高机动性飞机大迎角全局稳定性研究

为了研究高机动性飞机大迎角时的全局稳定性,建立了飞机的非线性动力学方程和气动力模型。应用微分方程定性理论,通过对算例的分析,讨论了目前飞机设计中常用的稳定性判据(如(C_(nβ))_(dyn)和LCDP等)与平衡点处线化矩阵特征值之间的关系。结果表明这些判据可以从飞机在特定状态下的线化矩阵经简化得到。因此,它们只能预测大迎角的局部稳定性。若考虑全局稳定性则必须用本文所给的方法进行。最后,指出了平衡点和周期吸引子与机翼摇晃、偏离和尾旋之间的关系,并用时间过程作了验证。     

多输入气动伺服弹性系统抗阵风不灵敏性研究

针对飞机多输入 /多输出气动伺服弹性系统的抗阵风不灵敏性进行理论分析与计算验证。对于耦合的多回路飞行控制系统 ,建立弹性结构、非定常气动力和控制系统构成的气动伺服弹性分析模型 ,根据现代控制理论中的鲁棒分析技术 ,以系统回差矩阵的奇异值理论为基础 ,应用系统抗干扰不灵敏性的判据 ,确定气动伺服弹性系统对阵风外干扰保持不灵敏性的能力。以某型飞机横侧向耦合控制系统为对象 ,采用 Dryden谱形式的大气紊流模型作为外部阵风干扰 ,对系统的不灵敏性及阵风响应进行计算、分析和比较     

Aeroservoelastic modeling and analysis of a canard-configured air-breathing hypersonic vehicles

Air-breathing hypersonic vehicles (HSVs) are typically characterized by interactions of elasticity, propulsion and rigid-body flight dynamics, which may result in intractable aeroservoelastic problem. When canard is added, this problem would be even intensified by the introduction of low-frequency canard pivot mode. This paper concerns how the aeroservoelastic stability of a canard-configured HSV is affected by the pivot stiffnesses of all-moveable horizontal tail (HT) and canard. A wing/pivot system model is developed by considering the pivot torsional flexibility, fuselage vibration, and control input. The governing equations of the aeroservoelastic system are established by combining the equations of rigid-body motion, elastic fuselage model, wing/pivot system models and actuator dynamics. An unsteady aerodynamic model is developed by steady Shock-Expansion theory with an unsteady correction using local piston theory. A baseline controller is given to provide approximate inflight characteristics of rigid-body modes. The vehicle is trimmed for equilibrium state, around which the linearized equations are derived for stability analysis. A comparative study of damping ratios, closed-loop poles and responses are conducted with varying controller gains and pivot stiffnesses. Available bandwidth for control design is discussed and feasible region for pivot stiffnesses of HT and canard is given.     

Helicopter Motion: Equation Linearization

The nonlinear set of equations which represent helicopter motion is linearized about a prescribed nominal state. Once the linearized system is obtained it is validated by comparing the output of the nonlinear system to that of its linearized counterpart. Having obtained a linear model, linear system theory may then be applied in order to investigate the stability and control characteristics of the aircraft.     

大展弦比飞机运动稳定性建模及分析

随着飞机设计技术的提高和新型材料的应用,现代飞机气动弹性效应越来越显著,刚体运动和弹性体振动互相解耦的关系不再明显,在分析弹性飞机相关问题时,应该综合地、统一地考虑飞行动力学与气动弹性力学问题。本文在一般体轴系下综合考虑了飞机刚体运动自由度和弹性振动自由度,建立弹性飞机刚弹耦合运动的方程,通过在配平状态下引入小扰动运动假设,以及利用有理函数拟合技术将偶极子格网法计算所得到的频域非定常气动力转化为时域形式,建立了大柔性飞机刚弹耦合小扰动状态空间方程,并对纵向运动稳定性进行分析计算。针对文中构型飞机,运用本文所采用的方法计算所得结果与传统的线性及刚体计算所得结果进行对比分析,非线性计算方法所得颤振速度更小,而且失稳模态亦发生变化,对于飞行力学模态而言,长周期模态稳定性变差,在计算速度范围内出现失稳的现象,由此说明机翼大变形对飞机稳定性所带来的影响。     

考虑刚体/弹性耦合的飞机动稳定性研究

以大展弦比无人机为例,研究刚体运动对飞机纵向气动伺服弹性(ASE)稳定性以及弹性运动对飞机纵向飞行力学稳定性的影响。针对带有反馈控制的飞机,建立了时域的刚体/弹性耦合运动方程;以刚体运动频率与一阶弹性频率的比值为参数,分别研究有无刚体模态时的ASE系统稳定裕度和有无弹性模态时的飞行力学稳定裕度。数值仿真表明,当飞机刚体运动频率与一阶弹性频率的比值大于0.01时,飞机刚体运动对ASE稳定裕度的影响逐渐显现;当该频率比值大于0.03时,飞机弹性运动对飞行力学稳定裕度的影响开始突现;当该频率比大于1时,刚体弹性耦合效应不可忽略。     

Control law design for transonic aeroservoelasticity

Existing computational transonic aeroservoelastic researches focus on directly coupling the structural dynamic equations, CFD solver and servo system in time domain, study the effect of the given feedback control laws on the responses of the aeroelastic system. These works have not involved the design of the flutter active control law. The non-linearity of transonic flow brings great difficulties to aeroservoelastic analysis and design. Recent research of the unsteady aerodynamic reduced order models (ROM) based on CFD provides a challenging approach for transonic aeroservoelastic analysis and design. Coupling the structural state equations with the aerodynamic state equations of the wing and the control surface based on the ROM, we construct a transonic aeroservoelastic model in state-space. Then the sub-optimal control method based on output feedback is used to design the flutter suppressing law. The study first demonstrates the open loop of the Benchmark Active Controls Technology (BACT) wing. The computational results of the CFD direct simulation method and the ROM analysis method are both agree well with the experimental data. Then both the closed loop time responses and the flutter results by ROM technique are compared with those of numerical aeroservoelastic simulation based on Euler codes to validate the correctness of the design method of the control law and aeroservoelastic analysis method. An increase of up to 20% of the speed index can be achieved by the control law designed by sub-optimal control method for this model.     

Nonlinear dynamics of a two dimensional airfoil with freeplay in an inviscid compressible flow

A state space model for the aerodynamic loads calculated by means of a linearized CFD code which models the inviscid compressible flow is used to analyze the dynamic behavior of a two dimensional airfoil with a freeplay structural nonlinearity in pitch. The state space aeroelastic equations of motion are utilized to study the existence of Hopf bifurcation points and of limit cycle oscillations. Comparisons between the solutions of the coupled Euler-structural equations and the state space equations are provided.     

三升力面飞机扰动运动线性化模型

以一种新型运输机Ecolifter为参考模型，由飞机的一般运动方程，在扰动飞行条件下．对空气动力和空气动力矩、推力和推力矩进行线性化，结合稳态飞行运动方程，推导出稳态直线飞行条件下，Ecolifter的纵向和侧向线性化扰动运动方程。该方程是三升力面气动布局飞机飞行性能和动稳定性研究及一体化设计的基础。     

推力协同的大柔性飞行器纵向姿态控制

为了研究大柔性飞行器飞行/结构耦合动力学特性,提出了改进的面向控制的大柔性飞行器多体模型,开展了大柔性飞行器纵向动力学耦合特性分析与推力协同下纵向姿态控制律设计。采用二面角动态近似描述大柔性飞行器结构动力学特征,并推导了纵向耦合动力学模型。根据改进模型在配平点处的线性化模型,分析了飞行/结构耦合系统的纵向稳定性与结构变形量之间的关系。针对大柔性飞行器姿态稳定与跟踪,设计了纵向姿态控制器。与常规飞行器相比,大柔性飞行器飞行过程中会发生大变形,当载荷较大时,配平构型近似“U”形,此时纵向动力学具有长周期不稳定的特征。分析结果表明:大柔性飞行器各模态之间的耦合程度随着变形的增大而增大。此外,纵向姿态控制需要升降舵与推力协同控制速度和俯仰角并且考虑结构动力学的影响,否则飞行/结构的耦合作用会导致姿态跟踪误差衰减缓慢甚至发散。      

民机极限飞行状态的动态气动力试验与建模

飞行失控是造成民机灾难性航空事故的重要因素,飞行失控中飞机难以避免超出正常飞行包线范围,进入具有复杂非线性和非定常动态气动特性的极限飞行状态。本文开展典型民机布局飞机极限飞行状态的动导数、大振幅试验,对大迎角动态气动力的参数影响规律以及非线性、非定常特性进行分析和建模。结果表明,在飞机失速到过失速区域,飞行姿态快速变化过程中动态气动力的非线性和非定常特征显著;在动导数试验和建模中,考虑运动角速率的影响,可以预示气动力非线性的迎角范围,并捕捉到关于飞机动稳定性演化的关键特征;利用Goman-Khrabrov状态空间模型结合大振幅试验,可以确定模型中表征非定常特征的关键时间常数,获得特定极限飞行状态运动中的非定常动态气动力特性。研究方法和结果为开展民机极限飞行状态的动态气动力风洞试验设计与建模提供了一个可行途径,能改进飞机飞行失控预防、极限状态改出、飞行模拟训练和飞行事故分析等。     

新型倾转旋翼机气动弹性稳定性分析模型

针对新型半铰接式、摆振柔软的倾转旋翼机,建立了旋翼/动力舱/机翼气动弹性稳定性分析模型,并导出了系统的动力学方程。分析模型采用了当量铰、弹性约束的刚性桨叶和弹性机翼假设,采用Pitt-Peters动力入流模型计入了非定常气动力的影响。所建模型为进行新型倾转旋翼机大速度飞行时气弹稳定性分析奠定了基础。     

Gain self-scheduled H_∞ control for morphing aircraft in the wing transition process based on an LPV model

This article investigates gain self-scheduled H 1 robust control system design for a tailless fold- ing-wing morphing aircraft in the wing shape varying process. During the wing morphing phase, the aircraft’s dynamic response will be governed by time-varying aerodynamic forces and moments. Nonlinear dynamic equations of the morphing aircraft are linearized by using Jacobian linearization approach, and a linear parameter varying (LPV) model of the morphing aircraft in wing folding is obtained. A multi-loop controller for the morphing aircraft is formulated to guarantee stability for the wing shape transition process. The proposed controller uses a set of inner-loop gains to provide stability using classical techniques, whereas a gain self-scheduled H 1 outer-loop controller is devised to guarantee a specific level of robust stability and performance for the time-varying dynamics. The closed-loop simulations show that speed and altitude vary slightly during the whole wing folding process, and they converge rapidly after the process ends. This proves that the gain self-scheduled H 1 robust controller can guarantee a satisfactory dynamic performance for the morphing aircraft during the whole wing shape transition process. Finally, the flight control system’s robustness for the wing folding process is verified according to uncertainties of the aerodynamic parameters in the nonlinear model.     

Active Control Law Design for Flutter/LCO Suppression Based on Reduced Order Model Method

 Active stability augmentation system is an attractive and promising technology to suppress flutter and limit cycle oscillation (LCO). In order to design a good active control law, the control plant model with low order and high accuracy must be pro-vided, which is one of the most important key points. The traditional model is based on low fidelity aerodynamics model such as panel method, which is unsuitable for transonic flight regime. The physics-based high fidelity tools, reduced order model (ROM) and CFD/CSD coupled aeroservoelastic solver are used to design the active control law. The Volterra/ROM is applied to constructing the low order state space model for the nonlinear unsteady aerodynamics and static output feedback method is used to active control law design. The detail of the new method is demonstrated by the Goland+ wing/store system. The simu-lation results show that the effectiveness of the designed active augmentation system, which can suppress the flutter and LCO successfully.     

Conceptual design and flight test of two wingtip-docked multi-body aircraft

To overcome the drawbacks such as large wing deformations, poor performance encountering gusts, limits in taking off and landing, inconvenience of transportation of High-Altitude Long-Endurance (HALE) Unmanned Aerial Vehicles (UAVs), a new conceptual aircraft called wingtip-docked Multi-Body Aircraft (MBA) has attracted lots of attentions. Aiming to investigate the feasibility of this concept, two UAV models were designed, manufactured and connected by a wingtip-docking mechanism, which only allows the relative roll motion between the two aircraft. The trim solution of the two connected aircraft is firstly obtained by solving the developed nonlinear flight dynamic equations, followed by the stability analysis based on the linearized model. The results show that the connected aircraft is inherently unstable and cannot fly without a reasonable flight control system. A set of Proportional-Integral-Derivative (PID) control laws was then developed and implemented in the two experimental aircraft. The success of the flight tests show that the flight control can effectively eliminate the unstable motion and the wingtip-docked MBA is controllable and feasible.     

Dihedral influence on lateral–directional dynamic stability on large aspect ratio tailless flying wing aircraft

The influence of dihedral layout on lateral–directional dynamic stability of the tailless flying wing aircraft is discussed in this paper. A tailless flying wing aircraft with a large aspect ratio is selected as the object of study, and the dihedral angle along the spanwise sections is divided into three segments. The influence of dihedral layouts is studied. Based on the stability derivatives calculated by the vortex lattice method code, the linearized small-disturbance equations of the lateral modes are used to determine the mode dynamic characteristics. By comparing 7056 configurations with different dihedral angle layouts, two groups of stability optimized dihedral layout concepts are created. Flight quality close to Level 2 requirements is achieved in these optimized concepts without any electric stability augmentation system.     

Stability analysis of the nonlinear dynamics of flexible aircraft

A method for nonlinear analysis of dynamic systems and its application to aircraft is presented. In extension of existing averaging methods this procedure follows the linearizing eigenvalue analysis, but defines the eigenvalues and eigenvectors as functions of the deflection from the examined equilibrium. This approach allows the determination of the systems stability properties beyond the close vicinity of the equilibrium. Its application to the model of a large aircraft with flexible degrees-of-freedom is shown. The analysis shows very good agreement with the results of bifurcational analyses and allows a quantitative prediction of the systems stability behavior even between the branches of the bifurcation diagram. The unstable spiral mode of the flexible aircraft shows the existence of additional stable equilibria in the vicinity of the examined equilibrium due to system nonlinearities.     

高超声速飞行器非线性建模及开环特性分析

根据美国NASA Langly研究中心提供的有关Winged-Cone飞行器的文献资料,建立和完善了一类近空间高超声速飞行器的六自由度非线性模型.开环仿真分析结果表明,所建模型能够体现出高超声速飞行器复杂的非线性、强耦合性以及快速时变性等特点,可以为开展近空间高超声速飞行器控制问题的研究提供测试平台.     

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.     

非线性反馈线性化方法在飞控系统中的应用

应用非线性反馈精确线性化方法进行了飞机自动着陆系统设计。首先,用非线性微分方程组表示的飞机纵向控制系统,在经过输入输出反馈线性化以后,可等效为线性系数;然后,用线性的ＰＩＤ控制方法对变换后的线性系统进行了设计;最后,考虑风的干扰,对所设计的飞机自动着陆控制系统进行了数字仿真。仿真结果表明,用反馈线性化方法设计的飞控系统具有良好的性能。