AEROELASTIC TAILORING OF AERONAUTICAL COMPOSITE WING STRUCTURES

This paper deals with the aeroelastic tailoring of aeronautical composite wing surfaces. The objective function is structural weight. Multi constraints, such as displacements, flutter speed and gauge requirements, are taken into consideration. Finite element method is used to the static analysis. Natural vibration modes are obtained by the spectral transformation Lanczos method. Subsonic doublet lattice method is used to obtain the unsteady aerodynamics.The critical flutter speed is generated by V-g method.The optimal problem is solved by the feasible direction method.The thickness of the composite wing skin is simulated by bicubic polynomials, whose coefficients combined with the cross-sectional areas or thicknesses of other finite elements are the design variables. The scale of the problem is reduced by variable linkage. Derivative analysis is performed analytically.Two composite wing boxes and a swept-back composite wing are optimized at the end of the paper.     

NUMERICAL SOLUTION OF NAVIER-STOKES EQUATION OF UNSTEADY SEPARATED FLOWS DUE TO A SPOILER''''S OSCILLATION

The finite difference method (FDM) is applied in the present paper to solve the unsteady NHS equations for incompressible fluids. ADI and SLOR methods are served for the vorticity equation and the Poisson equation for ψ respectively. The upwind scheme is used for the convective terms. The moving boundary conditions are specially treated, and the effects of outlet conditions on the flow field are abo examined. Numerical results obtained show that the spoiler's oscillation induces forming, growing and shedding of the vortices. The shedding frequency of vortices is equal to that of the spoiler's oscillation. The forced unsteady separated flows under the present investigation depend mainly on the reduced frequency. At low reduced frequency, the vortices shed from the spoiler interact weakly with each other, and move downstream at an almost uniform speed of 038 V∞. At high reduced frequency, the interaction between the adjacent vortices strengthens. They close up to and rotate around each other, and eventu     

Gust response analysis and wind tunnel test for a high-aspect ratio wing

A theoretical nonlinear aeroelastic response analysis for a flexible high-aspect ratio wing excited by harmonic gust load is presented along with a companion wind tunnel test. A multidisci-plinary coupled numerical calculation is developed to simulate the flexible model wing undergoing gust load in the time domain via discrete nonlinear finite element structural dynamic analysis and nonplanar unsteady vortex lattice aerodynamic computation. A dynamic perturbation analysis about a nonlinear static equilibrium is also used to determine the small perturbation flutter bound-ary. A novel noncontact 3-D camera measurement analysis system is firstly used in the wind tunnel test to obtain the spatial large deformation and responses. The responses of the flexible wing under different static equilibrium states and frequency gust loads are discussed. The fair to good quanti-tative agreements between the theoretical and experimental results demonstrate that the presented analysis method is an acceptable way to predict the geometrically nonlinear gust response for flex-ible wings.     

Characteristic analysis of lock-in for an elastically suspended airfoil in transonic buffet flow

Numerical simulations are performed to study the aeroelastic responses of an elastically suspended airfoil in transonic buffet flow, by coupling the unsteady Reynolds-averaged Navier-Stokes (RANS) equations and structural motion equation. The current work focuses on the char-acteristic analysis of the lock-in phenomenon. Great attentions are paid to studying the frequency range of lock-in and the effects of the three parameters, namely the structural natural frequency, mass ratio and structural damping, on lock-in characteristic of the elastic system in detail. It is found that when the structural natural frequency is close to the buffet frequency, the coupling fre-quency of the elastic system is no longer equal to the buffet frequency, but keeps the same value as the structural natural frequency. The frequency lock-in occurs and stays present until the structural nature frequency is near the double buffet frequency. It means that the lock-in presents within a broad range, of which the lower threshold is near the buffet frequency, while the upper threshold is near the double buffet frequency. Moreover, the frequency range of lock-in is affected by mass ratio and structural damping. The lower the mass ratio and structural damping are, the wider the range of lock-in will be. The upper threshold of lock-in grows with the mass ratio and structural damping decreasing, but the lower threshold always keeps the same.     

Bifurcation in a 3-DOF Airfoil with Cubic Structural Nonlinearity

Limit cycle oscillations (LCOs) as well as nonlinear aeroelastic analysis of a 3-DOF aeroelastic airfoil motion with cubic restoring moments in the pitch degree of freedom are investigated.Aeroelastic equations of an airfoil with control surface in an incompressible potential flow are presented in the time domain.The harmonic balance (HB) method is utilized to calculate the LCO frequency and amplitude for the airfoil.Also the semi-analytical method has revealed the presence of stable and unstable limit cycles,along with stability reversal in the neighborhood of a Hopf bifurcation.The system response is determined by numerically integrating the governing equations using a standard Runge-Kutta algorithm and the obtained results are compared with the HB method.Also the results by the third order HB (HB3) method for control surface are consistent with the other numerical solution.Finally,by combining the numerical and the HB methods,types of bifurcation,be it supercritical,subcritical,or divergent flutter area are identified.     

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.     

A ROBUST FLOW CALCULATION TECHNIQUE WITH MULTIPLE FINITE CONTROL VOLUMES

The control volume method gives the forces which act on the system, but not necessarily the wall pressure of the system. The author has made an attempt to develop a control volume method which makes it possible to obtain the wall pressure of the control volume. The 2-D inviscid incompressible steady duct flow is considered. The conservation equations in integral form are discretized for a control volume. The circulation along the control surface is expressed as a nonlinear function of the vertical velocity component at the inlet and is set equal to zero for the inviscid flow. The equation is solved by the Newton method, and the other aerodynamic properties can be obtained. The calculated results have been compared to the experiment and the agreement has been found fairly satisfactory.     

Numerical study of unsteady starting characteristics of a hypersonic inlet

The impulse and self starting characteristics of a mixed-compression hypersonic inlet designed at Mach number of 6.5 are studied by applying the unsteady computational fluid dynamics (CFD) method. The full Navier-Stokes equations are solved with the assumption of viscous perfect gas model, and the shear-stress transport (SST) k-x two-equation Reynolds averaged Navier- Stokes (RANS) model is used for turbulence modeling. Results indicate that during impulse starting, the flow field is divided into three zones with different aerodynamic parameters by primary shock and upstream-facing shock. The separation bubble on the shoulder of ramp undergoes a generating, growing, swallowing and disappearing process in sequence. But a separation bubble at the entrance of inlet exists until the freestream velocity is accelerated to the starting Mach number during self starting. The mass flux distribution of flow field is non-uniform because of the interaction between shock and boundary layer, so that the mass flow rate at throat is unsteady during impulse starting. The duration of impulse starting process increases almost linearly with the decrease of freestream Mach number but rises abruptly when the freestream Mach number approaches the starting Mach number. The accelerating performance of booster almost has no influence on the self starting ability of hypersonic inlet.     

利用气动弹性的柔性振翼模型建立及分析(英文)

Making use of modal characteristics of the natural vibration of flexible structure to design the oscillating wing aircraft is proposed. A series of equations concerning the oscillating wing of flexible structures are derived. The kinetic equation for aerodynamic force coupled with elastic movement is set up, and relevant formulae are derived. The unsteady aerodynamic one in that formulae is revised. The design principle, design process and range of application of such oscillating wing analytical method are elaborated. A flexible structural oscillating wing model is set up, and relevant time response analysis and frequency response analysis are conducted. The analytical results indicate that adopting the new-type driving way for the oscillating wing will not have flutter problems and will be able to produce propulsive force. Furthermore, it will consume much less power than the fixed wing for generating the same lift.     

颤振边界高效精确预测分析

高效精确地确定多种飞机构型的颤振边界在飞机设计过程中具有重要意义。为了提高计算效率和计算结果的准确性,针对亚声速和跨声速两种马赫数区域,提出分别采用线性和非线性方法进行非定常气动力分析。非线性分析在引入精确的定常气动力的基础上,采用高效率跨声速小扰动方程进行求解;颤振求解统一采用g 法。对大型飞机的梁架—减缩刚度组合模型的空机及三种典型燃油构型进行涵盖飞行包线的全马赫数变高度颤振分析,结果表明：四种构型的颤振边界与颤振试飞边界一致,与其他分析方法相比,效率有明显提高,尤其是对多种飞机构型能够高效地获得准确的颤振边界,即说明本文采用的方法是目前适用于工程上的一种高效精确的预测大型飞机颤振边界的方法。     

操纵面对跨声速机翼气动弹性特性的影响

 运用基于非定常CFD的气动力辨识技术，得到跨声速非定常气动力降阶模型。耦合结构动力学方程，建立了基于状态空间的跨声速气动弹性分析模型。分析了典型三自由度二元机翼的颤振边界，分析结果与CFD/CSD直接耦合方法吻合。然后研究了操纵面结构参数（固有频率和重心位置）对跨声速气动弹性特性的影响。研究发现,一些传统的结构设计准则和颤振排除技术未必适用于跨声速状态;操纵面偏转模态常常成为诱发跨声速颤振的主要模态;经典的质量平衡技术可能会降低跨声速气动弹性系统的稳定性。     

跨声速偶极子网格法的改进和在颤振计算中的应用

本文改进了跨声速偶极子网格法,且在颤振计算中能自动求颤振点速度,从而较大地提高了计算速度。     

带操纵面机翼的非定常跨声速流数值解法

本文介绍一种带操纵面机翼的非定常跨声速流的有限差分计算方法。采用的方程是三元非定常跨声速修正小扰动位势方程,使用时间积分法,求解格式是近似因式分解交替方向隐式(ADI)格式。使用这种方法计算了F-5机翼在来流马赫数为0.9和0.925时的定常气动力和操纵面振荡的非定常气动力。计算结果与国外的NLR试验结果和XTRAN3S方法的计算结果进行了比较,表明计算是成功的。     

二维非定常跨音速时间线化积分方程的推导

1.引言 求解非定常跨音速流动的主要困难是非线性问题。对于微幅翼型振动问题(同时引起激波的微幅振动)可做时间线化简化处理,得到时间线化微分方程定解问题。时间线化积分方程是由时间线化微分方程推导出来的。     

用单调相关格式计算跨音速定常和非定常翼型绕流

在跨音速非定常势流计算方面,人们一直在努力寻求处理包括激波运动的非线性问题的能力和计算效率的提高。Ballhaus和Goorjian（1977年）用交替方向隐式差分法（ADI）成功地直接求解了时间域的非定常跨音速小扰动方程（此法称为“LTRAN2”）。文献[5]引用坐标变换也做了类似LTRAN2的工作。     

计算绕薄翼型跨音速非定常流的积分方程法

 本文从小扰动方程出发推导出绕薄翼型的跨音速非定常流动的积分方程,为使此方程适用于具有激波的流场引入人工粘性,并对其数值求解。通过算例讨论了人工粘性和计算域大小对计算结果的影响,与实验及其它数值计算结果比较,表明本方法的准确度令人满意。计算量小,收敛性好是此法的特点。     

LTRAN2的一种改进方案

本文提出了LTRAN2的一种改进方案BTRAN2用来分析翼型的各种频率的非定常运动。用E-O调转换格式和ADI方法求解了完全的二维非定常跨音速小扰动位流方程,并用单调转换的AF2格式计算了定常跨音速小扰动方程,以此作为非定常计算的初场。本文给出了带后缘正弦振荡襟翼的NACA 64 A006翼型的绕流和做正弦俯仰振荡的NACA 64 A010翼型的绕流计算结果,它们与Euler方程解或实验数据很吻合。     

任意翼型有小分离气泡跨音速绕流的迭代算法

 本文给出一种带小分离气泡的任意翼型粘性跨音速绕流的计算方法,采用有粘-无粘干扰迭代的概念。无粘流的全速势方程用AF差分格式在保角变换法生成的计算网格中求解,粘流附面层方程用C-S盒式法求解,用逆算法消除分离点处的奇性。本文对Ma_∞=0.8,Re_∞=2×10~6,迎角α=3.5°和4°的NACA64A010翼型粘性绕流进行了计算,结果与实验相比较,吻合良好。     

全速位积分方程跨音速绕流的数值解

从全速位方程出发,利用Green公式将其化为激波捕获积分方程和激波装配积分方程,然后离散进行数值解。流场出现激波时,对激波捕获积分方程应用上风技术捕捉到激波,然后应用激波装配技术计算,得到了满意的结果。经算例考核,该方法具有计算区域小,收敛快和CPU时间较少等优点。