铺层参考方向选取对大展弦比机翼静气弹特性影响

以大展弦比机翼为研究对象，利用流固耦合方法对复合材料机翼铺层参考方向进行了数值模拟研究，分析了铺层参考方向轴偏角的改变对大展弦比机翼静气动弹性的影响．研究表明：铺层参考方向轴偏角的改变会对机翼气动弹性产生显著的影响．机翼的总体变形与扭转变形随着参考方向轴偏角的改变呈现周期分布；沿着机翼各个方向的挠度也会因为参考方向轴偏角的改变而产生不同的响应．     

超音速气流中受热曲壁板的非线性颤振特性

基于von Karman 大变形理论及带有曲率修正的一阶活塞理论, 用Galerkin方法建立了超音速气流中受热二维曲壁板的非线性气动弹性运动方程; 采用牛顿迭代法计算得到由静气动载荷和热载荷引起的静气动弹性变形; 根据李雅谱诺夫间接法分析了壁板初始曲率与温升对颤振边界的影响; 对二维曲壁板的非线性气动弹性方程组进行数值积分求解,分析了动压参数对受热二维曲壁板分岔特性的影响, 给出了典型状态下曲壁板非线性颤振响应的时程图与相图. 分析结果表明对小初始曲率的曲壁板, 温升对其静气动弹性变形影响较大, 且随着温升的增加其颤振临界动压急剧减小; 对具有较大初始曲率的曲壁板, 温升对其静气动弹性变形的影响较弱, 且随着温升的增加颤振临界动压基本保持不变. 初始几何曲率与气动热效应使得曲壁板具有复杂的动力学特性, 不再像平壁板一样, 经过倍周期分岔进入混沌, 而会出现由静变形状态直接进入混沌运动的现象, 且在混沌运动区域中还会出现静态稳定点或谐波运动, 在大曲率情况下, 曲壁板不会产生混沌运动, 而是幅值在一定范围内的极限带振荡.      

基于POD和代理模型的静气动弹性分析方法

静气动弹性问题考虑弹性结构与定常气动力间的相互耦合作用,对飞行器的性能和安全具有显著的影响.在现代飞行器设计阶段,计算流体力学(CFD)/计算结构力学(CSD)直接耦合方法是精确考察静气动弹性影响的重要手段.然而,基于CFD技术的气动力仿真手段在耦合过程中计算量大且耗时长,难以满足设计阶段的需求.因此,为了兼顾计算精度与效率,文章采用本征正交分解(POD)和Kriging代理模型相结合的模型降阶方法,替代CFD求解过程并耦合有限元分析(FEA)方法,建立了高效、准确的静气动弹性分析框架.相较于传统的以模态法为主的静气动弹性分析方法,该方法能够解决更为复杂的静气动弹性问题以及提供静气动弹性变形过程中的气动分布载荷.针对典型三维跨声速HIRENASD机翼模型开展的马赫数、迎角变化的算例验证表明:由建立的静气动弹性分析方法与CFD/CSD直接耦合方法计算得到机翼翼梢处的静变形量间的相对误差在5%以内;同时该方法预测静平衡位置处的气动分布载荷的误差在5%以内,静气动弹性分析的计算效率至少提升了6倍.     

静气动弹性计算方法研究

对基于结构网格的Euler方程及N-S方程求解器和基于非结构网格的Euler方程求解器,采用结构模态分析方法和柔度矩阵方法,对无人机大展弦比机翼在Ma=0.6,α=2?, 飞行高度20km的巡航状态下的静气动弹性特性进行了数值模拟. 验证了两种求解器对静气动弹性模拟的准确性. 同时,对模态分析方法和柔度矩阵方法进行了对比研究,发现柔度矩阵方法更适用于静气动弹性数值模拟. 另外,对应用物面法向偏转方法替代网格变形技术模拟静气动弹性进行了研究,计算表明物面法向偏转方法可以大大提高静气动弹性计算效率和克服机翼结构变形过大时动网格技术无法处理的不足.      

西南某大型水电站拱肩槽边坡开挖的变形响应研究

对基于结构网格的Euler方程及N-S方程求解器和基于非结构网格的Euler方程求解 器,采用结构模态分析方法和柔度矩阵方法,对无人机大展弦比机翼在Ma=0.6, α=2?, 飞行高度20km的巡航状态下的静气动弹性特性进行了数值模 拟. 验证了两种求解器对静气动弹性模拟的准确性. 同时,对模态分析方法和柔度 矩阵方法进行了对比研究,发现柔度矩阵方法更适用于静气动弹性数值模拟. 另外, 对应用物面法向偏转方法替代网格变形技术模拟静气动弹性进行了研究,计算表明 物面法向偏转方法可以大大提高静气动弹性计算效 率和克服机翼结构变形过大时动网格技术无法处理的不足.     

复合材料机翼/外挂系统颤振特性分析

综合考虑复合材料机翼的材料耦合和几何耦合的特点,利用Hamilton原理建立了非定常气动力作用下复合材料机翼/外挂系统的气动弹性运动方程。采用微分求积法进行离散,运用MATLAB语言编程对系统响应进行数值仿真。计算结果表明:用微分求积法计算复合材料机翼的颤振特性是可行的;随材料耦合刚度的增加,机翼的颤振临界速度出现了双峰值;而对于复合材料机翼/外挂系统,颤振临界速度随材料耦合刚度的变化规律较为复杂,与外挂参数的设置相关。     

高速小展弦比机翼颤振的微分求积法分析

引入微分求积法,分析高速小展弦比机翼的气动弹性问题。将小展弦比机翼等效为悬臂板,基于一阶活塞气动力理论建立机翼颤振偏微分方程,采用微分求积法将偏微分方程转化为常微分方程,根据频率重合理论对颤振问题进行求解。分析了机翼的固有频率及颤振速度,并与有限元软件计算结果进行比较,误差在2％以内,很好的验证了微分求积法求解小展弦比机翼颤振问题的有效性。分析了机翼面积、展弦比及厚度对颤振速度的影响,结果表明,小展弦比机翼的颤振速度受结构尺寸的影响较大,颤振速度随面积和展弦比的增大而减小,随机翼厚度的增大而增大。     

热过屈曲功能梯度壁板的气动弹性颤振

功能梯度材料的宏观物理性能随空间位置连续变化,能充分减少不同组份材料结合部位界面性能的不匹配因素.功能梯度壁板用作高速飞行器的热防护结构,能有效消除气动加热带来的壁板内部热应力集中.本文考虑热过屈曲变形引入的结构几何非线性,分析功能梯度壁板的气动弹性颤振边界.基于幂函数材料分布假设,采用混合定律计算功能梯度材料的等效力学性能.根据一阶剪切变形板理论、冯·卡门应变-位移关系和一阶活塞理论,基于虚功原理建立超声速气流中受热功能梯度壁板的非线性气动弹性有限元方程.采用牛顿-拉弗森迭代法数值求解壁板的热屈曲变形,分析超声速气流对热屈曲变形的影响机理.在壁板热过屈曲的静力平衡位置分析动态稳定性,确定了壁板的颤振边界.研究表明,当陶瓷-金属功能梯度壁板的组份材料沿厚度方向梯度分布时,会破坏结构的对称性导致壁板在面内热应力作用下发生指向金属侧的热屈曲变形.超声速气流中壁板热屈曲变形最大的位置随气流速压增大向下游推移,并伴随屈曲变形量的减小.热过屈曲壁板的几何非线性效应会提高壁板的颤振边界,这种影响在高温、低无量纲速压且壁板发生大挠度热屈曲变形时表现显著.较高无量纲气流速压下由于壁板的热屈曲变形被气动力限定在小挠度范围,几何非线性效应不明显.      

含间隙非线性的大展弦比机翼/外挂系统的气弹响应分析

综合考虑大展弦比机翼的柔性变形和外挂与机翼连接处的间隙非线性,建立了机翼/外挂系统的气动弹性力学模型。采用非定常气动力,根据Lagrange方程推导了大展弦比机翼/外挂系统的运动微分方程。运用伽辽金法进行了离散,通过数值模拟研究了系统的气动弹性响应及其稳定性。结果表明:系统极限环振动的临界速度随间隙的增大而减小,随间隙初偏值的增加而增大;几何非线性可大幅降低系统极限环振动的幅值;随流速的增加,系统呈现出复杂的响应,周期运动与混沌运动相间出现;最后系统发生屈曲。     

适用于几何非线性气动弹性分析的结构动力学降阶模型方法研究

几何非线性是壁板颤振和大展弦比机翼气动弹性等问题的一个主要特征，在进行数值仿真分析时往往需要采用商业非线性有限元求解器，存在计算量大和耦合迭代策略不易控制等问题。本文发展了一种适用于几何非线性的结构动力学降阶模型(CSD-ROM),利用广义坐标的非线性多项式表征非线性内力，采用参数识别方法获取多项式系数，并通过增加额外的线性模态来改善模型预测精度。基于此方法，分别针对壁板颤振、切尖三角翼的CFD/CSD-ROM非线性颤振问题开展了时域响应分析。计算结果表明，通过CSD-ROM计算出的壁板颤振速度为590 m/s,颤振频率为174 Hz,与有限元结果误差分别为0.8%和1.7%。马赫数0.879时切尖三角翼的颤振动压预测结果为2.25 psi,与非线性有限元相比的误差为3.8%。本文采用的非线性和线性模态基底组合方法，在保证计算精度的基础上可有效降低训练样本数量，一定程度上可替代非线性有限元开展气动弹性分析。     

Aeroelastic stability of a cantilevered plate in yawed subsonic flow

The aeroelastic stability of cantilevered plates with their clamped edge oriented both parallel and normal to subsonic flow is a classical fluid–structure interaction problem. When the clamped edge is parallel to the flow the system loses stability in a coupled bending and torsion motion known as wing flutter. When the clamped edge is normal to the flow the instability is exclusively bending and is referred to as flapping flag flutter. This paper explores the stability of plates during the transition between these classic aeroelastic configurations. The aeroelastic model couples a classical beam structural model to a three-dimensional vortex lattice aerodynamic model. The aeroelastic stability is evaluated in the frequency domain and the flutter boundary is presented as the plate is rotated from the flapping flag to the wing configuration. The transition between the flag-like and wing-like instability is often abrupt and the yaw angle of the flow for the transition is dependent on the relative spacing of the first torsion and second bending natural frequencies. This paper also includes ground vibration and aeroelastic experiments carried out in the Duke University Wind Tunnel that confirm the theoretical predictions.     

Reduced order nonlinear aeroelasticity of swept composite wings using compressible indicial unsteady aerodynamics

Nonlinear dynamic aeroelasticity of composite wings in compressible flows is investigated. To provide a reasonable model for the problem, the composite wing is modeled as a thin walled beam (TWB) with circumferentially asymmetric stiffness layup configuration. The structural model considers nonlinear strain displacement relations and a number of non-classical effects, such as transverse shear and warping inhibition. Geometrically nonlinear terms of up to third order are retained in the formulation. Unsteady aerodynamic loads are calculated according to a compressible model, described by indicial function approximations in the time domain. The aeroelastic system of equations is augmented by the differential equations governing the aerodynamics lag states to derive the final explicit form of the coupled fluid-structure equations of motion. The final nonlinear governing aeroelastic system of equations is solved using the eigenvectors of the linear structural equations of motion to approximate the spatial variation of the corresponding degrees of freedom in the Ritz solution method. Direct time integrations of the nonlinear equations of motion representing the full aeroelastic system are conducted using the well-known Runge–Kutta method. A comprehensive insight is provided over the effect of parameters such as the lamination fiber angle and the sweep angle on the stability margins and the limit cycle oscillation behavior of the system. Integration of the interpolation method employed for the evaluation of compressible indicial functions at any Mach number in the subsonic compressible range to the derivation process of the third order nonlinear aeroelastic system of equations based on TWB theory is done for the first time. Results show that flutter speeds obtained by the incompressible unsteady aerodynamics are not conservative and as the backward sweep angle of the wing is increased, post-flutter aeroelastic response of the wing becomes more well-behaved.     

LCO flutter of cantilevered woven glass/epoxy laminate in subsonic flow

The paper presents a cantilevered composite wing, aeroelastic characteristics of idealized as a composite flat plate laminate. The composite laminate was made from woven glass fibers with epoxy matrix. The elastic and dynamic properties of the laminate were determined experimentally for aeroelastic calculations. Aeroelastic wind tunnel testing of the laminate was performed and the result showed that flutter, a dynamic instability occurred. The cantilevered laminate also displayed limit cycle amplitude, post-flutter oscillation. The experimental flutter velocity and frequency were verified by our computational analysis.     

EFFECTS OF STEADY ANGLE OF ATTACK ON NONLINEAR GUST RESPONSE OF A DELTA WING MODEL

The effects of a steady angle of attack on the nonlinear aeroelastic response of a delta wing model to a periodic gust have been studied. For the theoretical analysis, a three-dimensional time-domain vortex lattice aerodynamic model and a reduced order aerodynamic technique were used and the structure was modelled using von Karman plate theory that allows for geometric strain–displacement nonlinearities in the delta wing structure. Also, an experimental investigation has been carried out in the Duke wind tunnel using a rotating slotted cylinder gust generator and an Ometron VPI 4000 Scanning Laser Vibrometer measurement system to measure deflections (velocities) of a delta wing test model. The fair to good quantitative agreement between theory and experiment verifies that the present analytical approach has reasonable accuracy and good computational efficiency for nonlinear gust response analysis in the time-domain. The results also contribute to a better physical understanding of the nonlinear aeroelastic response of a delta wing model to gust loads when the steady angle of attack is varied.     

A theoretical and experimental investigation of the effects of a steady angle of attack on the nonlinear flutter of a delta wing plate model

Limit cycle oscillations (LCO) of wings on certain modern high performance aircraft have been observed in flight and in wind tunnel experiments. Whether the physical mechanism that gives rise to this behavior is a fluid or structural nonlinearity or both is still uncertain. It has been shown that an aeroelastic theoretical model with only a structural nonlinearity can predict accurately the limit cycle behavior at low subsonic flow for a plate-like wing at zero angle of attack. Changes in the limit cycle and flutter behavior as the angle of attack is varied have also been observed in flight. It has been suggested that this sensitivity to angle of attack is due to a fluid nonlinearity. In this investigation, we study the flutter and limit cycle behavior of a wing in low subsonic flow at small steady angles of attack. Experimental results are compared to those predicted using an aeroelastic theoretical model with only a structural nonlinearity. Results from both experiment and theory show a change in flutter speed as the steady angle of attack is varied. Also the LCO magnitude increased at a given velocity as the angle of attack was increased for both the experiment and theory. While not proving that the observed sensitivity to angle of attack of LCO in aircraft is due to a structural nonlinearity, the results do show that a change in the aeroelastic behavior at angles of attack can be caused by a structural nonlinearity as well as a fluid nonlinearity. In this paper, only structural nonlinearities are considered, but an extension to include aerodynamic nonlinearities would be very worthwhile.     

Aeroelastic analysis of large horizontal wind turbine baldes

A nonlinear aeroelastic analysis method for large horizontal wind turbines is described. A vortex wake method and a nonlinear finite element method (FEM) are coupled in the approach. The vortex wake method is used to predict wind turbine aerodynamic loads of a wind turbine, and a three-dimensional (3D) shell model is built for the rotor. Average aerodynamic forces along the azimuth are applied to the structural model, and the nonlinear static aeroelastic behaviors are computed. The wind rotor modes are obtained at the static aeroelastic status by linearizing the coupled equations. The static aeroelastic performance and dynamic aeroelastic responses are calculated for the NH1500 wind turbine. The results show that structural geometrical nonlinearities significantly reduce displacements and vibration amplitudes of the wind turbine blades. Therefore, structural geometrical nonlinearities cannot be neglected both in the static aeroelastic analysis and dynamic aeroelastic analysis.     

Aeroelastic analysis and active flutter control of nonlinear lattice sandwich beams

Nonlinear aeroelastic characteristics of sandwich beams with pyramidal lattice core are investigated, and the active flutter control of the nonlinear structural system is also studied using the piezoelectric actuator/sensor pair. In the structural modeling, Reddy’s third-order shear deformation theory is applied. Aerodynamic pressure is evaluated by the supersonic piston theory. Hamilton’s principle and the assumed mode method are used to derive the equation of motion. The proportional feedback and the optimal H _∞ control methods are performed to design the controller. In the robust control, the uncertainty caused by omitting the nonlinear terms of the control equation is taken into account, and the mixed sensitivity method is used to solve the problem. The nonlinear aeroelastic property of the sandwich beam is analyzed and is compared with that of the equivalent isotropic beam with the same weight to show the superior aeroelastic characteristics of the lattice sandwich beam. Controlled vibration responses under the two different controllers are calculated and compared. Simulation results show that the robust controller is much more effective than the proportional feedback controller in the flutter suppression of the nonlinear sandwich beam.