计入分离涡时飞机机翼振动的极限环现象

通过同时求解机翼弹性运动方程和空气动力方程，对机翼的流固耦合问题进行了数值仿真，在空气动力计算中仔细地模拟了机翼在较大迎角时从机翼前缘和侧缘处卷起的分离涡。通过不同机翼、不同状态的仿真结果，发现如下结论：（１）机翼前缘或侧缘分离涡将使机翼颤振界速度降低，且机翼的基本迎角越大，颤振临界速度越低；（２）机翼前缘后掠角减小，分离涡对颤振性质的不利影响有所减缓；（３）在一定条件下，会出现稳定的极限环。     

低雷诺数反齐默曼机翼的气动特性

通过PISO方法求解非定常不可压N-S方程,研究小展弦比反齐默曼机翼在低雷诺数下的流场特征,并分析其对气动特性的影响.结果显示前缘分离涡在反齐默曼机翼上表面形成一对集中涡.分析表明这对集中涡是影响反齐默曼机翼气动特性的主要因素,给机翼提供了较大的非线性升力和较大的失速攻角,前缘涡之间的相互影响使得机翼出现非定常现象和大攻角的非对称现象.与矩形翼相比,反齐默曼机翼有较好的稳定性.     

波动翼非定常流场IB-LBM数值研究

经过亿万年自然选择, 鱼类进化出非凡高效的游动能力, 研究其游动机理, 对改善现有潜水器的性能具有重要指导意义.针对类鳗鱼游动问题, 采用浸入边界法-格子Boltzmann方法(immersed boundary-lattice-Boltzmann, IB-LBM)对三维波动翼进行1×108网格大规模数值模拟.在广州超算中心天河-2上模拟了不同振动幅度下正弦波动翼的非定常运动, 给出了流场涡系结构及其产生的非定常力, 清晰捕捉了仿生翼非定常涡系演化过程.仿真结果表明IB-LBM方法能在较大运动边界情况下保持算法稳定性, 也能在较大网格下高效运行.同时精细捕捉不可压非定常流场涡系结构细节, 是一种较为理想的仿生运动数值模拟方法.      

三角翼涡破裂的高精度数值模拟

采用5阶精度的加权紧致非线性格式(WCNS-E-5)数值模拟65°后掠角尖前缘三角翼的大攻角跨声速绕流流场,考察低耗散、高分辨率的WCNS-E-5格式对于三角翼涡破裂模拟的适用性,及激波旋涡干扰对涡破裂点位置的影响,重点研究三角翼大攻角旋涡破裂点的突然前移.通过求解任意坐标系下的非定常雷诺平均N-S方程,采用WCNS-E-5和SST两方程湍流模型,与试验结果和文献计算结果对比,表明既有高阶精度又能光滑捕捉激波的WCNS格式在模拟三角翼旋涡破裂方面具有一定优势,其数值结果与试验结果吻合较好,三角翼大攻角旋涡破裂点的突然前移是由于跨声速流场的激波旋涡干扰.     

边条翼后掠角对钝头细长旋成体非对称流动的影响

为探究大迎角时边条翼前缘后掠角对钝头细长旋成体导弹绕流特性的影响,开展了模型表面测压和粒子图像测速的风洞实验,研究了亚临界Reynolds数Re=150000、迎角α=50°条件下不同前缘后掠角固定边条翼的钝头细长旋成体非对称绕流特性.结果表明,在边条翼上游区,后掠角增大使边条涡涡位更靠近前体物面且对称性更好,导致前体...     

振荡机翼跨音速非定常粘性绕流的数值计算

以非定常Ｎ－Ｓ方程为主管方程，采用ＬＵ－ＮＮＤ混合差分格式，Ｃ和Ｃ－Ｈ型贴体运动网格，Ｂ－Ｌ双层代数紊流模型，求解绕振荡翼型和三维机翼的跨音速非定常粘性流场，分别计算了ＮＡＣＡ００１２翼型和Ｍ６机翼作俯仰振荡时跨音速非定常粘性绕流流场。研究了非定常绕流的气动特性，部分计算结果和风洞实验值作了比较。     

利用N-S方程模拟机翼气动弹性的一种计算方法

利用一种双时间方法求解三维非定常N-S方程,得到与任意非定常运动对应的气动力,在求解非定常气动力的同时,在时间域内用二阶龙格 库塔方法求解机翼弹性运动方程,从而模拟粘性流动中的气动弹性全过程.为保证网格生成效率,采用无限插值理论生成O-H型代数网格,考虑了机翼变形时的网格生成问题,并得到计算结果.     

蜗壳非轴对称性对离心压缩机内部非定常流动影响的数值研究

本文采用非定常数值模拟方法研究了蜗壳几何非轴对称性对离心压缩机内部非定常流场的影响。结果表明:流场的非定常波动主要受叶轮影响。远离蜗壳设计工况时,蜗壳非对称性的影响可传播至叶轮主叶片前缘。蜗壳非对称性对扩压器叶片压力面瞬时压力的影响在小流量和设计流量工况下主要体现叶轮主叶片与分流叶片通过频率的影响,而在大流量工况下叶轮主轴旋转频率也占一定比重。叶轮主叶片压力面的非定常压力波动主要受叶轮主轴旋转频率及其倍频与扩压器叶片通过频率及其倍频的影响。     

合成双射流对下游声压级影响试验

在Ma=0.4的来流条件下, 利用安装在主翼后缘处的合成双射流激励器对襟翼上的流动进行控制, 在风洞中开展了合成双射流对下游声压级影响的研究. 基于脉动压力测量结果, 结合油流显示试验, 得到了合成双射流对下游不同流动状态区域声压级影响的一些结论. 对于附着流, 在其峰值频率附近激励会明显提高其声压级; 对于受旋涡主导的流动, 恰当的合成双射流控制可以降低声压级, 激励频率较为关键. 在俯仰运动过程中, 对于附着流, 激励提高了声压级, 但不改变其迟滞特性; 对于受旋涡主导的流动, 激励对声压级的影响与攻角有关, 能够减弱其迟滞特性, 但激励强度对迟滞特性的影响较小, 减小声压级的最佳激励与运动历程有关.      

风力机翼型在失速工况下非定常流场的本征正交分解分析

风力机气动性能受静态失速与动态失速影响很大,对风力机翼型的失速问题研究具有重要意义。本文通过计算流体力学方法得到的风力机翼型在固定大攻角工况,以及大攻角震荡工况下的非定常流场,来研究翼型静态失速与动态失速。采用本征正交分解方法(POD),对非定常流场降阶,得到流场的POD模态以及对应的系数。POD模态结果表明在静态失速下,主要非定常流动结构是尾迹区域交替脱落的涡结构;在动态失速下,除了尾迹区域,前缘和整个吸力面都存在流动分离结构。     

Controlled Vortex Breakdown on Modified Delta Wings

This paper studies the effect of perturbation to the breakdown of the leading-edge vortices over delta wings. The passive perturbation in the normal direction is achieved by installing the hemisphere-like bulges on the delta wing along the projection of the vortices. The key purpose of this perturbation is to delay or suppress vortex breakdown over delta wings according to the self-induction mechanism theory. The design of bulge-like surface for delta wings offers a minimization of initial vorticity gradient and an elimination of linearly mutual induction within the vortex core. Three delta wings with swept angles of 60°, 65dg and 70° have been used. Dye flow visualization and force measurement in different water tunnels are performed at the water speed of U=0.10, 0.15, 0.20 and 0.25 m/s. In flow visualization, the results show contributions of bulges as perturbation to leading-edge vortices. The best outcome of perturbing the vortex core occurs in the case of the 65° delta wing. The breakdown positions on the 65° delta wing are delayed in almost the entire range of angles of attack, and that, the results are presented here.     

光学炮塔气动载荷的非稳态特性

采用耦合J-B模型的IDDES模型与双时间步LU-SGS方法开展炮塔非稳态气动载荷的数值仿真研究。炮塔流动会发生分离,天顶位置的分离角大于90°;当流动绕过炮塔时,形成马蹄涡、脱落涡街等非稳态流场结构,导致气动载荷也具有非稳态特性;炮塔顶点的脉动静压功率谱在1.6~40.0 kHz进入各向同性均匀湍流的惯性子区,基本满足Kolmogrov的-5/3定律;气动力以阻力为主,横向力的脉动幅值大,气动力矩则以俯仰力矩为主,滚转力矩的脉动幅值大,偏航力矩可以忽略不计;气动力和力矩的功率谱主要集中在1 kHz以下,存在多个尖峰频率,主频约为230 Hz (斯特劳哈尔数为0.15)。在ATP系统设计之初,需要考虑光学炮塔所受气动载荷的非稳态特性,并规避尖峰频率尤其是主频的谐振破坏问题。     

Flow visualization and aerodynamic-force measurement of a dragonfly-type model

An unsteady flow visualization and force measurement were carried out in order to investigate the effects of the reduced frequency of a dragonfly-type model. The flow visualization of the wing wake region was conducted by using a smoke-wire technique. An electronic device was mounted below the test section in order to find the exact position angle of the wing for the visualization. A load-cell was employed in measuring aerodynamic forces generated by a plunging motion of the experimental model. To find the period of the flapping motion in real time, trigger signals were also collected by passing laser beam signals through the gear hole. Experimental conditions were as follows: the incidence angles of the foreand hind-wing were 0° and 10°, respectively, and the reduced frequencies were 0.150 and 0.225. The freestream velocities of the flow visualization and force measurement were 1.0 and 1.6m/sec, respectively, which correspond to Reynolds numbers of 3.4 × 10³ and 2.9 × 10³. The variations of the flow patterns and phase-averaged lift and the thrust coefficients during one cycle of the wing motion were presented. Results showed that the reduced frequency was closely related to the flow pattern that determined flight efficiency, and the maximum lift coefficient and lift coefficient per unit of time increased with reduced frequency.     

The impact of the pitching motion on the structure of the vortical flow over a slender delta wing under sideslip angle

The primary purpose of this investigation is to observe the effect of the pitching motion on the vortical flow structure and bursting of leading-edge vortices over a delta wing under the sideslip angle, β using a dye visualization technique. In the current work, a delta wing with a sweep angle of Λ = 70° was oscillated in upstroke and downstroke direction to be able to discover the influence of pitching motion on the flow characteristics of the delta wing. The values of mean angles of attack were selected as α_m = 25° and α_m = 35°, and the sideslip angle was altered from β = 0 to 16°. The delta wing oscillated with the various periods of T_e = 5 s, 20 s, and 60 s, respectively. Amplitude of motion was adjusted as α_o =  ± 5°. It is found that the pitching motion of the delta wing under the sideslip angle β varies the location of the vortex bursting and vortical flow structure substantially.

     

LIMIT CYCLE FLUTTER OF AIRFOILS IN STEADY AND UNSTEADY FLOWS

The limit cycle flutter of a two-dimensional wing with non-linear pitching stiffness is investigated. For modelling the aerodynamic forces of the wing steady linear and non-linear models as well as an unsteady model were used. The flutter speed was calculated using the harmonic balance method and by predicting Hopf bifurcation. Analytical solutions based on the centre manifold theory and normal forms were obtained as were results given by the harmonic balance method. The analytical solutions were compared with those obtained by numerical integration. The results show that the harmonic balance method can forecast flutter speed with a good accuracy while analytical solutions based on centre manifold theorem are accurate only in a small neighbourhood of the bifurcation point. The oscillation of the airfoil after flutter for two different models, linear and non-linear pitching stiffness were compared with each other and the flutter speeds for two linear steady and an unsteady aerodynamic model calculated. The obtained results show that flutter analysis based on the linear steady model is conservative only for the ratios of plunge frequency to pitch frequency lower than 1.     

Flow field measurements of leading-edge separation vortex formed on a delta wing with vortex flaps

Wind tunnel tests are carried out using a 70 delta wing model with leading-edge vortex flaps. The structure of the leading-edge separation vortex over the leading-edge vortex flap is measured by use of a 5 holes pitot probe, surface pressure measurement technique and oil flow visualization technique. Separation vortices formed on a plain delta wing, on a vortex flap and inboard the vortex flap hinge line are clearly visualized. Results indicate that the flow around the vortex flaps is classified into several different cross flow patterns. The streamwise flap deflection angle is defined to discuss the vortex flap performance. The optimum lift to drag ratio is attained when the amount of the wing angle of attack is not far different from that of the streamwise flap deflection angle, as long as the vortex flap is deflected modestly.     

Quantitative visualization of the leading-edge vortices on a delta wing by using pressure-sensitive paint

Leading-edge vortices on a simple delta wing were visualized by using pressure-sensitive paint (PSP). PSP is an optical pressure measurement technique based on oxygen quenching of luminescent molecules. In the present study, we used PSP composed of platinum octaethylporphyrine (PtOEP) and fluoropolymer (poly-IBM-co-TFEM [Poly (isobutylmethacrtlate-co-trifluoroethylate)]). This new paint has higher sensitivity to pressure and lower sensitivity to temperature than previous ones, reducing an error due to temperature variation during a wind tunnel test. A thin coating of PSP was applied to a delta wing model with 70-degree leading-edge sweep. The coating was excited by Xenon light and emission from the coating was detected by a high-resolution CCD camera. Tests were done at subsonic speeds in the 0.2-m Supersonic Wind Tunnel at the National Aerospace Laboratory in Japan. Complicated flow structures on the delta wing including primary and secondary vortices were clearly visualized using pressure-sensitive paint. An a priori calibration technique was used to convert measured luminescent intensity into pressure. The obtained pressure distributions were in good agreement with pressure tap data. Pressure maps were obtained for various Mach numbers, Reynolds numbers and angles of attack. It was found that an increase in Mach number delayed vortex breakdown while Reynolds number had little effect on the vortex formation.     

Analytical prediction of the non-linear response of a self-excited structure

The time-dependent amplitude response of the self-excited harmonium reed vibrating at finite amplitudes is investigated both analytically and experimentally. The analysis contained herein is based on the assumption that all of the significant non-linear forces that act on the reed are of aerodynamic origin and that these forces influence the reed behavior through the system damping. The analysis is carried out, without the aid of empirical techniques, by deriving the induced aerodynamic pressure force as a non-linear function of the amplitude of the reed motion, via an unsteady potential flow field analysis, and then applying the resulting forcing function to the equation of motion of the reed. An approximate solution of the equivalent lumped parameter equation of motion yields functional relationships that predict various observable phenomena, including the limit cycle amplitude. A comparison of the results with experimental data serves to substantiate the analysis. Also presented is a brief discussion of some of the phenomenological similarities that exist between the results of the harmonium reed analysis and the observed behavior of a flat-plate wing undergoing torsional stall flutter.     

弹性载机对外挂物动力学响应的影响

在外挂物投放过程中, 载机对外挂物具有气动干扰效应, 产生附加气动力.对于弹性机翼, 在外挂物分离投放时, 相当于给机翼一个初始扰动, 机翼将发生弹性振动, 该振动也会对外挂物带来气动干扰效应.通过耦合求解非定常N-S方程刚体六自由度方程和基于模态法的结构动力学方程, 对考虑弹性变形的载机外挂物分离投放过程进行模拟, 研究了弹性机翼对外挂物的气动干扰效应.研究结果表明:在外挂物分离初期, 弹性机翼的干扰对外挂物气动力响应产生显著影响, 机翼的主要结构模态频率决定了外挂物气动力的变化频率, 并且由载机机翼动弹性变形引起的干扰气动力能占到外挂物总气动力的一半左右.      

叶轮机械叶片流体激振安全性的全功能分析

本文提出了对叶轮机械叶片流体激振安全性实行全功能数值分析的新概念和方法。实现气动力学和弹性力学的弱耦合计算,由弹性力学方法确定三维振型,用振荡流体力学方法确定三维条件下的三维振荡流场,确定在各种振幅条件下的非定常气动力作功与阻尼功,引入气动功等于阻尼功条件下的平衡振幅概念。根据平衡振幅确定叶片流体激振的动应力,并进行寿命评估。这种全功能分析将使叶轮机械流体激振的安全性分析更加符合工程实际的需要,也能够解释大量的流体激振破坏不属于瞬时突然破坏,而是一种疲劳破坏。