Skin Satellite Attitude Determination And Control System Design And Implementation,

Pico-satellites, based on 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, considering all of these factors, we choose one design scheme that fits the system constraints and mission demands. Meanwhile, we realize the scheme by hardware and test the function as well.In this thesis, mission analysis is firstly executed towards a pico-satellite. According to the mission target, a feasible scenario of ADCS design for the pico-satellite is outlined, that is: three axis magnetometers incorporated with the solar cell arrays as the attitude determination subsystem (ADS), and three magnetic torques incorporated with a momentum biased reaction wheel as the attitude control subsystem (ACS).In ADS design 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. The attitude control system is compared with a micro momentum-biased reaction wheel and three-axis magnetic-coil, to realize the followingfunctions: (1) after release from launcher, by B control law, this mode can damp the body rate down;(2) After that ADCS switches to the three-axis-stabilizing mode. This mode can stabilize the satellite attitude to a small deviation. Besides, we design the parameters of these two apparatuses.The second part of the thesis, the simulations are taken to model the operation mode of pico-satellites. Simulation results reveal that after about 3 orbit periods, the attitude angle velocity can be slow down and realize three-axis stable. Such performance meets the requirements of mission and satellite platform.In the third section, we discuss the realization of low energy-cost, three-axis stabilizedattitude controlling system. The system chooses DSP as main control CPU, an MEMS bias momentum flywheel and a three-axis magnetic-coil as main executing apparatus. The system directly chooses the six solar chip-batteries as solar sensors and a three-axis magnetometer to measure three axis's magnetic strengths. The main control board, which takes charge of gathering information and determining satellite's attitude and computing attitude deviation and driving various apparatuses to adjust and stabilize satellite, is designed based on DSP and other digital circuit techniques.Finally, Ground system simulation is taken to testify the hardware and the controlling methods. By using half physical experiment, we use a simulation computer to model the satellite kinematics and dynamics, model solar vector and magnetic vector as well. Taking the real interface and process, we make a half physical experiment. It is verified that, the attitude deviation is minimized obviously after 2-orbit period. After that ADCS switches to the three-axis-stabilizing mode. This meets our system requirement.