Thermal Mathematical Models (TMMs) are used to predict the thermal behavior of satellite structures in orbit. However, due to inherent uncertainties in these models, physical testing is necessary to achieve reliable predictions. While these tests are critical, they often introduce uncontrolled uncertainties, such as heat leaks and measurement errors, making the correlation process complex and time-consuming. To address these challenges, this thesis proposes an alternative testing methodology that reduces uncertainties in thermal test data by using the phase shifts between temperature responses from oscillatory heat loads for TMM correlation. By comparing the measured and predicted phase shifts, thermal model parameters such as conductance or capacitance can be effectively correlated. The results demonstrate that this methodology is largely insensitive to uncontrolled conductive heat leaks, allowing the correlation process to focus only on key thermal parameters. This approach improves the reliability of thermal test data and streamlines the correlation process. Moreover, its potential application in ambient conditions offers a promising testing solution for early-phase model correlation.

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.

The use of small satellites, enabled by the standardization of the CubeSat specifications and miniaturization in electronics, has seen a rapid increase in the past decades. The low-cost and short development time of these satellites has made them an attractive option for both commercial and academic applications, making space exploration more accessible. However, these small satellites are prone to failures, leading to lost scientific potential. Mitigation of these failures forms the motivation for this thesis. Recent advances in neural networks have shown promise in the field of anomaly detection. The black-box nature of such models, however, makes it challenging to understand the reasoning behind their predictions.

Constraining the data-driven models with known physics can not only help us understand the reasoning behind their predictions, but also ensuring the model is consistent with the real-world behavior of the system. The work presented in this Master's thesis aims to demonstrate the advantages of such first-principles neural networks over purely data-driven models in thermal behavior modeling of small satellites. Baseline performance of data-driven Long Short-Term Memory (LSTM) networks is established using FUNCube-1 telemetry data, quantifying the temperature prediction accuracy of the models under ideal conditions. The limitations of these models, especially with sparse data, are then investigated, to highlight the need for more robust models.

First-principles models, based on a physics-informed curve-fit and simplified thermal network models, are then developed to constrain the data-driven model predictions. The first-principles models are shown to be more robust to sparse data, with the predictions on data not seen during training being more consistent with the real-world thermal behavior of the satellite. Methods to relate the first-principles model parameters to the physical properties of the satellite are also proposed and explored, to help extract the evolution of the thermal behavior of the satellite over time.

This thesis presents the development, analysis, and testing of a comprehensive testing methodology for determining the mass, Center of Gravity (CoG), and principal Moments of Inertia (MoI) of small rocket stages. The methodology comprises five sequential tests, encompassing both static and dynamic procedures. The two static tests introduce innovative adaptations of established techniques. The multi-point weighing (MPW) method for the measurement of mass and horizontal CoG coordinates is modified into a suspended version, broadening its applicability, while the suspension method, measuring the CoG height, evolves into the Bifilar Suspension (BS) method, enhancing robustness and safety with the inclusion of an additional rope. The dynamic tests measure the body's MoI values, with the Bifilar Pendulum (BFP) method for roll MoI and the Compound Pendulum (CP) method for pitch and yaw MoI. Efficiency and cost-effectiveness are integral to the methodology, resulting in a streamlined testing campaign requiring minimal time and resource investment.

An analytical uncertainty analysis assessment explores the propagation of uncertainties in the employed methods, highlighting the impact of several parameters on the combined uncertainty on the results. Additionally, an analytical expression for the amplitude-dependent errors in dynamic tests is derived, providing a useful tool to predict such nonlinear effects. A simulation study numerically verifies the results from the uncertainty analysis, as well as the solution equations used for the methods.

The methodology's validation is carried out through three consecutive test campaigns. The results demonstrate the capability of the static tests to consistently determine mass and CoG coordinates with limited uncertainties. The BFP method achieves satisfactory accuracy, although unexpected deviations from the numerical predictions are observed. As for the CP method, multiple factors exert a large influence on the accuracy of the final results. Among the ones analyzed in this work are: the length of the ropes, the radius of gyration of the body, and the accuracy in the frequency measurement. Moreover, in both dynamic tests the type of suspension system is found to have an effect on the accuracy of the measurements.

While not all the intended objectives have been achieved, this thesis contributes to the understanding of testing methodologies for rocket stages, and offers insights into achieving accurate and precise results with simple and cost-effective methods.

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.

In this work, it is demonstrated that kirigami deformation behaviour can be controlled by tailoring the anisotropic material that is used to create the kirigami structure. The ultimate goal is to find a relationship between the material pattern and the deformation behaviour of the kirigami structure. An anisotropic thermoplastic polyurethane carbon fiber material was steered by using the fused filament fabrication method. Kirigami samples were manufactured and tested with varying geometry and material orientation. Each tested sample corresponded to a finite element (FE) model in Abaqus/CAE with identical features. Results show that the FE model is able to accurately simulate the impact of material orientation on the deformation behaviour of the kirigami structure. The FE model is validated by comparing the kirigami tip deflection and -shape to experimental data. Results show that the shape of fibre steered kirigami can be estimated by following the thin plate buckling theory.

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.

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.

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.

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.

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.

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.

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...

The need for Earth observation telescopes with high spatial and temporal resolution is constantly increasing, however, current state-of-the-art telescopes are costly. The Deployable Space Telescope project at TU Delft aims to provide a solution with an Earth observation telescope that has the same optical performance as today's best ones, but at a reduced cost. A baffle is required to surround the telescope to provide stable thermal environment, limit stray-light, and protect the optical components from debris. A deployable baffle consisting of pantographic arms has been designed that has only one degree of freedom, therefore the whole structure follows the configuration change if one angle is changed in it, and the diameter and height change happen synchronously, successfully reducing the required number of actuators. The proposed thermal solution decreases the thermal gradients and temperature extremes within the baffle considerably, and successfully shifts the overall temperature of the telescope towards colder regions.