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.

Splicing is a method for joining Fibre Metal Laminate (FML) panels in which, at the location of the joint, metal sheets are overlapped to form a Damage Tolerant (DT) bond. This joining method is successfully applied in aircraft structures, but has one main disadvantage; it cannot be used in structures with significant double curvature. To qualify FMLs for these specific locations as well, a different joining method is required. With this in mind, the idea of Friction Stir Welded (FSW-ed) FMLs was born. By FSW-ing individual metal sheets together before lay-up, it is possible to create FML structures with the required double curvature. This research was aimed at developing the DT design principles for successful application o f t h i s concept in aircraft structures. In order to reach this goal, a step-by-step DT compliance plan was created. This plan started with the identification of analytical methodologies for predicting the DT performance of the joint. In the following steps, the relations between the DT properties of FSW-ed Glare 3-3/2-0.4 and standard Glare 3-3/2-0.4 were determined experimentally, and used to design different concepts that met the requirements from the compliance plan. Out of the four concepts, the idea with the highest value in terms of cost-effectiveness was determined, which was subsequently compared to other materials typically used in the aerospace industry. Experimental results show that cracks grow faster inand residual strength is lower ofthe FSWed version of the laminate. To bring the DT performance back to the level of the standard laminate, additional layers of material are necessary at the location of the joint. The concept that does this in the most cost-effective way for Glare 3-3/2-0.4, is a joint with an inter-laminar aluminum doubler of the same thickness as the other metal layers. Compared to the spliced version of the laminate, the FSW-concept is significantly more expensive. When looking at the DT performance compared to other materials, only certain carbon-fibre reinforced plastics have better properties.