Research On Attitude Control Of Nano-pico Satellites Based On Geomagnetic Control

In recent years,Based on rapid development of microelectronics,new materials and precision manufacturing technology,nano-pico satellites have the outstanding characteristics of light quality,small volume,low cost,strong functions and so on.They have great potential and good application prospect in the civil and military aspects.However,the limitations of the design of pinna satellites make them very strict to the size,power consumption,quality and other requirements of the configuration components of the star,and the selection of the actuator is limited.The magnetic moment equipment which make full use of the earth's magnetic field near the earth has advantages of low quality,low energy consumption,simple structure and high reliability and so on.Based on the magnetic control,the attitude control of pinna satellites without rotating mechanism and jet propulsion has become an important research direction.Therefore,for the above mentioned problems,this paper studies the pina satellite attitude control system based on geomagnetism.Firstly,the working principle of the selected solar sensor,magnetometer and micromechanical gyro is studied in the attitude determination of pinna satellite,and the measurement model is established respectively.In full consideration of the nonlinear factors of the satellite attitude determine system and the storage capacity,computing speed and other constraints of on-board processor,the extended kalman filter method is adopted to improve the sensor data processing and attitude estimation.the two conditions of“sun sensor+magnetometer+gyro”filter in the light zone and“magnetometer+gyro”filter in the shadow zone are studied and analyzed.Simulation shows that when the three measurement information fuses effectively,attitude Angle estimation precision<0.5~°,three-axis angular rate estimation accuracy<0.02~°/s,when the sun sensor is unavailable in the shadow zone,attitude Angle estimation precision<1~°,three-axis angular rate estimation accuracy<0.04~°/s,this level of accuracy can satisfy pinna satellite posing accuracy requirements.Secondly,the working principle and related characteristics of the selected actuator magnetic torquer are studied for the attitude control module of pinna satellite.Magnetic control moment is constrained by magnetic field and can only be generated in the vertical plane of the geomagnetic field intensity,and the magnitude is usually small.Then,the corresponding magnetic control law is designed for different work characteristics and goals of three stages including velocity damping separated from rocket,attitude capture and attitude stabilization.In the initial stage of satellite orbit,the rate of change of the geomagnetic vector in the star system is used to rapidly reduce the angular velocity of the star.The simulation example shows that the algorithm can attenuate the angular velocity to 0.3~°/s in three orbital periods.In view of the large Angle maneuver attitude capture phase,control law based on sliding mode without switching is designed,the simulation shows that the control law has better robustness and practicality,under the initial large Angle deviation attitude Angle can converge to 8~°within five orbital cycles.in view of the small Angle maneuver Triaxial stable stage,according to the characteristics of magnetic control satellite,linear periodic system model of small Leo rigid satellite is established,The LQR theory is applied to the attitude magnetic control system,and the simulation results show that the control strategy designed can stabilize the attitude Angle error to 2~°and satisfy the requirement of attitude stability accuracy.Finally,two control processes are chose from the specific function of satellite attitude control system,including the orientation to earth after reaching satellite orbit and directional switching from the earth to the sun,the overall closed-loop system simulation analysis including three stages switch is made.The simulation results show that the working effect is good when the control law of the three stages is in the same closed-loop system.In both cases,the attitude error can be stable to 3~°,which can meet the requirement of satellite pointing accuracy.