级环境下静叶开槽抑制附面层分离的机理研究

王浩; 张皓光; 荆风玉; 冯奕鸣; 肖劲航

doi:10.13224/j.cnki.jasp.20230520

级环境下静叶开槽抑制附面层分离的机理研究

doi: 10.13224/j.cnki.jasp.20230520

西北工业大学动力与能源学院，西安 710072

基金项目: 国家自然科学基金（51006084）；国家自然科学基金重点项目（51536006）；国家科技重大专项（2017-Ⅱ-0005-0018）

详细信息

作者简介:
王浩（1999-），男，硕士生，研究领域为叶轮机械气动热力学。E-mail：hao1999@mail.nwpu.edu.cn

通讯作者:
张皓光（1981-），男，副教授、博士生导师，博士，研究领域为叶轮机械气动热力学。E-mail：zhg@nwpu.edu.cn

中图分类号: V231.3
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出版历程
- 收稿日期: 2023-08-10
- 网络出版日期: 2025-02-14

Study on mechanism of slotted stator blade inhibiting boundary layer separation in a single-stage axial flow compressor

School of Power and Energy，Northwestern Polytechnical University，Xi’an 710072，China

摘要

摘要:
对跨声速单级轴流压气机NASA Stage 35开展了静叶开槽控制附面层分离的机理研究，基于3个前槽出口位置和两个后槽出口位置，组合设计了6种双槽方案。数值结果表明：在中小质量流量工况下，6种开槽方案均提高了压气机总性能，前槽出口位于13%C_a（C_a表示叶顶轴向弦长）、后槽出口位于60%C_a处的研究方案对压气机总性能的改善效果最好，在没有明显降低设计点效率的前提下，将近失速工况压气机效率的绝对量提升了1.3%。流场分析发现：在99%叶高处，较高的射流动量和更靠近分离涡核心的射流方向能更好的消除由附面层分离产生的低能流体。但低叶高未开槽范围的叶片表面附面层厚度有所增加，这是因为在径向压力梯度和原有分离区逆向流动的双重作用下，近槽道下壁面处速度较低的射流向叶根区域和叶片前缘迁移所导致的。
- 单级轴流压气机 /
- 静叶开槽 /
- 槽道射流 /
- 附面层分离 /
- 流动机理
Abstract:
The mechanism of slotted stator blade controlling boundary layer separations was studied, and the transonic single-stage axial flow compressor NASA Stage 35 was selected as the research object. Based on three front slot outlet positions and two rear slot outlet positions, six double-slot schemes were designed. The numerical calculation results indicated that under small and medium mass flow rate conditions, all the schemes improved the compressor total performance. When the outlets of front slot and rear slot were located at 13%C_a (C_a was the blade tip axial chord length) and 60%C_a, the compressor total performance was the best. And without significantly reducing the design point compressor efficiency, the scheme exhibited the highest compressor efficiency improvement of 1.3% under the near stall condition. The internal flow field analysis showed that at 99% blade span, higher jet momentum and the jet direction closer to the separation vortex core better eliminated the low-energy fluids generated by the boundary layer separation. However, under the interaction of radial pressure gradient and reverse flow in the separation zone, the jet with lower velocity near the slot undersurface migrated to lower blade span and the leading edge of the blade, resulting in an increase in the thickness of the boundary layer at the lower blade span.
- single-stage axial flow compressor /
- slotted stator blade /
- slot jet /
- boundary layer separation /
- flow mechanism

HTML

图 1 NASA Stage 35几何示意图

Figure 1. Geometric model of NASA Stage 35

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图 2 计算域网格示意图

Figure 2. Calculation domain gird

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图 3 不同网格数目的压气机总性能对比

Figure 3. Compressor total performance comparison for different grid numbers

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图 4 不同湍流模型下的压气机总性能对比

Figure 4. Compressor total performance comparison for different turbulence models

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图 5 数值结果与实验数据对比

Figure 5. Comparison of numerical results with experimental data

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图 6 不同叶高范围的绝对马赫数分布

Figure 6. Distribution of absolute Mach number at different blade spans

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图 7 损失源分析区域划分示意图

Figure 7. Diagram of division domain of loss source analysis

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图 8 不同工况下损失源分析

Figure 8. Loss sources analysis at different operating points

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图 9 不同工况下各个损失源所占的比例（单位：%）

Figure 9. Proportion of individual loss sources at different operating points （unit：%）

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图 10 槽道位置及结构示意图

Figure 10. Diagram of slot position and construction

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图 11 压气机总性能对比曲线

Figure 11. Comparison curves of total compressor performance

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图 12 近失速工况下不同研究方案95%叶高截面的绝对速度矢量分布图

Figure 12. Contours of absolute velocity vector at 95% blade span for different schemes at near stall point

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图 13 槽道射流无量纲流量和轴向动量对比

Figure 13. Comparison of dimensionless jet mass flow rate and axial momentum

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图 14 不同研究方案40%、70%和90%叶高截面内绝对马赫数分布云图

Figure 14. Contours of absolute Mach number at 40%, 70% and 90% blade spans for different schemes

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图 15 近失速工况U13D60方案槽道射流流线分布

Figure 15. Slot-jet streamlines for U13D60 scheme at near stall point

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图 16 近失速工况原型静叶吸力面极限流线分布图

Figure 16. Limiting streamlines of stator blade suction surface for datum scheme at near stall point

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图 17 不同轴向截面内的总压损失系数分布云图

Figure 17. Contours of total pressure loss coefficient in different axial sections

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图 18 近失速工况不同研究方案损失源分析

Figure 18. Loss source analysis for different schemes at near stall point

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表 1 NASA Stage 35设计参数^[23-24]

Table 1. Design parameters of NASA Stage 35^[23-24]

参数	数值
设计质量流量/（kg/s）	20.19
设计总压比	1.82
设计绝热效率	0.83
转速/（r/min）	17188.8
叶尖速度/（m/s）	454.46
进口轮毂比	0.7
转子轮毂比	1.19
静子轮毂比	1.26
动叶数目	36
静叶数目	46

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表 2 网格点分布

Table 2. Grid point distribution

参数		数值
压气机通道网格数目/万		127
转子通道内网格数目/万		73
静子通道内网格数目/万		31
转子通道内的O型网格的网格数目	轴向	25
	径向	101
	环向	209
叶顶间隙内的O型网格的网格数目	轴向	9
	径向	25
	环向	209
静子通道内的O型网格的网格数目	轴向	25
	径向	57
	环向	201

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表 3 槽道几何参数

Table 3. Slot geometry parameters

参数	数值
参数	槽U			槽D
（进口宽度/C_a）/%	8			11
（出口宽度/C_a）/%	10			14
（进口位置/C_a）/%	6			12
（出口位置/C_a）/%	8	11	13	60	70
出口气流角度/（°）	60	52	47	24

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表 4 不同方案的${{\boldsymbol{\sigma}} ^*}$和${\boldsymbol{\psi }}$的数值

Table 4. Values of ${{\boldsymbol{\sigma}} ^*}$and ${\boldsymbol{\psi }}$ for different schemes

方案	${\sigma ^*}$	$\psi $/%
原型	0.93681
U8D60	0.94146	0.496
U8D70	0.93955	0.293
U10D60	0.94533	0.909
U10D70	0.94464	0.835
U13D60	0.94679	1.065
U13D70	0.94637	1.020

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