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

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

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

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

Caves on Mars are a tantalizing target for exploration, as they could harbor evidence of extinct or extant life, clues to the planet's geological history, and even potential future human habitation. However, the inherent challenges of navigating unknown and challenging terrain, coupled with limited communication capabilities, pose significant obstacles for robotic exploration in these caves. This paper presents the results of an effort to evaluate different operational concepts for a mission dedicated to exploration of caves on Mars. We conducted a series of analog exploration experiments in lava tubes on Earth, testing two hypotheses: (1) that two robots are more effective than one, and (2) that high levels of autonomy are more effective than low levels of autonomy. Our findings suggest that two robots are indeed more effective, except in low autonomy cases, where more operational resources are required. We also found that full autonomy is more efficient than low autonomy, as it enables quicker exploration and detection of targets of interest. However, the low autonomy cases benefited from the operator input to acquire higher quality data, an area of autonomy requiring further development. This paper provides insight into the design of our experiments, as well as detailing the results and implications for the design of future missions to explore caves on Mars. By shedding light on the operational concepts tested and their corresponding outcomes, we contribute to the knowledge base required to formulate optimal strategies for the realization of successful cave exploration missions on the Red Planet. [Display omitted] • We report on analog experiments for a multi-robot Martian cave exploration mission. • Full autonomy was shown to be more efficient than low autonomy (operator waypoints). • Two robots were shown to be more effective than one for scientific exploration. • Work is needed to improve autonomy for detailed scientific measurements. [ABSTRACT FROM AUTHOR]

There is growing momentum, as articulated through the visions of space agencies, to establish permanent settlements on planetary bodies such as the Moon and Mars. One key hazard that structures will be subjected to will be hypervelocity impacts of micrometeoroids and orbital debris. Shielding against such impacts is essential to ensure the safety of structures and their human inhabitants. There has, however, been limited research on such shields and no field applications to date. In order to inform the composition and design of protective shields for planetary structures, recourse can be made to protective shields utilised in spacecraft and space stations that have benefitted from years of research and development. This paper reviews the existing shielding practices for spacecraft and space stations as a starting point for the shielding of planetary structures. Shield configurations are firstly introduced. Both experimental tests and numerical simulations required to evaluate the performance of the shields are then discussed. Analytical prediction models for shields are reviewed, and the predictions of these models are compared with test results of corresponding shields from the literature. Finally, the effect of parameters such as projectile diameter, impact velocity, impact angle, and shield geometry on the overall shield performance is analysed. This paper ultimately provides technical knowledge about the shielding systems for spacecraft and space stations that can be used as a point of reference for scientists and engineers to develop shields for structures on planetary bodies. • Review of shielding systems on spacecraft and ISS for MMOD protection. • Analysis of theoretical models for predicting protective shield performance. • Recommendations provided for shielding systems of Lunar and Martian structures. [ABSTRACT FROM AUTHOR]

This paper presents a comprehensive framework for designing in-space propulsion systems, integrating four criteria: global propulsive performance, environmental impact, cost efficiency, and architectural reliability. The study focuses on the emerging class of Orbital Transfer Vehicles to illustrate the application of this method. By examining the synergistic potential of OTVs and greener propellants, the paper addresses different mission scenarios, including LEO, GEO, and lunar missions, with both scientific and commercial objectives. The proposed framework aims to go beyond traditional cost-centric approaches, offering a more complete evaluation method for early design phases. A case study comparing three liquid bipropellant options, pressure-fed MON-3/MMH, 98%-HTP/RP-1, and self-pressurizing N 2 O/Ethane, demonstrates the utility of the tool. Findings suggest that scientific missions benefit most from 98%-HTP/RP-1, while traditional propellants remain preferable for cost-driven commercial missions to GEO and the Moon, though greener alternatives are competitive for less demanding LEO missions. This innovative framework aims to guide the selection of propulsion systems to achieve greener space missions, aligning traditional performance figures with environmental responsibility. • The transition towards greener propellants introduces new requirements beyond pure performance. • A holistic early-design framework is proposed to identify optimal future propulsive solutions. • This approach integrates environmental impact and reliability alongside traditional cost-efficiency and performance targets. • Greener propellants demonstrate a competitive advantage over legacy ones for specific space mission profiles. [ABSTRACT FROM AUTHOR]

Compared with Earth, the lunar surface's environment is more complex, which suggests that the nuclear energy Helium–Xenon closed Brayton cycle (NHCBC), supplying electricity to the lunar base, may operate under off-design conditions. A thermodynamic model coupled with the lunar surface environment is established in this paper to analyze the effects of lunar environment exposure, including solar radiation exposure, lunar dust toxicity, and meteorite impact, on a given NHCBC. The results show that the output capacity of a given NHCBC is suppressed during the daytime and enhanced during the nighttime. The lunar dust will improve the system's output capacity at night but will also significantly inhibit the output capacity of the cycle during the daytime, resulting in a fluctuation of 824 % of the system without lunar dust. The output fluctuation of NHCBC is reduced after meteorite impact, while the output capacity of the NHCBC is suppressed, destroyed even worse. The optimization of cycle parameters can alleviate the influence of the lunar environment exposure on NHCBC to a minimal extent. This paper aims to analyze the impact of lunar surface environment exposure on the cycle parameters and output work characteristics of NHCBC and to provide some guidance for the subsequent application of nuclear energy in the lunar base. • A mathematical model of a nuclear energy cycle for a lunar base was established. • Lunar surface environment threats pose economic and safety risks to the energy cycle. • Optimizing a single-cycle parameter is unlikely to enhance energy cycle robustness. • Reducing radiator area to enhance cycle robustness sacrifices thermal efficiency. [ABSTRACT FROM AUTHOR]

This paper analyses the circumstances and events that led the Italian scientific community to encourage the astronautical sector in Italy at the end of the Second World War. Particularly important was the contribution of the A.I.R., Associazione Italiana Razzi, the Italian Rocket Association, founded in 1951 for the purpose of extending rocket and space culture. At the end of World War I, modern rocket science was already going forward in a number of countries. In the U.S.A., Robert Goddard launched the first liquid propellant rocket in 1926; in the U.S.S.R., Mikhail K. Tikhonravov launched the first Soviet liquid rocket in 1933; in Germany, Hermann Oberth, Wernher von Braun, Karl Becher and Eugen Sänger worked in the Peenemünde and Trauen centres on developing the V-2 rockets. In Italy, during the fascist regime activities focused on the aeronautical sector. In 1935 the Royal Academy of Italy organized the Fifth Volta Congress in Rome and since 1938 eminent scientists Gaetano Arturo Crocco, Antonio Ferri, Luigi Crocco and Antonio Eula had been conducting research in the subsonic and supersonic aerodynamic wind tunnels at Guidonia, near Rome, at the D.S.S.E., the Air Force Ministry's High Directorate for Studies and Tests. In order to increase technical development, in 1937 Adolf Hitler and Benito Mussolini created the Achse Berlin-Rom auch in der Luftfahrttechnik, the Rome-Berlin Axis in the sector of aeronautical technologies. This pact brought about a closer link between Italy and Germany with the exchange of technicians and experts. At that time, the most prominent organizations in the field were the two major aeronautical associations: the Lilienthal-Gesellschaft, in Germany, and the A.I.D.A., the Italian Association of Aerotechnics in Italy. At the end of World War II, Italy was in a critical economic and social situation that also had repercussions on the aeronautical sector; some scientists, Antonio Ferri, Luigi Crocco, Angelo Miele and Carlo Riparbelli among them, were forced to move to the U.S.A. It was precisely at this tragic time that other scientists, among them Gaetano A. Crocco and Antonio Eula, were courageous enough to found a new cultural association called the A.I.R., the Italian Rocket Association, to inject new vigour into the Italian rocket and space sector. This paper recalls these years, before and after World War II, in order to highlight the work done by the A.I.R., a small association that succeeded in giving important impetus to development in the astronautical field, and not in Italy alone. Antonio Eula and Eugen Sänger, whose firm friendship went back to the years of World War II, were among those who, in 1951, founded the I.A.F., the International Astronautical Federation of which Sänger was the first President. • History of the Italian Astronautical Sector after the Second World War. • The A.I.R., Associazione Italiana Razzi, the Italian Rocket Association. • Foundation of the A.I.D.A., Italian Association of Aerotechnics. • Creation of "The City of the Air – Guidonia". [ABSTRACT FROM AUTHOR]

This paper presents a mission concept designed for satellite-to-cell connectivity through a fractionated CubeSat approach, aiming to fulfill the objectives set by the BlueWalker-3 specifications. The envisioned system consists of 41 CubeSats, each adopting a 16U form factor, to provide dedicated coverage across specified sub-regions, as a reasonable solution. The pre-Phase-A design capitalizes on commercially available components with proven flight heritage, ensuring reliability and cost-efficiency. A genetic algorithm (GA)-based method has been introduced to tackle the complex challenge of designing a large antenna's folding pattern, with an analysis on how this design influences the overall system cost. The study conducts an in-depth analysis of system architecture and operational scenarios. Employing Monte-Carlo simulations, it assesses the Life Cycle Cost (LCC) and the system's resilience to failures, comparing it with similar configurations differing in CubeSat size and number. The results demonstrate that a fractionated approach significantly enhances the mission's scalability and robustness, providing substantial operational flexibility. The comparison highlights a cost-effective configuration within fractionated systems, emphasizing the proposed system's potential advantages. Furthermore, the paper addresses the environmental implications of deploying a fractionated spacecraft constellation, especially concerning astronomical studies. It underscores the CubeSats' minimized brightness and their reduced potential to interfere with telescopic observations, positioning the system as a more compatible option with nighttime sky studies. This consideration is particularly pertinent amidst increasing concerns over the impact of satellite constellations on astronomical research and night sky quality. • Introduction of a novel mission concept employing fractionated spacecraft for satellite-to-cell connectivity, showcasing an innovative design utilizing 41 CubeSats. • Comparative analysis demonstrating the proposed system's performance and cost advantages over the BlueWalker-3 mission, with findings supported by Monte-Carlo simulations. • Detailed discussion on the use of commercially available off-the-shelf components, highlighting the system's feasibility and reliability. • Exploration of the fractionated system's scalability and robustness, revealing potential for enhanced mission flexibility and economic efficiency. • Addressing the impact on astronomical observations, presenting a significant reduction in brightness and interference, enhancing compatibility with observational studies. [ABSTRACT FROM AUTHOR]

This paper first formulates a novel long-term beam power allocation (BPA) problem to tackle the harmful co-linear interference issue in the geostationary earth orbit (GEO) and low earth orbit (LEO) co-existing satellite system. This BPA problem intends to optimize the long-term weighted sum rate of the LEO system while ensuring that GEO user's received interference from the LEO satellite system is lower than a pre-fixed threshold. To solve it in a real-time manner, a deep reinforcement learning (DRL) framework based on the proximal policy optimization (PPO) algorithm is proposed, named as drlBPA. In addition, for the existing most relevant baseline, the fractional optimization (FO)-based BPA scheme, on the one hand, this paper improves it via a greedy strategy to fully exploit time resource. On the other hand, to further reduce the computational complexity stemming from its iterative solving procedure, a deep neural network approximation scheme is also developed. Simulation results demonstrate that (i) The trained DRL model of the proposed drlBPA scheme has good convergence and generality performance. (ii) Compared with the three FO-based benchmarks, the drlBPA scheme not only achieves the highest throughput of the LEO system within a significantly reduced computation time, but also yields the best system stability. • Maximizes the LEO system throughput while ensuring the GEO system service quality. • Builds a PPO algorithm based deep reinforcement learning framework. • Constructs two benchmark schemes: improved FO (IFO) and the DNN-accelerated IFO. • Achieves the highest LEO system throughput with the lowest computational complexity. • Yields the best system stability compared with the benchmarks. [ABSTRACT FROM AUTHOR]

An Environmental Control and Life Support System (ECLSS) meets the environmental and metabolic needs of the crew by providing a breathable atmosphere, potable water, food, and waste management. It also handles metabolic outputs, including carbon dioxide, respired and perspired humidity, urine, which includes dissolved solids, perspired solids, feces including both solids and water, and metabolic heat. While the International Space Station (ISS) has been continuously occupied, many space habitats beyond low Earth orbit will operate with intermittent human occupancy. Without the crew, mission control may suspend ECLSS operations and transition the system to a dormant state. However, no holistic guidelines exist that identify the gaps and needs to perform research on ECLSS dormancy. This paper reviews the literature on space habitat dormancy and relevant terrestrial examples. Results show that dormancy may have a larger impact on humidity control, water recovery subsystems, aqueous-based thermal control systems, and waste management, whereas the impact on atmosphere revitalization subsystems may be minimal. The paper also identifies potential anomalies that could adversely affect the ability to maintain ECLSS health during dormancy. We utilize the findings to recommend guidelines for ECLSS technology development and operations. • Aqueous-based ECLSS subsystems are likely to experience large number of anomalies during dormancy. • Atmosphere revitalization subsystems will likely experience less anomalies during dormancy. • The article serves as a guideline to identify system and subsystem-level needs for dormant ECLSSresearch and development. [ABSTRACT FROM AUTHOR]

OMOTENASHI is a 6U CubeSat launched by the first NASA SLS rocket. Its mission was to demonstrate that a CubeSat can make a semi-hard landing on the Moon. Tight resource constraints make it difficult to mount a heater and a radiator in a CubeSat, so passive thermal control is a prominent design feature of OMOTENASHI that resolves the technical difficulty. In November 2019, a thermal vacuum test was conducted. This paper discusses the thermal design and the test results. OMOTENASHI was launched by SLS rocket on November 16, 2022. After the SLS rocket separation, OMOTENASHI's attitude control malfunctioned so it could not generate power from its solar cells. Eventually, communications were lost and have not been restored. Brief telemetry data confirmed that the temperature monitor functioned correctly for 30 min. This paper reports the flight results of the thermal status during the first ground station pass. Also, a comparison between the analysis model and telemetry data is shown. Finally, based on its estimated trajectory, the future thermal profile of the spacecraft and the result of the search operation are discussed. • OMOTENASHI was developed as the world's smallest moon lander that was launched by a NASA SLS rocket. • Passive thermal control design without a thermal controller is adopted by using a white paint and a graphite sheet. • Passive thermal control design can be utilized for future CubeSat missions under tight resource constraints. • This paper introduced the experiment and analysis method that achieves the thermal design for lunar missions by CubeSat. • The flight data after the launch of the SLS rocket showed the feasibility of this passive thermal control design. [ABSTRACT FROM AUTHOR]

This paper discusses the processes which should be taken to establish debris indices that assess the orbital environmental impact of spacecraft or missions in orbit. The first process is establishing a methodology and criteria for evaluating an impact on the orbital environment. This study proposed an assessment method to evaluate the cumulative environmental changes during an evaluation period. It evaluated the impact of an uncontrolled LC mission abandoned in orbit. This study used Near-Earth Orbital Debris Environment Evolutionary Model (NEODEEM), an orbital environmental evolutionary model that Kyushu University and JAXA developed. The second process is the debris index formulation. This study considered four steps to establish the debris index. To reveal the establishing process of a debris index, this study used debris indexes as candidates which consider expected number of fragments generated from an object. The calculated debris index values were compared to the impact on the environment assessed by the evolutionary model to see if there was any correlation. Further, this study conducted numerical simulations of un-controlled large-constellation satellites abandoned in low earth orbit in different scenarios, and of active debris removals of large intact objects by NEODEEM. This paper discusses the applicability of proposed environmental assessment methods and the debris index. • Two processes to establish a debris index are proposed. • A methodology and criteria for evaluating environmental impacts is discussed. • Four steps to formulate a debris index is discussed. • Correlations between debris indexes and environmental impacts are compared. [ABSTRACT FROM AUTHOR]

Many proposals are being made for cleaner, more sustainable forms of energy production. Terrestrial solar photovoltaic farms (SPFs) could potentially deliver large quantities of energy to the grid, although these are limited to daytime use. The output from these SPFs could be enhanced, particularly around dawn and dusk, by the use of orbiting solar reflectors (OSRs) in near-polar orbit. These would reflect an image of the solar disk, or solar image (SI), onto the SPFs to augment their energy output. Pointing requirements are therefore to ensure that reflected sunlight is delivered to the terrestrial SPF, avoiding the losses incurred by an offset of the SI and the SPF itself. The SI would typically be of order 10 km for a reflector in a 1000 km orbit. Given the potentially large size of the reflectors, this presents a challenge for the attitude determination and control system (ADCS) to ensure that the maximum quantity of energy can be delivered to a SPF, typically requiring large control moment gyro actuators. In addition, there exist numerous sources of error in the ADCS which can cause further degradation in the quantity of energy delivered to the SPF. These errors can manifest in the resolution of the various sensors, flexible structural modes, manufacturing inaccuracies, and misalignments due to vibration during launch. This paper will investigate the effects of pointing error sources (PES) on the reflector ADCS and so on the quantity of energy delivered to the SPF. With the application of a PD controller with feedforward compensation, and the set of noise characteristics defined in this paper, numerical simulations will show the typical losses in energy delivered to SPFs of 0.015% when the model accounts for PES in onboard sensors, actuator uncertainty and flexible structural modes. • Orbiting solar reflectors. • Sensor and actuator noise modelling. • Establishing the energy delivery losses due to noise in OSRs and finding the minimum losses using PD controllers. [ABSTRACT FROM AUTHOR]

In-Orbit Demonstration (IOD) missions represent flight opportunities to acquire data on spacecraft and space environment and to demonstrate new techniques at system, subsystem, and payload level. Realizing this type of mission means accepting a greater level of risk. However, when very compact and relatively simple platforms like CubeSats are adopted, this becomes admissible since costs can be kept low. Distributed Synthetic Aperture Radar (SAR) is a novel Earth Observation (EO) technique for microwave imaging. It uses multiple cooperating receivers and suitable processing to enable performance and products that exceed those of the single components of the system. RODiO (Radar for earth Observation by synthetic aperture Distributed on a cluster of CubeSats equipped with high-technology micro-propellers for new Operative services) is an IOD of the distributed SAR concept. The mission is based on four, 16 U, CubeSats, that embark a receiving-only X-band SAR instrument able to collect bistatic echoes exploiting a monostatic SAR as an opportunity illuminator, i.e PLATiNO-1 mission. Distributed SAR demonstration is thus pursued in RODiO combining the principles of both multistatic and bistatic SAR. The paper, moving from mission goals, individuates different families of EO products that RODiO can demonstrate, e.g. in the contest of ground motion, digital elevation models, and ship detection. The main result is the definition of the RODiO system requirements for product generation that drive satellites and payload design. • RODiO is an IOD mission of DSAR principle based on four 16U CubeSat receivers. • The combination of mono and bistatic images enables the generation of new products. • These products provide advantages with respect to the monostatic-only counterparts. • Preliminary performance assessment of multi-baseline products is shown in the paper. • Requirements for product generation have been derived to drive RODiO design. [ABSTRACT FROM AUTHOR]

Experiments and numerical simulations were conducted across a range of contraction ratio to investigate the detonation combustion characteristics of liquid aviation kerosene/air rotating detonation combustor. The research indicates that extremely low and high contraction ratio adversely affect the stable operation of the two-phase rotating detonation combustors. At a contraction ratio that is too low (19.03 % in the RDC discussed in this paper), the atomization and mixing properties of the two-phase fuel are poor, with a lower reaction pressure in the fresh mixture, hindering the initiation of detonation waves. On the other hand, a contraction ratio (85.7 % in the RDC discussed in this paper) that is too high enhances the energy of the detonation wave; it triggers oblique shockwave feedback, preventing the fresh mixture from entering the combustor and thus leading to the quenching of detonation waves. With an increase in the contraction ratio of outlet configurations, there is a noticeable rise in both the peak pressure and combustion efficiency of detonation waves. In contrast, the wave height gradually decreases, and the wave velocity shows a pattern of initially increasing and then decreasing. Along the axial position inside the two-phase rotating detonation combustor, the upstream detonation wave is caused by the reflected shockwaves increasing the pressure of the upstream fresh mixture. In contrast, the downstream detonation wave results from increased fresh mixture pressure and enhanced fuel atomization and mixing due to the increased contraction ratio of outlet configurations and wave feedback. So, the detonation wave exhibits a strong-weak-strong distribution characteristic. • Three-dimensional numerical simulation and experiment in a RDC with separate feeding of kerosene and air is achieved. • Two atomization and mixing structures ("Kerosene" and "Enhanced Kerosene" region) is found. • An optimal contraction ratio size enables the RDC to achieve the best operational characteristics. [ABSTRACT FROM AUTHOR]

The series of celebrated and grand accomplishments and progress of mankind in space exploration make people more intelligent, quixotic, imaginative, creative and realistic. Also, these imply vision for the future, further exploration and taking advantage of the Universe, for human benefits and sustainability, gaining scientific knowledge and developing engineering structures to that end. The most modern Earth and Universe observation spacecraft to date are now equipped with large lightweight and flexible structures, such as antennas, telescopes, and extendable elements, including the incorporation of more complex and larger appendages with high precision for scientific applications. These demand more stringent structural integrity requirements to minimize performance degradation due to structural damages. In addition, spacecraft structures weight and their associated manufacturing costs should be minimized, most appropriate materials should be utilized, assembly costs should be reduced and fewer parts should be opted, or one-piece more complex configuration parts should be produced. Various sensors, actuators and control technology could be employed to reduce structural damage. Associated with these, appropriate sensors configuration and strategies for their deployment are required to accurately detect failures in the most critical areas of the structure. Spacecraft structural integrity is subject to a series of prevailing loading due to thermo-structural dynamics and environmental effects from design, manufacturing, launching and operation during their pathways to meet a multitude of objectives under a myriad of requirements and constraints. To that end, the objective of the present paper is to provide an overview on spacecraft structural integrity covering a host of selected concepts, design, methods and engineering to maintain structural integrity of space vehicles and associated sub-structures. In spite of citing examples drawn from small satellite situations under vibro-acoustic and environmental thermal loading, the overriding issues addressed could be extended to more general situations. An infographic is presented to map out the state of affairs to be addressed for structural integrity considerations. [Display omitted] The present paper is to provide an overview on spacecraft structural integrity covering a host of selected concepts, design, methods and engineering to maintain structural integrity of space vehicles and associated sub-structures, subject to a series of prevailing loading due to thermo-structural dynamics and environmental effects from design, manufacturing, launching and operation during their pathways to meet a multitude of objectives under a myriad of requirements and constraints, such as various loading and environmental effects. The Rationale and Key Technical Structural Integrity Challenge are first identified. In spite of citing examples drawn from small satellite situations under vibro-acoustic and environmental thermal loading, the overriding issues addressed could be extended to more general situations. An infographic is presented to map out the state of affairs to be addressed for structural integrity considerations. [ABSTRACT FROM AUTHOR]

Musical expression is a characteristic of humanity on Earth. This paper highlights the historical relationship between music and spaceflight, including its use in the representation of space, and its role in space culture. The paper documents circa 57 musical instruments: from a miniaturized harmonica to a 1.5 m Digeridoo that are known to have been taken into space – usually as part of a crew members personal allowance. Due to mass and volume restrictions numerous classical and ethnographic instruments have been 'spaceflight-adapted', in addition to the use of digital technologies for personal musical pleasure and public relations. Furthermore, it charts how access to personalised music playlists and shared musical experiences have been deemed to be vital by crew, promoting individual wellbeing including a connection to home, to each other and playing a part in the marking of success, and tragedy. The value of music making in space is briefly described leading to the proposition that as humanity ventures with greater regularity to Low Earth Orbit and beyond, consideration of how music can be effectively utilised to support individual and crew wellbeing, whilst contributing to the cannon of human artistic expression is warranted. • Music is a characteristic of humanity and social connection. • Circa 57 musical instruments have been taken into space. • Numerous instruments have been 'spaceflight-adapted'. • Personalised music playlists are deemed critical by crew to link to 'home'. • Music expression warrants consideration for missions beyond Low Earth Orbit. [ABSTRACT FROM AUTHOR]

On-Orbit Servicing is increasingly crucial with the growing number of spacecraft. The foundation for On-orbit servicing is space docking mechanisms. As On-Orbit Servicing tasks become more complex, the demands for docking capture mechanisms increase in both function and performance. Therefore, it is crucial to design a docking mechanism with a large tolerance. Among various docking capture mechanisms, the multi-claw capture mechanism has the advantages of low impact and reliable capture. This paper analyzes the Common Berthing Mechanism and a docking mechanism with an underactuated two-degree-of-freedom claw-type capture mechanism as its core is proposed. Subsequently, this paper provided a detailed design of the mechanism based on the principle of underactuated capture. Furthermore, this article conducts numerical analyses of key mechanisms such as the capture mechanism and guide mechanism in the design process, verifying the functionality of the mechanisms. After completing the analysis of capturing dynamics, the tolerance ability of the mechanism was also verified, various key parameters were determined, and the parameters of components such as the torsion spring were also determined. Finally, the prototype of the proposed docking mechanism is developed, and the experiment is completed to validate its functional principle and tolerance capability. • This paper presents a docking mechanism for large spacecraft (weighing over 8 tons). • The mechanism employs an underactuated two-degree-of-freedom claw-type capture mechanism. • Simulations and simple laboratory tests were performed for design and validation. • The maximum lateral tolerance of this mechanism reaches 220 mm. • The developed port has reached a TRL of about 4. [ABSTRACT FROM AUTHOR]

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

This paper investigates the orbital pursuit strategy for spacecraft swarms, where both pursuers and evaders possess limited delta-v maneuvering capabilities. Given that swarms can consist of numerous members, ranging from dozens to thousands, the swarm-to-swarm problem has significant coupling effects that cannot be solved by merely superimposing one-to-one or few-to-few problems. Otherwise, one would cause a dimension explosion in decision-making space. To avoid this issue, the swarm orbital pursuit problem is divided into two coupled problems: swarm target allocation and orbital maneuver. First, considering the relative orbit state between pursuers and evaders and estimated orbital maneuver costs of pursuit, the swarm target allocation problem is formulated into three coupled sub-problems, i.e., target allocation order, target allocation for sub-swarms, and target allocation for individuals. These sub-problems are solved by the presented double-layer contract net protocol algorithm to maximize task completion efficiency. Then, the orbital maneuver problem is decomposed into two coupled problems to maximize pursuit benefits, i.e., long-range rendezvous, and close-distance game maneuvers. For the long-range rendezvous, the optimal Lambert algorithm is developed to efficiently obtain the discrete optimal pursuit time and corresponding orbital transfer maneuver. To address close-distance games more proficiently, a hybrid orbital maneuver algorithm, combining the proximal policy optimization and the Lambert algorithm is proposed. Finally, comparison simulations verified the effectiveness and superiority of the proposed method. • Satellite swarm pursuit problem is transformed to the coupled problem with swarm target allocation and orbital maneuvers. • Swarm target allocation is formulated as a coupled problem with target order, allocations for subgroups and individuals. • An improved DLCNP algorithm is proposed to pursue a reasonable target allocation solution for higher task efficiency. • Optimal Lambert algorithm is developed to obtain the optimal maneuver strategyfor long-range rendezvous. • A hybrid algorithm combining PPO and the Lambert algorithm is presented to get closed-loop control for close-distance games. [ABSTRACT FROM AUTHOR]

As for blunt-body hypersonic vehicles, the extreme drag and aero-heating pose great challenges to the conceptual design. Different from the conventional configurations, a blunt body slotted with a channel is proposed in this paper. The channel is further combined with a jet on the shoulder to reduce drag and aero-heating. Simultaneously, based on the position of the separation point and vortex edge, a modified effective body approach is developed to accurately investigate the drag reduction mechanism. Specifically, the more slender the modified effective body is, the more obvious the drag reduction is. The simulation results show that the spike-channel-jet can push the shear layer on the spike away from the wall. In this way, the position of shear reattachment point moves downstream, which can enlarge the recirculation zone and weaken the shock-shock interaction. Additionally, the reattachment shock is pushed far away from the blunt body, which is also weakened. For drag, net power is introduced to evaluate the practical value of spike-channel-jet for drag reduction. Compared with a simple spike (spike length L / D = 1) of the same size, the spike-channel-jet can reduce the total drag by 24.68 %, the peak pressure coefficient of the blunt body by 38.54 %, the peak temperature of the blunt body by 65.75 %, and the peak temperature of the spike by 10.63 %. The maximum net power can reach 13.30 kW. Considering the fluid-thermal interaction, the findings reveal the influence of the spike length and jet mass flow rate on aerodynamic performance, by modified effective body further analyze the drag reduction mechanism with the jet, and validate the different reasons for the peak temperature and pressure of the blunt body. • Drag and aeroheating reduction mechanism of the spike-channel-jet concept under hypersonic conditions is investigated by the numerical simulation. • The influence factors are analyzed and compared comprehensively. • Combined spike-channel-jet concept is a promising drag and aeroheating reduction technique. [ABSTRACT FROM AUTHOR]

Ground-based full-scale physical experiments on spacecraft have become an essential part of ground simulations. However, achieving a vertical translational degree of freedom for zero-gravity simulation has become a key technological bottleneck for six-degree-of-freedom ground simulations. The combined active–passive vertical zero-gravity system discussed in this paper can achieve zero-gravity simulation in the vertical direction; however, it exhibits deficiencies such as inaccurate modeling, low control precision, and significant control delays. Addressing the issue of inaccurate modeling, this study uses interpretable ODE neural networks to identify the dynamic model of the system. For resolving the problems of poor control precision and significant delays, this study proposes a hybrid control algorithm based on the ODE dynamic model. The controller includes nominal model feedforward, PID feedback, and learning-based feedforward components. Finally, comparative experiments are designed to validate the proposed model-based hybrid control algorithm. Compared to PID-Feedforward controller, the hybrid controller exhibits a higher response speed, while compared to model-based feedforward controller, it exhibits a better control accuracy. Therefore, the proposed hybrid control algorithm based on the ODE dynamic model effectively enhances the control response speed and improves the control precision of the system. • A satellite simulator that realizes vertical zero-gravity simulation by combining active and passive methods. • Higher precision modeling of nonlinear dynamics by interpretable ODE neural networks. • Smaller control error compared with model-based feedforward controller, using model-based hybrid control algorithm. • Faster response speed compared with PID-Feedforward controller, using model-based hybrid control algorithm. [ABSTRACT FROM AUTHOR]

The EPFL Spacecraft Team (EST) is working on the Constellation of High-Performance Exospheric Science Satellites (CHESS) mission, which aims to launch a constellation of two CubeSats in 2027. These CubeSats will study the chemical composition of Earth's exosphere using time-of-flight mass spectrometry. To downlink the collected scientific data, the satellites will utilise a custom X-band transceiver, presented in this paper, operating in the amateur frequency range of 10.45 GHz to 10.50 GHz. The proposed transceiver architecture employs Software-Defined Radio (SDR) technology and can be easily adapted for different use cases. This modular approach has the potential to be suitable for future spacecraft missions, such as active debris removal or lifetime extension services, where flexible communication systems are required. • Hybrid SDR architectures enable versatile telecom systems for space missions. • IOD/IOV missions let students follow the entire life cycle of a space project. • Short development/prototyping cycles maximise the lessons learned in student teams. [ABSTRACT FROM AUTHOR]

NASA and other space agencies have been planning the construction of a space station in lunar orbit called the Gateway. They have also begun studies for its resupply missions using an HTV-X. There are several ways to transition to NRHO, a candidate orbit. In-Direct Transfer (IDT) has a short transition period of seven days, but the required ΔV is inefficient at more than 400 m/s. The Weak Stability Boundary (WSB) transfer requires a small ΔV, less than 100 m/s, but the transition period is more than 100 days. This paper proposes the Perilune Rendezvous Method (PRM), recently discovered in 2022. Its transition period is two weeks longer than IDT, but the period is much shorter than WSB. The required ΔV is 40 m/s lower than IDT but still larger than WSB. The PRM has an intermediate ΔV and a transition period between IDT and WSB. Since fuel consumption and transition duration requirements depend on the mission, the PRM expands the scope of future cargo transportation. To adopt the PRM in a mission, an efficient trajectory correction plan was found based on a Monte Carlo analysis that considered navigation and control errors. It confirmed that the PRM could rendezvous with Gateway with higher accuracy than IDT. • Perilune Rendezvous Method (PRM), an intermediate transfer method of ΔV and transfer period is well-suited for transferring cargo to Gateway. • PRM extends the transition period to 18 days, reducing ΔV and maximizing cargo better than In-Direct transfer for manned transport by Orion. • Because the PRM does not use the solar gravity, PRM has no constraints for the launch window as WSB trajectory employed in NASA's Capstone mission. • The spacecraft can correct errors during the lunar orbit in the PRM. The position and velocity errors upon arrival at the gateway can be kept small. • Since the requirement of the fuel consumption and the transition period depends on each mission, the PRM can expand the scope of cargo missions. [ABSTRACT FROM AUTHOR]

The increasing space debris population in critical orbits due to spacecraft failure dictates the need for action. On-Orbit Servicing (OOS) has been proposed as a method for mitigating this trend by repairing existing space assets. The development of large servicing spacecraft has been given significant attention. In this paper, we are approaching the problem with a new paradigm for a servicing spacecraft by proposing a CubeSat class servicer equipped with a one-degree-of-freedom (DoF) robotic arm. Considering the OOS context, we choose the Orbital Replacement Unit (ORU) installation on a target spacecraft as a challenging benchmark task for the proposed servicer configuration. To carry out this task, we formulate and compare performance of four controllers to achieve coordinated control of both the CubeSat and the robotic arm for the installation task. We start with a basic PD controller and progress to the more advanced impedance controller, Model Predictive Control (MPC) and Model Predictive Impedance Control (MPIC) designs. The controllers are evaluated and compared across several relevant metrics for the ORU installation and their merits for practical implementation are discussed. • Comparative analysis of model-based controllers against traditional control methods. • Monte Carlo simulations used to test robustness of the controllers. • Novel, mechanically simple CubeSat servicer for simulated ORU installation. • MPC, PD controllers show similar success rates, MPC requires higher control efforts. [ABSTRACT FROM AUTHOR]