Measurements of the space radiation environment in Low Earth Orbit (LEO) are critical for satellite
safety and operations. However, the inherent shielding of a spacecraft alters the incident radiation
field, complicating efforts to reconstruct the true external environment from measurements taken by internal detectors. This thesis investigates the feasibility of developing shielding correction factors for proton radiation measured by a Timepix3 (TPX3) detector.
The research was conducted using the Geant4 Monte Carlo toolkit to model the transport of protons
through a 5 mm aluminium shield. This simulation framework was first validated against data from
ground-based proton accelerator experiments. Empirical models for correcting kinetic energy reduction and particle transmission were then successfully derived from the simulation data.
The validation process confirmed the simulation’s accuracy for high-energy protons (>70 MeV) but
revealed a systematic overestimation of energy loss at lower energies (<40 MeV). The investigation
into applying the correction factors uncovered a more basic limitation: an inherent ambiguity exists
in the relationship between a proton’s deposited energy (𝐸𝑑𝑒𝑝 ) and its kinetic energy (𝐸𝑘𝑖𝑛 ), which prevents a reliable, direct conversion from the detector’s measurements.
It is therefore concluded that while theoretical correction models can be formulated, their practical
application to shielded detector data is impractical due to the main challenge of reconstructing the
incident energy of detected particles.

This thesis investigates the feasibility of the usage of space-based Laser Ablation Propulsion (LAP) to perform orbit maintenance on CubeSat in Low Earth Orbit (LEO). The thesis is motivated by the growing need for efficient, compact propulsion systems for small satellites and explores LAP as a method that eliminates the need for onboard fuel by generating momentum via directed laser beams from orbital laser stations. A detailed simulation framework is developed to model the physical environment in LEO, including atmospheric drag, solar radiation pressure, and third-body gravitational effects. The propagation of high-powered lasers through space is modeled considering diffraction, absorption, and current limits of aiming precision. The propulsion interaction between laser and CubeSat is simulated through a dynamic model based on material properties and laser beam parameters. Power generation and energy storage are analyzed in the context of space-based constraints, comparing solar arrays, battery technologies, and their degradation over time. A financial model evaluates system-level trade-offs, including station cost, maintenance intervals, and replacement economics. Optimizations are performed to understand how propulsion efficiency, energy storage capacity, and environmental variability affect system performance and scalability. Simulation results show that LAP is viable under current technological constraints, provided stations are deployed in sufficient numbers and at the right altitudes. The system can service a meaningful number of CubeSats while remaining competitive with the cost of satellite replacement. Technological limitations such as aiming accuracy and beam control are identified as key challenges, with recommendations for future work including the expansion of in-orbit LAP to networks, and further applications.

Launcher manufacturing is undergoing a rapid transformation, driven by the need for increased production robustness, cost reduction, and enhanced sustainability. In particular, the adoption of automation and digitalisation in production processes has become essential to remain competitive in today’s space industry. This thesis contributes to this transition by developing a data-driven anomaly detection system for robotic drilling and riveting, a critical step in the assembly of launcher structures. Drawing inspiration from the field of audio classification, in which transformer-based models have recently achieved state-of-the-art results, this research proposes a novel adaptation of the spectrogram transformer architecture to the domain of industrial time-series sensor data. The resulting deep learning-based approach enables automated detection of anomalous behaviour through sensor signals, offering a scalable solution to process monitoring.

This thesis investigates methods to enhance the robustness of space object cataloguing pipelines, focusing on tracklet correlation and orbit estimation using angular measurements from short observation arcs. The cataloguing robustness is defined as achieving high true positive and negative rates for tracklet correlation to allow for the build-up of an accurate object catalogue. The study addresses the main research question: How can the robustness of the cataloguing pipeline be improved when applying orbit estimation methods to the full angle set of short observation arcs?

A baseline tracklet correlation approach, based on the Boundary Value Problem (BVP) within the Admissible Region framework, is implemented. This method uses angular observations and hypothesized ranges to estimate an object's state, with correlations evaluated via a cost function based on the Mahalanobis distance. Classical IOD methods are employed to investigate their application toward validation of tracklet correlation when reconsidering the full angle set. The considered methods include the angles-only Gauss method, a multiple angles least-squares Gauss approach, Gooding’s method, as well as a Batch Least Squares (BLS) orbit determination (OD) method. The BVP and IOD methods consider two-body dynamics, and the BLS Earth's zonal harmonics and third body effects from the Sun and Moon. Simulated measurements are derived from Two Line Element sets (TLE) for initial reference states for LEO, MEO, and GEO objects, propagated with the SGP4 model accounting for Earth's atmospheric drag, zonal harmonics and third body Sun and Moon effects, providing the test data.

Results show that the BVP method performs best for GEO, achieving ~90% true positive rates with reasonable uncertainty gating. For LEO and MEO, higher thresholds and cost-function minima are required due to greater observation complexity and force-model discrepancy. Gooding’s method, making use of a Lambert solver, demonstrated robust performance across multiple orbital revolutions, while Gauss’ methods were less effective for large time gaps. Additionally, BLS struggled with sparse data and large time steps, offering limited state refinement despite higher computational expense.

The findings suggest gating based on chi-squared distribution thresholds for GEO and higher magnitudes for LEO and MEO to optimize true negative rates. While the BVP method provides sufficient accuracy for re-observation scenarios, classical IOD methods and BLS exhibit limitations under sparse tracklet conditions. This work highlights challenges in cataloguing lower-altitude objects, for ground-based optical observations, and suggests the application of the BVP method on lower altitudes requires inclusion of force models for the primary perturbations.

Due to the increasing interest of the aerospace industry and the scientific community in missions targeting halo orbits, a specific family of periodic solutions within the circular restricted three-body problem, it is highly desirable to find ways to reduce the cost of transfers between these orbits. To achieve this, depending on the mission characteristics, one could opt for reducing the required propellant mass or the time of flight. As such, with two objectives to be minimized, an ideal implementation for a mission designer would provide a Pareto front of feasible transfers instead of the single trajectory commonly obtained with conventional methods. Moreover, the necessary propellant mass can be minimized even further by means of electric low-thrust propulsion due to the significantly larger exhaust velocities. Therefore, the purpose of this research is to develop, implement, and study a suitable approach to obtain a collection of low-thrust transfers between halo orbits optimized in terms of both propellant mass and time of flight.

The approach employs an optimal control indirect method as thrust law which, combined with a heuristic optimizer based on differential evolution, can find a wide variety of trajectories that minimize the aforementioned objectives within the circular restricted three-body problem. Heuristic optimization is employed to remove the dependency of the solution on the provided initial guess and find trajectories in the region of the global minima. To satisfy the demanding boundary conditions characteristic of indirect methods, these constraints are included as a third objective for the optimizer to minimize as well. Due to the tolerance allowed on the constraints, the trajectories are subsequently refined with direct collocation methods. Then, they can be transitioned to a high-fidelity model, taking advantage of the versatility of direct methods. Furthermore, the implementation allows for the inclusion of invariant manifold phases arising from the departure and target orbits to obtain a wider set of Pareto-optimal solutions.

To assess the suitability of the proposed procedure, a specific transfer between two halo orbits around different Lagrange points of the Earth-Moon system was optimized. The results consisted of 100 Pareto-optimal transfers spanning more than 60 days, offering significant mission design freedom. Moreover, the Pareto front outperforms the trajectory found in the literature for a comparable use case by 30\% in all objectives. Next, an optimized trajectory was first successfully verified with the mission analysis software ASTOS and then refined with direct collocation. The refined trajectory exhibited negligible changes in performance, rendering the trajectories obtained with this approach as promising initial guesses for further optimization with direct collocation methods. Moreover, several major perturbations not included in the simplified dynamic system were also corrected with direct collocation. The next steps include: accounting for the eccentricity of the Moon's orbit to fully transition the trajectories to a high-fidelity model, assessing the optimization quality with different use cases, and implementing transfers between different periodic solutions and dynamic systems, such as the Sun-Earth system.

There is a growing demand for nano-satellite missions that require advancements in micro-propulsion capabilities to facilitate a wider range of orbital maneuvers. Among these capabilities, the ability to accurately control thrust would unlock new possibilities for nano-satellite applications, including missions such as space debris removal and orbit transfer.
Delft University of Technology is currently pioneering the development of an innovative green propellant-driven micro-propulsion system based on micro-electro-mechanical (MEMS) technologies for its PocketQube, known as Delfi-PQ, which features a compact form factor of 5x5x5 cm. While the thruster itself is in development, the interfacing and integration with other components are still ongoing.
Given the stringent mass, volume, and power limitations imposed by PocketQube satellite requirements, there is a pressing need for micro-scale components to realize a highly integrated propulsion system.
This thesis focuses on the design of a MEMS-based microvalve for proportional flow control in micro-resistojets. The design is conceptualized as comprising three components working in harmony: a valve seat with inlet and outlet, a flexible membrane, and a piezoelectric actuator. The valve seat with inlet and outlet, as well as the flexible membrane, utilize MEMS manufacturing techniques and are based on a silicon chip. And, a new design for piezoelectric actuators employing the d31 mode for contraction strokes is proposed.
The proposed preliminary design is a normally closed microvalve designed for a flow rate of 5g/hr, with the flexibility to accommodate higher flow rates if needed. It offers proportional flow control, ensuring precise regulation of fluid flow. Additionally, it promises a low power consumption of less than 1W and a low response time. Furthermore, this thesis provides a detailed outline of the MEMS fabrication process flow available at TU Delft’s Else Kooi Laboratory for the device. It also features a comprehensive test plan aimed
at assessing the feasibility of this design in future studies. Additionally, the thesis includes an elaborate risk analysis to help identify and mitigate potential risks during the manufacturing and testing phases.
The design shows promise and could pave the way for future developments of the microvalve within the department.

The Jovian system is one of the most attractive destinations for scientific space exploration, with many missions having flown to Jupiter and its moons. Among the latter, Io stands out due to its intense volcanic activity, with plumes extending up to hundreds of kilometers above its surface. Scientists believe that sampling, returning, and analyzing the particles ejected from Io’s volcanoes has the potential to unlock valuable information about Earth’s formation, and the history of the Solar System. For this reason, the concept of an Io Sample Return mission was born at the Jet Propulsion Laboratory, California Institute of Technology, where this thesis was carried out.
The project focuses on the mission design and trajectory optimization of a Discovery-class Io Sample Return concept. It investigates which geometry, maneuvers sequence, and flyby trajectories, can enable the sampling of Io’s Prometheus plume through a single flyby, before returning the material back to Earth.
Firstly, a broad-search of feasible patched-conics round-trip trajectories to Jupiter is conducted, using a simplified two-bodies model. The search incorporates launch, entry, time of flight, mission delta-V, and Io encounter constraints. Two algorithms for the reconstruction of ballistic and targeted flybys are developed and integrated within the trajectories-search workflow. This phase highlights the infeasibility of the 14 years flight time constraint and the 8 km/s maximum Io-relative speed during sampling. The two thresholds are increased to 18 years and 10 km/s, respectively. The solutions-space is progressively filtered with the aid of a primer-vector optimizer, and a single candidate trajectory is chosen for further studies. The solutions from the broad-search are also used to conduct a sensitivity study on alternative plume targets, which highlights Prometheus’ ideal position for sampling missions that avoid Jupiter orbit insertion.
The selected patched-conics candidate is used as initial guess for the numerical propagation and optimization of the end-to-end mission. A high-level trade-off is conducted to select the most suitable approach to propagate the trajectory. Two optimization approaches are then compared. An arc-wise scheme, in which each interplanetary transfer is optimized individually, and an all-arcs method, where the Earth-Jupiter and Jupiter-Earth journeys are each optimized in their entirety. Self-Adaptive Differential Evolution (SADE) and Generational Multi-Objective Evolutionary Algorithm by Decomposition (GMOEA/D) are the two optimizers of choice.
Optimization runs conducted using the patched-conics initial guess prove unable to converge to acceptable solutions. Moreover, single-objective optimization struggles to satisfy position discontinuities requirements, when adopting the all-arcs approach. The patched-conics solution is therefore extended to multi-conic, leading to significant performance improvements. Final results show that GMOEA/D outperforms SADE using the all-arcs approach, and finds a solution that satisfies delta-V, launch C3, entry speed, and position discontinuities constraints. It also improves the total delta-V of the baseline multi-conic solution by about 70 m/s, leaving over 700 m/s of margin on the mission delta-V budget.

This thesis work has been motivated by the growing interest in small-spacecraft small-body missions and more specifically inspired by a proposed mission to the largest Martian Trojan (5261) Eureka. This binary asteroid represents a promising target for a planetary mission, as a better understanding of its formation and evolution would provide valuable insights into the history of both the Martian system and binary asteroids. This work has focused on identifying orbital designs which would optimise the geodetic parameters estimate, and thus best help characterise the asteroid's internal structure. The orbit determination solutions have been computed from simulated tracking data in various orbital configurations. To compensate the highly perturbed nature of the binary asteroid's environment, pseudo-periodic solutions identified around Eureka have been used as stable initial conditions for the spacecraft orbits. As Eureka's shape, gravity field and rotational state are not well characterised yet, large resulting uncertainties have been accounted for. They have been propagated to assess the robustness of the spacecraft orbits and subsequent orbit determination solutions to dynamical mismodelling, in order to eventually limit the science risk of the mission. For single CubeSat configurations, an optimal semi-major axis has been identified, being the one closest to the asteroid before orbital instability would start deteriorating the orbit determination solution. Non-equatorial retrograde orbits have been found to provide the lowest formal errors for geodetic parameters. Moreover, the benefit of simultaneously deploying two or even three CubeSats around the asteroid has been demonstrated. In both the two and three CubeSats configurations, promising orbital designs have been identified. They provide an estimation of the geodetic parameters which is accurate enough to bring insights into Eureka’s internal structure, and has been proven to be robust to dynamical mismodelling.
The methodology developed in this work is of interest to design other small-spacecraft small-body missions, and optimise their achievable geodetic parameters estimate. Additionally, Eureka's dynamical model and the large uncertainties assigned to it have been mostly based on models available for other binary asteroids. This brings confidence in the possible applicability of the identified optimal orbital designs to CubeSats missions targeting other binary asteroids.

This Thesis presents the design and implementation of a preliminary Fault Detection Isolation and Recovery (FDIR) architecture for LUMIO, a CubeSat mission to the Moon. The baseline of the FDIR is the Failure Modes Effects (and Criticality) Analysis (FMEA/FMECA) of the mission, through which the potential failure scenarios are classified. The FDI design is based on the use of simple checks, mainly cross-checks. The FR strategy is based on the application of a sequence of recovery levels. The algorithm was implemented in Simulink and a simulation model of the satellite was developed to perform verification. The study paved the way for the advancement of the project, since the critical items were identified, and compensating provisions were proposed. The tests performed proved the ability of the FDIR to detect and recover the preliminary FMECA failures; hence, to increase the autonomy of LUMIO and the overall reliability of the mission.

Ballistic capture is a transfer method which was first applied in 1990. It allows a spacecraft to approach a target celestial body and enter a (temporary) orbit around it without requiring manoeuvres in between. Ballistic capture is a promising concept, as it is expected to be safer, cheaper, and more flexible in terms of launch windows than a traditional Hohmann transfer. Currently, a computationally efficient method which simulateneously allows for an inisghtful description of the dynamics of the ballistic capture problem remains to be found. A potential solution lies within the field of Lagrangian Coherent Structures (LCS). LCS is defined as a separatrix of regions in a flow with distinct dynamics. It may be possible that LCS around a planet have some correspondence to results found using stable set manipulation, a classic technique for obtaining capture trajectories. In this research three new areas within the field relating LCS to ballistic capture are explored. Firstly, it has not yet been shown what LCS can be found in an area around a planet, without making use of a priori stable set information. Furthermore, it is unclear what the effect is of changing the integration time in the procedure of extracting LCS. Finally, there has not yet been an analysis to show how the LCS relate to stable sets with different number of revolutions n. In this work two algorithms for extracting LCS have been developed. One is based on the simple but efficient computation of the Finite Time Lyapunov Exponent (FTLE). Another is based on the more involved Variational Theory. Both algorithms are validated on a toy problem used frequently in LCS extraction studies, and are then applied to the Elliptic Restricted Three Body Problem (ERTBP). It is shown that LCS around a planet yield resemblance with stable set results. The FTLE-based algorithm is able to quickly and efficiently identify the shape of the stable set. The Weak Stability Boundary, however, can not be extracted distinctly. The Variational Theory-based algorithm yields more distinguishable results for the Weak Stability Boundary. It is shown that large and constant integration times are beneficial. It is shown that extracted LCS form an approximation of the average resulting WSB for all stable sets.

Micropropulsion is universally considered to be a key technology enabling nano- and pico- satellites to perform more complex missions. However, past research has shown that nozzle efficiencies at the microscale are far inferior to their macro scale counterparts. These low efficiencies can be attributed to the relatively high viscous losses associated with this microscale. This thesis conducted a three-dimensional numerical study to investigate the impact of the nozzle geometry on the viscous losses. Furthermore, the designed micronozzles are fabricated and the ground work is laid for experimental testing of these nozzles. Results of the numerical study show that by application of a new double depth micro aerospike design nozzle efficiencies can be improved by as much as 41.2%.

Geolocation of Radio Frequency (RF) emitters, is finding the geographical location of said emitter. This thesis explores the concepts of geolocation, specifically, Space-Based geolocation of earth-surface emitters by means of measuring the Angle-of-Arrival using a Uniform-Rectangular Antenna Array. Initially the literature is explored, the significance of Geolocation is made clear and the research outline and scope is defined. Afterwards, Angles-of-Arrival finding algorithms are investigated and implemented, specifically the subspace algorithms: "MUltiple SIgnal Classification" (MUSIC) and "Estimation of Signal Parameters via Rotational Invariance"(ESPRIT). These algorithms are thoroughly explained, a model is obtained and implemented into a Simulation framework. The Simulation Framework is used to create a Graphical Tool used to calculate performance figures of Space-Based Geolocation. By using the obtained Simulation, a sensitivity analysis is produced. This sensitivity analysis observes the effect of design parameters of a satellite platform and geolocation payload on the geolocation performance. This sensitivity analysis is used to find the optimal approach when designing a geolocation payload. This approach can be used to synthesize a geolocation payload design that maximizes performance.

The increasing interest in solar electric propulsion techniques, that enable significant propellant mass savings for a wide class of transfers, has brought about a revolution in the approach to trajectory design and optimization, as a result of the complexity and diversity of the problem. To date, solutions of different nature exist, but numerical methods that require significant computational effort and user experience are typically used already in the early stages of mission design, due to the limited availability of reliable medium-to-low fidelity design tools for SEP transfers. This research project proposes a novel method that computes transfer performance parameters for Earth-Mars mass-optimal SEP transfers, by means of empirically derived analytic relations. The method is intended for applications such as concurrent engineering and early-phase concept development, which require the fast characterization of a broad design space. Besides accommodating a wide range of currently available systems, the method successfully deals with modelling the effect of non-zero infinity velocity at departure and/or arrival.

The methodology that has been applied consists of a first phase of generation and characterization of the transfers, and of the subsequent selection of the model variables, model functions and architecture. Regarding the generation of the transfers, it is assumed that the transfers are coplanar and that the initial and target orbits are circular. Hundreds of transfers are optimized in a semi-automatic way and characterized in terms of thrust profile and transfer performance parameters. In the investigated design space, different regimes are identified, but approximately 90% of the acceleration range of interest falls into the thrust-coast-thrust profile for any combination of departure and arrival infinity velocity. For a proper description of the underlying trends in the transfer parameters, three key variables have been identified, namely the average acceleration, the total infinity velocity and the infinity velocity at arrival (expressed as a function of the total infinity velocity). By means of curve-fitting, analytic relations are derived that successfully describe those trends, limited to the thrust-coast-thrust class of transfers.

The method that is presented computes near-optimal transfers in terms of ΔV cost, transfer time, transfer angle and departure date. While the first three parameters are the outputs of the mentioned curve-fit model above, the departure date is computed by solving analytically the problem of the phasing with Mars, in a subsequent step. The fit functions that are derived model circle-to-circle planar transfers with an accuracy in the order of 0.1% with respect to the ΔV , 1.5% to the transfer time and 1.2% to the transfer angle, successfully dealing with the dependence on the departure and arrival infinity velocities and generating instant estimates for all relevant transfer parameters. When the model performance is considered in relation to transfers derived in the full ephemeris model, the errors are within 1% for the ΔV , within 15% for the transfer time and within 12% for the transfer angle, which, together with the demonstrated efficiency and simplicity of implementation, make it suitable both for early-stage assessments and for generation of suitable first guesses.