AbstractThis study investigates the feasibility of using space debris as a supplemental resource for Lunar infrastructure, with a particular focus on the mission design and energy requirements of debris transfer operations. While recycling methods themselves remain at a conceptual stage, this work establishes a technical baseline for how orbital debris—specifically upper stages in GTO could be captured and transported for Lunar processing. The analysis highlights the central challenge of orbital transfer alignment under long-term perturbations and evaluates multiple capture and transfer scenarios, comparing them against direct material delivery missions. Both chemical and electric propulsion architectures are assessed, demonstrating potential energy savings of up to 30 % per kilogram of material, with further reductions when rideshare configurations are employed. By quantifying the mission energy expenditure, this study clarifies the role that efficient transfer design can play in making debris recycling a viable supplement to In-Situ Resource Utilization and reducing reliance on costly terrestrial launch. The results are intended to inform future research on processing methods by first establishing the transfer architectures under which recycling missions could realistically operate.

Green SWaP (Green Solar-to-propellant Water Propulsion) is a project funded by the European Innovation Council (EIC) Pathfinder program which aims at developing the core technologies for a new class of in-space water propulsion. Water offers unparalleled handling and storage advantages both on Earth and in orbit, and through in-situ resource utilization on the Moon and other bodies, can become a renewable source of propellant. Moreover, water provides dual-use benefits in future human outposts by serving as radiation shielding and as a working fluid in life support systems. Green SWaP pioneers the direct onboard conversion of water into hydrogen peroxide and gaseous hydrogen using solar energy, yielding a propellant combination with superior storability compared to conventional water electrolysis systems. The project aims to increase the technology readiness levels of all key subsystems, including microgravity water conversion systems, concentration of hydrogen peroxide to rocket grade levels, and the safe storage of hydrogen in an inflatable tank. These propellants are then employed in two innovative thrusters: a 1 N hydrogen Solar Thermal Thruster (STT) for precise attitude control and a 200 N bipropellant engine that uses High-Test Peroxide (HTP) and hydrogen for main propulsion. By integrating these building blocks, Green SWaP lays the foundation for renewable, self-sustaining mobility in space, extends water-based propulsion to higher thrust regimes, and enables new mission architectures leveraging in-situ resources.

The EU-EIC Pathfinder project Green SWaP (Green Solar-to-Propellant Water Propulsion) develops a sustainable in-space mobility architecture that directly converts water into hydrogen (H₂) and hydrogen peroxide (H₂O₂) using solar energy. This approach enables a reusable propulsion system combining a 200 N chemical thruster for primary manoeuvres with 1 N solar-thermal thrusters (Isp ∼500 s) for attitude control. By harvesting energy in orbit and producing propellant onboard, the system enhances operational safety, supports water circularity in space, and reduces dependence on Earth-supplied resources. Such a capability extends spacecraft lifetime, enables in-orbit refueling and in-situ resource utilization (ISRU), and broadens the feasibility of reusable orbital stages. To evaluate this potential, a dedicated mission analysis was performed after a selection process, focusing on a reusable kick-stage concept as a case study. The results provide preliminary sizing of key enabling technologies, such as bi-modal propulsion, inflatable hydrogen storage, and solar-to-fuel conversion, and demonstrate the transformative impact of Green SWaP on sustainable space logistics and future mission architectures.

Spacecraft and launcher development has recently focused on new design concepts employing intelligent propulsion systems, enabled by advanced AI-based paradigms for operations and condition monitoring (OCM) and structural health monitoring (SHM). The main challenge, however, remains providing abundant sensing data points to ensure reliable OCM and SHM processes for effective onboard systems control. This paper presents a case study on a smart spacecraft propellant tank prototype manufactured through carbon fiber filament winding and additive manufacturing of carbon fiber-reinforced polymer, using a Distributed Optical Fiber Sensor (DOFS) system. The optical fiber sensors (OFS) embedding technique is discussed, highlighting methods to optimize temperature isolation from strain variation effects. Composite structure post-processing considerations are also addressed for compatibility with acrylate-coated fibers. Thermal test results, using a high-backscattering OFS interrogated by a Luna ODiSI-6000 OFDR system, are presented.

Modern spacecraft and launch vehicle design is more oriented towards reducing system-level design and assembly complexities. In order to maintain high overall system performance while reducing these complexities, the use of smart materials and smart structural components is a well-known practice and is currently of rising interest to space systems' designers. The paper discusses a concept of smart space structures, in particular, a carbon fiber composites structure embedded with Optical Fiber Sensors (OFS) for spacecraft and launch vehicle applications. This study highlights the operational requirements for such tank and the smart features enabled by the optical fiber sensors. For the latter aspect, a quantitative comparison between Fiber Bragg Grating sensors (FBGs) and Distributed Optical Fiber Sensors (DOFS) based on Optical Frequency Domain Reflectometry (OFDR) is presented to state their core performance parameters, such as the sensitivity, sensing range, dynamic measurement capability, and spatial resolution. The increased performance and reliability in harsh environments associated with fiber optic sensors come with a reduction in size, mass, and power consumption compared to the conventional electronic sensors. Optical fiber sensors embedded in carbon fiber structures have proven their capability in providing accurate real-time measurements of temperature and monitoring structural integrity while detecting precisely possible points of rupture and failure as discussed and demonstrated in the literature review. The applications of fiber optic sensing in smart propellant tanks may extend to detecting fluid leakage, also providing increased precision in propellant gauging through temperature mapping, and can be used in on-ground qualification, pre-flight testing, as well as in-orbit operation, condition, and structural health monitoring. The article presents a statement for an optimal FOS embedding approach in composite pressure vessels and discusses the related placement and orientation method for the fiber optic sensors, coupled with a one component simplified analytical stress-strain transfer model deriving the stress component along the maximum principal direction (i.e., σ_MaxPrincipal). The novel approach is believed to serve the optimal employment of embedded FOS in composite structures, e.g., pressure vessels and light-weight structures in spacecraft, among other applications. The simplified model is believed to pave the way for effective data interpretation and processing, utilizing the available limited computational resources on-board the spacecraft.

This research evaluates Laser Powder Directed Energy Deposition (LP-DED) for producing fine feature internal microchannels. This study is focused on enhancing and characterising the surfaces of microchannels produced using techniques such as abrasive flow machining, chemical milling, chemical mechanical polishing, electrochemical machining, and thermal energy method to modify internal surfaces of microchannels made from NASA HR-1 Fe-Ni-Cr alloy. Flow testing for discharge coefficient measurement is conducted on processed microchannel samples, followed by characterisation through optical microscopy, Scanning Electron Microscopy (SEM), and Computed Tomography. Findings reveal variations in surfaces due to powder adherence, melt pool undulations, and polishing mechanisms. The study emphasises the significance of removing material equivalent to the mean powder diameter to reduce surface roughness and impact the discharge coefficient. The research proposes a ratio for planarising roughness and waviness peak height and density, offering insights for tailored surface adjustments in specific applications requiring reduced flow resistance. Highlights Internal microchannels with thin-walls were fabricated using the laser powder directed energy deposition process. Various surface enhancements and polishing processes were developed to modify the surface texture of the LP-DED channels. Flow testing was conducted to determine the discharge coefficient. Post-test characterisation was completed to obtain cross sectional area, perimeter, surface texture, and general surface condition to analyse results. Ratio of roughness and waviness peak and density (Spk/Spd and Wp/WPc) is proposed as a relevant surface characterisation parameter. Tailored surface modifications for specific end-use applications.

This study proposes the concept of recycling space debris as a novel means of supplying material resources for the establishment of a permanent Lunar presence while simultaneously cleaning up Earth's orbital environment. Upon the creation of a space debris dataset and characterizing debris objects as resources and reserves, spent Ariane 5 upper stages in GTO are identified as prime candidates for recycling. However, orbital transfer alignment poses a critical challenge due to orbit perturbations over time. Mission scenarios, including debris capture, transfer and Lunar processing, are analyzed, with global mission energy expenditure used to compare them to direct material delivery missions. Both chemical and electric propulsion transfer architectures are highlighted as enabling feasible and efficient recycling mission scenarios, with potential energy savings of up to 30% per kg of material. The significant reduction in launch mass as a direct consequence of capturing the mission payload in orbit allows for the inclusion of rideshare configurations, increasing efficiency to over 60% less energy investment per kg.

Lunar meteoroid impacts have caused in the past a substantial change in the lunar surface. With no atmospheric shield, the Moon is subject to many impacts from meteoroids, ranging from a few grams to a few kilograms. The high impact rate on the lunar surface has important implications for future human and robotic assets that will inhabit the Moon for significant periods of time. Therefore, a better understanding of the meteoroid population in the cislunar environment is required for future exploration of the Moon. Moreover, refining current meteoroid models is of paramount importance for many applications, including planetary science investigations. Studying meteoroid impacts can help deepening the understanding of the spatial distribution of near-Earth objects in the Solar System. The ability to predict impacts is therefore critical to many applications, both related to engineering aspects of space exploration, and to more scientific investigations regarding evolutional processes in the Solar System. The Lunar Meteoroid Impacts Observer (LUMIO) is a CubeSat mission to observe, quantify, and characterise lunar meteoroid impacts, by detecting their impact ashes on the far-side of the Moon. This complements the information available from Earth-based observatories, which are bounded to the lunar near-side, with the goal of synthesising a global recognition of the lunar meteoroid environment. LUMIO envisages a 12U CubeSat form-factor placed in a halo orbit at Earth-Moon L2. The detections are performed using the LUMIO-Cam, an optical instrument capable of detecting light ashes in the visible spectrum (450-950 nm). LUMIO has successfully passed the PDR and is currently moving towards Phase C. We present the latest results on the modelling of the meteoroid environment in the Earth-Moon system, including an estimate of LUMIO's potential impact on our existing knowledge of meteoroids, supported by high-fidelity simulation data. An overview of the present-day LUMIO CubeSat design is also given, with a focus on the latest developments involving both the ongoing/planned scientific activities and the development of the payload.

With a growing trend in the miniaturisation of satellites, there is an increasing need to develop micro-propulsion systems for these satellites. Since scaling down conventional propulsion systems is challenging and not always possible, new concepts need to be developed. These concepts, although often based on already known systems and principles, require significant modifications to make them meet the requirements of miniaturized propulsion. One such concept is the micro-resistojet, using an electrical resistance to increase the propellant temperature. At Delft University of Technology, two thrusters based on this concept were developed: the Vaporized Liquid Micro-resistojet (VLM) and the Low-Pressure Micro-resistojet (LPM), which were specifically designed for being demonstrated on-board a PocketQube satellite. To determine the operating regime of the thrusters for this demonstration, there is a need to develop simplified analytical models that accurately predict their performance without significant computational expenses. Although there were previous attempts to model these thrusters, they did not provide a complete representation of their performance. For the VLM thruster, the focus of the model presented in this paper was on coupling the heating chamber and the nozzle, to obtain a more accurate value for the mass flow rate through the thruster. The heating chamber section was discretized into finite one-dimensional cells and convective heat transfer equations were used to model parameters such as density, pressure, wall temperature and heat transfer coefficient. The nozzle was modelled based on ideal rocket theory corrected with adequate loss factors. The mass flow rate was calculated iteratively by coupling the two sections until it reached convergence. For the LPM thruster, the focus was on including an accommodation coefficient to account for heat transfer efficiency between thruster walls and propellant. Rarefied gas dynamics equations were used to calculate performance parameters due to the low-pressure conditions within the thruster. The models proved to produce realistic results when compared to available numerical and experimental values, although still with some limitations in modelling heat transfer, which could not be fully overcome yet due to the lack of available data for validation. Optimal operating points were determined for both thrusters by maximizing an objective function based on performance parameters such as thrust-to-power ratio, specific impulse, and mass flow rate. Constraints included thrust, power, and temperature requirements, which led to different optimal points for the thrusters under varying operational conditions.

Establishing a permanent lunar base has gained increasing attention since it offers opportunities for international cooperation and the commercialization of space, forming the foundation and testing ground for a human existence independent from Earth. Essential to future missions beyond cislunar space is the exploration and in situ processing of the Moon's resources, especially the sustainable production of energetic resources and propellants. Utilizing in situ generated propellants can dramatically reduce transportation costs by removing the need to source propellants from Earth. Resources on the Moon are limited, and the extraction of available resources are energy-intensive processes demanding advanced techniques and technologies. Consequently, one of the biggest challenges lies in developing process architectures with a positive energy balance, for which comprehensive analyses are still missing. The focus currently lies on the extraction of water ice from lunar regolith and the production of hydrogen and oxygen through water electrolysis. However, alternative fuel and process options may reduce the energy cost while providing equivalent energetic revenue. In the scope of this research, the infrastructure and technologies required for extraction, refining, and storing are assumed to exist in cislunar space; therefore, only the operating cost is considered. Exergy analyses of in situ extraction methods are conducted to investigate whether the required energetic budget allows sustainable implementation. The analysis includes extraction methods and propellant options to reveal the extent to which alternatives to hydrogen are feasible. Exergy analyses determine thermodynamic losses of energy flows giving the ground for process optimization. The exergy destructed represents the margin of improvement within the process architecture and thus reflects the process's thermodynamic and economic value while allowing a more distinct examination of energy use. Assuming the availability of water and carbon dioxide ice in permanently shadowed regions, the analysis shows that choosing methane instead of hydrogen in combination with oxygen as propellants can reduce the required exergy input by up to a third. An example mission allows to directly compare the operating cost of the extraction processes for the different propellant options. The mission entails a spacecraft propelled by a liquid bipropellant engine utilizing the extracted propellant and transporting a payload of the same propellant to a depot located in lunar near-rectilinear halo orbit (NRHO). Although abundant in space, the results suggest that hydrogen may not be the only or even energetically cost-effective resource for developing cislunar and Martian space infrastructures. Likewise, sustainable extraction of propellants suitable for current and future propulsion systems will foster humanity's reach further into the solar system.

This paper aims to investigate the capabilities of exploiting optical line-of-sight navigation using star trackers. First, a synthetic image simulator is developed to generate realistic images, which is later exploited to test the star tracker's performance. Then, generic considerations regarding attitude estimation are drawn, highlighting how the camera's characteristics influence the accuracy of the estimation. The full attitude estimation chain is designed and analyzed in order to maximize the performance in a deep-space cruising scenario. After that, the focus is shifted to the actual planet-centroiding algorithm, with particular emphasis on the illumination compensation routine, which is shown to be fundamental to achieving the required navigation accuracy. The influence of the center of the planet within the singular pixel is investigated, showing howthis uncontrollable parameter can lower performance. Finally, the complete algorithm chain is tested with the synthetic image simulator in a wide range of scenarios. The final promising results show that with the selected hardware, even in the higher noise condition, it is possible to achieve a direction's azimuth and elevation angle error in the order of 1-2 arc sec for Venus, and below 1 arc sec for Jupiter, for a spacecraft placed at 1 AU from the Sun. These values finally allow for a positioning error below 1000 km, which is in line with the current non-autonomous navigation state-of-the-art.

This manuscript aims to present and evaluate the applicability of combining optical line-of-sight (LoS) navigation with crosslink radiometric navigation for deep-space cruising distributed space systems. To do so, a set of four distributed space systems architectures is presented, and for each of those, the applicability of the combination is evaluated, comparing it to the baseline solutions, which are based on only optical navigation. The comparison is done by studying the performance in a circular heliocentric orbit in seven different time intervals (ranging from 2024 to 2032) and exploiting the observation of all the pairs of planets from Mercury to Saturn. The distance between spacecraft is kept around 200 km. Later, a NEA mission test case is generated in order to explore the applicability to a more realistic case. This analysis shows that the technique can also cope with a variable inter-satellite distance, and the best performance is obtained when the spacecraft get closer to each other.

The paper presents the initial outcomes of a project, currently ongoing under the supervision of the European Space Agency, having the main objective to specify and design a Fault Detection Isolation and Recovery (FDIR) system by making use of relevant RAMS (Reliability, Availability, Maintainability, Safety) analyses for missions in non-deterministic environment with limited resources. The initial project tasks have been to select a study case represented by a CubeSat complex mission, analyse in detail both its mission and system requirements and, based on them, define a set of relevant RAMS analyses to be carried out in the second phase of the project, as inputs for the development of a FDIR concept aimed at a careful balance of the limited spacecraft resources in case of critical failures. Two possible study cases have been identified: LUMIO, a 12U CubeSat mission for the observation of micro-meteoroid impacts on the Lunar farside, and M-ARGO, a 12U deep-space CubeSat which will rendezvous with a near-Earth asteroid and characterize its physical properties for the presence of in-situ resources. Although both missions are characterized by a high level of autonomy and complexity in a harsh environment, LUMIO has been eventually selected as study case for the project. In the paper, the challenges and features of this mission are shortly presented. The specificities of the RAMS analysis and FDIR concept for this specific class of small satellite missions (including the selected study case) are highlighted in the paper, looking in particular at aspects such as the improvement of reliability while maintaining the CubeSat philosophy, the tuning of mission and system requirements in view of facilitating the design and implementation of the FDIR concept, and the current gaps within the RAMS/FDIR body of knowledge. The conclusions drawn during this first project phase provide a real view of how systems engineering must work in tandem with RAMS analyses and FDIR to achieve a more robust and functional mission architecture, thus improving the mission reliability.

Spacecraft systems monitoring is crucial for early fault detection and troubleshooting of various subsystems and components. To detect unexpected performance degradation or anomalies during the mission lifetime, multiple spacecraft subsystems require in-situ real-time monitoring and delicate acquisition of the operations' data. Subsystems on a satellite or a probe that would require such monitoring include structures and mechanisms such as the spacecraft bus and the deployable antennas, the power generation and storage subsystem represented in the solar-panels and the batteries, as well as the propulsion system and its propellant storage, fluid management system, and thrusters. As high-performance systems are intrinsically sought, the spacecraft design complexity increases and onboard allowable volume decreases. Fiber-optic sensing figures prominently in such scenarios due to its significantly reduced size, mass, and power consumption coupled with its higher performance and reliability when compared to conventional electronic sensors. The article aims at surveying the current trends in optical fibre sensors and their interrogation systems and critically reviewing the state-of-the-art. The fundamentals and working principles are discussed for point sensors, quasi-distributed, and distributed sensors based on Fiber Bragg Grating, Rayleigh, Raman, and Brillouin scattering, among others. As the opportunities and advantages of the photonic sensing systems based on optical fibres are highlighted, a major focus is put on studying the current technical challenges facing the utilization of this technology in space applications. A case study of the PROBA-2 mission's Fiber-optic Sensor Demonstrator (FSD), majorly relying on FBG sensors, is presented to draw the future opportunities in innovative space systems enabling technologies such as the fibre-optic sensor networks utilizing the modern radiation-hard photonic integrated circuits (PICs) miniaturized interrogation units.