Design of Attitude Control System for satellites with flexible appendages

Nowadays, spacecrafts are employed for a wide variety of missions: Earth's observation and monitoring, deep universe investigation, communication and much more. For the success of these missions, it is essential that the satellite is correctly oriented during space flight and maintains this orientation with high pointing accuracy in spite of various perturbations. This objective is pursued by Attitude Determination and Control Systems (ADCS), which is responsible to determine the true attitude and to control the actuators to track desired orientation. In contemporary space missions, the design of ADCS is complicated by the fact that modern satellites are equipped with different large devices featuring low stiffness and moving parts, such as antennas and solar arrays. In this case, the effects of these devices on the overall system dynamics must be taken into account and the rigid-body model of attitude dynamics must be dismissed. Indeed, the coupling between the flexibility of the structure and the attitude dynamics causes vibrations of the flexible bodies and the internal torques acting on the satellite which disturb the attitude dynamics. Hence, the object of this work is the design of a three-axis Attitude Control System (ACS) to adequately handle the coupling effects due to the flexible appendages. The satellite model is based on ICEYE-X1, which was launched in January 2018 and consisted in a proof-of-concept satellite mission of the Finnish startup ICEYE, whose aim is a Synthetic Aperture Radar (SAR) equipped commercial constellation of 48 satellites. The simplified structural model of the spacecraft has been built in MSC Patran and consists of a main rigid body to which the flexible appendages are attached. Next, a Finite Element Method (FEM) analysis is performed by solver MSC Nastran to evaluate natural modes of the structure. The latter are the used to build an appropriate mathematical model for flexible spacecraft attitude dynamics, that captures the coupling effects between structural dynamics and attitude dynamics. Different spacecraft configurations are analysed and the above mathematical model is derived for each of them. This allows to establish the parameters which mainly impact the magnitude of the coupling effects between flexible and attitude dynamics. The comprehension of these aspects is crucial for generalising the procedure for estimating the significance of the coupling effect in other types of satellites. The second part of this thesis is focused on the design of a robust attitude control technique, i.e the Quaternions Feedback Control (QFC), to handle external perturbations and dynamics uncertainties. The control strategy is tested in a three degree-of-freedom orbital simulator including the mathematical model for attitude dynamics of the flexible spacecraft and the control algorithm. Extensive numerical simulations are performed in Matlab/Simulink environment to study the realization of inertial pointing maneuvers with different spacecraft configurations. The results show that SAR panels heavily influence the dynamics response of the spacecraft during the attitude maneuvers. This is most evident in the first seconds of simulation, when the transient excites the coupling between attitude and flexible dynamics and vibrations reaching their maximum values. However, the QFC controller proves its ability to counteract and nullify these disturbances by providing an appropriate attitude control actions.