Active Nutation Damping Utilizing Spacecraft Mass Properties

Active nutation damping generally requires a unique element to provide the necessary damping torques, plus a nutation-sensing system to properly monitor and control the active element by closed-loop action. This paper describes and analyzes, by computer- assisted mathematical techniques, a nutation damper for a dual-spin spacecraft which does not require an active ?element?. Damping is achieved through the inherent cross-axis torques generated by the spacecraft products of inertia and the existing motor controlling the relative spin rates of the two dynamic components comprising the spacecraft.     

Attitude Stabilization and Control of Dual-Spin Spacecraft

This invention relates to an attitude stabilizing system for adrodshaped satellite in which stabilization is achieved by use of a spinning ing energy dissipator. The system can operate by being mounted on the rotor of a dual spin satellite of by being connected to the main body of the satellite through a motor and bearing assembly. In the latter situation, the stabilizer acts both as a nutation damper and a momentum source and thereby eliminates the need for an additional rotor element to provide the spacecraft momentum. The spinning energy dissipator may be designed to use four identical al wheels mounted coaxially with and at the ends of the arms of a cruciform structure. The structure may then be spun about the axis perpendicular to the arms, to provide angular momentum along the spacecraft spin axis. In the event of any spacecraft nutation, n, the wheels experience the torques produced by inertial forces. The necessary damping and restoring (spring) torques are provided by means of a torsional arrangement built around the center of each wheel. In order to provide damping at lower threshold of spacecraft nutations, the dissipator wheels may also include viscous ring dampers mounted on or around each wheel.     

The Use of a Spinning Dissipator for Attitude Stabilization of Earth-Orbiting Satellites

This paper considers a novel attitude stabilizing system which acts both as a passive nutation damper and a momentum source for an Earth-orbiting satellite. The system may be designed to use four identical wheels mounted coaxially with and at the ends of the arms of a cruciform structure. The structure may then be spun about the axis perpendicular to the arms to provide angular momentum along the spacecraft spin axis. The wheels experience torques produced by inertial forces. The necessary damping and restoring (spring) torques are provided by means of a torsional arrangement built around the center of each wheel. The performance characteristics of the proposed stabilizing unit are evaluated by considering spacecraft parameters used in one phase of development of the Communications Technology Satellite (CTS). The results obtained for a specific design of the stabilizer indicate that global stability of the desired attitude motion of the spacecraft can be guaranteed with associated damping time constant as low as one second.     

Rotational maneuver of nonlinear uncertain elastic spacecraft

The question of attitude control and elastic mode stabilization of a spacecraft (orbiter) with beam-tip-mass-type payloads is considered. A three-axis moment control law is derived to control the attitude of the spacecraft. The derivation of the control moments acting on the spacecraft does not require any information on the system dynamics. The control law includes a reference model and a dynamic compensator in the feedback path. For damping out the elastic motion excited by the slewing maneuver, an elastic mode stabilizer is designed. The stabilization is achieved by modal velocity feedback using force and torque actuators located at the payload end of the elastic beam. Collocated actuators and sensors provide robust stabilization. Simulation results are presented to show that rotational maneuvers and vibration stabilization can be accomplished in the closed-loop systems despite the presence of model uncertainty and disturbance torque in the system     

A Nutation Damper for a Dual-Spin Spacecraft

This paper is concerned with a high-performance nutation damper for use in dual-spin spacecraft. The damper described here makes use of a wheel of uniform mass distribution. The wheel experiences torques which are produced by inertial forces. Damping and restoring (spring) torques are provided by means of a torsional arrangement built around the center of the wheel. The performance characteristics of the above nutation damper are evaluated by considering the small astronomy satellite (SAS-A) spacecraft parameters, and the results for an optimum design of the damper are presented. Finally, the result for the optimum performance is checked with the help of an approximate analysis.     

Nutation instability of spinning solid rocket motor spacecraft

The variation of mass,and moment of inertia of a spin-stabilized spacecraft leads to concern about the nutation instability.Here a careful analysis on the nutation instability is performed on a spacecraft propelled by solid rocket booster (SRB).The influences of specific solid propellant designs on transversal angular velocity are discussed.The results show that the typical SRB of End Burn suppresses the non-principal axial angular velocity.On the contrary,the frequently used SRB of Radial Burn could amplify the transversal angular velocity.The nutation instability caused by a design of Radial Burn could be remedied by the addition of End Burn at the same time based on the study of the combination design of both End Burn and Radial Burn.The analysis of the results proposes the design conception of how to control the nutation motion.The method is suitable to resolve the nutation instability of solid rocket motor with complex propellant patterns.     

The IMAGE Observatory

The Imager for Magnetopause-to-Aurora Global Exploration (IMAGE) mission will be the first of the new Medium-class Explorer (MIDEX) missions to fly. IMAGE will utilize a combination of ultraviolet and neutral atom imaging instruments plus an RF sounder to map and image the temporal and spatial features of the magnetosphere. The eight science sensors are mounted to a single deckplate. The deckplate is enveloped in an eight-sided spacecraft bus, 225 cm across the flats, developed by Lockheed Martin Missiles and Space Corporation. Constructed of laminated aluminum honeycomb panels, covered extensively by Gallium Arsenide solar cells, the spacecraft structure is designed to withstand the launch loads of a Delta 7326-9.5 ELV. Attitude control is via a single magnetic torque rod and passive nutation damper with aspect information provided by a star camera, sun sensor, and three-axis magnetometer. A single S-band transponder provides telemetry and command functionality. Interfaces between the self-contained payload and the spacecraft are limited to MIL-STD-1553 and power. This paper lists the requirements that drove the design of the IMAGE Observatory and the implementation that met the requirements.     

克服旋转飞行器螺旋运动的方法研究

旋转飞行器虽然具有内在稳定性，但在飞行过程中一些不确定因素会导致飞行器产生螺旋运动，而非理想的关于角动量矢量纯粹的旋转运动。结合旋转飞行器自身的特点，借鉴自旋卫生的被动章动阻尼的方法，提出了一种通过主动控制飞行器内部质量块的运动使得飞行器碑的惯量主轴发生偏移，惯量矩阵由原来的对称阵变为含有非零非对角元素的矩阵，从而改变飞行器的姿态角运动来克服螺旋运动的方法，这将有助于提高旋转飞行器的资态控制精度。以某型旋转飞行器为背景进行仿真研究，计算结果证实了该方法的有效性和可行性。     

Parameter analysis of PAF for whole-spacecraft vibration isolation

Whole-spacecraft vibration isolation, which is implemented by modification of the existing PAF (payload attach fitting), is a direct and effective approach toward improving the dynamic environment that a spacecraft experiences during its journey to the orbit. In this paper, based on Craig–Bampton component modal synthesis and theory about modal effective mass, both the condensed model and the simplified model of the whole-spacecraft vibration isolation system are obtained. By these models, effects of the PAF's parameters, i.e. stiffness and damping, on the transmissibility from the bottom of the PAF to the bottom of the spacecraft and acceleration response of the bottom of the spacecraft are analyzed. Results show that merely increasing the damping of the PAF can effectively attenuate the peak transmissibility, decreasing the stiffness of the PAF can further improve the vibration isolation performance, and can avoid resonance with the launch vehicle by adding enough damping in the PAF. Furthermore, the natural frequency of first lateral or longitudinal mode and their peak transmissibility of the spacecraft-PAF structure can only be estimated by the modal effective mass and the residual mass rather than the rigid body mass of the spacecraft.     

太阳帆航天器被动姿态控制研究

针对由有效载荷、太阳帆和4个控制叶片所组成的太阳帆航天器系统,从物理模型出发,利用欧拉方程建立了姿态动力学模型,并通过数值仿真对太阳帆航天器基于控制叶片的对日定向性能进行了研究。研究结果表明,在行星际飞行中太阳帆航天器可以通过使控制叶片保持偏置来保持姿态稳定。     

The IBEX Flight Segment

IBEX provides the observations needed for detailed modeling and in-depth understanding of the interstellar interaction (McComas et al. in Physics of the Outer Heliosphere, Third Annual IGPP Conference, pp. 162–181, 2004; Space Sci. Rev., 2009a, this issue). From mission design to launch and acquisition, this goal drove all flight system development. This paper describes the management, design, testing and integration of IBEX’s flight system, which successfully launched from Kwajalein Atoll on October 19, 2008. The payload is supported by a simple, Sun-pointing, spin-stabilized spacecraft with no deployables. The spacecraft bus consists of the following subsystems: attitude control, command and data handling, electrical power, hydrazine propulsion, RF, thermal, and structures. A novel 3-step orbit approach was employed to put IBEX in its highly elliptical, 8-day final orbit using a Solid Rocket Motor, which provided large delta-V after IBEX separated from the Pegasus launch vehicle; an adapter cone, which interfaced between the SRM and Pegasus; Motorized Lightbands, which performed separation from the Pegasus, ejection of the adapter cone, and separation of the spent SRM from the spacecraft; a ShockRing isolation system to lower expected launch loads; and the onboard Hydrazine Propulsion System. After orbit raising, IBEX transitioned from commissioning to nominal operations and science acquisition. At every phase of development, the Systems Engineering and Mission Assurance teams supervised the design, testing and integration of all IBEX flight elements.     

Momentum transfer-based attitude control of spacecraft with backstepping

A nonlinear control law design based upon the backstepping approach is addressed for attitude maneuver control of spacecraft by momentum transfer (MT) in the presence of disturbance. For MT, a traditional method usually applies constant torque as an input, which tends to produce significant residual oscillation. Enhanced methods such as optimal control can somewhat reduce the residual oscillation, but may not be enough for minimum residual motion. Feedback linearization technique can drive the final nutation angle small enough, but it is rather sensitive to parameter uncertainty. The proposed method here takes advantage of nonlinear control approach with small steady-state nutation angle. Sensitivity about parameter uncertainties by feedback linearization can be reduced by the backstepping technique. Stability of the resulting control law is guaranteed by the Lyapunov stability theory. Boundedness of the control law is presented to validate practical merit of the proposed control law.     

基于离子电推进系统的航天器有效载荷能力分析

为预估与提高航天器有效载荷能力,结合航天运输系统理论与离子推力器放电模型,对深空探测任务中以离子电推进系统为主要动力来源的航天器有效载荷能力进行了分析。通过理论推导,构建并揭示了有效载荷分数与深空探测任务参数和电推进系统性能参数的函数关系与潜在联系。结果表明:动力装置单位质量越小,航天器所能达到的最佳有效载荷分数越大;有效载荷分数的高低与离子引出份额、原初电子利用率参数的大小以及任务时间的长短呈正相关;当离子电推进系统可以达到更高的载荷比时,则需要更高的工质利用率作为支持。     

卡尔曼滤波在自旋卫星章动控制中的应用

细长体结构卫星的自旋运动是不稳定的,章动有发散趋势,须进行实时的章动监视,且章动角的测量值包含有误差.通过引入一个离散控制过程的系统,介绍了卡尔曼滤波器的基本原理,建立了自旋卫星自旋运动简化的动力学模型,给出了运动方程和观测方程,详细地推导了适用于自旋卫星章动控制的卡尔曼滤波处理过程,利用该模型对某自旋卫星章动进行仿真计算,并对姿态章动联合控制和应急章动控制情况下的章动控制应用情况进行分析,得出结论：基于卡尔曼滤波的卫星章动角计算的方法正确、结果准确.     

Electrical power distribution on space based radar satellites

The radar payload on a space-based radar (SBR) satellite could require tens of kilowatts of power distributed to many small loads over a large area. This poses special problems for the power distribution and control system (PDCS). A study that examined the power requirements of an SBR spacecraft is reported. A baseline prime power system, generating about 30 kW, was derived. The proposed distribution network would transmit 240 V at 20 kHz. The voltage would be downconverted in one converter for about 100 transmit/receive modules. The design considerations are discussed, and the baseline PDCS is described     

基于正负刚度并联的低频隔振器研究

为了能够抑制微振动对在轨运行航天器造成的不利影响,提出了一种由螺旋柔性弹簧（SFS）与磁性弹簧并联构成的低频隔振器;其中,螺旋柔性弹簧作为正刚度元件为系统提供承载能力,并通过有限元建模对其轴向刚度特性进行了数值分析;由两两相吸、同轴配置的3块环形永磁体构成的磁性弹簧为该隔振系统沿轴向提供负刚度以降低系统的固有频率;基于等效电流模型分析讨论了该磁性弹簧的磁刚度特性,并在平衡位置附近对磁刚度作了线性近似。通过分析所设计的低频隔振器与其相应线性隔振系统的位移传递率,比较了2种隔振器的隔振性能。研究结果表明：所提出的低频隔振系统能够有效地隔离低频微幅振动,降低系统的固有频率,进而拓宽了隔振频带;同时,进一步改善了隔振系统在共振区域附近的阻尼特性,有效地降低了共振峰值。     

挠性航天器动态输出反馈控制中的“溢出”管制方法

研究挠性空间飞行器控制系统鲁棒设计的动态输出反馈控制系统中的“溢出”管制问题,给出了抑制“溢出”的正交空间法和参数优化方法,而且通过在设计指标中引入对“溢出”的管制项,使控制系统最佳设计和“溢出”管制结合起来,达到了减少“溢出”,改善系统鲁棒性的目的。     

Experimental study in adaptive tracking control

The authors describe an experimental study of adaptive pointing and tracking control for flexible spacecraft conducted on a complex ground experiment facility. The algorithm used is based on a multivariable direct model reference adaptive control law. Several experimental validation studies performed using this algorithm for vibration damping and robust regulation are extended by addressing the pointing and tracking problem. As is consistent with an adaptive control framework, the plant is assumed to be poorly known to the extent that only system level knowledge of its dynamics is available. Explicit bounds on the steady-state pointing error are derived as functions of the adaptive controller design parameters. It is shown that good tracking performance can be achieved in an experimental setting by adjusting adaptive controller design weightings according to the guidelines indicated by the analytical expressions for the error     

航天器控制系统智能健康管理技术发展综述

健康管理作为智能自主控制亟待突破的关键技术之一,是提升航天器安全可靠稳定运行能力的有效手段。结合人工智能技术的发展趋势,基于前期已建立的新型航天器智能自主控制系统通用架构,详细综述航天器控制系统的智能健康管理技术现状与发展趋势。首先,根据现有航天器设计、研制和在轨的具体情况,梳理出航天器控制系统健康管理技术所面临的挑战;然后,分别从故障预警、故障诊断和寿命评估3个方面,详细阐述基于人工智能的健康管理技术研究现状及其在航天领域的应用情况;最后,提炼出航天器控制系统健康管理技术的发展方向。