2025年5月16日 22:42 星期五
首页 > 过刊浏览>2021年第53卷第5期 >801-812. DOI:10.16356/j.1005-2615.2021.05.018
PDF HTML阅读 XML下载 导出引用 引用提醒
旋翼翼型动态失速等离子体流动控制试验研究
CSTR:
[cstr]
作者:
作者单位:

1.中国空气动力研究与发展中心低速空气动力研究所，绵阳621000;2.中国空气动力研究与发展中心旋翼空气动力学重点实验室，绵阳 621000

通讯作者:

杨永东，男，研究员，E-mail： 1141534997@qq.com。

中图分类号:

V211.7


Experimental Research on Dynamic Stalled Plasma Flow Control of Rotor Airfoil
Author:
Affiliation:

1.Low Speed Aerodynamics Institute， China Aerodynamics Research & Development Center， Mianyang 621000， China;2.Key Laboratory of Rotor Aerodynamics， China Aerodynamics Research & Development Center， Mianyang 621000， China

  • 摘要
    • |
  • 图/表
    • |
  • 访问统计
    • |
  • 参考文献
    • |
  • 相似文献
    • |
  • 引证文献
    • | |
  • 文章评论
    摘要:

    翼型动态失速是指机翼或叶片的当地迎角呈现周期或急剧变化时绕流附面层大范围分离带来的一种强烈的非线性、非定常流动现象。动态失速涡脱离翼型后缘流向下游时，会引发升力急剧下降、阻力迅速增大的失速和颤振问题。基于旋翼翼型两自由度动态试验装置和高频高速振荡试验装置，以典型旋翼翼型为研究对象，利用纳秒脉冲激励电源和介质阻挡放电等离子体激励器，在FL-11风洞和FL-20风洞开展了翼型动态失速等离子体流动控制试验研究，试验最高雷诺数突破1.7×106，模型最高振荡频率突破10 Hz。试验结果表明，等离子体气动激励能够有效控制翼型动态失速，改善平均气动力，减小俯仰力矩负峰值，减小气动力/力矩随迎角变化的迟滞区域。

    Abstract:

    Airfoil dynamic stall refers to a strong nonlinear and unsteady flow phenomenon caused by the large-scale separation of the surrounding surface layer when the local angle of attack of the wing or blade changes periodically or sharply.The dynamic stall vortex breaks away from the airfoil. When the trailing edge flows downstream， it will cause stall and flutter problems with a sharp drop in lift and a rapid increase in drag. Based on the two-degree-of-freedom dynamic test device of the rotor airfoil and the high-frequency and high-speed oscillation test device， the typical rotor airfoil is used as the research object， the experimental research on dynamic stall plasma flow control of airfoil is carried out in FL-11 and FL-20 wind tunnels by using nanosecond pulse excitation power supply and dielectric barrier discharge plasma exciter. The maximum Reynolds number in the test exceeds 1.7×106， and the maximum oscillation frequency of the model exceeds 10 Hz. The test results show that the plasma aerodynamic excitation can effectively control the dynamic stall of the airfoil， improve the average aerodynamic force， reduce the negative peak of the pitching moment and the hysteresis area of the aerodynamic force/torque with the angle of attack.

    图1 低速风洞旋翼翼型动态试验模型三维结构图Fig.1 3-D structure diagram of rotor airfoil dynamic test model in low speed wind tunnel
    图2 低速风洞翼型模型测压孔分布图Fig.2 Pressure tap distribution of airfoil model in low-speed wind tunnel
    图3 高速风洞旋翼翼型动态试验模型结构图Fig.3 Structure diagram of rotor airfoil dynamic test model in high speed wind tunnel
    图4 高速风洞翼型模型测压孔分布图Fig.4 Pressure tap distribution of airfoil model in high-speed wind tunnel
    图5 FL-11两自由度低速风洞动态试验装置Fig.5 FL-11 two-degree-of-freedom dynamic test device in low-speed wind tunnel
    图6 FL-20风洞旋翼翼型动态试验装置Fig.6 FL-20 wind tunnel rotor airfoil dynamic test device
    图7 8510B动态压力传感器外形及尺寸图Fig.7 Shape and size of dynamic pressure sensor 8510B
    图8 XCE-062传感器外形及尺寸Fig.8 Shape and size of sensor XCE-062
    图9 典型介质阻挡放电等离子体激励器布局图Fig.9 Layout of typical dielectric barrier discharge plasma exciter
    图10 纳秒脉冲电源Fig.10 Nanosecond pulse power
    图11 不同雷诺数静态气动力曲线Fig.11 Static aerodynamic curves at different Reynolds numbers
    图12 试验翼型动态气动力与静态气动力对比Fig.12 Comparison of dynamic and static aerodynamic forces of test airfoil
    图13 不同风速下耦合振荡试验结果对比Fig.13 Comparison of coupled oscillation test results with different wind speeds
    图14 不同沉浮振幅下耦合振荡试验结果对比Fig.14 Comparison of coupled oscillation test results with different ups and downs
    图15 不同激励电压下耦合振荡试验结果对比Fig.15 Comparison of coupled oscillation test results with different excitation voltages
    图16 不同激励频率下耦合振荡试验结果对比Fig.16 Comparison of coupled oscillation test results with different excitation frequencies
    图17 不同马赫数下俯仰振荡试验结果对比Fig.17 Comparison of pitch oscillation test results with different Ma
    图18 不同激励电压下俯仰振荡试验结果对比Fig.18 Comparison of pitch oscillation test results with different excitation voltages
    图19 俯仰振荡试验不同激励频率结果对比Fig.19 Comparison of pitch oscillation test results with different excitation frequencies
    图20 下行程18 °时，等离子体控制前后翼型表面压力系数分布(α0=10.44 °，α1=6.94 °，f=0.499 Hz，V=10 m/s)Fig.20 Comparison of pitch oscillation test results with different excitation frequencies (α0=10.44 °，α1=6.94 °，f=0.499 Hz，V=10 m/s)
    参考文献
    相似文献
    引证文献
引用本文

吴霖鑫,李国强,杨永东,张鑫,陈磊,赵光银.旋翼翼型动态失速等离子体流动控制试验研究[J].南京航空航天大学学报,2021,53(5):801-812

复制
分享
文章指标
  • 点击次数:
  • 下载次数:
  • HTML阅读次数:
  • 引用次数:
历史
  • 收稿日期:2020-10-07
  • 最后修改日期:2021-06-09
  • 在线发布日期: 2021-10-05
文章二维码
您是第位访问者
网站版权 © 南京航空航天大学学报
技术支持：北京勤云科技发展有限公司
请使用 Firefox、Chrome、IE10、IE11、360极速模式、搜狗极速模式、QQ极速模式等浏览器，其他浏览器不建议使用!