A general model for trailing edge serrations simulation on wind turbine airfoils

Trailing edge serrations(TESs)are capable of noticeably suppressing the turbulent trailing edge noise induced by rotating wind turbine blades and become an integral part of a blade.However,the challenges involved in the dimensional design of serration height 2 h,wavelengthλand flap angleϕare yet to be dealt with in a satisfactory manner.To address the problem,a general model for simulating the effects of serrations on the hydrodynamic and aeroacoustic performance is proposed due to its ease of use and relatively low requirements for user input.The solid serrations are replicated by momentum sources calculated by its aerodynamic forces.Then,a case relevant to wind turbine airfoil is examined,a hybrid improved delay detached eddy simulation(IDDES)method coupled with FW-H integration is deployed to obtain the flow features and far-field sound pressure level.It is found that the modeling method could reproduce the flow field and noise as serrated airfoil.

Airfoil Trailing Edge Noise Generation and Its Surface Pressure Fluctuation

In the present work, Large Eddy Simulation (LES) of turbulent flows over a NACA 0015 airfoil is performed. The purpose of such numerical study is to relate the aerodynamic surface pressure with the noise generation. The results from LES are validated against detailed surface pressure measurements where the time history pressure data are recorded by the surface pressure microphones. After the flow-field is stabilized, the generated noise from the airfoil Trailing Edge (TE) is predicted using the acoustic analogy solver, where the results from LES are the input. It is found that there is a strong relation between TE noise and the aerodynamic pressure. The results of power spectrum density show that the fluctuation of aerodynamic pressure is responsible for noise generation.

Design and test of pneumatic artificial muscle driven variable trailing-edge camber wing

A variable camber wing driven by pneumatic artificial muscles is developed in this paper. Firstly, the experimental setup to measure the static output force of pneumatic artificial muscle is designed and the relationship between the static output force and the air pressure is investigated. Experimental results show that the static output force of pneumatic artificial muscle decreases nonlinearly with the increase of contraction ratio. Secondly, the model of variable camber wing driven by pneumatic artificial muscles is manufactured to validate the variable camber concept. Finally, wind tunnel tests are conducted in the low speed wind tunnel. It is found that the wing camber increases with the increase of air pressure. When the air pressure of PAMs is 0.4 MPa and 0.5 MPa, the tip displacement of the trailing-edge is 3 mm and 5 mm, respectively. The lift of aerofoil with flexible trailing-edge increases by 87% at AOA of 5°.

Modeling and Experiment of a Morphing Wing Integrated with a Trailing Edge Control Actuation System

Morphing wing has attracted many research attention and effort in aircraft technology development because of its advantage in lift to draft ratio and flight performance.Morphing wing technology combines the lift and control surfaces into a seamless wing and integrates the primary structure together with the internal control system.It makes use of the wing aeroelastic deformation induced by the control surface to gain direct force control through desirable redistribution of aerodynamic forces.However some unknown mechanical parameters of the control system and complexity of the integrated structure become a main challenge for dynamic modeling of morphing wing.To solve the problem,a method of test data based modal sensitivity analysis is presented to improve the morphing wing FE model by evaluating the unknown parameters and identifying the modeling boundary conditions.An innovative seamless morphing wing with the structure integrated with a flexible trailing edge control system is presented for the investigation.An experimental model of actuation system driven by a servo motor for the morphing wing is designed and established.By performing a vibration test and the proposed modal sensitivity analysis,the unknown torsional stiffness of the servo motor and the boundary condition of the actuation mechanism model is identified and evaluated.Comparing with the test data,the average error of the first four modal frequency of the improved FE model is reduced significantly to less than 4%.To further investigate the morphing wing modeling,a wing box and then a whole morphing wing model including the skin and integrated with the trailing edge actuation system are established and tested.By using the proposed method,the FE model is improved by relaxing the constraint between the skin and actuation mechanism.The results show that the average error of the first three modal frequency of the improved FE model is reduced to less than 6%.The research results demonstrate that the presented seamless morphing wing integrated with a flexible trailing edge control surface can improve aerodynamic characteristics.By using the test data based modal sensitivity analysis method,the unknown parameter and boundary condition of the actuation model can be determined to improve the FE model.The problem in dynamic modeling of high accuracy for a morphing wing can be solved in an effective manner.

Performance Assessment of Darrieus Turbine with Modified Trailing Edge Airfoil for Low Wind Speeds

Darrieus wind turbines are simple lift based machines with exceptionally high efficiencies in terms of power coefficient compared to similar drag based vertical axis turbines. However, in low Reynolds numbers, a notable performance loss was reported. As a potential solution, truncated NACA 0018 airfoil (NACA 0018TC-39) has been introduced with baseline cavity modification to achieve better start-up characteristics and to enhance the low wind speed performance. The baseline cavity will provide an additional benefit of reverse drag at low TSR which is obligatory for low wind speed start-up. Numerical optimization has been carried out on the conceived airfoil NACA 0018TC-39 to find out the effective truncation percentage in terms of the chord. The numerical study has been extended to compare NACA 0018 and NACA 0018TC-39 airfoil for their aerodynamic performances in terms of lift, drag coefficients and separation characteristics. The NACA 0018TC-39 airfoil was incorporated within a non-swept straight bladed Darrieus turbine miniature to experimentally evaluate the performance in terms of dynamic power coefficient, dynamic torque coefficient and static torque coefficient and compared with conventional NACA 0018 airfoil at six different Reynolds numbers 178917, 193827, 208737, 223646, 238556 and 268376. The experimental contrast implied that NACA 0018TC-39 airfoil turbine yielded almost double power coefficients at low Reynolds number compared to conventional NACA 0018 airfoil without hampering its performance at higher Reynolds number.

TR-PIV Analysis of Turbulent Wake of Hydrofoil with Beveled Trailing Edge

The turbulent wakes behind trailing edge are analyzed for understanding of the flow mechanisms responsible for the generation of trailing edge noise. The TILS （turbulence integral length scale） of the turbulent wake of hydrofoil with blunt trailing edge is calculated from TR-PIV （time-resolved particle image velocimetry） data. The temporal auto-correlation method based on Taylor hypothesis and spatial correlation method are used to get the TILS information of the turbulent wake of hydrofoil, respectively The comparison of results by two methods indicates that the spatial correlation method is independent on Taylor hypothesis and suitable to strong turbulence and non-isotropic turbulence.

Aeroacoustic Investigation of Trailing-Edge Finlets

Wind tunnel testing and embedded large eddy simulations are employed to study the noise reduction of trailing-edge finlets on an airfoil. Trailing-edge finlets are shown to increase the distance between the highly energetic fluid particles and the sharp trailing edge. Experiments were conducted at different angles of attack. Wind tunnel measurements confirm that finlets reduce the broadband noise radiated by the airfoil. Results also reveal that the noise reduction of finlets is dependent on the airfoil angle of attack, and that the highest noise reduction is obtained at the largest angle of attack tested.

Seamless morphing trailing edge flaps for UAS-S45 using high-fidelity aerodynamic optimization

The seamless trailing edge morphing flap is investigated using a high-fidelity steady-state aerodynamic shape optimization to determine its optimum configuration for different flight conditions,including climb,cruise,and gliding descent.A comparative study is also conducted between a wing equipped with morphing flap and a wing with conventional hinged flap.The optimization is performed by specifying a certain objective function and the flight performance goal for each flight condition.Increasing the climb rate,extending the flight range and endurance in cruise,and decreasing the descend rate,are the flight performance goals covered in this study.Various optimum configurations were found for the morphing wing by determining the optimum morphing flap deflection for each flight condition,based on its objective function,each of which performed better than that of the baseline wing.It was shown that by using optimum configuration for the morphing wing in climb condition,the required power could be reduced by up to 3.8%and climb rate increases by 6.13%.The comparative study also revealed that the morphing wing enhances aerodynamic efficiency by up to 17.8%and extends the laminar flow.Finally,the optimum configuration for the gliding descent brought about a 43%reduction in the descent rate.

小尺寸切开式尾缘锯齿对翼型气动噪声的影响

研究小尺寸切开式尾缘锯齿对翼型气动噪声的影响,并分析齿高与尾缘边界层厚度之间的关系对降噪效果的影响.以NACA0012翼型为研究对象,采用混合计算气动声学方法进行仿真.在20 m/s风速下对NACA0012翼型进行噪声实验,验证仿真方法的准确性.通过控制齿高和齿宽研究不同形状尾缘锯齿分别在0°和10°攻角下对翼型气动噪声的影响.研究结果表明,尾缘锯齿的降噪效果与齿高和齿宽均成正比.0°攻角下最大降噪量为3.3 dB,10°攻角下最大降噪量为2.6 dB.切开式尾缘锯齿结构会增加流向噪声的高频分量,尤其在0°攻角下显著.相比嵌入式尾缘锯齿,切开式尾缘锯齿具有更小的极限降噪齿高.

吹风比和肋板对叶片尾缘开缝气膜冷却特性的影响

为了阐明吹风比和肋板对燃气透平叶片尾缘开缝区域气膜冷却性能的影响,采用延迟-分离涡模拟方法求解了尾缘开缝模型的流量系数、非定常流场结构和气膜冷却效率,采用实验数据考核了延迟-分离涡模拟方法对流量系数和气膜冷却效率预测的有效性,获得了使尾缘开缝壁面气膜冷却效率最佳吹风比。结果表明:流量系数随吹风比增加而增大,但吹风比大于0.65后,流量系数几乎不受吹风比影响;在吹风比0.20~0.65范围内,尾缘开缝壁面气膜冷却效率随吹风比增加而增加;在吹风比0.80~1.25范围内,冷热气掺混剧烈,冷却效率略微下降;肋板结构增加了冷气通道的阻塞效应,并限制了开缝区域冷气旋涡的发展,导致展向涡提前扭曲、变形和分解;相对于无肋板结构,带肋尾缘开缝结构的流量系数下降了约5%,气膜冷却性能提升了约10.8%。开缝下游的旋涡脱落和冷热气流间的掺混是影响开缝壁面冷却效率的主要原因,综合考虑冷气消耗和气膜冷却效率,无肋板时最佳吹风比为0.65,带肋板时最佳吹风比为0.5。

多参数变弯度翼型气动优化设计方法

为了获得机翼最优的气动外形,提出了一种多偏转参数与非均匀B样条结合的翼型前后缘连续偏转变形方法。以NACA3412翼型为研究实例,对翼型曲线进行参数化重构。以翼型前、后缘偏转位置和偏转角度为变化参数,模拟分析了不同变形状态对翼型气动特性的影响规律。以翼型升力系数和升阻比为优化目标,以前、后缘的偏转角度、偏转位置以及过渡段长度等六个偏转参数为设计变量,采用多目标遗传算法和ANSYS软件进行优化设计。结果表明,与基础翼型相比,优化后的变弯度翼型的升阻比提升约19.26%,升力提升约44.43%,明显改善了翼型的气动性能。

中/低雷诺数翼型后缘纯音噪声物理机制与主/被动控制研究综述

基于国内外研究成果,本文总结了中/低雷诺数下翼型后缘纯音噪声的物理机制与主/被动控制的研究现状,阐述了目前该研究还需回答与解决的部分问题,并对可能的新研究思路与方法进行了展望。在物理机制方面,边界层内声反馈和后缘涡脱落是公认的中/低雷诺数下翼型后缘纯音噪声的2种主要发声机制。前者由T–S不稳定波散射噪声向上游传播、在边界层与翼型后缘之间形成,产生的纯音噪声具有典型“阶梯状”离散频谱特性;后者由翼型尾流整体不稳定性引起,纯音噪声具有单一频率特性。2种发声机制之间如何竞争以及噪声频率如何选择等问题仍未完全解决。在研究方法方面,主要采用线性稳定性分析、风洞实验测量和DNS/LES数值模拟等研究方法。由于现有研究的翼型较为单一,所得结果无法满足翼型低噪声设计的需求。在噪声控制方面,被动控制方法多采用翼型前/后缘锯齿、多孔材料和柔性壁面,主动控制方法多采用表面吹/吸气和等离子体技术;在基于机理分析有效抑制翼型后缘纯音噪声方面,还有较多可改进之处。

空气涡轮火箭发动机燃烧室异形尾缘波瓣混合器掺混、燃烧特性研究

为了促进空气涡轮火箭发动机(ATR)燃烧室内涵富燃燃气与外涵空气之间的掺混,提高燃烧室燃烧效率,本文通过对波瓣进行尾缘修形设计,探究C形尾缘结构对波瓣下游流动及燃烧特性的影响规律。研究发现:(1)在波瓣混合器尾缘增加C形结构,会诱导产生多源副涡结构;(2)径向高度越高,径向速度梯度越大,造成副流向涡和副展向涡的强度和尺度越大;(3)由于受到C形尾缘结构诱导的副流向涡和副展向涡的影响,燃烧室展向截面温度分布出现“串状”局部高温区,燃烧得到强化;(4)当燃烧室距离与波瓣直径<1时,内、外涵气流掺混强烈,热混合效率、燃烧效率以及总压损失迅速增大;(5) C形尾缘结构数量与燃烧效率不成单调递增关系,当C形结构数量≤2时,副涡强化掺混起主导作用,燃烧效率随着副涡量的增大而增大;当C形结构数量≥3时,出口面积增大导致的涡量衰减起主导作用,燃烧效率随着副涡量的增大而减小。

小攻角下锯齿尾缘翼型噪声控制与机理分析

受猫头鹰寂静飞行能力的启发,锯齿尾缘设计被认为是一种有效的控制湍流边界层-尾缘干涉噪声的方法。本文采用隐式大涡模拟法,详细研究了嵌入式锯齿尾缘对NACA 0012翼型绕流的近场流动和噪声特性的影响,雷诺数为9.6×10^(4),远场马赫数为0.1631,攻角为4°,计算采用的非结构化网格具有约7000万的自由度。在实际计算时,为促进流动快速转捩,在直尾缘和锯齿尾缘算例的翼型表面均布置了锯齿形粗糙元转捩带。研究结果表明:相比于0°攻角状态,4?攻角下的噪声辐射增强,主辐射方向发生偏转,在该方向上锯齿尾缘实现了约2.5 dB的降噪,且在小攻角(4°)下,锯齿也会诱导出有利于降噪的侧边涡对结构。针对壁面压力脉动的分析表明:锯齿主要改变了尾缘附近的时空关联特性,且压力场不能直接由现有针对速度场的Taylor或椭圆近似模型定量描述;此外,锯齿在抑制尾缘噪声的同时,对翼型气动性能造成了一定损失。

连续变弯度后缘飞机的滚转机动载荷减缓

变弯度后缘机翼具有变形连续、阻力较小、气动噪声较低等优势,越来越多地应用于新概念飞行器的设计之中。基于此,提出一种基于变弯度后缘的飞行器刚弹耦合动力学建模方法,并针对变弯度后缘飞机缩比模型开展了滚转机动仿真分析与风洞试验。结果表明:变弯度后缘可以操纵飞机在2 s内进行180°滚转机动。相较于外侧后缘单独变形,通过内外后缘协同变形可以降低30%以上的滚转机动附加载荷。另外,对比表明滚转角仿真结果与风洞试验数据误差小于6%,验证了所提方法的准确性。

变弯度柔性后缘机翼动响应减缓试验

变体飞行器能显著提升飞行器的气动性能,而变弯度柔性后缘是实现变体飞行器的重要方式之一。为探究柔性后缘动态偏转下机翼动响应特性及减缓效率,设计了一个变弯度柔性后缘机翼模型并开展了风洞试验。该机翼模型由承弯翼梁和6个3D打印的翼段组成。其中,2个翼段后缘分别设计2个变弯度柔性后缘舵面,这2个舵面分别用于动响应激励和动响应减缓控制。变弯度柔性后缘舵面由数字舵机、柔性索、波纹板结构和聚二甲基硅氧烷(PDMS)柔性蒙皮组成。对变弯度柔性后缘进行地面静态偏转试验和地面动态偏转试验测试,以研究后缘弯度变形规律及舵机动态时滞特性。在此基础上,在低速风洞试验中研究变弯度柔性后缘机翼动响应规律和基于变弯度柔性和闭环反馈控制的动响应减缓效率。风洞试验结果表明:机翼的翼尖加速度响应及翼根弯矩在频率为1.5~4 Hz时先增大后减小,并在接近机翼一弯频率时达到峰值。采用PID控制律和变弯度后缘进行闭环反馈控制后,在风速20 m/s、扰动频率2.2 Hz时翼尖加速度最大减缓效率达到70.18%,翼根弯矩的最大减缓效率为68.14%。此外,还提出了动响应减缓效率为正值的理论公式并分析了动响应减缓效率的影响机制和因素。

基于变桨与襟翼协同的降载特性研究及机理分析

研究了尾缘襟翼长度、偏转角和攻角对翼型降载特性的影响,并分析其影响机理。结果表明:变桨与尾缘襟翼的协同控制可有效实现挥舞方向上的降载;尾缘襟翼长度越大,尾缘偏转角越小,攻角越远离临界值,降载性能越佳;偏转角增大会导致临界攻角位置前移;为实现快速降载,在正偏转角、大攻角下优先选择尾缘控制,其他情况下均优先选择变桨控制;降载特性与翼型及翼缝流体分布密切相关,压差越小,降载性能越好。

电火花成型电极修整航空发动机叶片的加工误差分析

针对航空发动机叶片电解加工后前后缘存在较大加工误差这一问题,采用电火花加工技术对叶片前后缘进行修整,并分析前后缘的加工误差。使用自研的六轴电火花成型机床,采用成型电极对叶片后缘进行修整加工,并利用三坐标测量机完成叶片点云数据的测量。采用遗传算法和迭代最近点算法,完成包络加工后叶片点云数据与理论模型的配准,通过最小二乘椭圆以及NURBS曲线拟合处理点云数据。针对拟合的叶片加工后轮廓曲线,通过在该曲线上等间距取点并计算这些点到叶片理论轮廓线的距离来反映叶片后缘的加工误差和加工误差的整体趋势。结果显示:经过电火花成型电极修整后叶片后缘的面轮廓度误差在-16.4~88.5μm之间,平均误差为53.7μm,叶片后缘的缘头处轮廓度误差最小,验证了电火花加工技术包络修整叶片前后缘的可行性。

翼面形状对翼伞气动性能的影响

为了研究不同翼面形状对翼伞气动性能的影响,文章以某型冲压翼型为对象,建立了翼伞滑翔阶段稳态绕流流场数值模型,分析了翼面后缘形状对稳态滑翔阶段翼伞流场结构及气动特性的影响,在此基础上探究了后缘前掠角大小对翼伞气动性能的影响。研究结果表明,后缘前掠型翼面的上翼面前缘压力降低,有利于减小翼面压差阻力,升阻性能最好,翼伞设计时宜采用后缘前掠型翼面设计;随着前掠角增大,升力系数和阻力系数均先减小后增加,翼伞升阻比随着前掠角增大呈现先增后减的趋势;当前掠角过大时,上翼面流动分离区变大,压差阻力增加会导致升阻比降低;综合流场结构和气动特性考虑,前掠角为6°时翼伞气动性能最优。研究成果可为高滑翔翼伞设计提供一定参考。

风力机翼型尾缘迎风降载工况挥舞振动风洞试验研究

停机变桨降载是极端台风天气下保障大型风力机结构安全的重要措施,风力机叶片尾缘迎风(攻角180°)时,翼型的气动力系数相对较小,是一种有利的降载工况,但有可能发生挥舞方向单自由度振动。本文基于风洞试验,研究了典型风力机翼型在尾缘迎风降载工况下的挥舞单自由度气动弹性振动特性。研究发现,叶片不同截面的翼型在特定的来流攻角和结构无量纲减缩频率下均会触发挥舞失稳振动,总体的失稳攻角范围为155°~167°。失稳区域呈倒“V”字形,结构减缩频率越低,诱发挥舞振动的攻角范围越大。这种挥舞振动表现出锁频特性,即振动频率锁定于结构固有频率,且振动幅值随着减缩频率(与来流风速成反比)的降低而增加。对于普遍具有预扭设计的现代风力机叶片,叶片在进行尾缘迎风的降载设计时,应尽量避免主要叶片截面进入挥舞失稳敏感攻角区域,以免造成结构损坏。