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In recent years small satellites have shifted from being secondary items to dominate the space market thanks to, but not only, the development of the large LEO constellations. For small satellites to be highly effective, particularly when arranged in such constellations, each satellite has to be placed in a specific orbital plane and orbital position. Currently, three main options are available to reach a specific orbit. The first, more straightforward solution is to employ a dedicated launch with a small launch vehicle. Plenty of small launchers are in development around the world (with few already operational); however, their cost per kilogram is predicted to be much higher than for large launchers. The second possibility is the exploitation of a rideshare option, where the satellite is transported by a large launcher together with other payloads on a general predetermined orbit. Afterwards the satellite needs to be transferred to its designed orbit. This can be done in two ways: through the use of a satellite carrier (also called self-propelled dispenser) or with the satellite own propulsion system. In both last cases, the mass transported by the launcher into the release orbit is higher than the final one necessary for the nominal mission, impacting total costs. In this paper these three possibilities are compared considering the need to reach various specific orbits, starting from a different release one in the case of the rideshare options. First of all, the change in velocity for different orbital parameters (altitude, eccentricity, phase, argument of perigee, inclination, RAAN) is computed. Afterwards the propulsion mass budget is calculated. Everything else being equal, it is demonstrated mathematically that a dispenser is inherently less efficient than a group of autonomous satellites, particularly for the RAAN change and a large number of carried satellites. However, it is not always possible or convenient to provide small satellites, particularly the smallest ones (nanosats/cubesats/microsats) with comparable propulsion capabilities of a larger dispenser, making the latter still an attractive option in several situations. A cost analysis also shows that, particularly for sophisticated small satellites, when the final orbit is far from the release one, a dedicated small launch vehicle can be cost competitive with the nominally much cheaper large launcher. • Satellites can reach a specific orbit through dedicated launch or rideshare plus orbital transfer. • Sat carriers are mathematically inherently more inefficient than autonomous propulsion. • Penalty increases particularly as the ΔV and the number of satellites increase. • Small launch vehicles are inherently more responsive and expensive than rideshare options. • Launch cost penalty impacts only a fraction of total costs, particularly for unserved orbits. [ABSTRACT FROM AUTHOR]

Space robots are inseparable components of many space missions. However, capabilities of space robots with rigid manipulators are restricted in confronting with unknown and confined environments and situations require high safety. To overcome these defects, the potential of bio-inspired soft robots for space applications has recently been addressed in the literature. Dynamic modeling and control of soft manipulators are challenging due to their infinite degrees of freedom. In this paper, a general comprehensive three-dimensional dynamic modeling framework is developed for a floating soft manipulator-spacecraft system employing the Euler-Lagrange method. This framework derives the coupled equations of motion adapting the generalized Jacobian approach. Consequently, the opportunity of exploiting model-based control methods will be provided. In addition, the proposed model avoids numerical singularities once the bending angles are near zero. Furthermore, the generalized Jacobian matrix is defined for any arbitrary point of the soft manipulator essential for the contact dynamics. Eventually, to verify the proposed dynamic modeling procedure, the dynamic response of a floating soft manipulator-spacecraft system released from an initial configuration is investigated. Then, a sliding mode controller is designed for the floating soft manipulator to track the desired shape and compared to the proportional-derivative control scheme. The obtained results follow the physical principles and fulfill the control objective. • To overcome defects of usual rigid manipulators, soft space robots are concentrated. • A thorough 3D modeling frame is presented for soft manipulator-spacecraft systems. • Thorough dynamic modeling process is established using the Euler-Lagrange approach. • The generalized Jacobian matrix is derived for any optional point of the soft arm. • Application of the sliding mode control is proven for the soft space manipulator. [ABSTRACT FROM AUTHOR]

Measurements of Martian magnetic field, atmosphere and sampling are core missions of Martian exploration. Since the successful demonstration of the Ingenuity Helicopter, a large amount of researches have emerged on Mars rotor-craft. However, Mars rotor-crafts have short endurance and limited detection range. This paper proposes a solar-powered tilt-rotor unmanned aerial vehicle (UAV) named Archaeopteryx-I, which can perform vertical take-off and landing (VTOL) and long-endurance level flight. Its general concepts are first elucidated. Then, a multidisciplinary modeling and optimization method is proposed. Considering the harsh magnetic field and atmospheric environment of Mars, the detailed designation of mission system is carried out to further enhance the performance of Archaeopteryx-I, including avionics, energy flow, control system and electromagnetic compatibility (EMC) design. Finally, the conceptual design scheme is validated through simulation and ground testing. The results illustrate the efficaciousness and applicability of Archaeopteryx-I for Mars exploration. • A novel solar-powered tilt-rotor UAV for Martian medium granularity exploration. • The proposed conceptual design framework can deal with multidisciplinary constraints. • Avionics, energy, control, and EMC can enable the UAV to flight in high-performance. • In each cycle, the UAV can cruise on Mars at 33.24 m/s for 0.5 h and hover for 2 min. • The tilt-rotor UAV can achieve 7 to 8 complete mission cycles per Mars day. [ABSTRACT FROM AUTHOR]

In recent years, the remarkable progress in exoplanet science has unveiled a stunning diversity of exoplanets. Technological advancements have significantly expanded our capacity to investigate planetary objects beyond our solar system, which exhibit diverse properties including their sizes, compositions, and orbital characteristics. Although understanding the complex chemical-climate responses occurring in planetary-body atmospheres plays a pivotal role in determining potential habitability, much work is still required in this field. To interpret forthcoming spectral measurements obtained from exoplanetary atmospheres, there is a concerted effort to model these worlds. However, model parameters are often inferred from limited data sets. Consequently, laboratory-based atmospheric simulations are necessary for understanding complex gas-phase interactions and to complement atmospheric models through experimental examination of input parameters. In this paper we introduce the new Berlin Atmospheric Simulation Experimental chamber (BASE), present its current capabilities and provide an overview of experimental scenarios with applications ranging from planetary objects in the solar system (e.g. Mars, Venus, Titan) to Earth-like exoplanets. • Exoplanetary science demands experimental validation of atmospheric processes. • Laboratory simulations are fine tuned to a specific set of atmospheric parameters. • UV photons and electrons initiate important chemical reactions in the gas phase. • BASE simulates pressure, composition and radiation levels of planetary atmospheres. • BASE will provide important laboratory data to improve current atmospheric models. [ABSTRACT FROM AUTHOR]

Human exploration of Mars and beyond will demand unprecedented levels of self-sufficiency due to the exceedingly far distances from Earth and lengthy mission durations. As such, astronaut crews must be able to function in a progressively autonomous manner, as they will be unable to rely on timely logistical resupply and/or operational guidance from ground-based mission support teams. One means of increasing autonomy is by incorporating intelligent (i.e., smart) technologies into the mission architecture and habitat design. As plans for sending humans to deep space continue to develop and mature, identifying current and emerging smart technologies and which have the potential to enable these missions or maintain habitability for the crew can be used to explore design options prioritized for autonomous operations. This paper surveys select commercial-off-the-shelf (COTS) technologies currently available and marketed terrestrially for use in integrated smart homes and aligns them for consideration by their functional relevance toward potential Smart Habitat (SmartHab) applications in deep-space. • Self-sufficiency becomes increasingly necessary as distance from Earth prohibits timely communication and resupply capability. • Habitability affects a crew's physical, physiological, psychological, and social health, as well as human performance. • Self-sufficiency and habitability for deep space missions may benefit from utilizing 'smart systems'to enable autonomy. • Over 90 smart technologies were reviewed resulting in 97 space-relevant applications for deep space missions. [ABSTRACT FROM AUTHOR]

Satellites engaged in observation missions often face severe pointing constraints, which can limit the ability of their solar panels optimize the generation of electrical power during payload operations. Some approaches to maximize power generation involve either using orientable solar panels, or performing orbital maneuvers to point fixed or deployable (non-orientable) solar panels towards the Sun, which can compromise the payload's ability to meet its pointing requirements. In this study, an alternative approach that enables high-consuming remote sensing payloads to operate for extended periods without requiring orientable solar panels is presented. The satellite under consideration is equipped with fixed or deployable (non-orientable) solar panels mounted on up to three different perpendicular faces. The on-board payload is required to remain within a fixed angle from a target direction, which can be either nadir or a specific geographical position. To derive an optimal tracking law, the projected solar array area at each instant is maximized, resulting in optimal electrical power generation. This methodology allows the satellite to maintain its pointing constraints and efficiently operate high-consuming payloads for longer periods. The proposed tracking law is validated using an actual mission scenario, comparing the power generation achieved by the proposed method against other techniques. The approach presented in this paper shows improved performance, enabling satellites equipped with non-orientable solar panels to fulfill their observation requirements while optimizing electrical power generation. This work offers significant benefits for satellite operators, reducing the need for costly orientable solar panels and enhancing the overall efficiency of satellite missions. • Guidance law for optimal power generation while complying with payload pointing constraints. • Simple and easy formulation, suitable for real-time systems. • Methodology validated and compared against other attitude strategies. • Integration and simulation of the guidance law into a control scheme. [ABSTRACT FROM AUTHOR]

The assessment of casualty risks associated with re-entry necessitates a comprehensive analysis of trajectories and the examination of pertinent safety-related quantities such as the ground impact area and ground reaching velocity. In practical scenarios, the presence of uncertainty in input conditions introduces variability in safety-related quantities. Consequently, employing stochastic re-entry trajectory analysis becomes crucial in overcoming the limitations of conventional deterministic analyses. Conducting sensitivity assessments during the break-up phase is imperative to gain more insights into how to manage the variability of safety-related measures. Therefore, this paper conducted a surrogate-based global sensitivity analysis and employed explainability machine learning techniques to unveil the complexities of the relationship between input uncertainty conditions and three key measures: ground-reaching velocity, falling range, and falling time, with the object of interest being the Apollo-type capsule. A three-step polynomial chaos expansion-based strategy was devised to efficiently approximate the discontinuous relationships. The results show that the relationship is characterized by severe discontinuity that separates two modes: low- and high-ground reaching velocity caused by the presence of two distinct trim points, with precautionary measures that should be taken to prevent the occurrence of the latter. From this set of procedures, three key input conditions that significantly affect the safety-related measures were identified, namely, the altitude, pitch rate, and path angle of the capsule during the breakup. Subsequently, explainability techniques were utilized to give suggestions on how to control the input variability and avoid the high ground reaching velocity mode, aiming to achieve more dependable predictions for the three safety parameters mentioned earlier. • Performed surrogate-based sensitivity analysis for stochastic re-entry trajectory. • Utilized explainability machine learning techniques to decipher relationships. • Identified key input conditions influencing safety measures for Apollo-type capsules. • Contributed insights to minimize hazardous re-entry modes and mitigate risks. • Uncovered complex discontinuity in response surface with explainable surrogate models. [ABSTRACT FROM AUTHOR]

With the aim of enhance the designability of burning surface, star and wagon wheel fuel types are applied extensively in hybrid rocket motors. However, mechanism of nozzle ablation in hybrid rocket motors with these complex fuel types is unclear, and this increases the difficulty of predicting ablation laws and thrust. This paper intends to investigate spatial distribution characteristics of nozzle ablation in the hybrid rocket motor with various fuel types. According to the principle that the grain port area and grain length are equal, tube, star, single-port wagon wheel, and multi-port wagon wheel grains are designed. Numerical models for combustion flow and thermochemical erosion are established, and three-dimensional numerical simulations are conducted. A firing test for nozzle erosion is conducted, and hydroxyl-terminated polybutadiene and 95 % hydrogen peroxide are employed. Errors of throat erosion rate and combustion chamber pressure among simulation results and experimental data are 4 % and 4.25 %, respectively. Calculation results demonstrate that the circumferential distributions of nozzle wall temperature and erosion rate show an obvious correlation with fuel types. The maximum throat erosion rates for star, single-port wagon wheel, and multi-port wagon wheel grains occur at the circumferential position of 0° (star slot, spoke slot, and midpoint of outer edge of wheel hole), while the minimum throat erosion rates locate at the circumferential position of 36° (star root, spoke root, and midline of adjacent wheel holes). The circumferential non-uniformity of throat erosion rate of multi-port wagon wheel grain is the highest. When the grain port area and oxidizer mass flow rate are the same, the throat erosion rate of multi-port wagon wheel grain is the highest, followed by the star grain, and the throat erosion rate of single-port wagon wheel grain is the lowest. • Effect of complex fuel types on spatial characteristics of nozzle erosion is studied. • Firing test of carbon-based nozzle erosion is performed to verify numerical models. • Three-dimensional simulations on erosion in hybrid rocket motors are conducted. • Simulation results agree well with test data. • Ablation mechanism of star, single-port, and multi-port wagon wheel grain is revealed. [ABSTRACT FROM AUTHOR]

This paper proposes an integrated architecture for navigation and attitude control of low-cost suborbital launch vehicles, propelled by a solid motor. Single-nozzle, two Degrees-of-Freedom (DoF) Thrust Vector Control (TVC) actuation is adopted. For architecture design purposes, a non-linear, unstable, 6 DoF model for the generic thrust-vector-controlled launcher dynamics and kinematics is deduced, and a linear state-space representation is proposed. The navigation system provides full-state estimates resorting to novel complementary kinematic filters, whose design allows to establish an explicit relation with steady-state Kalman filtering. A globally stable estimation solution is obtained, apart from the singularities arising from the use of Euler angles. The attitude control law is derived from the Linear Quadratic Regulator (LQR) using the state-space models for each linearization point of the reference trajectory, with an integral action (LQI) added to improve robustness and to provide null steady state attitude tracking error. A correction method is proposed to allow for pitch and yaw control in the presence of spinning motion, precluding the need of a supplementary roll control system. The control system is implemented through gain scheduling, resorting to an altitude-based linear parametric varying method. The architecture is implemented in a realistic simulation environment, composed by the 6 DoF non-linear model, the Navigation and Control solutions, and the environmental disturbances, to assess its performance through Monte Carlo simulations. The navigation system is able to provide accurate estimates of the state vector, while the control system satisfies attitude tracking performance and robustness to both external disturbances and model parametric uncertainties. • An integrated architecture for navigation and attitude control of suborbital launch vehicles is proposed. • An original time-varying state-space representation is deduced. • Novel complementary kinematic filters are designed to provide accurate full-state estimation. • LQI-based control allows for accurate pitch and yaw reference tracking via thrust vectoring. • The architecture performance is assessed through extensive Monte Carlo simulations. [ABSTRACT FROM AUTHOR]

The accumulation of space debris is an ever-increasing problem. Systems that can reliably de-orbit (i.e., decrease the time to re-entry for) an inactive resident space object (RSO) are highly sought after. In this paper, we investigate the feasibility of utilizing the forces owing to the interaction of Earth's magnetic field with a permanent magnet that has been attached to an RSO in low Earth orbit. We consider an actively controlled permanent magnet that is optimally oriented to remove energy from the system, and the most favorable magnet-to-RSO mass ratio. We show that, even under these best-case assumptions, the interaction between a permanent magnet and Earth's magnetic field is not a viable means of de-orbiting an RSO. [Display omitted] • The interaction between a permanent magnet and Earth's magnetic field can be used to remove energy from an object in orbit. • The orientation of the permanent magnet can be controlled to optimize the energy loss. • This optimal orbital energy loss is not sufficient to de-orbit the object in a reasonable amount of time. • Interaction between a permanent magnet and Earth's magnetic field is therefore not a viable means for de-orbiting space debris. [ABSTRACT FROM AUTHOR]

A significant challenge for interplanetary small satellites is fully propulsive orbit insertion due to limited total velocity change capabilities. At destinations with significant atmospheres, this challenge can be circumvented via aerocapture, a technique that uses a single atmospheric pass to convert a hyperbolic approach trajectory into a captured elliptical orbit. This paper investigated small satellite aerocapture at Mars with the morphable entry system, a deployable entry vehicle concept that uses shape morphing for trajectory control. A Monte Carlo dispersion analysis was conducted to assess the vehicle's robustness with respect to uncertainties in entry state, atmospheric density, and vehicle aerodynamics. Under the combined effects of these uncertainties, the morphable entry system achieved a 100% aerocapture success rate. Furthermore, the total velocity change required to place the vehicle on the target orbit for 99% of the simulated trajectories was 166.5 m/s, which is within the current capabilities of small satellite propulsion systems. The results suggest that the morphable entry system is a viable option for smallsat aerocapture at Mars. • Morphable entry system (MES) uses shape-morphing for aerocapture trajectory control. • MES achieved 100% aerocapture success rate at Mars. • MES may be a viable option for small satellite aerocapture at Mars. [ABSTRACT FROM AUTHOR]

Lunar regolith on the surface of the Moon has gained great attention as a resource for many proposed lunar exploration projects. Not only is it serving as the ground support for construction projects on the Moon, but it is also the most accessible construction material for building the human habitat and a key component in in-situ resource utilization efforts. Therefore, gaining a fundamental understanding of the characteristics of lunar regolith is a critical step in the success of lunar exploration projects. However, it is challenging and expensive to obtain samples of lunar regolith directly, or conduct physical experiments on lunar regolith samples in situ. Hence, investigative alternatives, such as laboratory experiments and numerical simulation tools are very important means of investigating the properties of lunar regolith. This paper presents a comprehensive study on lunar highlands simulant (LHS-1), using both laboratory experimentation and discrete element method (DEM) simulations. In the laboratory experiment, triaxial tests were conducted on samples of LHS-1 to explore the stress–strain behaviors of the lunar regolith simulant. Correspondingly, 3D DEM simulations of the triaxial tests were conducted to generate the initial model parameters and the Taguchi method was performed to obtain the optimal calibrated input parameters of the DEM model. With the calibrated parameters, 3D DEM model of the cone penetration into LHS-1 was then built to study the responses of the regolith simulant during the penetration process. The effects of porosity and vertical stress on the tip resistance and sleeve friction profiles were investigated for both terrestrial and lunar environments to study the correlation between the in-situ state of the regolith simulants with the force reactions during the penetration. • Geotechnical properties of LHS-1 simulants were characterized using a set of laboratory tests. • Triaxial tests were used to calibrate contact parameters of DEM models using Taguchi method. • Cone penetration simulations were performed using discrete element modeling. • Correlation between penetration responses and in-situ state of regolith simulants was evaluated. [ABSTRACT FROM AUTHOR]

The recent proliferation of resident space objects (RSOs) in low Earth orbit (LEO) threatens the sustainability of space as a resource and requires persistent monitoring to avoid collisions involving valuable space assets. State-of-the-art ground-based space surveillance techniques, due to their susceptibility to atmosphere, weather, and lighting conditions, tend to focus on RSOs with characteristic length greater than 10 cm or 1 dm. Consequently, millions of smaller LEO RSOs remain untracked by ground-based methods, which reduces overall space situational awareness. Onboard satellite sensors offer a space-based method for tracking RSOs. Prior research has investigated the feasibility of using commercial star trackers (CSTs) — optical sensors prevalent on most active spacecraft — to observe, detect, and estimate the position and velocity of RSOs larger than 10 cm. In a recent effort, we expanded on these feasibility studies by assessing the capabilities of CSTs to detect debris particles smaller than 10 cm in characteristic length. We modeled the particles as Lambertian spheres with zero phase angle and ten percent reflectivity and found that typical CSTs can detect properly illuminated debris of characteristic length between 1 cm and 10 cm even at distances of tens of kilometers. More sensitive CSTs can characterize decimeter-scale RSOs hundreds of kilometers away; alternatively, they can track centimeter-scale and smaller RSOs at closer distances. In this paper, we summarize these key results and extend our previous study by relaxing its zero-phase-angle assumption and characterizing the effect of Sun-RSO-CST phase angle on debris detection range. We identify a number of representative CSTs with publicly available optical characteristics and consider the effects of properties such as pixel pitch, focal length, aperture diameter, and field of view (FOV) on the capability of each CST to detect debris at a given distance and relative velocity (in the form of streak speed). We find that, for debris particles modeled as diffuse Lambertian spheres, Sun-RSO-CST phase angles as high as 57° result in no more than 20% reduction to the useful RSO-CST detection range. In addition, we find that pixel pitch and focal length, rather than aperture diameter, tend to determine the capability of a given CST to resolve two distinct RSOs. Furthermore, streak speed may serve as a stronger limiting factor for detection of smaller debris particles than for larger ones. Despite these limitations, the overall results indicate that CSTs have the potential to substantially enhance space-based debris detection capability. • Prior state-of-the-art in space domain awareness focuses on objects exceeding 10 cm. • Commercial star trackers — CSTs — can detect smaller objects like sub-decimeter debris. • Resident space objects — RSOs — are modeled as diffuse Lambertian spheres. • CST sensitivity and RSO size and reflectivity determine maximum range of detection. • Sun-RSO-CST phase angle, CST resolution, and streak speed limit detection potential. [ABSTRACT FROM AUTHOR]

Shape Memory Alloys (SMA) are materials used to design smart structures with intrinsic functional properties and improved efficiency. This is a key aspect of aerospace industry and makes SMA good candidates in this field. One of the most widespread SMA is the equiatomic NiTi alloy which, however, has the strong limitation of poor machinability, so only simple shapes can be obtained. Additive Manufacturing processes allow to overcome this limit and to design complex shapes. Compared to other metallic materials, the optimization of the process for NiTi alloy is complicated because, beside mechanical properties and presence of defects, considerable attention needs to be dedicated to the material functionality. The high temperatures involved in the additive process significantly affect the material properties due to possible evaporation of Ni and formation of precipitates that enable a shift of the phase transformation temperatures. This paper is focused on the optimization of the process parameters of the NiTi alloy printed through the Laser Powder Bed Fusion (L-PBF) to ensure optimal pseudo-elastic behaviour, which is essential for the design of structural dampers. This was accomplished starting from simple structures and then designing a damper that couples the pseudoelasticity of NiTi with load support capacity. The L-PBF is a powder-bed technique that selectively melts layers of micrometric metal powder. A pseudoelastic NiTi powder with 50.8 at. % of Ni content was selected and characterized through scanning electron microscope (SEM) and observations connected to an Energy Dispersive X-ray Spectroscopy (EDX) probe. After that, some cubic samples were manufactured, with the dimension of 3 × 3 × 15 mm3. A set of different laser powers and scanning speeds were used to find the set of process parameters that optimize the functional properties of the printed parts. Near fully dense specimens with density higher than 99.5 % were selected for further investigations. Differential scanning calorimetry (DSC) and mechanical tests were performed on as-built and heat-treated samples. Quasi-static mechanical tests were accomplished in compression mode, at different strains, up to 8 %. It was observed that the residual strain for cyclic loading at 4 % is lower than 1 %, so good recovery of the deformation was shown. Moreover, numerical analyses that mimic the pseudoelastic behaviour in compression tests were implemented. Finally, the best set of parameters was selected on the basis of the material's ability to recover deformations and its loss factor. • Optimization of process parameters for functional behaviour of SMA. • Design of damping structure using pseudoelastic effect of Nitinol. • Internal defects observation of manufactured specimens. • Compression tests on as built and thermal treated specimens with strain up to 8 %. • Numerical simulation of compression tests with experimental correlation. [ABSTRACT FROM AUTHOR]

The agility of satellites equipped with large and flexible structures, such as solar panels or antennas, poses a significant challenge due to the influence of elastic dynamics on attitude motion. This results in reduced pointing accuracy and dimensional stability of the payload. To address this issue, an Active Collocated Attitude Control (ACAC) system has been developed, utilizing actuators for both the elastic dynamics (distributed piezoelectric devices) and attitude control (ON-OFF thrusters). In the initial phase of the project, the algorithm underwent extensive verification using a high-fidelity numerical model of a spacecraft with two very-low frequency flexible solar panels and a payload consisting of two long booms and a transversal antenna's reflector. The paper focuses on the second phase of the project, which revolves around the experimental verification of the control architecture concept. To achieve this, a scaling procedure employing the dynamic analogy approach was implemented to determine the design parameters of a free-floating platform equipped with elastic appendages. This ensured that the resulting elastic multibody system displayed similar dynamic behavior to the full-scale satellite. Based on the scaling parameters, a test rig was manufactured, and actuators and sensors were optimally positioned. The performance of the ACAC system was then compared to that of traditional benchmark controllers. The results show that the ACAC system demonstrates superior capabilities in performing fast slew maneuvers without developing unstable interactions with the elastic dynamics. • Agile maneuvers of large satellites could excite rigid-flexible interactions. • A robust control of attitude and elastic dynamics is implemented to improve agility. • Performance obtained with numerical models is verified with scaled-down test-rig. • Simplified dynamic equations are developed to determine scaling conditions. • Experimental results of the robust control are consistent with simulation. [ABSTRACT FROM AUTHOR]

The new generation of large-thrust rockets commonly employ cryogenic liquids such as liquid hydrogen (LH2), liquid oxygen (LOX), and methane as propellants. Conventional analytical methods established under subcritical (specifically, low-pressure) conditions are inadequate for addressing the atomization issues under trans -critical and supercritical conditions. Therefore, conducting experimental research on atomization of cryogenic fluids under supercritical conditions holds significant practical value. This paper employs liquid nitrogen (LN2) as the working fluid and focuses on conducting injection atomization experiments on LN2 under supercritical pressure conditions. The research involves the design and construction of an experimental system for supercritical cryogenic fluid injection, with the aim of measuring and studying the spray characteristics of centrifugal nozzles under various conditions. Three major categories of experiments were designed for low-temperature LN2 injection: subcritical nitrogen into subcritical environment (BB), subcritical nitrogen into supercritical environment (BP), and supercritical nitrogen into supercritical environment (PP). The experimental results ultimately reveal that for jets in BB, both the spray cone angle and the breakup length increase with higher injection pressure drop. In the case of BP, the spray cone angle decreases while the penetration distance increases with higher pressure drop. As for the jets in PP, the spray cone angle increases and the penetration distance decreases with higher injection pressure drop. Across different geometric parameters of the centrifugal nozzle, it was observed that with an increase in the geometric parameter, the rate of decrease in the spray cone angle with respect to pressure drop amplifies, and the penetration distance exhibits an increasing trend as the geometric parameter grows. • An experimental system designed for the atomization of cryogenic fluids, such as liquid nitrogen (LN2), under supercritical pressure conditions is designed. • The research provides an in-depth analysis of how different environmental and operational conditions (subcritical/subcritical, subcritical/supercritical, and supercritical/supercritical environments) affect key parameters of LN2 spray, including spray cone angle, penetration distance, and breakup length. • The characteristic parameters of trans -critical jets are significantly influenced by centrifugal nozzles with different geometric characteristic coefficients. [ABSTRACT FROM AUTHOR]

The main advantages of ceramic materials are their excellent thermal, chemical, and mechanical properties. They also have a much lower density than metals, making them ideal for lightweight construction, which can also be interesting for space components. However, ceramics have not been widely used due to their low ductility and the challenges of mechanical processing using conventional methods like milling and turning. Additionally, integrating different functions into a single component has been limited. The introduction of additive manufacturing (AM) technologies has revolutionized this field. It is now possible to manufacture highly complex ceramic components with unprecedented functionality. By hybridizing different manufacturing technologies and materials, additional functions such as electrical contacts, sensors, and actuators can be directly integrated, leading to improved manufacturing costs and component properties. Three different demonstrator examples are presented in this paper to give a first impression of what will be possible in the field of space applications in the future. These examples are a ceramic igniter, a ceramic reactor for thermal decomposition of H 2 O 2 , and a ceramic aerospike engine. • Additive manufacturing of ceramic components allows geometrical functionalization. • Adding of further materials to AM components allows further functionalization. • Sequential & simultaneous manufacturing available for AM of multi-material component. • Hybridization of AM technologies allows optimization of component's quality and costs. [ABSTRACT FROM AUTHOR]

The objective of this paper is to analyze the performance of two possible architectures of tethered satellite systems, used as a platform for a distributed radar sounder. The first architecture consists in a cross-track oriented tethered satellite system, controlled and stabilized by exploiting the aerodynamic forces generated by the interaction with the rarefied atmosphere in Low Earth Orbit. The second architecture involves a tethered satellite system controlled through gyroscopic stabilization, obtained by spinning the system about an axis contained in the orbital plane. After a brief survey of radar sounding techniques, the methodology is introduced for describing the geometry of the systems and their characteristics, the performance of the two architectures are then compared with each other and with the current state of the art. By analyzing the modeled nominal behavior, it is shown that the two proposed architectures can achieve continuous or multiple observations, respectively, at maximum cross-track resolution, during one orbit, minimizing clutter noise. This is a considerable improvement of performance versus a formation flight architecture which can typically achieve only up to four observations per orbit. Finally, the advantages and disadvantages of each architecture are studied, and their possible mission scenarios are discussed. • A Tethered Satellite System is proposed, flying with tether in cross-track direction. • The tether tension is generated by aerodynamic lift forces on plates at its extrema. • Antenna elements distributed along the tether enable radar sounding applications. • Top radar sounding performances are achieved everywhere along the orbit. • An alternative gyroscopic tether system is studied for planets with no atmosphere. [ABSTRACT FROM AUTHOR]

In recent decades, the interest of the scientific community in plant space biology increased based on the certainty that cultivating higher plants in space is imperative for prolonged human missions. The cultivation of plants in space assumes great relevance moved by the consideration that there is a pivotal need to improve resource regeneration and onboard production of plant-based food, thus reducing the supply of resources from Earth. Therefore, the hereby paper provides an overview on the evolution of the interest in plant space biology from 2000 to 2023, basing the search on specific keyword combinations. The study encompasses (i) a proper quantitative search on the Scopus database; (ii) keyword frequencies and clustering analysis, (iii) a keyword network and time-gradient analysis and (iv) an elaboration of knowledge gaps comparing our results and recent reviews. Results highlighted an increase in publications in the last 10 years and a specific time-gradient related shifting on subtopics such as in situ resource utilization (ISRU) application and food production in a space mission context. Currently, knowledge gaps were identified in four research areas related to water dynamics in extra-terrestrial environments, plant physiology and biochemistry in response to growth on regolith, interactions with fertilizers, and plant/soil microbiome. Identifying these gaps will allow to canalize future efforts of space research where needed, facing the unsolved questions. Our analysis pointed out that space plant biology turned out to be oriented towards short and long-term space mission applications for improving resource regeneration and onboard/extraterrestrial-grounded production of plant-based food, thereby diminishing reliance on Earth for resupply. • The scientific community's interest in plant space biology increased in the perspective of prolonged human missions. • The bibliometric analysis highlighted a general increase in publications in the last ten years. • Publications shifted their focus on subtopics such as ISRU application and food production in a space mission context. • Main identified knowledge gaps are water dynamics in space, plant growth on mars-like soils, and plant/soil microbiome. [ABSTRACT FROM AUTHOR]

Fixed-wing aircraft with potential for long-duration flight and efficient manoeuvrability is expected to be the next frontier of Mars surface exploration. However, the feasibility of such an aircraft demands for wing airfoil suitable for low Reynolds flight conditions. In this framework, the paper deals with the computational study of two wing sections specifically designed for Mars exploration aircraft. Two-dimensional steady Reynolds-averaged Navier–Stokes simulations, using the computational fluid dynamics tool SU2 with the γ − R e θ transition model and the XFoil panel flow solver, are performed to address airfoils' aerodynamics. Computational tools are validated by simulating flow past the Eppler 387 airfoil at R e ∞ = 60 × 1 0 3 , and comparing the results with experimental data collected in wind tunnel experiments. Aerodynamic performances of optimal wing sections are investigated at R e ∞ = 3. 4 × 1 0 4 by considering pressure coefficients and skin friction coefficients. The application of a literature-based correlation, specifically tailored to identify transition and reattachment locations, is extended to separation point detection. Computational Fluid Dynamics analysis conducted on the studied airfoils revealed that separation occurs within the laminar regime. Additionally, examination of turbulent shear stresses highlighted the role of airfoil curvature in counterbalancing the suction defect produced by the laminar separation bubble. This curvature-induced effect was found to play a crucial role in optimizing airfoil efficiency. • Aerodynamics issues of flying in Martian atmosphere. • Computation of flows past airfoils for future Mars exploration aircraft. • Detection of transition and separation at low-Reynolds number. [ABSTRACT FROM AUTHOR]

Over the past fifty years, the form and function of space habitat modules have remained largely unchanged. While cylindrical modules are relatively simple to transport and deploy, their limited size poses a challenge to their efficiency and usability. In contrast, large space habitat construction projects are costly and present logistical difficulties. The development of polyhedral modules could bridge the gap between these two approaches, allowing for modular, polyhedral units to be assembled and linked together to create larger habitats more efficiently and with less risk. However, identifying the most suitable polyhedron for use in space remains a critical question. Previous research has explored the construction of polyhedral modules, but little rigor has been applied to identifying the most optimal form. This is a crucial issue because the first module to be used will likely set the standard for subsequent modules and any disadvantages present at that time will be perpetuated. Therefore, it is essential to identify the optimal form prior to constructing prototypes. This research paper employs multi-criteria decision analysis and sensitivity analysis to compare various candidate polyhedra across several evaluation metrics, including number of faces, volume to surface area ratio, and joint stress, among several other quantitative and qualitative metrics. The results demonstrate that the Rhombic Dodecahedron is a particularly suitable candidate compared to other forms analyzed. Thus, the Rhombic Dodecahedron should be considered the standard polyhedral form for future research involving the development of polyhedral modules. • Rhombic Dodecahedron identified as optimal polyhedron for space habitat design. • Polyhedral modules surpass cylindrical counterparts in multiple criteria. • Novel multi-criteria analysis method allows cross-metric comparison. [ABSTRACT FROM AUTHOR]

This paper concentrates on robust attitude control issues for spacecraft formation in gravitational-wave detection missions. Notably, with the large-scale communication delays brought by the million kilometers-scale inter-satellite distance, the exponential convergence of the closed-loop delayed system is firstly achieved via a new delay-dependent nominal attitude coordinated control scheme. Then, a finite-time disturbance observer is deployed to reconstruct the unknown lumped disturbance from internal uncertainties and the external environment. Meanwhile, an integral sliding manifold is embedded in the delay-dependent attitude controller for superior systems' robustness. Finally, two simulation examples illustrate the feasibility and effectiveness of the proposed attitude control strategy for in-orbit gravitational-wave detection spacecraft systems. • Spacecraft formation attitude control in gravitational-wave detection is studied. • Large-scale communication delays (LCD) among formation members are firstly modeled. • An attitude coordinated controller is put forward to handle the LCD. • An integral sliding manifold with disturbance observer is embedded for superior robustness. [ABSTRACT FROM AUTHOR]

Aotearoa New Zealand's emerging New Space economy provides an opportunity for key actors to focus on space and sustainability issues beyond space debris. The conflict between competing definitions and paradigms of sustainability highlights the importance of diverse values, assumptions, and drivers of change that shape the normative understanding of space sustainability issues. This paper recognises that Indigenous knowledges and practices are in parallel with systems-thinking and transdisciplinary approaches to space and sustainability. The aim of this paper is to describe how current actions can have long term impacts on using and accessing space commercially, scientifically, and culturally. • Space sustainability requires a transdisciplinary approach to explore the definitions and values at various scales. • Space sustainability in New Zealand involves addressing issues beyond space debris. • Space sustainability requires addressing businesses in New Zealand, where sustainability frameworks are poorly implemented. • Space sustainability requires policy interventions beyond compliance with environmental legislation. [ABSTRACT FROM AUTHOR]

This paper aims to describe and assess the dispute settlement mechanisms available to States in the context of space and telecommunications activities. Disputes either between private actors or between private actors and States are beyond the scope of this research. The choice to examine mechanisms and procedures for the settlement of international disputes in these specific areas stems from an elective affinity between these sectors, given the existing interaction between them. At a first level, the paper highlights the importance of mechanisms to prevent the creation of international disputes, both in the field of space law (mainly through Art. IX of the Outer Space Treaty) and in the field of international telecommunications (Arts. 45 para. 3, 48 of the ITU Constitution). Further, the general scheme of peaceful settlement of international disputes, as enshrined in the UN Charter not only applies to the disputes in question but is further specified in the context of space and telecommunications law (Liability Convention, Art. 56 of the ITU Constitution), with a clear orientation towards the use of diplomatic rather than judicial means of resolution. Last but not least, if the settlement of disputes through diplomatic means fails, the preferred judicial mechanism is that of arbitration, as it is clearly demonstrated by the Claims Commission of Art. XIV LIAB, the PCA "Optional Rules for Arbitration" of space disputes, the mechanism of Art. 41 of the ITU Convention as well as the relevant "Optional Protocol". The specific options for dispute settlement show that, in view of the common interest of States in "international goods" of a technical nature, the actors involved are mainly interested in the non-occurrence of disputes rather than in their resolution through judicial means. And when they do so, they choose to resort to flexible resolution mechanisms (arbitration) rather than to ordinary international courts. • There is an elective affinity between telecommunications and space activities. • States are mainly interested in preventing disputes than to resolve those arising. • States opt for diplomatic resolution mechanisms instead of jurisdictional ones. • Ιf diplomatic settlement fails, the preferred judicial mechanism is arbitration. [ABSTRACT FROM AUTHOR]

The Surface Water & Ocean Topography (SWOT) mission is a joint NASA/CNES mission launched on December 16th, 2022 dedicated to oceanography and continental hydrology. Once the scientific mission is over, SWOT will have to perform a controlled re-entry to comply with the French Space Operation Act (FSOA). Indeed, the launch authorization is subject to the respect of the FSOA, fully applicable since 2020. The main requirement applying to the end-of-life of the satellite consists in limiting the casualty area and then the probability of having a victim on ground under the threshold of 10−4 in case of a natural re-entry. With a mass over two tons and a casualty area of about 47 m2, SWOT cannot comply with this requirement. Thus, the platform has been designed with a specific propulsion system to carry out end-of-life maneuvers leading to an atmospheric re-entry and impact in the South Pacific Ocean Uninhabited Area (SPOUA). CNES is in charge of defining the re-entry maneuvers strategy, performing end-of-life operations and evaluating the risk of casualties in case of failure during nominal strategy. This paper addresses these topics. The computation of the ten maneuvers used to decrease the perigee altitude and finally target the SPOUA is performed using the DOORS tool. Before describing the maneuvers strategy, this paper presents an overview of the current functionalities of this tool, which has been already used during the five ATV missions and ASTRA-1K de-orbit twenty years ago. DOORS is mainly used for mission analysis purposes, but for end-of-life operations it can be activated by the operational Flight Dynamics System (FDS). Finally, the casualty risk estimation relies on the use of two other CNES tools, DEBRISK and ELECTRA, whose methods are presented. The list of surviving fragments has been computed using the new DEBRISK-V3 methodology in case of controlled and uncontrolled re-entry. The risk of human casualty is evaluated using ELECTRA and its Monte Carlo simulations. This tool generates numerous trajectories of the satellite and its fragments with various dispersed variables. The impact locations compared to an up-to-date grid of the World population lead to the probability of victims. The complete risk of SWOT controlled re-entry encompasses both the risk related to a failure during the last re-entry maneuver and the risk associated with the natural re-entry in case a failure occurs before this last burn (during the mission or during the first re-entry maneuvers). Both of them are weighted by their probability of occurrence. • SWOT controlled re-entry strategy to fulfill French regulation. • Robustness test of the nominal strategy made of 10 maneuvers. • Latest aerothermodynamic models used to build the list of surviving fragments. • Probability of victims considers contingencies during the whole re-entry operation. • Combination of several CNES softwares: DOORS, PSIMU, DEBRISK, ELECTRA. [ABSTRACT FROM AUTHOR]

The National Aeronautics and Space Administration (NASA) has been involved in the development and maturation of metal additive manufacturing (AM) for space applications since the late 2000's. Several efforts have focused on the understanding of AM processes through material characterization and testing, standards development, component fabrication, and infusion into propulsion development and flight applications. NASA matured commonly used aerospace alloys from various alloy families (Nickel, Copper, Stainless and Steel, Aluminum, and Titanium-based) through detailed AM process and heat treatment characterization, in addition to mechanical and thermophysical testing. While these alloys are actively used in many propulsion applications, there is a need for ongoing AM optimized alloys using integrated computational materials engineering (ICME) and process development for high performance applications. The applications targeted are liquid rocket engines; advanced propulsion systems; and in-space propulsion with high heat fluxes, high pressure, and/or that use propellants that can degrade alloys (e.g., hydrogen). This paper highlights the characterization and physical properties of the more common AM alloys using laser powder bed fusion (L-PBF) and laser powder directed energy deposition (LP-DED) processes. Additionally, this paper discusses some of the ongoing novel alloy development and maturation using AM for use in these harsh environments, such as GRCop-42, GRCop-84, NASA HR-1, GRX-810, and C-103. The results from these processes demonstrated that AM could enable rapid development, and that optimized alloys could be developed using ICME, yielding higher performances. These alloys have undergone modeling, fundamental metallurgical evaluations, heat treatment studies, detailed microstructure characterization, and mechanical testing campaigns. This, combined with direct application-specific component fabrication and hot-fire testing, enabled the increase of the Technology Readiness Level (TRL) through high duty-cycle testing. A background and overview of these novel AM-enabled alloys and AM processing developments including metallurgical and mechanical property studies is presented here. The latest advancement in the parallel component development and hot-fire testing and future developments for these alloys is also discussed. • Additive manufacturing (AM) has enabled development of novel alloys for extreme propulsion environments. • NASA has matured alloys including GRCop-42, GRCop-84, NASA HR-1, GRX-810, C-103. • Characterization, material properties, and hot-fire test application data has been made available. [ABSTRACT FROM AUTHOR]

The purpose of this paper is to study the development of space-age aesthetics culture and arts in seven decades since 1950s. The first known use of "space-age" was in 1946, and was defined as "of, relating to, or befitting the age of space exploration." A more popular definition for the "space age" is "The space age is the time period in human history related to the space race, space exploration, and space technology. Most people say this started when Sputnik 1 was launched in 1957, and continues to today. This has influenced culture such as in movies, television, music, art, and architecture." Culture, in the anthropology, is the patterns of behaviour and thinking that people living in social groups learn, create, and share. Culture distinguishes one human group from others. It also distinguishes humans from other animals. A people's culture includes their beliefs, rules of behaviour, language, rituals, art, technology, styles of dress, ways of producing and cooking food, religion, and political and economic systems. Therefore, the space-age art is one of the major elements of the space-age culture of human beings. On the other hand, the space exploration has significant inspiration and influence on the aesthetics art. In particular, the following aspects shall be studied in this paper: architecture and exhibition craftworks. • The space-age architecture and exhibition craftworks were investigated. • Five space-age architecture shapes investigated are disc, rocket, tower, spherical and combined. • Three space-age installation craftworks studied are sky waves, embroidery, planet models, etc. • Many space-age architectures are destroyed due to many reasons. • Hopefully the young generations can be inspired to carry forward these precious heritages. [ABSTRACT FROM AUTHOR]

Partially reusable launch vehicle design has raised a great interest in order to minimize the costs of space transportation by recovering and refurbishing the first stage. Several configurations such as toss-back architectures are now operational. These concepts require additional rocket propellant in order to carry the boost-back and landing maneuvers, that induce several losses in terms of payload mass performance for the ascent mission. In order to limit these losses, this paper focuses on the design of two types of architectures to carry out the return-to-launch-site mission. The first, named glide-back, uses the main propulsion system of the first stage to perform a boost-back burn and returns to the landing site by a gliding mode using additional lifting surfaces. The second, named, fly-back, performs the return-to-launch-site mission using several air-breathing engines located in the nose of the first stage combined with lifting surfaces. In this paper, the design of a reusability kit, allowing to provide the first stage with both expendable and reusable capabilities, is investigated. This kit is composed of the lifting surfaces, the nose including the air-breathing propulsive system (for fly-back configuration), landing gears and additional avionics. This paper addresses the design of such reusability kits and presents the mission specifications, the design process relying on Multidisciplinary Design Optimization techniques, the vehicle performance and the optimal trajectories. Different analyses (aero-propulsive, optimal control, operational considerations, etc.) are detailed and trade-offs between the studied configurations are assessed. [Display omitted] • Reusable launch vehicles. • Multidisciplinary design and optimization of reusability kit. • Inclusion of aeronautical technologies in launch vehicle design. [ABSTRACT FROM AUTHOR]

The detection and observation of space debris in Low Earth Orbit is generally carried out through the use of ground based radars and telescopes. These instruments allow for a precise reconstruction of the space debris trajectories, and therefore represent a key asset for planning avoidance maneuvers when threats of collisions are predicted. The recent deployment of mega-constellations, with the consequent increase of the number of satellites, imposes new challenges in terms of simultaneous tracking capability and readiness of the current space situational awareness systems. This adds to the current need to track small and dull objects to further mitigate the probability of triggering cascade collisions. However, ground based observations are limited due to their intrinsic sensibility to atmospheric refraction, their diurnal inoperability and their dependence on meteorological hazards. This paper proposes to study the feasibility and the benefits of a potential deployment of a constellation of CubeSats in Low Earth Orbit, to acquire optical observations of space debris with enhanced accuracy, as part of the ORCA mission: Orbit Refinement for Collision Avoidance. Here, the focus is on the optical design of the payload instrument to be integrated onboard of the orbiting platforms. The study trades-off the current state-of-art of optical detection technologies, by assessing their performance against a set of specific requirements: (a) the minimization of the uncertainty associated to the image resolution; (b) a field of view that maximizes the extent of the monitored area; (c) an optimal exposure time to avoid under or overexposure of the image; (d) minimization of the effects of light diffraction and above all, (e) the maximization of the signal to noise ratio to detect the smallest and dullest objects possible. Several configurations of optical systems are then chosen as suitable for the ORCA constellation, also considering the system design implications of its integration into a CubeSat, such as size requirements. Commercial Off-The-Shelf hardware are explored and performances of the optical system are evaluated through numerical simulations in order to estimate the detectable sizes of space debris while taking into account their potential distances from the sensor. This paper concludes with estimates of the impact of the ORCA mission on space situational awareness for decades to come. • We defined requirements for an optical system for space debris surveillance. • We formulated conditions allowing detection of space debris from optical images. • We explored feasible design of optical systems for a CubeSat size implementation. • We compared our results with available commercial-off-the-shelf optical systems. [ABSTRACT FROM AUTHOR]

In this paper, different compressible correction strategies for k-ω turbulence model are proposed. The dilatation dissipation term, which is significant in compressible flow, is taken into account and modeled in the Favre-averaged k-equation and ω-equation. We propose correction model for the dilation dissipation term based on the gradient Mach number in this paper, the coefficients of which are calibrated from experimental data on compressible turbulent mixed layers, and compare it to the Wilcox, Sarkar, and Heinz correction models. In slope-cavity flow and oblique shock wave/boundary layer interactions, the performance of these corrections is evaluated and compared to experimental data. The new correction has a similar or even better effect in separation prediction, heat flux prediction, and other aspects, and the uncertainty of model form and coefficients can be investigated further. • A general compressible correction model is conducted with the consideration of dilatation and structural compressibility. • The experimental data of compressible turbulent mixed layers is employed to calibrate the model coefficients. • Verification and validation are carried out in slope-cavity flow and incident oblique shock wave/boundary layer interactions. • The performance of the model and its comparison with other correction models are discussed. [ABSTRACT FROM AUTHOR]

This paper treats the problem of positioning and navigation in the absence of GNSS. Given recently raised vulnerabilities for GNSS both on Earth and in space, the work revisits the old problem of the sextant in new light. The combination of stars and planetary horizons was the popular tool for autonomous navigation on-board spacecraft, but given the rise of GNSS receivers, this solution has been largely disregarded. New spacecraft missions beyond Earth have made new progress in visual-based navigation. The work utilises these new methodologies in the scenario that RF-derived positioning is unavailable, achieving a performance below 100 m. An additional important consideration is the development of new navigation infrastructure in LEO and the Moon. Current methods seek to use GNSS and RF-derived sources for orbit determination, however, to be seen as 'redundancy infrastructure' for critical Earth and beyond applications, these signals cannot be the primary and only source of navigation reference. This paper utilises derived performances for visual-based methods and applies them to the ranging service problem. Most user scenarios in maritime, aviation and lunar domains are satisfied. This work illustrates implementation using the simple architecture of a star tracker camera. Known as CROSS, the technology is a new navigation tool in development by the University of Sydney. As a star tracker is common device to many spacecraft platform, it simplifies implementation, given that redundant systems are always not a priority to the manufacturer. • Vision-based approach for orbit determination using only stars and the horizon. • Application of this approach to future LEO PNT mega constellations. • Comprehensive review of requirements by different market segments. • Proposed methods to overcome key challenges including stray light and planetary atmospheres. [ABSTRACT FROM AUTHOR]

The cooperative orbit determination (OD) using inertial inter-spacecraft measurements is of great value for the constellation. Previous studies concluded that the cooperative OD system in the two-body dynamics is observable using the inertial angle measurement. This paper performs an observability analysis for the inertial angle-only cooperative OD system based on the observability matrix (OM), and finds that the previous conclusions do not hold for some special configurations. Eight special configurations are proposed, and for each configuration, the unobservable orbit elements are analytically revealed by analyzing the linear relationship among different columns of the OM. Moreover, a method of improving the observability of the two-spacecraft system is proposed by adding another spacecraft. An extended Kalman filter is employed to numerically illustrate the observability of the proposed special configurations and to show the relationship between the observability and the convergence ratio of estimations. Monte Carlo simulations show that the convergence ratios of the observable orbit elements and state combinations are higher than 97%, while convergence ratios of the unobservable states are lower than 70%. With one non-planar elliptic orbit added, all the elements of the three-spacecraft system are observable, and the estimated errors converge within 6 h. The results found in this paper can provide a reference for the design and management of the constellation. • Observability analysis for the angle-only cooperative OD is performed. • Eight unobservable configurations are revealed using the observability matrix. • A method of improving the observability is proposed by adding another spacecraft. [ABSTRACT FROM AUTHOR]

The use of modular, yet reusable and interchangeable components offers flexibility for system designers and thus leads to a higher degree of reconfigurability with cost effective, adaptable and scalable solutions. The development of a robotic system remains, however, a complex task with increasing demands depending on the targeted capabilities, functionalities and operational constraints of the robot. This requires expert knowledge in mechanical, electrical and software engineering, as well as tools that can collect, process and use relevant data. As a result, the amount of available and managed information is increasing, while its relevance, usability, and knowledge that can be extracted from it are consequently decreasing: a situation described as information overload paradox. The design of a modular system can be based on the use of standardized modules with universal interconnects, but it should be aligned with a strong formalism, e. g. the semantic representation of components, to allow for a better management of the overall complexity. An ontology-based knowledge representation, for instance, which provides domain and application specific knowledge for the design of robotic systems, can be used as a method of storing and sharing data in a uniform, machine-readable and standardized way. In this paper, we introduce the Knowledge-based Open Robot voCabulary as Utility Toolkit (korcut) as a core component to support the ontology-driven development of robotic systems as part of a reference implementation of the Q-Rock development cycle. Q-Rock gathers methods to assist users in designing robot configuration by: (a) automatically exploring robot capabilities based on robot hardware, (b) proposing robot designs to meet the user's needs, (c) refining the proposed robot designs. We use korcut to improve the robot design process in orbital and planetary robotics from the requirements definition to the development phases of complex modular robotic structures (e. g. composition of an unmanned ground vehicle with robotic arms). This is implemented by embedding semantic component descriptions in a state-of-the-art open-source 3D modeling software. Furthermore, the korcut ontology family includes various sub-ontologies to address specific space-related tasks, such as the modeling of a Standard Interconnect (SI) for On Orbit Services (OOS) and Orbital Factory. The ontology design follows criteria that have been derived from a survey conducted as part of this work. This paper also covers an evaluation of its current applications from a methodological, knowledge representation, and software tool perspectives. In the application part, we introduce the use of the ontology to model a multi-functional SI that enables mechanical connections between various (modular) robotic components, as well as transfer of power and data. Finally, we provide a critical analysis of our work and outline its future work. • The development of complex robotic systems in space domain needs expert knowledge. • Increasing information about a robotic system can lead to an information paradox. • The information paradox can be improved by standardized and inferable knowledge. • Ontology is used in a case study to facilitate the robot development workflow. • Ontology is used to provide a metric for evaluating a space interface. [ABSTRACT FROM AUTHOR]

On-Orbit-Servicing missions aim to improve the operative life of orbiting satellites. Among the studied technologies, the most promising are satellites equipped with one or more robotic arms (space robots). One robotic arm might be employed to establish a rigid connection with the target satellite, while the other performs the servicing operations. Nowadays, the control of the capture tool is managed by the computer of the space robot. However, it is possible to distribute the control tasks between the capture tool and the space robot. For this purpose, this paper presents the development of the SMArt Capture Kit (SMACK). SMACK is a standalone capture system equipped with sensors, actuators, and an integrated computer. In this fashion, the capture tool executes the guidance and navigation algorithms, and the control of the actuators; so that the space robot acts as the "positioning" system for the capture tool. The paper provides a description of all the subsystems of SMACK and their experimental validation. The tests of SMACK at the system level, confirm its capabilities to autonomously perform the capture the target interface. • Smart capture tool. • Autonomous tool with integrated sensors, actuators and computer. • Distributed control. • Use of fuzzy logic to develop the model of a custom sensor. • Use of fuzzy logic to create the approach trajectory. [ABSTRACT FROM AUTHOR]

By the time of the 2022 IAC, NASA's Mars2020 Perseverance rover will have spent 18 months on the surface of Mars, during which time the MOXIE experiment (M ars OX ygen I SRU E xperiment) will have produced oxygen at night and in the day during both the annual maximum and minimum atmospheric density periods, as well as at many other times during the year. MOXIE is the first demonstration of the use of indigenous resources (ISRU = In Situ Resource Utilization) on the surface of another planet. This paper will present a summary of what MOXIE has accomplished, how its performance on Mars has changed with time, and plans for the future. The paper will also present results from an optimization study of a human-scale MOXIE-type system capable of providing the oxidizer for a 6–person Mars Ascent Vehicle. As an experiment carried inside the rover, MOXIE had to satisfy many constraints that would not apply to an independent, full-scale system. Other potential oxygen-producing technologies should be compared to the optimized human-scale system results summarized in this paper rather than to a simple linear scaling of the mass, power consumption, and oxygen production rate of MOXIE. • MOXIE (M ars Ox ygen I SRU E xperiment) has successfully demonstrated a solid oxide electrolysis technology for extracting oxygen from the carbon dioxide-rich Martian atmosphere. • The in-situ production of oxygen is necessary for sustainable human Mars exploration, mainly for the combustion process powering a rocket to return from the Martian surface to space. • Bringing all this oxygen to Mars from the Earth would require multiple launches of costly Saturn 5-class rockets, which would be a serious impediment to sustainable human exploration of Mars. • In-Situ Resource Utilization (ISRU) for planetary exploration (i.e., "living off the land") has been discussed and written about for decades, but MOXIE is the first experiment to actually demonstrate ISRU on the surface of another planet. • The article summarizes the purpose and method of operation of MOXIE, shows the results of oxygen production so far, and shows how operating MOXIE informs the design of future larger-scale oxygen-producing systems. [ABSTRACT FROM AUTHOR]

This paper discusses a system for Titan exploration enabled by nuclear fusion power. Titan is one of the most interesting locations in the solar system with a thick atmosphere, surface oceans, under-ice oceans and complex terrain. This paper provides a conceptual design of a fusion-powered system to explore many parts of Titan and enable the use of high-power instruments. The design includes a fusion-powered orbital transfer vehicle and an electric Titan science aircraft. A Direct Fusion Drive (DFD) propulsive engine could bring a sizable spacecraft to Titan orbit in less than two years. A second fusion reactor, configured as a closed-loop power generator, would be used for an electric Titan science aircraft. Both reactors are based on the Princeton Field-Reversed Configuration (PFRC) concept which combines an FRC with a magnetic mirror. PFRC uses a novel radio-frequency plasma heating system and deuterium-helium-3 fuel. A lower temperature plasma flows around the closed-field FRC region removing the fusion products. In the DFD propulsive configuration, this secondary flow permits direct and variable thrust and exhaust velocity. The science aircraft would do a powered entry to Titan and then have the capability to fly anywhere on the moon at subsonic speeds. The DFD-powered transfer vehicle would allow the in-orbit transfer stage to change inclination as needed to cover different areas of the surface. • A PFRC fusion rocket could deliver 1 MT payloads to Titan in about two years. • A PFRC fusion reactor could power an electric aircraft to explore Titan's surface. • Fusion power allows the aircraft to utilize novel high-power instruments. • The PFRC transfer vehicle doubles as a science platform and communications relay. [ABSTRACT FROM AUTHOR]

All current rehydratable food and drink items prepared for U.S. astronauts on missions to the International Space Station (ISS) must be compatible with the station's potable water dispenser, called the rehydration station. This highly validated system ensures that water is compartmentalized and safely transferred to custom-made food and beverage pouches to prevent free-floating food particles and water in microgravity. However, the food and beverages packaged in these single-purpose pouches require complex, time-consuming, and costly procedures before flight, which is not feasible for future commercial space stations seeking more affordable life-support systems. Furthermore, even though a vast number of rehydratable food and beverage items available at supermarkets have appropriate nutritional content and shelf life for spaceflight, their incompatible consumer packaging renders them unusable by astronauts. In this paper, we present a low-cost solution that draws on NASA technology to adapt these commercially developed food and drink items for spaceflight and makes them compatible with the water dispensing unit and consumption in a microgravity environment. The 3D-printed adapter is based on the NASA septum adapter with the addition of a piercing end and one-way barb that stabilizes the adapter inside the food pouch and ensures it is not released due to the force generated by the water when it is dispensed. The dimensions of the adapter are 20.00 (L) x 12.50 (W) x 37.24 (H) mm and the mass is approximately 1.5 g. The design is compatible with at least three types of packaging commonly used for commercially available rehydratable foods: kraft paper and polymer laminate (KP), polymer aluminum laminate (PA), and compostable polymer laminate (CP). Based on these three packages, we present a table of commercial items that could be available to astronauts if the adapter were to be used. To validate the adapter's use in the space environment, we also present data on the structural properties against loads experienced during launch and re-entry, as well as two functionality tests to determine piercing and rehydration effectiveness. The results show that the prototype of the adapter meets structural specifications for spaceflight and can effectively pierce KP and PA materials with minimal tear propagation. The rehydration test with KP showed that the adapter can effectively rehydrate a typical dehydrated food package. In summary, the adapter is an effective method to increase menu diversity while simultaneously eliminating costs associated with space food preparation and re-packaging. • A new proof-of-concept adapter for food and beverage rehydration in space was demonstrated. • Adapters can be made compatible with U.S. galley equipment on the ISS. • Enabling commercial food on the ISS could significantly reduce costs of human spaceflight. [ABSTRACT FROM AUTHOR]

Tether deployment of the space tether system is the first step of related on-orbit experiments, which requires an advanced and reliable control scheme to realize rapid deployment without overshoot. In this paper, we propose a fractional-order fixed-time controller based on hierarchical sliding mode control to be applied to the deployment process of the tether system. The application of fractional calculus improves the local effect of the controller. Through strict mathematical derivation, we prove that the proposed control law satisfies the global fixed-time stability. So the closed-loop system has the performance of fixed-time convergence. Part of the simulation uses a fractional sliding mode control law for comparison, which further illustrates the superiority of the control law proposed in this paper. In view of the uniqueness of tether system dynamics, the research ideas in this paper are also of reference value to similar under-actuated systems. • Aiming at the underactuated dynamics, the hierarchical sliding mode is adopted. • Multiple sliding surfaces are used, and they all satisfy the fixed time convergence. • The rapid deployment of tethered satellite without overshoot is realized. [ABSTRACT FROM AUTHOR]

This paper provides a concise yet informative overview on future operations in near-Earth space including in-orbit servicing, assembly, manufacturing, and the use of space tugs, in the frame of the space traffic management context. After providing a general definition of these activities, the significant impact that they will have in the space domain from an economic perspective is highlighted. Past achievements in terms of orbital demonstrations are then recalled, leading to the identification of the most critical technological challenges which need to be addressed to improve reliability, frequency, and robustness of servicing operations. A summary of the most critical next steps and recommendations is finally provided. • A general definition and classification of In-Orbit Servicing activities is provided. • The economic and environmental impact of IOS activities is discussed. • Past achievements are reviewed and critical technical challenges are identified. • A summary of the most critical next steps and recommendations is provided. [ABSTRACT FROM AUTHOR]

A long-duration microgravity environment has numerous detrimental effects on the human physiology. The most obvious solution for this problem related to long duration space exploration missions is to remedy the lack of gravity. This could be done using short arm human centrifuges but they do not seem to be sufficiently effective, perhaps because of the short duration exposure of this countermeasure and/or the huge body gravity gradient. New views have to be investigated, such as to see if a (very) long-arm rotating system generating a continuous 1 g or partial gravity field might resolve this issue. Besides the expected benefits regarding astronauts' microgravity pathologies, additionally the spacecraft itself, its on-board (sub-)systems and procedures might benefit from a rotating configuration. In this paper we address very briefly the medical issues, but the work is mainly focused on the advantages regarding engineering, operations, life support, safety and budget of having a constantly rotating spacecraft first in Low Earth Orbit and later for long duration missions to Mars. A large rotating spacecraft is feasible and affordable to build, operate and maintain. It has advantages for governmental and commercial use but also in light of the expected increase in space tourism. It will also save crew time and billions of dollars now being spent to counteract the effects of microgravity. A continuous rotating spacecraft providing Artificial Gravity will: • Improve crew health and wellbeing. • Improve mission safety and success. • Reduce costs significantly. • Simplify numerous operations in flight. • Provides the use of 1 g verified systems. [ABSTRACT FROM AUTHOR]

Iodine has proven to be a promising alternative to xenon as a propellant for Hall and ion thrusters. High atomic mass and favourable ionization characteristics, comparable to xenon in terms of propulsive performance, together with lower cost compared to xenon makes it an attractive option. However, its reactivity and corrosive nature raise concerns. Tests to assess material compatibility and determine whether thruster components are susceptible to corrosion are unavoidable for its safe adoption as a propellant. This paper describes the main results of an experimental campaign aimed at a qualitative assessment of the compatibility with iodine of different materials. The study included more than fifteen materials that find application in electric propulsion devices, such as cathodes, thrusters and feeding systems, including ceramics such as LaB 6 , polymers, iron and aluminum alloys, and graphite. The specimens were exposed to iodine vapor environments at temperatures reaching approximately 200 °C for extended periods (from three to over 30 h) or were directly exposed to liquid iodine. The experimental setup and results of morphological and elementary analyses are presented and discussed. The research provides some insights into the viability of using iodine as a propellant and the potential challenges associated with its corrosive properties. [ABSTRACT FROM AUTHOR]

Laser inter-satellite links (LISLs) are crucial for enabling rapid information exchange within low earth orbit (LEO) constellations. Previous research on laser link establishment often relied on the idealized probability density function (PDF) of the pointing uncertainty regions (PURs) when designing scan-and-capture methods, neglecting the impact of various practical uncertainties. To address the above issues, there is an urgent need to analyze the practical uncertainties to enhance the stability and reliability of LISL. This paper first provides a comprehensive introduction to the alignment process in LISL and mathematically formulates typical sources of uncertainty errors. Subsequently, a generalized system model (GSM) for laser link establishment was established at three stages: Orbit, Attitude, and Laser terminal. Finally, we integrate uncertainty models into the GSM and propose a PUR generating method to obtain precise PDFs. The research method proposed in this study can consider the influence of multiple uncertainties and generate the PDF for the PURs. Compared with traditional research, the PDF obtained by fitting the oscillation range considering dynamics is more detailed. By analyzing the PDF generated by typical uncertainty errors, it is observed that the orbit initial uncertainty may be the predominant factor influencing the PUR. • A generalized system model for the laser alignment process is established. • A method to generate the model for the pointing uncertainty region is proposed. • Envelope curve is used to describe the oscillation range for deviation samples. • The impact of different uncertainty errors on the uncertainty region is explored. [ABSTRACT FROM AUTHOR]

Refractive and diffractive solar sails have been cited to yield benefits in both performance and utility over reflective sails, but their range of viable flight regimes and future applications have not been fully explored. In this paper, a flight model is developed to test and compare these transmissive sail designs under realistic conditions. Raw performance is translated into tangible flight characteristics within a range of flight regimes, such as rate change of orbital energy and minimum operational altitude, and used to make comparison with reflective sails and contemporary thrusters. Additionally, the sensitivity of these flight characteristics to certain orbital parameters is explored when operating under either a locally optimal or simplified Sun-pointing steering law. The developed flight model focuses on solar radiation pressure, atmospheric drag and the effects of eclipse and orbital precession; locally optimal steering laws are numerically generated for every flight regime using a ray tracing-derived performance sensitivity profile. Relative to an idealised reflective sail, the sensitivity of transmissive sail performance is found to be lower for altitude, but higher for orbital inclination. High performance transmissive sail designs are found to outperform idealised reflective ones in every flight regime nonetheless. Meanwhile, only certain lower performance designs demonstrate this trait; others retain advantage only within high inclination, low altitude orbits. 36 m 2 transmissive sails performing an orbit-raising manoeuvre from low Earth orbit are shown to generate transit times comparable to mid-range electric thrusters. In light of these findings, potential applications for transmissive sails are discussed, as well as several practical considerations and potential limitations. • Solar sail performance and operational flexibility improved by transmissive membranes. • Non-meta transmissive sails find advantage at low altitude and/or high inclination. • Transmissive metasails find advantage in every flight regime. • Transmissive sails may fulfil contemporary thruster roles in low Earth orbit. • Transmissive sails may provide feasible space debris mitigation strategy. [ABSTRACT FROM AUTHOR]

In this paper, a novel post-capture takeover control strategy of non-cooperative target with model uncertainties by the dual-arm space manipulator is proposed. Two mission objectives of target detumbling and pose adjusting are comprehensively considered. First, the motion trajectory of the target is parameterized with normalized time by a Bezier shape-based function. On this basis, by considering the resultant external force and moment constraint and the model uncertainties of the target, the mission duration is optimized to obtain the time optimal detumbling and pose adjusting the trajectory of the target. Furthermore, the desired trajectory in manipulator joint space is obtained by introducing the arm angle inverse kinematics method, and a joint trajectory tracking controller is designed based on the feedback linearization theory to track the desired joint trajectory. Simulation results show that detumbling and pose adjusting can be simultaneously achieved by one control action, which verifies the effectiveness of the proposed integrated control strategy. • A time-optimal control strategy is proposed to simultaneously achieve target detumbling and pose adjusting. • The model uncertainty of non-cooperative target is considered in post-capture takeover control. • The arm angle parameterized inverse kinetics is introduced to avoid singularity of the inverse Jacobian matrix. [ABSTRACT FROM AUTHOR]

The space community is moving forward with the development of in-orbit servicing and removal technologies to enable the circular economy in space and improve future space missions' scientific and commercial return and space environment exploitation. The application of nanosatellites to carry out proximity operations presents an additional opportunity to support various novel and cost-effective distributed mission architectures. The SpEye CubeSat mission, funded by the Agenzia Spaziale Italiana in the framework of the Alcor Program, aims at demonstrating inspection and rendezvous guidance navigation and control capabilities of a nanosatellite in close-proximity of the satellite carrier in a low Earth orbit environment. This paper focuses on the design of the SpEye mission and on the guidance and control strategies to enable its technological demonstrations. Particular attention is placed on the safety aspects of trajectory design, considered of paramount importance for the robust autonomous operations of a low-cost CubeSat in proximity to another satellite. • CubeSat mission for proximity operations demonstration. • In-orbit experiments applicable to the In-Orbit Servicing (IOS) and Active Debris Removal (ADR) domains. • Mission design and ConOps of a nanosatellite mission released by a carrier for proximity operations purposes. • Safe inspection design considering the optimisation of information cost metric. • Close range rendezvous solved with sequential convex programming and with enhanced safety applied. [ABSTRACT FROM AUTHOR]

The major purpose of this paper is to study the important role of historical painting art on the creation of space-age era eight decades ago in 1950s. According to the Merriam-Webster Dictionary , the first known use of "space age" was in 1946. In the Collins Dictionary , the space age is the present period in the history of the world when travel in space has become possible, and people use "space-age" to describe something that is very modern and makes you think of the technology of the space age. In American English, space age is the period in modern history characterized by space exploration, usually considered as beginning on October 4, 1957, when the Soviet Union launched the first artificial satellite, Sputnik I, into orbit around the Earth. In the ancient time, many artists including Giovanni di Paolo, Albrecht Dürer, Étienne Léopold Trouvelot and Vincent van Gogh created masterpieces of space related paintings. Even in the extreme ancient era, the cave art already evidenced humans' astronomical knowledge. Besides the technological research and development for space explorations, groups of workers and artists in the areas of culture, painting, architects, installation artwork, etc. had done many contributions in the development of space-age cultural and aesthetic art before 1950s. They not only created so many masterpieces with plenty of space-age aesthetics, but witnessed and recorded the development history of the space age. It goes without doubt that historical space painting art played an important role on the creation of space-age era in 1950s. Therefore, it is mandatory to investigate and reveal the fact. • To study the important role of historical painting art on the creation of space-age era eight decades ago in 1950s. • The first known use of "space age" was in 1946, considered as beginning on October 4, 1957 when Sputnik I launched. • Giovanni di Paolo, Albrecht Dürer, Étienne Léopold Trouvelot and Vincent van Gogh created masterpieces of space related paintings. • In the extreme ancient era, the cave art already evidenced humans' astronomical knowledge. • Groups of artists in the areas of culture, painting, architects, etc. contributed space-age cultural and aesthetic art before 1950s. • Historical space painting art played an important role on the creation of space-age era in 1950s. [ABSTRACT FROM AUTHOR]

Bioregenerative Life Support Systems (BLSS) are a key facet of plans for long term habitats on other celestial bodies. With the Artemis program headed to the moon and SpaceX pushing towards Mars the roadmap is about 10–20 years. Therefore it is vital to consider how BLSS will work with the traditional physicochemical Environmental Control and Life Support Systems (ECLSS). Biological systems are complicated. Their inputs and outputs cannot be turned on and off the way physicochemical systems can. Therefore the accurate monitoring and prediction of these systems is fundamental to integrating the BLSS into the life support system as a whole. This paper will summarize the history of BLSS research, the system requirements for integrating a BLSS into the ECLSS, and the research needed to meet these requirements. As the world looks towards the future of humans living on other celestial bodies, there is a lot of work to do to support keeping them safe and healthy. Bioregenerative Life Support Systems have potential to provide massive support, if they are effectively integrated with the physicochemical systems. • Explores the systems requirements of physicochemical and bioregenerative systems. • Provides updated equivalent system mass estimates using LED lighting technologies. • Calculates return on investment time for bioregenerative food production. • Suggests future research directions for bioregenerative life support systems (BLSS). [ABSTRACT FROM AUTHOR]

On-orbit servicing and on-orbit assembly represent very appealing mission concepts that could facilitate the exploitation of the space environment. Autonomy is a critical requirement for systems aimed at the execution of these operations. The development of enabling technologies for this type of missions is a focal research topic for the global space community. The AUTOMA project (University of Padua, Italy) aims at the development of technologies to enable the on-orbit assembly of a standardized modular unit by means of a robotic arm. The goal is the development of a capture interface, composed of a gripper mechanism and a suite of close-range navigation sensors (a navcam, four time-of-flight sensors, two custom matrix sensors based on a set of phototransistors), and the mock-up of a modular assembly unit. The paper prevalently focuses on the conspicuous series of tests that has been performed at subsystem level for the characterization of the two custom matrix sensors, both in terms of their resolution and range of application. In particular, the test campaign has proven how the in-plane matrix sensor provides information about lateral misalignment and relative distance with a resolution of a few millimetres while the roll matrix sensor provides information about the angular roll misalignment with a resolution dependent on the distance between the two bodies. In addition, an analytical relation for the computation of the relative distance through the information provided by the in-plane matrix has been investigated. • Capture interface composed of a gripper mechanism and a suite of close-range navigation sensors. • Design of a standardized unit mock-up for modular assembly. • Custom sensors based on a matrix of phototransistors activated by an infrared LED beacon counterpart. • Testing campaign aimed at the characterization of the custom matrix sensors. • Analysis of a relation for the retrieval of the relative distance based on the number of phototransistors activated. [ABSTRACT FROM AUTHOR]

The asteroid impacts can be devastating for the life on the Earth. Thus, mitigation of this unwanted event is of crucial importance, and nuclear devices are state-of-the-art choice. The presented paper focuses on nuclear data validation of silicon and oxygen. Nuclear data validation is an essential task because nuclear data are the deciding factor for energy deposition profiles, which serve as input for subsequent hydrodynamic simulation, thus directly affecting the description of asteroid deflection. The silicon and oxygen in the form of silicate materials are abundant in asteroids. The set of neutron and gamma transport experiments using a primary neutron standard 252Cf neutron source was performed. The fast neutron spectrum in the energy range of 2–10 MeV was measured using a stilbene detector. The best agreement for fast neutron leakage spectra was reached using INDEN evaluation for 28Si and ENDF/B-VIII.0 16O evaluation. Activation foils sensitive in different energy ranges were utilized. The results depend on the examined reaction and distance from the source. Finally, the prompt gamma rays from capture and inelastic scattering were measured by an HPGe detector. With the exception of 28Si(n,n′) 1779 keV peak, other prompt gamma production peaks show significant discrepancies. The calculations for comparisons of different transport libraries were performed using Monte Carlo MCNP6.2 code. The results show the necessity of further improvement of evaluations for oxygen and silicon. • Energy deposition profile highly depends on used transport library. • Effect of different transport libraries on calculations. • Measurement of leakage neutron spectra from sand cylinder • Measurement of reaction rates of activated materials. • Measurement of prompt gammas. [ABSTRACT FROM AUTHOR]

Cameras are rapidly becoming the choice for on-board sensors towards space rendezvous due to their small form factor and inexpensive power, mass, and volume costs. When it comes to docking, however, they typically serve a secondary role, whereas the main work is done by active sensors such as lidar. This paper documents the development of a proposed AI-based (artificial intelligence) navigation algorithm intending to mature the use of on-board visible wavelength cameras as a main sensor for docking and on-orbit servicing (OOS), reducing the dependency on lidar and greatly reducing costs. Specifically, the use of AI enables the expansion of the relative navigation solution towards multiple classes of scenarios, e.g., in terms of targets or illumination conditions, which would otherwise have to be crafted on a case-by-case manner using classical image processing methods. Multiple convolutional neural network (CNN) backbone architectures are benchmarked on synthetically generated data of docking manoeuvres with the International Space Station (ISS), achieving position and attitude estimates close to 1 % range-normalised and 1 deg, respectively, an established rule of thumb for the navigation measurement accuracy during final approach. The integration of the solution with a physical prototype of the refuelling mechanism is validated in laboratory using a robotic arm to simulate a berthing procedure. • A novel AI-based solution for spacecraft docking and refuelling is presented. • First demonstration of deep learning for vision-based pose estimation in docking. • Various CNN backbones are benchmarked on photorealistic ISS refuelling datasets. • ResNet-101 reports average 1.19% error in position, 0.29 ∘ in attitude per trajectory. • Prototypes of the developed mechanism are validated in lab with 7-DOF manipulator. [ABSTRACT FROM AUTHOR]