Control Of Satellite Attitude Maneuvers Using Control Moment Gyros

New satellite missions, such as high-precision earth observation, space surveillance, satellite surveillance, distributed platform, constellation, satellite rendezvous and interplanetary exploration, etc., require large-angle and rapid manuverability. Momentum exchange devices such as reaction wheels can not meet these requirements, since their torque output is rather small compared to control moment gyroscopes(CMG). CMG, with its excellent performance in exchanging mementum, has the ability to fulfill such missions. The miniaturization of CMG makes it possible to mount several CMGs on one small satellite. Some deep research has been done both in theory and application about the decision-making and control for large-angle satellite attitude maneuver using single gimbal control moment gyroscopes(SGCMG). The main contents of this dissertation are summarized as follows:To solve the inherent singularity problem in a system consisiting of several SGCMGs is the most difficult work in the steering logic design. The singularity problem will affect energy consumption and rapid actuation of SGCMG if it is not handled properly. Therefore the model of a satellite using SGCMG is made at first, based on which the distribution and characteristics of singularity in the system of pyramid-type SGCMGs is analyzed. Several singularity avoidance and escape steering strategies based upon gimbal angular velocity are designed and testified to make sure that SGCMGs will produce the comand torque output even in singular gimbal angle configrations.Although the steering logic designed provides an effective means of avoiding and escaping singular states, disturbing torques are innevitably brought in when SGCMGs are close to or reaching the singular states. Therefore, this thesis designs a path planning method based on pseudospectral method, that is, the guidance law to significantly reduce the possibility of reaching the singular states as well as execute the maneuver task even though the initial and final velocities of the satellite are nonzero.Thanks to the guidance law, the maneuver problem is converted to an attitude tracking problem. For attitude tracking control problem of satellite with disturbances and uncertain inertia matrix, a nonlinear satellite model based on error quaternions and velocity errors is builded, according to which we derive the adaptive sliding mode tracking controller to track the desired satellite variables, such as the quaternions, produced by the guidance law. The adaptive control logic can eliminate the influence of parameter uncertainties while the sliding mode control logic is robust to external disturbances. Consequently, the cooperation of these two logic ensues good performance of attitude tracking.