圆形坚冰影响翼型气动性能的数值分析

本文用FLUENT软件模拟了结冰后NACA 0012翼型周围流场的变化,并与结冰前NACA 0012翼型的气动性能进行了对比.工作中首先以未结冰的NACA 0012翼型(干净翼型)为标准模型进行了数值验证计算,再以经过检验的方法计算结冰模型,并与结冰风洞试验数据进行了对比.本文计算攻角为0°～20°,温度为250.37K,雷诺数为2,400,000,冰型为圆形坚冰.通过对比升力阻力性能,发现与干净翼型相比,结冰翼型的最大升力系数大约减少了50%,阻力系数增加了约65%,失速攻角降低了4°.结冰后翼型提前失速是造成气动性能恶化的主要原因.     

UNSTEADY SEPARATION OF FLOW AROUND AIRFOIL WITH LOCAL ELASTIC STRUCTURE

对在低雷诺数下局部弹性翼型绕流中，局部弹性导致的自激振动所产生的复杂非定常流动分离现象和描述方法进行了分析．采用ALE—CBS方法数值模拟了具有可动边界的绕流流场问题，同时采用Galerkin方法求解局部弹性结构的控制方程．着重研究了翼型的局部弹性对流动分离和翼型性能的影响，并分别从Eulerian和Lagrangian的角度分析了局部弹性结构导致的不同非定常分离现象，其中Lagrangian角度可以方便地揭示出局部弹性翼型大幅度提高升力的机理和流动中的能量迁移．结果表明翼型的局部弹性对非定常分离和分离泡的演化过程有着明显的影响，可以使得流体质点由主流获取动量实现再附，并且在一定的攻角下可以将固定分离转变为移动分离，从而明显地提高了翼型的升力．     

局部弹性翼型非定常分离的动力学特性

对在低雷诺数下局部弹性翼型绕流中, 局部弹性导致的自激振动所产生的复杂非定常流动分离现象和描述方法进行了分析. 采用ALE-CBS方法数值模拟了具有可动边界的绕流流场问题, 同时采用Galerkin方法求解局部弹性结构的控制方程. 着重研究了翼型的局部弹性对流动分离和翼型性能的影响, 并分别从Eulerian和Lagrangian的角度分析了局部弹性结构导致的不同非定常分离现象, 其中Lagrangian角度可以方便地揭示出局部弹性翼型大幅度提高升力的机理和流动中的能量迁移. 结果表明翼型的局部弹性对非定常分离和分离泡的演化过程有着明显的影响, 可以使得流体质点由主流获取动量实现再附, 并且在一定的攻角下可以将固定分离转变为移动分离, 从而明显地提高了翼型的升力.     

低雷诺数翼型蒙皮主动振动气动特性及流场结构数值研究

针对低雷诺数(Re)翼型气动性能差的特点,文章通过对翼型柔性蒙皮施加主动振动的方法,提高翼型低Re下的气动特性,改善其流场结构.采用带预处理技术的Roe方法求解非定常可压缩Navier--Stokes方程,对NACA4415翼型低Re流动展开数值模拟.通过时均化和非定常方法对比柔性蒙皮固定和振动两种状态下的升阻力气动特性和层流分离流动结构.初步研究工作表明在低Re下柔性蒙皮采用合适的振幅和频率,时均化升阻力特性显著提高,分离泡结构由后缘层流分离泡转变为近似的经典长层流分离泡,分离点后移,分离区缩小.在此基础上,文章更加细致研究了柔性蒙皮两种状态下单周期内的层流分离结构及壁面压力系数分布非定常特性和演化规律.蒙皮固定状态下分离区前部流场结构和压力分布基本保持稳定,表现为近似定常分离,仅在后缘位置出现类似于卡门涡街的非定常流动现象.柔性蒙皮振动时从分离点附近开始便产生分离涡,并不断向下游移动、脱落,表现为非定常分离并出现大范围的压力脉动.蒙皮振动使流体更加靠近壁面运动,大尺度的层流分离现象得到有效抑制.     

平板大攻角绕流升力和阻力系数的计算

二维平板或二维对称薄翼型大攻角绕流升力和阻力系数与攻角之间存在的函数关系一般用数据表格的形式给出。本文根据垂直平板绕流阻力实验数据和对称薄翼型全攻角绕流实验数据,分析得到了平板大攻角绕流总压力及其升力分量和阻力分量系数的近似计算公式。结果表明：平板总压力系数约等于攻角正弦值的2倍;总压力的阻力分量系数约等于攻角正弦值平方的2倍;升力分量系数约为攻角2倍的正弦值。计算结果与两组试验数据具有较好的一致性。     

基于涡粘性模型的翼型湍流流场熵产率计算

利用有限体积法实现了基于非正交同位网格的SIMPLE算法。基于熵分析方法,采用涡粘性模型求解湍流熵产方程,系统研究了湍流模型对二维翼型绕流流场熵产率的影响。通过计算NACA0012翼型在来流雷诺数为2.88×106时,0°攻角~16.5°攻角范围内的翼型表面压力系数分布和升阻力特性,验证了算法及程序的正确性。结果表明,选择不同湍流模型时,翼型流场熵产的计算结果存在差异,湍流耗散是引起流场熵产的主要原因;翼型流场的熵产主要发生在翼型前缘区、壁面边界层和翼型尾流区域,流场熵产率与翼型阻力系数线性相关;当产生分离涡时,粘性耗散引起的熵产下降。     

《力学学报》2006年分类目录

流体力学对称冀型低雷诺数小攻角升力系数非线性现象研究..............……白鹏崔尔杰李锋周伟江2006，38(l):1弯管内爆轰波传播的流场显示和数值模拟........................……王昌建徐胜利费立森郭长铭2006，38(l):9高猫性流动有限元模拟的迭代稳定分步算法，.....................................……韩先洪李锡夔2006，38(l):16各向同性湍流内顺粒碰撞率的直接模拟研究...........……李瑞霞柳朝晖贺铸陈溉密郑楚光2006，38(l):25三角典前缘涡破裂形式及特性研究....................................……吕志咏祝立国张明禄2006，…     

一种升浮一体环形浮空器气动特性研究

选取填充轻质气体的环形浮空器为研究对象,采用数值模拟方法开展高空风力发电机用浮空器气动特性研究.采用有限体积法求解不可压N-S方程和S-A湍流模型来数值模拟风力机气流场,分别对翼型,安装角,长细比,雷诺数及风力机等因素对引入浮力后浮空器气动特性影响进行研究,对比分析引入浮力后布局外形气动力特性随各外形特征参数的变化规律.数值结果表明,截面翼型弯度越大,最大升阻比越小,其出现位置有一定前移;截面厚度越大,三维效应越强,最大升阻比出现有一定的滞后性;增大安装角,相当于增大攻角,使得升力系数和阻力系数随攻角变化曲线均有一定前移;引入浮力后,最大合升阻比增大,并且存在一个明显前移;长细比越小,浮空器升阻比越大,随着长细比增大,浮空器最大升阻比出现越滞后;一定范围内,雷诺数增大,浮空器动升阻比增大,引入浮力后,当来流风速变化时,浮空器合升阻比随雷诺数增大先迅速减小然后趋于平缓,但基于浮空器尺寸变化时,合升阻比则随雷诺数增大而增大;风轮转速增大,浮空器阻力增大,升力有一定下降.     

基于涡粘性模型的翼型湍流流场熵产率计算

利用有限体积法实现了基于非正交同位网格的SIMPLE算法。基于熵分析方法，采用涡粘性模型求解湍流熵产方程，系统研究了湍流模型对二维翼型绕流流场熵产率的影响。通过计算NACA0012翼型在来流雷诺数为2.88×10⁶时，0°攻角~16.5°攻角范围内的翼型表面压力系数分布和升阻力特性，验证了算法及程序的正确性。结果表明，选择不同湍流模型时，翼型流场熵产的计算结果存在差异，湍流耗散是引起流场熵产的主要原因；翼型流场的熵产主要发生在翼型前缘区、壁面边界层和翼型尾流区域，流场熵产率与翼型阻力系数线性相关；当产生分离涡时，粘性耗散引起的熵产下降。     

低雷诺数下柔性翼型气动性能分析

基于流固耦合方法对吸力面5%至95%弦长处为三段柔性结构的NACA0012翼型绕流进行了数值模拟,研究了不同弹性模量下柔性翼型的气动性能和结构响应.结果表明:在大攻角下,翼面变形影响着翼型表面的非定常流场,起到延缓失速和提高升力的作用;失速后柔性翼的升力系数下降得较为缓慢,且柔性越大,升力系数下降得越平缓;适当减小弹性模量能够提高翼型的气动性能,然而弹性模量过小反而不利于翼型气动性能的提升,并且翼面会产生大幅度的振动.     

等速上仰翼型动态失速现象研究

翼型大迎角绕流的静态失速将造成升力突降和气动性能急剧恶化，但利用非定常运动所产生 的动态失速效应，可以大大地延缓气流分离和失速现象的发生. 采用Rogers发 展的双时间步Roe格式，求解拟压缩性修正不可压N-S方程. 数值模拟了低雷诺数 ($Re=4.8 \times 10^{4}$)条件下NACA0015翼型作等速上仰($\alpha =0^{\circ} \sim 60^{\circ}$)的动态失速过程，同Walker的试验结果比 较，验证了计算结果的正确性. 研究了该过程中主涡、二次涡和三次涡的发展，升 力系数随攻角变化，以及不同上仰速度对动态失速效应所造成的影响.     

横摆振荡翼型动态气动掠效应试验研究

大型风力机设计对获取翼型更加全面、准确的动态载荷提出更高要求, 研究翼型横摆振荡动态气动特性具有重要意义. 借助"电子凸轮"技术和动态数据同步采集手段, 针对翼型动态“掠效应”首次开展了横摆振荡风洞试验研究, 研究表明: 横摆振荡翼型的气动曲线存在明显迟滞效应, 吸力面压力周期性波动是主要诱因, 且随着振荡频率、初始迎角和振幅的增大, 气动迟滞特性均增强; 升力和压差阻力随横摆角变化的迟滞回线呈"W"形, 俯仰力矩迟滞回线呈"M"形, 升力差量迟滞回线呈"$\infty$"形; 负行程下翼型气动力相对于正行程下的更高, 且负行程下翼型气动力随振荡频率的增大而略有增大, 正行程下则明显减小; 升力系数功率谱密度分布在振荡频率倍频处的能量集中的幅值随着振荡频率增大有增大趋势; 吸力面1.2%和40%弦长处压力的滞回特性较强, 是由于翼面剪切层涡和动态分离涡周期性发展、运动、破裂和重建; 振幅为$10^{\circ}$时, 升力迟滞曲线呈"$^{\wedge}$"形, 振幅为$30^{\circ}$ 时, 升力迟滞曲线呈"$^{\wedge\wedge\wedge}$"形.      

Parametric study of separation and transition characteristics over an airfoil at low Reynolds numbers

Time-resolved surface pressure measurements are used to experimentally investigate characteristics of separation and transition over a NACA 0018 airfoil for the relatively wide range of chord Reynolds numbers from 50,000 to 250,000 and angles of attack from 0° to 21°. The results provide a comprehensive data set of characteristic parameters for separated shear layer development and reveal important dependencies of these quantities on flow conditions. Mean surface pressure measurements are used to explore the variation in separation bubble position, edge velocity in the separated shear layer, and lift coefficients with angle of attack and Reynolds number. Consistent with previous studies, the separation bubble is found to move upstream and decrease in length as the Reynolds number and angle of attack increase. Above a certain angle of attack, the proximity of the separation bubble to the location of the suction peak results in a reduced lift slope compared to that observed at lower angles. Simultaneous measurements of the time-varying component of surface pressure at various spatial locations on the model are used to estimate the frequency of shear layer instability, maximum root-mean-square (RMS) surface pressure, spatial amplification rates of RMS surface pressure, and convection speeds of the pressure fluctuations in the separation bubble. A power-law correlation between the shear layer instability frequency and Reynolds number is shown to provide an order of magnitude estimate of the central frequency of disturbance amplification for various airfoil geometries at low Reynolds numbers. Maximum RMS surface pressures are found to agree with values measured in separation bubbles over geometries other than airfoils, when normalized by the dynamic pressure based on edge velocity. Spatial amplification rates in the separation bubble increase with both Reynolds number and angle of attack, causing the accompanying decrease in separation bubble length. Values of the convection speed of pressure fluctuations in the separated shear layer are measured to be between 35 and 50% of the edge velocity, consistent with predictions of linear stability theory for separated shear layers.     

圆角化对方柱气动性能影响的流场机理

方形截面柱体的圆角化处理是常用的流动控制方法,但其流场作用机理尚未被澄清.采用大涡模拟方法,在雷诺数为2.2$\times$10$^{4}$时,考虑风攻角的影响,对均匀流作用下的标准方柱和圆角方柱的气动性能和流场特性进行了研究,定量分析了圆角化气动措施和风攻角变化对分离泡特性的影响规律,从流场角度澄清了圆角化气动措施对方柱气动性能的影响机理.研究表明:与标准方柱相比,圆角方柱的表面风压、气动力和涡脱强度呈整体下降的趋势,但圆角方柱的斯特劳哈尔数更高;圆角方柱的"分离泡流态'发生在更小的风攻角范围内,分离泡的出现会进一步造成方柱的尾流变窄,涡脱强度减弱;随着风攻角的增大,分离泡的长度会逐渐减小直至消失,分离泡的中心会逐渐向方柱前角(迎风向)和方柱壁面移动;与标准方柱相比,圆角方柱的气流发生初次分离的位置向下游移动,分离后的剪切层更贴近方柱,因而更易发生再附现象;方柱尾流宽度的减小和涡脱强度的减弱是导致圆角方柱气动力减小和斯特劳哈尔数增大的主要原因.      

Time-resolved volumetric particle tracking velocimetry of large-scale vortex structures from the reattachment region of a laminar separation bubble to the wake

The present paper presents time-resolved volumetric Particle Tracking Velocimetry measurements in a water towing tank on a SD7003 airfoil, performed at a Reynolds number of 60,000 and a 4° angle of attack. The SD7003 airfoil was chosen because of its long mid-chord and stable laminar separation bubble (LSB), occurring on the suction side of the airfoil at low Reynolds numbers. The present study focuses on the temporal resolution of unsteady large-scale vortex structures emitted from the LSB. In contrast to other studies, where only the observation of the flow in the transition region was examined, the entire flow from the leading edge to the far wake of the airfoil was investigated here.     

Laminar separation bubble on an Eppler 61 airfoil

Laminar separation bubble that occurs on the suction side of the Eppler 61 airfoil at Re=46000 is studied. The incompressible flow equations are solved using a stabilized finite element method. No turbulence model is used. The variation of the bubble length and its location, with the angle of attack (α), is studied in detail. An abrupt increase in the lift coefficient is observed at α∼4.5°. It is found to be related to a sudden decrease in the separation bubble length at the trailing edge of the airfoil. Significant differences are observed in the results from the 2D and 3D computations. Stall is observed in 3D simulations, but is found to be absent in 2D. The laminar bubble, which fails to reattach in 3D for α>14°, continues to reattach for α as large as 20° in the 2D computations. Reynolds stress calculations in both 2D and 3D indicate the extent to which the outer flow is affected by the presence of bubble. It is found that the Reynolds stress components ${\over{u{^\prime}}{v{^\prime}}}$ and ${\over{u{^\prime}}{w{^\prime}}}$ are of comparable order of magnitude indicating that spanwise fluctuations are significant. The effect of the time window used to compute the time‐averaged aerodynamic coefficients is studied. The time‐averaged and root mean square (rms) value of the aerodynamic coefficients are calculated for both 2D and 3D computations and compared with the previously published experimental results. The 3D computations show good agreement with the earlier data. The variation of the rms value of the aerodynamic coefficients with angle of attack shows certain peaks. The cause of their appearance is investigated. The effect of Reynolds number is studied. The increase in Re at α=10° is found to reduce the bubble length and cause it to move closer to the leading edge. Copyright © 2009 John Wiley & Sons, Ltd.     

High-fidelity simulations of moving and flexible airfoils at low Reynolds numbers

The present paper highlights results derived from the application of a high-fidelity simulation technique to the analysis of low-Reynolds-number transitional flows over moving and flexible canonical configurations motivated by small natural and man-made flyers. This effort addresses three separate fluid dynamic phenomena relevant to small fliers, including: laminar separation and transition over a stationary airfoil, transition effects on the dynamic stall vortex generated by a plunging airfoil, and the effect of flexibility on the flow structure above a membrane airfoil. The specific cases were also selected to permit comparison with available experimental measurements. First, the process of transition on a stationary SD7003 airfoil section over a range of Reynolds numbers and angles of attack is considered. Prior to stall, the flow exhibits a separated shear layer which rolls up into spanwise vortices. These vortices subsequently undergo spanwise instabilities, and ultimately breakdown into fine-scale turbulent structures as the boundary layer reattaches to the airfoil surface. In a time-averaged sense, the flow displays a closed laminar separation bubble which moves upstream and contracts in size with increasing angle of attack for a fixed Reynolds number. For a fixed angle of attack, as the Reynolds number decreases, the laminar separation bubble grows in vertical extent producing a significant increase in drag. For the lowest Reynolds number considered (Re _c  = 10⁴), transition does not occur over the airfoil at moderate angles of attack prior to stall. Next, the impact of a prescribed high-frequency small-amplitude plunging motion on the transitional flow over the SD7003 airfoil is investigated. The motion-induced high angle of attack results in unsteady separation in the leading edge and in the formation of dynamic-stall-like vortices which convect downstream close to the airfoil. At the lowest value of Reynolds number (Re _c  = 10⁴), transition effects are observed to be minor and the dynamic stall vortex system remains fairly coherent. For Re _c  = 4 × 10⁴, the dynamic-stall vortex system is laminar at is inception, however shortly afterwards, it experiences an abrupt breakdown associated with the onset of spanwise instability effects. The computed phased-averaged structures for both values of Reynolds number are found to be in good agreement with the experimental data. Finally, the effect of structural compliance on the unsteady flow past a membrane airfoil is investigated. The membrane deformation results in mean camber and large fluctuations which improve aerodynamic performance. Larger values of lift and a delay in stall are achieved relative to a rigid airfoil configuration. For Re _c = 4.85 × 10⁴, it is shown that correct prediction of the transitional process is critical to capturing the proper membrane structural response.     

Effect of Local DBD Plasma Actuation on Transition in a Laminar Separation Bubble

This work examines the effect of local active flow control on stability and transition in a laminar separation bubble. Experiments are performed in a wind tunnel facility on a NACA 0012 airfoil at a chord Reynolds number of 130 000 and an angle of attack of 2 degrees. Controlled disturbances are introduced upstream of a laminar separation bubble forming on the suction side of the airfoil using a surface-mounted Dielectric Barrier Discharge plasma actuator. Time-resolved two-component Particle Image Velocimetry is used to characterise the flow field. The effect of frequency and amplitude of plasma excitation on flow development is examined. The introduction of artificial harmonic disturbances leads to significant changes in separation bubble topology and the characteristics of coherent structures formed in the aft portion of the bubble. The development of the bubble demonstrates strong dependence on the actuation frequency and amplitude, revealing the dominant role of incoming disturbances in the transition scenario. Statistical, topological and linear stability theory analysis demonstrate that significant mean flow deformation produced by controlled disturbances leads to notable changes in stability characteristics compared to those in the unforced baseline case. The findings provide a new outlook on the role of controlled disturbances in separated shear layer transition and instruct the development of effective flow control strategies.     

基于雨燕翅膀的仿生三角翼气动特性计算研究

针对低雷诺数微型飞行器的气动布局, 设计出类似雨燕翅膀的一组具有不同前缘钝度的中等后掠($\varLambda =50^{\circ}$)仿生三角翼. 为了定量对比研究三角翼后缘收缩产生的气动效应, 设计了一组具有同等后掠的普通三角翼. 为了深入研究仿生三角翼布局的前缘涡演化特性以及总体气动特性, 采用数值模拟方法详细地探索了低雷诺数($Re=1.58\times 10^{4})$流动条件下前缘涡涡流结构和气动力随迎角的变化规律. 分析结果表明, 前缘钝度和后缘收缩对仿生三角翼前缘涡的涡流强度和涡破裂位置有显著影响. 相对于钝前缘来说, 尖前缘使仿生三角翼上下表面的压力差增大, 涡流强度也更大, 增升作用也更显著. 相对于普通三角翼构型, 仿生三角翼的前缘斜切使其阻力更大, 但后缘的收缩使涡破裂位置固定在此位置, 因此整个上翼面保持低压, 总的升力更大. 由于小迎角时升力增大更明显, 因此仿生三角翼的气动效率在小迎角时明显大于普通三角翼. 这些结论对于揭示鸟类的飞行机理以及未来微型仿生飞行器的气动布局设计具有重要的研究价值.