Designing high-performance and aerospace-grade heat exchangers requires detailed characterization of the as-manufactured geometry, including cross-sectional area and surface texture, to reduce uncertainties in performance prediction and issues regarding subsequent system integration. This paper presents experimental testing and analysis of microchannels fabricated using the laser powder directed energy deposition (LP-DED) additive manufacturing (AM) process. Research has shown that as-built surfaces result in differential pressure higher than what is predicted with current correlations and surface enhancements may be required for heat exchangers built using AM to meet the desired pressure drop specifications. Various surface enhancement techniques including abrasive flow machining (AFM), chemical milling (CM), and chemical mechanical polishing (CMP), were applied to the internal surfaces of the channels to tailor flow dynamics and induce variations in pressure drop. Based on experimental flow testing, channels processed with surface enhancements provide a tenfold reduction in differential pressure compared to the as-built channels. After testing, the samples were destructively sectioned to obtain geometric and detailed surface texture information. This characterization helped to inform a new prediction method for determining hydraulic diameter and equivalent sand grain roughness, thus reducing the uncertainty of predicted friction factors. The new correlation allows to estimate friction factor and pressure drop with a deviation from the experimental data that is within 20% of their value. The identification of the mechanisms at the basis of the formation of surface texture allowed to categorize distinct aspects related to friction factor ranges: roughness peaks, peak smoothing/reduction, minimized roughness, and combined waviness and valley reduction.@en

The 16U4SBSP mission concept is based on using a swarm of CubeSats to perform a scaled demonstration of SpaceBased Solar Power (SBSP) from Earth orbit. In this demonstration mission, seven identical spacecraft of 16U format are used to provide wireless energy in the kW-scale using Radio-Frequency (RF) Wireless Power Transfer (WPT), and the spacecraft in the swarm are designed to be suitable to both space-to-ground or space-to-space WPT applications. The main objective of the mission is to validate the general concept of providing SBSP using a swarm of satellites instead of a monolithic configuration, as well as some of the involved miniaturized technologies, in view of full-scale missions which could serve users in remote areas with low power requirements or support emergency operations in blackout zones affected by unpredicted hazards (e.g. natural disasters). More in general, the mission would represent a low-cost precursor towards MW-GW scale SBSP to supply clean and affordable energy from space to large areas on the Earth surface. A pre-Phase A study of the mission, funded by the European Space Agency (ESA) through the Sysnova campaign “Innovative Missions Concepts enabled by Swarms of CubeSats”, has led to encouraging results on the feasibility of the mission concept. This paper summarizes the main results of the 16U4SBSP pre-Phase A study, including the mission design and the possible way forward to the following steps in mission implementation (Phase A and later). A summary of the spacecraft system design and mission analysis is presented, as well as a short description of the mission CONOPS (concept of operations). Particular focus is given to the available options and objectives of the SBSP demonstration, and to the proposed solutions for RF power beaming, formation flying maintenance and end-of-life operations for compliance to the new ESA regulations on space debris mitigation.@en

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.

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The 16U4SBSP mission concept is based on using a swarm of CubeSats to perform a scaled demonstration of Space-Based Solar Power (SBSP) from Earth orbit. In this demonstration mission, seven identical spacecraft of 16U format are used to provide wireless energy in the kW-scale using Radio-Frequency (RF) Wireless Power Transfer (WPT), and the spacecraft in the swarm are designed to be suitable to both space-to-ground or space-to-space WPT applications. The main objective of the mission is to validate the general concept of providing SBSP using a swarm of satellites instead of a monolithic configuration, as well as some of the involved miniaturized technologies, in view of full-scale missions which could serve users in remote areas with low power requirements or support emergency operations in blackout zones affected by unpredicted hazards (e.g. natural disasters). More in general, the mission would represent a low-cost precursor towards MW-GW scale SBSP to supply clean and affordable energy from space to large areas on the Earth surface. A pre-Phase A study of the mission, funded by the European Space Agency (ESA) through the Sysnova campaign “Innovative Missions Concepts enabled by Swarms of CubeSats”, has led to encouraging results on the feasibility of the mission concept.

This paper presents in detail the final outcome of the pre-Phase A design effort for the 16U4SBSP spacecraft. The trade-off studies conducted to select all sub-systems and components are presented and their final outcomes detailed and justified, together with the technical budgets and the main areas of attention for the spacecraft design. Particularly critical for the success of the mission are the choices related to: the power transmission payload (DC-RF converter, transmitting antenna and heat dissipation system); the ADCS subsystem and in particular the sensors required to provide sufficient accuracy in the knowledge of the 3-axis attitude (both absolute and relative to the other spacecraft in the swarm); the relative navigation system, based on inter-satellite link between the spacecraft in the swarm and on a beacon link to the receiving station on ground, for efficient beaming coordination; the main propulsion system for continuous formation flying control through the whole mission lifetime; the electric power system, based on orientable solar arrays by means of a SADA mechanism and a set of batteries with sufficient capacity for beaming the required amount of power while in eclipse conditions.@en

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.

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In the original version of the book, on page xi, one of the authors listed for Chapter 2 is “R. Schnmehl”, which should be “R. Schmehl”. On page 21, the same correction needs to be made twice, in the listed authors at the top of the page and also in the footnotes. “Schnmehl” should become “Schmehl” in both the cases. This has now been rectified and the author’s name has been corrected. The correction to this book have been updated with the changes.@en

In order for off-Earth top surface structures built from regolith to protect astronauts from radiation, they need to be several metres thick. In a feasibility study, funded by the European Space Agency, Technical University Delft (TUD aka TU Delft) explored the possibility of building in empty lava tubes to create rhizomatic subsurface habitats. With this approach natural protection from radiation is achieved as well as thermal insulation because the temperature is more stable underground. It involves a swarm of autonomous mobile robots that survey the areas and mine for materials such as regolith in order to create cement-based concrete reproducible on Mars through in-situ resource utilisation (ISRU). The concrete is 3D printed by means of additive Design-to-Robotic-Production (D2RP) methods developed at TUD for on-Earth applications with the 3D printing system of industrial partner, Vertico. The printed components are assembled using a Human--Robot Interaction (HRI) supported approach. The 3D printed and HRI-supported assembled structures are structurally optimised porous material systems with increased insulation properties. In order to regulate the indoor pressurised environment a Life Support System (LSS) is integrated, which in this study is only conceptually developed. The habitat and the D2RP production system are powered by an automated kite power system and solar panels developed at TUD. The long-term goal is to develop an autarkic, automated and HRI-supported D2RP system for building autarkic habitats from locally obtained materials.@en

The 16U4SBSP mission aims to demonstrate Space-Based Solar Power (SBSP) using a CubeSat (CS) swarm from Earth orbit. This mission employs seven 16U CSs to deliver 1 kW-scale wireless energy via Radio-Frequency (RF) beaming, adaptable for space-to-ground and space-to-space applications. The goal is to validate SBSP provision using a satellite swarm and to explore miniaturized technologies for future large-scale missions. A pre-Phase 0 study funded by the European Space Agency (ESA) through the Sysnova campaign has shown encouraging feasibility results. This paper presents a study on the formation flying and orbital dynamics of a CS mission, using a model that includes Earth’s gravitational perturbations, solar radiation pressure (SRP), atmospheric drag, and lunar and solar gravity. The swarm configuration consists of seven CSs, with one at the center and six in a hexagonal arrangement. The Concept of Operations (CONOPS) is divided into three phases: deployment and acquisition, maintenance, and disposal. CSs are deployed at 30-second intervals, followed by a one-day Launch and Early Orbit Phase (LEOP) for subsystem checks. A 1000-meter formation is initially established, then reduced to 100 meters for the first half of the mission and 10 meters for the second half, maintained by a bang-bang limit-cycle controller. A disposal strategy compliant with ESA’s Space Debris Mitigation Requirements is outlined. The analysis characterizes propellant consumption at various altitudes, proposes optimal initial conditions and launch dates, and performs a trade-off analysis, resulting in a detailed mission characterization and baseline definition. The work presented in the paper proves the feasibility of the 16U4SBSP mission, which would supply clean energy from space through wireless power transfer.@en

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.

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This chapter discusses the current status of chemical and cold gas micro-propulsion systems for small satellites, with a particular focus on CubeSats. Initially, a short historical background is presented, focusing specifically on the precursor cold gas systems developed and demonstrated on small spacecraft in the decade 2000-10. A brief overview of the principle of operation of systems based on the thermodynamic expansion of a gas in a nozzle is then provided, including the simplified equations for characterizing their performance based on the one-dimensional, isentropic Ideal Rocket Theory. In particular, the specific technical challenges associated with the miniaturization of this class of propulsion systems are highlighted and shortly explained. The current state-of-the-art available commercial-off-the-shelf chemical and cold gas micro-propulsion systems is then described in detail, including comparative tables with their main design characteristics and performance parameters. Based on these tables, performance charts are presented that define a range of possible applications of this class of micro-propulsion systems to small satellite missions and tasks. Finally, some of the future expectations in the development of these systems are briefly discussed.

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Autonomous Line-of-Sight navigation represents an appealing technique that can be exploited by deep-space spacecraft, particularly miniaturized, to estimate the state during cruising. It is based on the observation of visible bodies' directions, processed onboard to estimate the spacecraft' 6D heliocentric state. Its applicability has been preliminarily investigated by feeding the navigation filter with simplified measurements, simulated by adding a certain noise on the actual direction, based on star-tracker characteristics. However, while this approach is convenient and appropriate for a preliminary study, it is not sufficient to dive into the characteristics and performance of the method, and later on to definitely prove its applicability to real missions. This is because the measurement error cannot be considered relying exclusively on the star tracker's characteristics, as the observation scenario (e.g. planet apparent size, illumination condition, stars background) does play an important role in the measurement error. For this reason, in this work, we include image processing in the simulation loop. First, we define how to simulate realistic and reliable synthetic space images as a function of hardware characteristics and observation scenarios; then we use the generated images within the simulation to compute the measurements. Thanks to this approach, it is also possible to further improve the navigation filter design. In fact, we developed an Adaptive Extended Kalman Filter to cope with variable measurement errors and dynamics conditions. This filter allows the automatic tuning of both the process noise covariance and the measurement noise matrices as a function of the scenario. With this work, we add two important pieces to the road map for fully autonomous deep-space spacecraft: performance evaluation refinement including image-processing, and design of AEKF for the technique.

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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.

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Line-of-Sight (LoS) navigation is an optical navigation technique that exploits the direction to visible celestial bodies, obtained from an onboard imaging system, to estimate the position and velocity of a spacecraft. The directions are fed to an estimation filter, where they are matched with the actual position of the observed bodies, retrieved from onboard stored ephemerides. As LoS navigation represents a really promising option for the next-generation deep-space spacecraft, the objective of this work is to provide new insights into the performance. First, the information matrix is analyzed to show the influence of the geometry between the spacecraft and the observed planet(s). Then, a Monte Carlo approach is used to investigate the influence of measurement error (ranging from 0.1 to 100 arcsec), and tracking frequency (ranging from four observations per day to one observation every two days). The effect on navigation performance is quantified by two indicators. The first is the 3D position and velocity Root-Mean-Square-Errors, computed once the estimation is considered to be steady-state. The second is the convergence time, which quantifies the required time for the estimation to reach the steady-state behaviour. The simulation is based on a set of four planets, which do not follow the common heliocentric dynamics but rotate around the Sun with the same (distance-independent) angular velocity of the spacecraft. This approach allows the separation of scenario-dependent behaviours from navigation intrinsic properties, as the same relative geometry between observer and observed objects is maintained during the whole simulation. The results provide a useful guide for the next-generation autonomous navigation system, for both the definition of hardware requirements and the design of an appropriate navigation strategy. Considerations are then applied to Near-Earth Asteroid fly-by mission scenarios for the definition of the navigation strategy and hardware requirements. It is shown the importance of relative angles between the spacecraft and the planets. In the single-planet observation scenario, when the angle between the position vectors of the spacecraft and planet approaches a null value, the estimation error decreases. In the double-planets observation scenario, when the separation angle between the two LoS directions gets close to 90°, the estimation error decreases. The main influence on the performance is driven by the measurement error, which with current technologies is shown to be able to provide a position error in the order of a few hundred kilometers, while with a lower measurement error (0.1 arcsec) it would be possible to have a position error below 100 km. Finally, it is demonstrated that tracking frequency plays a secondary role in the performance, and only influences tangibly the convergence time.

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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.

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Metal additive manufacturing (AM) is being used for mission-critical applications in both developmental and production components, driven by economic and technical benefits. Laser powder directed energy deposition (LP-DED) allows manufacturing of thin wall geometric features for various components at diameters larger than 2 m. The characterization of geometric capabilities and limitations is critical for establishing guidelines for end users of the technology. Within this study, several samples of enclosed vertical tracks were fabricated and characterized using LP-DED, with 1 mm-thick walls and varying inclination angles up to 45° using the NASA HR-1 alloy (Fe-Ni-Cr). The wall thickness, melt pool, and surface texture, inclusive of waviness and roughness, were evaluated and results presented. The experimental results indicate that the wall thickness increases exponentially above 30°. The surface texture was shown to be dependent on 1) excess powder adherence, 2) melt pool irregularities causing material droop, and 3) excess material. The experiment revealed that the mean roughness reduces with increasing wall angle for the downskin surface. The upskin roughness reaches a maximum peak at 20° and slowly reduces as powder adheres within the valleys. Both the downskin and upskin surface textures are dominated by irregular waviness generated by the melt pool.@en

The propellant storage compartments (propellant tanks) have undergone noted evolution in the design nature (mainly the shape and the structural properties) as well as the development process. To achieve high system performance for a given propulsion system, inert mass reduction as well as efficient volume utilization can be considered as the main attributes to concern the designer. Additive manufacturing (AM) techniques, on the other hand, have played a major role in recent years in altering the propulsion system design process to achieve higher overall propulsion performance due to the technical advantages of AM in reducing mass, enhancing heat transfer through enabling complex geometries and using high performance alloys. With new chances of increasing propulsion performance come new challenges on using AM propellant tanks, and chemical compatibility with green propellants is one. The relevant chemical properties of several green energetic ionic monopro-pellants are addressed, as well as an assessment of their compatibility with the main materials used in AM processes. This article is published with the permission of the authors granted to 3AF; Association Aéronautique et Astronautique de France (www.3AF.fr) organizer of the Space Propulsion International Conference.

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To allow the self-consistent modelling of plasma transport throughout the multiple flow regimes encountered inside Helicon Plasma Thrusters (HPTs) and other magnetically enhanced plasma thrusters, a coupled model is developed. The resulting model, MUPETS (MUlti-regime Plasma Equilibrium Transport Solver), is based on the self-consistent coupling between fluid and Particle-In-Cell (PIC) solvers and can be applied to the full domain of a plasma thruster and it's surrounding space. Running the fluid and kinetic solvers iteratively, found solutions are passed from each model to the other, thus eliminating the need for assumptions on the boundary conditions at the interface between the numerical domains. To achieve this both the boundary conditions as well as the numerical domains are coupled self-consistently. From the fluid model the plasma species' solved properties are coupled to the species' particle distribution function (PDF) and velocity distribution function (VDF) injected in the kinetic model. Meanwhile the solved gradient of the plasma potential ϕ in the kinetic model is coupled back to that of the fluid model. The numerical domains are distributed to match the respective fluid and kinetic regimes occurring in the thruster, with the plasma inside the source chamber assumed to be in the fluid regime, the plasma exhaust in the magnetic nozzle to be in the kinetic regime, and the interface to be at the thruster's throat at the outlet of the source chamber. The coupled model has been tested for a simulated thruster operating on Argon at 50 W input power and 0.05 T magnetic field strength. The resulting plasma profiles converge within the first 2 iterations of the coupled model. At the coupled model's domain interface, the average densities of the ions and electrons differ less than 1% and 4% between the fluid and kinetic models.

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CubeSats suffer from low reliability and have little to no Fault Detection, Isolation and Recovery mechanisms onboard. Advanced CubeSat missions such as the Lunar Meteoroid Impact Observer (LUMIO), will use more complex attitude determination and control systems, increasing the need for advanced fault detection. Traditionally this requires model-based fault detection methods which are complicated, computation heavy, and highly sensitive to disturbances. Machine learning has proven proficient at fault detection in several non-space related applications, but training data including spacecraft faults is not readily available. In this research an especially lightweight unsupervised learning method for fault detection is designed for the LUMIO attitude determination and control components. The result is a system capable of detecting artificially induced bias, calibration error, and drifting measurement faults on the scale of 0.1 milliradians per second in the IMU, with no false alarms being raised. The method was tested on simulated LUMIO telemetry from the IMU and reaction wheels as well as on real spacecraft telemetry from the OPS-SAT sun sensor, star tracker, reaction wheels, and IMU. In both cases, excellent fault detection and false alarm performance was observed, highlighting the potential of this method for application in CubeSat AOCS fault detection and isolation.

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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.

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Modern spacecraft systems 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 becoming of more interest to space systems’ designers. The paper will present a concept of smart space structures, in particular carbon fiber composites embedded with Optical Fiber Sensors (OFS) for spacecraft and launch vehicle applications. Although the application is of interest as well to electric and hybrid propulsion with suitability for a variety of relevant propellants, the case study presented is on smart propellant tanks to be used in launch vehicle’s upper-stage chemical propulsion, utilizing gaseous and liquid fluids with specific interest in green propellants (i.e., EIL and NOx classes). This study will discuss the typical mission and operational requirements for such tanks 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 highlight their core performance parameters, such as the sensing range and spatial resolution. The proposed smart tanks utilizing optical fiber sensors are found to be capable of providing precise readings of static and dynamic strain as well as temperatures on a wide range. The increased performance and reliability 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 proved capability in providing accurate real-time measurements of temperature and monitoring structural integrity while detecting precisely possible points of rupture and failure. The applications of fiber optic sensing in smart propellant tanks may extend to detecting fluid leakage and providing increased precision in propellant gauging and can be used in on-ground qualification, pre-flight testing, as well as in-orbit operation and health monitoring.@en