切向孔隙流动对大折转角压气机叶栅气动性能的影响
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摘要
现代先进航空发动机高效、高负荷设计指标对其主要气动部件—风扇/压气机的效率和负荷水平提出了越来越高的要求,而叶片负荷的大幅度增加将使得叶栅流道内部横向压力梯度和二次流动加强,附面层内的低能流体在强逆压力梯度下势必造成大尺度非定常流动分离,叶栅气动损失迅速增加,从而限制了叶片负荷水平的进一步提高,压气机失速裕度和效率急剧下降,工作稳定性无法得到保证。因此,在深入研究压气机/风扇叶栅流道、尤其是端部角区内复杂流场结构和损失产生机理的基础上,探索降低损失,特别是端部损失的方法和途径,开发利用弯曲叶片、孔隙流动等控制附面层流动分离的综合流动控制技术是改善压气机气动性能的关键。
本论文对采用切向孔隙流动控制技术的大折转角直、弯扩压叶栅气动性能进行了数值研究,探索孔隙流动控制技术在降低大折转角扩压叶栅流动损失的流体动力学机理。切向孔隙流动控制技术是指在叶栅吸、压力面之间设计合适的切向削涡孔或缝,利用叶栅吸、压力面压差产生的射流来增加吸力面分离区内低能流体的动能及湍流度,使得分离区内的流体能够进一步克服强逆压力梯度而避免或推迟大尺度分离,从而顺利实现整个叶栅流道的扩压流动。文中首先数值模拟了不同孔径、不同轴向及径向位置的单孔型削涡孔对直叶栅气动性能的影响,然后开展了具有多孔组合型式的直叶栅气动性能及流动机理研究,并提出削涡缝的设计思想,在上述研究基础上,深入探索了削涡缝流动控制技术改善直、弯曲扩压叶栅流动特性的机理。
数值研究结果表明,单孔型削涡孔的孔径及其径向、轴向位置对扩压叶栅气动性能有较大影响,存在着最佳孔径及径向、轴向位置。当这三个主要设计参数处于最佳匹配时,即孔径D=2~4mm,径向位置为10%~15%叶高,轴向倾斜角为30°,吸力面处位置为70%轴向弦长时,通流能力最大可增加1.4%,出口总压损失下降17%,叶栅气动负荷及扩压段长度也有所增加,流动分离显著减弱。采用多孔组合设计时叶栅流场特性强烈的依赖于孔径大小,孔径较小如D=2mm时,不同的多孔组合型式下栅内气动性能改善的程度差别较大,且组合孔的数目越多叶栅通流能力越强;大孔径如D=4mm时,任意的组合方式均能有效降低叶栅出口总压损失,组合孔越多效果越明显。组合式削涡孔能使叶栅0~15%H叶展叶栅出口扩压因子小于0.6,有效地改善吸力面/端区的气动性能,而25%H~50%H叶展的负荷增
The trend to reduce compressor size and weight by reducing the number of stages leads to higher amounts of diffusion per stage, which will cause large endwall loss in the compressor, or even largely reduce compressor efficiency and surge margin. Since the improvement in the aerodynamic performance of aero-engines depends on further increases in efficiency of each component, especially compressor and turbine, many researchers are addressing this issue through the use of varying flow control techniques, such as compound lean blade or air injection.
In this paper, an extensive numerical study has been performed to investigate the effects of air injection on the performance of the straight blade compressor cascade and compound lean compressor cascade as well. Air injection was implemented via the hole/slot between the pressure and suction surfaces at appropriate locations. Injected air due to the pressure difference between the pressure and suction surfaces was used to energize the low energy fluid within the endwall/suction corner to increase its ability to overcome the adverse pressure gradient so as to avoid or delay flow separation. First, single-hole configurations with different hole diameters, at different axial or radial locations were simulated in the straight blade cascade. Second, the mechanism of multi-hole configurations were studied in the straight blade cascade, thus the concept of injection slot was developed. Third, the mechanism of injection slot on the performance improvement of straight blade cascade and compound lean blade cascade were discussed in detail.
The numerical results show that the aerodynamic performance of compressor cascade changes significantly with the variation of the single-hole diameter, its axial and radial locations. The optimum hole diameter, or the optimum axial or radial hole location, therefore exists. When the three parameters are in a good match, for example, the hole with a diameter of 2~4mm placed at 70% of the axial chord between 10% and 15% of the blade height, the capacity of through-flow is increased 1.4%, and the outlet total pressure loss is reduced 17%. The blade loading and the axial diffusion range are also increased
本论文对采用切向孔隙流动控制技术的大折转角直、弯扩压叶栅气动性能进行了数值研究,探索孔隙流动控制技术在降低大折转角扩压叶栅流动损失的流体动力学机理。切向孔隙流动控制技术是指在叶栅吸、压力面之间设计合适的切向削涡孔或缝,利用叶栅吸、压力面压差产生的射流来增加吸力面分离区内低能流体的动能及湍流度,使得分离区内的流体能够进一步克服强逆压力梯度而避免或推迟大尺度分离,从而顺利实现整个叶栅流道的扩压流动。文中首先数值模拟了不同孔径、不同轴向及径向位置的单孔型削涡孔对直叶栅气动性能的影响,然后开展了具有多孔组合型式的直叶栅气动性能及流动机理研究,并提出削涡缝的设计思想,在上述研究基础上,深入探索了削涡缝流动控制技术改善直、弯曲扩压叶栅流动特性的机理。
数值研究结果表明,单孔型削涡孔的孔径及其径向、轴向位置对扩压叶栅气动性能有较大影响,存在着最佳孔径及径向、轴向位置。当这三个主要设计参数处于最佳匹配时,即孔径D=2~4mm,径向位置为10%~15%叶高,轴向倾斜角为30°,吸力面处位置为70%轴向弦长时,通流能力最大可增加1.4%,出口总压损失下降17%,叶栅气动负荷及扩压段长度也有所增加,流动分离显著减弱。采用多孔组合设计时叶栅流场特性强烈的依赖于孔径大小,孔径较小如D=2mm时,不同的多孔组合型式下栅内气动性能改善的程度差别较大,且组合孔的数目越多叶栅通流能力越强;大孔径如D=4mm时,任意的组合方式均能有效降低叶栅出口总压损失,组合孔越多效果越明显。组合式削涡孔能使叶栅0~15%H叶展叶栅出口扩压因子小于0.6,有效地改善吸力面/端区的气动性能,而25%H~50%H叶展的负荷增
The trend to reduce compressor size and weight by reducing the number of stages leads to higher amounts of diffusion per stage, which will cause large endwall loss in the compressor, or even largely reduce compressor efficiency and surge margin. Since the improvement in the aerodynamic performance of aero-engines depends on further increases in efficiency of each component, especially compressor and turbine, many researchers are addressing this issue through the use of varying flow control techniques, such as compound lean blade or air injection.
In this paper, an extensive numerical study has been performed to investigate the effects of air injection on the performance of the straight blade compressor cascade and compound lean compressor cascade as well. Air injection was implemented via the hole/slot between the pressure and suction surfaces at appropriate locations. Injected air due to the pressure difference between the pressure and suction surfaces was used to energize the low energy fluid within the endwall/suction corner to increase its ability to overcome the adverse pressure gradient so as to avoid or delay flow separation. First, single-hole configurations with different hole diameters, at different axial or radial locations were simulated in the straight blade cascade. Second, the mechanism of multi-hole configurations were studied in the straight blade cascade, thus the concept of injection slot was developed. Third, the mechanism of injection slot on the performance improvement of straight blade cascade and compound lean blade cascade were discussed in detail.
The numerical results show that the aerodynamic performance of compressor cascade changes significantly with the variation of the single-hole diameter, its axial and radial locations. The optimum hole diameter, or the optimum axial or radial hole location, therefore exists. When the three parameters are in a good match, for example, the hole with a diameter of 2~4mm placed at 70% of the axial chord between 10% and 15% of the blade height, the capacity of through-flow is increased 1.4%, and the outlet total pressure loss is reduced 17%. The blade loading and the axial diffusion range are also increased
引文
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95 G. R. Gosterlow. Cascade Aerodynamics. Pergamon Press, 1984:153-315
96 L. Reid, R. D. Moore. Experimental Study of Low Aspect Ratio Compressor Blading. ASME Paper. 1980, 80-GT-16
97 A. J. Wennerstrom. Highly Loaded Axial Flow Compressor: History and Current Development. Trans. of the ASME, J. of Turbomachinery. 1990, Vol.112,pp.567-578
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86 E.Dennis, M.Culley,M.Bright. Activective Flow Separation Control of a Stator Vane Using Surface Injection in a Multistage Compressor Experiment. ASME Paper. 2003, 2003-GT-38863
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88 K.R. Kirtley, P. Graziosi, P. Wood, B. Beacher, H.-W. Shin. Design and Test of an Ultra-Low Solidity Flow Controlled Compressor Stator. ASME Paper. 2004, 2004-GT-53012
89 Bhaskar Roy, Manish Chouhan, Kota Venkata Kaundinya. Experimental Study of Boundary Layer Control through Tip Injection on Straight and Swept Compressor Blades. ASME Paper. 2005, 2005-GT-68304
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95 G. R. Gosterlow. Cascade Aerodynamics. Pergamon Press, 1984:153-315
96 L. Reid, R. D. Moore. Experimental Study of Low Aspect Ratio Compressor Blading. ASME Paper. 1980, 80-GT-16
97 A. J. Wennerstrom. Highly Loaded Axial Flow Compressor: History and Current Development. Trans. of the ASME, J. of Turbomachinery. 1990, Vol.112,pp.567-578
98 轴流压气机气动设计.国防工业出版社,1975
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