To obtain unambiguous ground truth of water-ice residing in permanently shaded regions of the Moon and to characterise the local regolith, ESA considers a mission involving an instrumented penetrator implanted there by high-speed impact. Released into lunar orbit, the European Lunar Penetrator (ELUPE) Descent Module will autonomously traverse a controlled trajectory to its designated target. The associated attitude control problem involves highly nonlinear large-angle slew manoeuvres and unstable minor-axis spin manoeuvres. To establish a benchmark, a controller based on classical control techniques was designed, verified and tested in a simulation. Legacy control algorithms were implemented and extended. A thruster management function was developed to translate the control commands into thruster actions. For the simulator, accurate models of the descent module and its environment were created.

A Monte Carlo simulation was run to determine the success rate of the ELUPE mission from a descent-and-landing perspective, and to assess the performance of the controller under off-nominal conditions. From the results, it was found that the success rate was 58.5% for a surface slope of 20°, and 74.2% for a slope of 10° or lower. Key factors affecting the success rate were identified to be the centre-of-mass offset and the solid rocket motor thrust misalignment angle. As further constraining these parameters would be unrealistic, it was recommended to modify the thrust curve of the solid rocket motor to improve the success rate.

Analysis of the attack angle and the nutation angle just prior to impact, revealed their success criteria were met in 98.7% and 99.8% of all cases, respectively. These successful results confirmed the attitude control problem could be satisfactorily solved by a 'classical' controller. However, despite its good global performance, the proposed controller was found to also exhibit some serious shortcomings. For this reason, it was recommended to explore the possibilities for a different controller.

To the best of the author's knowledge, this thesis represented the first known attempt at designing a comprehensive controller for a fully actuated, thruster-controlled penetrator mission targeted for an airless body. In addition, it was the first known study to provide insight into the feasibility and success rate of such a mission from a descent-and-landing perspective.

Current and future space missions require agile and reliable spacecraft capable of trailing and keeping the required attitude. Most of the agile spacecraft missions are near-Earth based but some are placed far away from Earth and its influence. One example of such missions is the Athena mission, which requires the spacecraft to perform fast and large-angle attitude slew manoeuvres. Such manoeuvres often imply simultaneous use of multiple actuators such as thrusters and reaction wheels (RWs). A fault in any of these actuators might lead to partial or full damage of sensitive spacecraft instruments. In this research project, a novel model-based Fault Detection and Isolation (FDI) strategy is proposed, which is able to detect and isolate various actuator faults, such as stuck-open/closed thruster, thruster leakage, loss of effectiveness of all thrusters, and change of RW friction torque due to change of Coulomb and/or viscosity factor. Moreover, the proposed FDI strategy is also able to detect and isolate faults affecting the RWs tachometer. The design of the FDI algorithm is based on a multiplicative extended Kalman filter, a generalised likelihood ratio thresholding of the residual signals, and a logic algorithm which unequivocally link the faults to the symptoms. The performance and robustness of the proposed FDI strategy are evaluated using Monte Carlo simulations and carefully defined FDI performance indices. In addition, the influence of faults’ magnitudes, times of fault occurrence, and uncertainties’ magnitudes on the FDI system performance are evaluated. Preliminary results suggest promising performance in terms of detection/isolation times, miss-detection/isolation rates, and false alarm rates. Also, uncertainties on the spacecraft inertia seem to have a negative impact on the FDI performance. In order to fully understand the research project presented here, graduate-level knowledge on rigid body dynamics and kinematics, control theory, and filters applied to estimation might be required. If any of these areas are not known by the reader, it is recommended to read some of the associated literature referenced in the bibliography.

In recent years there has been an increasing interest in formation flying with many lightweight spacecraft, as satellite missions can potentially become cheaper and more flexible. An example of such mission is the Silicon Wafer Integrated Femtosatellites mission, consisting of 100 to 1000 spacecraft with a mass of 0.1 kg. Due to modest control capabilities of the spacecraft and a higher risk of collisions, the requirements on trajectory optimization algorithms are increased. In this thesis a distributed trajectory optimization algorithm is developed which minimizes the required fuel for collision-free reconfiguration trajectories. First, the balance between cooperation in the formation and fuel consumption is investigated: with less cooperation the trajectory optimization algorithm can easily be distributed but the resulting fuel consumption is higher. The problem can also be distributed using dual methods, which can result in the same solution as a centralized algorithm. In literature both dual decomposition and the Alternating Direction Method of Multipliers (ADMM) in consensus form are proposed to solve this specific problem. In this thesis it is demonstrated that the Jacobian decomposition of the Augmented Lagrangian Method outperforms both dual decomposition and ADMM in terms of convergence rate. Furthermore it is shown that this algorithm does also converge in an asynchronous setting. Finally, the synchronous algorithms are significantly accelerated using Heavy Ball acceleration, the Fast Iterative Shrinkage-Threshold Algorithm and Anderson Acceleration.

Nowadays, a satellite is considered as out of range forever once it is launched. This may soon no longer hold thanks to on-orbit servicing. However, the complexity of that topic makes it highly challenging on many aspects. On-orbit servicing using a robotic arm is nonetheless the most investigated technology in order to achieve the berthing of two satellites in orbit. This research project aims at developing a robust controller for the deployment of a robotic arm, which is mounted on a servicer satellite, and for the reaching of a defined handle on the target satellite under mission constraints. While operating the space manipulator, the combined control of both the manipulator and the spacecraft base has a great potential in terms of achievable performance and may bring robotic on-orbit servicing one step closer to reality. In order to reveal the potential of this innovative control strategy, a model of the system dynamics was developed, before being used to design a robust controller. The latter was then successfully implemented in different simulators to test its performance and robustness to a variety of constraints including hardware, environment and mission constraints. The simulations proved that combined control can be used for the entire maneuver. Finally, this controller was successfully implemented in the on-board computer of the on-orbit servicing simulator at the German Aerospace Center. From now on, the maneuver can be completed from the ground control room. The implementation of the robust and combined controller has increased the technology readiness level of the reach phase during an on-orbit servicing mission.