During its lifetime a space launcher is subjected to extreme heat loads. To protect space vehicles from these heat loads and their resulting temperatures they are equipped with a Thermal Protection System (TPS). Heat loads are typically the largest during flight through the atmosphere. The research question addressed in the thesis is: how robust is a thermal protection system design for a spaceplane wing-body configuration to variations with respect to the design parameters or trajectory taking into account heat transfer through radiation and conduction in three dimensions? To answer the research question a tool was developed which is capable of designing a TPS and optimizing the insulation layer thickness. Furthermore, a trajectory simulation was made for the reentry phase. From the trajectory specifics the heat flux over the vehicle over time could be generated, which is the source of the temperature increase. In the TPS design tool the first major task is to divide the vehicle into different TPS areas, based on the temperatures that are experienced at each of these areas of the skin surface. A thermal analysis is performed, taking into account both conduction in three dimensions and radiation to outer space as well as to the inner subsystems of the vehicle. From this analysis the maximum experienced temperature can be deduced for each area. When the TPS design is found, it is aimed to optimize the insulation layer thicknesses in all TPS areas. The goal is to find a design that is as light as possible, thus with a minimum insulation layer thickness, while not exceeding the maximum temperatures of the TPS types and underlying structure. A sensitivity study was performed to investigate the robustness of the TPS design resulting from the tool. The performance of the TPS design was tested when small changes were made to it, for the nominal reentry trajectory of the reference vehicle. Furthermore the TPS designs performance was analyzed for small changes in the trajectory. From the analysis of the results of the developed tool it was found that a TPS design can be developed for a simple wing-body configuration, under the specified conditions. However, the functionality of the TPS design is limited. Improvements must be made to the developed tool to increase its performance, so that it can come to an acceptable TPS design. It is suspected that with suggested improvements the tool will work properly, and a well functioning TPS design can be made. However, further research is required to ensure this.

Spacecraft are under constant threat of structural damage from hypervelocity impacts by micrometeoroids and orbital debris. To bolster the shielding used for protection against these impacts, ballistic materials can be employed. Aramid-based materials are currently used aboard the International Space Station (ISS), but Ultra High Molecular Weight PolyEthylene (UHMWPE) fibres are a common alternative for ballistic protection on Earth. In this report, the suitability of UHMWPE-based composites for spacecraft impact shielding is investigated. Hypervelocity impact simulations using smoothed particle hydrodynamics discretisation form an essential part of the design and analysis of such protection systems. Two formulations of nonlinear orthotropic hydrocode models are proposed for this purpose, which are validated using footage from hypervelocity impact experiments on Dyneema® HB26 targets. One of the proposed models yields good prediction of residual impactor velocities, generally being within 10% of experimental data. The other reproduces both residual velocities and debris cloud shape well for the highest considered impact velocities, but suffers from decreased performance as the ballistic limit is approached. Numerical comparison between UHMWPE- and aramid-based composites shows comparable ballistic performance for the considered cases.

The space environment is ever-changing with space structures getting larger and the orbits getting increasingly crowded with time. This creates a need for removal of large defunct satellites to avoid the disastrous Kessler syndrome, which poses a major threat to the future of space exploration. This research examines the dynamics and control involved in the active removal of a large space debris - Envisat. European Space Agency's e.deorbit mission aims to de-orbit Envisat using a chaser satellite, which synchronises, docks, detumbles and deorbits it. The presence of large flexible appendages makes the configuration prone to elastic perturbations leading to complex dynamics that cannot be represented using rigid body dynamics. Therefore, a unique multibody approach based on the absolute interface coordinates in the floating frame formulation is used to model the Flexible Multibody Dynamics. The novel method proves to provide a good balance between computation time and efficiency for the control application. The controllability characteristics of two phases of the e.deorbit mission are analysed using a linear PD controller and an Incremental Nonlinear Dynamic Inversion controller. For the first phase, both controllers successfully synchronise the chaser with the target debris tumbling at the rate of 3.5 deg/s about all axes. However, during the detumbling phase, the large appendage (14.2 m) in the stacked configuration introduces complex dynamics, which could not be stabilised completely by applied controllers. Nonetheless, interesting relationships could be established between the dynamics and control of the system, which will facilitate robust control design in future work.