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One of the challenges of nanosatellites use is the limited payload volume. In this regard, transformable structures are frequently used in the development of nanosatellites. Most modern deployment mechanisms are based on thread burning. However, they have a significant disadvantage, which consists in a difficult procedure of thread replacement. The paper presents the novel design of a deployment system for CubeSat nanosatellites, which is based on Rose's alloy. In the "cold" state, the Rose's alloy holds the deployable structure in a closed position, and when temperature reaches 92–98° Celsius, the alloy melts and a structure deploys. On the basis of the experiments, the required heating power for deployment was determined. The heating time and kinematic characteristics of the system were also estimated experimentally. Our paper presents electronic circuitry and the deployment algorithm of the proposed system. We also consider abnormal operation of the deployment system. In addition, a set of environmental tests was carried out: thermal tests, vacuum tests, vibration and shock tests. The proposed system has successfully passed all the experiments and will be launched to space in 2022. • A new deployment system for transformable structures is presented. • The proposed system has no destructible elements. • Deployment process occurs with temperature less than 100 °C. • The system passed thermal-vacuum and vibration tests. [ABSTRACT FROM AUTHOR]

Safe operation of orbital platforms relies heavily on stability of fluid flows in microgravity. The inherent instability of two-phase flows under microgravity conditions could be a threat to fluid transporting equipment operation in Space thus altering Space flight safety conditions. The paper considers a multiphase flow in a Hele-Shaw cell. A dimensionless criterion characterizing the flow in such cells is described. If one changes the parameters of the experiment in such a way that the dimensionless criterion retains its value, then one can observe similar processes in microgravity and in terrestrial gravity. The paper describes the results of experiments on the flow in a Hele-Shaw cell in microgravity during parabolic flights, as well as laboratory experiments in normal gravity. It is shown that the generally accepted mathematical model used to model flows in the Hele-Shaw cell is not suitable for describing multiphase flows. A modified semi-empirical mathematical model is proposed, which gives a satisfactory coincidence of the experimental data with the analytical solution. Non-linear analysis of three fluids displacing each other is performed demonstrating the growth of viscous fingers on both interfaces and their interactions. • Safe operation of orbital platforms relies heavily on stability of fluid flows in microgravity. • The Hele-Shaw cell can be used to simulate microgravity flows in an Earth laboratory. • The existing theory does not correspond to the experimental data. • A new modified mathematical model with a correction factor is proposed. • The results of numerical simulation of the flow in the Hele-Shaw cell are presented. [ABSTRACT FROM AUTHOR]

The development of telescopes with large apertures is driven by the increasing demand for higher imaging resolution and coverage. Large membrane diffraction space telescopes are particularly attractive due to their lightweight nature, ease of folding and unfolding, and high-precision imaging capabilities. However, their substantial size and flexibility make them susceptible to long-duration, low-frequency vibrations during attitude maneuvers, which degrade attitude and imaging accuracy and risk instrument damage. Consequently, an urgent need exists for a high-precision integrated control scheme. In this paper, the integrated control of vibration and attitude in a large space telescope is studied using cable actuators and control moment gyroscopes (CMGs). Since cables can only withstand limited tension, the unilateral and constrained characteristics of the control input are taken into account during the control design process. Firstly, a rigid-flexible coupling dynamic model of the space telescope is established using the velocity variation principle with hybrid coordinates. Additionally, a two-body astrodynamic model is developed to assess the impact of gravity gradient torque. Next, combining the computational torque method and H∞ control theory, an integrated control scheme is proposed, which achieves vibration and attitude control during maneuvers. Then, this paper optimizes the cable actuator positions via the discrete particle swarm optimization (PSO) algorithm and plans various attitude maneuver paths. Finally, numerical simulations are adopted to demonstrate the effectiveness of the control scheme. The results show that this integrated control scheme swiftly suppresses structural vibrations during attitude maneuvers, enabling high-precision attitude control of the space telescope. • A vibration-attitude integrated control for a novel large membrane diffraction space telescope is studied • 3 types of attitude maneuver paths are planned for the two requirements of time optimal and minimum vibration excitation. [ABSTRACT FROM AUTHOR]

Imaging systems larger than the James Webb Space Telescope will enable unprecedented astronomy in the future. Larger aperture space-based imagers will also support next-generation Earth Observation missions for national security and disaster monitoring. However, launching and operating such large telescopes in the poses several practical challenges. A key challenge is that mirrors with apertures larger than 3 m cannot be monolithically manufactured so a segmented design is utilized to achieve primary mirror apertures. Even if larger mirrors could be made, it would be impossible to stow them in fairings of most existing and future launch vehicles. Folded-wing designs to deploy segmented primary mirrors, as done with the James Webb Space Telescope, are one approach to overcoming this volumetric challenge but inappropriate for 25 m apertures considered in this study. This paper presents the concept of operations and mission architectures/analysis for on-orbit assembly of a 25 m aperture telescope operating in the visible waveband of the electromagnetic spectrum capable of 1 m spatial resolution from geostationary orbit. Further, a technology demonstration roadmap towards maturing the robotic assembly technology stack is then presented as precursors to the 25 m imaging system. • Robotic on-orbit assembled large telescopes will enable next-gen imaging. • The paper designs a 25 m aperture space telescope. • In-orbit mass of telescope is ∼ 50 tonnes. • Robotic assembly is architected in a NewSpace lens. • Precursor technology demonstrations are then presented. [ABSTRACT FROM AUTHOR]

The construction of Gateway in lunar orbit is planned as a post-ISS mission by NASA, and JAXA is considering a cargo supply mission to provide service. The Near-Rectilinear Halo Orbit (NRHO) is a candidate for a Gateway orbit, and several methods for the transfer trajectory are being researched. The In-Direct Transfer (IDT) has a short transition period of seven days, but the ΔV application is inefficient. The Weak Stability Boundary (WSB) transfer maintains a small ΔV, but the transition takes over 100 days. However, a trajectory intermediate between IDT and WSB has not yet been actively studied. This paper proposes the Extended Perilune Rendezvous Method (EPRM) of intermediate transfer to NRHO. This method reduces ΔV by orbiting in an elliptical lunar orbit for over two weeks. EPRM has several transition methods corresponding to ΔV and transition periods. This paper systematically reveals the transition mechanisms. Moreover, an effective TCM, considering several error types, is presented for an actual mission. Finally, the rendezvous accuracy with Gateway is discussed. • Extended Perilune Rendezvous Method (EPRM), an intermediate transfer method is well-suited for transferring cargo to Gateway. • Multiple EPRM trajectories can further reduce the ΔV and extending the transfer period. These can maximize the cargo volume. • The position and velocity errors upon arrival at the gateway can be efficiently reduced during the lunar orbit in the EPRM. • This database of the ΔV and transition period can greatly broaden the scope of future cargo resupply missions. • The multiple EPRM trajectories have not yet been found. Further reduction of ΔV may be possible by extending transfer period. [ABSTRACT FROM AUTHOR]

Crater identification aims to match the craters of a planetary body observed in an image to the corresponding entries in a catalogue of known craters of the same body. It is a key step in crater-based navigation algorithms for planetary missions. Many crater identification methods extract descriptors from the observed craters and compare them to an index of descriptors constructed a priori from the crater catalogue. The need for viewpoint invariance motivates descriptors defined over crater tuples (e.g. , crater triads for projective invariance). However, this implies a descriptor index that grows rapidly (e.g. , cubically) with the size of the catalogue, which practically limits the spatial resolution and/or coverage of the catalogue. Another serious drawback is the sensitivity of the descriptors and the matching accuracy to noise and outliers in the crater detection result, which are inevitable due to imperfect crater detection algorithms. In this paper, we propose a novel descriptorless crater identification technique. At its core, our method solves the perspective cone alignment problem with geometric verification iteratively over the crater catalogue. We show that our simple algorithm, with time and space complexities that are linear in the catalogue size, is substantially more accurate than state-of-the-art descriptor-based methods in the presence of noise. Moreover, the availability of multi-core onboard processors raises the prospect of speeding up our method through parallelisation. • This paper introduces the first descriptor-less crater identification method. • It comprises a perspective cone alignment solver and a geometric verification scheme. • It achieves 2–3 orders of magnitude memory improvement over descriptor-based methods. • It demonstrated superior robustness against input noise and challenging conditions. [ABSTRACT FROM AUTHOR]

This paper addresses the problem of guiding a rocket-powered vehicle to land using aerodynamic and thrust controls. The vehicle initially glides towards the landing site, flying at an unprecedentedly high angle-of-attack to decelerate, thereby reducing the propellant needed for a later retro-thrust landing. The high angle-of-attack pure aerodynamic flight leads to highly nonlinear dynamics that have not been addressed by existing powered descent guidance methods. In particular, real-time optimization of the vehicle's landing trajectory has been very challenging. In this paper, a closed-loop guidance method is proposed that optimizes a restricted 6-degrees-of-freedom (6-DoF) trajectory in real-time. To remedy the difficulty caused by the highly nonlinear dynamics, the general propellant-optimal problem is regularized with two modifications. The first is to augment the angular velocity to the performance index, removing possible singularity arcs that cause numerical difficulties. The second is to parameterize the thrust magnitude with a piecewise-constant form, reducing the number of control variables to be optimized. The regularized problem is then solved with successive convexification to update the restricted 6-DoF landing trajectory in each guidance cycle. Monte Carlo experiments with aerodynamic and thrust dispersions are conducted to examine the proposed guidance method. • A new guidance method is proposed that uses both aerodynamic and thrust controls. • The guided vehicle can fly at unprecedentedly high angle-of-attack to decelerate. • The guidance method optimizes a restricted 6-degree-of-freedom trajectory in real-time. • The guidance problem is regularized twice to address the highly nonlinear dynamics. [ABSTRACT FROM AUTHOR]

The primary objective of this paper is to construct an analytical model for determining the total duration of eclipse events of satellites. The approach is based on the assumption that the trace formed in the orbital plane, cutting body's shadow under the classical conical shadow model, can be described as an ellipse. This allows for the derivation of its parameters through the use of classical orbital elements and solar position in the body's centric frame. Consequently, the problem of identifying satellite occultation events is simplified to searching for points of intersection between two ellipses: the satellite orbit and the shadow. The developed model has been demonstrated to be applicable for a wide range of inclinations, with the exception of cases where the orbital plane of the satellite's motion coincides with that of the sun's in the body-centric frame. It is shown that the shadow function constructed under the aforementioned assumptions can be simplified in the cylindrical shadow model. Several simplifications are provided in the research. The paper also considers the extension of the model to account for the non-spherical shape of celestial bodies and presents an algorithm for accounting for changes in eclipse duration caused by the heliocentric motion of celestial bodies. The model has been validated through numerous tests with a numerical algorithm and satellite real data, as well as with other analytical models. • An analytical method for calculating the total duration of satellite eclipses has been developed. • Developed approach can be applied to wide range of satellite orbits either hyperbolical or elliptical ones. • The equations that depend on the satellite's orbital elements and solar longitude to predict satellite eclipses are obtained. • Simplification of the problem in certain scenarios, allowing for rigorous and simple algebraic solutions are provided. [ABSTRACT FROM AUTHOR]

Currently, considerable attention is being directed toward active debris removal (ADR) using small, cost-effective satellites. From high altitude to re-entry, ADR missions use thrusters that consume significant amounts of fuel. The volume of the propulsion system in these satellites can be quite large, and it is difficult for a small satellite to accommodate a large-scale propulsion system. Thus, active utilization of atmospheric drag is being considered for small satellites with low capability. However, the principal challenge for small satellites to remove large debris is to deal with the large disturbance torque caused by strong atmospheric drag on the debris. This paper proposes attitude guidance around an equilibrium (balance) point in attitude dynamics acted on by gravity gradient and aerodynamic torques. The desired attitude is set to optimize two conflicting objectives: maximizing the projection area to generate large aerodynamic force and minimizing the disturbance torque. In this paper, we develop an aerodynamic database (ADDB) by rigorously calculating the aerodynamic torque for each altitude and solar array paddle (SAP) rotation based on an aerodynamic characteristic model of free molecular flow. Using the ADDB, the SAP angle and orbit altitude effects on the torque equilibrium attitude (TEA) are clarified. Then, the mean disturbance torque map at each orbit altitude is introduced to obtain the candidate TEA paths. The effectiveness of the proposed paths for debris descent is demonstrated through numerical simulations. • Active debris removal mission by small satellites was considered. • Aerodynamic database was developed using an aerodynamic characteristic model. • The effect of SAP angle on aerodynamic moment was evaluated. • Candidate TEA paths were obtained from the proposed torque disturbance maps. • The optimal path was designed considering both duration and amount of torque. [ABSTRACT FROM AUTHOR]

NASA HRP considers developing radiation protection for human space flight and surface habitation as one of the critical technologies for successful deep space exploration. Although the danger of radiation exposure is recognized as a potential show-stopper for deep space exploration, and radiation effects on humans are now understood better than before, shielding strategies for different stages of space flight and habitation are still not addressed in a fully comprehensive manner. Habitats and all space structures planning and design are guided by essential requirements and constraints associated with safety, manufacturing and assembly procedures, propellant and construction costs, maintenance, crew and/or passenger satisfaction. This paper reviews radiation shielding options for space habitats and discusses their feasibility in relation to habitability needs, including the potential for outside viewing in distinct types of space habitats. Typically, radiation protection depends on the thickness of the exterior structure that consists of a pressurized shell, multilayer insulation (MLI) and any applied radiation shielding material. The mass of the external protection shell is the primary factor of radiation shielding effectiveness. Nevertheless, using materials with low atomic numbers (e.g., Boron (5), Carbon (6), and H 2 O) helps to lower secondary radiation hazards. Water, which is rich in hydrogen and has the lowest atomic number, can also be used for other mission needs. Other materials, including biomaterials, can be considered when their inclusion in the structure is possible and appropriate. The paper presents a comprehensive strategy for radiation shielding selection that includes investigation of complications associated with a type of space habitat structure, mission needs, duration, destination, and crew requirements. The approach presented in the paper aims to establish an evaluation methodology for defining the feasibility of the integration of diverse radiation shielding types into habitat structures. The paper summarizes by reviewing radiation shielding proposals with selected case studies including water, regolith, hydrogen-rich polymers, biotechnology, polyethylene/boron nitride composites, and active strategies. • Case studies of different materials and methods used for radiation shielding purposes, • Pros and cons of using referenced materials and methods in habitat architectures • Criteria for evaluation and selection of appropriate radiation protection strategy • Comparisons of radiation shielding methods by space architecture categories • Comparisons of radiation shielding methods by structural integration options [ABSTRACT FROM AUTHOR]

The satellite attitude error is transmitted to the gravity field through KBR antenna phase center correction and non-conservative force conversion, which affects the accuracy of gravity field inversion. Therefore, obtaining high-precision satellite attitude is one of the important research contents of gravity field inversion. Each satellite of GRACE-FO is equipped with three SCs and an IMU for satellite attitude measurement, and fusion of these two types of data can obtain high-precision satellite attitude. In this paper, we use kalman filter to fuse the original GRACE-FO 1A observation data, and take into account the calibration of the installation matrix of the SCs during the fusion process. Meanwhile, a quaternion integrator that takes angle increments as inputs is derived from the equivalent rotation vector differential equation to address the computational inefficiency of first-order quaternion integrator, the accuracy and computational efficiency are verified using measured data. The experimental results show that after the calibration of the installation matrix of the SCs, the system bias of 2 0 ′ ′ caused by switching from three SCs observations to two SCs observations can be eliminated. The standard deviation of the attitude difference between the obtained attitude in this paper and the attitude calculated by JPL is at most 6 ′ ′ and the proposed incremental quaternion integrator has the same accuracy as the first-order quaternion integrator but with computational efficiency improved by at least 35%. • Attitude determination is critical for gravity field inversion and Orbit determination. • The installation matrix calibration method is simple and effective. • The incremental quaternion integrator is high-precision and computationally efficient. [ABSTRACT FROM AUTHOR]

The global community, with coordination from the United Nations, is energized to pursue the Sustainable Development Goals (SDGs), a list of 17 important aspirations that summarize the key challenges of our era. The SDGs apply to every nation and represent an international effort to eliminate extreme poverty, ensure access to safe drinking water, strengthen food security, and produce clean and reliable energy, among other pursuits. Space technology is already being used around the world to advance progress toward the SDGs and monitor their related Indicators. This paper explores how six technologies related to space—satellite Earth observation, satellite communication, satellite navigation and positioning, human spaceflight and microgravity research, space technology transfer, and basic scientific research—are being used to realize the vision that the SDGs represent. The paper also discusses the obstacles that limit the application of these technologies for the SDGs and provides an overview of potential paths to overcome these barriers. The paper finally studies the historical and potential roles that four distinct types of entities involved in global sustainable development—governments, non- and inter-governmental organizations, entrepreneurial companies, and universities— have played or may play in the application of space technologies towards the SDGs. • Six space technology areas are shown to support the U.N. Sustainable Development Goals. • Examples of space projects that support the SDGs are found in many sectors. • New designs for technology are needed to improve effectiveness of space applications for the SDGS. [ABSTRACT FROM AUTHOR]

The regional coverage satellite constellations are in demand for fulfilling earth observation needs of specific user regions of interest as opposed to global coverage thus enhancing the revisit period for the area. In this paper, we present a few key mission aspects of the regional coverage constellation designed and operated by the Indian Space Research Organization (ISRO) in low earth orbit. ISRO developed three X-Band Synthetic Aperture Radar (SAR) Earth Observation satellites and placed them in three inclined orbit planes to enhance the Indian region's coverage. This is India's first inclined orbit constellation designed for remote sensing applications in low earth orbit. The satellites were built around a standard bus carrying a Synthetic Aperture Radar payload with a deployable Radial Rib Antenna (RRA). They were designed to achieve a high agility capability. As SAR sensors are capable of all-weather, day, and night imaging; thus, most suitable in inclined orbits where the illumination conditions vary continuously for imaging. Though the inclined orbits provide more revisit opportunities to the region of interest, their characteristics pose challenges in mission design and operations. Though the mission design and planning involved a host of elements, this paper brings out the details of a few key mission analysis elements, planning aspects and operations that were critical in the course of various mission phases. The mission analysis included mission-specific requirements like launch windows selection for the constellation and reference attitude generation specific to inclined orbit geometry. Mission planning had to take care of Radial Rib Antenna (RRA) deployment requirements which were very critical for the mission, apart from taking care of individual subsystem requirements. The SAR payload data product quality and targeting accuracy are very sensitive to the satellite image point range (vis-a-vis SAR operation data window). Extensive post-launch calibration activity was carried out to meet all these mission requirements and data product specifications. The challenge was in fixing the payload beam alignment and fixing the uncertainties in onboard and ground computations. The paper also describes the methods of fixing all possible error sources affecting the high-resolution SAR payload performance. • The regional coverage satellite constellations are in demand for area targets coverage. • Inclined orbit constellations are more suitable for regional coverage using SAR sensors. • Mission operations plans need to consider all the specific requirements of spacecraft systems. • Pre-launch and post -launch calibrations of sensors is mandatory for achieving system performance. [ABSTRACT FROM AUTHOR]

This paper examines the effect of shortening the post-mission disposal (PMD) orbit lifetime (remaining orbit lifetime after PMD) from 25 years to 5 or 1 year. Using NEODEEM, the debris evolutionary model developed by Kyushu University and JAXA, the change in the orbital environment at various altitude bands for long-term stability and short-term safety are discussed. It was confirmed that the PMD compliance rate is more important than shortening the orbit lifetime for improving the long-term stability of the orbital environment. In the short term, the 5- or 1-year rule reduced the collision rate and collision avoidance frequency at an altitude below 700 km, more than the 25-year rule, thus shortening the PMD orbit lifetime improves short-term safety, especially at low altitudes. The paper also evaluates the effect of large constellations (LCs) and it was shown that shortening the PMD orbit lifetime of LCs improves short-term safety, but the long-term impact is dominated by the number of LCs that remain at high operational altitudes due to PMD failure. It shows that the post-mission disposal lifetime and required compliance rate should be set according to the number and size of objects expected to be launched in the future. • The effect of shortening post-mission-disposal lifetime from 25 years are shown. • Short-term safety can be improved by 5-year rule or 1-year rule. • PMD compliance rate is more important for the long-term stability. • Difference between an elliptical and a circular orbit for PMD is shown. • PMD requirement changed according to the characteristics of the large constellation. [ABSTRACT FROM AUTHOR]

The threat of debris collisions in Low Earth Orbit (LEO) has driven many researchers to develop Active Debris Removal (ADR) missions to manage large debris in LEO. A modern method of capturing debris with a non-rigid tether has been a recent research focus because of difficult post capture dynamics associated with controlling an uncooperative debris through a tether. To facilitate these control schemes, this paper proposes a new two-stage inertia estimator that estimates all principal and products of inertia for an uncooperative space debris using LiDAR and tether force measurements made on-board a CubeSat sized chaser satellite. The method proposed in this work estimates the debris center of mass with a novel pseudo measurement Kalman Filter so that the full inertia tensor can be subsequently estimated using an additional iterative parameter identification algorithm. In addition, the paper achieves online inertia estimation without assuming the tether connection point on the debris by approximating its location using tether tension measurements. The proposed control and estimation scheme is shown in this paper to estimate the full debris inertia tensor even with frequent tether slackness. Two simulation scenarios are presented in this paper, one where the connection point of the tether on the debris is approximated using noiseless measurements of tether tension, and a second where the location of the tether connection point is approximated using noisy measurements of tether tension. For the simulation conditions used in this research, it is shown that the proposed two-stage estimator (TSE) is successful for both cases that the debris is inertially asymmetrical and tri-inertial with non-zero products of inertia. • Determinable input torque without assumption of Body Frame tether connection points. • Novel Two-Stage Estimator for all principle and products of inertia for space debris. • Responsive Kalman Filter formulations under frequent tether slackness. • Center of mass estimation and tracking for debris object using LIDAR sensor model and Kalman Filter. • Station keeping control law for CubeSat chaser satellite. [ABSTRACT FROM AUTHOR]

Demand for imaging of large area targets by space-based assets such as constellations of Earth Observation Satellites (EOS) is on the rise for different applications. The area targets coverage cannot be achieved by a single satellite in a single observation opportunity, as the region for imaging spreads wider than the coverage swaths of imaging sensors. Hence, the imaging satellites are scheduled in a cooperative way to achieve maximum coverage of the area in a given planning time horizon. The scheduling algorithms also aim to maximize or minimize some specific objectives of interest while simultaneously complying with all resource constraints for an efficient use of precious space resources. As the current generation satellites are highly agile, there are many ways of choosing an imaging location in the target area for each single observation opportunity of a satellite. Even though the possible imaging region in continuous space is discretized into a finite number of strips for each observation, the huge combinatorial search space formed with all satellites' imaging opportunities makes the problem intractable. Thus, the area target imaging scheduling is a complex combinatorial NP-hard optimization problem. Many exact and heuristic methods were evolved to provide a solution to this scheduling problem. Even though the exact methods solve the problem to optimality, they are limited to smaller problem instances because of the high computational complexities involved in those methods. More focus is seen in the literature on heuristic approaches as they guarantee scheduling performance with feasible solutions within an acceptable computational complexity. This paper presents our efficient heuristic approach proposed for the scheduling problem for imaging a large area target using a constellation of multiple satellites. Our approach is an extension of a widely used greedy approach in combination with insights from dynamic programming. In this paper, first we addressed the scheduling problem by describing it in detail and formulating it into a non-linear integer programming problem, considering the multi-objectives of achieving maximum coverage and minimum imaging resolution. The solution to the problem comprises two phases, namely the area decomposition phase and the scheduling phase. In the area decomposition phase, the area target is divided into strips dynamically for each observation opportunity of a satellite. Then exact observation period of each strip is computed using a simplified semi-analytical computational method. In the scheduling phase, we applied our heuristic search strategy, namely Forward and Backward Heuristic Search (FBHS) to obtain a near optimal solution within an acceptable computational time. Through the extensive simulations conducted under various scenarios, the effectiveness of the proposed method is verified in comparison with the baseline greedy approach. • The large area target imaging scheduling for multiple satellites is a complex combinatorial NP hard optimization problem. • Greedy algorithm, which guarantee the scheduling performance, is a widely used method. • Forward and Backward Heuristic Search method is proposed for area target image scheduling. [ABSTRACT FROM AUTHOR]

Electric propulsion, which is an advanced form of thrust production, is widely used in small satellites. Measuring the thrust of electric propulsion in an orbit is difficult because it is very small (μN∼mN). However, the thrust of electric propulsion can be determined by evaluating the change in the speed of a momentum wheel. The contributions of this paper mainly include two aspects: 1) a robust PID controller and 2) a thrust inversion method. The torque of the electric propulsion is regarded as the disturbance torque in this paper. To achieve finite-time attitude stabilization of a satellite during electric propulsion, a robust PID controller is designed. Furthermore, the thrust of the electric propulsion can be determined in the stable attitude of the satellite by only measuring the speed derivative of the momentum wheel. Compared with the traditional thrust inversion method, the designed thrust inversion method is simple and effective and does not use a filtering algorithm. In the HIL experiment, the error rate between the inverted thrust (31.287 μN) and the designed thrust (30 μN) is 4.29%. Most importantly, the thrust inversion method is verified in orbit based on the small satellite APSCO SSS-1, which is consistent with the HIL experimental results and ground test results. • The torque of electric propulsion is regarded as the disturbance torque. • The controller can fulfill finite-time attitude stabilization of the satellite. • Thrust can be calculated by evaluating the speed change of the momentum wheel. • Thrust inversion method is verified in orbit based on the small satellite SSS-1. [ABSTRACT FROM AUTHOR]

Low-thrust trajectory design in the context of the three-body problem, particularly in the vicinity of liberation points, is addressed in this paper with a focus on efficiency and accuracy. The Fourier series, previously successful in approximating near-coplanar low-thrust trajectories within two-body dynamics, is applied to the challenging landscape of the three-body problem, marked by instability and nonlinearity. This paper introduces a novel approach for constructing an initial guess trajectory, that combines coast arcs from periodic orbits and intermediate trajectory arcs from other natural dynamical structures. The resulting initial guess trajectory is instrumental in generating efficient initial estimates for the coefficients of the Fourier series. The study encompasses the analysis of homoclinic and heteroclinic transfers in both two-dimensional and three-dimensional scenarios. The approximated trajectories derived from this methodology serve as invaluable starting points for high-fidelity optimization solvers, offering promise for enhancing the precision and efficiency of low-thrust transfers in cislunar space. • The shape method based on Fourier series is applied to the initial design of the transfer trajectories between periodic orbits in cislunar space. • An innovative approach to construct a continuous initial guess trajectory is introduced to derive initial guess values for the coefficients of Fourier terms. • An accurate approximation of the optimal transfer trajectory can be rapidly obtained. [ABSTRACT FROM AUTHOR]

Comet Interceptor (Comet-I) was selected in June 2019 as the first ESA F-Class mission and aims to explore a pristine comet, which will visit the inner Solar System for the first time. Comet-I will hitch a ride to a Sun-Earth L2 quasi-halo orbit, as a co-passenger in ESA's M4 ARIEL's launch. At the time of writing, this is scheduled for 2029. It will then remain there awaiting the right departure conditions to leave definitively and intercept a newly discovered comet. Comet-I will be the first mission to be designed and launched without an already identified target. Given the latter, the Comet-I Target Identification Working Group is tasked with the continuous follow-up of newly discovered comets, to analyse their characteristics as virtual targets for Comet-I. The goal of the paper is to inform the decision-making process that will eventually be necessary to commit Comet-I spacecraft to its final target. Within this context, this paper presents a trajectory optimization process based on Primer Vector theory that considers intercept trajectories with multiple revolutions and multiple deep space manoeuvres (DSM). The process is applied to 21 long period comets, which were discovered since July 1, 2020 and whose characteristics are of interest for the Comet-I mission. In fact, one of the latest of these recently discovered comets, C/2023H2, is the most accessible virtual target discovered at the time of writing, with delta-V intercept costs well within Comet-I capabilities. A complete launch window analysis is thus presented for C/2023H2, which accounts not only for the most optimal delta-V intercept transfer, but also considers all launch options that satisfy all Comet-I's current boundary conditions and flyby constraints. A systematic launch window analysis is made possible by a grid search approach on which departure and fly-by dates are fixed and the remaining design variables, including the times for manoeuvres and their Δ v impulse vector, are optimized via Primer Vector theory. This allows to understand the latest possible departure date, but also the sensitivity to departure and arrival dates of key encounter characteristics such as relative fly-by speed, the solar aspect angle and the Earth distance. • A Primer Vector Theory-based trajectory optimization tool is shown, tailored for ESA's Comet Interceptor (Comet-I). • The method supports trajectories with multiple revolutions around the Sun and multiple impulses. • The method allows rapid grid search over the search space, finding the lowest cost trajectory and other notable solutions. • A novel analysis of the launch window is presented, contributing to the decision-making process of Comet-I. • A multi-manoeuvre strategy benefits Comet-I's trajectory design, lowering the delta-V cost and improving mission flexibility. [ABSTRACT FROM AUTHOR]

In order to accurately estimate the period of X-ray pulsars to improve navigation performance, this paper proposes a pulsar period estimation method based on photon energy using reverse Hilbert coding and double Convolutional Neural Networks (CNNs) discrimination. This method uses photon energy information to perform epoch folding. It converts one-dimensional contour information into two-dimensional image signals through reverse Hilbert curve and recognizes contour image features through CNN models to determine the optimal period. According to the simulation results, the period estimation accuracy of this method is 0.5350 ns, which is 65.71% higher than the χ 2 epoch folding method. The contour phase accuracy is 0.0004767, which is 41.93% higher than the χ 2 epoch folding method. At the same time, the observation time and noise interference that may affect the accuracy of period estimation are also quantitatively analyzed. In addition, in order to simulate real navigation scenarios, this paper also conducts dynamic period estimation experiments on PSR B0531+21 pulsar in a specific period of time. The method proposed in this paper has the advantages of high precision, fast calculation speed, strong anti-interference ability and accurate dynamic period estimation, which can significantly improve the navigation performance of X-ray pulsars. • The proposed period estimation method uses photon energy information for contour weighted folding. • The proposed method uses reverse Hilbert curve to convert one-dimensional contour information into two-dimensional image information. • Convolutional Neural Network models are used to optimize pulsar period estimation. [Display omitted] [ABSTRACT FROM AUTHOR]

Over the past decade, the number of member states in the United Nations Committee on the Peaceful Uses of Outer Space (UNCOPUOS) has risen by 40%. The UNCOPUOS continues to be one of the largest committees in the United Nations, with recent additions representing many emerging space nations including the Dominican Republic, Rwanda, Angola, Guatemala, and Bangladesh, among many others. This paper addresses the role of emerging space nations in updating and refining current policies and norms of behavior related to the long-term sustainability of the space environment. The paper provides examples of the recent implementation of long-term space sustainability design and operational guidelines in the national space strategies of several emerging space nations, highlighting the importance nations give to the development of legal mechanisms to regulate the peaceful use of the space environment. Examples of both national and regional initiatives are presented including Thailand's 2021 Draft National Space Act, aimed at creating a national legal regime and establishing a governmental agency dedicated to developing space policies for the registration of objects launched into outer space and space debris mitigation measures, and the National Space Law Initiative (NSLI) study group consisting of Australia, Indonesia, India, Japan, Malaysia, the Philippines, the Republic of Korea, Thailand, and Vietnam to create a framework that aims to promote information sharing and mutual learning in relation to the participants' respective national regulatory frameworks for long-term space sustainability. More recently, new initiatives have been developed that celebrate the efforts of satellite mission operators with the aim to reduce the likelihood of space debris and collisions among space objects. The Space Sustainability Rating (SSR) was created by the World Economic Forum, the European Space Agency, the University of Texas at Austin, BryceTech, and the Massachusetts Institute of Technology, and is now hosted by the EPFL Space Centre. The SSR is a rating system to assess and recognize missions that are designed to be compatible with sustainable and responsible operations, reducing the potential harm to the orbital environment and the impact on other operators. The SSR comprises six modules aimed at assessing missions for their compatibility with sustainable and responsible operations. This paper specifically focuses on evaluating the Detection, Identification and Tracking (DIT) scores for satellite missions launched by emerging space nations. The DIT module of the SSR serves as a standardized measure for assessing space sustainability. The paper adopts an exploratory multi-case study approach. Through this focused study, the paper identifies barriers and unique challenges emerging space nations face, including experiences of operator organizations, launch options, financial constraints, technical options available, and other relevant factors. • Paper title: Role of Emerging Nations in Ensuring Long-term Space Sustainability. • UNCOPUOS membership rises 40%, highlighting emerging nations in space policy. • Emerging space nations prioritize guidelines for peaceful and sustainable space use. • SSR serves as an effective tool for evaluating space sustainability metrics. • Case study on THEOS, CBERS, and SumbandilaSat reveals barriers for new players. [ABSTRACT FROM AUTHOR]

Secondary and High School students are usually adolescents, most of them in the urban areas and metropolitan cities are aware about the existence, various opportunities, career paths and advancements of the space industry. They are constantly updated, given they have access to quality and organized education models and a competitive environment. Although, the students in the rural areas who are equally ambitious and talented yet are not fortunate enough to have access to the state of art facilities like the counterparts. Even if they try and gain access with their self-interest, they do not have the motivation to sustain the passion, given their lack of clarity in subject, lack of awareness about engaging in various ways and fields in space other than technological related and mainly lack of driving force. The purpose of this paper is to develop and present an education model, exclusively for the high school students of the rural areas and inculcate sustainable passion for space. As opposed to the primary school students, the secondary school students have a consciousness of the space industry. This paper will talk about various plans and strategies, that will allow them to gain exposure, guide them about the existence of different career paths which are not entirely technological related, yet are still essential in the space like, administrative framework for example. This opens a new dimension in their brain helping them realize how they can be part of the industry in countless ways, and this shall assist them in being on the right track to space. In addition to, hosting creative space themed exhibitions, fun and engaging model design competitions, workshops, all of these would widen their knowledge boundaries and help them challenge their skills. A real time case study example from a school located in the rural southern of India shall be focused on this paper. Including data from a survey thereby the results, with action plans shall be presented in detail, which will aid in the analysis of the need, feasibility, and long-term sustenance of this approach and the requirement to scale it over different rural areas globally. • Need for improvised education models in rural areas elucidated, with data collection. • Data analysis and correlation of space education awareness, for high school students across two geographical locations. • Readily executable action plans age groupwise, for real-time education models in remote areas globally. [ABSTRACT FROM AUTHOR]

This paper reviews manufacturing progress in making materials that have the strength necessary to form the tether for the space elevator. These fall into two categories, nanotubes and 2D materials. The aim is to raise awareness of the properties of these new materials and the state of the art of manufacturing methods. A NASA Innovative Advanced Concepts (NIAC) space elevator feasibility study concluded in 2003 that the "space elevator could be built in the near future with acceptable risk and less funding than some current space programs". The key technology that needed development was the tether component. The study concluded that carbon nanotubes were the material of choice for the tether. Considerable effort has been invested in manufacturing carbon nanotubes. Publicly disclosed information reveals that the longest carbon nanotube made is 0.5 m. The current state of the art is to manufacture carbon nanotubes in bundles called forests. Waseda University has shown that these can be grown at lengths of 0.14 m, at speeds of 1 m in 186 h. One year after the NIAC feasibility study was published, a new class of materials was isolated for the first time. In 2004, graphene was isolated as the world's first two-dimensional material. In 2010 the Nobel prize was awarded to the researchers who isolated and characterised the material, discovering it to have a similar tensile strength to carbon nanotubes (130 GPa). In the past decade much academic and industrial research effort has resulted in the manufacture of graphene as a large area material that can be manufactured at speeds of 2 m per minute and lengths of up to 1 km. While the manufacture of tether-quality material still needs more development, the trajectory to a high-quality industrial product is clear. This paper will detail the current state of the art of manufacturing for these tether candidate materials. • There is considerable progress in manufacturing 2D materials suitable for a space elevator. • The emphasis is on single-crystal graphene and graphene super laminate. [ABSTRACT FROM AUTHOR]

The increasing number of man-made objects in near-earth space is becoming a serious problem for future space missions around the Earth. Among the proposed strategies to face this issue, and due to the passive and propellant-less character, electrodynamic tethers appear to be a promising option for spacecraft in low Earth orbits thanks to the limited storage mass and the minimum interface requirements to the host spacecraft. This work presents the roadmap that the Electrodynamic Tether Technology for Passive Consumable-less Deorbit Kit (E.T.PACK) is following to develop a prototype of a deorbit device based on electrodynamic tether technology with Technology Readiness Level 4 by the end of 2022. The paper illustrates the roadmap of the activities carried out at the University of Padova, where software and hardware have been prepared to validate some of the critical elements of the deorbit device. Specifically, the software tools include: (a) the software called "DEPLOY" that allows the computation of a reference trajectory for the deployment of the tether and the completion of sensitivity analysis of the deployment trajectory to key error sources; (b) the software called "FLEXSIM" that predicts the performances of electrodynamic tethers as a function of the system configuration employed; and (c) the software called "FLEX" that includes the dynamical effects of tether flexibility and provides important information on the dynamic stability of the system during deployment and deorbiting phase. The paper describes in detail the three software tools and provides results of a simulation showing how it is possible to deorbit a 24-kg satellite from an initial orbital altitude of 600 km in less than 100 days using a 500-m long tape-like bare tether. The team has also developed laboratory mock-ups and performed experimental activities to: (a) determine the tether mechanical properties; (b) test the functionality of mechanisms used to deploy the tether; (c) test the functionality of the attitude control assembly used during the deployment phase; and (d) validate a passive damper designed for dissipating the longitudinal oscillations of the tether and thus guarantee the stability of the system during both deployment and deorbiting phase. The paper provides a description of both the laboratory setup and the experimental activities performed to validate EDT technologies, including the damping capability of a compact passive-damping mechanism, showing how it can reduce consistently the peak forces up to about 80%. • Optimization of full-length deployment trajectory. • High fidelity simulations of deorbiting with electrodynamic tethers. • Experimental validation of technologies for tape electrodynamic tethers. • Close-proximity deployment testing with frictionless table and trajectory control. • Development and testing of passive in-line damper. [ABSTRACT FROM AUTHOR]

Aerospace solar arrays serve as the primary means for spacecraft to collect energy. Flexible solar arrays have the advantages of a high deployment ratio, light weight, a large span, and high specific power. Currently, they are primarily used on space stations. Ultra-thin, multi-stepped variable-section flexible solar arrays are a novel design solution. Due to extremely low stiffness, flexible solar arrays exhibit nonlinear behavior such as large deformations under bending during the deployment process driven by external forces, which is the main topic of this paper. To quickly obtain the morphology of the flexible solar array after unfolding and improve design efficiency, this paper first transforms the unfolding morphology of the flexible solar array into a nonlinear large deflection problem of a fourth-order discontinuous variable-section composite beam and gives the equivalent methodology. Subsequently, an efficient nonlinear difference algorithm is proposed. A set of differential equations for large deflections of multi-stepped variable-section beams under complex loading conditions is derived. The finite difference method is employed to discretize this system, and four different categories of node processing methods are derived in detail. Newton's iterative technique is used to solve for the curvature of each point, and explicit formulas to calculate the point coordinates, strain, and stress are developed. Ground deployment experiments, numerical simulations of on-orbit deployment, and comparisons with benchmark examples are used to verify the validity of the ND algorithm. Finally, the morphological analysis of the on-orbit deployment of the flexible solar array is performed, and some design advice is given. [Display omitted] • This approach can be used in the analysis of various flexible spacecraft. • It is applied to a certain flexible solar array. • The on-orbit morphology can be calculated quickly, which facilitates the design. • It is clear, highly reliable, and can be easily generalized to complex problems. • Some design advice is given for flexible solar arrays. [ABSTRACT FROM AUTHOR]

Advances in artificial intelligence (AI) and automated robotics will profoundly influence space operations. By utilising machine learning and deep learning approaches, AI-enabled systems may accomplish tasks as well as improve their own performance. These capabilities are useful in the often-remote settings of outer space and will grow in value as automated space operations become more widespread. As AI extends throughout the space domain, automated algorithms will take on many of the roles that have historically been handled by humans. Artificial intelligence is progressing from theory to implementation in the space environment by exposing new satellites and orbital autonomous vehicles to new data. Even though all initial computational parameters are provided, such systems' outputs can be very unpredictable, putting people, property, and the environment at risk. This paper investigates the application of United Nations space treaties, selected regional AI regulations, and various 'soft-law' instruments and industry initiatives focusing on responsible AI system development to space-based AI systems. Following that, reforms are proposed to clarify the practical relationship between AI systems and the international legal regime that governs space, as well as a 'bottom-up' regulatory approach to better facilitate the future development of regulation governing the use of AI by the global space sector. While this work does not purport to provide a conclusive resolution to these multifaceted matters, its objective is to underscore significant obstacles that arise at the convergence of space law and AI, serving as a preliminary foundation for subsequent discussions on this issue. • Advances in AI and automated robotics will have a profound impact on space operations. Automated algorithms will take on roles traditionally handled by humans as AI becomes more widespread in space. • However, the unpredictable outputs of AI systems can put people, property, and the environment at risk, raising questions about liability. • The paper investigates the application of UN space treaties, regional AI regulations, and industry initiatives to space-based AI systems. • Reforms are proposed to clarify the relationship between AI systems and the international legal regime governing space. • A 'bottom-up' regulatory approach is suggested to facilitate future regulation of AI in the global space sector. [ABSTRACT FROM AUTHOR]

In order to solve the deficiencies of traditional board-level assembly CubeSat, such as low modularity, complex assembly process and poor scalability, a sandwich module assembly structure and method with no drive components is proposed, which can achieve quick connection and flexible assembly between modules. First, the proposed CubeSat standard module can be compatible with the standard interface of traditional CubeSat. The 4-in-1 interface of Mechanical-Electrical-Thermal-Data does not occupy the internal space of the standard module, and the electrical and thermal interfaces meet the requirements of CubeSat in-orbit work. Secondly, in order to improve the reliability of the connection between the modules, a plug-in connection scheme of the mechanical interface was designed. Finally, a genetic algorithm was used to optimize the interface size, so that it has better plug-in performance. The simulation results show that the performance of the male interface of titanium alloy material is the best, and the error between the finite element analysis and genetic algorithm results is less than 4.8 %, which is consistent with the actual tensile test results of the optimized prototype, and meets the mechanical requirements of the launch process and space orbit operation for CubeSat. The test results can guide production and provide a reference for subsequent engineering applications of CubeSat modular assembly. • In this paper, a standardized modular assembly method of CubeSat sandwich structure is designed. The assembly structure does not include driving parts, is simple in form, can be quickly connected, is flexible in assembly mode, and does not occupy the internal space of the standard module. • A standard 4-in-1 connection device was designed in this paper, including four interfaces: mechanical, electrical, thermal and data, which can achieve "plug and play". • This paper uses genetic algorithm to optimize the structure of the mechanical connection male joint, and obtains the optimal configuration of the male joint. This ensures the reliability, firmness and stability of mechanical connections, and meets the mechanical requirements of CubeSats during launch and in orbit. [ABSTRACT FROM AUTHOR]

With an expected global revenue exceeding 1 trillion USD by 2040, the space industry is one of the world's fastest-growing sectors. Given the booming investment in the space industry and the anticipated space tourism era, it is crucial to assess the impact of already operative launch assets as well as to adopt design-to-sustainability strategies for the under-development and future launchers. This paper discloses an integrated methodology to estimate nitrogen oxides emissions of a hydrogen-fuelled air-breathing concept. The proposed strategy combines multi-fidelity propulsive system modelling with 0D chemical-kinetic air/hydrogen combustion numerical simulations to define a high-fidelity emissive database representative of various on-ground and in-flight operative conditions. The propulsive and emissive databases are then used to suggest an analytical formulation able to predict nitrogen oxides emission indices in any in-flight conditions, perfectly fitting the conceptual design needs. Throughout the paper, the Skylon spaceplane and its Synergetic Air Breathing Rocket Engine are used as case study. The methodology disclosed allows proving the high competitiveness of this air-breathing space launchers with respect to famous past and current competitors, as the Space Shuttle and the Falcon 9. [Display omitted] • Integrated design methodologies can support NOx estimations in conceptual design. • 0D chemical-kinetic simulations with Z24 scheme provides reliable E I N O x estimation. • Air-breathing launchers prove to be environmentally friendlier than conventional ones. • Upgraded P 3 T 3 method can support E I N O x estimation for hydrogen-fuelled launchers. [ABSTRACT FROM AUTHOR]

Space-age technology makes it possible to travel faster than the speed of sound from one continent to another one or even to make suborbital flight as a space tourism. The existing technology challenge is to bring this innovative way of travel without damaging the environment to ensure human safety and to preserve environment balance. The sonic boom issue is one of the concerns when flying faster than sound. The issue must be taking into account at the stage of developing environmentally friendly high-speed vehicle. The next generation supersonic transport is fundamentally different from subsonic medium- and long-range aircraft in a breakthrough feature: it ensures a long-term supersonic cruise flight in the range of Mach numbers 1.6–1.8. For a successful implementation of the civil supersonic project at a new stage of technology development, it is necessary to solve a number of scientific and technical problems. In the present paper some of them are under discussion. In particular, the addressed issues are as follows: how to achieve a high level of aerodynamic efficiency at supersonic cruise along with low sonic boom; how to fit into low noise requirements at take-off and landing modes of supersonic aircraft with a moderate-bypass-ratio engine; how to implement a wing with high load factor and fuselage with long nose part to fit into required airframe structural stiffness, etc. Also, new technology approaches in multidisciplinary optimization task to achieve aircraft design goals of low drag, low boom and low noise characteristics are discussed. A few examples of optimal supersonic aircraft configurations along with passenger cabin are given. The paper was presented at the X-th IAA Symposium "Space Flight Safety" in St.-Petersburg on June 2023. • Preliminary aerodynamic design,. • Advanced supersonic transport,. • Sonic boom,. • Multifactor optimization,. • Aerodynamic configuration. [ABSTRACT FROM AUTHOR]

This paper analyzes the change in the heliocentric orbital elements due to the implementation of aerogravity and powered aerogravity assist maneuvers. Three cost functions were selected to maximize the planetocentric latitude, longitude, and velocity at the end of the atmospheric flight. An optimal control problem was solved to guide a spaceplane passing above the three inner planets: Venus, Earth, and Mars. The planets were selected because they have been relevant for gravity assists, and previous studies suggest that taking advantage of their atmospheres could increase the effects of the close approach. The research aims to analyze the variation in the heliocentric orbital elements of the trajectories after performing optimal aerogravity-assists and powered aerogravity-assisted maneuvers at high atmospheric altitudes. The trajectories were calculated via nonlinear programming to find the history of the control variables: angle of attack, bank angle, and thrust. Low-fidelity gravity and atmosphere models were implemented to evaluate this proof of concept. The optimization converged for all cases, and results show that maximization of the longitude increases the duration of the atmospheric flight, increasing the turning angle by almost twice the value of a single gravity assist. Furthermore, it reduces the energy and semimajor axis in half for approach angles between 20 and 90°, saving more than 10 km/s on propulsion. Maximizing the velocity with propulsion increases the semimajor axis by more than 1.1 times the semimajor axis of the respective planet, and the optimal latitude presents a change in inclination of more than 1.0°. This maneuver could be applied to collect atmospheric samples or increase the coverage above the planet of interest. Those kinds of applications are described at the end of this paper, including a brief discussion on the maturity of this technology. • Optimal control in aerogravity assists maneuvers at high atmospheric altitudes. • The maneuvers increase the values of the heliocentric orbital elements. • Multiple advantages over the gravity assist are observed. • Spaceplanes are feasible for performing those maneuvers. • Atmospheric passages at high altitudes reduce the risks associated with low altitudes. [ABSTRACT FROM AUTHOR]

Autonomous spacecraft relative navigation via monocular images became a hot topic in the past few years and, recently, received a further push thanks to the constantly growing field of artificial neural networks and the publication of several spaceborne image datasets. Despite the proliferation of spacecraft relative-state initialization algorithms developed, most architectures adopt computationally expensive solutions relying on convolutional neural networks (CNNs) that provide accurate output at the cost of a high computational burden that seems unfeasible for current spaceborne hardware. The paper addresses this issue by proposing a novel pose initialization algorithm based on lightweight CNNs. Inspired by previous state-of-the-art algorithms, the developed architecture leverages a fast and accurate target detection CNN followed by a line segment detection CNN capable of running with low inference time on mobile devices. The line segments and their junctions are grouped into complex geometrical groups, reducing the solution search space, and subsequently, they are adopted to extract the final pose estimate. As a main outcome, the analyses demonstrate that the lightweight architecture developed scores high accuracy in the pose estimation task, with a mean estimation error of less than 10 cm in translation and 2.5°in rotation. The baseline algorithm scores a mean SLAB error of 0.04552 with a standard deviation of 0.22972 in the test dataset. Detailed analyses demonstrate that the uncertainties on the overall pose score are driven mainly by errors in the relative attitude, which gives the highest contribution to the pose error metric adopted. The analyses on the error distributions point out that the uncertainties on the estimated relative position are higher in the camera boresight axis direction. Concerning the relative attitude, the algorithm proposed has higher uncertainties in estimating directions of the target x and y axes due to ambiguities related to the target geometry. Notably, the target detection CNN trained in this work outperforms the previous top scores in the benchmark dataset. The performances of the proposed algorithm have been investigated further by analyzing the effects on the accuracy due to the relative distance and the presence of background in the images. Lastly, the paper delves into the possibility of adopting a sub-portion of the 2D-to-3D match matrix made by the most complex perceptual groups identified that positively affects the overall run-time, pointing out the performances in terms of accuracy of the estimates and providing a comparison of both the baseline and the reduced match matrix versions against state-of-the-art algorithms concerning relative position and attitude errors and solution availability, highlighting the high accuracy and solution availability of the proposed architectures. • New relative pose initialization algorithm based on the detection of line segments. • Detailed literature review and description of available spaceborne image datasets. • Improvement of state of the art in spacecraft detection using lightweight CNN. • Assessment of accuracy of the proposed architecture and error sources analyses. • Proven feasibility of the approach via comparison with other architectures. [ABSTRACT FROM AUTHOR]

Planetary defense is an emerging field among the aerospace community. As the rate of near-Earth object (NEO) discoveries increases, so too does the probability that we will discover a NEO on a collision course with Earth. This paper documents a planetary defense mission campaign design to defend against the hypothetical Earth impactor 2023 PDC. At its 90th percentile mass estimate, the asteroid could have a diameter of approximately 1120 m. The asteroid makes no close approaches until Earth impact and has the potential to affect over a billion people, depending on its actual size. This paper describes a mission campaign consisting of both reconnaissance and mitigation missions that will be deployed to neutralize the threat posed by the asteroid. The reconnaissance orbiter named THEO (Terrestrial Hazard Exploration Orbiter) will survey the asteroid to clarify its morphological, gravitational, and dynamical properties. This information will be used to inform the design of the subsequent mitigation mission. THEO will be designed to have a mission lifetime of 12 years which will enable it to observe the mitigation mission and confirm its success. Two 21-day launch periods have been identified that minimize post-launch Δ V requirements and include current launch vehicle capabilities. For the purposes of this paper, the mitigation mission has been designed assuming a geocenter impact for the 90th percentile mass estimate of the asteroid because a precise impact location and an accurate mass estimate will not be available until THEO reconnoiters 2023 PDC. The mitigation mission, named MUFN (Mitigation Using a Fission Nuclear device), will send three spacecraft to rendezvous with 2023 PDC and alter the asteroid's trajectory prior to Earth impact through a series of standoff nuclear detonations. The MUFN spacecraft are identical in design, each employing a hypergolic bi-propellant thruster. Each MUFN spacecraft launches on a SpaceX Falcon Heavy Expendable launch vehicle between 2028 and 2030 and carries four 340 kt nuclear explosive device (NED) equipped orbiters, named BELA-Os ("Bull's Eye, Later Asteroid" Orbiters). The BELA-Os are designed to sequentially deploy, maneuver to an optimal standoff distance anti-parallel to the asteroid velocity vector, and detonate the NED to impart a change-in-velocity to the asteroid. After arrival at 2023 PDC, each MUFN spacecraft will station-keep above the asteroid. The NED-equipped BELA-Os will deploy from the MUFN spacecraft and maneuver 180 degrees out of phase from MUFN to the 'front' of the asteroid. Once there, each BELA-O will detonate and impart 2.32 mm/s of velocity change to the asteroid. After the ablation debris has cleared, MUFN will complete a single orbit around the asteroid to survey the blast site and assess the effectiveness of the detonation. After MUFN has completed its survey, it will release another BELA-O. This process will repeat until all BELA-Os have been released by all MUFN spacecraft. All denotations will occur between 2029 and 2031. After all BELA-Os have detonated their NED payloads near the 2023 PDC asteroid to deflect its orbit, the asteroid will safely fly by Earth at a minimum altitude of 3500 km instead of colliding on October 22, 2036. • A planetary defense mission campaign design is presented to defend the Earth against the hypothetical Earth impactor 2023 PDC. • A reconnaissance mission design with a 12-year lifetime is presented. • A mitigation mission design using nuclear devices is presented. • The use of nuclear devices is favorable due to the large size of the hypothetical impactor. [ABSTRACT FROM AUTHOR]

Leveraging advancements in micro-scale technology, we propose a fleet of autonomous, low-cost, small solar sails for interplanetary exploration. The Berkeley Low-cost Interplanetary Solar Sail (BLISS) project aims to utilize small-scale technologies to create a fleet of tiny interplanetary spacecraft for rapid, low-cost exploration of the inner solar system. This paper describes the hardware required to build a ∼10 g spacecraft using a 1 m2 solar sail steered by micro-electromechanical systems (MEMS) inchworm actuators. The trajectory control to a NEO, here 101955 Bennu, is detailed along with the low-level actuation control of the solar sail and the specifications of proposed onboard communication and computation. Two other applications are also shortly considered: sample return from dozens of Jupiter-family comets and cometary nuclei imaging. The paper concludes by discussing the fundamental scaling limits and future directions for steerable autonomous miniature solar sails with onboard custom computers and sensors. • An autonomous 10−2 kg spacecraft to navigate the solar system is proposed. • Low-cost spacecraft composed of mostly off-the-shelf parts to return images of NEOs. • Propelled by a 1 m2 solar sail controlled by micro-electromechanical systems motors. • Autonomous trajectory control to a NEO, communication, and computation is detailed. [ABSTRACT FROM AUTHOR]

In the past, space trajectory optimisation was limited to the design of space transfers to single destinations. More recent mission concepts however present the challenge to target multiple destinations; be it for science, exploration or even exploitation. This paper focuses on space trajectories that aim to visit multiple main-belt asteroids with one single spacecraft. The trajectory design of these missions is complicated by the fact that the asteroid sequence is not known a priori, but it is the objective of the optimisation problem. Usually, these problems are tackled as global optimisation problems under the formulation of Mixed-Integer Non-Linear Programming, on which the design variables assume both continuous and discrete values. However, beyond the aim of finding the global optimum, mission designers are usually interested in providing a wide range of mission design options reflecting the multi-modality of the problem. The paper describes a multi-fidelity design pipeline that enables a consistent exploration of all feasible asteroid tours with a given set of mission boundary conditions (e.g., launch window, mission duration, Δ v). The consistency of the exploration is ensured by a deterministic search scheme which retains the n -best paths at each stage of a dynamic programming process. Bellman's principle of optimality is made applicable by transcribing the search space into a series of discretized events defined by asteroid index and fly-by epoch, as well as ensuring optimal substructure property with the choice of transfer model and design variables. The method is deployed to explore asteroid tour opportunities visiting 10 or more main belt objects and mission boundary conditions compatible with medium or discovery class missions. 79 million such trajectories are found with launch date in 2037. • A multi-fidelity pipeline is proposed to deal with the challenge of trajectories that fly-by many main belt asteroids. • A novel transcription of the mixed-integer problem allows for an efficient pruning of the search space. • A dynamic programming-based exploration is proposed, keeping the n best solutions for expansion at each unique node. • Novel mission scenarios are presented in the framework of ESA's M-class mission envelope. [ABSTRACT FROM AUTHOR]

Our understanding of the formation and interaction processes of astronomical structures in the Universe and in galaxies has significantly expanded due to recent progress in improving the quality of observations, brought about by the emergence of new telescopes. Such processes are associated with the release of a significant amount of energy and may influence our vital activities. One example of energy flow release is magnetic storms, which are associated with processes occurring on the Sun. Such storms lead to disruptions of communication facilities, impact people, and can cause satellite and aircraft failures. There are objects in both near and deep space that are much more powerful than our star. In such objects, enormous energy flows are generated and released in the form of space jets. These energy flows occur in quasars, black holes, protostars, and during supernova explosions of types Ia, Ib, Ic, and II. During a supernova explosion, an immense amount of energy is released in the form of energy and high-energy particles. Such processes can have a significant impact on the safety of space flights. In this paper, the processes of formation of stellar systems within a turbulent interstellar medium are considered. This process is the result of complex processes that occur in the interstellar medium and denser gas formations – molecular clouds. At the same time, there are nonlinear interactions between turbulence and gravity, collisions between molecular clouds, interactions with shock waves, and so on. Supersonic turbulence is one of the main reasons for the formation of dynamic prestellar structures. The evolution of superdense substances formation begins when they gather in turbulent flows or are formed by supersonic collisions between molecular clouds. This process continues until these areas reach prestellar density. Depending on various factors, these superdense formations can either collapse and form new star systems or disintegrate, returning the substance to the interstellar medium. Until recently, the prevailing opinion was that a significant portions of molecular clouds were not gravitationally bound. Conclusions about this were made based on the virial theorem, which expresses the relationship between gravitational and kinetic energy. However, recent studies have refuted this. In this paper, the results of a simulation carried out to study the collision of large ellipsoidal and spherical molecular clouds in a three-dimensional setting. The simulation was conducted on high-resolution grids and utilized the adaptive mesh refinement (AMR) method. • Morphological analysis of supersonic turbulence arising from the collision of spherical and prolate molecular clouds. • Modeling processes leading to the possible formation of new stellar systems. • Study of complex shock wave loading leading to the emergence of supersonic turbulence, Kelvin-Helmholtz instability and the formation of clumps. • Comparison of Chandrasekhar's analytical and numerical solution for an prolate molecular cloud. [ABSTRACT FROM AUTHOR]

Balloon-based exploratory platforms, or aerobots, are a budding planetary exploration technology for Venus to collect scientific data in-situ while keeping a safe distance from the highly inhospitable surface of Earth's "evil twin". Such a mission is expected to deploy and inflate the aerobot during the system's initial descent through the Venusian atmosphere, rather than launched from the surface. The analysis of the entry, descent and float (EDF) sequence of the aerobot is accordingly a delicate and fundamental part of the aerobot design, and generally where the envelope undergoes the largest loads. The inflation phase in particular plays an essential role, especially for variable-altitude aerobots that can have multiple-envelopes with different inflation requirements. This paper addresses a collection of preliminary considerations about the inflation phases of a Venus aerobot – starting from the end of deployment to completion of reaching an equilibrium altitude – where the nominal mission flight profile would then begin. We include two different inflation sequences, highlighting issues and threats, utilizing a simulation of the inflation dynamics of the aerobot conducted in MATLAB. The paper also presents results from a set of rapid inflation experiments, conducted on a subscale prototype of the aerobot, aimed at an initial assessment of feasibility and response of the materials involved. • Inflation sequence design for a variable-altitude aerobot at cloud-level on Venus. • Sequence modification to reduce peak loads on the aerobot. • Constraint-driven analysis to evaluate the impact of the model's initial parameters. • Laboratory demonstration of the inflation utilizing a subscale aerobot prototype. [ABSTRACT FROM AUTHOR]

Virtual reality (VR) and augmented reality (AR) are rapidly developing technologies that may have important healthcare-related applications, including the possible evaluation of the health and performance of astronauts. Advances in the fields of VR/AR and space exploration have independently progressed; however, over the last decade synergies and overlapping applications have emerged between spaceflight and extended reality. This paper is intended to be a resource to update the medical community on the current and future integration of VR and AR into space exploration, particularly as this integration relates to astronaut health. We highlight how these technologies are used in terrestrial healthcare with specific attention to visual function. Furthermore, we elaborate on emerging VR applications being developed for future spaceflight. The paper discusses the current state of the art, such as the Apple Vision Pro headset, and outlines key challenges like mass and volume constraints, power limitations, and the heightened cognitive and physical demands of space travel. It also addresses the potential of VR/AR technologies to mitigate these issues through efficiency, stress reduction, and exercise motivation. Finally, we emphasize the benefits that these technologies provide to overcome possible barriers to long duration spaceflight and manned exploration missions, including spaceflight associated neuro-ocular syndrome (SANS). • Virtual reality (VR) and augmented reality (AR) are rapidly developing technologies that may have important healthcare-related applications, including the possible evaluation of the health and performance of astronauts. • This paper is intended to be a resource to update the medical community on the current and future integration of VR and AR into space exploration, particularly as this integration relates to astronaut health. • We highlight how these technologies are used in terrestrial healthcare with specific attention to visual function. Furthermore, we elaborate on emerging VR applications being developed for future spaceflight. [ABSTRACT FROM AUTHOR]

Recent startups and university space activities extend to cis -lunar and near-Earth interplanetary field. One of the biggest hurdles in those missions is in the orbit determination based on the range measurements, while the radio transmission and reception capability becomes within those small organizations' facilities. Also, about the navigation in cis -lunar space, conventional GNSS (Global Navigation Satellite System) real time positioning is hardly available and some alternative methods have been sought. The most straight forward approach is to be supported by the space agencies' facilities for the range measurement with the transponders onboard the probes. However, it depends on the facilities availability and also on the relatively expensive, old and analogue transponders. This paper presents a new method deriving from the GNSS based devices. It does not require the transponders but inexpensive radio circuits easily built. And both the probes and the ground stations have only to be equipped with the common and simple inexpensive devices. The method enables the small organizations to perform autonomous operation of the probes in a distance beyond the moon's orbit. The paper first presents the latest test data obtained during the development using the Software Defined Radio (SDR) boards. • Low-Cost Ranging Measurement for Deep Space beyond Cis-Lunar Region. • Real Test Results using Software Defined Radio Boards, Hardware. • Semi-Real-Time and Simultaneous Range Measurement and Clock Synchronization. • Absolute Clock Relay Capability via A Pair of Spacecraft. [ABSTRACT FROM AUTHOR]

Space-based data provide essential information on the past, current, and future state of climate change. Satellite data widely contribute to scientific research conducted by recognized institutions such as the IPCC, which underpin most climate policies in Europe. In addition, half of the Essential Climate Variables (ECVs), which are necessary to monitor climate change, directly come from satellite data. However, the role of space is not always accurately reflected in Europe's climate policies, with only a few direct references to ESA's or EU's space programmes and data. Against this background, this paper aims to provide an in-depth assessment on the actual role played by space-based data in the definition and implementation of European climate policies. Based on the outcomes of a year-long research conducted in collaboration with the European Space Agency, the paper will first analysed the use and references to space-based data in over 500 climate policy documents of the 22 ESA Member States. Subsequently, it will address some potential blocking points to the use of space in the climate policymaking process in the analysed European countries. By doing so, the paper will also look into aspects related to science communication, the science-to-policy divide as well as the climate governance in ESA Member States. Finally, this paper will provide recommendations to increase the use of space-based data in the climate policymaking process. • Climate policies in ESA Member States show that references to Essential Climate Variables are abundant. • The origin of the data or the systems used to generate variables are not referenced despite half of variables being only measurable from space. • All ESA Member States reference space in their climate policies. • 17 ESA Member States have reference ESA in their climate policies. • A blocking point to the integration of space-based data in climate policies pertain to the science to policy divide. [ABSTRACT FROM AUTHOR]