 NerurkarA C et al .Design study of turbomachine blades by optimization and inverse techniques AIAA-96-2555, Lake BuenaVista ,FL: AIAA,1996
16 Shahpar, S., Lapworth, L., Depablos, T., and Taylor, M.. A linear approach to the Multiparameter Design of Three-Dimensional Turbomachinery Blades.AIAA 99-0363, 37th AIAA Aerospace Sciences Meeting and Exhibit, Reno,NV
17 June Chung, Ki D.Lee. Shape Optimization of Transonic Compressor Blades Using Quasi-3D Flow Physics. ASME 2000-GT-489
18周正贵.压气机/风扇叶片自动优化设计的研究现状和关键技术.航空学报. 2008年3月第29卷第2期.Vol129 No12
19汪光文,周正贵,胡骏.基于优化算法的压气机叶片气动设计.航空动力学报. 2008年7月第23卷第7期
20 S. Goel, J.I. Cofer IV and H. Singh(1996). Turbine airfoil design optimization. ASME Paper 96-GT-158
21 S. Pierret and R.A. Van den Braembussche. Turbomachinery Blade Design Using a Navier-Stokes Solver and Aritificial Neural Network. ASME Paper 98-GT-4
22 JUNE CHUNG,LEE KI D. Shape optimization of transonic compressor blades using quasi - 3D flow physics[R].ASME Paper ,2000 - GT- 489
23 ERNESTO BENINI. Three-dimensional multi2objective design optimization of a transonic compressor rotor [J ] . Journal of Propulsion and Power ,2004 ,20 (3) :559 - 565
24 SHAHROKH SHAHPAR DR. A comparative study of optimization methods for aerodynamic of turbomachienry blades[ R] . ASME Paper ,2000 - GT - 523
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