| 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.
|
PDF Full Text Request