An Atmosphere-Breathing Electric Propulsion (ABEP) system is a novel concept in Satellite propulsion that involves capturing and compressing the rarefied atmosphere at Very Low Earth Orbit and using the captured gas to generate thrust using conventional electric propulsion methods. For such a system, the gas intake system needs to provide sufficient compressibility and mass flow to sustain the operation of the thruster. A numerical investigation into the flow characteristics and performance parameters of an active ABEP device has been conducted using the Direct Simulation Monte Carlo (DSMC) method. The Intake presented here comprises of a passive converging part and an active Knudsen Pump. The presented work initially recreates two studies one each of a passive intake and a Knudsen Pump from literature to validate the DSMC simulation methodology. Subsequently additional simulations are done based on the results of these benchmark cases.

Knudsen Pumps are usually operated in the transition flow regimes for its superior performance compared to the free-molecular regime. This study initially investigates the operation of Knudsen Pumps in free-molecular flow regime. This is done because in ABEP applications, the gases are usually in the free-molecular flow regimes. The channel lengths and the ratio of the size of the narrow channel to that of the wide channel is varied in the presented work. The variation of macroscopic properties along the axis of the Knudsen Pump is obtained from the results of the DSMC simulations. The results are compared to 1D analytical solution for some limit cases. It was concluded that increasing the channel lengths and maintaining a transition flow regime inside the wide channel of a Knudsen Pump brings up its compression characteristics while operating in a free-molecular flow regime.

In the second stage of this work, a hybrid intake consisting of a passive converging diffuse wall connected to an active Knudsen pump is investigated. A numerical DSMC model for such a setup is realised and verified using convergence studies. Two different configurations of single stage and five stage Knudsen Pump is studied. In the former case, the mass flow rate through the intake was varied using a given transmissivity for the intake exit and the temperature difference within each channels is varied. It was found that higher compressibility was observed when the total mass flow rate through the thruster was lower than the maximum thermal creep mass flow rate. Whereas the compressibility increased when the temperature difference was increased from Δ𝑇 = 500 to Δ𝑇 = 1000. However, further increase in temperature difference did not result in increased compressibility. In the five stage Knudsen Pump case, two different outlet transmissivities were studied. It was observed that the pressure and mass density drops down after consecutive stages of the Knudsen pump when the mass flow rate through the Knudsen Pump was higher than the maximum thermal creep mass flow rate. This resulted in no significant pressure or density compressibility. However, when the mass flow rate through the thruster was less than the maximum thermal creep mass flow rate, there was progressive increase in the mass density while the pressure continued to drop after consecutive stages. It was concluded that in-order to obtain significant compressibility the mass flow rate through the hybrid intake needs to be very low which in-turn might not yield sufficient thrust. A closed configuration where there is only mass flow during intermittent thruster firing could be ideal for obtaining higher compressibility using the Knudsen Pump

What if repairing a damaged composite panel could be as precise and repeatable as manufacturing a new one? Despite the widespread use of composite materials in the aerospace, automotive, and renewable energy sectors, in situ repair methods remain largely manual and often fail to restore the original mechanical integrity. The increasing use of fibre-reinforced polymers as a result of their high strength-to-weight ratio, fatigue, and corrosion resistance increases the need for more advanced and reliable repair solutions. This demand motivates the development of semi-automated systems to improve both repeatability and performance in composite repair.

This thesis investigates the feasibility of a semi-automated handheld robotic system for in situ material deposition, leveraging human guidance alongside robotic precision. To this end, a prototype handheld extrusion device was developed, integrating onboard optical flow sensors for real-time positional correction. Its control architecture enables user-guided motion while autonomously compensating for tracking errors, combining human adaptability with the repeatability of robotic control to ensure consistent material deposition.

To evaluate system performance, a surface-level defect modelled as a 175 mm long crack was introduced into a wooden substrate. Material was deposited into the defect under controlled conditions using the prototype, completing the process in 35 s. The device successfully performed real-time correction and deposition, demonstrating the feasibility of augmented handheld additive repair. Average accuracies of 1.05 mm in x and 2.45 mm in y were achieved after movements of about 300 mm in global x and 70 mm in global y. This level of precision demonstrates the basic capabilities of the system, but targeted improvements will be required to meet the demands of industrial deployment.

While developed for composite repair, the underlying technology is versatile and applicable across various domains. Its ability to deposit functional materials with spatial precision suggests potential use cases in structural health monitoring and field-deployable additive manufacturing. Future extensions, such as support for curved surfaces, vision-based localisation, and integrated non-destructive testing methods, could significantly enhance system capability. These developments may enable intelligent, semi-autonomous platforms that combine human intuition with robotic accuracy for material deposition in complex field environments.

This thesis presents the structural and thermal design of a CubeSat-based rover system for lunar surface exploration, developed with standard 3U CubeSat as foundation. The study focuses on the integration of a mobility platform, capable of being deployed from a lunar lander, on to the 3U CubeSat. Detailed finite element analysis (FEA) was conducted to ensure structural survivability under launch conditions, including quasi-static, random vibration, and shock loads. The first natural frequency was found to be 267 Hz, exceeding the 115 Hz threshold across major spacecraft launchers. Maximum von Mises stress reached 300 MPa under the worst-case shock loading, yielding a minimum margin of safety of 0.10. Thermal analysis using passive insulation strategies demonstrated that all subsystem components remained within -20°C to +65°C under lunar surface conditions. Mobility performance was validated through terramechanics modelling and field testing on Earth's sandy terrain, where the prototype exhibited static and dynamic sinkage within 11‰ and 9‰ of analytical predictions, respectively. These results demonstrate that with appropriate design strategies and material selection, a CubeSat-Based Rover configuration can be feasibly integrated within a 3U CubeSat. While further studies—such as radiation tolerance, dust mitigation, lunar analogue terrain testing, and structural testing—are required, the findings highlight the potential of this architecture as a standardized, low-cost platform for short-duration lunar surface missions.
Dataset: https://doi.org/10.4121/uuid:6bd1d545-e071-4f7b-a457-e213909b878f

The Starfixers project addresses the urgent challenge of space debris mitigation in Low Earth Orbit (LEO), where inactive satellites from mega-constellations such as Starlink threaten sustainable space operations. Our mission is to design a cost-effective and environmentally responsible Active Debris Removal (ADR) spacecraft capable of de-orbiting at least ten failed satellites within one year.
The selected removal method employs plume impingement: a novel, contactless technique in which high-momentum gas jets are directed at target satellites to induce controlled trajectory changes. This avoids complex capture mechanisms and enables re-entry without requiring physical contact. The mission architecture ensures each target is individually approached and guided toward a controlled re-entry trajectory, before the "shepherd" spacecraft returns to the initial orbit for the next operation. The complete mission concept involves performing numerous controlled approach manoeuvres with each target debris gradually changing its trajectory using a custom-developed momentum transfer and transfer efficiency simulations until the debris reaches an elliptical orbit with a 381km perigee. At its final orbit the debris passively de-orbits within 7.5 months. The developed strategy is proven to be viable for debris ranging from 250- 500 kg and altitudes of 550-630 km. However, the detailed subsystem design is performed for 10 Starlink v1 satellites at 600 km circular orbit as this scenario was deemed most common for an ADR mission considering the abundance of Starlink. The mission is scheduled for deployment in January 2030 and is designed to stay within a 100 million total budget, covering all subsystem development, testing, and launch operations.
Subsystems have been developed in detail: the propulsion subsystem uses bi-propellant thrusters on a two-axis gimbal for precise plume control; the GNC and ADCS subsystems combine LIDAR, IR cameras, reaction wheels, and IMUs for autonomous attitude control, tracking, and detumbling. Power is provided by solar arrays and lithium-ion batteries to operate in the eclipse. Communication is handled via the ESA Estrack network, and all systems are designed for modularity and sustainability. During the final weeks of the DSE exercise, the team will finalise all subsystem designs, refine the plume-based and orbital targeting control algorithms, and integrate all components into a fully verified and validated spacecraft configuration.

Recent advances in AI, driven by increased computational power, are expanding capabilities in aerospace. While machine learning (ML) reveals hidden data patterns, it typically requires large datasets. Numerical simulations offer insights but are costly and sometimes unreliable due to model limitations. Experimental testing, though essential, is expensive and yields limited data. Physics-Informed Neural Networks (PINNs) integrate physical laws into ML models, reducing data needs and improving accuracy. This study presents a PINN tool developed in Python with PyTorch, supporting flexible architectures and adaptable cost functions. PINNs were applied to aerospace use cases, including temperature prediction, computational fluid dynamics (CFD), and Solid Rocket Motor modelling. Compared to traditional numerical and purely data-driven methods, PINNs demonstrated improved accuracy and data efficiency, though with higher computational costs. They also show promise as surrogate models. Future work will focus on optimising for dedicated hardware, deeper architectures, and broader applications to further explore their potential in aerospace simulation and modelling.

Architected beam lattices—ultra-light truss networks fabricated by additive manufacturing—have emerged as a promising material class for protecting different structures against impact, including defending spacecraft against micro-meteoroid and orbital debris (MMOD) impacts. Their high specific strength, energy absorption and tunable mechanical properties make them ideal for various high performance applications including impact protection. Current work on lattice metamaterials has regarding their energy-absorption potentially has identified three distinct impact modes: bounce-back, which is characterized by the impactor rebounding from the lattice; capture, where the impactor becomes lodged in the lattice structure; and penetration, where the impactor breaches the lattice and continues through it. In this work the 3 regimes are reproduced using a Discontinuous Galerkin framwork that integrate Kirchoff-Love beams, Cohesive Zone Model and Penalty Contact Law to model the impact of a rigid impactor on a lattice metamaterial. The framework is used to predict large-deformation contact and fracture in high-fidelity lattice structures struck by rigid projectiles at up to 350 m/s. As a result trends are identified with regards of impact velocity, lattice density and impact regimes. Additionally, key issues are identified with the current setup that give insight into metamaterial impact simulations.

Research in miniaturization of satellite technology has enabled the development of a new class of satellites- PocketQubes. The extremely small form factor greatly limits power generation capabilities, with body-mounted arrays expected to generate around 1W average orbital power for the 3P size. Deployable solar arrays present an ideal solution for small satellites, enabling more demanding payloads onboard. Numerous designs for CubeSats have already been developed commercially and in academia.
This research focuses on adapting solar array designs available in literature to 3P PocketQubes, addressing their unique size, power and deployment constraints imposed by the deployer. A 4-panel array design in a wing configuration was designed, facilitating a peak power of 19.7 W. This was supported by a burn-wire mechanism and a torsion spring hinge. Structural analysis was conducted using LS-DYNA and ANSYS Mechanical to simulate deployment impact and launch vibrations, respectively. The analysis evaluated the design’s response to dynamic launch loads and mechanical impacts during deployment. The final assembly weight was 204.8 g, contributing to a specific power of 96.2 W/kg, comparable to many COTS deployable solutions for CubeSats. The procurement cost was found to be €1480 (excluding solar cell assemblies), and can be greatly lowered with higher order quantities for formation flying/distributed missions with multiple PocketQubes.
Additionally, this research examined the impact of solar array deployment configurations on power generation and orbital lifetime. Power generation was assessed using a simplified Python model, later verified through AGI STK simulations. Results indicated that the 𝛽 angle significantly influences power output across different configurations. Moreover, for peak 𝛽 angles representing noon-midnight and dusk-dawn orbits, seasonal variations were studied. This yielded an approximate 30% reduction during summer solstice for dusk-dawn orbits, but no visible change for noon-midnight orbits. For velocity aligned PocketQubes, configurations with panels mounted on the 5 × 5 cm face at a 135° deployment angle were found to be optimal. In contrast, for PocketQubes with pointing capabilities and high peak power requirements, the previously designed wing configuration was recommended.
The study of orbital lifetimes was facilitated by ESA’s DRAMA software and CROC was used to identify the minimum, average (random tumbling scenario), and maximum cross sections, contributing to a wide range of expected orbital lifetimes. In the velocity-aligned case, configurations featuring panels attached long edges (parallel to the drag force) yielded lifetimes suitable for long Earth observation missions. The latter configurations were found to be suited for shorter, technology demonstration missions. The launch date significantly impacted the results of the study, especially for configurations with higher ballistic coefficients. The results from this study however are highly idealistic, as the PocketQube angle of attack is expected to vary, largely influencing the drag area experienced. Assumptions regarding the drag coefficient and the use of NRLMSISE-00 drag model within DRAMA also limit the accuracy of the results, as seen when comparing the simulation and real observations from Delfi-PQ.

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

With the advent of Reusable Launch Vehicles (RLVs) on the European launch market, the German Aerospace Centre (DLR)’s In-Air Capturing (IAC) recovery technique promises a competitive reduction in launch costs. Central to this IAC procedure is the Aerodynamically Controlled Capturing Device (ACCD), which is deployed from a Towing Aircraft (TA) by means of a tether. It intercepts a returning RLV, and provides a structural interface between the TA and RLV during tow-back. Once in the vicinity of a landing site, the ACCD releases the towing connection, and the RLV lands autonomously. Whereas the aerodynamic shell of the ACCD has been defined and studied in the past, an in-depth exploration of its electromechanical design space has not been performed previously. This work proposes a systems engineering approach to fill the corresponding design gap, based on the definition of key requirements and risks characterizing the ACCD’s functionalities, and its operational environment.

First, the behaviour of the ACCD as part of a larger towing system is analysed. For this, a two-dimensional (2D), quasi-steady-state Towing Model is developed, enabling a swift exploration of the ACCD’s design space, as well as a preliminary estimation of extreme tow-back loads. This Towing Model shows that the ACCD’s towing position relative to the TA is highly sensitivity to the ACCD’s Centre of Gravity (CoG); in order to stay clear of the TA’s disturbing wake zone, a maximum allowable CoG position of 1 m behind the vehicle’s nose is defined. Additionally, a significant pitch control authority is demonstrated, indicating substantial manoeuvrability. Finally, a study of extreme tow-back conditions results in the definition of a 300 kN axial towing design load - representing a 72% increase compared to previous estimates.

Next, the ACCD’s relative navigation system is studied, used to estimate the position and attitude of the RLV. A comparison between state-of-the-art systems is made, and the so-called VisNav solution is proposed - with a sensor attached to the RLV, and active beacons mounted on the ACCD. In order to identify a representative configuration for this system, a geometric VisNav Model is developed, analysing the visibility of beacons from the sensor’s Point of View (PoV). Based on comparative studies performed with this model, a design with three asymmetric rings of twelve beacons is proposed. Combined with Global Navigation Satellite System (GNSS) and Inertial Navigation System (INS) data, this configuration enables relative position and attitude estimation with an accuracy of 20 cm and 1°.

A representative electromechanical design for the ACCD is then proposed, based on results from the foregoing design space exploration. On the one hand, the selection of Commercial-off-the-Shelf (COTS) avionics is covered, as well as a preliminary dimensioning of the vehicle’s power system. Here, the ACCD’s actuators are identified to be the main drivers of its electronic design in terms of Size, Weight and Power (SWaP) footprint. Additionally, the feasibility of a battery-powered design is demonstrated. On the other hand, a Computer-Aided Design (CAD) study of mechanical subassemblies is performed, which are further analysed in terms of extreme operational loads. Most noticeably, a separation of the docking and release systems is proposed, as combining both into a single mechanism is deemed infeasible, due to the continuous presence of significant towing loads. Based on the presented design, a Bill of Materials (BoM) for the ACCD is established, yielding a total functional mass of 159.14 kg. Because of the substantial impact of the vehicle’s wings and docking system on its overall CoG, an additional 16.3 kg trim mass is required to reach the desired CoG position. Compared to previous studies, the final estimate for the ACCD’s total mass has increased by 33%.

Concluding the study, compliance with the vast majority of the postulated design requirements is demonstrated, while a number of future improvements and complementary studies are identified. These include an aerodynamic redesign and shortening of the ACCD’s shell, to eliminate the need for a trim mass. Additionally, prototyping of this unique vehicle is highly recommended, in order to further study its behaviour, and validate the theoretical design proposed in this work.

Stiff and transparent materials are essential across industries such as aerospace, defense, and consumer electronics. Traditional materials like glass and ceramics, while effective, have limitations due to brittleness, high density, and low impact resistance. Consequently, there is a demand for lightweight, durable materials with favorable optical properties. Fiber-reinforced polymers (FRPs) offer high strength-to-weight ratios and greater impact resistance than glass and ceramics. Given the low light absorption of glass fibers in the visible spectrum, transparent glass fiber-reinforced polymers (TGFRPs) show promise as replacements for traditional transparent materials.

This thesis explores two main aspects: (1) optimizing TGFRP transparency through manufacturing parameters and (2) leveraging TGFRP transparency for damage detection and stress analysis. In the first phase, findings demonstrate that achieving high surface smoothness and maximizing light transmission within the green spectrum (520–600 nm) are crucial for enhancing TGFRP transparency. This is best controlled via post-curing, a more stable approach than adding methyl methacrylate (MMA) as suggested in existing literature. Sizing was also found to play a critical role in transparency, ensuring optimal bonding between glass fibers and the matrix. Additionally, minimizing the number of glass fiber fabric layers improved transparency by reducing transmission losses across visible wavelengths.

The second phase investigated the transparency of TGFRP for visualizing internal damage and stress distributions. Confocal microscopy was used to observe cracks at various depths within TGFRP samples, while image processing techniques, such as thresholding, were employed to calculate crack density. The results indicate that a higher crack density corresponds to reduced light transmittance across the visible range. However, some limitations were encountered, as cracks and scratches on the upper layers cast shadows on the lower layers, complicating the detection of deeper damage.

During tensile testing of open-hole specimens, significant opacity changes were observed in high-strain regions, suggesting that opacity correlates with localized stress. Digital image correlation (DIC) and grayscale histogram analysis provided insights into the strain thresholds at which opacity changes begin. Future work could build on this by examining the extent to which opacity changes are due to refractive index variations alone, independent of crack formation, using these strain levels as a baseline.

Furthermore, verifying the reversibility of opacity changes through post-test spectrophotometry could position TGFRPs as highly effective materials for developing innovative and reliable non-destructive testing, damage detection, and stress visualization methods in glass fiber-reinforced polymer composites.

Recent studies have shown the feasibility of (quasi-)pole-sitter orbits at Mars and Venus, which involves a satellite positioned along or near the polar axis of a planet in order to have a continuous, hemispherical view of the planet's polar regions. In order to further demonstrate the feasibility of this mission concept, this thesis investigates time-optimal solar sail transfers to these (quasi-)pole-sitters. In particular, (quasi-)pole-sitters which are achievable when assuming solar sail technology expected in a near- to mid-term time-frame. To reduce mission operational cost, the objective of this research is to minimize the time required for the transfer, which requires the solution to an optimal control problem. Initial guess solutions for this optimal control problem are provided through two completely different techniques, in order to compare and validate the individual performances: first, a technique derived from dynamical systems theory (a type of grid search) and second, a genetic algorithm. Subsequent optimization using a direct pseudospectral algorithm results in time-optimal transfers to the considered Mars (quasi-)pole-sitters that span 2.61 and 2.72 years, and 1.07 and 1.19 years to the considered Venus (quasi-)pole-sitters. Effects due to variations in performance of the ideal sail, non-ideal sail properties, and Earth departure orbit are investigated. In addition, this paper demonstrates that a genetic algorithm is well suited to generate initial guesses for similar interplanetary transfers in the inner solar system. It provides initial guesses that outperform the more conventional grid search technique, in terms of feasibility of the initial guess transfers, as well as in computation time and ease of implementation.

The use of Taylor Series Integration (TSI) for the propagation of spacecraft trajectories and its combination with the relatively unknown Unified State Model (USM) coordinate system, resulting in the TSI-USM propagator, is known to allow fast propagation of interplanetary low-thrust trajectories. Unfortunately, the TSI-USM propagator is a problem-specific method, which is detrimental for its use in evolutionary algorithms.

In this thesis, a method was developed to generalize the TSI-USM propagator. The TSI-USM was generalized to allow the propagation of a spacecraft trajectory in a central gravitational field subject to any pre-set thrust profile, without the need to adapt the internal working of the propagator. The three-dimensional thrust profile should be specified a priori, and with respect to time, by the user or by an evolutionary algorithm. The method used to generalize the TSI-USM requires the coefficients of the cubic spline interpolation of the pre-set thrust profile to compute the recurrence relations of the TSI-USM.

Results show that, although the method indeed generalizes the TSI-USM, it is less accurate and requires more CPU time for a given accuracy than existing RK8(7)13M-based propagators. From this result, it can be expected that future attempts to generalize the TSI-USM propagator, without sacrificing its accuracy and computational speed, will be challenging.

the aim of this thesis is to regard all aspects concerning mass optimization with respect to small satellite structures and deployers, specifically aimed at the DelfiPQ PocketQube mission.