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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]

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]

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

Interest in the exploration of other planets has led to the success of Ingenuity (Mars helicopter)'s first flight in the Martian atmosphere. Despite this achievement, the flight of aircraft under ultra-low air density, such as Mars, is still limited. A flapping-wing aircraft that mimics nature's flyers can provide an alternative solution to conventional aircraft. Successful flapping flight on Mars, where the air density is very thin, requires an understanding of aerodynamics, and prediction of the dynamic changes of the flapping wing under very-low air density. In this paper, we constructed a simple dynamic model that can predict the changes in aerodynamics and dynamic behavior under various low air densities. The predicted changes were compared with the measured ones in a chamber that could simulate the low air density with a modified flapping-wing system based on our flapping-wing micro-air-vehicle (FW−MAV) models, the KUBeetle. Both prediction and measurement showed that when the air density decreased, the flapping frequency tended to increase due to drag reduction, even without consuming additional power. The model reasonably predicted a large decrease in the lift for the 10 % air density, where the flapping frequency increased by 10 %. Under the 50 % and 25 % low air densities, the measured lift coefficients did not change much. However, under the 10 % air density, the lift coefficient decreased by about 16 % due to less wing rotation and wing twist. While the aerodynamic power was predicted to decrease significantly due to the large decrease in aerodynamic forces under low air density, an increase in the angular velocity and peak angular acceleration caused an increase in the inertial power. Therefore, the amount of reduction in the measured mechanical power consumption under the 10 % air density was relatively small, compared to that under the 100 % air density. This work may provide insights into the design of an FW−MAV for the ultra-thin air density, such as Mars. • The proposed dynamic model reasonably predicted the flapping motion under low air density. • The measured flapping frequency increased by about 15 % under the 10 % air density, consuming about 9 % less power. • The reduction in measured power consumption under low air density was small, due to the large inertial power. • The measured lift coefficient decreased by about 16 % under the 10 % air density, due to less wing Q4 rotation and wing twist. [ABSTRACT FROM AUTHOR]

Cislunar L 4 or L 5 axial orbit families are continuous asymmetrical periodic orbits from the figure-8 orbit around L 4 or L 5 Lagrange point of the Earth–Moon system to the halo-family around the Moon. They provide special locations of lunar relay satellites for the lunar south pole due to their offset locations. In this paper, we investigate the orbital characters and applications of cislunar L 4 and L 5 axial orbits. A construction method of cislunar L 4 and L 5 axial orbit families is proposed according to their topological feature in the torus space. Then, lunar relay performances, including coverage and eclipse, are analyzed for different cislunar L 4 and L 5 axial orbits. Numerical computation indicates that 3:1 resonant L 4 and L 5 axial orbits are the optimal options in terms of the stability, coverage and eclipse avoidance. Lunar relay performances of a two-satellite constellation deployed in these orbits are compared with that in a near rectilinear halo orbit (NRHO). Numerical results show that the constellation of axial orbits can provide longer contact duration to the lunar equator and southern hemisphere than that of the NRHO. Finally, a station-keeping strategy and a transfer design of cislunar L 4 and L 5 axial orbits are developed, and their fuel costs are studied. • Cislunar axial orbits are studied for relay mission of the lunar exploration. • The 3:1 cislunar axial orbit is selected for the eclipse avoidance. • A constellation on axial orbits has longer contact duration than NRHOs. • The corresponding station-keeping method and transfer design are proposed. [ABSTRACT FROM AUTHOR]

Surface movement exploration techniques are an important approach to further the understanding of small celestial bodies. Hopping movement is employed to take advantage of the weak gravitational field of small celestial bodies, facilitating long-distance mobility of the rover. The rover exhibits a strong attitude-orbit coupling characteristic during the hopping movement. Different contact attitudes contribute to the challenge of determining the rebound trajectory, rendering it difficult for the rover to maintain continuous movement towards the target position along the nominal trajectory. In this paper, a dual-stage control method for continuous surface hopping movement is developed, wherein the control is applied separately during the flight stage and the collision stage. During the flight stage, applying attitude control enables the rover to establish contact with the surface in a specific attitude, ensuring accuracy in movement direction and creating favorable conditions for the control process in the collision stage. The flywheel is controlled in the collision stage to correct the state change and energy loss due to the collision, enabling the rover to attain the desired hopping state. Considering the energy consumption of the rover, the minimum hopping velocity sequence and angular velocity sequence are designed based on the hopping angle and heading angle. The intricacies and uncertainties inherent in the surface environment of small celestial bodies pose challenges for the rover during hopping and flying. In response to the issue, a finite-time adaptive sliding-mode control law is devised. The control law facilitates the rapid adjustment of the rover to the target attitude in flight, demonstrating strong robustness. In addition, by integrating the changes in contact point velocity with the impulse contact model, the accuracy of the rover's rebound state updates is enhanced. Finally, a set of simulations is performed, and simulation results indicate that the dual-stage control method can achieve a final landing position error of less than 0.2 m for a continuously hopping. The finite-time adaptive sliding-mode control law can stabilize the rover's attitude within 10 s. Conducting 500 Monte Carlo simulations of the continuously hopping rover in three small celestial body surface simulation environments shows that the position endpoint error is within 0.3 m. • An innovative control method for continuous hopping on the surface of small celestial bodies with low gravity is proposed. • The attitude-orbit coupling challenge is addressed through a staged control. • A sequence of hopping states with minimum velocities is devised. • A finite-time adaptive sliding-mode control law is designed to effectively address unknown disturbances. • A corrected sliding-mode control law is developed for precise hopping state updates. [ABSTRACT FROM AUTHOR]

In this article, a fault-tolerant tracking control technique is developed for the ascent of hypersonic vehicles (HSV) experiencing wing damage and time-varying state constraints. The difficulty of control problem lies in handling state constraints associated with both state and time, while determining the unknown control directions of the multi-input multi-output (MIMO) nonlinear system. The presence of constraints in both state and time complicates the design of the control algorithm. This paper proposes the construction of a novel state transition function to establish a new unconstrained system, thereby eliminating the conservative feasibility conditions of Barrier Lyapunov Function (BLF) and integral BLF methods when dealing with time-varying constraints. The fusion of time-varying scaling functions and Nussbaum-type functions not only resolves the issue of unknown control directions in the MIMO nonlinear system but also enhances the transient and steady-state performance of the closed-loop system. Neural Networks are employed to approximate the unknown nonlinear functions. Using Lyapunov stability theory, it is proven that state constraints are not violated, and all signals in the closed-loop system are bounded. Simulation results demonstrate the merits of the designed fault-tolerant control (FTC) algorithm. • A unified fault-tolerant control scheme for the ascent hypersonic vehicle is designed. • Wing damage and function constraints on states are considered. • Nonlinear mapping eliminates the feasibility conditions in barrier Lyapunov functions. • The unknown control direction issue in multivariable nonlinear systems is resolved. [ABSTRACT FROM AUTHOR]

The rapidly growing number of objects encircling our planet is an increasing concern. Collisions between these objects have already occurred and pose a potential threat in the future, resulting in the creation of countless debris fragments. In particular, the Low Earth Orbit (LEO) region is densely populated and highly contested. This underscores the critical importance of space surveillance in this area. Moreover, the utilization of Medium Earth Orbit (MEO) and Geosynchronous Earth Orbit (GEO) is also on the rise. To ensure the safety and functionality of operational satellites, it is paramount to accurately determine and continuously monitor the orbits of space objects, mitigating the risk of collisions. Precise and timely predictions of future trajectories are essential for this purpose. In response to these challenges, this survey paper provides a comprehensive review of various methods proposed in the literature for Orbit Determination (OD). It also identifies research gaps and suggests potential directions for future studies, emphasizing the pressing need for adequate Space Situational Awareness (SSA). [Display omitted] • Admissible Region is a robust, reliable cornerstone method for orbit determination. • Machine learning transforms orbit determination, boosting accuracy and efficiency. • Explainable AI is crucial for transparency and trust in orbit determination. • The recent focus on cislunar regime demands novel orbit determination methods. • Undervalued, space-based orbit determination is crucial for accurate cislunar missions. [ABSTRACT FROM AUTHOR]

In the framework of Space Traffic Management (STM) the growing number of space debris and the consequent increased risk of collisions have become a topic of great interest. Moreover, the difficulty of detecting and characterizing a Resident Space Object (RSO) in terms of orbital parameters makes this research field very challenging. In most cases the available observations have a very short duration and an accurate estimation of the orbit is very difficult. This paper introduces the Initial Orbit Determination via Artificial intelligence (IODA) algorithm, a reliable and accurate orbit estimator combining metaheuristic and artificial intelligence frameworks to perform precise orbit determination from Very Short Arcs (VSAs). Starting from an initial orbit estimate derived through an Artificial Bee Colony (ABC) algorithm, a Physics Informed Extreme Learning Machine is employed to increase the accuracy of the predicted orbit. Results on simulated data have shown the robustness of the algorithm to random parameters initialization and a percentage of success of 89% of retrieving the reference orbit with an error under 1 km and 0.01 km/s on the initial target's state. • Description of an AI-based orbit determination tool from very short arcs. • Combination of meta-heuristic and machine learning techniques for accurate orbit estimation. • Tests on simulated data for robustness against random parameters initialization. • Montecarlo simulation on simulated data for robustness against random mission scenarios. [ABSTRACT FROM AUTHOR]

Laser ablation technology enables contactless debris removal, providing the advantages of safety and cost efficiency. To deorbit space debris using laser ablation, the attitude motion of the debris must be controlled in advance. Although a previous study derived an attitude control procedure, it empirically determines control gains and is easily affected by assumed preconditions. To solve this problem, this paper proposes an improved attitude control law using laser ablation. Previous methods use feedback control based on the angular velocity error and error quaternions to obtain a reference torque. By analogy with a magnetic attitude control law, this study designs a reference torque based on the angular momentum errors and presents theoretical condition for designing the control gains. Numerical examples using the proposed control law are conducted to control the target attitude from a random initial rotation of 1 rpm to an arbitrary attitude, which verify the effectiveness of the proposed method. Furthermore, the position, direction, and magnitude uncertainties of laser irradiation are introduced to numerically examine their effects on the control accuracy under environmental disturbances. • Conventional attitude control method by laser ablation is improved. • The effective control gains can be set theoretically by applying the concept of magnetic attitude controller. • The designed control gains depend on the rate of target rotation. • The method is effective under laser uncertainties and environmental disturbances. [ABSTRACT FROM AUTHOR]

The Mach 6-capable turbine based combined cycle engine is a potential first-stage propulsion for TSTO mission. In order to improve thrust-to-weight ratio, an innovative tandem combined ramjet engine-FABRE (Fan Augmented Air-Breathing Ramjet Engine) is designed utilizing a versatile multi-mode combustor working at both high-speed ramjet mode and low-speed turbocharged mode. Employing the strut as mixer and flame holder in multi-mode combustor presents theoretical advantages. Nonetheless, a thorough investigation into its practical feasibility and combustion characteristics is essential. This paper experimentally verifies the feasibility of strut approach in multi-mode combustor at various speeds and evaluates the effects of different strut schemes on flow field structure, thermodynamic characteristics, and combustion performance via simulation. The experimental results affirm that employing the strut scheme as a combustion organization method enables the multi-mode combustor to operate effectively under diverse inflow conditions at both high and low speeds. Simulation results further validate that these strut solutions consistently demonstrate superior combustion performance. Additionally, the influences of the strut geometry on heat release and high-temperature area are highly consistent with its effects on combustion efficiency. At low speed, wider struts lead to increased combustion efficiency, while at high speed, they initially rise before declining with further widening. Regardless of speed, increasing strut slant angle consistently reduces efficiency. At low speed, widening the strut lowers the total pressure recovery coefficient, peaking at 45° with varying strut slant angle. Conversely, at high speed, changes in strut width minimally impact, while increasing strut slant angle reduces the coefficient. These conclusions provide valuable insights for the optimization of combustion organization in the design of combined ramjet engine with a wide range of operating conditions. • An innovative Mach 6-capable combined ramjet engine is designed to enhance the thrust-to-weight ratio for TSTO. • Experimentally validating the strut scheme for combustion organization in multi-mode combustor. • Combustion characteristics are numerically investigated using different strut schemes. [ABSTRACT FROM AUTHOR]

This paper presents a novel six degree-of-freedom guidance and control algorithm for free-flyer robots equipped with a robotic manipulator operating on the International Space Station. The proposed algorithm combines a sliding mode controller for rotational motion and a model predictive controller for translational motion. A novel prediction-twisting sliding mode controller (PT-SMC) is introduced to stabilize configuration changes induced by the manipulator and correlated disturbances and to achieve a trade-off among control law flexibility, robustness, and accuracy. The proposed PT-SMC is shown to have improved stability compared to a first-order sliding mode controller. Model predictive control is investigated as a position controller to achieve formation flight between multiple free flyers in a constrained environment, and a leader–follower approach is used to move a triangular formation of robots. A control allocation problem is solved to address the subdivision of attitude and position controls between twelve thrusters. Simulation results of multiple robots approaching a target object demonstrate the improved performance of the proposed control approaches. • A Predictive Twisting Sliding Mode Controller was designed and verified. • An effective guidance and control algorithm was developed for formation flight. • The combination of the proposed control strategies was demonstrated. [ABSTRACT FROM AUTHOR]