The management of toxic leaks in confined environments, such as CERN's experimental caverns and space habitats, is crucial due to potential risks to human life. This thesis examines aggressive carbon dioxide releases from CERN's 2-Phase Accumulator Controlled Loop (2PACL) system, used in both terrestrial (ATLAS, CMS, LHCb) and extraterrestrial (AMS) experiments. Computational fluid dynamics (CFD) simulations were performed to assess the severity of various failure modes. A two-dimensional analysis of the immediate leak area showed that shock waves travel up to 30 centimeters from the failure point. Subsequent three-dimensional simulations in the CMS experimental cavern (UXC55) revealed that while the upper cavern floors are well-protected, the lower floors pose significant risks due to semi-confined spaces and worse ventilation. Recommendations include the use of oxygen deficiency monitors, self-rescue masks, and detailed evacuation plans to enhance worker safety.

Beam-powered thermal propulsion, utilizing an external energy beam to heat a propellant, is proposed as a promising alternative to traditional chemical and electrical propulsion systems. This research aims to contribute to this technology’s numerical and experimental development, offering a comprehensive analysis of the receiver-absorber cavity
(RAC) performance and its optimization.
The work begins with a detailed background study of the principles of beam-powered propulsion. This sets the stage for the subsequent chapters, which investigate the numerical modelling and experimental testing of the thruster.
The numerical part of this thesis focuses on developing accurate prediction tools for the total absorbed beam power by the RAC. This information is added to a previously
developed RAC performance prediction tool and the development of a new tool considering spatial information is started. The optimization process regarding efficiency is
presented with an example case study with optimal values above 80% with a margin to improve.
The experimental component involves the design, manufacturing, and testing of a prototype beam-powered thermal thruster. Leak testing, test bench calibration and a
cold flow test are conducted. Detailed test procedures and setups for these are presented, alongside their results. The experimental findings are compared with the numerical predictions to assess the accuracy and reliability of the models.
The thesis concludes with evaluating the research questions, and discussing the successes and challenges encountered. Recommendations for future work are provided,
aiming to guide further advancements in beam-powered thermal propulsion technology. The combination of numerical and experimental insights gained through this research contributes to the feasibility and optimization of beam-powered thermal propulsion systems.

On-orbit servicing and active debris removal missions are becoming increasingly important to limit the growth of space debris in Earth’s orbit. SmallSats offer a promising option for reducing the cost and risk of these missions, allowing for their expanded use on a wider scale. For these missions to be successful, SmallSats are required to operate at very-close ranges to a target object. Relative visual navigation at this distance still presents a hurdle that needs to be overcome prior to implementing these platforms. This work analyses the suitability of Simultaneous Localization and Mapping (SLAM) for navigation at very-close ranges and proposes a modified algorithm to increase performance. Characterization of this algorithm is performed against a custom generated lab-based image dataset simulating final approach to a satellite. Algorithm performance shows promising results for SLAM use in very-close range environments and presents a first step in the development of this SmallSat navigation technology.

A laser-sustained plasma (LSP) generator capable of operating as a laser-thermal thruster was designed and tested. The apparatus was powered by a 3-kW 1070-nm fiber-optic laser and was operated with argon gas. The laser absorption, radiated spectrum, and change in static pressure were recorded, and high-speed video footage of the LSP was acquired. Of special interest is the heat-deposition efficiency into the working gas resulting from LSP ignition, which was estimated by tracking the pressure change in the test section. Laser transmission through the plasma was also measured to quantify two energy loss mechanisms: incomplete laser absorption and radiation of heat to the walls of the test section. Minimizing these losses would be critical in the realization of a practical and efficient laser-thermal propulsion (LTP) thruster, as most of the laser energy should be deposited as heat in the propellant to maximize thruster specific impulse.

With the development of space research into novel areas, new complex problems arise. The interest in solving space routing problems considering large numbers of targets has recently grown. This paper proposes a novel method to solve the optimal trajectory in such combinatorial space routing problems. This paper focuses on a global optimisation algorithm implemented to solve the problem posed in the 4th Global Trajectory Optimisation Com- petition (GTOC4). The solution is a trajectory of multiple legs, where each leg links two targets and has a specific flight time. To enable the use of the powerful mixed-integer linear problem solver software, Solving Constraint Integer Programs (SCIP), the routing problem concerned with visiting as many target bodies with a predetermined fuel and time budget is split into linear sub-problems. The Fixed Budget sub-problem selects a subset of the given set of targets. The Full Tour sub-problem orders the targets in the subset, and the Fixed Tour sub-problem optimises the flight time for every leg of the given trajectory to find the solution with the lowest total fuel consumption. Each of these sub-problems is formulated in a linear form and is solved using SCIP. The global optimisation algorithm evolves a population where every individual exists of a set of initial guesses for the time of flight values. Analysis shows that initialising this population with a mix of randomly generated individuals and individuals containing a constant value for all entries leads to the fastest convergence towards the optimal solution. In a population of 20, seeding ten individuals is found to be optimal. It is also found that the algorithm performance can be further increased by evolving individuals with infeasible solutions instead of iterating them until a feasible solution is found and eliminating the Full Tour sub-problem. These simplifications allow for an increase in the cost budget multiplier, which leads to finding better objective values without further increasing computational time. The best-performing setup, which uses a cost budget multiplier of 10, can find the optimal solution to the test problem in 100% of the runs, on average in 9 iterations, with a computation time of 5.82 seconds per evaluation. The results show that the global optimisation algorithm produces results that closely match known results for GTOC4 consistently and accurately.

The electric-pump cycle is a rocket engine configuration that uses an electric motor to power the pumps instead of a turbine. This offers several expected advantages such as simpler design, lower development costs, and easier restartability, but comes with reduced performance compared to conventional cycles. Previous research has primarily compared the electric-pump cycle to the gas generator cycle and has been limited to direct comparison. To extend the previous research a Rocket Cycle Analysis Tool was developed called RoCAT. It models the last-named cycles as well as the open expander cycle for a broad scope of thrusts, burn times, and chamber pressures, and several propellants. In addition, RoCAT optimizes several inputs for each cycles individually for a fairer comparison. Besides analyzing the electric-pump cycle's current performance, this research also estimates its performance in the future based on historic trends in its key technologies like the battery and electric motor.

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.

Using renewable energy to power a Mars habitat is a technological challenge because resources such as solar and wind are significantly weaker than on Earth. This work investigates the feasibility of using airborne wind energy (AWE) systems in combination with solar photovoltaic (PV) modules to power a Mars habitat. The Luchsinger model and the higher fidelity QSM are used to simulate the performance of the AWE system and compared. This thesis builds upon two earlier design synthesis exercise (DSE) projects by implementing a version of the quasi-steady model (QSM) that accounts for the transition phase, models a realistic retraction trajectory, and accounts for the mass of the airborne components. Additionally, the results of the first DSE indicate that the Luchsinger model used did not appear to consume any energy during the reel-in phase, which is not realistic and led to an over-prediction of mean cycle power. However, this thesis aims to implement the Luchsinger model correctly.

Creating a road map for sizing AWE kite systems on Mars is the main objective of this thesis. Since the performance of the AWE system is varying in time and space, the Mars Climate Database (MCD) is used to retrieve atmospheric and surface solar flux data including wind Weibull probability distribution functions (PDF). The second DSE used the MCD and validated the results against wind data from various Mars landers. The MCD is based on numerical simulations of the Martian atmosphere using a general circulation model and validated with available observational data. Seasonal vertical wind profiles are generated from the meteorological data to characterise the boundary layer over time. A scaling study assesses how AWE on Mars differs from that on Earth, performing dimensional analysis. In the system characteristics chapter, the initial sizing of the kite area and mass is computed using the scaling study. The performance models create the power curves, which together with the wind PDFs and surface solar flux data are used in the habitat energy model to verify whether the power requirements are met. Due to an insufficient amount of quantitative information on the energy consumption of the robotic construction of the habitat, the design of the microgrid is covering only the use of the habitat, which is 10 kW of continuous power. This is similar to remote off-grid solutions on Earth, with the additional challenge of having lower resource availability, both for wind and solar. This thesis concludes that various configurations of a hybrid power plant can continuously provide 10 kW of power throughout the entire Martian year. Moreover, the results indicate that using kites alone could generate sufficient power for the habitat without using solar PV.

Rhizome project: http://www.roboticbuilding.eu/project/rhizome-development-of-an-autarkic-design-to-robotic-production-and-operation-system-for-building-off-earth-habitats/

Published poster on thesis for Airborne Wind Energy Conference in Milano, Italy,: https://repository.tudelft.nl/islandora/object/uuid%3A4013ca0b-5508-4413-b8cb-5251ca1e7781

Video made to illustrate how Mars flight would look like.: https://www.youtube.com/watch?v=YS20vHhqgKk&ab_channel=AirborneWindEnergyOnline

Synthetic Aperture Radars (SAR) have demonstrated to be great instruments for space-based Earth observation in the microwave frequency. Metasurface antennas, on the other side, are a new type of planar antennas with a great potential for space applications, as they are compact and easy to produce. Several methods have been developed in the past years addressing one or many parts of the design and applications of these antennas, which rely in more or less measure in locally developed tools. However, some parts of these methods can take a great amount of time to implement, making this process more difficult than it should especially if the user has commercial electromagnetic simulation tools available that could accelerate this process. In this work, different procedures have been adapted to the design of a high-gain SAR anisotropic metasurface antenna for a multi-static Earth observation mission, showing that the proposed methodology produces accurate results.

Quasi-satellite orbits (QSOs) are orbits around the smaller primary in three-body systems. The MMX mission (developed by JAXA, to be launched in 2024) will use QSOs to orbit Phobos, before landing on its surface. This work studies the feasibility of using the invariant manifolds of three-dimensional QSOs for designing landing trajectories, using the Mars-Phobos system as a case study.

Applying continuation and bifurcation-analysis techniques, different families of QSOs were computed and their invariant manifolds propagated. Using a set of favorable manifolds as a reference, multiple maneuvers were designed via multiple-shooting and optimization techniques, producing feasible landing trajectories. The robustness of these trajectories was analyzed using a stability index and Monte Carlo simulations.

It was found that using the invariant manifolds of QSOs to land is only possible for some families of QSOs. However, for these, the manifolds allow generating robust and propellant-efficient landing trajectories that are able to reach most of Phobos' surface.

Since the introduction of Hydrazine to the list of substances of very high concern, REACH (Registration, Evaluation, Authorisation and Restriction of Chemicals), there has been an increase in the investigation of “green” propellant alternatives. At the DLR Institute of Space Propulsion, a mixture of Nitrous Oxide
and Hydrocarbons (Ethylene or Ethane), called HyNOx, has been under development since 2014. These monopropellants offer bipropellant-like performance, the ability to be used in a self-pressurised propulsion system, and are non-toxic. HyNOx propellants have the drawback of high combustion temperatures (up to 3400 K). Therefore, it is necessary to study what cooling methods can be applied to rocket engines using HyNOx propellants.
During this project, the test setup, HTMS (Heat Transfer Measurement Section), has been designed, built and tested in the M11.2 test bench at the DLR Lampoldshausen to study heat transfer in regeneratively cooled channels. The design was based on what was learned from studying past experimental
setups. A test campaign was performed to study the use of Nitrous Oxide and Ethane as coolant in stainless steel, copper and aluminium alloy pipes of different diameters. The setup diagnostics system was improved by procuring new measurement instrumentation. Additionally the instruments were
installed in different ways in order to understand which method would provide the most accurate and reliable results. The results of the thesis show that Ethane has better coolant properties than Nitrous Oxide due to its higher specific heat. However, the carbon present in Ethane leads to the formation of carbon deposits at high temperatures which can lead to clogging of the pipes and therefore Nitrous Oxide could be a the preferred coolant for situations where the temperatures are high enough for carbon deposition to occur. Additionally, it was determined that pure copper and the aluminium alloy AlMgSi 0.5 are not adequate materials for the construction of cooling channels. In the course of this project and to complement the experimental testing, a Matlab-based numerical simulation tool was developed that simulates the temperature profile in the HTMS setup using empirical Nusselt correlations to determine the heat transfer coefficient.

The Brik-II satellite is a Nano-Satellite which has the main objective to be a technology demonstrator and prove the use of self owned military satellites. One of the payloads onboard the Brik-II satellite is called the Store and Forward payload. It has the capability to Store and Forward messages from and to different military assets. To demonstrate this capability it has been purposed to design and install a communication system onboard the Royal Netherlands Navy submarines enabling communication with Brik-II. Subsequently, this would be a good way to demonstrate the performance and operational relevance of such system.
The first step was to identify the problem that should be solved. The main objective of this MSc thesis was to design, manufacture and test an operationally deployable communication system for the submarines operated by the Royal Netherlands Navy. Based on the main objective and consultation with the Navy the requirements of the system were compiled. From the requirements it was concluded that there are a number of similarities between the Remote Radio Station (RRS) and the submarine communication system. However, the antenna system installed on the RRS cannot meet the size requirements. Therefore a new antenna systems had to be design that would comply with the requirements. This has been the main focus of this MSc thesis.
To find the optimal design for the antenna, a simulation tool for the communication between the submarine and Brik-II was developed. In the model different antennas were tested to identify the optimal design. The type of antenna used in the optimisation is a helical antenna. Once the optimal design was identified it was manufactured used the 3D printing facilities at the 982 Squadron in Dongen. Hereafter, measurements were conducted to verify that the radiation pattern of the antenna corresponds with the predicted radiation pattern. The similarities in shape between the predicted and measured pattern were evident. However, the measured gain of the antenna was lower than the predicted gain. To meet the requirements the losses will have to be mitigated. Multiple solutions were purposed to reduce the losses. A communication test between the antenna and the satellite could not be conducted due to technical difficulties during testing.
Part of the main objective was to develop an operationally deployable system. This first prototype developed in this MSc thesis provides a basis for further iteration. The current design imposes two constraints on the deployability. First the antenna gain losses have to be mitigated since the lower gain results in less than required data throughput. The losses can be reduced by implementing the proposed solutions. Second, the choice of a helical antenna with an omnidirectional radiation pattern results in an increased chance of detection by possible adversaries. This functional constraint cannot be remedied due to the size requirements. In conclusion, the antenna design proposed in this thesis gives a basis for a system that can provide added value to the operations of the submarine.

Electric propulsion is generally characterized by higher specific impulse than traditional chemical rockets, which gives them propellant saving benefits. These have been applied in deep-space missions and micro-satellites. A Pulsed Plasma Thruster consists of two electrodes around a block of teflon. A pulsed discharge ablates the teflon and ionizes the vapor. This plasma forms a current bridge between the electrodes which is accelerated outwards by the self-induced magnetic field, producing thrust.

The current research project has developed and used a 1D model to investigate the current bridge with a focus on the plasma electrode interaction. The influence of electron temperature and density, cathode temperature, voltage, geometry and propellant type on the current density and magnetic field has been explored. The influence of thermionic electron emissions on the current-voltage characteristic has been demonstrated.

Resonance ignition devices utilize the gas dynamic phenomenon to heat propellant gases to their auto-ignition temperature in order to initiate combustion. These devices are a promising alternative to current rocket engine ignition systems. In this thesis a python code was developed to generate a resonance ignition system design. The design was tested and parameters characterized using ANSYS Fluent. This resulted in a detailed design of a GOX-Ethanol resonance ignition device that is successful in heating the resonance gas to above its auto-ignition temperature. This project presents the python and Fluent tools developed to conduct this study, analyses the effect of different input parameters on the device, and presents a working detailed design.

The Kuiper Belt is considered to be formed by remnants of the original Solar System, that is why exploration missions to that region are susceptible of having an enormous scientific impact. Nonetheless, it is far away from Earth, missions to KBO targets a significant challenge. Additionally, current methods to identify trajectories that encounter multiple KBOs are computationally intensive with impractically long run times on the order of months.

The present work deals with applying a rapid low-fidelity technique to identify candidate preliminary KBO sequences, to be used as a starting point for futuremission designers to identify trajectories tomultiple KBOs. This pathfinding approach uses two consecutive grid search types.

First, awell-tuned grid search is implemented using Lambert arcs to reach a first KBO, including the introduction of a new, systematic approach to identify the step sizes in target-body arrival date for the Lambert-based grid search. Then, second and third KBO encounters are assessed from the results of the first grid search and second KBO search respectively. Both second and third KBO searches are performed with a new algorithm consisting of a time grid along with STM linear propagation for maneuver calculation. Finally, an example trajectory resulting from this technique that encounters two KBOs is given as a potential flyable route.

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.

The conducted study investigates the potential of a newly released multi-phase solver to simulate atomization in liquid rocket injectors. The "VOF-to-DPM" solver was used to simulate primary and secondary atomization in an air-blast atomizer with a coaxial injector-like geometry. The solver uses a hybrid Eulerian/Eulerian-Lagrangian formulation with a geometric transition criteria between the two models. The conducted study assumed isothermal, non-reacting flow at room temperature. The primary focus was predicting Sauter Mean Diameter and droplet velocity data at a sampling plane downstream of the injection site. The results showed that the solver is able to produce the expected data and to predict trends similar to those found in experimental measurements. The accuracy of the produced droplet diameters was roughly a factor 2 off compared to experiment. This is attributed to mesh resolution. Measurements were obtained via a cooperative agreement between TU Delft and The University of Sydney. It was concluded that the solver has the potential to predict atomization at a reasonable computational cost, but further study is needed to confirm its full capabilities.

Current and future space missions require agile and reliable spacecraft capable of trailing and keeping the required attitude. Most of the agile spacecraft missions are near-Earth based but some are placed far away from Earth and its influence. One example of such missions is the Athena mission, which requires the spacecraft to perform fast and large-angle attitude slew manoeuvres. Such manoeuvres often imply simultaneous use of multiple actuators such as thrusters and reaction wheels (RWs). A fault in any of these actuators might lead to partial or full damage of sensitive spacecraft instruments. In this research project, a novel model-based Fault Detection and Isolation (FDI) strategy is proposed, which is able to detect and isolate various actuator faults, such as stuck-open/closed thruster, thruster leakage, loss of effectiveness of all thrusters, and change of RW friction torque due to change of Coulomb and/or viscosity factor. Moreover, the proposed FDI strategy is also able to detect and isolate faults affecting the RWs tachometer. The design of the FDI algorithm is based on a multiplicative extended Kalman filter, a generalised likelihood ratio thresholding of the residual signals, and a logic algorithm which unequivocally link the faults to the symptoms. The performance and robustness of the proposed FDI strategy are evaluated using Monte Carlo simulations and carefully defined FDI performance indices. In addition, the influence of faults’ magnitudes, times of fault occurrence, and uncertainties’ magnitudes on the FDI system performance are evaluated. Preliminary results suggest promising performance in terms of detection/isolation times, miss-detection/isolation rates, and false alarm rates. Also, uncertainties on the spacecraft inertia seem to have a negative impact on the FDI performance. In order to fully understand the research project presented here, graduate-level knowledge on rigid body dynamics and kinematics, control theory, and filters applied to estimation might be required. If any of these areas are not known by the reader, it is recommended to read some of the associated literature referenced in the bibliography.

Large-scale liquid rocket engines require high-speed turbopumps to inject cryogenic propellants into the combustion chamber. Modeling the impact of secondary path and leakage flows on the performance and axial thrust of the pump is critical during the early design phase. Previous studies derived simplified models and empirical correlations to model leakage loss and determine the pressure distribution in the side gaps. The prediction capabilities of these models are subject to limitations and uncertainties mainly due to the interaction between impeller exit flow and sidewall gaps especially at partload operation. The effects of the leakage injection on the impeller flow are not fully understood and not captured by analytical correlations from literature. These can have destabilizing and detrimental effects on the hydraulic performance and the head curve, leading to a divergence betweenmodel predictions and experimental results. In this project a reduced order model is developed to capture the effects of the sidewall gaps on turbopump performance and determine the axial thrust on the rotor. High fidelity calculations of the pump and volute are performed to identify the key physical mechanisms associated with loss generation in leakage paths and improve upon existing analytical models found in literature. Axial thrust and leakage losses are modeled numerically with a single passage geometry to capture the interaction of impeller flow and leakage flow through the side gaps at multiple operating points. It is found that the recirculation of pressurized fluid, leakage mass flow, is the most impactful effect on the performance. The volumetric efficiency is captured within 16% with the analytical model and 6% with the single passage numerical calculations at nominal operation. At partload operation however, the accuracy of both models reduces. Disk friction is under predicted by the investigated analytical model and axial thrust is over predicted. For both effects the reduced numerical model shows promising results that improve the prediction capability around nominal operation and offer a higher flexibility required for the early design phase as compared with the high fidelity, full annulus calculations. For the investigated cases the leakage injection to the impeller flow does not show a destabilizing impact on the characteristic behavior.

With the introduction of Interplanetary Laser Ranging (ILR), the acquired tracking measurements accuracy of planets can be improved drastically (up to a few millimeters) which can result in an improvement in planetary ephemerides of not only the planet which is the target of laser ranging but also other bodies in the solar system due to dynamical coupling between the bodies. A quantitative analysis is performed to analyze how much improvement in the current planetary ephemerides can be achieved if highly accurate laser ranging observation were to be introduced in the current planetary ephemerides generation processes. The thesis achieved this by developing an ephemerides generation model using TU Delft's Astrodynamic Toolkit (i.e TUDAT) in which planetary ephemerides are generated for two cases. Once using simulated laser ranging observations to Mars between 2020 and 2023 and once without utilizing laser ranging observations. By comparing the estimated planetary ephemerides of the two cases, it is concluded that laser ranging to Mars resulted in a more stable and close to a factor of two improvement in ephemerides uncertainty for most of planets. The ephemerides of the Mars itself sees nearly a factor of ten improvement in uncertainty which resulted in a significant improvement in the knowledge of asteroid masses. The estimated mass parameter of 11 of the most perturbing asteroids in the solar system see more than a factor of ten improvement in their uncertainty which is unprecedented. Laser ranging to Mars and its cascade effect on the orbit of Mercury resulted in providing better constraints on PPN parameter γ and Sun's oblateness parameter J₂ allowing γ to be determined with an uncertainty of 5.5 x 10^-8 which is 3 order of magnitude better than the constraints provided from Cassini experiment. An order of magnitude improvement is also observed is Sun's J₂ parameter estimating it to an uncertainty level of 2.0 x 10^-9.