The next step in spaceflight, among others, is the development of a reusable launch vehicle (RLV) suitable for manned missions. All existing designs of re-entry vehicles are either ballistic or low L/D vehicles. In the future, the aim is to develop a high L/D vehicle, because such vehicles experience lower g-loads, making the flight comfortable for manned flights. The development of these vehicles can be useful for both, space (for example, space tourism) and military applications (for example, long-range missiles). Thermal protection system (TPS) is deemed critical to the RLV development, as high L/D vehicles are expected to experience higher thermal loads. Existing TPS solutions are not suitable for this purpose. Therefore, arises the need to find reusable TPS solutions that can sustain the desired thermal loads. Flight testing is a crucial step for developing any hypersonic system, this applies to the TPS design as well. This study aims to investigate the influence of TPS on designing the mission and vehicle for a test flight.
From the various developed and proposed TPS designs in literature, a TPS solution is selected, to meet the requirements identified for future RLVs. The solution is an active cooling concept (cooled metallic TPS), named Enhanced radiation cooling, that uses water as a coolant. Based on a preliminary design investigation, suitable materials are selected and the thickness of the layer is determined. The performance of the proposed TPS design, i.e., the thermal load-bearing ability, is assessed using a transient thermal analysis tool developed for this purpose. The outcome of this analysis proves that the proposed design does not meet the thermal load requirements identified in this study. Nonetheless, the concept shows substantial improvement in performance, it can sustain almost double the heat flux as compared to an uncooled system.
To meet the heat flux requirements, a few modifications in the design are proposed and analysed. These modifications increase the complexity of the system and have other adverse consequences that must be addressed. However, thermal analysis proves that the heat flux requirements, identified at the start of this study, are satisfied. A sensitivity analysis of this modified cooling system shows that the performance of the design is not influenced by material uncertainties. Additionally, the design is found to be robust under varying mission and design parameters, within reasonable bounds. These results can serve as a preliminary input for designing a test vehicle and mission.

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.