基于CFD／CSD的大展弦比机翼气动弹性分析

大展弦比飞机在飞行过程中受气动载荷影响,其大展弦比机翼产生弯曲和扭转变形,这种弹性变形严重影响飞机的飞行性能和飞行安全,不能将此种飞机机翼当作传统的刚性机翼进行气动分析。针对一大展弦比机翼,采用气动／结构耦合分析方法,利用计算流体动力学（CFD）软件CFX和计算结构动力学（CSD）软件ANSYS联合求解,研究了在不同载荷情况下大展弦比机翼静气动弹性变形对机翼气动特性的影响。结果表明,大展弦比无人机机翼受载变形后升阻比降低,升力下降明显,阻力略有上升,机翼翼尖容易失速。     

基于径向基函数的网格变形及非线性气动弹性时域仿真研究

为开展非线性气动弹性研究,基于非线性结构有限元软件NASTRAN和自主研制的多块结构化计算流体力学(CFD)求解器,开发了一套基于计算流体力学/计算结构动力学(CFD/CSD)耦合求解方法的气动弹性时域仿真程序.该程序采用径向基函数(RBF)交换两套求解器之间的数据并进行网格变形.为提高RBF方法的效率,构造了基于多次插值的空间待插值点精简算法.在多次插值过程中,每次插值的对象为上次插值的误差,并同时限制插值区域,以此实现了空间待插值网格数的精简.数个网格变形的算例表明该方法可支持大变形运动,并且具有较高的计算效率.采用此程序开展了AGARD 445.6机翼颤振计算、大展弦比机翼的静气动弹性计算与切尖三角翼极限环振荡(LCO)现象的动气动弹性仿真,结果揭示了当机翼展弦比较大或者响应幅值较大时,结构非线性对于气动弹性有显著影响.     

基于CFD/CSD时域耦合方法的多通道叶栅颤振分析

为了对叶轮机多通道叶栅颤振进行研究分析,基于时域计算流体力学(CFD)方法,耦合简化的结构动力学方程,运用杂交的线性多步方法实现了对多通道叶栅的时域耦合颤振仿真.该耦合方法通过分析叶片的动力响应特性,获得叶栅的气动弹性稳定性特征.以标准叶栅颤振模型——STCF4 (Standard Test Configuration 4)为算例,计算结果表明该耦合方法与经典非耦合方法在对失稳区域的预测结果吻合得很好,证实了该耦合方法的正确性.运用该耦合方法计算了不同通道数下的自由响应,对比发现真实的叶轮机颤振与外流机翼颤振失稳的不同之处在于叶轮机叶排颤振失稳在很大程度上是一个多模态同时失稳问题;且由于出现不稳定的分支频率和阻尼都接近,叶片的时域响应会出现典型的葫芦节型发散.     

Static aeroelastic analysis of very flexible wings based on non-planar vortex lattice method

A rapid and efficient method for static aeroelastic analysis of a flexible slender wing when considering the structural geometric nonlinearity has been developed in this paper. A non-planar vortex lattice method herein is used to compute the non-planar aerodynamics of flexible wings with large deformation. The finite element method is introduced for structural nonlinear statics analysis. The surface spline method is used for structure/aerodynamics coupling. The static aeroelastic characteristics of the wind tunnel model of a flexible wing are studied by the nonlinear method presented, and the nonlinear method is also evaluated by comparing the results with those obtained from two other methods and the wind tunnel test. The results indicate that the traditional linear method of static aeroelastic analysis is not applicable for cases with large deformation because it produces results that are not realistic. However, the nonlinear methodology, which involves combining the structure finite element method with the non-planar vortex lattice method, could be used to solve the aeroelastic deformation with considerable accuracy, which is in fair agreement with the test results. Moreover, the nonlinear finite element method could consider complex structures. The non-planar vortex lattice method has advantages in both the computational accuracy and efficiency. Consequently, the nonlinear method presented is suitable for the rapid and efficient analysis requirements of engineering practice. It could be used in the preliminary stage and also in the detailed stage of aircraft design.     

Investigation of rotor control system loads

This paper concentrates on the aeroelasticity analysis of rotor blade and rotor control systems. A new multi-body dynamics model is established to predict both rotor pitch link loads and swashplate servo loads. Two helicopter rotors of UH-60A and SA349/2, both operating in two critical flight conditions, high-speed flight and high-thrust flight, are studied. The analysis shows good agreements with the flight test data and the calculation results using CAMRAD II. The mechanisms of rotor control loads are then analyzed in details based on the present predictions and the flight test data. In high-speed conditions, the pitch link loads are dominated by the integral of blade pitching moments, which are generated by cyclic pitch control. In high-thrust conditions, the positive pitching loads in the advancing side are caused by high collective pitch angle, and dynamic stall in the retreating side excites high-frequency responses. The swashplate servo loads are predominated by the rotor pitch link loads, and the inertia of the swashplate has significant effects on high-frequency harmonics of the servo loads.     

非线性气动弹性模型参考自适应控制

 首先将含有前/后缘双控制面二元机翼的动态方程以状态空间形式描述,然后考虑俯仰方向的迟滞非线性模型存在参数不确定性的情况下,利用Lyapunov稳定性理论进行了结构化模型参考自适应控制律设计。仿真结果显示：所设计的控制律能够使开环不稳定的气动弹性系统快速地达到稳定状态。由于最大控制面偏转的存在,来流速度较高时闭环系统仍会发生颤振;根据控制面最大偏转的不同取值,文中给出了闭环临界颤振速度的变化曲线。     

复合材料大展弦比机翼动力学建模与颤振分析

新一代航空结构广泛采用复合材料，对复合材料机翼的气动弹性工程化建模和分析是飞机设计的重要任务。应用气动弹性分析理论和方法，对复合材料大展弦比机翼进行了结构有限元建模、模型修正、固有振动特性计算、部件发散与颤振工程分析。本文使用MSC／NASTRAN软件，在复合材料大展弦比机翼的初步静力分析模型基础上，依据结构图纸、相关试验结果反复修改得到合理的机翼结构动力学有限元模型，固有振动计算中采用动力减缩方法消除局部模态并提高计算精度，采用亚音速偶极子格网法求解非定常气动力，并对单独机翼进行了发散和颤振计算分析。     

一种面向工程的气动弹性剪裁技术

以COMPASS为平台研究了蒙皮铺层主方向角对机翼气动弹性的影响,以及采用不同形式的多项式函数表示蒙皮厚度的方法,分析了设计变量的敏度计算方法,并运用优化的方法对两个复材机翼结构进行了位移、频率、发散和颤振速度约束下的剪裁设计,获得了结构响应与主方向角的变化曲线和机翼表面的厚度分布等高线.研究结果表明本方法拓宽了复材机翼的设计手段,具有广阔的应用前景.     

基于非定常气动力辨识技术的气动弹性数值模拟

选择离散型输入输出差分模型,运用最小二乘方法进行非定常气动力建模,并将辨识得到的降阶模型用于气动弹性的数值模拟。1个马赫数下的颤振临界点的计算仅需调用一次非定常流场求解器。计算精度保持与非定常欧拉方程计算方法相当的同时计算效率提高了1~2个量级。计算了跨声速具有S型颤振边界的气动弹性标准算例－Isogaiwing和三维气动弹性标模算例AGARD445.6,辨识模型计算边界与非定常Euler方程计算结果吻合。证明非定常气动力辨识技术可以提供高效的高精度的气动弹性分析。     

基于非线性试验气动力的飞机静气动弹性响应分析

分别使用线性和非线性静气动弹性分析方法,对某飞机纵向静气动弹性响应特性随飞行动压、攻角、平尾偏度和纵向过载变化的趋势进行了分析,并将分析结果和飞行试验进行了比较。线性方法的气动力计算使用亚音速偶极子格网法。非线性方法由风洞试验提供刚体气动力,并使用线性气动力影响系数矩阵对其进行弹性化处理。通过对两种方法的计算结果进行对比分析可以看出:①使用刚体试验气动力的非线性分析方法能够获得和飞行试验比较一致的结果;②在线性方法所提供的结果中,翼面的弯曲变形及剪力、弯矩和扭矩等部分参数能为初步设计提供大致参考,但气动力系数的弹性增量、翼面的扭转变形、翼面的剪力、弯矩和扭矩的弹性部分可能不能为型号设计提供正确的指导;③不论刚体气动力为线性还是非线性,气动力系数随动压变化的趋势均呈线性。