Advancements in the space sector have driven the democratization of planetary exploration. As a longstanding target of scientific interest, Mars will consequently witness an increasing demand for surface missions of various sizes. Its thin atmosphere, however, is a challenging environment for entry, descent, and landing (EDL). An incoming spacecraft will encounter extreme heating while experiencing low levels of atmospheric deceleration. This atmospheric constraint places limitations on payload mass. Lightweight approaches to entry vehicle design beyond current technologies are critical to allow safe and precise landings for a variety of Mars missions. Thermal protection systems (TPS) constructed from low-density structural materials with high-temperature capabilities are a promising solution for rigid aeroshells. This is because the need for a separate load-bearing carrier structure can be reduced, thus conserving vehicle mass and internal volume. Among the materials currently available, ceramic matrix composites (CMC) such as C/C and C/SiC are essential to the realization of thermally-resistant lightweight structures. They have attracted international interest for Mars entry applications and offer potential versatility for components within various EDL architectures. Novel ultra-high temperature ceramic matrix composites (UHTCMCs) also emerge as good candidates as their capabilities extend beyond the operational temperature limits of traditional ceramic matrix composites.

This thesis explores the use of (UHT)CMCs in the design of a demonstrator heat shield for a low-mass Mars mission delivering a wind-driven spherical rover. Five CMC-based TPS concepts were proposed. Based on a high-level trade-off, a hot structure solution with internal lightweight insulation was selected due to its mission suitability and potential for minimal weight. A ballistic entry trajectory model was used to define the thermomechanical loads. These loads were used as inputs for thermal and structural simulations using the FEA package Ansys Workbench. The TPS layers were sized for minimum mass and appropriate temperature limits, and thermomechanical stress responses were analyzed.

Based on mass and stress margins of safety, a comparison between sized heat shields utilizing a baseline CMC and a UHTCMC was made across two aeroshells with different size configurations. It was found that a traditional CMC provides a heat shield that does not only save mass, but shows a noticeably higher thermostructural performance compared to a UHTCMC; it is better suited for Mars entry unless a degree of reusability and longer entry times are involved. Within a limitation of vehicle mass and base diameter, heat shields with higher vertex angles are noticeably lighter. This work aims to provide a first step towards the exploration of novel structures for Mars entry, enabling a range of robotic and human missions to the red planet.

In recent times, there has been a growing interest in spending extended amounts of time in outer space. To enable these endeavors, living and storage spaces will have to be much larger than current technologies can transport and set up in space. An interesting solution to this problem is the usage of a deployable structure that occupies a small volume in a launch vehicle but can be deployed to larger volumes in space.

This thesis explores the concept of origami–known for its compactness, ease of deployment, scalability, and structural integrity– to create a deployable structure by developing a technology demonstrator that fits in a 12U CubeSat. Engineering origami has been used across many fields such as medicine, architecture, and most recently, space. For space applications like instrument booms and antennas, it is used to save space in the satellite housing them. The current work aims to translate this space-saving strategy to a larger application by creating an origami-inspired small-scale deployable structure that, if successful, can be scaled up for different purposes.

Based off requirements set by the mission, research objective, and constraints, four design
concepts are proposed. For the chosen concept, suitable materials and a compatible manufacturing technique are chosen. The structure is then modelled and prototyped to test its feasibility and determine the configuration that provides maximum compactness and inner volume–two criteria that are critical for sizeable modules to be transported in rockets with limited space. Finally, the structural performance of the deployable unit during folding and deployment is studied.

This thesis work lays the foundation for developing Kresling origami-based deployable structures with insights into optimal configurations, material and manufacturing options, and the development of a parametric model to rapidly check the geometric feasibility of a proposed structure.

As high pointing accuracy spacecraft are being subjected to more stringent requirements on their micro-vibration environment, the reduction and mitigation of these disturbances has become of great importance. The reaction wheel assembly significantly impacts this environment, and as such improvements in bearing technologies have the greatest potential for performance gains. A current line of investigation are magnetic bearings, which generate a soft suspension mechanism thus damping vibrations and eliminating bearing wear. This study aims to explore the concept of Multipole Magnetic Bearings, focusing on radially segmented Halbach design, which introduce the ability of controlling the external magnetic field of the bearing to limit its effect on adjacent equipment.

Maintaining a high speed, secure and robust connection to the world is of paramount importance
in modern society. This proved especially true during a recent volcanic eruption in the small island nation of Tonga, where the only optical fibre line to the country was severed due to the cataclysm, completely disconnecting the island from the rest of the world. The lack of communication made the disaster relief to the eruption more complex. The Emergency Communications Node (ECN) can be deployed within 48 hours to enable alternative means of communication that replace damaged communication lines...

At the Technical University of Delft(TU Delft), a Kirigami inspired method of spirally deploying a coiled up band into a parabolic reflector is being studied for its use in space.
This thesis performs a thermo-elastic analysis on the spiral dish Antenna (SDA) using a combination of ESATAN-TMS, Abaqus, and python scripting. After analysing a reflector for different low earth orbital cases in ESATAN, the heat fluxes are transferred to Abaqus to perform a second thermal analysis. A base SDA is modelled with an Aluminum zipper and Carbon-fiber-reinforced polymers (CFRP) band. The SDA is shown to experience temperatures ranging from 175K to 375K with temperature changes of 100K in less than 100s.
The reflector can experiences maximum temperature deltas of 93K across the reflector at a single moment during orbit. These temperature changes cause the SDA to have a maximum displacement of 20mm with a maximum root mean square (RMS) of 6mm. This thesis also shows the creation of a cross pattern on the SDA most likely created by the coefficient of thermal expansion (CTE) mismatch between the zipper and band material. An improved SDA is modelled after an initial parametrization, showing an improvement to the displacement field with the max RMS being around 1mm.
This thesis shows how the base model of the SDA deforms into an cross pattern and experience high displacement regions near the end of the spiral interface. This version of the SDA is shown to not be able to perform in space. A short parametrization analysis with regards to materials choice, spiral geometry, and
thickness does show a path in which the SDA can perform similarly to a normal reflector with the same design parameters.