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.

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%.

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 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 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.

There is a trend in the miniaturisation of satellites. They not only become smaller, but lighter and more powerful as well. A large growth in the nano-satellite sector is present in the form of CubeSats. These milk carton size satellites use commercial of the shelf components in order to keep the costs low and to allow for technology demonstrations for future larger space missions. In order to increase their current mission lifetime from 1 - 2 years, electric micro propulsion systems are being developed. These systems are able to provide sufficient velocity increments in order to maintain the CubeSat within the required orbit. Electric micro propulsion systems make use of Coulomb forces to accelerate charged particles. The method is to inject a propellant gas such as xenon into a cylindrical discharge chamber. The propellant is ionised by means of electron bombardment. The electrons that are required in order to do so are pulled from the cathode into the thruster system. Inside the discharge chamber they experience an electromagnetic field, where the resulting Lorentz force makes them gyrate in the projection plane that is perpendicular to the magnetic field lines. As the propellant is inserted in the discharge chamber, a neutral gas pressure builds up. Next, the electrons that are emitted from the cathode are colliding with the neutral gas particles, which causes them to be ionised. These charged particles get accelerated out of the discharge chamber by the electrostatic field in order to create the required thrust. The goal is to develop an affordable, low complexity and efficient cathode for a high performancemicro propulsion system that can be used on small satellites. In the future, the whole system can be miniaturised in order to realise an electric thruster system that can be implemented on CubeSats in order to expand their mission lifetime. The current graphite needle tip cathode can be successfully used as an electron source for propellant ionisation and plasma plume neutralisation in combination with the xenon fed engineering model electric thruster. The thermionic LaB6 cathode is capable of achieving an emission current up to 64 mA with 56 Watts of input power. This is realised by using graphite needle tips on the end of molybdenum posts that heat the LaB6 emitter pellet. The micro High Efficiency Multistage Plasma Thruster is operated with a mass flow of 2.0 sccm. It uses an anode voltage and anode current of 700 V and 63 mA respectively. The cathode current is equal to 5 mA. Furthermore, the nominal power to thrust ratio is equal to 25.5 W/mN and the specific impulse that is achieved equals 930 s, dependent on the applied settings. The divergence efficiency ranges from 80 - 90 % and the total system efficiency is currently equal to 80 %. This total efficiency needs to be improved in the subsequent iterations of the design.

Progress in Laser powder bed fusion manufacturing has led to greater use of Additive Manufacturing in combustion chambers/nozzles. This allows for the creation of more intricate and efficient solutions, but generates the need for a better understanding of how surface roughness in micro-channels used for engine cooling relates to build angle and performance. The research presented in the framework of this thesis assesses the friction factors and internal surface roughness of 18 additively manufactured channels, through the means of flow testing and microscope inspection. Positive correlations between build angle and friction factors were found, and further investigation revealed relative roughness levels outside the applicability range of Moody charts. Comparison of open and closed contour channels confirmed roughness similarity, providing valuable input for the manufacturing of witness channels in cost/time sensitive applications. Finally, the relation between surface roughness parameters and sand-grain roughness were investigated, revealing their dependency on flow properties.

Future Mars exploration missions require safe and controlled landing on the planet’s surface. Conventional entry, descent and landing (EDL) technologies, such as parachutes and rigid aeroshells, face limitations in meeting increasing demands for heavier payloads, harsher entry conditions, and desired landing locations due to their limited deployment windows or geometry constraints set by current launchers. The stacked-toroid inflatable aerodynamic decelerator (IAD) has emerged as a promising EDL technology which has the potential to enable new and ambitious applications. Unlike conventional aeroshells, it utilizes flexible, high-temperature resistant materials that can be folded during orbital injection and transportation, and subsequently deployed before entering the Martian atmosphere. To address the complex interdependence of design variables and the multidisciplinary nature of stacked-toroid analysis, this research proposes a novel Multidisciplinary Design Analysis and Optimization (MDAO) framework. The framework integrates aerodynamics, aerothermodynamics, structural analysis, mass estimation, and Flexible-Thermal Protection System (F-TPS) sizing with trajectory simulations for a parametrized stacked-toroid. The major design variables are parameterized to trace model responses back to the design space. An additional novel contribution is the inclusion of a smaller torus on the IAD’s shoulder. For aerodynamics and aerothermodynamics modelling, the local-inclination panel method is implemented, separately addressing the continuum, rarefied, and transitional flow regimes using well-established analytical methods and bridging functions. The scalloping phenomenon of the deflected surface is accounted for through semi-empirical expressions and experimentally-fitted correlations, capturing the additional aerodynamic and aerothermal contribution. F-TPS sizing employs a 1D Finite Difference Method (FDM) with harmonic weighted averaging of material characteristics to accommodate abrupt thickness variations. Results from each discipline are compared to experimental, flight, and high-fidelity numerical data, showing consistent agreement under various conditions. All disciplines present mean percentage errors in the order of 10-20% which are deemed acceptable for early design stages. The framework is applied to the ESA MiniPINS study, to demonstrate its applicability to a novel EDL architecture for a penetrating probe. The proposed environment efficiently evaluates the stacked-toroid optimum design for minimum mass, weighing only 3.72 kg whilst complying with mission requirements and optimisation constraints. The design remains robust against variations in entry parameters and atmospheric density, requiring minor adjustments for the penetrator impact speed. The framework enables design space exploration, revealing trends favoring stacked-toroids with low inner torus radii and large numbers of tori to minimize aerothermal loads while ensuring sufficient aerobraking. The inflated radius remarkably results in a global variable that can further simplify the design space to a single input for near-optimum rapid evaluations. The feasible parameter ranges identified through the design space search expedite the evaluation and optimization of stacked-toroids’ multidisciplinary performance, aiding decision-making in early design stages of future Mars missions.

Micro-propulsion technology is under rapid development and considerably extends the mission capabilities of small spacecraft. However, the flow in conventional nozzles on a microscale is relatively viscous, which, combined with the bounding nozzle walls, leads to low efficiencies. Aerospike nozzles, on the other hand, offer free-flow expansion and mitigate the boundary layer effects.  The research objective of this project is to investigate the losses and output efficiency of micro-nozzles by optimizing a 3D axisymmetric aerospike in the μN-mN range. Planar nozzles, typically found in micro- propulsion subsystems, limit the flow radial expansion and considerably suffer from viscous losses, as well as momentum loss. In addition, in lower thrust settings, the boundary layer massively limits the performance of conventional nozzles – conical and bell layouts –, whereas aerospikes do not bound the flow and, subsequently, inherit better gas expansion. The design optimization relies on numerical analyses, with the free and open-source OpenFOAM’s solver, rhoCentralFoam, and self-developed Python scripts that automate the simulations in various machines.  This study comprises three design iterations. The first contour parameters derive from Delft University of Technology’s Vaporizing Liquid Micro-thruster (VLM) requirements and the preceding work. However, since this work considers axisymmetric geometries, the aerospike’s throat width is reduced from 45 μm to 15 μm to preserve the original thrust magnitude (< 10 mN).  The initial results show that, at the same throat Reynolds number, the aerospike outperforms the bell nozzle, especially towards lower Reynolds, where the specific impulse efficiency variation tops 25%. However, at equal thrust levels, the conventional design surpasses the aerospike thrust and specific impulse efficiencies. The Mach contours reveal that the small throat width and high area ratio ineffectively mitigate the viscous losses and lead to extreme momentum loss.  The following iterations with four truncation percentages (20%, 40%, 60%, and 80%) prove the first hypothesis right: when decreasing the area ratio from ~17 to ~3 and raising the throat width to 30 μm, the efficiency of an 80% long aerospike reaches ~94% for the specific impulse and ~89% for the thrust. In addition, the aerospike yields the best performance when it is the least truncated (highest truncation percentage, i.e., 80%).  Finally, a ±100 K temperature sensitivity study shows that the aerospike performance oscillates up to 3%, with a maximum thrust efficiency of 91% and specific impulse efficiency close to 98%, rivaling macroscale performance. With a small area ratio and a wide throat, the aerospike nozzle outperforms an equivalent bell nozzle by more than 10% in terms of specific impulse and thrust efficiencies.

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.

During the last decades, large deployable structures are starting to be seen as a plausible configuration to multiple space missions, such as solar sailing, LEO (Low Earth Orbit) deorbiting missions or solar power generation. Technological advances in key areas, such as thin film solar cells or new deployment methods, as well as the miniaturization of satellites and their components, have considerably increased the usefulness of this design. The research presented in this document aims to contribute to these technological developments, helping to unfold the whole potential of this structural solution. This research comprises the MSc thesis of the author, in partial fulfillment of the MSc in Aerospace Engineering by the Technical University of Delft. The research was conducted during an internship at the department of Guidance Navigation and Control (GNC) Systems, Institute of Space Systems, German Aerospace Center (DLR). The research focuses in the study of the behavior of large thin flexible structures (LTFSs) in space from the perspective of the attitude determination and control system (ADCS). This work was performed in close relation to the DLR’s mission GoSolAR (Gossamer Solar Array). The contributions of the research can be classified in two areas, related to the modeling of the structure and to the design of the ADCS. In relation to the first one, a methodology to integrate the flexibility of a structure into its equations of motion is explained from a theoretical perspective and implemented over a flexible spacecraft. This methodology is based in Lagrange’s equations and can be applied to diverse structures, allowing modeling accurately the dynamics of the system while maintaining the model comprehensive and of relatively low order. The implementation of this approach leads to equations of motion containing both rigid and flexible motion, enabling to conduct an evaluation of the impact of the flexibility of a structure on its dynamics in a space environment. The second major area in which this thesis aims to contribute to the body of knowledge is related to the design and evaluation of attitude controllers targeting flexible satellites. Different controller’s designs are proposed, based in: 1) a linear-quadratic regulator (LQR) approach, 2) robust control theory, particularly in the minimization of HInf and H2 norms, and 3) a control approach based in analytical dynamics, known as the Udwadia–Kalaba approach. The performance of these designs was evaluated not only in relation to control instructions related to the rigid motion of the satellite, e.g. angular rate, but also to the capability of each controller of damping the system, reducing the vibrations that appear due to its flexibility. The particularities that need to be considered when implementing each controller are also studied, in order to give a clear idea of when would each controller be an adequate control solution. The result is a set of different control designs, including for each of these designs: 1) the methodology to derive the controller, 2) the performance of the controller and 3) the particularities and limitations of its implementation.

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.

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.

AE faculty at TU Delft is currently developing a PocketQube satellite under the name of Delfi-PQ. Following the previous study, performed on a high level systems engineering on Delfi-PQ propulsion sub-system, there is a need to verify the propulsion sub-system's design, its final operational envelope values, and propulsion sub-system's requirements by using TEST and ANALYSIS methods. Leak, volume and mass, solenoid valve PWM, and heater efficiency tests have been performed. The verification tests of the propulsion sub-system design provided a lot of valuable information. Not only some of the its requirements have been verified, but also the underlying issues have been observed, registered, and described for future researchers who are going to test such a micro-propulsion system and optimize its design.

Space exploration could be significantly aided by combining the disciplines of machine learning and computer vision, but these disciplines need to be developed further for specific space-related applications to have merit. One of the applications for space exploration is the detection of certain structures designating areas of interest. This thesis demonstrates a method of structure-detecting that is applied to staircases. In addition to incorporating certain physical features, like other algorithms have done, the proposed algorithm (Step-1) also takes into account the spatial relation between these features, in order to increase its robustness. Looking at a staircase from the front, the distances between each step become warped, as they are further away from the observer. This exponential spatial distortion is known as a ’chirp’. Step-1 tries to match a chirp-waveform to every edge along a straight line randomly drawn through an image, and based on that match classify the image as containing a staircase or not. The random lines are then weighted based on their effectiveness using Adaboost, which are finally combined to obtain a final classification. The results show potential but there are still some issues to be addressed. However, once the algorithm has been upgraded it could aid space exploration by being applied to satellite images and autonomous rovers.

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.

As the CubeSats have been getting more energy dense over the years, a greater need has come for thermal control on CubeSats. Phase Change Materials (PCMs) are researched in this thesis as an energy dense passive Thermal Storage Device (TSD). This energy density comes from the melting of the PCM. To determine the feasibility of this TSD. The thermal storage device needed to be able to keep certain components from overheating while keeping other components from getting too cold. This thesis successfully tested that the PCM was able to store heat from a radio and use that heat to keep a CubeSat battery warm. The TSD is meant to stabilize the temperature of all components connected to it by maintaining a temperature around the melting temperature of the PCM. Prototypes were made using eicosane wax, it is a paraffin wax with a melting temperature of 36.7∘C. This PCM was chosen due to its safe and predictable properties, and because of its previous use in the space industry. Two different casings where tested with two different filler materials and casing sealing methods. Although the filler materials had a relatively small effect on the thermal response of the casing at the heat fluxes tested. The fin filler material used less mass and was simpler to manufacture than the honeycomb filler. There was a large difference in the sealing method of the casing that was glued together sealed better than the one laser welded together, the casing which had fill ports that were clamped with a thread with a rubber filament O-ring between them sealed better than the one that was glued on. Although the degree to which this could be tested was limited and is therefore difficult to draw definitive conclusions over. The tests proved that the TSD was much more effective than traditional heat sinks at storing heat. This will allow energy dense CubeSats with large intermittent heat-loads to be designed for their average temperatures and not for extreme peaks and valleys. The complexity this casing adds to the CubeSat is more than traditional heat sinks but much less complex than active thermal control systems. The casing thermal response could be predicted within a 5∘C accuracy using a simple Finite Element Method (FEM) model using the built-in conduction and Phase Change Material modules in Comsol. The remaining obstacle with this research project is that more work needs to be done to be able to prove that the casing is completely hermetically sealed and will not contaminate other hardware while in operational.

This thesis studies the radiation dose deposited on a 6U CubeSat in interplanetary orbit. The European Space Agency’s radiation data tool SPENVIS has been used to determine the proton flux in interplanetary space during a solar particle event which occurs only once every one hundred years. SPENVIS has further been applied to obtain a dose depth curve for this specific 1-in-100 years proton flux. Simultaneously, 6U simplified models of several materials are created as well as a realistic 6U CubeSat structure which represents the state-of-the-art in CubeSat industry. Using radiation simulation tool Fastrad, two simulations have been performed: ray-tracing simulation and forward Monte Carlo simulation. Ray-tracing simulation provides a sector shielding analysis of the models, giving an overview of the received radiation dose for each specific point in the models. Forward Monte Carlo gives an overview of the deposited radiation for each part of the model, which has shown especially beneficial for the state-of-the-art model, in which deposited dose has been obtained for each part in the model. This has led to a dataset that shows the radiation dose deposited at each location and in each part of the 6U CubeSat structure during a 1-in-100 years solar particle event that strikes in interplanetary space.