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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]

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]

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]

Konstantin Tsiolkovsky famously stated that while the Earth serves as the birthplace of humanity, it is not a place where mankind can indefinitely remain. Perhaps during that period, the prospect of exploring the mysteries of outer space appeared to be an unattainable aspiration. However, in the present day, there are no longer any limitations, not even the sky, since human ingenuity has facilitated access to outer space for humanity. This access is not just for the purposes of research and exploration but also for economic endeavours. Until now, the commercial utilisation of outer space has advanced at a very sluggish rate. However, firms including SpaceX, Orion Span, Virgin Galactic, and Blue Origin have achieved significant advancements in the growth of the space industry. The revenue generated by various space-related endeavours has experienced a significant 73% increase over the last ten years. The global space economy, estimated to be valued at USD one trillion in the coming years, is primarily driven by commercial activities. This presents a formidable challenge to the existing national and international taxation systems. Similar to the open seas, space is also considered res communis omnium, meaning it belongs to the entire community, and presents comparable taxing challenges with potentially uncertain solutions. The three fundamental elements of every taxation regulation, such as the Organisation for Economic Co-operation and Development or the United Nations Model Double Taxation Convention, are the taxpayer's place of residence, the origin of their income, and the methods by which they generate their money. The current tax system does not have the necessary concepts and provisions to adapt to the rapid advancements in commercial space technology. This paper examines the legal issues surrounding commercial activities conducted in space, including the nature and handling of the income generated in various legal systems. It also addresses concerns such as tax avoidance and excessive taxation, emphasising the necessity for a globally coordinated approach to effectively tax commercial activities in space. • Human ingenuity has facilitated access to outer space for humanity, for the purposes of research and exploration and also for economic endeavours. • Space-related endeavours has experienced a significant 73% increase over the last ten years. • Commercial activities conducted in space require concerted action at international level. [ABSTRACT FROM AUTHOR]

In contrast to the rising technological applications of China's satellite systems, their applications and impacts in modern Chinese Society, particularly in the legal and administrative forums, have largely escaped under the radar. Specifically, the use of satellite outputs can now be found within a number of administrative and legal activities within China - including policy and rule-making, administrative investigations, and the production of court evidence. At the same time however, outdated legal practices, ambiguous status, and lack of guidance has prevented China's legal system from fully capitalizing on the advantages of satellite technologies. Therefore, to fill these legislative gaps, this review provides a comprehensive overview on the applications of satellite technologies within China's legal system, and provide important insights as to their current developments, legal framework, and legal issues. Additionally, having identified the aforementioned issues, this paper offers policy recommendations for the purposes of legislation. • Satellite applications in China's policy and rule-making processes. • Investigatory applications of satellite technologies within China's administrative organs. • Emerging evidentiary applications of satellite technologies in Chinese court systems. [ABSTRACT FROM AUTHOR]

A comprehensive approach to creating a control system for the shape of the radio-reflecting surface of a large-size space reflector is considered. Large-aperture space antennas are used for satellite communications and space exploration. The design of these structures is one of the most promising areas. Ensuring the specified characteristics of the reflector is directly related to maintaining the specified shape of the radio-reflective netting. Developing a control system for the shape of the active antenna surface is an urgent task. Optimizing this process makes it possible to accomplish the mission during the operational period effectively. The paper evaluates the effects of perturbations on a large-scale structure operating in outer space. A method for adjusting the geometry of the frontal network is proposed. This allows for improved communication between spacecraft and ground stations, which is very important for the safety of space flights. The conductive parts of the reflector are considered to transmit energy and control signals to the actuators. The application results of various control algorithms of the executive device making it possible to correct the netting considering the minimization of energy costs and oscillations of the structure are shown. According to the simulation results, the optimal algorithm of the target criteria hierarchy showed greater stability in terms of convergence in the presence of measurement noise and disturbances, which allows it to be applied in real time. This algorithm is supposed to be used in the future when creating a prototype of the system for controlling the shape of the radio-reflective surface of a large-size space reflector using its current-conducting structural parts. • The design of a system for setting up and maintaining a radio-reflective network fabric has been developed. • Algorithms are developed to minimize energy when setting up a radio-reflective mesh. • The optimal algorithm of the control allows solving the problem in real time. • The method for transmitting energy and control signals through the use of conductive structural parts of the reflector is proposed. [ABSTRACT FROM AUTHOR]

With the dramatic increase in the number of resident space objects, the risk of collision between space objects and debris is growing. To protect spacecraft from such collisions, it is crucial to determine the generation patterns of space debris and analyze the evolution of debris clouds. In this paper, a continuous distribution formula is derived for each parameter of the breakup debris based on the China Aerodynamics Research and Development Center spacecraft breakup model (CARDC-SBM). Using CARDC-SBM, we theoretically establish a normal distribution for the separation velocity of the breakup debris in relation to the breakup mass center. Based on this foundation, the mean and standard deviation of the debris separation velocity distribution are empirically modeled using polynomial and exponential functions. Consequently, an empirical formula with a broad range of applicability is derived, which significantly reduces the computation time of debris cloud simulations. The initial velocity distribution given by this empirical formula is employed as the initial condition for the evolution of debris clouds from the perspective of a boundary value problem. Simulations using CARDC-SBM and subsequent analysis of the debris cloud evolution resulting from a breakup event reveal that the distribution of each parameter is determined solely by the relative collision velocity and the spacecraft's nominal density in the impact breakup scenario. Our results show that the empirical formula for the debris cloud distribution provides good agreement with Monte Carlo simulations. • A continuous distribution formula is derived for each parameter of the impact breakup debris based on CARDC-SBM. • We point out and validate that normal distribution can characterize the velocity of debris produced by CARDC-SBM. • Empirical formulas are developed to compute the debris separation velocity distribution with low computational burden. • The proposed method can provide initial condition for space debris clouds evolution. [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]

Lunisolar resonances shape the dynamics in the MEO region and have the effect of increasing an object's eccentricity possibly up to a value so high that the orbit's perigee reaches the atmosphere and friction determines the re-entry. The focus of this paper is the eccentricity growth phenomenon observed for navigation satellites. In particular, starting from a secular Hamiltonian model, analytical estimates for the maximum eccentricity reachable along an inclination-only dependent lunisolar resonance are derived as a function of the initial inclination. Also, trajectories that minimize an object's lifetime are studied, and analytical estimates for such a minimum lifetime are provided as a function of the initial eccentricity of the object. In addition, lifetime cartography maps are presented in the plane of the argument of perigee and the longitude of the ascending node. • Analytical estimates for the maximum eccentricity along lunisolar resonances. • Analytic formulae of minimum satellite lifetime at GPS and Galileo altitudes. • Litetime maps at GPS and Galileo altitudes in (ω)-(Ω) plane. [ABSTRACT FROM AUTHOR]

Collisions involving space objects pose significant threats to space sustainability. Within these collisions, catastrophic collisions which generate a larger amount of debris are particularly harmful. The quantity and characteristics of fragments generated from a collision are usually modeled using the NASA standard breakup model (SBM). However, on several occasions, the NASA SBM has incorrectly estimated the effects of collisions, including the Yunhai 1-02 breakup event. Using outer probability measure theory and information theory, we develop a correction methodology to estimate the Yunhai 1-02 event based on NASA SBM. This methodology successfully passes a statistical test using historical data. After correction, the characteristic diameter and relative velocity distributions of fragments generated from this event exhibit greater similarity with observations. The work in this paper facilitates the analysis of the properties of the debris cloud resulting from collisions with limited information, improving the predictions of future collisions. [Display omitted] • We condense and analyze the historical collision data. • We obtain a corrected parameter to adjust the SBM result, and this adjusted result passes the statistical test. • We improve the precision of characteristic diameter and relative velocity of fragments generated from Yunhai event. • This work is important to estimate the effect of collisions with limited information. [ABSTRACT FROM AUTHOR]

Accurately measuring real economic output in the space economy is made difficult by the rapid increase in capabilities and decrease in prices of launch and satellite technologies achieved over the past two decades. Nominal measures of output in space will tend to underestimate real growth if customers are paying lower prices for better services over time. In this paper, we use price indexes that apply techniques to the space economy that have been used for decades to adjust nominal measures of output in sectors such as information technology, including matched-model and explicit hedonic quality adjustment. We find that adjusting for price and quality changes in the space economy has substantial effects on estimated growth in economic output, especially across its sectors. Price increases over time in the space information sector, which is dominated by direct-to-home (DTH) satellite television, mean that real growth has been slower than nominal growth since at least 2012, the typical pattern in industries not undergoing rapid technological change. In sharp contrast, rapid price decreases and quality increases in other areas—especially launch and manufacturing of satellite, earth observation, and positioning, navigation, and timing (PNT) equipment—have meant that nominal growth rates substantially understate real growth rates in those industries. Given the central importance of space to our modern economy, having accurate information on growth in the space economy is vital for efficiently allocating the substantial private and public investment being devoted to its development. • Adjusting space output for price and quality changes impacts economic growth rates. • Growth in the US space sector masks underlying heterogeneity across its subsectors. • US space manufacturing products exhibit rapid price decreases and quality increases. • US space information product prices are rising even as overall output is falling. [ABSTRACT FROM AUTHOR]

Advancements in bulk metallic glass thermoplastic forming unleash the potential to fabricate microscale metallic features with unparalleled precision and durability, i.e. microspine gripping structures for mobility platforms on planetary or in-space robots. Four designs of grippers were fabricated and rigorously tested. Coefficient of friction on four test surfaces were measured and compared to SiC paper and Velcro. The hardness and elasticity of the bulk metallic glasses make them a game-changer in the field, offering a superior alternative to conventional nonmetallic grippers. • Microspine grippers were manufactured via thermoplastic forming of bulk metallic glasses using silicon-on-insulator molds. • The microspines excelled on textured surfaces, where sharp points anchor on to pores or softer regions. • Mechanical behavior of microspines were studied using nanoindentor-based micro-compression and computer simulations. [ABSTRACT FROM AUTHOR]

Two-phase flows under microgravity conditions play a crucial role in Space technology and science. Studying the behavior of dispersed droplets suspended in a gas flow and their influence on the characteristics of the carrier flow is the main task of computational modeling of turbulent two-phase flows. To understand the features and predict the behavior of such flows, along with experimental studies, it is necessary to develop mathematical models, which adequately describe all the basic physical and chemical processes in a gas-droplet environment at different scales. In this paper we obtain solutions for the problem of two droplets collision, using solutions for cumulative jets, and propose droplet collision outcome criterion. This research will focus on the process of two interacting droplets coalescence into one and coalescence with subsequent separation into two or more droplets. A semi-analytical method has been developed for application of the solution of the cumulative jet problem to the problem of droplet collision. A differential equation has been obtained that describes the following possible outcomes of droplet collision: coalescence and coalescence with subsequent separation. A criterion for determining the outcome of on-head droplet collision has been obtained. In accordance with the developed criterion, the number and size of formed droplets are obtained. The effect of collision on moving droplet evaporation in a heated atmosphere was regarded. • Solutions for the problem of two droplets collision and fragmentation are obtained. • A differential equation has been obtained that describes droplet coalescence or atomization. • A criterion for droplet fragmentation has been obtained. • The effect of collision on moving droplet evaporation in a heated atmosphere was regarded. [ABSTRACT FROM AUTHOR]

In this paper, cyclic triaxial tests are performed by using discrete element method (DEM) to investigate the dynamic behaviors of QH-E lunar soil simulant. The influences of number of cycles, loading frequency, waveform, cyclic stress ratio (CSR) and confining pressure on the dynamic behaviors of lunar soil simulant are discussed. The results indicate that the axial strain and dynamic modulus first increase and then tend to be stable with increasing number of cycles. The dynamic modulus increases with increasing the loading frequency, while the axial strain decreases. The damping ratio is not particularly sensitive to the confining pressure and CSR, it only decreases slightly with increasing confining pressure or decreasing CSR. Under the same conditions, rectangular wave dynamic load can produce a larger axial strain and a smaller dynamic modulus than those of sine and triangular waves, and a higher loading frequency will lead to a higher dynamic modulus and a lower axial strain, these characteristics are consistent with other lunar soil simulants and sandy soils tested in the experiment. However, the effect of confining pressure on the dynamic modulus depends on the magnitude of the CSR. For the lower CSR (CSR<0.6), the dynamic modulus increases with increasing confining pressure. While for the higher CSR(CSR>0.6), the dynamic modulus decreases significantly with increasing confining pressure, which demonstrates the unique dynamic characteristics that differs from other lunar soil simulants and sandy soils, and shows the importance of considering the magnitude of the CSR under cyclic triaxial tests. • Dynamic characteristics of QH-E lunar soil simulant under cyclic triaxial tests are studied by DEM. • Dynamic modulus and axial strain first increase and then gradually tend to be stable with increasing number of cycles. • The damping ratio is not particularly sensitive to the confining pressure and CSR. • Effect of confining pressure on the dynamic modulus depends on the magnitude of CSR. [ABSTRACT FROM AUTHOR]

The Discrete Element Method is uniquely suited to modeling the extreme fragmentation typical of impact-induced satellite breakups. In this paper, we introduce a discrete orthotropic material model for simulating carbon fiber reinforced polymers (CFRP). We calibrate and validate the CFRP model with hypervelocity impact experiments performed at Fraunhofer EMI on thin CFRP face sheets. We demonstrate the suitability of the model for simulating spacecraft breakups by comparing simulation results to a ground-based experiment of a 15 x 15 × 15 cm CFRP micro-satellite from literature. Fragment size distributions show excellent agreement between experiment and literature. Comparisons are also made with the NASA breakup model. • Introduce a discrete material model for simulating carbon fiber reinforced polymers. • Calibrate and validate the model with hypervelocity impact experiments. • Simulate the breakup of a 15 x 15 × 15 cm CFRP microsatellite. • Compare the simulation results to experimental fragmentation results from literature and NASA breakup model. • Simulation and experiment show excellent agreement. [ABSTRACT FROM AUTHOR]

Near Earth asteroids with diameters less than 50 m possess distinct characteristics including a large population, a high likelihood of Earth impact, and limited warning time. While small asteroid impacts may not induce global disasters, they can still result in localized hazard. This study focuses on investigating the feasibility and strategies for disrupting small-sized asteroid structures through kinetic impact under short-term warning scenarios. Initially, a dataset comprising ten asteroids ranging from 20 m to 50 m in diameter is compiled by correcting their orbital parameters to simulate potential impacts on Earth. Subsequently, a space-based interception constellation is designed and deployed at the L3, L4, and L5 points of the Cislunar space. The study further examines the trajectory design methodology for achieving rapid interception of asteroids, enabling timely response within a warning period of 7 to 14 days. Additionally, the paper discusses the feasibility of impactors in disrupting asteroids through analytical models and Adaptive Smoothed Particle Hydrodynamics (ASPH) numerical simulation method. Lastly, the defense capability of the space-based interception constellation against Earth-impacting asteroids approaching from various directions is assessed. The simulations show that C and S-type asteroids below 40 m can be effectively fragmented using an impactor with an initial mass of 2 tons within the specified 7 to 14-day warning time. Furthermore, the deployment of impactors at L3, L4, and L5 allows for comprehensive coverage of asteroids approaching from diverse directions. • Design of interception constellation for small-size asteroid defense under short warning time. • Fragmentation efficiency analysis of small size asteroids via multiple criterions. • Defense capability of the interception constellation against asteroids approaching from various directions. [ABSTRACT FROM AUTHOR]

The aerospace sector is evolving due to reduced launch costs and standardization of small satellite platforms. This research, aligned with European Guidelines for Thermo-Elastic Verification, addresses the pointing precision gap in small satellites by assessing space telescope performance using uncertainty propagation in thermo-elastic models. The methodology will be directly applied to an Earth observation space telescope, VINIS, currently under development by the Instituto de Astrofísica de Canarias (IAC). This procedure helps to identify key design elements impacting its functionality. Thirteen elements were identified as main contributors to the deformations in the optical bench. Due to the bench's crucial role in the telescope's performance, this paper also explores how results vary with different sandwich panel modelling techniques and the enhancements from design modifications. While the focus is on space telescopes, this approach has broader applicability to thermo-elastic analysis of various space instruments. • Research aligned with the thermo-elastic European guidelines. • Applied to the high-accuracy VINIS Earth observation space telescope. • Performance assessment through uncertainty-based method. • Identification of the features contributing to optical degradation. • Evaluation of different sandwich panel modelling techniques. [ABSTRACT FROM AUTHOR]

With interstellar mission concepts now being under study by various space agencies and institutions, a feasible and worthy interstellar precursor mission concept will be key to the success of the long shot. Here we investigate interstellar-bound trajectories of solar sails made of the ultra lightweight material aerographite. Due to its extremely low density (0. 18 kg m − 3 ) and high absorptivity (A ∼ 1), a thin shell can pick up an enormous acceleration from the solar irradiation. Payloads of up to 1 kg can be transported rapidly throughout the solar system, e.g., to Mars and beyond. Our simulations consider various launch scenarios from a polar orbit around Earth including directly outbound launches as well as Sun-diver launches towards the Sun with subsequent outward acceleration. We use the poliastro Python library for astrodynamic calculations. For a spacecraft with a total mass of 1 kg (including 720 g aerographite) and a cross-sectional area of 1 0 4 m 2 , corresponding to a shell with a radius of 56 m, we calculate the positions, velocities, and accelerations based on the combination of gravitational and radiation forces on the sail. We find that the direct outward transfer to Mars near opposition to Earth results in a relative velocity of 65 km s − 1 with a minimum required transfer time of 26 d. Using an inward transfer with solar sail deployment at 0.6 AU from the Sun, the sail's velocity relative to Mars is 118 km s − 1 with a transfer time of 126 d, where Mars is required to be in one of the nodes of the two orbital planes upon sail arrival. Transfer times and relative velocities can vary substantially depending on the constellation between Earth and Mars and the requirements on the injection trajectory to the Sun diving orbit. The direct interstellar trajectory has a final velocity of 109 km s − 1 . Assuming a distance to the heliopause of 120 AU, the spacecraft reaches interstellar space after 5.3 yr. When using an initial Sun dive to 0.6 AU instead, the solar sail obtains an escape velocity of 148 km s − 1 from the solar system with a transfer time of 4.2 yr to the heliopause. Values may differ depending on the rapidity of the Sun dive and the minimum distance to the Sun. The mission concepts presented in this paper are extensions of the 0.5 kg tip mass and 196 m 2 design of the successful IKAROS mission to Venus towards an interstellar solar sail mission. They allow fast flybys at Mars and into the deep solar system. For delivery (rather than fly-by) missions of a sub-kg payload, for example medical supply or replenishment of essential materials, the biggest obstacle remains in the deceleration upon arrival. • Aerographite solar sails were investigated as precursor propulsion concept. • Possible interplanetary/interstellar trajectories were simulated. • Aerographite has low density (0.18 kg m−3) and high absorptivity (A ∼ 1). • A spacecraft with a mass of 1 kg reaches Mars after 26 days. • The spacecraft becomes interstellar after 4 years with a velocity of 150 km/s. [ABSTRACT FROM AUTHOR]

This paper proposes an innovative approach for constrained rendezvous trajectory optimization (CRTO) using stochastic optimal feedback control and unscented transform (UT) uncertainty quantification. The method overcomes limitations of traditional deterministic CRTO solutions that suffer from uncontrolled terminal state error growth and severely deteriorated path constraint satisfaction when initial state uncertainty, dynamics uncertainty, and navigation uncertainty are considered. The approach involves constructing a stochastic optimal feedback control (SOFC) problem with chance constraints and introducing linear feedback control to regulate both the mean and variance of terminal state errors. UT is employed to approximately quantify the state errors and their propagation, transforming the SOFC problem into an unconstrained deterministic optimization (UDO) problem. Differential dynamic programming (DDP) is then used to solve the UDO problem. The obtained stochastic optimal solution provides a robust rendezvous trajectory and a corresponding explicit closed-loop guidance law, improving terminal accuracy and path constraint satisfaction in the presence of various considered uncertainties. The influence of different term weights in the objective function on terminal accuracy and path constraint satisfaction is also studied. The effectiveness of the proposed approach is verified through Monte Carlo simulation, demonstrating the robustness of the closed-loop control strategy. The results highlight the potential of the method in enhancing terminal accuracy and path constraint satisfaction in uncertain rendezvous scenarios. • The uncertainty-resilient rendezvous trajectory optimization algorithm provides not only a nominal trajectory but an explicit optimal state-feedback control policy. • Rendezvous trajectory uncertainty caused by initial state errors, dynamics errors, and navigation errors is propagated using an unscented transformation method. • The proposed method considers common input constraints and path constraints in space rendezvous missions. [ABSTRACT FROM AUTHOR]

For many years, space human factors research has focused on the impact of microgravity on human physiology and the health risks astronauts face during space flights of various durations. Our review paper reveals that microgravity leads to both reversible and non-reversible effects on astronauts' oral health. In addition to oral health concerns, astronauts often experience muscle degeneration during space missions, which is closely tied to bone loss. Despite our knowledge of high bone resorption and rapid mineral and calcium loss in microgravity, the exact underlying mechanisms remain unclear. However, to fully understand these processes and develop effective strategies to counteract the consequences of microgravity, more research is needed. This includes a deeper exploration of how these molecules and other biomarkers work, identification of early mechanotransduction sensors, and a comprehensive understanding of the physiological effects of weightlessness. Current solutions primarily involve administering antioxidants, regular exercise, and appropriate nutrition. Additionally, innovative approaches such as experimental drug therapies using anti -myostatin antibodies and r-irisin have gained attention. Further large-scale, long-term studies are required to understand the effects of microgravity on the oral cavity and prevent its adverse consequences. • Impact of Microgravity on Dental Health: Discuss the physiological changes experienced by astronauts in microgravity environments, including fluid redistribution, decreased mechanical loading, and altered bone metabolism, and their implications for oral health. • Weight-Bearing vs. Non-Weight-Bearing Bones: Examines the relationship between microgravity and bone loss, with a focus on distinguishing its effects on weight-bearing and non-weight-bearing bones. • Comprehensive Impact Assessment: Calls for comprehensive research on how microgravity affects various bones in both animals and humans, including analyses of bone structure, density, and strength. • Bone Loss and Tooth Mobility: Examine how microgravity-induced bone loss affects the integrity of the jawbone, leading to increased tooth mobility and potential dental complications such as loosened crowns, dental implants, or even tooth loss. • Differential Recovery: Investigates how weight-bearing and non-weight-bearing bones recover following bone loss in microgravity conditions and the underlying mechanisms at play. • Addressing Inconsistencies: Highlights the need to resolve inconsistencies in existing literature regarding the influence of microgravity on osteogenic proliferation and differentiation. • Osteocytes and Osteoclasts in Microgravity: Emphasises the importance of understanding the roles of osteocytes and osteoclasts in microgravity-induced bone loss to identify potential therapeutic targets. [ABSTRACT FROM AUTHOR]

Feature extraction and matching play integral roles in autonomous vision navigation. As a typical morphological feature, craters can be used as navigation landmarks. However, the spin of small bodies introduces substantial variations in illumination and viewing angles within acquired images, thereby posing challenges to feature extraction and matching algorithms. Addressing these challenges, a deep learning-based approach emerges as pivotal in mitigating the limitations of existing navigation algorithms. This paper presents an algorithm designed to segment noteworthy craters and irregular regions on such small bodies, thereby furnishing crucial autonomous navigation data for missions involving uncharted terrains. By seamlessly integrating deep learning techniques with conventional computer vision-based feature description methods, the algorithm achieves matching segmentation results of noteworthy precision. The experimental results show that the algorithm has better results compared to the existing mainstream segmentation networks. And it can achieve the accurate matching of segmentation results. • Accurate segmentation of small-body craters via deep-learning Algorithms. • Gradient descriptor construction method based on segmentation results. • The proposed method has viewpoint, illumination, scale, and rotation invariance. [ABSTRACT FROM AUTHOR]

For impact velocities larger than a threshold velocity of 3 km/s, aluminium Whipple Shields present an enhanced protection capability with respect to monolithic protections with the same areal density. In particular, in the range between 3 and 7 km/s the projectile partially fragments and melts after impacting the bumper; modelling this transition might be complex due to the high number of parameters affecting the collision. In particular, limited data is available in literature with a systematic evaluation of such parameters. In this context, a campaign of 22 experiments was performed to assess the response of aluminium Whipple Shields to normal impacts of aluminium projectiles in the transition range up to 5 km/s. In the tests, the projectile diameter was fixed at 2.9 mm and the bumper thickness and impact velocity were systematically varied respectively at 1, 1.5, and 2 mm and between 2.6 and 5 km/s. Collected data included high-velocity videos of the debris cloud generated by the impacts at both low resolution (for all shots) and high resolution (for 19 out of 22 tests); in addition, for 10 experiments high-resolution images of both the bumper and the wall were acquired after the tests. In this paper, the experimental campaign is described and the main collected results are presented; in particular, the influence of the different impact parameters is discussed. Experimental results are finally compared with numerical simulations conducted with a smoothed particle hydrodynamics method available in IMPETUS Solver. • 22 impact experiments were performed on Aluminium Whipple Shields. • The projectile diameter was fixed at 2.9 mm for all tests. • Bumper thickness and impact velocity varied in the ranges 1–2 mm and 2.6–5 km/s. • The influence of the different impact parameters is discussed. • Experimental results are compared with numerical SPH simulations. [ABSTRACT FROM AUTHOR]

Until recently space travel was considered a privilege of governmental astronauts, with the exception perhaps of a few billionaire space tourists, who had been accommodated on the International Space Station on some rare occasions. Nonetheless, with major space pioneers announcing plans for space tourism activities and after the successful launch in orbit of space tourism vehicles from the Virgin Galactic and Blue Origin companies, academic interest about human spaceflight for purposes of leisure and recreation has ignited. Taking into account that space tourism is in fact an unregulated activity, this paper will firstly examine whether space tourism activities fall under the scope of space law or under the scope of international air law, depending on the planned trajectory of the space tourism vehicle. Furthermore, the authors will attempt to draw conclusions concerning the legal implications of space tourism activities through a comparative analysis with the 1974 Athens Convention on the Carriage of Passengers and the Carriage of their Luggage by Sea. This international convention provides a great example on the regulation of touristic ventures, as it refers to commercial touristic activities realized within a res communis omnium , i. e., the high seas, similarly with touristic activities in space. Moreover, said Convention provides for a specialized regime for touristic activities and is of particular interest, since it has achieved stability and a desirable balance of interests between private tourism operators and passengers. In this connection, the authors will evaluate the provisions of the 1974 Athens Convention in comparison with the space treaties applicable to space tourism activities. Furthermore, suggestions will be made as to whether it would be appropriate and feasible to introduce in the space law realm certain rules of law similar to those of the 1974 Athens Convention, in order to facilitate growth in the space tourism sector, without compromising the safety of human spaceflight for space passengers. • The 1974 Athens Convention in comparison with the space treaties applicable to space tourism activities. • Legal analogy between space law and maritime law. • Legal regime governing space tourism activities. [ABSTRACT FROM AUTHOR]

This paper investigates the dynamic behavior of a tethered satellite system in an elliptical orbit as the tethered subsatellite is deployed from a mother satellite at an out-of-plane angle while maintaining the tether at a constant in-plane pitch angle by a predefined control force acting on the subsatellite. A variable-length pendulum model is employed to describe the purely out-of-plane tether deployment dynamics, from which an analytical tether deployment control law is derived. The stability condition of the analytical control law is examined by Floquet theory. The corresponding analytical deployment law is asymptotically stable, and the requirement for a positive tension control input is guaranteed to ensure the controllability of the tether system. Following deployment, the critical conditions for the out-of-plane libration of the tethered subsatellite to become chaotic during the station-keeping phase are examined. Finally, simulation results reveal that the proposed approach can effectively and rapidly identify periodic, quasiperiodic, and chaotic motions using bifurcation diagrams, which provides valuable insights for enhancing the stability and reliability of such tethered satellite systems. • Proposed an analytical deployment law for a purely out-of-plane tethered satellite. • Revealed nonlinear phenomena of the system during the station-keeping phase. • Explored bifurcation characteristic and chaotic windows of the system. [ABSTRACT FROM AUTHOR]

This paper presents dynamics and control of orbits around an asteroid that is modelled as a uniformly rotating homogeneous rectangular parallelepiped (HRP). Aside from the non-uniform gravitational field of the asteroid, we include the solar radiation pressure (SRP) as a perturbing force. A simple model based on the polyhedral method is employed for calculating the gravity around the HRP. First, circular periodic orbits are identified both in the presence and absence of solar radiation pressure. Then, families of general non-circular periodic orbits in the presence of SRP are determined and subsequently analysed for stability. Next, for a general orbit in the equatorial plane, its stability against a catastrophic impact of an orbiting spacecraft onto the asteroid's surface is checked in the presence of SRP. Orbits, where the solar radiation pressure is less/more dominant over the central body's gravity, are identified with the aim of finding the best orbits for parking a spacecraft. Then, for long-period orbits, SRP is averaged and the time averaged evolution of the orbital elements is obtained. For a special set of initial condition, a closed form solution for the evolution of the orbital eccentricity is obtained. Using this solution, a lower limit for the life-time of a spacecraft that is left uncontrolled is derived in a closed form. Finally, active control strategies are developed for orbit-keeping ofspacecraft for any choice of orbital parameters. A linear-quadratic regulator (LQR) is designed to track a given Keplerian orbit with minimum fuel expenditure. Fuel expenditure on various orbits are calculated and compared to determine the orbits of cheapest operation. The work presented is applicable to small body (asteroids and comets) exploration missions which require online cancellation of the perturbing forces with minimal knowledge of the accurate shape, size, and composition of the small body prior to the mission. • Radiation perturbed periodic orbits around a parallelepiped-asteroid are obtained. • A boundary for stable non-crossing orbits is also analytically derived. • Closed form solutions for far-range orbits are obtained through Lagrange averaging. • Using the above, an analytical estimate for the lifetime of the orbit is presented. • A linear quadratic regulator (LQR) is employed to obtain ideal parking stations. [ABSTRACT FROM AUTHOR]

Operation to failure was the modus operandi of most satellite operators prior to the awareness of the impact on the space debris environment. While compliance to space debris mitigation guidelines, standards and regulations is clearly a priority, there remains a business advantage in extracting the maximum lifetime out of any in-orbit asset. For certain types of strategic assets, which provide unique societal benefits through their missions and are typically irreplaceable, due to their complexity, cost and public funding sources, it becomes imperative to ensure extraction of maximum mission value while fully respecting the need for a sustainable orbital environment. With reference to mission examples in EUMETSAT, this paper highlights: the techniques which have already been used to ensure compliant end-of-life strategies while extending mission lifetimes on satellites not designed to support disposal; the approaches made on new generation satellites to achieve compliance with the debris mitigation standards, highlighting what this brings, but also risks remaining for debris mitigation and lifetime value optimisation; how an Active Debris Removal service in LEO positively impacts the lifetime extension/debris mitigation risk trade-off, describing the potential business case for such services on high value, strategic assets. • Active Debris Removal provides economic benefits for operators through mission extension. • No need to use own fuel for self-disposal; longer operational life. • Operation to failure becomes acceptable again, reduced mission continuity risks. • One-off government satellite programmes could be major beneficiary. • Study on application to EUMETSAT Metop mission underway. [ABSTRACT FROM AUTHOR]

The Institute for Microelectronics and Microsystems of the Italian National Research Council (Italy) has designed and fabricated a water-fed MEMS vaporizing liquid microthruster (VLM) in collaboration with the University of Salento (Italy) and KU Leuven (Belgium). Previous studies have demonstrated the device's functioning and highlighted the presence of a strong thermoelectric-hydraulic coupling that severely affects the heating and propulsive efficiencies and reliability of the device. In this regard, the current work provides a preliminary experimental investigation of a pulsed heating system actively controlled by temperature using PI-D logic (proportional and integrative actions applied to the error; derivative action applied to the output). The overall performance assessment focuses on evaluating power consumption during steady-state operation and analyzing the dynamic response of the VLM undergoing such an actively controlled pulsed heating, which has been built using both manual and data-driven offline adaptive tuning of the PID parameters. Concerning the steady-state operation, results highlight a promising enhancement of the heating efficiency to values above 0.9. The dynamic response analysis shows that operating with a single set of fixed PID parameters is not feasible and that the manual tuning is a trial-and-error approach highly dependent on the operator's experience, operating conditions, and reference temperature profile. Furthermore, shorter rise and response times require a higher proportional gain, and the overshoot experienced when crossing the saturation temperature cannot be avoided. In this paper, we show that the response of the data-driven adaptive controller solves these issues, exhibiting a reduced maximum settling time from around 59 s to less than 10 s, while the overshoot is avoided thanks to the temporal adjustment of the PID parameters. • Thermoelectric-hydraulic coupling impacts VLM heating and propulsion efficiency. • PI-D controlled pulsed heating investigated for VLM propulsion. • Steady-state efficiency boosted beyond 0.9. • Adaptive controller reduces settling time. • Data-driven approach minimizes overshoot. [ABSTRACT FROM AUTHOR]

Risk assessment studies of local asteroid hazards traditionally simulate the physics of meteors with engineering models tailored to analyze tens-of-millions of scenarios. However, these simplified approaches still need to solve time-dependent ODEs to model the entry process and the resulting ground damage. With a computational cost of O(0.01 CPU.s) per scenario, simulating these large numbers of potential entry conditions can take several days on local computers. To improve computational efficiency, we propose in this paper an orthogonal approach based on machine learning models to predict the size of damaged areas given a list of entry parameters. We train 5 machine learning methods and compare the predictions to the outputs of NASA's state-of-the-art Probabilistic Asteroid Impact Risk (PAIR) model. PAIR's remarkable capability to handle a broad spectrum of entry conditions and a high volume of scenarios establishes it as our reference model, with PAIR estimates assumed ground truth and used to train, test, and validate the machine learning models. The ML models, initially trained with PAIR's primitive input variables, are then further enhanced with engineered features to improve the performance of interpretable models. We find that complex models like neural networks are well-suited to estimate blast hazards, while simpler, interpretable linear models can accurately assess thermal damage. For both types of hazards, the radii of damaged areas can be predicted with less than 10% average relative errors for radii greater than 100 km. The CPU time is decreased by a factor O(1 0 3 ) compared to the PAIR model, enabling the simulation of millions of scenarios within minutes on a local computer. Beyond serving as surrogate models, we incorporate the machine learning models into a comprehensive Shapley sensitivity analysis, yielding a ranking of entry parameters based on their contributions to ground damages. These results offer valuable insights for prioritizing observation missions. • Machine learning models enable accurate and rapid simulations of asteroid hazards. • Data-driven methods can assess the radius of damaged areas with 10% average errors. • Models provide insights on the relative importance of entry parameters. • Risk studies can be performed on local machines instead of supercomputers. [ABSTRACT FROM AUTHOR]

Flexible vibration has a great impact on flexible spacecraft attitude control, which will reduce the pointing accuracy of the spacecraft platform and bring severe challenges to the design of a high-precision attitude control system for flexible spacecraft. Aiming at the problem of attitude and vibration coupling control of flexible spacecraft, an anti-unwinding attitude coupling controller using the positive position feedback (PPF) control is proposed in this paper. The model compensation control term is constructed based on the prior information to improve the tracking performance. The proposed attitude coupling control method can effectively suppress the flexible vibration while achieving more accurate attitude control. The obtained closed-loop system under the proposed control scheme is asymptotically stable, which is proved by the Lyapunov stability theory. The numerical simulation illustrates that the proposed method can effectively improve the attitude control accuracy of the flexible spacecraft and is capable of large-angle attitude maneuver control because of its prominent performance in attitude tracking and its unwinding-free performance. • Coupling control for flexible spacecraft. • Active vibration control using positive position feedback. • Model compensation control term is constructed to improve the dynamic performance. • The obtained closed-loop system achieves the asymptotic stability and anti-unwinding performance. [ABSTRACT FROM AUTHOR]

Separation and reconfiguration of stacked satellites in orbit is an effective method for constructing large space structures. This paper investigates the dynamics of a stacked satellite system (SSS) during the separation and reconfiguration process. Firstly, the dynamic model of the SSS is established via the natural coordinate formulation (NCF) to conveniently handle the time-varying constraints between satellites. Then, the separation and reconfiguration strategies are proposed for the SSS including the reconstruction step between two configurations. To ensure collision-free satellite assembly, a proportional differential (PD) controller combined with a potential function is introduced. Additionally, a docking mechanism is designed, considering collision force constraints during satellite docking assembly. Finally, several numerical simulations are conducted to demonstrate the effectiveness of the separation and assembly strategies. • An orbital dynamic model of a stacked satellite system (SSS) is established. • A collision detection algorithm for the docking mechanism is introduced. • A PD controller law for the collision-free assembly of satellites is designed. • The dynamics for reconfiguration and assembly of the SSS is studied. [ABSTRACT FROM AUTHOR]

The pioneering objective of a permanent outpost on the lunar surface requires the development of technologies for in situ resource utilisation. Consolidation of lunar soil into products is imperative to avoid unaffordable transportation burdens. The paper addresses the feasibility of the Local Microwave Heating effect as a promising technology for sintering lunar regolith based on microwave radiation, with a simple solid-state generator requiring a relatively modest amount of energy. Experiments were conducted using terrestrial simulants with physical and chemical properties similar to those of lunar regolith. Laboratory equipment with a purpose-built microwave actuator was designed and manufactured with "Design for Embarking" in mind. Experiments have demonstrated the feasibility of the proposed technology with a reasonable degree of confidence. Solid artefacts were obtained with proper setting of microwave parameters and appropriate scanning strategy, proving the feasibility of regolith sintering with a localised microwave source A roadmap for future development of the technology has been established. • Preliminary study on Localized Microwave Sintering of Lunar Regolith: Highlights. • Localized Microwave Heating is effective to consolidate lunar regolith simulants. • A solid-state generator with an open-end antenna is a promising lightweight low-energy system for ISRU. • Microwave heating of regolith is feasible regardless of the simulant's ilmenite content. • The technique can be used to produce 3D parts using an additive approach. [ABSTRACT FROM AUTHOR]

Liquid crystal diffractive sails can generate controllable tangential radiation pressures, which have advantages in the Solar Polar Imager mission (SPI). This paper proposes a composite diffractive sail design consisting of an ideal passive diffraction grating sail and liquid crystal polarization gratings (LCPGs). The main diffractive grating sail generates thrust for orbital transfer with a sun-facing attitude. An optimal 3.8-year transfer trajectory is obtained by an improved direct method. Small-area LCPGs distributed around the edge are used for three-axis attitude control. The LCPGs can continuously modulate the force and torque by controlling the applied voltage based on the electro-optical characteristic of liquid crystal materials. A control strategy, including an adaptive attitude-tracking PD controller and a control allocation method, is designed to track the attitude profile generated by the transfer trajectory. Finally, the attitude-orbit coupling dynamics are used to simulate the whole process, and the results show that orbital insertion errors are minor. • A combined diffractive solar sail model using liquid crystal polarization gratings. • Independent and controllable three-axis attitude control of diffractive solar sails. • Optimizing transfer trajectory for solar polar mission based on the proposed sail. • An adaptive attitude-tracking PD controller and a control allocation method. [ABSTRACT FROM AUTHOR]

NASA's Probabilistic Asteroid Impact Risk (PAIR) model assesses the likelihood of potential damage for asteroid impact scenarios. Fast-running models are used to capture the effects of different hazards. This paper looks specifically at local ground damage hazards, including blast overpressure and thermal radiation damage, for large object impact scenarios. A sensitivity study is conducted to determine which parameters, and over what ranges, cause impact risks to become sensitive to thermal damage. Two additional thermal models with different approaches are used for comparison. The study determined the current thermal model is most sensitive to the luminous efficiency parameter that reflects the model's uncertainty in the amount of energy contributing to the thermal radiation damage. This sensitivity was most apparent for the highest severity damage levels. Comparisons of the three models showed that, in addition to sensitivities within the models, the impact risks are also sensitive to the choice of thermal model. The study results were applied to the 2023 Planetary Defense Conference hypothetical asteroid impact scenario (800 m diameter, 10.29 Gt energy), and parameter ranges of interest were determined. At the serious damage level, luminous efficiencies above 0.006 (0.6%) showed a small chance of thermal playing an important role, while luminous efficiencies above 0.0008 (0.08%) led to thermal playing a significant role at the unsurvivable damage severity level. Study results are used to identify key areas where additional model refinement and better knowledge of asteroid properties may be important for improving damage estimates. • Thermal radiation damage plays a more significant role in large object impacts. • Local ground damage hazards are sensitive to luminous efficiency. • Variation exists in available thermal radiation model approaches and results. • Further refinement in luminous efficiency range and distribution needed. [ABSTRACT FROM AUTHOR]

The primary objective of this paper is to study the detonation re-initiation through multiple reflections upon its entry into a confined space containing liquid fuel within a flow system, which holds significant implications for two-phase pulse detonation engines (PDE) and rotating detonation engines (RDE). Utilizing the Eulerian-Lagrangian method, this study investigates the disparities in the diffraction, quenching, and re-initiation processes of liquid n-heptane/air detonation upstream and downstream. Additionally, it explores the interaction between shock waves and droplets as well as determines the lean flammability limit for re-initiation within a flow system. The findings suggest that the incoming flow significantly contributes to the upstream decoupled detonation re-initiation, as it intensifies the leading shock, thereby expediting droplet fragmentation and evaporation. However, the compressive effect of the incoming flow is insufficient to initiate upstream detonation independently. To achieve re-initiation, it is still necessary for the detonation wave to be reflected by the wall. Moreover, the incoming flow expands the re-initiation boundaries, thereby facilitating the re-initiation of upstream detonation. As the equivalence ratio decreases, a higher inflow Mach number is required to achieve upstream re-initiation. • Upstream and downstream detonation re-initiation are compared. • Interaction between droplets and leading shock upstream and downstream are illustrated. • The detonation re-initiation boundaries beneath an incoming flow are found. [ABSTRACT FROM AUTHOR]

Controlled de-orbiting is crucial for a low earth orbiting (LEO) satellite or capsule intending to land in a desired location and to prevent damage to people and property on the ground caused by debris. Drag modulation is one possible mechanism to control de-orbiting, exploiting atmospheric drag variation to reduce the necessary orbital energy from the initial conditions to the re-entry interface. In this context, this paper proposes a novel targeted de-orbiting Artificial Neural Network (ANN) and drag-based guidance algorithm for a LEO artificial satellite. It relies on the combination of empirical and analytical relations with Deep Neural Networks (DNNs) to estimate the main parameters of the optimal control law in real time. The training set is composed of several optimal control solutions generated with a previously developed optimal control algorithm, exploiting the formulation in equinoctial modified orbital parameters to manage the large time scale of the problem and the computational cost. An innovative procedure for the training set generation led to the achievement of general results with a limited number of samples. The successful outcome of the guidance algorithm on over 1000 cases demonstrates its robustness and generalization of results. To conclude, a Linear Quadratic Regulator (LQR) feedback control is applied to one case to deal with a more realistic and uncertain density model. • Analytical optimal guidance for a satellite de-orbiting from Low Earth Orbit. • A novel approach based on Deep Neural Networks. • Prediction of the key parameters of the optimal control law in real time. • A Monte Carlo analysis showed the algorithm's robustness and generalization. • Successful feedback control thanks to a Linear Quadratic Regulator (LQR). [ABSTRACT FROM AUTHOR]

The launcher recovery method known as 'In-Air Capturing' is an innovative approach where a winged rocket stage is captured mid-air by an aircraft and subsequently towed back to the launch site. To achieve this, a capturing device attached to a tether is released from the towing aircraft. This device autonomously connects the reusable launch vehicle to the aircraft, while the two vehicles are in proximity. The tether's flexible dynamics have a strong influence on the maneuverability of the capturing device. Thus, it is critical to model the dynamics of the tether to get a realistic understanding and to evaluate feasibility of the concept. However, precise modelling and control of highly flexible systems with large deformations is both challenging and computationally intensive. For engineering applications which involve closed-loop simulations, finding a trade-off between accuracy and computational effort is crucial. In this paper, the tether is modelled as a discretized chain of rigid bodies connected by rotational springs at the joints. The tether along with the capture system is integrated in a simulation model and open loop tests are performed to analyse the system characteristics. The tether properties are studied to find a suited configuration for 'In-Air Capturing'. • Multibody modelling approach for long cable like bodies with large deformation. • Capture system dynamics for the unique 'In-Air Capturing' launcher recovery concept. • Dynamic modelling for the complete capturing system including towing aircraft, capturing device and tether. • Assessment of tether properties suited for 'In-Air Capturing' with sensitivity studies. [ABSTRACT FROM AUTHOR]

This paper addresses the fault-tolerant attitude tracking control problem for spacecraft with limited communication capability and input saturation. The design of a hysteresis quantizer aims to alleviate communication burden between the controller module and the actuator module. Then, the attitude tracking problem for spacecraft with input quantization, input saturation, actuator faults and external disturbances is transformed into an attitude control problem with uncertain input coefficients and bounded disturbances. Thereafter a dynamic loop gain function-based approach and a hyperbolic tangent function term with a time-varying boundary layer are introduced to address the uncertainties of input coefficients and the bounded disturbances, respectively. To facilitate the boundness analysis of the signals in the closed-loop system, a pertinent lemma about the dynamic loop gain function is proved. Finally, numerical simulations are employed to validate the effectiveness of the proposed scheme. • The spacecraft attitude tracking problem with disturbances and uncertainties. • The quantized control law alleviates communication burdens of spacecraft. • A dynamic loop gain function method is used to address the impact of uncertain. • A hyperbolic tangent term is devised to counteract bounded disturbances. [ABSTRACT FROM AUTHOR]

This paper studies the high telepresence control for the relay communication based space teleoperation. High telepresence requires real-time, accurate force perception and good position tracking, but all of them are destroyed or affected by time delay. To eliminate the influence of time delay, direct recursive generalized predictive control (DR-GPC) is quoted into this paper to realize position tracking. As for the other two properties, the pseudo position-force (Pseudo-PF) feedback strategy (composed of parameter identification and force reconstruction) is a good way as it can provide real-time and precise force feedback when environment change slowly and time delay is relatively small. However, complex and changeable environment and serious and complex time delay in space teleoperation make the environment force parameter available to force reconstruction is outdated and cannot realize accurate force rebuilding. To solve this problem, we propose a dual regression Pseudo-PF feedback strategy. Compared with the traditional one, its biggest change is the introduction of dual regression that consists of time delay prediction and parameter regression. The former predicts time delay and provides it to the latter to determine the forward prediction steps, while the latter regresses the actual latest and precise environment force parameter according to the forward prediction steps. With the cooperation of time delay prediction and parameter regression, the influence of fast changing environment or serious and complex time delay on force reconstruction can be effectively overcome, so as to realize real-time and accurate force feedback. Finally, DR-GPC and dual regression Pseudo-PF feedback strategies cooperate to achieve high telepresence control, and simulation is given to show its effectiveness. • A dual regression Pseudo-PF feedback strategy is proposed. • High teleopresence control for the relay communication based space teleoperation is first studied. • The introduction of dual regression makes the applicability of this method wider. [ABSTRACT FROM AUTHOR]

This paper presents an intelligent task planning strategy for an Earth Observing Satellite (EOS) monitoring the Ozone distribution using multi-step Bayesian Optimization (BO). Significant improvement in observation performance can be achieved by implementing an offset-pointing strategy using single-step or classical BO compared to the simple nadir-pointing approach, which does not consider the possible future outcomes when deciding the current action. To account for future considerations, this paper presents a modified Continuous Belief Tree Search (CBTS) algorithm that enables a non-myopic multi-step look-ahead optimization in the continuous action domain. The design utilizes a Control-effort based Upper Confidence Bound function as the acquisition function for BO. Simulation results demonstrate that the multi-step BO is capable of significantly reducing the model uncertainty compared to the single-step BO. Additionally, using BO provides a superior way to determine the first actions than obtaining them randomly. Moreover, sampling the training data set provides a way to reduce the computational time without reducing the computation tree size. • Intelligent task planning creates informative models from satellite data. • Task planning is designed based on Multi-step Bayesian Optimization. • Control-effort based acquisition function is constructed for action planning. • First action determination based on Bayesian Optimization is investigated. • Randomly sampled training dataset is devised to reduce computational time. [ABSTRACT FROM AUTHOR]

In recent years, small celestial bodies (SCBs) exploration missions have attracted extensive attention. For the purposes of space resource exploitation and earth defense, revealing the formation and evolution of the solar system and exploring the origin of life, sampling return and in situ resource utilization (ISRU) have become the focus of SCBs exploration missions. Based on this background, this paper attempts to comprehensively illustrate the advances in SCBs resources utilization and exploitation technologies. First, the status of the utilization and exploitation of SCBs is summarized. The classification and resource distribution of SCBs are introduced later and focus on the utilization prospects of SCBs resources. From the four aspects of resource prospecting, exploitation, extraction and storage, transfer and transportation, the technical progress of SCBs mining is discussed. Some typical sampling technologies, such as siliceous aerogel, impact projectile, auger drill and excavating mechanism, pneumatic apparatus and wheel brush sampler, cutter and abrasive tool, laser and electromagnetic device, as well as bionic mining technologies inspired by some animals and plants, are reviewed in detail. According to the research of space mining, the critical points and legal issues involved in the utilization and exploitation of SCBs resources are presented, and future development trends are proposed. Finally, the full paper is summarized. • The status of the utilization and exploitation of SCBs is summarized. • The classification and utilization prospects of SCBs resources are introduced. • The technical progress of SCBs sampling and mining is discussed. • The critical points and issues are presented. • Future development trends are proposed. [ABSTRACT FROM AUTHOR]

NASA recognizes Human Systems Integration (HSI) as part of the overall systems engineering and acquisition strategy for space systems. The Lunar Gateway Program is implementing HSI technical and management process across the lifecycle of the mission, as required by NPR 7123.1C NASA Systems Engineering Processes and Requirements, and led by the Gateway HSI team as required by NPR 8705.2C Human-Rating Requirements for Space Systems, now HEOMD-003 Crewed Deep Space Systems Human Rating Certification Requirements and Standards for NASA Missions. NASA has been using HSI principles for many years and has applied them to many of its previous human spaceflight Programs. As NASA returns to the Moon in a more sustainable manner, the Gateway Program is maturing the application of HSI by implementing it more visibly as part of Artemis, with guidance from the NASA/SP-20210010952 NASA HSI Handbook. This paper discusses how HSI is being implemented in the Gateway Program, challenges faced with its implementation during the development phase, and strategies/approaches used to overcome those. The paper also covers HSI implementation for flight systems, vehicle processing, and interfaces across the six identified NASA HSI domains: human factors engineering, operations, safety, training, maintainability and supportability, habitability and environment. The goal is to provide an overview of the implementation process of HSI in the Gateway Program as an example for other Programs/Projects/Missions that are looking to implement HSI. • Human Systems Integration (HSI) overview and Gateway scope. • HSI activities and products in NASA's Systems Engineering lifecycle. • Challenges with HSI implementation during design development. • Strategies and approaches to implementation of HSI in Gateway. [ABSTRACT FROM AUTHOR]

This paper focuses on mixed-integer trajectory optimization of no-fly zones avoidance for a hypersonic vehicle. When there are multiple no-fly zones, the trajectory optimization problem has many local optimal solutions, and the performance of the solution depends on the selection of an initial guess. Essentially, the selection of the initial guess can be abstracted as a discrete path decision. To deal with this issue, this paper first integrates the discrete path decision and continuous trajectory optimization for no-fly zones avoidance and models these as a mixed-integer optimal control problem (MIOCP). As the corner stone, a creative directed graph is established to model the path decision problem, which is then formulated by integer constraints. Combined with the minimum effort optimal control problem (Problem 1), we form the MIOCP (Problem 2), thus integrating the path decision and trajectory optimization to improve global performance. Additionally, to solve the MIOCP, an effective iterative mixed-integer convex programming (IMICP) algorithm is customized by adding two ad-hoc techniques, relaxation and a proximity term, thus remedying the artificial infeasibility. Simulation results show that this approach is practically independent of the initial guess and has better performance than existing methods. • An approach is proposed to improve the global performance of no-fly zones avoidance. • A creative graph is established for path decision and the mixed-integer problem. • Relaxation and proximity terms are used in the IMICP algorithm for stable iteration. • The convergence is not practically dependent on the initial guess. [ABSTRACT FROM AUTHOR]

The National Science Foundation's (NSF) decision to rename the Large Synoptic Survey Telescope (LSST) to the Vera C. Rubin Observatory in 2020 made it the first U.S. observatory to be named after a female astronomer. While this move fosters a heightened awareness of women's struggles for advancement in the sciences, it also indicates a need for additional historical studies of women in these fields. Our research seeks to create a new narrative for the historical experiences of women in astronomy and astrophysics through the poetry of second-wave feminist Adrienne Rich, author of Diving into the Wreck and Of Woman Born. Rich's 1968 poem "Planetarium" celebrates the career of Caroline Herschel, while also hinting at the more recent accomplishments of women astronomers like Dame Jocelyn Bell Burnell. Literary historiography is a theoretical methodology for understanding historical subjects like Herschel, through their representations in literary works. Despite the intimate format, Rich's retelling does not focus on Herschel's personal relationships but on her professional pursuits. Rich uses astronomy as a metaphor for her and other women scientists' experiences as they struggle to tune out the noise of obligations, expectations, and invisible labour, and strive to locate the signal representing individual motivation and discovery. This paper examines the professional methodologies of women in astronomy from the lens of feminist literary historiography to support science educators and policymakers in adopting gender inclusive practices. Using Rich's poem "Planetarium," alongside the scientific processes involved in astronomical discovery, this work explores the ways in which Rich's poem reimagines astronomers like Caroline Herschel, Jocelyn Bell Burnell, and Vera Rubin, and how the piece conjures from their lives a new narrative for women in pursuit of their own vocations. The exploration of the intertwining nature of history and culture allows for an analysis of how the professional methodologies of women in astronomy relate more profoundly than differing historical periods would suggest and provides additional insight into the unique challenges currently facing women of colour in these same scientific areas. This paper seeks to contribute a new perspective in space and science communication discussions and help mitigate the dissemination of gender exclusivity in research and education. • Uses astronomy as a metaphor for women scientists' experiences. • Examines professional methodologies of women in astronomy. • Applies feminist historiography to the lives of women in astronomy. • Attempts to mitigate the dissemination of gender exclusivity in research and education. [ABSTRACT FROM AUTHOR]

Modular self-reconfigurable spacecraft (MSRS) are composed of homogeneous or heterogeneous modules that can autonomously achieve configuration transformation without external intervention. In this study, the research status and key technologies of MSRS are reviewed. First, the concepts of modular reconfigurable spacecraft (MRS) and MSRS are described, and the advantages of the latter are analyzed. Typical MSRS are then introduced. Furthermore, the key technologies of MSRS are analyzed from four aspects: assembly structure design technology, mission configuration optimization technology, self-reconfiguration planning technology, and attitude cooperative control technology. Finally, the future trends and development of MSRS are presented. These spacecraft are expected to play a vital role in the aerospace field. • This paper is the first review of the modular self-reconfigurable spacecraft. • This paper gives a definition of a modular self-reconfigurable spacecraft. • This paper provide a detailed analysis of the self-reconfiguration planning technology. • A future ground management control system based on parallel systems is introduced. [ABSTRACT FROM AUTHOR]

Mars sampling is of great significance for the comprehensive study of Martian geological evolution. The successful demonstration of the Ingenuity Helicopter reveals the potential of rotorcraft for sampling tasks on Mars. However, the Ingenuity Helicopter cannot carry samplers, and large helicopters face difficulties in folding automatically. Thus, a multi-axis rotorcraft may be more suitable for Mars sampling tasks. This paper proposes an octocopter named MarsBird-V with a foldable rotor system and wheels for moving agilely. The design method for the structure and control system of MarsBird-V is illustrated, followed by the mass and power estimation approach. On the basis of the blade element momentum theory, the criterion is proposed for the aerodynamic performance evaluation of the coaxial rotors. The aerodynamic data of the coaxial rotors are obtained by the computational fluid dynamics method, which is applied to determine parameters. The propulsion capacity is calculated in the rotational speed range limited by the Mach number of the blade tip. As estimated, the minimum gross mass of the MarsBird-V is 6 kg. The proportion of payload in the gross mass increases as the power increases. The MarsBird-V achieves maximum durability of 420 s at a power of 1017 W and obtains a maximum payload efficiency of 0.43 g/W at a power of 2362 W. This paper also presents the deployment process and mission analysis. The MarsBird-V can perform sampling tasks in a cycle time of 2 to 5 days. • A foldable Mars octocopter is proposed to perform Mars flight-sampling tasks. • The structure, avionics, and GNC systems of the Mars octocopter are designed. • The collective pitch angles of the upper and lower rotors are 23° and 25°. • The minimum gross mass of the MarsBird-V is 6.6 kg. • The MarsBird-V achieves the durability of 420 s and a payload efficiency of 0.43 g/W. [ABSTRACT FROM AUTHOR]

There is an increasing demand in the space industry for methods allowing for gyroless spacecraft attitude determination to either provide low cost sensor fusion or to recover from gyro failures. The gyroless configuration used in this paper is a combination of magnetometers and Sun sensors. Sun sensors are subject to measurement discontinuities during solar eclipses. Gyroless attitude sensing is known to increase sensitivity to measurement noise and model uncertainty. This paper presents a new Laplace Particle Filter with a Singular Value Decomposition to enhance attitude determination performance under these circumstances, while also preventing filter degeneracy. An orthogonal Procrustes problem is formulated to minimise the Wahba cost function and determine the attitude matrix from the Sun and magnetic field vectors and a third virtual predicted measurement vector, which is also available during Sun occlusion. The proposed filter is compared to a regularised particle filter approach, which handles particle filter degeneracy. Both filters are applied to the attitude determination of a small Earth observation satellite using three Sun sensors available during the sunlit phase and three orthogonal magnetometers. A Monte Carlo numerical simulation analysis demonstrates that the proposed particle filter significantly outperforms the regularised particle filter in terms of accuracy and robustness to the Sun sensor measurement discontinuities during solar eclipse and is also more accurate when both types of sensor measurements are available. • Gyroless attitude determination was obtained with Sun sensors and magnetometers. • A Laplace particle filter uses a Procrustes algorithm to solve Wahba's problem. • A Procrustes problem is formulated to approximate the maximum a posteriori attitude. • Estimation accuracy and robustness is significantly enhanced during solar eclipse. [ABSTRACT FROM AUTHOR]

In recent years, the Moon, as the closest celestial body to the Earth, has become the foremost target of human exploration outward. Some lunar exploration missions and studies have shown that ice may exist in the polar regions of the Moon. Water is very important for human survival. The ice mining and utilization is based on the understanding of the sublimation mechanism. Simulation of ground-based tests is an effective means to explore the sublimation of ice. Therefore, in this paper, the thermal environment of Shackleton Crater and the particle size distribution of typical lunar soils are analyzed as the basis of parameter settings for the ground test. Four operating conditions are established for ground tests, including the influence of heat sink temperature, vacuum level, particle size distribution and dry soil density. The larger the radiation heat flow, the higher the heat sink temperature, the smaller the vacuum degree, the larger the average particle size, resulting in a larger sublimation rate. The decrease of density makes the sublimation process faster and then slower. The sublimation process proceeds fastest when the dry soil density is 1200 kg/m3. The heat sink temperature has the greatest effect on the sublimation process, and the vacuum has the least effect. When the radiation heat flow ranges from 39.01 W/m2 to 45.63 W/m2 at the heat sink temperature of 120 K, the maximum sublimation rate is 5.79E-15 kg/s, the maximum sublimation area is 1.93E-9 m2 and the maximum sublimation heat flow is 1.59E-8 W. However, the effect of vacuum degree and particle size distribution on the sample temperature variation is not significant, due to the small order of magnitude of the sublimation heat flow compared to the radiation and conduction heat flows. This paper gives a range of parameters for ground-based simulation experiments and points out the microscopic changes of the ice sublimation process, which provides ideas for the next ground-based experiments and the exploitation of water-ice resources on the Moon. • A ground-based environmental test setup that is closely linked to the real lunar surface environment. • Analysis of the effects of temperature, vacuum and particle size distribution on the sublimation process of water ice. • Calculation of sublimation area, sublimation percentage and sublimation heat flow. • Energy analysis of sublimation process. [ABSTRACT FROM AUTHOR]

Visual instance segmentation ability is one of the effective means to promote the autonomy and intelligence of space agents. However, due to the limited airborne computing capacity, current methods are difficult to deploy on space agents, because these methods are developed based on GPU. To this end, this paper proposes a novel framework of instance segmentation by introducing depth information and combining traditional computer vision techniques with an object detection method. This framework provides a new idea for the implementation of instance segmentation. The experiment results show that the proposed method achieves a real-time performance under a common laptop CPU platform. In addition, thanks to the introduction of depth information, the proposed method can obtain better segmentation results compared to Mask R–CNN and SOLOv2 in complex scenes (poor illumination and occlusion). Finally, because the semantic information is obtained by the object detection method in this paper, the model training adopts a weakly supervised manner from bounding-box annotations, which can reduce various costs of data labeling to a certain extent. • A depth-aided instance segmentation fuses mask extraction and object detection. • Real-time performance on consumer grade laptop CPU; friendly to airborne platforms. • Better segmentation results in complex scenes compared with Mask-RCNN and SOLOv2. • Training process adopts a weakly supervised manner from bounding-box annotations. [ABSTRACT FROM AUTHOR]

Lunar sun-synchronous roving is an affordable and promising manner of acquiring continuous power supply and warmth in robotic polar exploration. Sun-synchronous route planning is challenging to meet the spatiotemporal constraints of lunar topography and time-varying illumination. Moreover, in pursuit of spaceflight feasibility and safety, the safety-optimal planning method adapting to spatiotemporal maps remains to be addressed. This paper takes the first step to investigate this issue, in which the distribution of roving speed is considered a critical factor. Three-dimensional lunar spatiotemporal maps are constructed based on topography and illumination analysis. A spatiotemporal planning algorithm named MSST-A* is proposed, which takes a multi-speed node expansion mechanism, thus supporting speed distribution optimization. Afterward, route evaluation functions involving motion, energy, and integrated risk, together with the optimal heuristic function, are designed for safety-optimal planning. Two comparative tests are conducted, and results indicate that the proposed approach is practical and superior in terms of motion safety and energy reliability. Besides, an integrated-risk-minimal route circumnavigating the Shackleton Crater at the south pole in three lunar days is solved, demonstrating the existence of long-term lunar sun-synchronous routes with good engineering significance. • The paper is innovative in the following bullet points: • Safety-optimal lunar sun-synchronous route planning is firstly studied. • MSST-A* for optimizing the spatiotemporal distribution of speed is raised. • Polar sun-synchronous circumnavigating routes of fine safety are solved. • The existence of the sun-synchronous route of engineering reliability is proved. [ABSTRACT FROM AUTHOR]

This paper introduces a flexible framework to enable high-fidelity simulations of the fuel slosh dynamics for verification of the Guidance, Navigation and Control (GNC) algorithms. The sloshing phenomenon must be tackled during the control system design phase and the GNC software run and validated within appropriate simulation frameworks. Equivalent mechanical models can well approximate sloshing under specific assumptions, allowing the derivation of the multibody vehicle's equations of motion; however, while suitable for analysis and control synthesis purposes, this approach presents several limitations for implementation as a simulation model. In this work, the multibody plant embedding slosh dynamics is modeled by means of the DLR's Vertical Landing Vehicles Library (VLVLib) written using the object-oriented Modelica modeling language. The advantages of using Modelica are explored throughout the paper. Given that control synthesis, analyses and simulation campaigns are usually performed in the Matlab/Simulink environment, a core point of the proposed solution is to achieve a good synergy within the latter and the Modelica plant model with the least readaptation burden for the developer. The strategies to achieve a full integration and verification tool chain are illustrated. The potential of this approach is demonstrated via two different test cases: a lunar landing scenario of the ALINA spacecraft developed by PTS, and CALLISTO reusable rocket demonstrator. • A flexible framework to enable high-fidelity simulations of the fuel slosh dynamics for verification of the GNC algorithms is introduced. • The spacecraft model embedding the slosh dynamics is implemented exploiting Modelica language. • A GNC testing framework is obtained with the integration of the causal Simulink environment with the compiled Modelica model. • Two test cases are presented: the ALINA spacecraft lunar landing scenario and CALLISTO reusable Vertical-Takeoff–Vertical-Landing demonstrative mission. [ABSTRACT FROM AUTHOR]