Search

There is a growing realization on the part of lawmakers, policy experts and practitioners that cyber risks and threat vectors constitute a critical and strategic challenge for all outer space operations. Stakeholders' professional attention and resources are beginning to focus on measures for mitigating these risks and threats. Space-based activities depend on the reliable functionality of computer systems and the wireless communications connecting them through use of the electromagnetic spectrum resource – all of which constitute parts of the infrastructure of cyberspace both on Earth and in space. Thus, cyber risks and vulnerabilities arise in all aspects of outer space operations: satellite launches and their orbital telemetry, tracking, and control (TTC); operational communications (space-to-space, earth-to-space and space-to-earth); and other spectrum-dependent space transmissions. Additional cyber vulnerabilities are present in the terrestrial systems supporting space operations, including launch systems and supply chains for dedicated equipment and services. The mitigation of such cyber threats and vulnerabilities in the more traditional terrestrial context is a rapidly developing area of law and policy aiming to counter the activities of malicious cyber actors whether they are criminals, States or hostile ideological groups. On the one hand, the threat mitigation aspects of cybersecurity have by no means been resolved, as hostile actors use increasingly sophisticated means to achieve their ends. On the other, progress has been made with respect to measures and best practices recognized as critical to boosting cybersecurity for space activities. This article focuses on one of these: information sharing (IS) among trusted participants in a secure platform designed for sharing actionable cybersecurity data. The primary goal of such sharing is to reduce the informational asymmetry between the attacker and the target, whether the latter is a State, private company, or other entity. Cyber attackers leverage their knowledge of the target's vulnerabilities, and only need to be right on one occasion to cause operational, financial, and/or reputational damage that may impair the long-term functionality of the targeted entity. Yet when a group of potential targets can reliably share data on cyber risks and measures to counter them, the informational asymmetry is reduced, and cybersecurity improved. Several conditions exist for optimizing IS which provides actionable information, and for ensuring that trust is developed and maintained among IS participants. Finally, in the context of outer space-related cybersecurity threats, the paper argues that IS proves critical because of two elements: the aggravated asymmetry of cyber threat vectors in outer space, and their distinctive space-related characteristics. • There is a growing realization that cyber risks and threat vectors constitute a critical and strategic challenge for all outer space operations. • Stakeholders' professional attention and resources are focusing on measures for mitigating these risks and threats. • Space-based activities depend on the reliability of computer systems and their wireless connection, – which constitute the infrastructure of cyberspaceon Earth and in space. • Cyber risks and vulnerabilities arise in all aspects of outer space operations, andadditional. • Cyber vulnerabilities are present in the terrestrial systems supporting them. • Several conditions exist for optimizing IS which is actionable, and for ensuring that trust is maintained among IS participants. • Several conditions exist for optimizing IS which is actionable, and for ensuring that trust is maintained among IS participants. • Information sharing for outer space cybersecurity is critical because of the aggravated asymmetry of cyber threat vectors in space and their distinctive space-related characteristics. [ABSTRACT FROM AUTHOR]

The thin-walled tubular deployable composite boom (DCB) can realize folding and deploying functions, and it has a good application prospect in space field. This paper investigates the folding behaviour of the tubular DCB by analytical modeling. Based on the Archimedes' helix, the geometrical model of the DCB was established. By combining equilibrium equation and energy principle, an analytical model to predicted the folding moment versus rotational displacement of the DCB was presented. The failure indices in the equal-sense and opposite-sense folding processes were calculated utilizing Tsai‐Hill and maximum stress criteria. Analytical results agreed well with experimental and numerical results. At last, the influence of geometric parameters (i.e., radius, central angle and thickness of cross-section) on the folding behaviour of the DCB was further studied using the analytical model. [Display omitted] • An analytical model is presented to predict the folding behaviour of the deployable composite boom (DCB). • Failure indices are calculated using two criteria in equal-sense and opposite-sense folding processes. • The analytical model is validated against experimental and numerical results, showing good prediction accuracy. • Geometric parameters (radius, central angle, and thickness) significantly affect the folding behavior of the DCB. [ABSTRACT FROM AUTHOR]

Thin-walled tubular deployable composite boom (DCB) has gained significant attention due to its lightweight, simple structure and high packaging efficiency. The experimental and numerical simulation methods were employed in this paper to investigate the folding behavior of the DCB. Using T700/epoxy unidirectional reinforced prepreg as raw material, DCB specimens were prepared by the vacuum bag method. To complete the folding experiments of DCB specimens, a folding mechanism was designed and manufactured. Folding experiments of DCB specimens were conducted and folding moment versus rotational displacement curves were measured. In addition, a Finite Element Model (FEM) was established to predict the folding behavior of the DCB. Different failure criteria were considered in the numerical analysis, including the Tsai‐Hill criterion and maximum stress criterion. Prediction results using the FEM were compared with experimental results, and both sets of results were in good agreement. It is shown that the DCB can achieve the folding function without failure. The research results in this paper are of great significance to the practical engineering application of the DCB. [Display omitted] • Six tubular deployable composite boom specimens and a precise folding mechanism were designed and prepared. • Folding experiments were conducted to measure the folding moment and rotational displacement curves. • A finite element model for predicting the folding behavior of the tubular deployable composite boom was established. [ABSTRACT FROM AUTHOR]

Space Resource Utilization could reduce the mass, cost, and risks of future space activities. The many nonlinear and compounding effects mean that seemingly insignificant concept of operations choices can invalidate the conclusions of sophisticated analyses. Here we successively relax self-imposed constraints in a recent publication to demonstrate that the proximity of a market and the benefits of aerobraking cannot be omitted. • Earth Moon proximity allows a high cadence of space resource transportation. • Higher cadence allows costs to be amortized over a higher mass of delivered product. • A powerful economic advantage of the proximity of lunar resources is demonstrated. • Modifying a published model, we can show orders of magnitude reduction in costs. • Fair comparison of resource locations must incorporate multiple missions over time. [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]

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]

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]

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]

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]

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]

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]

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]

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]

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]

DART, the Double Asteroid Redirection Test, successfully impacted asteroid Dimorphos on September 26, 2022, becoming the first mission to demonstrate asteroid deflection. Shared live via a NASA broadcast, over a million concurrent viewers around the world watched as the DART spacecraft streamed images to Earth up to the final second before its impact with Dimorphos. In this paper, we share the final phases of the DART encounter and the latest results from the DART Investigation Team's analysis of ground and space-based data. This includes measurements on the amount of deflection that DART imparted on Dimorphos, a determination of the momentum transfer enhancement factor, and results on understanding the geology and surface characteristics of the impact site. The success of the DART mission paves the future for a bold international Planetary Defense program. We also share ideas on bolstering international coordination, developing new technologies that can mitigate different types of asteroid threats, and closing identified gaps in our overall preparedness. [ABSTRACT FROM AUTHOR]

This paper considers the problem of optimizing minimum-time and minimum-propellant many-revolution low-thrust geocentric orbit-raising maneuvers to an arbitrary target. The proposed direct single-shooting technique utilizes the regularized h , e element set as well as analytical models for ephemeris, shadow geometry, and thrust magnitude for ease of the optimization procedure. Third body and J 2 effects are included for higher fidelity of the computed solution. The gradient information supplied through adjoint sensitivities facilitates the convergence of the numerical optimization scheme. We also propose a new automated, fast and robust initial guess generation methodology by solving a sequence of unconstrained optimization orbit-raising sub-problems. The performance of the developed algorithm is demonstrated for orbit-raising to a variety of targets: the geosynchronous equatorial orbit, the Molniya orbit, and a near-rectilinear halo orbit. • A new low-thrust trajectory algorithm is developed using adjoint sensitivities. • The sequentially generated initial guesses are fast and robust to calculate. • An end-to-end optimization formulation computes minimum-time and minimum-fuel transfers. [ABSTRACT FROM AUTHOR]

Current interest in spacecraft constellations for government and commercial uses continues to drive advancement in sophistication, capability, and mission complexity. As these systems proliferate, they require increased levels of automation to properly orchestrate in achieving the mission objectives. This paper develops a constellation planning capability for cross-link enabled Geospatial Intelligence missions. The planning for constellation coordination is performed using a two-level network flow approach, where the first level creates a communication plan, and the second layer determines the Earth sensing activities. A network flow method is used to simultaneously obtain a schedule for each unique satellite within the constellation and provides an optimal operational schedule for the entire cooperative system. [Display omitted] • Network flow optimizes the constellation with individual satellite operations. • Resulting satellite plans exist within the constrained feasible resource space. • Decoupling variables improves planning and data transport performance. [ABSTRACT FROM AUTHOR]

Satellite formation plays an important role in space missions such as on-orbit services and distributed space facilities. In these missions, hovering in a bounded range is one of the core tasks, but hovering technologies for the chief satellite in an elliptical orbit have been rarely investigated. Given the boundary constraints of relative motion between the chief and deputy satellites, this paper proposes two models for the design and control of hovering formation near an elliptical reference orbit through impulse: (1) the revisiting hovering formation model, which forces the deputy satellite to revisit the same relative position and maintains the formation via impulses; (2) the sequential-impulse-control hovering formation model, which uses a sequence of impulses to ensure the deputy satellite flies within the given boundary. Unlike the revisiting hovering formation model, this model does not require a fixed revisiting point, so the satellite's trajectory is more flexible. The geometrical properties, formation maintenance strategy, and impulse-control solution for both models are investigated in detail. Finally, numerical simulations are conducted to verify the advantages of the proposed methods over other conventional methods, suggesting their wide applications in practical space missions. • Two impulsive-control models are proposed for bounded hovering formation in elliptical orbit. • The proposed methods support in-plane, out-of-plane, and three-dimension hovering formations. • The in-plane and out-of-plane controls are decoupled to decrease control complexity and difficulty. • Different factors on the fuel cost are analyzed to provide a reference for practical engineering. [ABSTRACT FROM AUTHOR]