摘要
金果园柔性航天器通常安装有太阳翼，以提供航天器正常工作所需的能量。由于太阳翼弹性振动不可避免地与航天器姿态运动相互耦合，因此柔性航天器是典型的刚柔耦合系统。当前，为了使航天器具有更强的功能和更长的寿命，太阳翼的尺寸不断增大，这导致系统刚柔耦合效应愈发显著。值得关注的是，目前多数的柔性航天器动力学与控制研究还依赖于单独的太阳翼的模态（假设模态）。这类模态无法反映柔性航天器的刚柔耦合效应，导致相关的研究方法应用于安装大型太阳翼的航天器时，存在一定的不足。因此，发展柔性航天器刚柔耦合模态的获取方法，在此基础上深入开展航天器姿态运动与太阳翼结构振动
本文以安装大型太阳翼的柔性航天器为研究对象，结合解析分析和数值仿真两种手段，提出了获取系统刚柔耦合模态的解析方法，并以此为基础，从固有特性、动力学建模及动力学特性、姿态运动-结构振动协同控制等方面进行了相关基础理论研究。主要研究内容及结论如下：
提出了一种直接获取柔性航天器刚柔耦合模态的解析方法——采用一组统一的广义坐标描述航天器姿态运动和太阳翼弹性振动，然后使用边界条件联立求解航天器姿态运动常微分方程和太阳翼弹性振动偏微分方程，快速且准确地计算柔性航天器刚柔耦合模态的振型表达式和频率值。收敛性分析表明，随着模型维数升高，基于假设模态及有限元的离散模型计算的系统频率逐渐收敛于这种解析方法得到的频率精
祝寿佳句确值。相比于传统的假设模态建模方法，所得的刚柔耦合模态可用于建立航天器的低维高精度离散动力学模型，基于此模型设计的低维控制器能够在更短的时间内使用更少的能量完成航天器姿态运动-太阳翼结构振动协同控制的目标。
基于柔性航天器的刚柔耦合效应，将航天器姿态运动表示为刚体模态（类比于太阳翼模态振型）与广义坐标（与太阳翼的相同）的乘积，拓展了广泛用于弹性体自由振动分析的Rayleigh-Ritz法，使其适用于“中心刚体-柔性梁/板”这类刚柔耦合系统的固有特性分析。拓展后的Rayleigh-Ritz法可以给出系统刚柔耦合模态的解析表达式，而且具有良好的收敛性和很高的计算精度。利用此方法，本文全面深入地分析了三轴姿态稳定航天器（安装有一对太阳翼）刚柔耦合模态的特性，给出了判断各阶模态耦合情况的判据。最后还讨论了航天器中心刚体平台的转动惯量以及太阳翼长度对系统模态特性的影响。
针对以蜂窝夹芯板作为基板的太阳翼，采用三层的层合板（包含蜂窝芯层和上下面板）而非单层各向同性板对太阳翼建模，随后基于刚柔耦合模态建立
研究了蜂窝板参数和航天器柔性对系统固有特性的影响；结合输入成形和比例微分控制，设计了算法
简单且适宜于实时控制的协同控制器，能完成具有不同柔性的以及受太阳热流照射的航天器的姿态机动-结构振动协同控制任务，并同时消除可能发生的热颤振现象。
针对受太阳热流照射的姿态机动柔性航天器，采用有限差分法推导了计算太阳翼（含有蜂窝板基板）截面温度分布的显式递推格式；在考虑热应力和热应变的基础上，使用刚柔耦合模态建立了航天器刚-柔-热耦合动力学模型。随后，设计了热-结构双向耦合求解流程，用于计算姿态机动航天器的热诱发刚柔耦合振动响应。数值结果表明：姿态机动航天器的太阳翼会受到时变热载荷作用，由此引发的动力学响应包含准静态变形以及叠加其上的振动分量两部分；如果热流的最终入射角（约等于初始入射角与姿态机动角之和）较大，则具有较小结构阻尼的航天器的热诱发响应可能是不稳定的，此时可能出现热颤振现象；若航天器阻尼较大，太阳翼热诱发动力学响应的振动分量将逐渐衰减，最终仅余下准静态变形。
关键词：柔性航天器；大型太阳翼；刚柔耦合模态；协同控制；热诱发振动
Abstract
车辆代办
Abstract
Flexible spacecraft are usually fitted with solar panels to provide energy for the normal operation. Th
ey are typical rigid-flexible coupling system becau of the coupling between the solar panel elastic vibration and the spacecraft attitude motion. To achieve stronger function and longer life of modern spacecraft, the solar panel size is growing, which leads to more conspicuous rigid-flexible coupling effect. At prent, most of the rearches on dynamics and control of flexible spacecraft rely on the modes of the individual solar panel (i.e., assumed mode). This kind of mode cannot reflect the rigid-flexible coupling effect, which results that tho studies have a certain shortage when ud to study the spacecraft with large-span solar panels. Therefore, it is of great theoretical and engineering value to obtain the rigid-flexible coupling modes (RFCMs) of flexible spacecraft to conduct coupled analysis and cooperative control of spacecraft attitude motion and solar panel vibration.
In this disrtation, the spacecraft with large-span solar panels is studied in detail by analytical and numerical approaches. Approaches are firstly propod to obtain rigid-flexible coupling modes. Then, investigations are conducted relevant to the natural characteristics, dynamic modeling and dynamic characteristics, and cooperative control of attitude-structure motion. The main contents and achievements derived in this disrtation are listed as following.
An analytical method is developed to directly obtain the RFCMs of flexible spacecraft. Describing the spacecraft attitude motion and solar panel vibration with a uniform t of generalized coordinates, th
眼镜的图片e analytical expressions and frequency values of RFCMs can be accurately solved from the rigid-flexible coupling dynamic equations of the continuous system of flexible spacecraft by using the boundary conditions. The convergence study reveals that, as the degree of freedom of discrete models bad on assumed modes or finite elements increas, the frequencies calculated from the discrete models converge to tho from the propod analytical method. Compared with the modeling approach with assumed modes, tho RFCMs can be ud to derive the low-dimensional but high-preci discrete dynamic model of the flexible spacecraft, and the low-dimensional controller designed by using this model can achieve the cooperative control-index for spacecraft attitude motion and solar panel vibration in a shorter time with less input energy.
The Rayleigh-Ritz method widely ud in free vibration analysis of elastic structures is extended to investigate the modal characteristics of the rigid-flexible coupling hub-beam/plate system by reprenting the spacecraft rigid-body motion with the product of rigid-body mode and generalized coordinate. The extended
Rayleigh-Ritz method can compute the analytical expressions and frequencies of RFCMs with high a
ccuracy, excellent convergence and high efficiency. Using this method, numerical simulations are conducted for a three-axis attitude stabilized spacecraft installed with a pair of solar panels. Bad on the analys of tho simulation results, the characteristics of RFCMs are investigated comprehensively, and then a criterion is given to determine the rigid-flexible coupling property of each mode. Finally, the influences of the solar panel length and the moment of inertia of the rigid body on the system’s modes are studied.
Modeling the solar panel compod of honeycomb sandwich panel by a three-layer laminated plate, i.e., the honeycomb core and two facesheets, rather than a single-layer isotropic plate, a high-preci discrete dynamic model which is clor to the real engineering is established by using RFCMs of a flexible spacecraft. Subquently, the effects of honeycomb panel parameters and spacecraft flexibility on the system’s natural characteristics are investigated. Applied the discrete dynamic model and combining the input shaping (IS) technique with the proportional-derivative (PD) controller, a real-time control scheme with simple algorithm is designed to achieve the cooperative control of attitude maneuver and panel vibration for spacecraft with different flexibility and to inhibit the possible occurrence of unstable thermally induced vibration for flexible spacecraft subjected to solar radiation.
For an attitude maneuvering flexible spacecraft under solar radiation, an explicit recursive scheme is derived with finite difference method to calculate the cross-ctional temperature distribution of solar panels compod of honeycomb panel, and the system's rigid-flexible-thermal coupling dynamic model is established in terms of RFCMs by considering the thermal stress and strain. Then, this model is ud to compute the spacecraft thermally induced respons by employing a coupled solving procedure. The numerical results reveal that: the solar panel of a maneuvering spacecraft is subjected to time-varying thermal loading, and the thermally induced respons consist of quasi-static displacement and superimpod vibration; if the final incident angle of heat flux (the sum of initial incident angle and the maneuver attitude angle) is large, the thermally induced respons of a spacecraft with small damping may be unstable and the thermal flutter may occur in this ca; the thermally induced vibration decays with time and only the quasi-static deformation exists finally in the ca of large damping.
吉他弹唱童年
Keywords:Flexible spacecraft, Large-span solar panel, Rigid-flexible coupling mode, Cooperative control of attitude-structure motion, Thermally
induced vibration
目录
目录
栈房
摘要 .......................................................................................................................... I Abstract ...................................................................................................................... I II
第1章绪论 (1)
1.1 课题背景及研究意义 (1)什么地告诉
1.2 柔性航天器动力学建模研究现状 (2)
1.2.1 刚柔耦合动力学建模 (3)
1.2.2 柔性构件连续位移离散 (6)
1.3 柔性航天器动力学特性研究现状 (8)
1.3.1 刚柔耦合动力学特性 (8)
1.3.2 热诱发振动 (9)
1.4 姿态运动-结构振动协同控制研究现状 (12)
1.5 本文主要研究内容 (14)
第2章单轴转动柔性航天器的刚柔耦合模态 (17)
2.1 刚柔耦合模态 (17)
2.1.1 动力学模型 (17)
2.1.2 刚柔耦合模态解析表达式 (19)
2.2 柔性附件连续位移离散 (21)
2.2.1 模态离散 (21)
2.2.2 有限单元离散 (23)
一团乱麻2.2.3 基于离散模型的系统频率方程 (24)
2.3 不同离散方式的对比研究 (24)
2.3.1 有效性验证 (25)
2.3.2 固有特性分析 (27)
2.3.3 姿态运动-结构振动协同控制 (31)
2.4 本章小结 (36)
第3章三轴稳定柔性航天器的刚柔耦合模态 (37)
3.1 解析方法 (37)
3.1.1 模型描述 (37)
3.1.2 动力学方程 (39)
