A method for static aeroelastic analysis based on the high-order panel method and modal method

A method for static aeroelastic analysis based on the high-order panel method and modal method is presented. The static aeroelastic characteristics of flexible wings are investigated using this method. Three-dimensional aerodynamic models of flexible wings are constructed based on the geometry of wing configuration, and the modal method is adopted to achieve the fluid-structure coupling. The static aeroelastic characteristics of the AGARD445.6 wing and a low-aspect-ratio wing are investigated in this study....     

Multiobjective evolutionary computation for supersonic wing-shapeoptimization

This paper discusses the design optimization of a wing for supersonic transport (SST) using a multiple-objective genetic algorithm (MOGA). Three objective functions are used to minimize the drag for supersonic cruise, the drag for transonic cruise, and the bending moment at the wing root for supersonic cruise. The wing shape is defined by 66 design variables. A Euler flow code is used to evaluate supersonic performance, and a potential flow code is used to evaluate transonic performance. To reduce the total computational time, flow calculations are parallelized on an NEC SX-4 computer using 32 processing elements. The detailed analysis of the resulting Pareto front suggests a renewed interest in the arrow wing planform for the supersonic wing     

Coupled adjoint aerostructural wing optimization using quasi-three-dimensional aerodynamic analysis

This paper presents a method for wing aerostructural analysis and optimization, which needs much lower computational costs, while computes the wing drag and structural deformation with a level of accuracy comparable to the higher fidelity CFD and FEM tools. A quasi-three-dimensional aerodynamic solver is developed and connected to a finite beam element model for wing aerostructural optimization. In a quasi-three-dimensional approach an inviscid incompressible vortex lattice method is coupled with a viscous compressible airfoil analysis code for drag prediction of a three dimensional wing. The accuracy of the proposed method for wing drag prediction is validated by comparing its results with the results of a higher fidelity CFD analysis. The wing structural deformation as well as the stress distribution in the wingbox structure is computed using a finite beam element model. The Newton method is used to solve the coupled system. The sensitivities of the outputs, for example the wing drag, with respect to the inputs, for example the wing geometry, is computed by a combined use of the coupled adjoint method, automatic differentiation and the chain rule of differentiation. A gradient based optimization is performed using the proposed tool for minimizing the fuel weight of an A320 class aircraft. The optimization resulted in more than 10 % reduction in the aircraft fuel weight by optimizing the wing planform and airfoils shape as well as the wing internal structure.     

Composite stacking sequence optimization for aeroelastically tailored forward-swept wings

A method for stacking sequence optimization and aeroelastic tailoring of forward-swept composite wings is presented. It exploits bend-twist coupling to mitigate aeroelastic divergence. The method proposed here is intended for estimating potential weight savings during the preliminary aircraft design stages. A structural beam model of the composite wingbox is derived from anisotropic shell theory and the governing aeroelastic equations are presented for a spanwise discretized forward swept wing. Optimization of the system to reduce wing mass is undertaken for sweep angles of ?35° to 0° and Mach numbers from 0.7 to 0.9. A subset of lamination parameters (LPs) and the number of laminate plies in each pre-defined direction (restricted to {0°,±45°, 90°}) serve as design variables. A bi-level hybrid optimization approach is employed, making use of a genetic algorithm (GA) and a subsequent gradient-based optimizer. Constraints are implemented to match lift requirements and prevent aeroelastic divergence, excessive deformations, airfoil stalling and structural failure. A permutation GA is then used to match specific composite ply stacking sequences to the optimum design variables with a limited number of manufacturing constraints considered for demonstration purposes. The optimization results in positive bend-twist coupling and a reduced structural mass. Results are compared to an uncoupled reference wing with quasi-isotropic layups and with panel thickness alone the design variables. For a typical geometry and a forward sweep of ?25° at Mach 0.7, a wingbox mass reduction of 13 % was achieved.     

太阳能无人机机翼颤振动力学建模与分析

太阳能无人机作为一种大展弦比轻质飞行器,其机翼的气动弹性效应显著,其中颤振问题尤为关键。此类飞机具有大尺寸和低刚度特点,通过风洞试验研究机翼颤振问题,成本高而且难度大,难以实现,因此仿真计算是分析此类飞机颤振问题的主要手段。针对国内某翼展为40米的太阳能无人机大展弦比机翼,首先对机翼有限元模型进行工程化处理,在此基础上开展结构动力学分析和颤振计算,重点计算了机翼上不同吊舱布置下的颤振速度。经过仿真计算,得到该太阳能无人机机翼颤振速度为26m/s, 满足设计要求,进一步分析表明,可以通过增加发动机连杆的长度、在发动机上增加配重以及改变吊舱在机翼上的展向站位等手段来提高此无人机的颤振速度。     

底部导流板形式对高速列车气动阻力的影响

为减小高速列车运行时的气动阻力,设计直式、斜式、内圆弧式和外圆弧式等4种转向架前后底部导流板的高速列车模型.通过风洞试验验证数值模拟方法的有效性,采用数值计算分析底部导流板对列车气动阻力和底部流场的影响.结果表明：不同形式底部导流板的列车总阻力相差可达20%,其中头车气动阻力因数极差值最大为0.062.导流板影响列车底部气流速度和转向架区域压力分布,其导流作用使得转向架区域气动阻力和转向架的阻力同时改变.转向架前后导流板的导流效果越好,转向架区域的气动阻力越小;同时,气流冲击使得转向架上的滞止压力增大;在二者的共同作用下高速列车的总阻力存在一个较小值.底部采用直式导流板对降低全车气动阻力的效果最好.     

Coupling of a nonlinear finite element structural method with a Navier-Stokes solver

A new three-dimensional viscous aeroelastic solver is developed in the present work. A well validated full Navier-Stokes code is coupled with a nonlinear finite element plate model. Implicit coupling between the computational fluid dynamics and structural solvers is achieved using a subiteration approach. Computations of several benchmark static and dynamic plate problems are used to validate the finite element portion of the code. This coupled aeroelastic scheme is then applied to the problem of three-dimensional panel flutter. Inviscid and viscous supersonic results match previous computations using the same aerodynamic method coupled with a finite difference structural solver. For the case of subsonic flow, multiple solutions consisting of static, upward and downward deflections of the panel are discussed. The particular solution obtained is shown to be sensitive to the cavity pressure specified underneath the panel.     

Aeroelastic tailoring and scaling using Bacterial Foraging Optimisation

Bacterial Foraging Optimisation (BFO) is investigated in an attempt to evaluate its use in solving complex optimisation problems for aeronautical structures. A hybrid variant of BFOA, which incorporates meta-modelling techniques, is also proposed and employed. The efficiency and effectiveness of the methods are tested for tailoring a rectangular composite wing, aiming to maximise the flutter speed and for scaling a joined-wing aircraft, targeting to match aeroelastic responses between the physical prototype and wind tunnel model. The obtained results are compared with those found using a range of other biologically inspired optimisation methods (GA, PSO, ACO), proving that the social foraging behavior of motile bacteria is an effective tool for aeroelastic optimisation.     

Structural optimization of lifting surfaces with divergence and control reversal constraints

In this paper, the static aeroplastic characteristics, divergence velocity, control effectiveness and lift effectiveness are considered in obtaining an optimum weight structure. Swept wing structures are used with upper and lower skins, spar and rib thicknesses, and spar cap and vertical post cross-sectional areas as the design parameters. The aerodynamic strip theory is used to derive the constraint formulations and aerodynamic load matrices. A Sequential Unconstrained Minimization Technique (SUMT) algorithm is used to optimize the wing structure to meet the desired aeroelastic constraints.     

大展弦比机翼μ控制

利用大展弦比机翼后缘不同位置上的操纵面进行颤振主动控制,通过将大展弦比机翼简化为包含弯曲和扭转两种模态的悬臂梁结构,根据片条理论,建立包含操纵面运动规律的大展弦比机翼气动弹性方程.由于简化的数值模型与实际模型之间存在一定的误差,通常模型的运动方程包含有不确定参量用来表示建模误差.鲁棒控制方法能够得到一个有效控制器,控制这种带有模型不确定参量的运动方程.文中论述利用鲁棒μ控制方法,研究有两个操纵面大展弦比机翼的鲁棒控制问题.仿真结果表明鲁棒μ控制可以有效地抑制大展弦比机翼的受扰振动,提高颤振临界速度,且两个操纵面共同控制效果比单操纵面显著.     

迷你翼梢小翼增升减阻效应

选取某窄体客机的翼梢小翼为研究对象,采用Spalart Allmaras模型对无翼梢小翼、全尺寸翼梢小翼和迷你翼梢小翼3种机翼构型进行数值模拟,通过流场分析和速度分解等手段,研究翼梢小翼的增升减阻机理。结果表明：迷你翼梢小翼有恢复涡核流速、减弱涡流掺混程度和梳理翼梢气流的作用;增升减阻的关键在于迷你翼梢小翼对气流方向的修正;翼梢小翼的局部流动差异会对整体机翼近场造成影响。由于尺寸较小,迷你翼梢小翼能在较大攻角范围内改善传统翼梢小翼的性能,具有一定的实践意义。     

Sliding Mode Observer Controller Design for a Two Dimensional Aeroelastic System with Gust Load

The aim of this paper is to develop state estimation and sliding mode control schemes for the vibration suppression of an underactuated wing aeroelastic system in the presence of a gust load disturbance. Ignoring structural elastic deformation and using the concentrated elastic system (spring) to simulate the overall elastic deformation, this aeroelastic model consists of a straight wing and spring system, describing flap and pitch freedoms. The corresponding dynamic motion equation is established using the Lagrange method, and the gust is modeled as a typical “1‐cosine” gust. The aerodynamic lift and moment on the wing are computed by strip theory. The open loop system exhibits the limit cycle oscillations (LCOs) at a certain freestream velocity. The objective is to design a control system for suppressing the LCOs. For the purpose of control, a single trailing‐edge control surface is used. It is assumed that only the pitch angle is measured and the remaining state variables needed for full state feedback are estimated by the designed observer. Then an integral sliding surface is put forward on the estimation space; a new continuous reaching law is proposed to reduce the chattering phenomena. The finite‐time reachability of the predesigned sliding surface is proved and guaranteed by the designed sliding mode control law. The sufficient condition for the asymptotic stability of the closed‐loop system composed of the sliding mode dynamics and the error dynamical system is derived in terms of linear matrix inequality (LMI). The effectiveness of the proposed strategy is finally demonstrated by simulation results.     

Efficient nonlinear static aeroelastic wing analysis

The objective of this work is to demonstrate a computationally efficient, high-fidelity, integrated static aeroelastic analysis procedure. The aerodynamic analysis consists of solving the nonlinear Euler equations by using an upwind cell-centered finite-volume scheme on unstructured tetrahedral meshes. The use of unstructured grids enhances the discretization of irregularly shaped domains and the interaction compatibility with the wing structure. The structural analysis utilizes finite elements to model the wing so that accurate structural deflections are obtained and allows the capability for computing detailed stress information for the configuration. Parameters are introduced to control the interaction of the computational fluid dynamics and structural analyses; these control parameters permit extremely efficient static aeroelastic computations. To demonstrate and evaluate this procedure, static aeroelastic analysis results for a flexible wing in low subsonic, high subsonic (subcritical), transonic (supercritical), and supersonic flow conditions are presented.     

Aircraft morphing wing design by using partial topology optimization

A morphing wing concept has been investigated over the last decade because it can effectively enhance aircraft aerodynamic performance over a wider range of flight conditions through structural flexibility. The internal structural layouts and component sizes of a morphing aircraft wing have an impact on aircraft performance i.e. aeroelastic characteristics, mechanical behaviors, and mass. In this paper, a novel design approach is proposed for synthesizing the internal structural layout of a morphing wing. The new internal structures are achieved by using two new design strategies. The first design strategy applies design variables for simultaneous partial topology and sizing optimization while the second design strategy includes nodal positions as design variables. Both strategies are based on a ground structure approach. A multiobjective optimization problem is assigned to optimize the percentage of change in lift effectiveness, buckling factor, and mass of a structure subject to design constraints including divergence and flutter speeds, buckling factors, and stresses. The design problem is solved by using multiobjective population-based incremental learning (MOPBIL). The Pareto optimum results of both strategies lead to different unconventional wing structures which are superior to their conventional counterparts. From the results, the design strategy that uses simultaneous partial topology, sizing, and shape optimization is superior to the others based on a hypervolume indicator. The aeroelastic parameters of the obtained morphing wing subject to external actuating torques are analyzed and it is shown that it is practicable to apply the unconventional wing structures for an aircraft.     

Aeroelastic tailoring considering uncertainties in material properties

The robustness of aeroelastic design optimization with respect to uncertainties in material and structural properties is studied both numerically and experimentally. The model consists of thin orthotropic composite wings virtually without fuselage. Three different configurations with consistent geometry but varying orientation of the main stiffness axis of the material are investigated. The onset of aeroelastic instability, flutter, is predicted using finite element analysis and the doublet-lattice method for the unsteady aerodynamic forces. The numerical results are experimentally verified in a low-speed wind tunnel. The optimization problem is stated as to increase the critical air speed, above that of the bare wing by massbalancing. It is seen that the design goals are not met in the experiments due to uncertainties in the structural performance of the wings. The uncertainty in structural performance is quantified through numerous dynamic material tests. Once accounting for the uncertainties through a suggested reformulation of the optimization problem, the design goals are met also in practice. The investigation indicates that robust and reliable aeroelastic design optimization is achievable, but careful formulation of the optimization problem is essential.     

伸缩机翼动力学实验研究

伸缩机翼作为一种可变形翼,变形过程中,由于其结构时变特性严重且具有柔性体特点,翼面横向振动特性实时变化从而影响飞行性能.基于伸缩机翼实验模型,分别开展了伸缩机翼模态实验研究和振动实验研究,即通过PLC控制系统控制伺服电机转速及转向,进而改变活动机翼外伸长度及伸缩速度,一方面利用锤击法获得五种活动机翼外伸长度下前三阶模态参数;另一方面利用ICP加速度传感器、LMS数据采集分析系统采集并处理活动机翼伸缩过程中横向振动加速度信号,获得伸缩速度与最大振动幅值、最大响应频率之间关系.结果表明,活动机翼外伸长度增加,结构固有频率值减小;伸缩速度作为参数激励,其等效激励与活动机翼某一外伸长度下某一阶固有频率值相近而导致颤振.为伸缩机翼翼面振动控制和伸缩速度合理选择提供理论依据和技术支持.     

Multi-fidelity wing aerostructural optimization using a trust region filter-SQP algorithm

A trust region filter-SQP method is used for wing multi-fidelity aerostructural optimization. Filter method eliminates the need for a penalty function, and subsequently a penalty parameter. Besides, it can easily be modified to be used for multi-fidelity optimization. A low fidelity aerostructural analysis tool is presented, that computes the drag, weight and structural deformation of lifting surfaces as well as their sensitivities with respect to the design variables using analytical methods. That tool is used for a mono-fidelity wing aerostructral optimization using a trust region filter-SQP method. In addition to that, a multi-fidelity aerostructural optimization has been performed, using a higher fidelity CFD code to calibrate the results of the lower fidelity model. In that case, the lower fidelity tool is used to compute the objective function, constraints and their derivatives to construct the quadratic programming subproblem. The high fidelity model is used to compute the objective function and the constraints used to generate the filter. The results of the high fidelity analysis are also used to calibrate the results of the lower fidelity tool during the optimization. This method is applied to optimize the wing of an A320 like aircraft for minimum fuel burn. The results showed about 9 % reduction in the aircraft mission fuel burn.     

Integrated aerodynamic-structural-control wing design

The aerodynamic-structural-control design of a forward-swept composite wing for a high subsonic transport aircraft is considered. The structural analysis is based on a finite-element method. The aerodynamic calculations are based on a vortex-lattice method, and the control calculations are based on an output feed-back control law. The wing is designed for minimum weight subject to structural, performance/aerodynamic and control constraints. Efficient techniques are developed to calculate the control-deflection and control-effectiveness sensitivities which appear as second-order derivatives in the control constraint equations. To suppress the aeroelastic divergence of the forward-swept wing, and to minimize the take-off gross weight of the design aircraft, two separate cases are studied: (1) combined application of aeroelastic tailoring and active controls; and (2) aeroelastic tailoring alone. For the particular example problem considered in this study, the aeroelastic tailoring was found to have a substantially greater influence than active controls on weight minimization and divergence suppression.     

跨声速机翼阻力分解的一种改进方法

提出黏性区域探测器的一种改进形式，并用于捕捉激波和翼梢涡的熵增阻力；给出尾迹平面的可压缩涡动力学诱导阻力表达式，并与基于热力学的诱导阻力对比。在跨声速来流状态下，对ONERA M6和某民用飞机巡航状态下的机翼阻力进行分解，同时分析该民用飞机机翼安装翼梢小翼前、后的远场阻力构成。结果表明：新的区域探测器合理可靠，黏性阻力与伪熵阻力的计算结果更加准确；2种诱导阻力计算方式的计算结果一致，但基于涡动力学的诱导阻力计算方法受积分平面位置的影响更小；安装翼梢小翼基本不影响整个流场的黏性阻力，减阻的主要效果体现为诱导阻力的减小。     

A numerical investigation of fluid flow over a rotating cylinder with cross flow oscillation

This paper studies a two-dimensional incompressible viscous flow past a rotating cylinder with cross flow oscillation using a finite element method based on the characteristic based split (CBS) algorithm to solve governing equations including full Navier–Stokes and continuity equations. Dynamic unstructured triangular grid is used employing lineal and torsional spring analogy which is coupled with the solver by an Arbitrary Lagrangian–Eulerian (ALE) formulation. After verifying the accuracy of the numerical code, simulations are conducted for the flow past a rotating cylinder with cross flow oscillation at moderate Reynolds numbers of 50, 100, and 200 considering different non-dimensional rotational speeds based on the free-stream velocity in the range 0–2.5, and various oscillating amplitudes and frequencies. Effects of the oscillation and rotation of the cylinder on the vortex shedding both in lock-on and non-lock-on regions, the mean drag and lift coefficients, and the Strouhal number are investigated in detail. It is found that similar to the fixed cylinder beyond a critical non-dimensional rotational speed the vortex shedding is highly suppressed. In addition, by increasing the rotational speed of the cylinder, the lift coefficient increases while decreasing the drag coefficient. However, in the vortex lock-on region both the lift and the drag coefficients increase significantly.