太阳帆航天器编队维持和重构的方法研究

针对高面质比航天器可以利用太阳光压进行轨道控制的特点,本文提出一种太阳帆航天器编队构型维持和重构的方法.该方法通过控制主从航天器太阳帆姿态角和反射系数,调整主从航天器之间的光压差,产生抵消编队成员间相对运动受到摄动差或进行轨道机动时所需的连续小推力,从而实现编队构型的维持和重构.仿真结果表明,在主航天器太阳帆的姿态角和反射系数相对固定的条件下,对于太阳同步轨道上的高面质比太阳帆航天器编队,使用滑模控制方法,能够调整编队中从航天器太阳帆的姿态角和反射系数产生推力抵消摄动力影响,达到长期维持太阳帆航天器编队构型的目的;通过开环控制方法,能够调整编队中从航天器太阳帆的姿态角和反射系数产生连续小推力,在较长时间周期内实现编队重构.     

铰链展开式构型航天器设计及其动力学仿真

基于航天器结构的模块化设计概念,设计了一种空间可展开航天器模块化结构构型,即铰链展开式构型。利用虚拟样机技术,建立了模块化本体和太阳翼虚拟样机模型。在空间失重环境下,分析了模块化本体和太阳翼在3种展开顺序下（本体各模块和太阳翼同时展开、太阳翼先展开本体各模块后展开和本体各模块先展开太阳翼后展开）对航天器姿态的影响,同时对比了不同扭簧参数对姿态角和展开时间的影响。仿真结果表明该模型可为未来高机动、多型态、多用途自适应变构型航天器的设计以及空间姿态控制提供技术支持和借鉴。     

射流动量控制器精细建模及其应用

文章研究了新型姿态控制执行机构——射流动量控制器的精细动力学模型.基于流场内的流体特性,分析了粘性使流场速度衰减的规律,通过对流动面各点的速度积分,推导了与工质特性相关的执行机构等效角动量,得到了能够反映射流动量控制器内部特性的精细动力学模型.然后,建立了含有三正交构型射流动量控制器组的航天器姿态动力学模型,并依据该动力学模型设计了姿态控制律.仿真结果表明,依据射流动量控制器精细模型得到的执行机构输出力矩反应速度快、误差小,应用在航天器姿态控制的过程中提高了对航天器控制的效率和精度.     

航天器太阳翼展开锁定分析

研究航天器太阳翼锁定过程的动力学特性。对展开锁定和姿态运动做耦合分析,给出了采用有限段模型划分的耦合动力学方程。针对某一卫星给出了锁定过程中各参量的响应曲线,包括卫星本体姿态以及BAPTA 上所承受的作用力和作用力矩由线。该模型也适用于航天器多体系统的展开锁定分析。     

载人航天器体装太阳电池阵有效发电面积计算方法

航天器的发电能力与太阳电池阵有效发电面积成正比。针对圆柱形载人航天器,提出了一种体装太阳电池阵有效发电面积计算方法。首先,将太阳电池阵沿圆周方向划分为多个子阵,通过坐标变换,计算太阳矢量与每个子阵法线的夹角（即太阳入射角）;然后,将每个子阵面积与对应太阳入射角的余弦相乘并求和,得到体装太阳电池阵的有效发电面积。对不同轨道日照角、不同飞行姿态下体装太阳电池阵有效发电面积进行仿真分析,仿真结果表明:三轴对地姿态下平均有效发电面积为安装面积的25%~32%,通过固定角度偏航可将有效发电面积提高至安装面积的30%~44%。      

日心悬浮轨道混合推进航天器编队构型保持研究

针对混合推进航天器编队日心悬浮轨道保持控制问题进行了研究.首先推导出在日心悬浮轨道附近的航天器编队相对运动方程,考虑到航天器间距离变化值较小且航天器间距离与航天器到太阳的距离的比值为小量,将其在悬浮轨道附近线性化.基于该线性化方程,设计了一种LQR编队控制方式,该控制方式可通过调节太阳帆的姿态及航天器间库仑力的大小对编队构型进行改变或保持,具有响应速度快和控制简单的特点.最后对控制律进行数值仿真,表明该控制方法能实现编队.     

大型复杂航天器的柔性附件展开的动力学分析

分析了复杂航天器在柔性附件展开过程中的几个关键问题,并提出了相应的解决方案,编制了应用软件。通过数值模拟详细分析了其附件的展开过程,讨论的内容包括卫星姿态角和姿态角速度的变化规律、柔性附件的展开运动及其弹性振动等。     

航天器太阳帆板展开过程的最优控制

本文讨论航天器太阳帆板展开过程中主体姿态的最优控制问题. 利用多体动力学与最优控制理论建立数学模型, 考虑系统的非完整约束特性, 提出太阳帆板展开过程的最优控制算法. 通过数值仿真, 表明该方法对太阳帆板展开姿态控制的有效性.      

航天器在轨服务接近策略研究

为解决航天器在轨服务中如何接近目标航天器的问题,研究了服务航天器的接近策略。从Hill方程出发,深入分析了航天器的相对运动,得到了椭圆型、振荡型、跳跃型和飞越型四种基本的相对运动类型;通过组合基本相对运动,提出了异面接近、盘旋接近、全向接近和共轨接近四种接近策略,并对其形成过程、速度增量需求和应用范围等方面进行了分析。分析结果表明:异面接近策略可以节省大量推进剂,但应用范围受到限制;其他策略具有较广的应用范围。     

单框架控制力矩陀螺系统的构型分析和对比研究

根据实际情况限定了大型航天器单框架控制力矩陀螺 (SGCMG)构型分析研究的对象 ;根据构型分析的主要指标 ,对常用的双平行构型、三平行构型、金字塔构型、四棱锥构型、五面锥构型和五棱锥构型等六种构型进行了分析 ,对比了相互的优缺点 ,得出了最优构型为五棱锥构型的结论 ,为大型航天器SGCMG系统选型提供了理论基础     

一种翻滚非合作航天器抵近绕飞避障轨迹规划和跟踪控制方法

针对空间无人在轨服务任务中翻滚非合作航天器抵近、绕飞和避障问题,在目标特征部位本体坐标系,建立了轨道和姿态相对运动模型.设计了抵近和绕飞策略,以抵近轨迹的燃料和时间最优为目标函数,考虑规避障碍物情况,结合动力学和路径等约束条件进行轨迹规划,最后采用高斯伪谱法对连续最优控制问题进行离散转化,对转化后的非线性规划问题进行求...     

Paths not taken – The Gossamer roadmap’s other options

Highly efficient low-thrust propulsion is increasingly applied beyond commercial use, also in mainstream and flagship science missions, in combination with gravity assist propulsion. Another recent development is the growth of small spacecraft solutions, not in size but in numbers and individual capabilities.Just over ten years ago, the DLR-ESTEC Gossamer Roadmap to Solar Sailing was set up to guide technology developments towards a propellant-less and highly efficient class of spacecraft for solar system exploration and applications missions: small spacecraft solar sails designed for carefree handling and equipped with carried application modules.Soon, in three dedicated Gossamer Roadmap Science Working Groups it initiated studies of missions uniquely feasible with solar sails such as Displaced L₁ (DL1) space weather advance warning and monitoring, Solar Polar Orbiter (SPO) delivery to very high inclination heliocentric orbit, and multiple Near-Earth Asteroid (NEA) rendezvous (MNR). Together, they demonstrate the capability of near-term solar sails to achieve at least in the inner solar system almost any kind of heliocentric orbit within 10 years, from the Earth-co-orbital to the extremely inclined, eccentric and even retrograde. Noted as part of the MNR study, sail-propelled head-on retrograde kinetic impactors (RKI) go to this extreme to achieve the highest possible specific kinetic energy for the deflection of hazardous asteroids.At DLR, the experience gained in the development of deployable membrane structures leading up to the successful ground deployment test of a (20 m)², i.e., 20 m by 20 m square solar sail at DLR Cologne in 1999 was revitalized and directed towards a 3-step small spacecraft development line from as-soon-as-possible sail deployment demonstration (Gossamer-1) via in-flight evaluation of sail attitude control actuators (Gossamer-2) to an envisaged proving-the-principle flight in the Earth-Moon system (Gossamer-3). First, it turned the concept of solar sail deployment on its head by introducing four separable Boom Sail Deployment Units (BSDU) to be discarded after deployment, enabling lightweight 3-axis stabilized sailcraft. By 2015, this effort culminated in the ground-qualified technology of the DLR Gossamer-1 deployment demonstrator Engineering Qualification Model (EQM). For mission types using separable payloads, such as SPO, MNR and RKI, design concepts can be derived from the BSDU characteristic of DLR Gossamer solar sail technology which share elements with the separation systems of asteroid nanolanders like MASCOT. These nano-spacecraft are an ideal match for solar sails in micro-spacecraft format whose launch configurations are compatible with ESPA and ASAP secondary payload platforms.Like any roadmap, this one contained much more than the planned route from departure to destination and the much shorter distance actually travelled. It is full of lanes, narrow and wide, detours and shortcuts, options and decision branches. Some became the path taken on which we previously reported. More were explored along the originally planned path or as new sidings in search of better options when circumstance changed and the project had to take another turn. But none were dead ends, they just faced the inevitable changes when roadmaps face realities and they were no longer part of the road ahead. To us, they were valuable lessons learned or options up our sleeves. But for future sailors they may be on their road ahead.     

Adaptive predefined-time control for liquid-filled flexible spacecraft attitude stabilization during orbit maneuver

A predefined-time attitude stabilization for complex structure spacecraft with liquid sloshing and flexible vibration is investigated under input saturation during orbital maneuver. First, the attitude dynamics model of liquid-filled flexible spacecraft is constructed. Meanwhile, the influence of solar panel vibration and liquid sloshing is treated as a disturbance in the controller design. Next, an adaptive predefined-time control scheme is proposed by applying sliding mode control theory. A predefined-time convergent sliding surface and reaching law are designed to ensure the predefined-time fast convergence rate. Furthermore, a novel adaptive algorithm is developed to handle the disturbances from liquid sloshing and flexible vibration, ensuring that the system converges to a small neighborhood of the equilibrium. Additionally, a new auxiliary system is constructed to deal with the effects of input saturation. At last, one simulation case is performed to verify the feasibility and advantages of the proposed algorithm.     

CubeSail: A low cost CubeSat based solar sail demonstration mission

CubeSail is a nano-solar sail mission based on the 3U CubeSat standard, which is currently being designed and built at the Surrey Space Centre, University of Surrey. CubeSail will have a total mass of around 3 kg and will deploy a 5 × 5 m sail in low Earth orbit. The primary aim of the mission is to demonstrate the concept of solar sailing and end-of-life de-orbiting using the sail membrane as a drag-sail. The spacecraft will have a compact 3-axis stabilised attitude control system, which uses three magnetic torquers aligned with the spacecraft principle axis as well as a novel two-dimensional translation stage separating the spacecraft bus from the sail. CubeSail’s deployment mechanism consists of four novel booms and four-quadrant sail membranes. The proposed booms are made from tape-spring blades and will deploy the sail membrane from a 2U CubeSat standard structure. This paper presents a systems level overview of the CubeSat mission, focusing on the mission orbit and de-orbiting, in addition to the deployment, attitude control and the satellite bus.     

CubeSats detumbling using only embedded asymmetric magnetorquers

After deployment from a rocket, a CubeSat is detumbled using magnetorquer rods bringing the norm to the point where the reaction wheels take over to reduce the angular velocity to null. Therefore, utilizing reaction wheels for satellite detumbling at higher initial velocities is vital but they are heavy and occupy significant space on a spacecraft having challenging control. To address this challenge, this paper features a disruptive approach that conducts the control only by the PCB-integrated magnetorquers with various geometries using a diverse non-unity track width ratio. The trace widths are parametrized such that the optimal torque to power dissipation ratio is investigated. The optimizations are then simulated for various geometric distributions and validated through comprehensive measurement setups that establish a framework for selecting the best-case coil configuration according to mission requirements. The detumbling rates of multiple asymmetric coil configurations are compared with the embedded designs in published literature and state of the art. It is found that the proposed asymmetric embedded magnetorquers can detumble the vehicle at high initial angular velocities. Lastly, the simulation results of thermal analysis are validated for selecting the application-specific optimal coils configuration. At the end, the proposed system is compared with the embedded magnetorquers available in the literature and commercial attitude control systems.     

扇形太阳翼重复折展机构运动仿真及其功能试验

传统折叠式太阳翼体积与重量大, 采用一次性展开锁定机构易引起航天器调 姿或变轨时的颤振. 为此提出了一种新型扇形太阳翼重复折展机构. 基于 ProE/Adams联合仿真, 建立虚拟样机模型, 获取不同电机转速下扇形太阳翼转 动导板展开运动参数的变化规律, 对所研制的扇形太阳翼重复折展机构原理样 机进行展开功能试验. 对比仿真与试验结果可知, 在电机允许转速范围内调节 转速, 扇形太阳翼重复折展机构均可在规定时间内完全展开锁定, 具有重复折 展与锁解功能, 且仿真与试验数据高度吻合, 表明其符合设计要求.      

基于误差空间的航天器姿态反步容错控制

提出了一种基于误差空间的航天器姿态反步容错控制方法,以反作用飞轮作为航天器的执行器,在考虑反作用飞轮存在安装偏差及故障的情况下,仍可保证航天器姿态的稳定性。首先,基于Lyapunov稳定性原理,根据系统机械能变化构造了具有普遍性的Lyapunov方程。通过反步递推方法,得到了适用于航天器存在执行器偏差及故障情况的普遍性的容错控制方法;然后,通过误差空间拓扑所得的误差函数描述了势能误差。从几何层面上看,这是描述势能误差的最短路径选择,从而得到了基于误差空间的反步容错控制方法。因此,在对航天器进行姿态控制时,该方法可以迅速调整增益,使得系统姿态误差迅速收敛至零,从而有效减少系统响应时间;最终,通过对考虑执行器偏差及故障情况的航天器姿态控制系统使用不同的控制方法进行数值仿真,验证了该方法能够在执行器故障情况下依然保持系统姿态的稳定,且具备良好的响应速度。