Very Low Earth Orbit (VLEO) improves the geometric performance of optical payloads and increases communication link margins relative to conventional Low Earth Orbit (LEO), but exposes spacecraft to aerothermal heating, aerodynamic disturbances, and large environmental uncertainty. These effects are particularly challenging for CubeSat-class platforms, where mass, volume, and available thermal-control hardware are limited. This thesis assesses the thermal feasibility of a predominantly commercial-off-the-shelf 16U CubeSat platform operating down to a minimum perigee altitude of 130 km.
A simulation-based thermal-analysis framework is developed to construct conservative VLEO thermal cases and propagate them into spacecraft temperatures. Atmospheric, aerothermal, aerodynamic, and Earth-radiation inputs are defined for bounding hot and cold conditions. Analytical free-molecular aerothermal and aerodynamic models are benchmarked against DSMC reference data and used for rapid altitude- and attitude-dependent load evaluation. Earth albedo and infrared radiation are represented with a CERES-derived zonal-harmonics model after comparison with constant and ECSS-based alternatives. The resulting cases are analysed with a reduced-order LTspice thermal network and a spatially resolved COMSOL finite-element model. Morris screening, deterministic design sweeps, and surrogate-based uncertainty propagation are then used to size and verify the passive thermal configuration.
The baseline configuration is non-compliant in the hot case even without internal dissipation, identifying insufficient bus heat rejection as the limiting mechanism. A passive design combining a ram-facing heat shield, high-emissivity side panels, and thermally controlled external interfaces permits a deterministic payload duty cycle of 20 % at 130 km. The propagated maximum structure temperature is centred at 39.7 °C with a standard deviation of 0.85 °C and remains more than 7 °C below the 50 °C structure-to-payload interface requirement. No requirement exceedance is observed within the propagated uncertainty set. The results indicate that passive thermal control of a 16U COTS-based CubeSat is feasible for short-duration operation down to 130 km perigee, provided that interface conductances and subsystem-level thermal behaviour are verified by further modelling and test correlation. ... Read more

Thermal analysis of CubeSats during early design stages is often limited by lack of time, resources and reliable material property data, which leads to large uncertainties in predicted temperatures and potential thermal risks. Additionally, the use of high-fidelity tools makes iterative analysis difficult, especially for small satellite teams.

To address this, this work develops an integrated approach that combines experimental determination of thermo-optical and thermal properties with reduced-order thermal modeling. A database of material properties for commonly used CubeSat components is generated through laboratory testing. Further, simplified thermal models are created and validated against experimental results to capture the dominant heat transfer behavior with reduced complexity.

The combined framework allows faster and more reliable preliminary thermal analysis, helping identify thermal issues early in the design process and improving overall confidence in thermal design. ... Read more

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This thesis focuses on designing and optimizing a shock table using Finite Element Analysis (FEA) to predict and assess Shock Response Spectrum (SRS) levels. The goal is to create a cost-effective,  and accurate shock table capable of replicating launch vehicle shock environments. A structured methodology is adopted, including concept generation, simulation, and design space exploration. The results show that careful tuning of impactor characteristics, fixture geometry, and material selection significantly improves shock table performance. ... Read more

On the Delfi family of satellites, antenna deployment currently takes place using a burn wire system. This requires the antenna to be of the same length as the satellite body. To decouple this length, a new mechanism is needed. Neutrally stable tape springs actuated by shape memory alloys (SMAs) were proposed and studied, a concept previously proposed by Schultz et al in 2008. Thermal analysis, manufacturing tests, and lab testing of SMA actuated tape springs were performed. It was found that the concept can successfully deploy. Satellite integration and more fine-tuning of the mechanism is necessary for an in-flight application. ... Read more

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. ... Read more

The growing accumulation of space debris poses a significant threat to the long-term sustainability of space activities. A drag sail system presents a passive, post-mission disposal solution that utilizes the low atmospheric density in Low Earth Orbit to accelerate the re-entry and disintegration of satellites. This study focuses on the sizing and structural design of a drag sail system based on a preliminary design provided by Demcon High-tech Systems.

A numerical orbital decay model was developed in Python using Tudat libraries to determine the drag sail area that complies with the decay time requirement. The conservative perturbing forces, such as the Earth’s gravitational field (including oblateness effects) and third-body gravitational influences, as well as non-conservative perturbing forces, such as atmospheric drag and solar radiation pressure, were
included. The atmospheric density was estimated using the US76 neutral model and the NRLMSISE-00 dynamic model, with the latter yielding a more accurate result. Furthermore, two aerodynamic flow models: free molecular and continuum flow, were analyzed, showing a minor 4% difference in decay time. The results reveal that an accurate representation of atmospheric density is more critical than the drag coefficient due to the high variability in density predictions between models. Through validation of the model, it was found that using the average outer surface area of the CubeSat as the effective drag area offers a conservative yet reliable estimate for decay prediction.

The natural decay time of the host spacecraft was estimated at 31.4 years. The drag sail was designed to reduce this to 6.3 years, meeting the mission’s fivefold reduction target. Analysis of the area-to-mass ratio indicated a ratio of 54 cm²/kg, corresponding to a drag sail area of 0.2 m², will comply with the decay requirement time.

For the structural design of the booms, inflatable deployable structures were selected due to their compactness and long shelf-life, with strain rigidization as the post-deployment stiffening technique. Localized low rigidization near the boom end caps was observed. The sail configuration limited out-of-plane deflection to 2 mm, ensuring no effective drag area reduction or decay time increase. Stresses on the sail were well below failure thresholds, and its lowest natural frequency was 200 times greater than the highest frequency of the drag load, eliminating resonance risk.

Boom analysis revealed that bending is the primary failure mode. Lower inflation pressures enhanced bending performance, with the lowest pressure studied allowing the boom to withstand over five times the applied drag load before yielding. This significant safety margin accommodates potential imperfections, such as residual creases remaining after rigidization. Additionally, vibrational coupling between
the sail and booms was avoided, thanks to a sufficient frequency separation of four times between the subsystems.

A final full system analysis indicated the deorbit system could survive the mechanical stresses induced by the drag load down to an altitude of 120 km. However, further investigation is needed to account for thermal loading at lower altitudes. ... Read more

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. ... Read more

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. ... Read more

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... ... Read more

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. ... Read more