直升机涡环状态边界风洞试验研究

王畅; 马帅; 黄志银; 王浩文; 黄志远; 邓皓轩

doi:10.11729/syltlx20220055

直升机涡环状态边界风洞试验研究

王畅^{1, 2,},
马帅²,
黄志银²,
王浩文¹,
黄志远^2, ,,
邓皓轩²

1.
清华大学航天航空学院，北京　100084
2.
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

基金项目: 国家自然科学基金面上项目（11672323）

详细信息

作者简介:
王畅: （1985—），男，河北邢台人，博士研究生。研究方向：直升机空气动力学。通信地址：北京市海淀区清华大学蒙民伟科技大楼北楼航天航空学院N103（100008）。E-mail：chang-wa18@mails.tsinghua.edu.cn

通讯作者:
黄志远: E-mail：doheheyzh@sina.com

中图分类号: V212.4
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出版历程
- 收稿日期: 2022-06-26
- 修回日期: 2022-08-25
- 录用日期: 2022-10-12
- 网络出版日期: 2023-06-14
- 刊出日期: 2023-10-29

A wind tunnel investigation of the helicopter vortex ring state boundary

1.
School of Aerospace Engineering, Tsinghua University, Beijing　100008, China
2.
Low Speed Aerodynamics Institute, China Aerodynamics Research and Development Center, Mianyang　621000, China

摘要

摘要: 本文对直升机涡环状态边界进行了系统的分析与研究。首先，剖析了涡环状态事故的成因，阐述了其在飞行特性、旋翼性能、桨盘入流、涡系结构等方面的物理机制，指出涡环状态下安全隐患的主要诱因是桨尖涡受挤压形成集中涡，使桨盘面上诱导入流相对垂向来流占优，造成旋翼拉力负阻尼与性能损失，导致浮沉运动失稳。然后，对比了各类涡环状态边界的差异性和适用性，指出现有边界预测模型存在建模方式主观性强和试验数据离散度高的问题，并提出了改进思路。最后，设计并开展了模拟下降飞行的旋翼风洞试验。试验结果显示：涡环状态下出现了旋翼拉力负阻尼、拉力损失和功率沉陷现象，旋翼拉力损失最高达30%，旋翼产生维持机体重量拉力的需用功率约为悬停功率的 160%。以特情防范实践中关注的旋翼拉力负阻尼和拉力性能损失为指标，从试验结果中提取了涡环状态边界临界速度离散点。在涡环状态边界预测模型构建中区分水平来流、垂向来流和诱导入流对桨尖涡驱动作用的强弱，并计入不同前进比下动量理论的修正和桨尖涡运动阈值的差异，基于试验值采用最小二乘法确定了模型参数，建立了半经验化的涡环状态边界预测模型，模型预测结果与风洞试验结果吻合较好，且符合飞行试验规律。本文对认识涡环状态特情和预防涡环状态事故具有现实意义。
- 涡环状态 /
- 直升机 /
- 飞行特情 /
- 风洞实验 /
- 桨尖涡
Abstract: The helicopter vortex ring state boundary is systematically analyzed and studied in this paper. Firstly, the causes of vortex ring state accidents are analyzed, and the physical mechanisms in flight characteristics, rotor performance, rotor inflow, and vortex structure are expounded. The formation of concentrated vortex causes the induced inflow to dominate the vertical inflow on the rotor disk, causing negative damping of rotor thrust and loss of rotor performance, and as a result the instability of subsidence motion. Then, the differences and applicability of various vortex ring state boundaries are compared, the problems of the existing boundary prediction models with strong subjectivity in modeling methods and high dispersion of test data are concluded, and improvement ideas are proposed. Finally, on the basis of the above knowledge, a rotor wind tunnel test to simulate the descending flight was designed and carried out. The test results show that the rotor thrust negative damping, thrust loss and power subsidence phenomenon are presented in the vortex ring state, the rotor thrust loss is up to 30%, and the required power is about 160% of the hovering power when the rotor generates the same thrust as hover; using the negative damping of rotor thrust and the loss of thrust performance, which are concerned in the practice of flight emergency, as the defining index, the discrete points of the critical velocity at vortex ring state boundary are extracted from the test results; in the construction of the vortex ring state boundary model, the strength of the horizontal inflow, the vertical inflow and the induced inflow on the rotor tip vortex is distinguished, and the correction of the momentum theory under different advance ratios and the difference of the rotor tip vortex motion threshold are taken into account. On the basis of the model parameters determined by the least squares method based on the test values, a semi-empirical vortex ring state boundary prediction model is established, and the model is in good agreement with the wind tunnel test results and in line with the trend of flight test results.
- vortex ring state /
- helicopter /
- flight emergency /
- wind tunnel experiment /
- blade tip vortex

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图 1 3种涡环状态边界对比

Fig. 1 Comparison of three kinds of vortex ring state boundaries commonly used in Engineering

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图 2 CEV增大下降速度进入涡环状态的试飞方式^[18]

Fig. 2 Flight test mode with increasing descent speed to enter vortex ring state applied by CEV^[18]

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图 3 穿越涡环状态过程中总距、前飞速度、下降速度的时间历程^[18]

Fig. 3 Time history of collective pitch, forward flight speed and descent speed during vortex ring state^[18]

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图 4 CEV降低前飞速度进入涡环状态的试飞方式^[18]

Fig. 4 Flight test mode with reducing forward speed to enter vortex ring state applied by CEV^[18]

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图 5 平直前飞与涡环状态下操纵与加速度时间历程对比^[18]

Fig. 5 Comparison of control and acceleration time history when entering vortex ring state^[18]

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图 6 涡环状态下直升机对总距提升的响应^[18]

Fig. 6 Response of helicopter to collective pitch increase in vortex ring state

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图 7 垂直下降状态旋翼拉力和扭矩随等效下降速度变化（总距恒定）^[44]

Fig. 7 Variation of rotor thrust and torque with equivalent descent speed under vertical descent (constant collective pitch control)^[44]

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图 8 垂直下降阶段旋翼总距和功率随下降速度变化（拉力恒定）^[45]

Fig. 8 Variation of rotor collective pitch and required power with descent rate under vertical descent (constant thrust)^[45]

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图 9 旋翼在涡环状态的入流模型^[19]

Fig. 9 Inflow model of rotor in vortex ring state^[19]

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图 10 垂直下降状态旋翼剖面速度矢量图^[52]

Fig. 10 velocity vector diagram in vertical descent state^[52]

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图 11 下降飞行过程中桨尖涡的结构演化^[53-54]

Fig. 11 Structural evolution of blade tip vortex during descending flight^[53-54]

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图 12 Wolkovitch对桨尖涡运动速度的假设^[55]

Fig. 12 Wolkovitch's hypothesis on the velocity of blade tip vortex^[55]

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图 13 Wolkovitch与Dress 边界对比^[55-56]

Fig. 13 Comparison of Wolkovitch and Dress boundaries^[55-56]

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图 14 Peters对桨盘处流动的假设^[59]

Fig. 14 Peters′ assumption of flow at the disc^[59]

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图 15 Peters涡环状态边界^[59]

Fig. 15 Peters vortex ring state boundary^[59]

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图 16 Newman给出的涡环状态边界^[61]

Fig. 16 State boundary of vortex ring given by Newman^[61]

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图 17 高−辛涡环状态边界^[14]

Fig. 17 Gao−Xin vortex ring state boundary^[14]

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图 18 ONERA涡环状态边界^[18]

Fig. 18 ONERA vortex ring state boundary^[18]

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图 19 NASA涡环状态边界^[19]

Fig. 19 NASA vortex ring state boundary^[19]

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图 20 风洞试验照片

Fig. 20 Wind tunnel test photo

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图 21 试验风速与下滑角示意图

Fig. 21 Schematic diagram of glide angle between test wind speed and rotor

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图 22 旋翼在垂直下降状态的气动特性

Fig. 22 Aerodynamic characteristics of each rotor in vertical descent

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图 23 1号旋翼在不同下滑角下的气动特性

Fig. 23 Aerodynamic characteristics of No. 1 rotor at different glide angles

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图 24 2号旋翼在40°下滑角下的拉力特性

Fig. 24 Thrust characteristics of rotor 2 at 40 ° glide angle

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图 25 从风洞试验数据提取的涡环状态边界

Fig. 25 Vortex ring state boundary extracted from wind tunnel test data

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图 26 式（14）计算的涡环状态边界与风洞试验和CEV飞行试验结果对比

Fig. 26 Comparison of vortex ring state boundary calculated by equation 14 with wind tunnel test results and CEV flight test results

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表 1 国内外已开展的涡环状态旋翼性能测量试验

Table 1 Tests of rotor performance in vortex ring state

研究者	时间	参考文献	桨叶片数	旋翼半径 /mm	旋翼转速 /(r﹒min⁻¹)	扭转角/(°)	下降姿态	试验环境
Castles、Gray	1951	[45]	3	610、914	1200、1600	0、−12	垂直下降	直径2.74 m风洞
Mort、Yaggy	1963	[46]	3	1448、1829	700～1410、 700～1100	−22.4、 −46.6	斜下降、垂直下降	NFAC
Washizu、Azuma、Koo等	1966	[47]	3	549	1000	−8.33	斜下降、垂直下降	滑轨
Azuma、Obata	1968	[44]	3	550	1000	−8，0	垂直下降	直径3 m风洞
Empey、Ormiston	1974	[48]	2	162	13250	0	斜下降、垂直下降	2.13 m × 3 m风洞
辛宏、高正	1993-1996	[14-16]	2	549	1406	0、 −5.5、−9.22	斜下降、垂直下降	悬臂机
Betzina	2001	[49]	3	610	1800	−41	斜下降、垂直下降	NFAC

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表 2 旋翼模型参数

Table 2 Parameters of rotor models

编号	半径/mm	弦长/mm	扭转角/(°)	桨尖形状	翼型厚度	转速/(r﹒min⁻¹)
1	550	135	0	抛物线后掠	12%	3000
2	550	135	−8	抛物线后掠	11.3%	3000
3	550	135	−8	抛物线后掠	12%	2600
4	450	97	−6.7	矩形	14%	3700

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