Ensuring the sustainability of future space missions and maintaining continuous access to space necessitates a serious and proactive approach to addressing the space debris dilemma. One crucial measure that can be undertaken is actively protecting against space debris by implementing satellite collision avoidance manoeuvres. This thesis focuses on researching robust methods to perform such manoeuvres, with a particular emphasis on the guidance and control systems of manoeuvrable vehicles. The Starlink constellation is used in this research as a reference scenario, considering its popularity and extensive development, although the developed system is applicable to any satellite. The core of achieving successful manoeuvres lies in the creation of a robust guidance system. This system integrates convex optimisation algorithms with reference tracking mechanisms, enabling the autonomous execution of collision avoidance manoeuvres to be both robust and fuel-efficient. The outcomes of this research represent a significant advancement in space debris mitigation, enhancing space sustainability.

Automated control of Earth observation missions does not only allow for a more efficient use of onboard resources but also improved quality on the scientific return and cost reductions in the operations. Current developments in space hardware computational capabilities allow the implementation of onboard optimisation algorithms that can use the state estimations available onboard to dynamically generate control sequences in adaptation to disturbances or unpredictable behaviour.

This study proposes an onboard target management software with the aim of observing a set of targets on Earth’s surface in an automated manner. The development must be carried out within the framework of AAC Hyperion’s integrated attitude determination and control system. The proposed solution is designed to autonomously manage observations and optimise reorientation manoeuvers to designated ground targets, reducing the need for continuous ground intervention and enhancing operational efficiency. The method proposed combines the solving of a Mixed-Integer Linear Programming (MILP) optimisation problem for the scheduling of the observations, while
using successive convex programming to perform optimal control for spacecraft constrained reorientation and target pointing in a real-time manner. These two primary components, together with another task for computing the target visibility windows, operate within a modular framework, facilitating communication with the existing ADCS architecture and making efficient use of onboard computational resources.

The MILP scheduling subproblem selects targets based on priority, visibility constraints, and the required slew times between observations. Initial results identified limitations in MILP performance, particularly as the number of targets or proximity constraints increased, where large computational loads exceeded real-time processing capabilities. Alternative scheduling strategies are suggested for highly complex or densely populated target scenarios. Convex optimisation is found to be a suitable method for the solving of highly constrained manoeuvres in real-time. The reformulation of non-linear systems and complex, combined with the use of successive convexification were used to efficiently solve the spacecraft attitude control problem. This method’s robustness and
capability to detect infeasibility were especially beneficial for the modular onboard software.

The research tested and validated the proposed algorithms using a software-based real-time simulation environment, allowing a preliminary assessment of the observation manager’s performance. On the one hand, the method used for approximating the slew rate was found innaccurate, worsening the performance of the convex optimisation due to its interdependency with the MILP optimisation. On the other hand, results showed that the convex optimization method was effective for achieving precise target orientations, but achieving target accuracy of 0.1° often required refinement through additional onboard controllers, such as a PID controllers, to maintain
target lock during observation. Additionally, the trajectory optimiser demonstrated robustness at an update rate of 0.2 seconds, allowing the spacecraft to realign to targets successfully despite potential deviations in spacecraft state. However, the limited memory and slower processing speed of the ADCS hardware present challenges in extending these methods to onboard use.

Based on the results, the study suggests the modification of the slew approximation, and the use of alternative methods to the MILP, such as dynamic programming, to improve scheduling efficiency in future work. Moreover, testing on the ADCS hardware processor through the use of HIL simulations is essential to faithfully reproduce the onboard performance for analysis. A shift to higher-performance onboard processors or offloading some computational tasks to external processors is recommended to fully support the demands of automated Earth observation operations. This thesis demonstrates the feasibility and potential of integrating advanced optimisation methods onboard Earth observation satellites, providing a foundation for future advancements in autonomous spacecraft operations and contributing to the field of real-time, dynamic scheduling, and control in space systems.

With the price of launching satellites remaining high, many companies and universities are turning towards smaller satellites as more affordable options. This requires the miniaturisation of their different components, including the propulsion subsystem. At TU Delft, Vaporising Liquid Microthrusters (VLM) are being developed, with the latest iteration being a new generation Micro-Electromechanical System (MEMS) design. In this work, the updated design of this microthruster is remanufactured and prepared for future operational validation, heat transfer and instabilities studies. This is done by simplifying the manufacturing process, creating a testing interface, and performing measurements and extensive experimental characterisation of the thruster and test set-up. The newly-created interface allows for the study of heat transfer efficiency and convection heat transfer coefficient (HTC) from the silicon wall to the propellant, as well as multiple ways to study the instabilities in the system. One interface + thruster assembly was shown to survive to at least 180 °C temperature and 6-7 bar relative pressure. On the manufacturing side, the fabrication process was streamlined by reducing the number of steps by 13%, most significantly by the use of a soft photolithography mask instead of a hard silicon dioxide one. The deviations in horizontal dimensions are in the same order of magnitude as the ones obtained by the previous student working on this project, and lower than other previous attempts at the Faculty of Aerospace Engineering. The etching process encountered some difficulties, leading to inconsistencies in the feature depth between wafers. Across wafers, etching rate deviations as large as 8% were observed. Within the final production wafer, the range of deviation from the design values on throat depth was from +3.5% to −16.5%, showing a measurable higher etching rate at the edges of the wafer than in the centre. Recommendations were given on which features’ depths are most important to be measured, in order to increase the confidence in the predicted values. With these results, future researchers are ready to use the produced thrusters and testing interfaces. It is recommended to continue with the experimental studies, especially by measuring the heat transfer efficiency, thrust levels, HTC, and by studying any instabilities that could appear.

This study addresses the challenge of accurately modeling the multi-regime plasma flow through Ambipolar Plasma Thrusters (APTs), a type of Electric Propulsion (EP) system. Despite their advantages, understanding the behavior of plasma within APTs is complex due to the presence of two different flow regimes: the fluid regime inside the thruster’s source chamber and the kinetic regime inside the thruster’s magnetic nozzle. To enhance the precision of APT modeling, this thesis introduces a novel coupling methodology between fluidic and kinetic solvers, resulting in the MUlti-regime Plasma Equilibrium Transport Solver (MUPETS). MUPETS employs two models, the continuum OpenFOAM code for the fluidic regime and the Particlein- Cell (PIC) Starfish code for the kinetic regime, coupled through a closed-loop iterative coupling scheme. This approach allows for a self-consistent prediction of plasma transport within the entire thruster domain, including the interface between the numerical domains of each model. At this interface, which often coincides with the thruster outlet, the model’s iterative loop self-corrects the interfacing boundary conditions of each numerical domain, removing imposed boundary conditions such as assumed velocities. This has reduced plasma species discontinuities across the models’ interface to less than 4% for electrons and less than 1% for ions. The performance of MUPETS was validated against experimental measures from a laboratory thruster at the University of Padova, showing less than a 19% difference in predicted propulsive performance. The MUPETS code requires minimal alteration of the separate solvers, allowing for the fluid or kinetic solvers to be swapped out with higher-fidelity models or numerically better performing models as required. This flexibility enables the comparison of previously developed and validated models from literature to investigate their suitability for describing the plasma flow across the regime change. The study concludes that the developed coupling method presents an improvement to the state of multiregime plasma flow modeling while providing similar accuracy when used for predicting the propulsive performance of APTs.

Even though Deep Reinforcement Learning (DRL) techniques have proven their ability to solve highly complex control tasks, the opaqueness and inexplicability associated with these solutions many times stops them from being applied to real flight control applications. In this research, reward decomposition explanations are used to tackle this issue and augment DRL end-user explainability. A reward decomposition-based DRL controller is deployed in a longitudinal state-space model of the Cessna Citation 500 aircraft, and it is assessed on two attitude flight control tasks. Furthermore, a new explanation type called Dominant Reward eXplanations (DRX) is presented, which allows users to obtain more global insights than the ones generated by Reward Difference eXplanations (RDX). Results show that the explanations produced lead to straightforward and intuitive insights about the controller’s behaviour, capable of improving end-user explainability. Moreover, a small analysis seems to indicate that the decomposed method has similar performance to the one obtained without reward decomposition, however, training time increases considerably. To the author’s best knowledge, this is the first application of reward decomposition explanations to the flight control domain.

The plasma behaviour in an air-propelled Hall thruster (SITAEL HT5k) is investigated by means of numerical modelling and experimental measurements. The air plasma employed (nitrogen/oxygen mixture) is firstly chemically characterized, identifying the most relevant species produced: four neutrals (N₂, N, O₂, O), four positive ions (N₂⁺, N⁺, O₂⁺, O⁺) and electrons. Then, it is explored the kinetic theory of ion-ion streaming instabilities in the presheath, relevant to electropositive multispecies plasmas, extending it to the four-ions case under study. These results are used to adapt a non-stationary 1D fully fluid Hall thruster model to operation with an air plasma. Calibration with an experimental discharge current signal confirms the physical validity of the formulation in reproducing breathing mode oscillations, even in multispecies plasmas. The experimental data also consist of fast-diving triple Langmuir probe measurements, which are used to validate the model, showing reasonable accuracy of the spatio-temporal distributions of the plasma properties.

Observing that the majority of SSTO programs failed due to the propulsion system being far ahead of its time. Furthermore, acknowledging the that the private launch markets is increasingly dominated by partly reusable TSTOs. This thesis aims to evaluate whether next generation rocket engines make an SSTO viable within a foreseeable economical time frame, for the increasingly privatised space industry. From the market analysis it follows that the VT SSTO RLV developed by the private launch industry, with a payload capability of 15- 20 tons to a LEO with an altitude of 200-400 km, is most likely. The evaluation is performed by a two level trade-off. Initially, a literature trade-off is preformed on elaborate range of potential next generation rocket engine, which are subdivided into pure rocket engine and breathing engine. All engines are evaluated on the market requirements, performance, and achievability. Only three engine types are selected. Upon evaluating the propulsion literature it was noticed that no literature cross evaluated advanced propulsion systems with high performance propellant pallets. Therefore, a literature propellant trade-off is performed. For the aerospike, the pulse detonation engine and the precooled hybrid airbreathing rocket engine with the propellant pallets H2/O2, CH4/O2, C2H2/O2, and C2H4/O2 a performance analysis is performed. The performance analysis is done by both optimising the engine configurations and the ascent trajectory for a 15 ton payload delivery to an orbit of 400 km altitude. To simulate each engine, three performance analysis are developed, which are all derived from the continuity equations to ensure a consistent comparison. The aerospike is simulated as a continuous optimal convectional via the frozen equilibrium method, although it is found an alternation is needed to account for directional losses. The open-end of the PDE is simulated via a CJ-detonation wave followed by an exponential decaying blowdown phase. The DC nozzle expansion is iteratively simulated via the frozen equilibrium method. The intake and HPC of the precooled hyrbid airbreathing engine is simulated with convectional aero engine performance analysis, while the precooler is simulated ass a counter-flow heat exchanger. The CC and CD nozzle are simulated via the frozen equilibrium method. From the optimised trajectory the representative performance of each engine is extracted and five engine configurations are identified as viable VT SSTO RLV engines. These are the H2/O2 powered aerospike, the pulse detonation engine and the precooled hybrid airbreathing rocket engine operating on an H2/O2 pallet, and the C2H2/O2 powered aerospike and pulse detonation engine. The H2/O2 powered PDE is found to be the most promising, as it offers the best performance gain relative to its achievability. Therefore, it advocates that future VT SSTO RLV research utilises the H2/O2 powered pulse detonation engine as the main propulsion system

Detection of the openings in the Arctic sea ice pack, or leads, allow to sample instantaneous sea surface height (SSH) and this information is crucial for quantifying the impact of sea ice melting. It is therefore important to correctly detect as many leads as possible to obtain more SSH references. This paper studies 12 different classification methods including supervised-, unsupervised machine learning methods and thresholding method, being applied to the Sentinel-3 Synthetic Aperture Radar (SAR) altimetry data collected in March/April of 2017-2020 and June/July of 2020 from areas all across the Arctic Ocean. These are compared and assessed with respect to images taken by Ocean and Land Color Instrument (OLCI), also on board Sentinel-3, ensuring a perfect temporal alignment between the two measurements. The supervised Adaptive Boosting, Artificial Neural Network and Linear Discriminant classifiers showed excellent and robust results in March/April with overall accuracies up to 91.82%. The unsupervised K-medoid classifier produced excellent results achieving up to 91.51% accuracy and it is an attractive classifier as it does not require ground truth data. The classifiers perform poorly in the summer months, as sea ice returns show more ambiguous reflections due to melting. Therefore on summer data, classifications that are solely based on waveform data from SAR altimetry is unsuitable and auxiliary information is required. Furthermore, this paper attempted to identify off-nadir leads (ONL) by adding an extra class in supervised learning methods, intending to reduce the falsely detected leads. Most classifiers failed to detect leads and did not improve their false lead rate. However, as RUS Boost classifier was able to identify 61.6% of total ONLs, this can be used to initially reject these points for more conservative lead detection.

The Deployable Space Telescope (DST) is a telescope that uses foldable optical elements to reach state-of-the-art ground resolutions while minimising the launch volume and total mass of the system. One of the promising concepts that is part of the DST design is the Compliant Rolling-Element (CORE) Hinge. This hinge uses compliant links to let two semi-circular cams roll over each other with minimal friction. For this thesis, a thermo-mechanical analysis is conducted of the DST CORE hinges to determine whether the current design complies with the stability and in-orbit drift requirements of the DST. From this analysis, it can be concluded that this is currently not the case, but that a simple solution is at hand. By elongating the baffle of the DST, which protects the DST from solar irradiation, more solar radiation is blocked and the stability and in-orbit drift are kept within limits.

The Deployable Space Telescope (DST) project aims at producing a competitive telescope by using segmented deployable light weighted optics which reduces launch volume and mass, and thus launch costs. Due to the criticality of heat management to the performance of the telescope, a thermal model had to be built in order to determine the expected temperatures throughout the telescope in orbit. Further, the premature baffle design required a supportive thermal analysis such that its critical parameters could be determined.

River width is an important indication for water storage, with the development of remote sensing techniques, estimating river width in a large area becomes possible. In this report, fully-focused SAR altimetry data with an advantage of high spatial resolution is applied to compute river width for the first time. Methodology developed based on morphologic characters of rivers and data features is validated in two different areas, the Netherlands and Vietnam, with good statistical results.

In multi-stage launch vehicles, the adjacent stage re-enters the Earth atmosphere. Once the stage is re-entering the atmosphere, the low-melting-point materials melt and the inner parts will be exposed to excessive heat. With heat increasing, the parts will break into debris. Some of the fragments will burn in the atmosphere, others will crash into the Earth’s surface. Since there is a lack of knowledge about the re-entering physics, Demise Observation Capsule (DOC) project is started. The DOC collects information about the re-entry progress of launch vehicle stage, where it is mounted on. Several launchers and stages are considered to investigate the effect of the location on the DOC’s configuration. In an ideal situation, the DOC can be placed on all stages/parts of launch vehicles and at the same time it is optimized for that location. A generalized design concept that fits on several locations and stages, will reduce the cost and effort. Therefore, the thesis objective is to determine whether it is feasible to have a generalized DOC or an optimized DOC per location and load case. Additionally, the current DOC is designed for four locations and the second objective of this thesis is to optimize the current DOC and shift its center of gravity to the nose of the DOC. For this reason, the main research questions are as follows: Is it feasible to generalize the DOC concept design so that it will fit on every ESA launcher and stage and meet the predefined requirements and will withstand environmental loads while general operability of the DOC is guaranteed? How can the center of gravity be shifted on the DOC to its nose under the current designed loads such that the DOC can be made more stable?

Unlike a liquid rocket-engine, it is not possible to actively throttle the overall combustion of a solid rocket-motor. It is however possible to alter the thrust of an individual thruster by varying its throat area. The throat area can be altered by inserting a pintle of a conical or other shape into the nozzle throat. A setup which consist of multiple thrusters each with a pintle to alter its throat area, and which are connected to a single solid-propelled combustion chamber, can be used for attitude control of a rocket or spacecraft.

The goal of this thesis was to investigate the flow in a single thruster of such a system. For this purpose, a simple 1D simulation was developed. Additionally, multiple CFD simulations both static and dynamic were setup. The results of the simulations were then analysed in detail and compared amongst each other and with externally obtained test-data.