【作者】 李东；

【导师】 王跃林；

【作者基本信息】 中国科学院研究生院（上海微系统与信息技术研究所） ， 微电子学与固体电子学， 2005， 博士

【摘要】 皮卫星以微电子、微机电、纳机电、精密制造等技术为基础,具有成本低、功能密度高、研发周期短、在轨功能针对性强等优点。 姿态确定与控制子系统是卫星系统的重要组成部分,在一定程度上决定了卫星所能实现的在轨功能。皮卫星平台在质量、体积、功耗、数据存储和运算量等指标上要求严格,研究如何在当前技术水平下,寻找出一种满足系统总体约束和任务要求的ADCS设计方案,是本文的研究目标。 论文首先对“微系统与微器件功能验证皮卫星”进行任务分析,明确卫星总体对ADCS系统的要求,并在比较多种可能的姿态测量与控制方法之后,提出一种可行的皮卫星ADCS子系统总体方案:以“双矢量敏感（地磁强度矢量+太阳方位矢量）”作为基本姿态测量手段,以“俯仰偏置动量轮、俯仰微喷机构组合三轴磁力矩器”构成控制子系统。 在姿态测量子系统设计中,依照结构复用设计思想,首次提出将星体表贴太阳电池阵,复用为全向太阳敏感器解算太阳方位矢量,同时采用商用三轴磁强计和样点卡尔曼滤波算法,解决皮卫星在无角速率敏感条件下的姿态测量和姿态确定问题。理论分析和实验结果表明,太阳矢量测量精度<1.5°（1σ）,磁场强度矢量测量精度<1.275°（1σ）,双参考矢量夹角>31.5°时,可保证双矢量定姿算法精度<3°。启动UKF姿态滤波器后,双矢量可观测时,姿态估计精度<0.5°,角速率估计精度<0.0057°/s;仅磁场强度矢量可观测时,姿态估计精度<3°,角速率估计精度<0.0115°/s。 在姿态控制子系统设计中,依据设计轨道高度（400km）上的干扰力矩水平(气动干扰力矩约3.3×10^-8Nm,磁干扰力矩约1.0×10^-7Nm)和总体对ADCS子系统提出的要求,方案中实现了三种执行器件组合的控制系统,包括:微型偏置动量轮(额定转速8600rpm,额定功耗180mW,可提供偏置动量1.441×10^-3Nms)、俯仰微喷机构（推力69mN,最小推力脉宽10ms）、三轴磁力矩器(开关式控制,额定磁矩输出5.49×10^-3Am²,开启时瞬态功耗83.2mW)。这是国内首次在皮卫星平台上实现的三轴稳定姿态控制系统。仿真结果证明,星箭分离后,ADCS系统首先利用B磁控速率阻尼控制律（入轨模式）,在2个轨道周期内,可将星体角速率从4.7°/s衰减至0.01°/s,平均瞬时功耗320mW;之后,切换至三轴稳定模式,约1个轨道周期后,可将三轴姿态指向稳定到3°以内,该阶段平均功耗270mW;进入任务飞行模式后保持三轴姿态稳定指向,常规干扰水平下磁控平均功耗小于40mW,强干扰条件下磁控平均功耗小于70mW,稳态控制时执行机构总功耗可保证250mW以下。总体对ADCS系统提出的要求基本满足。 转台实验是验证姿态控制器件和控制方法的有效手段。在专为皮星开发的微型气浮台上,通过设计的单轴控制实验证明:电池片功率信号,能获得1.3°的测角精度;动量论转速在4000rpm～13000rpm范围内线性程度最好,能保证最低500rpm/s的加更多还原

【Abstract】 Pico-satellites, supported by advanced technologies such as Micro-Electronics, Micro-Electro-Mechanical Systems (MEMS), Nano-Electro-Mechanical Systems (NEMS), and Precision Machining, has many advantages such as low cost, high density of functionality, less Research&Developement time demanded, and mission-oriented features. Attitude Determination and Control Subsystem (ADCS) is one of the most important subsystem, which partly defines the orbital function of satellite. Pico satellite operates under stringent constraints on mass, volume, power, memory, and computational burden. Therefore, it is necessary to study the ADCS design techniques of Pico satellite under such system constraints and mission demands, without the ignorance of feasibility.In this thesis, mission analysis is firstly executed towards a pico-satellite, whose mission target is to achieve the space validation of some micro devices and systems. After that, the system requirements towards ADCS are specified. By analyzing and comparison of many possible sensing and control methods, a feasible scenario of ADCS design for the pico-satellite is outlined, that is: three magnetometers incorporated with the solar cell arrays as the attitude determination subsystem (ADS), and three magnetic torquers incorporated with a momentum biased reaction wheel and a set of micro propellers in pitch direction as the attitude control subsystem (ACS).In ADS degign section, the solar cell arrays are reused as omni-direction sun sensor to measure the sun vector. Intensity vector of geomagnetic field is obtained by three-axis magnetometers. These two measurements are integrated into a double-reference-vector attitude algorithm and attitude parameters can be computed. When the Sun vector error less than 1.5°, the magnetic vector error less than 1.275°, and the angle between the two reference vectors larger than 31.5°, the total attitude error can be confined to 3°. By introducing the Unscented Kalman Filter (UKF) into the ADS, better performance can be achieved: when double vector is available, the attitude precision better than 0.5° and the angular velocity estimation error below 0.0057°/s can be observed; when magnetic vectoris available only, the attitude precision 3° and the angular velocity estimation error 0.0115° Is can be achieved.In ACS design section, according to the disturbing level at 400km height (the disturbance torque caused by aero drag and residual magnetic moment is about 3.3x1O~8JV7?j andl.0xl0~7vVm respectively) and the system requirements, three attitude actuators are implemented : (1) a micro momentum-biased reaction wheel, mounted along the negative direction of pitch axis, with nominal power dissipation of 180mW and nominal momentum bias 1.441 X 10"3Nms at speed of 8600rpm; (2) a set of micro propeller in pitch, with nominal thrust 69mN and minimum pulse width 10ms; (3) three magnetic torquers, which operate in ON-OFF mode, with maximum magnetic moment ±5.49xlO~3Am2 and power consumption 83.2mW for each . This is the first instance of three-axis-stabilization ACS in Chinese pico-satellite. Simulation results reveal that: (1) After release from launcher, ADCS works under the rate-damping mode firstly. By B control law, this mode can damp the body rate from 4.7° Is to 0.01° Is in 2 orbit periods with averaged power consumption of 320mW. (2) After that ADCS switches to the three-axis-stabilizing mode. By using the nutation damping and precession control laws, this mode can stabilize the satellite attitude to 3° in 1 orbit periods, with averaged power consumption less than 250mW. Such performance meets the requirements of mission and satellite platform.Experiments on air bearing table is very straightforward and efficient to demonstrate the effectiveness of ADCS components and control law. By single-axis air bearing test, the precision of attitude determination using both voltage and current information of multiple solar panels is proved to be better than 1.3°. In 4000-13000rpm range, the momentum wheel has best linearity in voltage-speed relation, and acceleration of 500rpm/s can be ensured, corresponding to control torque of 8.38xlO°./vm to the satellite. MEMS gyro and infrared sensor are also tested. The result has shown that: the Gyro can sensing the rotation rate in effect, the infrared sensor can catch the pass-by of infrared emitter correctly. Most orbital operations of ADCS are validated through the air-bearing experiment.更多还原