Thermal Protection System Analysis and Sizing for Spaceplane Configurations

Kurzfassung

During its lifetime a space launcher is subjected to extreme heat loads. These can dramatically influence the design of the launcher, when parts of the vehicle heat up to very high temperatures. Every part of the vehicle has a temperature range in which it is functional, thus it is important that these limits are respected. Therefore, space vehicles are equipped with a Thermal Protection System (TPS). A properly designed TPS is vital to every reusable launcher vehicle, as an under-designed system will lead to mission failure, and an overdesigned system to an increase in mass, resulting in an increase in cost. Heat loads are typically the largest during the ascent or decent phase, where the spacecraft flies through the atmosphere. The large velocities combined with the atmosphere cause atmospheric drag, which results in aerodynamic heating. In this thesis the design of a TPS is considered, focused on spaceplane vehicles. The accompanying research question 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. For rocket powered launchers the largest heat loads typically occur during reentry, as in this mission phase the longest time is spent in the atmosphere. 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. Five different TPS areas will be defined, all assigned a passive TPS type. The TPS types are FRSI, AFRSI, TABI, AETB TUFI and CMC. 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 still meeting the given constraints. These constraints are the maximum reusable temperatures of the TPS types, and the limit functional temperature of the underlying structure. The structure is made of aluminum, and has a maximum functional temperature of 450 K. The maximum temperatures of the TPS types are 644 K, 922 K, 1400 K, 1600 K, and 1850 K respectively. The TPS design tool was applied to a reference vehicle. 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.