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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]

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]

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]

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]

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]

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]

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]

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]

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]

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]

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]

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]

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]

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]

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]

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]

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]

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

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

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

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

Local linearization is often used to enable analytical methods in spacecraft dynamics, navigation, and control. However, the linear approximation becomes inaccurate at large distances from the linearization point and/or when the underlying dynamical system is strongly nonlinear. As a result, linearization-based methods like the finite time Lyapunov exponent (FTLE) can underestimate stretching in phase space. Alternatively, this paper presents a semianalytical method to quantify nonlinear stretching: the eigenpairs of higher-order tensors are used to bound the maximum stretching of a state deviation over time. The tenors in question are derived from the Taylor series expansion of a nonlinear solution flow about a reference trajectory, i.e., the state transition tensor (STT) expansion. It is well known that the eigenvalues of the Cauchy–Green tensor (CGT) bound the stretching of state deviations over time for a linear system, and the largest eigenvalue can be related to the FTLE. This paper will present higher-order, nonlinear analogs to the CGT and the FTLE to more accurately bound nonlinear stretching for applications in guidance, navigation, and control. • Linearization is often used to study the time evolution of state deviations. • Linear methods are insufficient for strongly nonlinear systems or large deviations. • Tensor eigenpairs can be used to bound the nonlinear growth of state deviations. [ABSTRACT FROM AUTHOR]

How does a country embark on large-scale R&D programs? Among R&D programs, this is an intriguing question for large scale Science and Technology (S&T) projects such as space programs, because of the necessity for large resource commitments, and the difficulty of being accountable to the taxpayers when the outcome is not always known. It can be assumed that engineers, scientists, bureaucrats, and politicians play respective roles. This paper aims to shed light on the specific role of politics, through the examples of policies of and decisions by NAKASONE Yasuhiro. Nakasone was the minister for science and technology (1959–1960, 1972), and later became the prime minister (1982–1987). He played a crucial role in establishing the Science and Technology Agency and the National Space Activities Council in the 1950s, which would be the foundational institutions responsible for space policy for the next half century. As prime minister, he decided to participate in the Space Station program. Within his vision for Japan for the 21st century, he emphasized the need for radical change in economic policy, education, defense and S&T investment. He was one of the few politicians that had vision to promote S&T and recognized the strategic value to the nation. He was able to explain to the nation why Japan should promote the development of a launch vehicle, when the engineers faced development difficulties in doing so. Thus, he was influential both in the policy layer and in promoting specific programs to support his national vision. Politics intersects with S&T projects because of the uncertainty and tradeoffs associated with these efforts. Japan did not address space programs for national security and commercial space for a long time. It is a politician's role to create institutions, to allocate resources and, more importantly, to explain to the taxpayers why policy measures or programs are necessary. R&D investment is a societal activity: it reflects the social, political, and economic situation, as well as being important for diplomacy and international relations. Bearing in mind that it would be misleading to attribute the initiation of largescale S&T projects to the leadership of a single individual, this paper will attempt to show the driving forces behind, and fundamental objective of Japan's space policy, in addition to the historical shift in that policy by examining the influence of this notable politician. • Nakasone played the key role in establishing institutions promoting space development in post-war Japan. • Nakasone made the best use of close relationships with the U.S. in promoting Japan's space activities. • Nakasone made decisions on space policy process, resource allocation with vision, thus integrating the stakeholders. [ABSTRACT FROM AUTHOR]

In order to help control the spread of space debris and junk, deorbit sail devices have been used for deorbiting techniques; however, this approach is still considered risky due to potential failure or malfunction as it need to be exposed to the harsh space environment for a prolonged period of time, particularly in low-Earth orbit. China has made new strides forward in managing space junk, as it has successfully unfolded a 25-square-meter deorbit sail developed by the Shanghai Academy of Spacecraft Technology in-orbit for the payload capsule of a recently launched rocket. Accordingly in this paper, a full life-cycle modelling based on Monte Carlo method is developed to reveal the damage behavior of the 25-square-meter deorbit sail surface under the combined effects of atomic oxygen erosion and deorbit sail defects. The simulation results show that as the atomic oxygen fluence increases and the material defect width widens, the interaction effect of the two factors becomes more and more pronounced, which decisively influences the erosion rate of the sail surface. Morphological characterization of aluminized film PET material, activated silanisation modified material and plasma-polymerization coating material before and after atomic oxygen erosion combined defects is carried out and the atomic oxygen erosion of the deorbit sail surfaces is quantitatively assessed. The assessment results show that compared to the PET material and activation silanisation surface modification material, the plasma polymeric coated material showed much better flexibility, folding resistance and irradiation resistance to prevent geological fold and to shield from the atomic oxygen erosion in space. The atomic oxygen erosion combined defects assessment method studied in this paper provides valuable reference data for the subsequent serialisation of deorbit sail products and other typical lightweight film materials for space applications. • Introduction of Monte Carlo method for material surface damage simulation analysis. • Compare the performance of off-orbit sail surface materials under different protection treatments. • Standardize the process of space flexible film material performance evaluation system. [ABSTRACT FROM AUTHOR]

As part of the Artemis era of space exploration, space agencies will be working together with their industry partners to establish systems and infrastructure that enable sustained Lunar missions and develop capabilities for Mars. Lockheed Martin has a longstanding history of designing, building, and operating systems for deep space applications, from Viking to Orion. Lockheed Martin develops concepts, trades, and early system designs for Moon and Mars exploration to enable future missions. In 2016, Lockheed Martin presented a vision for achieving crewed exploration of Mars. Known as Mars Base Camp (MBC), this design reference mission envisioned a crewed vehicle in Martian orbit from which astronauts could perform excursions to Phobos and Deimos and perform telerobotic exploration of the Martian surface, including sample return. This concept served as an "existence proof" for a novel, practical, and affordable path to enable human exploration of the Martian system. In 2017, an update was published that included the production of propellant from water and a single-stage fully reusable Mars lander concept. Then in 2021, Lockheed Martin introduced a nuclear thermal propulsion (NTP) configuration of Mars Base Camp, a technology that will be greatly enabling for crewed missions to Mars. This paper matures the design of the NTP configuration of Mars Base Camp and explores the Mars mission capabilities that the NTP technology will enable. These investigations include optimized trajectories for faster transit to Mars and an analysis of abort scenarios that are made possible by the high thrust and performance of NTP. The results provide insight into the performance and capabilities that are realized when NTP technology is matured and integrated into a Mars transit vehicle. The paper will also describe initial surface infrastructure capabilities that significantly enhance sortie science operations, including mobility and power. The resulting updated Mars Base Camp serves as an example of what is possible with current technologies and technologies under development. The MBC reference mission builds on the capabilities that will be developed and demonstrated through Artemis missions, enabling a viable, sustainable long-term Mars exploration program, with the first crewed Mars mission possible in the 2030s. • Matures design of nuclear thermal propulsion configuration of crewed Mars mission • Explores Mars mission capabilities and trajectories enabled by nuclear propulsion • Describes initial infrastructure to support sustainable long-term Mars exploration [ABSTRACT FROM AUTHOR]

The Korea Pathfinder Lunar Orbiter (KPLO), officially named as Danuri, is the Republic of Korea's first spacecraft to orbit the Moon. The goal of KPLO is to secure core deep space technology for future space exploration in Korea and contribute to lunar science by gathering data from the lunar orbit for one year. This paper presents the results of KPLO's Lunar Orbit Acquisition (LOA) phase in detail, with a particular focus on the operational results of Flight Dynamics (FD). During the LOA phase, the KPLO FD team experienced many expected and unexpected issues. These issues were mitigated by making real-time updates to the LOA phase Design Reference Mission (DRM). The paper also presents the results of the Orbit Determination (OD) and performance of each Lunar Orbit Insertion (LOI) maneuver conducted during the actual flight operation. The importance of considering operational constraints when designing the DRM is emphasized by presenting lessons learned based on actual flight experiences. Unlike the original KPLO LOA phase DRM, KPLO successfully achieved the final mission orbit using only three LOI burns, instead of the planned five LOI burns and an Orbit Trim Maneuver (OTM). Despite the extensive revisions made to the DRM, all mission requirements were still fulfilled during the final lunar orbit insertion, taking into account the significant operational constraints. • Real flight operation results of the Korea Pathfinder Lunar Orbiter's lunar orbit acquisition phase are presented. • Orbit determination and lunar orbit insertion maneuver performance are analyzed based on actual flight data. • Unexpected issues during the actual flight operation were mitigated through real-time updates on the design reference mission. • Considering real-world operational constraints while designing the design reference mission is very critical. [ABSTRACT FROM AUTHOR]

OMOTENASHI is a CubeSat that was launched by a NASA SLS rocket. Its mission was to demonstrate that a CubeSat could make a semi-hard landing on the Moon. A chemical propulsion system which provides high thrust and can be re-ignited is large and difficult to install on a CubeSat. Therefore, a small cold-gas jet makes orbit corrections with only a small ΔV. Moreover, a rocket motor with a high Isp and structural mass ratio is used for a lunar landing deceleration. The spacecraft consists of an orbiting module, a rocket motor, and a surface probe. This paper analyzes the trajectory design of the Moon landing and a method that increases the landing success rate under the tight constraints of the CubeSat. To reduce fuel and power consumption, the maneuver conditions during the lunar transition to reach the target landing site are evaluated based on several error factors. This study considers the trade-off between vertical and horizontal landings. Because active control is not possible for decelerating a rocket motor for the lunar landing, the optimal altitude at the starting maneuver is determined by considering such errors as attitude and timing. Finally, the paper discusses mission feasibility and the landing success rate. • OMOTENASHI was developed as the world's smallest moon lander that was launched by a NASA SLS rocket. • OMOTENASHI uses Rocket Motor to decelerate Surface Probe for a semi-hard landing. • This paper shows an innovative trajectory design method of the moon landing for a CubeSat as a secondary payload. • For a high landing success rate, the CubeSat must accommodate many constraints in selecting the landing site. • This method can be utilized for future CubeSat missions under tight resource constraints, such as a thrust and a power consumption. [ABSTRACT FROM AUTHOR]

ESA Solar Orbiter was launched from Cape Canaveral in collaboration with NASA on the February 10, 2020 to start its journey around the Sun, with the commissioning phase ending in April 2020. After commissioning, ESA set a tighter pointing requirement goal to further improve the performance in Fine Pointing Mode, in preparation of the upcoming start of Science Phase. The first part of this paper outlines the end-to-end process which allowed - in just six months – to update the Fine Pointing Mode Design, re-tune the controller with refined assumptions taken from flight data and robust methods and execute full validation of the new AOCS Software on the Avionics Test bench at Airbus Defence and Space. The second part of the paper describes the upload in flight of the new Software and the analysis of the on-orbit pointing performance, which showed a significant improvement with respect to the previous one, thus contributing to the success of the Mission's scientific goals. • Solar Orbiter Payload Instruments required a better pointing jitter performance (Relative Pointing Error) after launch • Root cause analysis on flight telemetry to identify driver contributors to pointing error • Fine Pointing Mode controller design and tuning updated with refined inputs based on flight data and with robust control methods • Controller uploaded in flight significantly improved the Spacecraft pointing performance [ABSTRACT FROM AUTHOR]

Safe operation of orbital platforms relies heavily on stability of fluid flows in microgravity, which are strongly affected by capillary forces. The paper considers multiphase fluid seepage through a porous medium. Questions of instability of the interface between liquids are discussed. The effect of capillary instability is shown, which was revealed during the processing of experimental data on the imbibition of a porous structure in microgravity conditions during parabolic flights. A mathematical model is proposed to describe the process of multiple imbibition of a porous medium under conditions of variable gravity. The model takes into account hysteresis, various models of capillary pressure are used during imbibition and drainage. The paper also considers a method for modeling the instability by transforming the constitutive equations. The method allows modeling additional blurring of the imbibition front. The results of modeling based on this method for capillary imbibition of a porous medium under microgravity conditions, as well as for the case when chemical interactions occur between flowing fluids, are considered. • The instability of the phase boundary during seepage is considered. • Flows are studied controlling safe operation of spacecraft fluid supply systems. • A new mathematical model for describing the multiple imbibition is proposed • The seepage with chemical interactions between the phases is considered. [ABSTRACT FROM AUTHOR]

Academic and industry experts build their careers on accurate and precise definitions and meanings of domain-specific terminology, whereas novice practitioners and lay-persons often conflate similar or similar-sounding terms to describe concepts in the same domain of interest. This is common in the space sector when individuals of different academic or practitioner backgrounds discuss "commercial" topics. In English, as in other languages, the word "space" is often followed by words such as "industry," "industry segment," "industrial base," "market," "ecosystem" and others. The practice of casually using words with ambiguous meanings may often be appropriate for informal discussion if it facilitates the communication and exchange of ideas and all parties generally understand the implications of the casual usage. This practice, however, is problematic for detailed and analytical applications requiring precision and order of thought. Grounded in contemporary organization theory and economic practice, this paper provides background discussions of different levels of analysis that characterize embedded hierarchical social systems of organizations as they pertain to space. This paper then recommends a hierarchical taxonomy for defining specific terms that describe organizations and activities within the space domain for use by practitioners, analysts and academics requiring more precise language to communicate their market analysis ideas. This taxonomy places specific terms within each level of analysis discussed and provides definitions of each, with the intention of stimulating critical discussion and potential modifications or improvements. • Using phrases like "space industry" can be vague and problematic for precise analytical tasks. • For accurate space terminology, understanding analysis levels in organizational systems is key. • This paper offers a taxonomy for clearer language in space market analysis communication. • Proper terminology aids in strategic decisions for everyone in the space sector. • Feedback on this proposed taxonomy is encouraged for potential enhancements. [ABSTRACT FROM AUTHOR]

With NASA's Artemis program currently underway, the necessity to have lunar structures that can house astronauts for months at a time is now greater than ever. This paper investigates current deployable lunar habitat concepts and finds a paucity of detailed papers in the literature. The hybrid deployable lunar habitat architecture considered here went through many design iterations to simplify and support its manufacturing. A novel deployment mechanism is developed utilizing the release of pressurized gas for an autonomous deployment sequence. A prototype was fabricated that demonstrated a successful proof of concept, leaving a robust foundation for future research. • delom Hybrid autonomously deployable structures for lunar habitats are studied. • Such structures have solid elements plus an inflatable component that deploys after landing. • Astronauts arrive at the Moon and have the structure waiting for them. • A key design challenge is the deployment mechanism. • A scale model is designed and manufactured to test concepts. [ABSTRACT FROM AUTHOR]

In recent years, thanks to advancements in commercial hardware technology miniaturization, small satellites have become more and more important and effective in the Space ecosystem. In previous papers it has been demonstrated that future smallest observation systems, operating at a lower altitude than traditional systems, have the potential for comparable or better performance, much lower overall mission cost, lower risk, shorter schedules, lower up-front development cost, more sustainable business model, mitigating the problem of orbital debris. Flying in very low Earth orbit requires the addition of a propulsion system capable of providing drag compensation. Thanks to its high specific impulse, electric propulsion is potentially well suited for this task. However, the use of an electric thruster often requires significant power that can be provided only with deployable solar panels. To minimize the drag, the solar panels should be aligned with the satellite orbital velocity. At the same time, to maximize the power input, the solar panel surfaces should point toward the Sun. In the majority of cases these two conditions are not met simultaneously, so an optimal trade-off is needed. This topic is investigated in this paper by the use of an orbital model of a generic satellite with deployable solar panels. The model is able to simulate any circular orbit of the satellite in any day of the year and calculate both the power input and the drag from the solar panels based on their inclination. An equivalent specific impulse is also defined taking into account the thrust that the electric thruster has to provide only to push the area of the solar panels required to power the propulsion system. The results show, as expected, that the sun-synchronous Dawn-Dusk (or Dusk-Dawn) orbit is the preferred one for the electric propulsion as the velocity vector and the Sun vector are almost orthogonal. In the other cases, at very low Earth orbit, even relatively small inclinations of the solar panels respect to the velocity vector have a significant impact on the drag produced, increasing the propellant consumption and the required power. Nevertheless, it is shown also that if the solar panels are aligned with the velocity, the power lost can be kept reasonable. The same analysis could also apply to a satellite in low Earth orbit that does not employ an electric thruster but still needs deployable solar panels because of other sources of high power consumption (e.g. the payload). • Deployable solar panels in Very Low Earth Orbit generally produce drag. • If aligned with velocity, ideally no drag is produced but power available reduces. • Penalty for best case (dawn-dusk) is 1–12%, for worst case (noon-midnight) is 47%. • For other panels positions part of the power absorbed and motor thrust are wasted. • Equivalent specific impulse and net power can be determined. [ABSTRACT FROM AUTHOR]

Existing exoatmospheric guidance laws become ineffective when a thrust loss fault (like one or more engines unexpectedly shut down) occurs in a launch vehicle, which makes the nominal orbit unreachable. This paper presents a new abort orbit determination process to determine an abort orbit which the launch vehicle is able to reach. It involves solving at most two optimization problems and the abort orbit may be circular or elliptical. We will present how to transform the nonconvex optimization problems into difference of convex (DC) programming problems which can be solved by existing methods with guaranteed convergence. The main contribution of this paper lies in proposing a relaxation technique to convexify a nonconvex control constraint and the combination of the introduction of new constraints and a relaxation technique to deal with the strict terminal constraints. Validity of the relaxation technique is ensured by theoretical analysis. A unique feature of the transformation process is that the feasible sets defined by both the control constraint and the strict terminal constraints are all enlarged. This brings a critical benefit that the resulting DC programming problems can be very efficiently solved. Thus, an abort orbit determination algorithm, which is based on DC programming, can be designed to very reliably and efficiently determine an abort orbit, which will be demonstrated by numerical examples. • A new concise and practical abort-orbit determination process is proposed. • Relaxation techniques and the introduction of new constraints are properly combined. • Validity of the relaxation techniques are theoretically proved. • An abort orbit can be determined by solving at most two DC programming problems. [ABSTRACT FROM AUTHOR]

Space Trafic Management (STM) has gained relevance in regional and global strategies when there is more significant congestion in space. Ongoing discussions in the United Nations Committee on the Peaceful Uses of Outer Space (UNCOPUOS) and other fora are reflecting the essence of STM in Europe, in the United States and others. In the US, the 2018 Space Policy Directive-3 on National Space Traffic Management Policy has become a focal eleemnt. More recently, STM was underlined in the European Union – European Space Agency Space Council Orientations of November 2020 as part of the European ambitions to achieve strategic autonomy. In February 2021, the European Commission launched STM as a flagship project under its' Action Plan on Synergies' between the civil, defence, and space industries, aiming to reinforce the European industry's competitiveness. In February 2022, a joint communication on a European Union approach on STM was published. In this regard, STM entails several technological, socio-economic, and regulatory dimensions with significant security implications as reflected in national and regional strategies. This paper focuses on a critical element: the cooperation between the civilian and the military and how organizations effectively address the dual-use aspect concerning STM development under the current legal and regulatory framework. Moreover, it examines how the dual-use element of STM influences space security of space technologies. Successful strategies with tangible outcomes require effective cooperation between the civilian and military actors. Hence, this paper seeks to identify potential legal and regulatory barriers and proposes solutions to accommodate divergent strategic interests and needs in STM future strategies. [ABSTRACT FROM AUTHOR]