Active Aerothermoelastic Control of Hypersonic Double-wedge Lifting Surface

Designing re-entry space vehicles and high-speed aircraft requires special attention to the nonlinear thermoelastic and aerodynamic instability of their structural components. The thermal effects are important since temperature environment brings dramatic influences on the static and dynamic behaviors of flight structures in supersonic/hypersonic regimes and is likely to cause instability, catastrophic failure and oscillations resulting in structural failure due to fatigue. In order to understand the dynamic behaviors of these ＂hot＂ structures, a double-wedge lifting surface with combining freeplay and cubic structural nonlinearities in both plunging and pitching degrees-of-freedom operating in supersonic/hypersonic flight speed regimes has been analyzed. A third order piston theory aerodynamic is used to estimate the applied nonlinear unsteady aerodynamic loads. Also considered is the loss of torsional stiffness that may be incurred by lifting surfaces subject to axial stresses induced by aerodynamic heating. The aerodynamic heating effects are estimated based on the adiabatic wall temperature due to high speed airstreams. As a recently emerging technology, the active aerothermoelastic control is aimed at providing solutions to a large number of problems involving the aeronautical/aerospace flight vehicle structures. To prevent such damaging phenomena from occurring, an application of linear and nonlinear active control methods on both flutter boundary and post-flutter behavior has been fulfilled. In this paper, modeling issues as well as numerical simulation have been presented and pertinent conclusions outlined. It is evidenced that a serious loss of torsional stiffness may induce the dynamic instability; however active control can be used to expand the flutter boundary and convert unstable limit cycle oscillations （LCO） into the stable LCO and/or to shift the transition between these two states toward higher flight Mach numbers.     

Aeroelastic response of a 2-D airfoil in a compressible flow field and exposed to blast loading

This paper deals with the generation and use of proper aerodynamic indicial functions toward the aeroelastic formulation of two-dimensional lifting surfaces in the subsonic compressible, linearized transonic, supersonic and hypersonic flight speed regimes. The indicial function approach enables one to treat in an unified way (i.e. in the time and frequency domains) the subcritical aeroelastic response and the flutter instability of lifting surfaces. Validations of the aerodynamic model are documented and excellent agreements are reported. In addition, closed form solutions and aerodynamic derivatives for different flight speed regimes are obtained; comparisons, and results displaying the aeroelastic response to blast loads are presented, and pertinent conclusions are outlined.     

Adaptive aeroelastic vibration suppression of a supersonic airfoil with flap

This paper concerns the flutter, post-flutter and adaptive control of a non-linear 2-D wing-flap system operating in supersonic/hypersonic flight speed regimes. An output feedback control law is implemented and its performance toward suppressing flutter and limit cycle oscillations (LCOs) as well as reducing the vibrational level in the subcritical flight speed range is demonstrated. This control law is applicable to minimum phase systems and we provide conditions for stability of the zero dynamics. The control objective is to design a control strategy to stabilize the pitch angle while adaptively compensating for uncertainties in all the aeroelastic model parameters. It is shown that all the states of the closed-loop system are asymptotically stable.     

Nonlinear aeroelastic analysis of the folding fin with freeplay under thermal environment

The nonlinear aeroelastic behavior of a folding fin in supersonic flow is investigated in this paper. The finite element model of the fin is established and the deployable hinges are represented by three torsion springs with the freeplay nonlinearity. The aerodynamic grid point is assumed to be at the center of each aerodynamic box for simplicity. The aerodynamic governing equation is given by using the infinite plate spline method and the modified linear piston theory. An improved fixed-interface modal synthesis method, which can reduce the rigid connections at the interface, is developed to save the problem size and computation time. The uniform temperature field is applied to create the thermal environment. For the linear flutter analyses, the flutter speed increases first and then decreases with the rise of the hinge stiffness due to the change of the flutter coupling mechanism. For the nonlinear analyses, a larger freeplay angle results in a higher vibration divergent speed. Two different types of limit cycle oscillations and a multiperiodic motion are observed in the wide range of airspeed under the linear flutter boundary. The linear flutter speed shows a slight descend in the thermal environment, but the effect of the temperature on the vibration divergent speed is different under different hinge stiffnesses when there exists freeplay.     

Nonlinear aeroelastic analysis of airfoils: bifurcation and chaos

Different types of structural and aerodynamic nonlinearities commonly encountered in aeronautical engineering are discussed. The equations of motion of a two-dimensional airfoil oscillating in pitch and plunge are derived for a structural nonlinearity using subsonic aerodynamics theory. Three classical nonlinearities, namely, cubic, freeplay and hysteresis are investigated in some detail. The governing equations are reduced to a set of ordinary differential equations suitable for numerical simulations and analytical investigation of the system stability. The onset of Hopf-bifurcation, and amplitudes and frequencies of limit cycle oscillations are investigated, with examples given for a cubic hardening spring. For various geometries of the freeplay, bifurcations and chaos are discussed via the phase plane, Poincaré maps, and Lyapunov spectrum. The route to chaos is investigated from bifurcation diagrams, and for the freeplay nonlinearity it is shown that frequency doubling is the most commonly observed route. Examples of aerodynamic nonlinearities arising from transonic flow and dynamic stall are discussed, and special attention is paid to numerical simulation results for dynamic stall using a time-synthesized method for the unsteady aerodynamics. The assumption of uniform flow is usually not met in practice since perturbations in velocities are encountered in flight. Longitudinal atmospheric turbulence is introduced to show its effect on both the flutter boundary and the onset of Hopf-bifurcation for a cubic restoring force.     

Match Point Solution for Robust Flutter Analysis in Constant-Mach Prediction

This paper presents a method for robust flutter computation which uses flight altitude as the perturbation variable in order to obtain a match point solution. The air density and sound speed of standard atmosphere model are approximated as the polynomial function of altitude, such that the flight altitude becomes the single perturbation variable that describes the aeroelastic system. The uncertainties of generalized stiffness and damping are considered and the uncertain aeroelastic system can be formulated as linear fractional transformation （LFT） representation which is suitable for/.t analysis framework. Finally, the match point solutions of robust flutter margins can be computed with structured singular value （SSV） theory. The robust flutter analysis method provided in this paper is suitable for constant-Mach flight flutter test and provides valuable reference for flight envelope expansion.     

带有气动及结构非线性的二元机翼颤振分析

研究了翼型在低马赫数条件下的非定常气动特性,从翼型表面气流运动的角度对Leishman-Beddoes(L-B)模型进行了修正,并在此基础上建立了适合低马赫数颤振研究且带有气动及结构非线性的二元机翼气弹系统分析模型.对比低马赫数翼型气动载荷试验结果表明对L-B模型的修正是有效的,且机翼颤振试验结果亦验证了二元机翼气弹分析模型.研究结果表明:二元机翼气弹系统的失速颤振与初始变距角和来流速度密切相关,且耦合的三次非线性变距和浮沉刚度是造成系统呈现准周期运动的主要原因.      

二维非定常欧拉方程的摄动解在超声速矩形翼上的应用

本文利用超声速线化理论给出的矩形翼载荷系数解析解作为权函数,对沉浮和俯仰谐振的二维平板翼压强系数进行加权修正,获得矩形翼谐振气动力的近似解,计算结果与线化理论数值结果符合得很好,本方法是超声速和高超声速矩形翼非定常气动力的一种统一和快速的计算方法。     

Nonlinear Aeroelastic Response of High-aspect-ratio Flexible Wings

The aeroelastic analysis of high-altitude, long-endurance (HALE) aircraft that features high-aspect-ratio flexible wings needs take into account structural geometrical nonlinearities and dynamic stall. For a generic nonlinear aeroelastic system, besides the stability boundary, the characteristics of the limit-cycle oscillation (LCO) should also be accurately predicted. In order to conduct nonlinear aeroelastic analysis of high-aspect-ratio flexible wings, a first-order, state-space model is developed by combining a geometrically exact, nonlinear anisotropic beam model with nonlinear ONERA (Edlin) dynamic stall model. The present investigations focus on the initiation and sustaining mechanism of the LCO and the effects of flight speed and drag on aeroelastic behaviors. Numerical results indicate that structural geometrical nonlinearities could lead to the LCO without stall occurring. As flight speed increases, dynamic stall becomes dominant and the LCO increasingly complicated. Drag could be negligible for LCO type, but should be considered to exactly predict the onset speed of flutter or LCO of high-aspect-ratio flexible wings.     

考虑间隙非线性的控制舵非线性气动弹性分析

间隙结构的气动弹性系统非线性颤振问题是飞行器气动弹性力学工程领域的研究热点和难点,研究 考虑间隙非线性的控制舵系统的气动弹性特性具有重要意义。基于最小状态拟合方法获得时域降阶气动力模 型,并通过Lagrange方程获得系统非线性气动弹性方程;对比分析三种不同非线性控制舵系统的极限环颤振 及非线性动力学响应特性,并与等效线化法和时域仿真的结果进行一致性对比。结果表明:俯仰和扑动弹簧刚 度的变化对系统颤振边界有显著影响,当俯仰和扑动两个方向同时含有间隙非线性时,系统在线性颤振速度内 存在倍周期、混沌等复杂非线性动力学现象。     

低速流场中柔性悬臂板的后颤振响应

建立了一个新的非线性气动弹性模型,对低速流场中柔性悬臂板的后颤振响应特性进行了分析。建模中考虑了结构几何非线性、气动力非线性以及两者之间的强耦合效应。通过实验数据对所建立的气动弹性模型进行了验证。发现在低速流场中柔性悬臂板可能会以周期加倍的方式进入混沌运动。结构几何非线性效应和翼尖涡引起的非定常气动力效应对柔性悬臂板的结构响应有显著影响,而尾涡变形引起的非定常气动力对结构运动的影响较小。还研究了不同耦合算法的差异,给出了小展弦比大柔性结构非线性气动弹性数值仿真时耦合策略的选择依据。     

考虑机翼几何非线性的气动弹性建模与分析

大展弦比柔性机翼结构重量轻、气动效率高,广泛应用于高空长航时无人机(UAVs)。飞行过程中,这类机翼在气动力作用下发生大变形,线性结构模型不再适用,需要建立考虑几何大变形的结构模型。采用牛顿力学方法推导了考虑结构几何非线性的机翼结构动力学模型,该方法推导过程简洁、物理意义明确,可以与Hodges基于哈密顿原理的推导方法相互补充,相互验证。为了能够更准确地求解大展弦比柔性机翼的非定常气动力,建立了能够考虑机翼三维效应且适用于机翼空间大变形的非定常气动力模型。基于建立的非线性结构模型和非定常气动力模型,采用松耦合方法建立了非线性气动弹性模型,并通过算例验证了气弹模型的准确性。研究结果表明,大展弦比柔性机翼颤振速度对来流迎角和机翼的展长均较为敏感;当来流速度大于颤振速度时,由于几何非线性,机翼振动并未发散而是形成稳定的极限环振荡(LCO);随着来流速度进一步增加,机翼再次穿过临界稳定点,由不稳定系统变为稳定系统,直到随着速度的增加系统再次达到临界稳定状态。     

压电驱动器的气动弹性应用

 随着压电智能材料与结构的发展,压电驱动器在气动弹性控制领域占据重要地位。使用压电驱动器控制翼面变形,利用而不是抵抗气动弹性效应可以控制升力、力矩以及它们的分布。采用基本相同的智能结构翼面控制系统,根据不同的控制目标需求,使用压电智能材料驱动器可以达到多种目的,包括静态的形状控制与动态的颤振抑制、抖振控制与阵风响应控制。静态控制方面例如改变翼面形状获得附加空气动力以增加升力、提供横滚力矩、改变升力分布以减小诱导阻力或减小翼根弯矩等;动态控制例如利用改变翼面形状产生的附加空气动力作为控制载荷,改变气动弹性系统的耦合程度,根据控制效果要求可作为气动阻尼、气动刚度或气动质量。这种控制方法可以减轻结构重量,提高操纵效率,扩大飞行包线,提高材料利用率,已成为可变形飞行器的重要研究内容。本文主要阐述压电驱动器气动弹性应用的动机与机理、发展与成就以及问题与展望。     

一种基于亚临界响应的颤振稳定性边界预测新方法

基于地面模态试验事先获取的结构模态参数,通过弹性机翼在气流中的亚临界响应,解算出作用在机翼上的非定常模态气动力系数。以弹性机翼的模态位移作为系统输入,模态气动力系数作为系统输出,通过辨识方法获得弹性机翼振动的气动力模型。在时域内耦合结构运动方程和气动力模型,建立基于试验数据的气动弹性稳定性分析模型。通过分析系统稳定性随动压的变化规律,获得弹性机翼的颤振稳定性特性。与经典颤振边界外推方法的主要区别在于该方法实质上只需要一次亚临界响应试验即可预测颤振临界点,可极大降低颤振试验的风险和成本。该方法既可用于颤振风洞试验,也可用于颤振试飞。     

压气机转子叶片颤振边界的预测方法

通过发展的基于能量法的颤振数值预测方法得到了压气机转子叶片的颤振边界.将计算结构动力学(CSD)分析得到的叶片表面节点位移插值到耦合面的流体网格点上,并将设计的多层动网格技术应用于计算流体动力学(CFD)方法,实现叶栅振荡作用下的非定常分析,得到叶片表面的非定常气动功以及模态气动阻尼比.以某第一级压气机转子叶片为例,对...     

不同振动形式下的翼型失速特性

通过求解雷诺平均Navier-Stokes方程,研究了翼型在强迫振动和自然结构振动下的大迎角流场特性,尤其是失速迎角附近的流场和气动特性.研究结果表明:在接近颤振临界速度情况下结构自然振动可以引起翼型大尺度的分离,导致失速分离涡提前出现;强迫性的沉浮运动和俯仰运动在一定幅度下也可以引起失速性质的大分离,而且沉浮和俯仰振动的频率和振幅都是影响翼型大尺度分离的重要因素.     

The flutter of an airfoil with cubic structural and aerodynamic non-linearities

The flutter of a two-dimensional airfoil in a supersonic flow field, with cubic structural and aerodynamic non-linearities, is investigated using an efficient algorithm of normal form, which combines the normal form theory and the center manifold theory together. First, the stability of the linearized system is analyzed in the neighborhood of an equilibrium point, which shows that the flutter instability is resulted by the Hopf bifurcation. Then the normal form of Hopf bifurcation is deduced by applying the symbolic procedure of the new normal form algorithm to the perturbation equations. Analyzing the obtained coefficients of normal form shows that for a given system, the Hopf bifurcation can change from super-critical type to sub-critical type, consequently the flutter instability changes from “benign” type to “catastrophic” type, as the flight Mach number increases. Numerical simulations verify the dependence of response on initial conditions. Finally, the effects of the structural and aerodynamic parameters on the character of flutter instability are analyzed.     

基于谐波平衡法和 V-g 法的高效颤振预测分析

提出了一种高效的颤振预测分析方法,将之应用于跨声速颤振边界分析及结构参数影响研究中。本方法基于频域颤振分析 V-g 方法,为气弹系统提供一定的人工阻尼使之保持简谐运动状态,从而将结构动力学方程转换到频域内。然后通过高效的谐波平衡法得到一系列简谐运动频率下的气动力描述函数矩阵,结合频域结构方程,将气弹系统的稳定性问题转化为广义特征值求解。结果表明：本方法计算得到的颤振边界与高精度的时间推进方法非常吻合,分析效率有了明显的提升,而且当结构参数发生变化后,只需进行若干次广义特征值求解即可得到新的颤振边界,无需像时间推进方法一样开展大量的气动/结构耦合数值模拟。     

叶轮机超音速非失速颤振边界研究

把叶片简化为具有对称截面、长、直、带预扭的非均匀弹性梁 ,并结合超音速状态下计算由振动引起的非定常气动载荷的气动模型 ,用哈密尔顿原理建立一组旋转状态下失调叶栅气弹稳定性方程 ;这组方程对弯曲—弯曲—扭转之间耦合严重的失调叶栅尤有意义。对一个风扇级进行的气弹稳定性分析结果表明 :失调及其幅度、阶次 ,扭心位置 ,耦合效应及其程度对叶栅气弹稳定性有明显的影响     

Nonlinear flutter wind tunnel test and numerical analysis of folding fins with freeplay nonlinearities

The flutter characteristics of folding control fins with freeplay are investigated by numerical simulation and flutter wind tunnel tests. Based on the characteristics of the structures, fins with different freeplay angles are designed. For a 0° angle of attack, wind tunnel tests of these fins are conducted, and vibration is observed by accelerometers and a high-speed camera. By the expansion of the connected relationships, the governing equations of fit for the nonlinear aeroelastic analysis are established by the free-interface component mode synthesis method. Based on the results of the wind tunnel tests, the flutter characteristics of fins with different freeplay angles are analyzed. The results show that the vibration divergent speed is increased, and the divergent speed is higher than the flutter speed of the nominal linear system. The vibration divergent speed is increased along with an increase in the freeplay angle. The developed free-interface component mode synthesis method could be used to establish governing equations and to analyze the characteristics of nonlinear aeroelastic systems. The results of the numerical simulations and the wind tunnel tests indicate the same trends and critical velocities.