Effect of multiple engine placement on aeroelastic trim and stability of flying wing aircraft

Effects of multiple engine placement on flutter characteristics of a backswept flying wing resembling the HORTEN IV are investigated using the code NATASHA (Nonlinear Aeroelastic Trim And Stability of HALE Aircraft). Four identical engines with defined mass, inertia, and angular momentum are placed in different locations along the span with different offsets from the elastic axis while fixing the location of the aircraft c.g. The aircraft experiences body freedom flutter along with non-oscillatory instabilities that originate from flight dynamics. Multiple engine placement increases flutter speed particularly when the engines are placed in the outboard portion of the wing (60–70% span), forward of the elastic axis, while the lift to drag ratio is affected negligibly. The behavior of the sub- and supercritical eigenvalues is studied for two cases of engine placement. NATASHA captures a hump body-freedom flutter with low frequency for the clean wing case, which disappears as the engines are placed on the wings. In neither case is there any apparent coalescence between the unstable modes. NATASHA captures other non-oscillatory unstable roots with very small amplitude, apparently originating with flight dynamics. For the clean-wing case, in the absence of aerodynamic and gravitational forces, the regions of minimum kinetic energy density for the first and third bending modes are located around 60% span. For the second mode, this kinetic energy density has local minima around the 20% and 80% span. The regions of minimum kinetic energy of these modes are in agreement with calculations that show a noticeable increase in flutter speed if engines are placed forward of the elastic axis at these regions.     

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.     

Flutter suppression of a plate-like wing via parametric excitation

The present work is motivated by the well known stabilizing effect of parametric excitation of some dynamical systems such as the inverted pendulum. The possibility of suppressing wing flutter via parametric excitation along the plane of highest rigidity in the neighborhood of combination resonance is explored. The nonlinear equations of motion in the presence of incompressible fluid flow are derived using Hamilton's principle and Theodorsen's theory for modeling aerodynamic forces. In the presence of air flow, the bending and torsion modes possess nearly the same frequency. Under parametric excitation and in the absence of air flow, each mode oscillates at its own natural frequency. In the neighborhood of combination resonance, the nonlinear response is determined using the multiple scales method at the critical flutter speed and at slightly higher airflow speed. The domains of attraction and bifurcation diagrams are obtained to reveal the conditions under which the parametric excitation can provide stabilizing effect. The basins of attraction for different values of excitation amplitude reveal the stabilizing effect that takes place above a critical excitation level. Below that level, the response experiences limit cycle oscillations, cascade of period doubling, and chaos. For flow speed slightly higher than the critical flutter speed, the response experiences a train of spikes, known as ‘firing,’ a term that is borrowed from neuroscience, followed by ‘refractory’ or recovery effect, up to an excitation level above which the wing is stabilized. The results of the multiple scales method are verified using numerical simulation of the original nonlinear differential equations.     

Flutter analysis of a viscoelastic tapered wing under bending–torsion loading

The dynamic stability of a tapered viscoelastic wing subjected to unsteady aerodynamic forces is investigated. The wing is considered as a cantilever tapered Euler–Bernoulli beam. The beam is made of a linear viscoelastic material where Kelvin–Voigt model is assumed to represent the viscoelastic behavior of the material. The governing equations of motion are derived through the extended Hamilton’s principle. The resulting partial differential equations are solved via Galerkin’s method along with the classical flutter investigation approach. The developed model is validated against the well-known Goland wing and HALE wing and good agreement is obtained. Different solution methods, namely; the k method, the p-k method, and the flutter determinant method are compared for the case of elastic wing. However, when the viscoelastic damping is introduced, the k and p-k methods become less effective. The flutter determinant method is modified and employed to carry out non-dimensional parametric study on the Goland wing. The study includes the effects of parameters such as the taper ratio, the density ratio, the viscoelastic damping of wing structure and many other parameters on the flutter speed and flutter frequency. The study reveals that a tapered wing would be more dynamically stable than a uniform wing. It is also observed that the viscoelastic damping provides wider stability region for the wing. The investigation shows that the density ratio, bending-to-torsion frequency ratio, and the radius of gyration have significant effects on the dynamic stability of the wing. Based on the obtained results, a wing with an elastic center and inertial center that are located closer to the mid-chord would be more dynamically stable.     

Distribution of the energy of a piezoelectric actuator between traveling waves excited in an elastic layer

The distribution of the energy of a piezoelectric actuator between normal modes (Lamb waves) as a function of the source parameters and frequency is studied by solving the dynamic contact problem of the interaction between a flexible piezoelectric patch and a flexible elastic substrate with explicit representations for the excited traveling waves. Zones of the maximum and minimum energy of the fundamental modes are determined in the “oscillation frequency–piezoelectric patch width” plane.     

Energy harvesting of a heaving and forward pitching wing with a passively actuated trailing edge

This study experimentally investigates the energy harvesting capabilities of an oscillating wing with a passively actuated trailing edge. The oscillation kinematics are composed of a combined heaving and forward pitching motions, where the pitching axis is well behind the wing center of mass. Passive actuation is attained by connecting the trailing edge with the wing body using a torsion rod. The degree of flexibility of the trailing edge is represented by the Strouhal number based on the trailing edge natural frequency. The trailing edge passive response is studied for oscillation Strouhal numbers of 0.017, 0.025 and 0.033. Instantaneous aerodynamic forces are measured in a closed loop wind tunnel at a Reynolds number of 40 000, based on the free stream velocity and the wing chord length. Measured results include the effective angle of attack induced by the trailing edge actuation as well as the lift and moment during the oscillation cycle. For the imposed kinematics in this study, the pitching motion has a positive contribution to the mean power output whereas the heaving motion has a relatively small but negative contribution. Additionally, by decreasing the natural frequency of the trailing edge closer to that of the imposed oscillation frequency, the magnitude of the lift and moment forces and hence the mean power output, increases. It is found that there exists a strong correlation between mean power output and the effective angle of attack, shown through the passive trailing edge response, resulting in an increase in energy harvesting potential.     

Nonlinear aeroelastic analysis of a two-dimensional wing with control surface in supersonic flow

Based on the piston theory of supersonic flow and the energy method, a two dimensional wing with a control surface in supersonic flow is theoretically modeled, in which the cubic stiffness in the torsional direction of the control surface is considered. An approximate method of the cha- otic response analysis of the nonlinear aeroelastic system is studied, the main idea of which is that under the condi- tion of stable limit cycle flutter of the aeroelastic system, the vibrations in the plunging and pitching of the wing can approximately be considered to be simple harmonic excita- tion to the control surface. The motion of the control surface can approximately be modeled by a nonlinear oscillation of one-degree-of-freedom. The range of the chaotic response of the aeroelastic system is approximately determined by means of the chaotic response of the nonlinear oscillator. The rich dynamic behaviors of the control surface are represented as bifurcation diagrams, phase-plane portraits and PS diagrams. The theoretical analysis is verified by the numerical results.     

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。     

Parametric aeroelastic modeling based on component modal synthesis and stability analysis for horizontally folding wing with hinge joints

To investigate the aeroelastic stability of a folding wing effectively, a parametric aeroelastic analysis approach is proposed. First, the fixed interface component modal synthesis is used to derive the structural dynamic equation for a folding wing, in which the elastic connection is considered. The unsteady aerodynamic model is established by the doublet lattice method (DLM), and the aeroelastic model is achieved from integration of the DLM with the component modal analysis. The generalized aerodynamic influence coefficient matrix is established by modes kept and constraint modes of each component. The aeroelastic stability of a folding wing is investigated based on the Gram matrix in control theory. The effectiveness of the proposed method is verified via comparison with traditional flutter eigenvalue analysis for both extended and folded configurations. The proposed method identifies coupled modes and improves computational efficiency when compared to classical aeroelastic stability analysis methods, such as the p–k method.     

Numerical simulation of force and wake mode of an oscillating cylinder

The turbulent flow behind a circular cylinder subjected to forced oscillation is numerically studied at a Reynolds number of 5500 by using three-dimensional Large Eddy Simulations (3-D LES) technique with the Smagorinsky model. The filtered equations are discretised using the finite volume method with an O-type structured grid and a second-order accurate method in both time and space. Firstly, the main wake parameters of a stationary cylinder are examined and compared in the different grid resolutions. Secondly, a transversely oscillating cylinder with a constant amplitude in a uniform flow is investigated. The cylinder oscillation frequency ranges between 0.75 and 0.95 of the natural Kármán frequency, and the excitation amplitude is moderate, 50% of the cylinder diameter. The flow characteristics of an oscillating cylinder are numerically examined and the corresponding wake modes are captured firstly in 3-D LES at Re=5500. A transition between different wake modes is firstly investigated in a set of numerical simulations.     

Flutter of a Wide Strip Plate in a Supersonic Gas Flow

The stability of an elastic plate in the form of a wide strip in a supersonic inviscid gas flow is investigated in the linear approximation. An expression for the dependence of the pressure on the plate deflection, asymptotically exact for wide plates, is used. Two qualitatively different instability types are obtained: flutter with respect to a single oscillatory mode due to negative aerodynamic damping and flutter of a related type due to the interaction of oscillatory modes. For each type the stability criterion and the frequency at which the oscillation amplitude grows most intensely are found.     

翼型厚度对气动弹性特性的影响

结合基于$k$-$\omega$的SST两方程湍流模型,求解雷诺平 均Navier-Stokes方程获得定常和非定常气动力,耦合翼型弹性运动方程,在时间 域内模拟了不同厚度对称翼型在不同迎角下的气动弹性动态过程, 并重点研究了较大迎角下的不同厚度翼型流场特征和气动弹性的性质,研究结果表明：在论 文所涉及的参数情况下,对于迎角从零到大迎角范围,翼型颤振临界速度随迎角的变化不是 单调的. 翼型颤振临界速度迅速下降的起始迎角比最大升力系数对应的迎角小很多.     

Physical mechanism of the destabilizing effect of damping in continuous non-conservative dissipative systems

The paper attempts to give a physical explanation of the mechanism behind the so-called destabilizing effect of small internal damping in the dynamic stability of Beck's column. Both internal (material) and external (viscous fluid) damping are considered. An energy equation is derived for the balance between the work done by the non-conservative ‘follower force’ and the energy dissipated by the internal and external damping forces. Evaluated at the critical load, where a flutter instability is initiated, this equation explicitly shows the influence of damping upon flutter frequency, phase angle, and vibration amplitude. The gradient of the phase angle, evaluated at the free end of the column, is found to be the ‘valve’, which controls how much work the follower force can do on the column during each period of oscillation. And a large change in this gradient with increasing—but still small—internal damping is found to be responsible for the destabilizing effect.     

On the design of an electromagnetic aeroelastic energy harvester from nonlinear flutter

A new energy harvester by coupling the electromagnetic induction and the pitch vibration of a rigid wing is built up in this paper. It is aimed: (1) to harvest energy from the pitch limit cycle oscillation (LCO) of the wing due to the preloaded free-play nonlinearity; (2) to introduce a theoretical analysis scheme based on the equivalent linearized method into the design of this harvester. With the equivalent linearized method, the domains of the single stable LCO and double stable LCOs are respectively obtained. Combining the analytical and numerical solutions, the single stable LCO along with the stable limit cycle amplitude greater than its corresponding unstable one is recognized as the better mode for harvesting, since the larger limit cycle domain is induced and the more energy are yielded. Based on such chosen mode, analyses of varying parameters are conducted with respect to the plunge stiffness, pitch stiffness, distance of elastic axis from center of gravity, distance of geometric center from elastic axis, load resistance and magnetic flux density. Meanwhile, three indicators are applied to reveal their effects on the harvesting performances: (1) the size of limit cycle domain, (2) the onset velocity of LCO, and (3) the energy output.     

Effects of multiple structural nonlinearities on limit cycle oscillation of missile control fin

Aeroelastic analyses are performed for a 2-D typical section model with multiple nonlinearities. The differences between a system with multiple nonlinearities in its pitch and plunge spring and a system with a single nonlinearity in its pitch are thoroughly investigated. The unsteady supersonic aerodynamic forces are calculated by the doublet point method (DPM). The iterative V-g method is used for a multiple-nonlinear aeroelastic analysis in the frequency domain and the freeplay nonlinearity is linearized using a describing function method. In the time domain, the DPM unsteady aerodynamic forces, which are based on a function of the reduced frequency, are approximated by the minimum state approximation method. Consequently, multiple structural nonlinearities in the 2-D typical wing section model are influenced by the pitch to plunge frequency ratio. This result is important in that it demonstrates that the flutter speed is closely connected with the frequency ratio, considering that both pitch and plunge nonlinearities result in a higher flutter speed boundary than a conventional aeroelastic system with only one pitch nonlinearity. Furthermore, the gap size of the freeplay affects the amplitude of the limit cycle oscillation (LCO) to gap size ratio.     

双稳态压电俘能器的簇发振荡与俘能效率分析

本文从理论上分析了双稳态压电俘能器在高频激励下的动力学行为和低频激励下的簇发振荡, 旨在为系统找到多条高能轨道从而提高俘能效率. 首先, 介绍了双稳态压电俘能器的结构以及一般模型. 与工程上研究俘能器的目的不同, 本文主要从动力学方面分析了俘能器的运动, 电压输出与效率, 包括高频激励下系统的低能阱内周期运动、阱间混沌运动等, 并说明了单个低频激励下双稳态压电俘能器会在阱间高能轨道上发生簇发振荡, 但在阱内低能轨道上只做周期运动. 同时, 结合振幅以及势阱深度等因素对簇发振荡的存在性和强度进行分析. 为了说明高能轨道与低能轨道对系统俘能效率的影响, 讨论了不同的等效阻尼、负载电阻下俘能器输出电压的变化, 找到了最优匹配. 最后, 对于多个低频外激励的情况, 从不同的轨道组合模式上得到了双高能簇发振荡模式输出的电压最大, 其次是单高能簇发振荡与单低能周期振荡的组合模式, 输出电压最低的是双低能周期振荡模式. 并与单个外激励进行对比, 表现了多个激励的良好性能.      

High-frequency plate flutter

Earlier, using the global instability method, the stability of a strip plate in a supersonic gas flow was investigated. In addition to the classical (low-frequency) flutter developing upon the interaction between the plate oscillation modes, a novel (high-frequency) flutter type in which the oscillations are unimodal was detected. In the present study, the effect on the high-frequency flutter of the plate width (earlier only an asymptotic analysis for a width tending to infinity was performed), its damping characteristics, and the presence of a gas at rest on the side opposite the flow is investigated.     

Influence of external loading on the stability of a fluid-conveying pipeline

A section of a fluid-conveying pipeline is analyzed for stability. The section is modeled by an elastic cylindrical shell of finite length under external static loading. Two qualitatively different instability modes are studied: “quasistatic” (divergence) and dynamic (flutter). The effect of uniform radial pressure and uniform axial pressure on the instability modes is analyzed     

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。     

Direct flutter and limit cycle computations of highly flexible wings for efficient analysis and optimization

The usefulness of flutter as a design metric is diluted for wings with destabilizing (softening) nonlinearities, as a stable high-amplitude limit cycle (subcritical) may exist for flight speeds well below the flutter point. It is thus desired to design aeroelastic structures such that the post-flutter behavior is as benign (i.e., supercritical) as possible, among the other constraints commonly considered in the optimization process. In order to account for these metrics in an accurate and efficient manner, direct tools are utilized to first locate the Hopf-point (flutter speed), and then to obtain a nonlinear perturbation solution via the method of multiple scales. The latter scheme provides a scalar variable whose sign and magnitude dictate the nature of the limit cycle. The accuracy of these methods is demonstrated with a high-aspect-ratio highly flexible wing, modeled with nonlinear beam finite elements and the ONERA dynamic stall tool. Stiffness and inertial design variables are allowed to vary spatially throughout the wing, in order to conduct gradient-based optimization of the limit cycle under flutter and mass constraints. The resulting wing structure demonstrates strongly supercritical behavior, as well as several design conflicts between linear (flutter) and nonlinear (limit cycles) sensitivities, which are not present in the uniform baseline wing.