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and complex time delay in space teleoperation make the environment force parameter available to force reconstruction is outdated and cannot realize accurate force rebuilding.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

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

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

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

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

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

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

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

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

As we continue exploration of the Lunar and Martian surfaces and push farther into unknown and unstructured environments, employing a team of distributed heterogeneous robots will increase the odds of success by enabling more complex task planning that utilizes the complementary capabilities and synergy of the team members. A requirement to reap these potential benefits is effective cooperation and collaboration between the members of a robot team. Defining a metric for effective cooperation is a difficult task that will depend on the composition of the team, the task to be performed, and the working environment. This paper establishes a method for determining the reward criteria (figures of merit) that can be used for training the robot swarm through reinforcement learning techniques. A hierarchical framework of rewards is used which, at the lowest level, measures the success of an individual robot in performing its task. The success of all robots performing different subtasks is then measured using the Quantified Cooperation Assessment (QCA) metric which was introduced in our previous research of multi-robot collaboration. Finally, the mission-level success and overall reward is determined by weighting each task using its priority within the overall mission context. A common reward for each of the robotic teammates is then applied within the learning process, which emphasizes group performance over that of an individual and encourages cooperative behavior. This cooperation framework is trained in a grid-based environment representing an exploration mission on a planetary surface by a heterogenous team of robots consisting of a landing craft, a traditional rover, and several small agile robots. The robots are trained concurrently, but with individual policies developed for each agent resulting in a decentralized control scheme. Once trained, the control policies were evaluated in several test environments consisting of novel terrain maps and regions of interest. An average success rate of 90 % was seen in the test environments demonstrating the robustness of the trained policies. The robots have been trained to not only complete the mission but also perform it in a cooperative manner as well as show the scalable and resilient behavior. • Robotic team trained to locate volatiles in a simulated planetary environment • Cooperation among team members assessed with a quantifiable metric • Success achieved in novel environments not used in training • Decentralized control provided robustness to adding and removing team members [ABSTRACT FROM AUTHOR]

The Moon is a highly valued destination for human space exploration because it is close and it contains a wide array of lunar and Solar System science targets, which include resources that could be used to sustain humans on the lunar surface. These resources have the potential to enable sustainable human space exploration, develop a vibrant cislunar economy, and directly benefit society here on Earth. However, recent rhetoric about the importance and value of these resources has used the term as if we know they are reserves. An immediate and vital next step has not yet been realized to define the reserve potential of such resources, and that is designing and implementing a coordinated international lunar resource evaluation (prospecting) campaign. This paper outlines the issues that need to be addressed by such a campaign, including why it needs to be international in nature and how the coordination can be started and evolved, as well as exploring the benefits that would come from prospecting on the Moon. • A distinction between lunar resources and lunar reserves is defined. • Roadblocks in the in-situ resource value chain are highlighted. • Reserve potential of lunar resources requires a coordinated prospecting campaign. • The prospecting campaign must be international due to the scale of the task. • Coordination can start at the grassroots level with current and scheduled missions. [ABSTRACT FROM AUTHOR]

This paper presents the high-fidelity physical modeling of a Thrust Vector Control (TVC) system operated by Electro-Mechanical Actuators (EMAs) for Reusable Launch Vehicles (RLVs). In contrast to simplified, often linear, models, high-fidelity physical models enable a better assessment of the actual system performance and a deeper verification of the on-board software robustness against disturbances, unmodeled dynamics or degradation. This aspect is essential to enable the reusability concept as driven by long-term reliability requirements. Moreover, critical systems like Fault Detection, Isolation and Recovery (FDIR) logic, whose robustness and efficacy are often difficult to prove, can be thoroughly tested without requiring hardware experiments with eventual invasive modifications for fault injections. This work captures the whole dynamics of the EMA and TVC components, including the power drive electronics, the electrical motor, and the mechanical transmission for the EMA, as well as the mechanical properties of the engine nozzle acting as a load. The resulting differential algebraic equation system is modeled in the Modelica acausal object-oriented modeling language. An ad-hoc framework is implemented to guarantee flexibility and modularity, which allows models with different fidelity levels to be easily exchanged to achieve the simulation objectives and needed accuracy level. The impact of the different models on the closed-loop TVC dynamics is analyzed in both the frequency and time domain, and then benchmarked with a realistic RLV mission. The results show that the high-fidelity physical models provide a better understanding of the more complex effects governing the TVC dynamics and can, in turn, effectively improve its requirement definition process. • The physical modeling and simulation of an EMA-based TVC system is presented. • High-fidelity physical models help assessing TVC system control software performance. • EMA component models with incremental fidelity levels are proposed. • EMA and TVC system models are integrated in a Modelica-based framework. • Different closed-loop responses are analyzed in the frequency and time domains. [ABSTRACT FROM AUTHOR]

The Mars rotorcraft is an aerial exploration platform capable of repeated take-offs and landings to extend the range of rover explorations. NASA's Ingenuity helicopter is a full y functioning coaxial helicopter for technology demonstration. Increasing the rotor diameter or the number of blades can enhance the load capacity of the coaxial helicopter but would make it difficult to be folded. A more suitable configuration for Mars exploration is the multi-rotor aircraft, which has a stronger load capacity with the same rotor diameter. Although the aerodynamic performance and operational capabilities of the Mars multi-rotor aircraft have been partially considered, mechatronic system design and autonomous flight control have not been fully addressed. Harbin Institute of Technology and China Academy of Space Technology jointly proposed a deployable Mars quadcopter for air patrol and sampling with a sample mass of 100 g. This paper presents a concept for a Mars quadcopter and its mission operations, including operational procedures and flight modes, to provide a framework for future Mars flight sampling missions. We describe the structure and avionics architecture of the Mars quadcopter. Then, we present the hardware and software design of the guidance, navigation, and control (GNC) system to illustrate the autonomous flight control method. Finally, we introduce the deployment mechanism of the Mars quadcopter. • A foldable Mars quadcopter improves the efficiency and flexibility of Mars sample return tasks. • Five-module avionics system augments the Mars quadcopter for survival and operation. • A navigation method combining inertial and binocular vision enables autonomous flight on Mars. • The guidance system based on mode commands can independently plan exploration operations. [ABSTRACT FROM AUTHOR]

Deployable structures inspired by origami have gained significant prominence in space applications, including solar arrays, reflector arrays, and satellite antennas. Traditionally, these structures consist of separate components that are folded during launch and later deployed in space using springs or actuator mechanisms. However, recent advancements in multi-material additive manufacturing have opened new possibilities for the fabrication of monolithic structures, integrating deployable mechanisms into a single entity. In this paper, we present a novel framework for designing unidirectional (UD) and bidirectional (BD) hinges utilizing viscoelastic materials, incorporated into rigid plates, forming deployable structures. These hinges exhibit the unique ability to store strain energy upon deformation and self-deploy to a flat configuration when released. The experimental results presented here demonstrate the effectiveness and feasibility of our hinge designs in real-world applications. To demonstrate the practical applicability of these hinge designs, we construct a Repeating Unit Cell (RUC) of a large origami-inspired deployable structure, showcasing their compatibility with popular deployable structure patterns. [Display omitted] • Designed Multi-material, monolithic structure with viscoelastic hinges. • Stored strain energy helps to self-deploy the hinge when released. • Designed topology of the hinge structure to allow for higher strain energy storage. • Designs also allow for minimal local buckling and easy self-deployment. • Experimentally verified hinges that are biased unidirectionally or bidirectionally. [ABSTRACT FROM AUTHOR]

A large fraction of spacecraft external surfaces or appendages often consists of solar arrays, which can be subjected to space debris impacts as in the case of the Sentinel 1 A event of August 2016. Therefore, it is of interest to understand how solar arrays respond to hypervelocity impacts and to investigate the generated fragments population. In this context, the University of Padova performed an impact experiment on a solar array consisting in a composite sandwich panel, coated with a Kapton layer, and provided with solar cells: a nylon cylinder of 0.039 g collided with the solar panel at a velocity of 4.86 km/s with an impact angle of 45 deg, detaching a solar cell and damaging the panel structure. More than 4500 fragments larger than 0.2 mm were collected and classified after the impact. In this paper the impact experiment is described and the fragments analysis is presented in terms of size and shape distributions; a comparison with a test on a composite sandwich panel shows that the distributions are strongly affected by material and manufacturing choices, in particular regarding the fragments generated by delamination. • An impact experiment was performed on a solar array sample. • More than 4500 fragments larger than 0.2 mm were collected and classified. • Fragments size and shape distributions were obtained and compared with older experiments. • Distributions are strongly affected by material and manufacturing choices. [ABSTRACT FROM AUTHOR]

This paper presents an algorithm for the preprocessing of observation data aimed at improving the robustness of orbit determination tools. Two objectives are fulfilled: obtain a refined solution to the initial orbit determination problem and detect possible outliers in the processed measurements. The uncertainty on the initial estimate is propagated forward in time and progressively reduced by exploiting sensor data available in said propagation window. Differential algebra techniques and a novel automatic domain splitting algorithm for second-order Taylor expansions are used to efficiently propagate uncertainties over time. A multifidelity approach is employed to minimize the computational effort while retaining the accuracy of the propagated estimate. At each observation epoch, a polynomial map is obtained by projecting the propagated states onto the observable space. Domains that do no overlap with the actual measurement are pruned thus reducing the uncertainty to be further propagated. Measurement outliers are also detected in this step. The refined estimate and retained observations are then used to improve the robustness of batch orbit determination tools. The effectiveness of the algorithm is demonstrated for a geostationary transfer orbit object using synthetic and real observation data from the TAROT network. • Uncertainty on the initial orbit determination is reduced using other sensor data. • Measurement outliers are detected and isolated. • Multifidelity methods are used for fast and accurate uncertainty propagation. • Robustness of batch orbit determination tools is improved. [ABSTRACT FROM AUTHOR]

Low gravity fluid management is key to a variety of spaceflight applications, from life support aboard the International Space Station to propellant management in satellite tanks. Capillary devices are traditionally employed to position fluids in a microgravity environment, but they fail to induce effective phase separation in the conformal tank geometries and saturated two-phase propellants adopted in modern CubeSat propulsion. In order to meet the stringent requirements of future small satellite missions, this paper explores the application of capillary, magnetic positive positioning, thermal, and dielectrophoretic fluid management devices to conformal propellant tanks. The electromagnetic and capillary devices rely upon body and surface forces, respectively, while the proposed thermal approach exploits phase change in saturated liquid–gas mixtures to position the fluid. Each concept is evaluated based on its ability to reorient the propellant and maintain its position during operation, as well as the resulting volume, mass, and power requirements. Ultimately, thermomagnetic and thermodielectrophoretic devices are deemed to be best suited for the case under analysis because of their low size, weight, and power demand in combination with their reliability and performance. These new systems offer a novel and robust solution to small satellite propellant management, a key enabling technology for the future of maneuverable CubeSats. • Capillary PMDs cannot ensure phase separation in saturated mixtures. • Electromagnetic devices can maintain fluid position in microgravity. • Thermoelectromagnetic PMDs offer improved performance with decreased mass and volume. [ABSTRACT FROM AUTHOR]

Hill's approximation models the motion of two small masses gravitationally interacting with each other and perturbed by a large central body. An application of the model is employed in this paper to manoeuvre the relative motion of two asteroids by a small impulse to capture them in bound binary motion in Earth's orbit. The initial conditions prior to the capture manoeuvre are restricted in a parameter space termed the gateway region. The gateway region is produced by applying the constraint that the capture impulse is a real-valued function, and the zero-velocity curve closes, enveloping both asteroids. A full mission scenario is designed with three impulses. The first impulse transfers one asteroid from a far field region to the gateway region, where there is mutual interaction with the second asteroid, which is assumed to be the origin of the relative motion. The second impulse changes the trajectory of the asteroid in the gateway region to a linear drift approaching the second asteroid. The third impulse is calculated by means of increasing the Jacobi integral to reach the critical value required to close the zero-velocity curve. It is demonstrated in principle that the triple impulses enable the candidate asteroids to be captured in a bound pair in Earth's orbit. It is also shown that the thruster on-time for the capture impulse as applied for the different small asteroids considered is short relative to the natural time-scale of the orbital dynamics of the problem, which shows the feasibility of the impulse approximation with the typical thruster forces applied. This strategy could provide a basis for parking small captured near-Earth asteroids in Earth's orbit. • A three-impulse scenario is proposed to form an asteroid pair in bound motion. • Hill's equations are modified with capture impulse to the transverse relative velocity. • An analytical investigation for position and velocity constraints is applied to the capture impulse. • A gateway map is generated to provide possible initial conditions for the capture impulse scenario. • The scenario is illustrated by forming an artificial binary pair in high Earth orbit. [ABSTRACT FROM AUTHOR]

Integrating a model-based and data-driven approach to generate sufficient number of training data that include time-varying operating conditions is needed to develop a prognostics and health management technique using machine learning for a reusable rocket engine. A data-driven approach was used to predict the engine behavior, which cannot be determined using a model-based approach. This paper proposes a clustered regression model to predict various operating conditions for a reusable rocket engine. The training data were divided in advance according to the operating conditions using the Gaussian mixture model, and multiple regression models were individually trained on each cluster. By switching the regression models according to the operating conditions, the time-series data comprising various operating conditions, such as the combustion and chill-down phases and transition between phases, were accurately predicted when compared with the prediction results of a single regression model. The prediction results of the present hybrid approach agree reasonably well with the static firing test results of a reusable rocket engine if the explanatory variables are adequately included. By integrating model-based and data-driven approaches, adequate training data can be generated for prognostics and health management of the reusable liquid rocket engine. • This study improves the prediction accuracy of the data-driven approach. • Training data for the data-driven approach comprised varying operating conditions. • The varying operating conditions were classified using a Gaussian mixture model. • The regression models were trained based on each classified operating condition. • The prediction accuracy was improved using multiple regression models. [ABSTRACT FROM AUTHOR]

The establishment of secure earth-independent long-term lunar habitats has been envisioned by numerous government agencies and private companies. Recent advancements in assessing seismic hazards caused by shallow moonquakes have highlighted the importance of incorporating this phenomenon into the design of robust and resilient lunar structures. However, further research is required to explore lunar habitat design that considers seismic loads. This paper proposes assessing the structural response of a lunar habitat made of sulfur concrete covered with a regolith layer. The numerical model of the structure is subjected to gravitational, internal pressure and seismic loads. The seismic analysis of the structure is carried out using spectral and nonlinear time history methods. Conditional mean spectra for shallow moonquakes with return periods of 75, 475, 970 and 2475 years are used in the seismic analysis. The records used for the temporal analyses were ground motions that agree with a preliminary seismic hazard on the Moon. The results of temporal analyses reveal that shallow moonquakes with return periods greater than 475 years can lead to the loss of the global stability of the structure. Consequently, the findings imply that seismic loads have the potential to impose unacceptable demands on lunar structures constructed from in-situ materials like sulfur concrete. Hence, it is imperative to incorporate seismic considerations in the design process for developing resilient and long-term lunar habitats. • Seismic analysis of a lunar structure was carried out using spectral and temporal methods. • A methodology was adapted to scale moonquake records to different hazard spectra obtained from probabilistic methods. • It was demonstrated that seismic demand might generate unacceptable deformations in lunar structures built with in-situ materials. [ABSTRACT FROM AUTHOR]

Recently proposed, the airframe/inlet integrated full-waverider vehicle design methodology introduces an innovative approach to the design of air-breathing hypersonic vehicles. This methodology, compared to traditional integrated waverider design methods, leverages the features of the waverider to enhance the compression ability of the precompression surface and the lift-to-drag ratio characteristics of the vehicle. However, existing full-waverider design methods, based on cone-derived waverider theory, result in a gas cross-flow phenomenon at the inlet lip due to the conical shockwave. This paper suggests an airframe/inlet integrated full-waverider vehicle design method that utilizes the osculating cone waverider theory to improve gas intake quality. Initially, we present the method to obtain profile lines and basic flow field models at different osculating planes of the vehicle. We then validate the osculating cone integrated full-waverider vehicle design methodology through computational fluid dynamics. Although the viscous effect slightly reduces the lift-to-drag ratio of the vehicle and the total pressure recovery coefficient at the isolator exit, the shock wave remains attached to the leading-edge. Finally, a comparison with the cone-derived integrated full-waverider vehicle reveals that the osculating cone integrated full-waverider vehicle achieves similar aerodynamic performance, but with superior gas intake quality. The results indicate that the methodology proposed herein holds promise for guiding future engineering applications of integrated full-waverider vehicles. • An osculating cone integrated full-waverider vehicle design method is proposed. • Aerodynamic analysis of an integrated vehicle obtained by this design method. • Numerical simulation investigates the effects of viscosity on this design method. • The present design method can improve the gas intake quality of the vehicle. [ABSTRACT FROM AUTHOR]

In this paper, an effective control and parameters identification scheme was proposed for high-precision trajectory tracking of space robotic manipulators. Our proposed control method employed a linear-extended-state-observer (LESO) based model predictive control (MPC) scheme. The feedback linearization approach was utilized to develop a model predictive controller. The optimization objective function was simplified to a standard quadratic programming (QP) problem for solution. To minimize the impact of the inaccuracy in the dynamics model, a LESO was designed to estimate the disturbance, and its convergence was demonstrated. In order to improve the accuracy of dynamic model, the physical consistency of inertial parameters was constrained using linear matrix inequality (LMI) techniques. Moreover, the dynamics parameters were estimated through semidefinite programming (SDP) techniques. To validate the effectiveness of the proposed control method, it was compared with existing controllers through simulations and experiments. The results demonstrated the superior performance of the proposed scheme, highlighting its potential for achieving high-precision trajectory tracking control in space robotic manipulators. • High-precision trajectory tracking control for space manipulator is investigated. • Feedback linearization based predictive model for MPC is established. • A LESO is designed and its convergence is proven. • A physical consistency parameters identification method is presented. • Both simulations and experiments are performed for validation. [ABSTRACT FROM AUTHOR]

Planetary defense (PD) is a well-established program that refers to activities and actions to predict and mitigate a potential impact by an asteroid or comet on the Earth. States are discussing possible deflection scenarios based on their scientific knowledge and technology capacities at the UN-endorsed Space Mission Planning Advisory Group (SMPAG) and at the General Assembly. Even though PD is being formally recognized as an inherently global challenge requiring international cooperation at the level of UN COPUOS, the current state of the global PD governance depends on particular space-faring states and their willingness to fund and execute planetary defense missions. Political representatives consider PD mostly as scientific efforts, respectively efforts conducted mostly by the community of scientists. However, this would change the day an asteroid on a collision course is detected and without solid PD governance in place, behavior of decision-makers could be contingent leading to the possibly unpredictable supranational policy above national security agendas but with inherent implications to the national security of sovereign states. This situation has historical reasons. This paper, firstly, studies the evolution of PD governance from the 80s up to today, to identify the reasons for the current state-oriented (or inter-governmental in the case of ESA) PD governance. Then, secondly, I propose a multilateral PD security regime as an answer to the current state of PD governance. As the threat of being hit by an asteroid is a low probability, high impact natural phenomenon involving the whole Earth, it requires a multigenerational, financially sustainable, and scientifically beneficial planetary-wide program that covers all without exception and does not cause national security concerns of any sovereign state. As asteroids do not discriminate, the planet Earth should consider establishing a multilateral planetary defense security regime to provide security assurances to all, develop trust in the related technology development that all can oversee, and provide an environment to which all entities contribute with their scientific and industrial capacities. • Study of the last 30 years of planetary defense governance development. • Should serve as a guide for policy discussions given the detailed historical insight. • Identifies policy bottlenecks for future policy discussions. • Proposes core multilateral principles for planetary defense governance. • That is a multigenerational, financially sustainable, and scientifically beneficial policy. [ABSTRACT FROM AUTHOR]

The logistic supply for Space Station is very expensive and difficult compared with any other human natural exploration activities due to launch costs and challenges, and the logistic problems of establishing bases on the Moon and Mars in the future will be more prominent. Especially for emergencies, the losses caused by ground support delay will be immeasurable. Considering the successful maintenance of the Hubble Telescope and operations on the ISS, one of the most effective methods to solve the logistic supply problem is to enhance the capability of self-maintenance and support of the facilities. Through the participation of astronauts and a series of tools and technical support, the capability of self-maintenance and support of the facilities can be effectively improved. For this purpose, the on-orbit Maintenance and Manipulation Workbench (MMW) has been developed and installed on the Mengtian Lab Module (MLM) of the Chinese Space Station (CSS) to provide an enclosed working area for on-orbit fault diagnosis, maintenance, and manual or mechanical operations for the payload. This paper introduces the development and functions of the MMW on the MLM. In the limited rack space and limited resources, the MMW provides a 380-L work volume, 1000-W DC power with a versatile supply interface (100, 28, and 12 Vdc), 800-W cooling capability, video and data recording and real-time downlink capabilities, ground commanding capability, wastes (waste gases and waste particles) treatment capabilities. With these capabilities, the MMW is an ideal platform to support the operation and test verification of fine mechanical operations, structural and electronic assembly of scientific experiments and technical experiments in the cabin, and to support the development of space robots and telescience technology experiments. Areas of research that will benefit from investigations in the MMW include fluid physics, materials science, combustion science, on-orbit maintenance and manipulation, and advanced life support technologies. The mechanical, electrical and thermal control subsystems are discussed separately. Special tests for the treatment of wastes have been carried out to verify the capabilities. The development of the MMW will surely enhance the space research in the forthcoming long-term activities of CSS. • The MMW has been successfully installed on the CSS. • A "space maintenance factory", an "assembly workshop", or a laboratory. • Astronauts can complete complex operational tasks more easily and intuitively. • Explore more orbital experiments and applied research technologies. • Support for the forthcoming long-term activities of CSS. [ABSTRACT FROM AUTHOR]

Development of efficient and safe energy systems for space objects includes ensuring fire safety. This problem requires a deep understanding of influence of gravitational forces on combustion processes, which explains the importance of studying this issue. The paper analyzes the dynamics of inverted conical methane-air flame under conditions of normal and reverse gravitational force with external acoustic excitation. Applied frequencies of acoustic excitation could be parasitic or be produced during proper work of different technical systems of aircraft. High speed recording of flame chemiluminescence and PIV analysis were conducted. Decomposition of flow vector pictures into the modes was calculated using POD method. Energy contribution into the flow was analyzed for its separate elements at the different frequencies of excitation. At the high excitation frequencies vortex generation intensity in the shear layer with reverse gravity is about the same as it was observed in the experiments with normal gravity direction, that could indicate dominance of acoustic mechanism of the vortexes stimulation over the mechanism of shear instability. At external excitation frequency equals 240 Hz significant growth of shear vortex diameters and increasing of perpendicular oscillations amplitude of flame branches were obtained. Thereby in such experiment with 240 Hz external acoustic excitation the most intensive large scale instability of flow was observed. • The smallest vortices sizes are observed at 0 and 60 Hz excitation frequencies. • At excitation of 160, 240 and 420 Hz the vortices are non-affected by the gravity. • Vortice sizes and transverse oscillation range approach maximum at 240 Hz excitation. [ABSTRACT FROM AUTHOR]

All structural materials contain certain microdefects. Various elements of aircraft and spacecraft made from these materials are subjected to high mechanical loads and thermal effects during operation, which, in turn, can lead to the evolution of defects and total fracture of individual parts of the mechanism. Therefore, a quantitative and qualitative estimation of the stress and temperature fields in different bodies with complex configuration of external loading and temperature disturbances is needed. This paper presents a technique for numerical modeling of the temperature field in a medium weakened by a large system of partially heat penetrable cracks, including the possibility of considering it as an infinite periodic system. The method is based on a numerical algorithm, which uses the expansion of the solution into a finite series. Every term of this series is a certain analytical solution of the heat conduction theory obtained by the authors. To verify the method, the numerical results were compared with the well-known analytical solutions both for a configuration with finite number of cracks and for the periodic system. Also the heat flux intensity factors for a doubly periodic system were determined. • The numerical modeling of the temperature field in a medium weakened by partially heat penetrable cracks is presented. • The possibility of considering an infinite periodic system of cracks is included. • The numerical results were compared with the well-known analytical solutions and showed good agreement. • The heat flux intensity factors for a doubly periodic system were determined. [ABSTRACT FROM AUTHOR]

This paper presents a distributed trajectory optimization algorithm for the reconfiguration of large-scale spacecraft formation, which uses a low thrust propulsion system to safely guide the spacecraft formation to a desired formation. Formation reconfiguration is formulated as a trajectory optimization problem with complex constraints, and the relative motion is accurately described by a nonlinear relative dynamics considering J 2 perturbation. The collision avoidance constraint is convexified into a tangent plane to ensure the accuracy of the solution. The resulting nonlinear trajectory optimization problem is solved by applying the h p -adaptive pseudospectral method to convert it into a nonlinear programming problem. In order to overcome the disadvantage of huge computation of centralized algorithm considering collision avoidance, "predicted trajectory" is introduced to transmit information between spacecraft, and the parallel computing is implemented. Finally, a numerical simulation is given to verify the computational efficiency and collision avoidance effectiveness of the proposed distributed algorithm. • Distributed trajectory optimization algorithm for massive formation reconfiguration. • The J 2 nonlinear relative dynamics is used to obtain accurate optimal trajectory. • Parallel computation is achieved by information interaction of neighbors. • Neighbor updating is proposed to reduce communication traffic. • Significantly decrease computation and achieve collision avoidance. [ABSTRACT FROM AUTHOR]

Model-based systems engineering (MBSE) is an emerging methodology not only in the field of spacecraft systems engineering. The key objective is to store the whole system design in one single source of truth system model. The continuously maintained and customizable open-source MBSE tool Virtual Satellite developed by German Aerospace Center (DLR) offers a framework for storing such a system model. Since the open-source 3D computer aided design (CAD) software FreeCAD has a dedicated Virtual Satellite workbench for importing the 3D model file generated by Virtual Satellite, it is possible to export the 3D model from Virtual Satellite, apply changes to it in FreeCAD, and afterwards reimport it in Virtual Satellite. This paper presents a workflow implementation to conduct automated steady-state and transient thermal analyses using the Virtual Satellite system model as a basis, utilizing the aforementioned features. Part of this effort was to extend Virtual Satellite by thermal modeling capabilities, such as material parameters, surface characteristics, thermal interfaces, boundary conditions and further relevant aspects. With custom Java applications, so-called "Apps", all aspects of the system model relevant for a thermal analysis are exported from Virtual Satellite and written to accordingly generated text files. Moreover, the before described 3D model export is used to export the model to FreeCAD. The FreeCAD Python console, in combination with the finite element method (FEM) workbench, is used to create a script that executes meshing, application of the boundary and initial conditions, as well as identification of contacts between parts and application of the specified contact conductance values. The obtained files form a complete thermal model and can be fed directly into the open-source FEM software CalculiX for executing the thermal analysis. Then, the result can be processed by an additional Virtual Satellite App. To account for the space thermal environment an additional feature was added to the FreeCAD script. The feature uses orbit information obtained from external mission analysis tools to calculate the approximate combined heat load from Sun, Earth infrared, and albedo radiation at each time point in the specified simulation time interval. This new workflow allows to execute quick thermal analyses at an early design stage and therefore opens new opportunities in the evaluation of different system designs, as early as in a concurrent engineering study. Configuration parameters are provided that also allow more sophisticated analyses in later development phases, such as custom mesh definition and local mesh refinement. • Thermal design process of spacecraft relies on engineer's experience in early stages. • Potential help for the thermal design process is thermal simulation. • Thermal simulations require significant manual setup effort. • Created workflow automates simulation setup and result feedback. • Data for simulation setup maintained in MBSE system model. [ABSTRACT FROM AUTHOR]

For the first time in history, Parker Solar Probe touched the Sun in April 2021 when it entered the Sun's corona—humanity's first mission ever that reached the Sun—for a scientific in-situ study to unlock the mystery of our star. To reach the Sun is extremely difficult because, unlike any other space exploration destination, the Sun is the solar system's central body, around which Earth and other planets revolve. Earth orbits the Sun at a velocity of about 30 km/s, so for a spacecraft to reach the Sun from Earth the spacecraft must cancel out its velocity in the Earth orbit so it can fall to the Sun. That requires an excessively large energy higher than that required to reach any destination in the solar system, which is beyond the capability of current rockets. This paper describes how Parker Solar Probe is delivered to the Sun through a unique V7GA trajectory that enables it to attain the excessive energy through planetary gravity assists of Venus, flying by Venus seven times to ultimately come within 9.86 solar radii (R S) of the Sun. An emphasis is on in-flight trajectory control of the probe through the intricate V7GA trajectory in actual operation in an unprecedented space environment around the Sun, where unpredictable orbit perturbations plus their ripple effects on the consecutive seven Venus flybys could easily drive the flight trajectory out of control. A novel systematic 2-level control method is developed and has been applied to solve the complicated trajectory control problem in flight operation, using in-flight trajectory re-optimization to select the targets of Venus flybys as trajectory control points and then using trajectory correction maneuvers to target a Venus flyby at the control point. This method enables effective and efficient control of the intricate V7GA trajectory and has resulted in significantly better than pre-launch anticipated flight performance, saving tremendous ΔV. Four years from launch, Parker Solar Probe has completed 5 of the 7 Venus gravity-assist flybys and 12 of the 24 planned solar encounters, passing by the Sun at 13 R S in its orbit. Flight results and performance of the completed flight to date are also described. • How Parker Solar Probe was flown and touched the Sun for the first time in history. • Getting to the Sun is extremely difficult and is enabled by a unique V7GA trajectory. • A novel systematic 2-level control method for complex in-flight trajectory control. • In-flight trajectory re-optimization, trajectory correction maneuver and performance. • Flight results of 5 Venus gravity-assist flybys and record-braking solar encounters. [ABSTRACT FROM AUTHOR]

This paper focuses on the application of meta-reinforcement learning to the robust design of low-thrust interplanetary trajectories in the presence of multiple uncertainties. A closed-loop control policy is used to optimally steer the spacecraft to a final target state despite the considered perturbations. The control policy is approximated by a deep recurrent neural network, trained by policy-gradient reinforcement learning on a collection of environments featuring mixed sources of uncertainty, namely dynamic uncertainty and control execution errors. The recurrent network is able to build an internal representation of the distribution of environments, thus better adapting the control to the different stochastic scenarios. The results in terms of optimality, constraint handling, and robustness on a fuel-optimal low-thrust transfer between Earth and Mars are compared with those obtained via a traditional reinforcement learning approach based on a feed-forward neural network. • In interplanetary space missions, the spacecraft trajectory is affected by multiple uncertainties. • Meta-reinforcement learning can be applied seamlessly to any uncertainty and dynamic model. • A recurrent neural network is used as a history-dependent closed-loop control policy. • The network is trained on a low-thrust transfer featuring dynamic uncertainties and control execution errors. • Meta-reinforcement learning shows improved performance and robustness compared to standard reinforcement learning. [ABSTRACT FROM AUTHOR]

This paper investigates the effects of model fidelity and parameter uncertainty on the performance of a hybrid model-based feedback-feedforward control scheme for attitude tracking of a satellite with flexible appendages. The feedforward component is an inverse model-based term produced through a computational approach known as inverse simulation (InvSim), which works by iteratively solving a discretised reference trajectory. The hybrid controller's feedback is proportional-derivative (PD) based, using body attitude and rate feedback to provide stability and robustness. Furthermore, to ensure that the flexible modes do not trigger instability, the PD control gains are tuned to give a closed-loop response that is significantly slower than the flexible modes. Additionally, excitation of the flexible modes is reduced by minimising jerk through polynomial rest-to-rest manoeuvres, following the shortest quaternion path using spherical–linear-interpolation (SLERP). The effects of the appendage flexing on attitude tracking are then compensated through the feedforward element of the hybrid controller, with performance being compared to a traditional PD tracking law. The effect of the model fidelity on the performance of the hybrid controller is investigated through the use of both rigid body and multiple-fidelity finite-element mathematical models. Additionally, the effect of uncertainties in the model parameters is investigated to determine the accuracy of the model required to obtain significant improvement in attitude tracking. It is found that in the absence of any model parameter uncertainty, the hybrid controller outperforms the PD tracking control law by at least one order of magnitude when the finite-element model is used. Increasing the number of finite elements was found to provide no significant improvement in performance, with one element being sufficient and favourable with its lower computational overhead. It was also found that to ensure good performance compared to the PD tracking controller, the uncertainty in the inertia tensor should be < 1%. Similarly, uncertainty in the first flexible modal frequency should be < 0.5 rad/s. • Inverse simulation (InvSim) used to produce model-based feedforward control. • Combined with PD feedback to improve attitude tracking. • Effects of model uncertainty and fidelity on model-based control investigated. [ABSTRACT FROM AUTHOR]

In the exploration of celestial bodies, including planets, moons, and asteroids, it is useful to identify stable orbits in which the orbital elements remain, on average, constant (referred to as frozen orbits). This paper aims to investigate the feasibility of frozen orbits for a small body, such as a probe, orbiting a main body under the perturbation due to the gravitational attraction of an inner third-body, on a circular (or slightly eccentric) orbit coplanar with the main body equator, by means of an approach based on mean element theory. The disturbing potential has been developed in Legendre polynomials up to order l = 5 (triakontadipole level of approximation). Thus, the double-averaged potential function and the associated constants of motion have been determined up to the same order. Then, by applying the Lagrange Planetary Equations, the secular and long-term orbital element variations have been obtained. Therefore, frozen orbital solutions have been evaluated, and the effects on them of the addition of main body oblateness perturbation have been considered. Furthermore, the procedure has been repeated, and the frozen solutions are re-evaluated in the case of a disturbing body with small eccentricity up to order l = 4 (hexadecapole level of approximation). These solutions offer significant advantages in exploring bodies subject to a notable perturbation from an inner third-body, such as binary asteroid systems. • Frozen orbits for the inner third-body problem. • Inner circular planar third-body: Double-averaged potential up to triakontadipole. • Inner slightly eccentric planar third-body up to hexadecapole. • Orbital solutions for binary asteroid observation. [ABSTRACT FROM AUTHOR]

Ionic liquids are regarded as promising propellants for green spacecraft propulsion. A well understanding of their fundamental properties is essential to accelerate the applications in practical propulsion. Therefore, the decomposition properties of an imidazole-based ionic liquid 1-ethyl-3-methyl imidazolium tetrafluoroborate ([EMIm][BF4]) were studied in this paper to help evaluate its chemical propulsion performance. Decomposition experiments were carried out with a high-temperature integrated thermal analyzer with heating rates of 5–20 °C/min. Results showed the decomposition process can be divided into four stages by three characteristic temperatures. The mass loss and heat absorption mainly occurred in the second and third stages. The activation energy of decomposition reactions was also obtained. Analyses indicated mechanisms in the first and second stages are multiple-step kinetics. But that of the third and fourth stages can be simplified as single-step kinetics. • Decomposition properties of [EMIm][BF4] were systematically studied. • Decomposition process was found to have four stages. • Activation energy of the decomposition reactions was obtained. • Decomposition process of [EMIm][BF4] should be described as multiple-step kinetics. [ABSTRACT FROM AUTHOR]

In this paper, we investigate the dual-channel event-triggered deployment control for the space tethered system with intermittent output. Two different dynamic event-triggered mechanisms are designed to implement the event-based state sampling and the event-based input sampling, in which the data transmission frequencies are reduced at the dual channel, namely sensor-to-controller and controller-to-actuator. Then, the neural network (NN) state observer based on the intermittent output is designed to estimate the unmeasurable state under external disturbance, and the observer-based sliding mode controller is designed. Furthermore, because of the non-periodic sampling of the state signal and the output signal, the closed-loop system is proved via the analysis of the hybrid system, and the Zeno behavior is avoidance under these two event-triggered conditions. Finally, the simulation tests are implemented to verify the effectiveness of the proposed scheme. • A control scheme incorporating a dual-channel event-triggered mechanism is investigated. • A neural network-based state observer on intermittent output is designed. • Two dynamic event-triggered mechanisms are designed to reduce the event-triggered numbers at two channels. [ABSTRACT FROM AUTHOR]

The Japan Aerospace Exploration Agency has developed a power processing unit for a 6 kW Hall thruster, a candidate for propulsion systems for geostationary satellites. Hall thrusters have a discharge instability of tens of kHz, which may damage thrusters, power processing units, and satellite systems. This paper proposes our system for detecting discharge current amplitudes that exceed the specified limit of the thruster. The discharge current is converted into a voltage by the current sensor and is output to a digital controller through an AD converter. The system's innovation includes a frequency prediction model and a damping correction algorithm, which allow the system to accurately detect the peak amplitude of the discharge current without testing to measure the frequency characters of the Hall thruster. The design of the amplitude detection system considers the frequencies of the ionization oscillation and the switching frequency of the power processing unit. The frequency prediction model estimates the frequency of discharge current oscillations from simple parameters: the output voltage of the discharge power supply and the thruster dimensions. The amplitude of the discharge current is damped slightly by the frequency characteristics of the system, but the algorithm corrects this. The results of the frequency prediction model agreed well with the actual measurements. The predicted value was 13.2 kHz, while the actual measurement was 8–15 kHz. The integration test showed that the amplitude detection system identified 36 peaks exceeding the specified limit of 80% peak-to-peak during a 100 ms segment, although the actual number was 34. This difference is acceptable. The proposed amplitude detection system can accurately count the peaks in the discharge current that exceed specifications. • An amplitude detection system detects discharge current oscillations that exceed the specifications for protecting satellites. • The algorithm and frequency response of the system are designed considering the frequency of ionization oscillations. • The frequency is estimated by a simple model with only the output voltage of the discharge power supply and thruster dimensions. • The system detected the amplitude of the discharge current oscillations with an error of 5.8%. [ABSTRACT FROM AUTHOR]

The success of a multi-kilometre drive by a solar-powered rover at the lunar south pole depends upon careful planning in space and time due to highly dynamic solar illumination conditions. An additional challenge is that the rover may be subject to random faults that can temporarily delay long-range traverses. The majority of existing global spatiotemporal planners assume a deterministic rover-environment model and do not account for random faults. In this paper, we consider a random fault profile with a known, average spatial fault rate. We introduce a methodology to compute recovery policies that maximize the probability of survival of a solar-powered rover from different start states. A recovery policy defines a set of recourse actions to reach a safe location with sufficient battery energy remaining, given the local solar illumination conditions. We solve a stochastic reach-avoid problem using dynamic programming to find an optimal recovery policy. Our focus, in part, is on the implications of state space discretization, which is required in practical implementations. We propose a modified dynamic programming algorithm that conservatively accounts for approximation errors. To demonstrate the benefits of our approach, we compare against existing methods in scenarios where a solar-powered rover seeks to safely exit from permanently shadowed regions in the Cabeus area at the lunar south pole. We also highlight the relevance of our methodology for mission formulation and trade safety analysis by comparing different rover mobility models in simulated recovery drives from the LCROSS impact region. • The safety of a solar-powered rover depends on its location, energy and time of day. • Maximally safe behaviours depend on when and where delays occur during a traverse. • The probability of survival of a rover varies non-monotonically with time. • Min–max dynamic programming can conservatively account for the temporal risk ambiguity. [ABSTRACT FROM AUTHOR]

Reflector antennas, operating at high frequencies, require a high surface accuracy: deviations from the target geometry accuracy result in a loss of RF performance (e.g., antenna gain and efficiency). As a result, the aperture may need to be larger, or a more complex backing structure must be engineered to achieve the reflector surface accuracy to deliver the equivalent gain of an efficient solid reflector. This would typically increase the mass, the stowed volume, and the design complexity. The design of larger aperture deployable antennas (e.g., Cassegrain, Offset) for SAR (Synthetic Aperture Radar) applications can be challenging for missions where platform size and payload mass are limited. Inspired by the Origami based techniques, Oxford Space Systems (OSS) is currently developing an innovative SAR Reflectarray antenna, targeting the X-Band frequency range, with the aim of achieving a relatively large aperture from a small, stowed volume suitable for a small satellite. The Reflectarray antenna architecture has several potential advantages over other antenna designs. It is a cost-competitive technology with excellent stowage efficiency, which can be further improved with origami-based folding mechanisms/kinematics. It requires a simplified deployment mechanism and in contrast to conventional parabolic reflectors, it does not need a separate backing structure. Until now, work on space deployable Reflectarray antennas (e.g., MARCO, OMERA and ISSARA missions from NASA) have focused on multi-degree of freedom structures, i.e., independent self-actuated hinge driven architectures. This paper presents an elegant single degree of freedom, origami-inspired Reflectarray architecture that does not require actuation on all the hinges. Thanks to its Hold-Down Release Mechanism (HDRM) and spring-loaded hinges, the OSS Reflectarray antenna is self-deployable. It maintains the required flatness accuracy achieved through its novel hinge integrated latching system. The OSS Reflectarray antenna has an 8:1 footprint stowage ratio, which enables installation on small satellite platforms (between 50 kg and 100 kg). The reflectarray has been designed to work in the X-band frequency with dual linear polarization for the concept study. However, the architecture can be scaled up or down to suit other frequency bands, making the OSS Reflectarray compatible with many applications and platforms. • Rigid origami folding ensure high deployed/stowed volume ratio. • OSS Origami Reflectarray design ensure high Performances/mass ratio. • OSS Reflectarray simulations results ensure a good correlation with tests. • Target for SAR applications in the design. • Kinematics and RF tests successfully prove the design of such antenna. [ABSTRACT FROM AUTHOR]

The use of electro-optical imaging system for commercial, military, or surveillance missions, as well as space scientific missions, dates back to the first successful operation aboard the US Explorer-6 spacecraft, which was the first spacecraft to capture low-resolution images of the Earth's surface, and the Luna-3 space probe, which was the first spacecraft to send images of the lunar far side, in 1959. The technology underlying EO system have evolved over time, with advances in optics, electronics, and materials science leading to the development of more sophisticated and capable imaging sensors. Furthermore, the technological advances in sensors, optics, mechanics, material, signal and image processing algorithms have allowed EO payloads to become smaller, lighter, and more efficient while still providing high-quality imagery. Over 840 small spacecraft weighing less than 500 kg were launched, with 80% in the constellation segment and 20% as monolithic satellites (i.e., single spacecraft equipped with an electro-optical imaging instrument launched into space for various missions). In this paper, the electro-optical imaging systems of these spacecraft were reviewed using data from a variety of sources, including peer-reviewed journal publications, conference proceedings, mission webpages, and other publicly available satellite databases. Furthermore, the technological and commercial expansion of small spacecraft (< 500 kg) equipped with electro-optical imaging systems, as well as EO-COTS parametric performance values such as GSD, spectral range, field of view, and sensor type, are discussed through 2023, divided by spacecraft type: SSO satellites, Small spacecraft constellation, Deep space spacecraft, and EO commercialization. • The evolution of electro-optical imaging system. • Small spacecraft missions equipped with EO sensors. • Constellation of small satellite missions can provide comparable capability. • The advantages of an EO-based deep space spacecraft mission. • The main factors driving the market trend toward small spacecraft optical payload. [ABSTRACT FROM AUTHOR]

The regolith on the Earth's Moon has been attracting attention from the perspective of understanding the planetary surface processes and procuring in-situ resources. However, very few studies have investigated the hydraulic properties of the lunar regolith, which should pose one of the most significant physical properties for understanding the lunar surface processes and evaluating regolith utilizations. In this paper, we study the water permeability of the surficial regolith layer on sunlit lunar highlands using the lunar highlands regolith simulant LHS-1. We measure the water permeability of LHS-1 at packing fractions of 0.49–0.60 by simulating lunar gravitational environments and evaluating the influences of freeze-thaw, which is expected to occur on the sunlit lunar highland surfaces. Our results indicate that the water permeability of lunar highland regolith is one to two orders of magnitudes lower than that of fine silica sands and monodisperse spherical glass beads at the given packing fraction. The low water permeability of lunar highland regolith attributes to the regolith particles having irregular and angular shapes with sharp edges and micron-sized chunks due to the least weathering, unlike soil particles on the Earth. The horizontal water permeability of lunar highland regolith simulant before freeze-thaw is 1.6 × 10−13 m2 at a packing fraction of 0.6. From the porosity-depth profile available from previous studies, we also obtained the water permeability-depth curve, representing the vertical water permeabilities of 6 × 10−13–2 × 10−12 m2 and 7 × 10−14–3 × 10−13 m2 at the lunar surface and the depth of 5 m within the surficial lunar regolith layer, respectively. We also found that the water permeability of lunar highland regolith increased after one freeze-thaw cycle, achieving 6.4 times higher than the permeability before the freeze, and the increase was more significant with lower packing fractions. We suggest that the pore water in the solid phase due to the freeze encourages the rearrangement of the regolith particles changing the total contact area between the particles. This study will help us better understand the abundance and distribution of molecular water on the lunar surface, lunar surface processes, and regolith utilizations. • Water permeability (WP) of lunar highland soils is estimated using simulant LHS-1. • WP before the freeze-thaw is 1.6 × 10−13 m2 with a packing fraction (PF) of 60%. • WP-depth curve was estimated; e.g., WP can be 7 × 10−14–3 × 10−13 m2 at 5 m depth. • The change in WP before and after one freeze-thaw cycle was greater with lower PFs. • The tortuosity with a PF of 49% decreased by about 60% with the freeze-thaw. [ABSTRACT FROM AUTHOR]

Future human exploration missions require new concepts of operations since the international roadmap foresees stations with uncrewed periods (e.g. Lunar Gateway) and deep space missions with communication delays making real-time monitoring and control impossible (e.g. any vehicle on its way to or in orbit of Mars). As part of the commitment to achieve human presence in lunar orbit and prepare for missions to Mars, the goal for future operations is to become independent from ground and enable crews to live and work autonomously. However, as a first step towards on-board crew autonomy, the ground segment itself needs to be automated and its functions eventually transferred on-board. For this purpose, this paper introduces a novel tool, based on Columbus operations, conducted at the Columbus Control Center (COL-CC), allowing the interaction between scheduled activities and the commanding infrastructure, by automatically generating a command sequence, transferring the command sequence to the operators' command interface for inspection and approval, and finally sending the commands to the receiver for execution. This tool is implemented as a ground prototype, in order to show how ground operations can be automated and how such a tool can ultimately lead to increased crew autonomy. • Crewed deep-space missions cannot use real-time commanding, due to signal delays. • Real-time commanding is shifted from the ground controller to an on-board system. • An interface allows interaction between crew, mission timeline and command system. • The system is capable of automatically generating commands, based on constraints. • An activity database stores all necessary information for activity execution. [ABSTRACT FROM AUTHOR]

The paper investigates the problem of fire safety in habitable space structures. The flame propagation over thermally destructing material is studied theoretically within the framework of the assumptions of the boundary layer theory and the presence of a diffusion flame. The solution is developed based on the concept of thermally thin fuel layer. Flame propagation velocity over structured material, on which different widths ribs are attached, in an opposed flow is determined analytically. PMMA sample of 10 mm thickness, on which ribs of widths 1–10 mm are attached in the direction of flow, is considered as an example. The analytical solution is compared with experimental data. • Flame propagation velocity over structured material in an opposed flow is determined analytically. • The flame propagation velocity decreases with an increase in the width of the rib. • Slowing down the blower speed can help reduce the speed of flame propagation and the intensity of combustion. • The ribs create more favorable conditions for flame propagation. [ABSTRACT FROM AUTHOR]

Solar sails are an attractive technology for propellant-free propulsion, especially in small satellites with tight power and mass budgets. Several recent missions have shown promising applications of solar sails for nanosatellite orbital control in low Earth orbit (LEO). They enable easier access to and removal from high-altitude orbits. Since solar radiation pressure produces torques in addition to forces, a solar sail can also be used for low-power attitude control. Few missions have exploited this effect. In response, this paper proposes a new nanosatellite mission concept, only achievable by combined orbit-attitude control via solar sail: the sunflower mission, for high-altitude wide-area Earth observation. A realistic sail design meeting the mission requirements is obtained via parametric analysis and comprehensive review of past missions. A pyramid-shaped sail is proposed for the mission, and its advantages and drawbacks are analysed. The dynamics are evaluated using a high-fidelity numerical simulator with all relevant physical perturbations in LEO. The imaging objectives are successfully achieved over a multi-year duration. The sail also provides propellant-free orbit insertion from a low-altitude launch orbit, and fast de-orbit at end-of-life. It uses less power and has a smaller stowed size than several competitor technologies for orbit-attitude control of LEO nanosatellites. [Display omitted] • Sunflower mission proposed for wide-area small satellite Earth observation in LEO. • A pyramid-shaped solar sail provides low-power orbit and attitude control. • Propellant-free orbit raising from a low deployment altitude reduces launch cost. • Passive Sun pointing enables low-power attitude control for remote sensing. • Propellant-free orbit lowering enables quick de-orbit at end-of-life. [ABSTRACT FROM AUTHOR]

As a novel green and high-performance propellants, H 2 O 2 /H 2 O/Al propellants have the potential to solve the pollution problem faced by composite solid propellants while maintaining high performance. However, the drawback that it cannot be ignited at a low oxidizer-to-fuel weight ratio (O/F) and a low pressure hinders its further development. Fortunately, this ignition defect was addressed by the mixture of hydrogen peroxide gel and micron-Al in our recent work. In this paper, the effects of Al content and particle size on the combustion characteristic of burning surface propagation and combustion products of the mixture were investigated. Combustion experiments were conducted in a windowed pressure vessel, using mixtures with three mixing ratios (fuel-lean, stoichiometric, and fuel-rich) and Al particle diameters between 1 and 12 μm. The burning surface propagation was captured by a high-speed camera. The particle size distribution, component, and morphology of combustion products were characterized by laser particle size analyzer, X-ray powder diffraction, and scanning electron microscopy, respectively. And it is the first time that the combustion properties of H 2 O 2 /H 2 O/Al propellants (including H 2 O/Al propellants) are studied from the prospective of particle size distribution of combustion products. Results showed that with the decrease of Al content and the increase of particle diameter, the combustion intensity of the mixture decreased, resulting in the agglomeration of melted Al droplets on the burning surface. When the Al content was decreased to O/F = 1.7, Al particles in the mixture were no longer further oxidized in the process of combustion. When the Al particle diameter was increased to 7–10 μm, burning rates of the mixture at O/F = 1.7–1.1 tended to be stable. The combustion of mixtures containing 3–7 μm was kinetically controlled. From the prospective of variable burning rate and efficient combustion, the appropriate Al content in the mixture was O/F = 1.7–1.1, and the Al particle diameter should be kept within 1–7 μm. • Two agglomeration mechanisms exist during the combustion of hydrogen peroxide gel and micron-Al mixtures. • Burning rates approach stable when Al particle diameter beyond 7–10 μm. • Al particles reach the oxidation limit during combustion when O/F = 1.7. • The morphology of combustion products is greatly affected by combustion intensity. [ABSTRACT FROM AUTHOR]

In this paper we develop a single ballistic limit equation for multi-shock shields that is applicable to materials other than those used in its original development, that is applicable to impact scenarios at impact velocities beyond the testable regime and includes an improved projectile density term. The predictions of this ballistic limit equation regarding whether or not a particular multi-shock shield would be perforated by a specific impact was found to compare favorably with both test results and the results of numerical simulation predictions. In the end, the ballistic limit equation developed herein is seen to be not only more applicable to a wider class of spacecraft designs, but also can be used for trade studies that are typically performed to select and optimize multi-shock shielding designs for future spacecraft. • The multishock shield is a big improvement over the traditional Whipple Shield • A new, more general, multishock shield ballistic limit equation is developed • This equation compares very favorably against experimental results • This new form of the BLE can be used in trade studies and probabilistic risk assessments [ABSTRACT FROM AUTHOR]

Rocket engines are prone to suffer from transverse combustion instabilities, manifesting as periodic oscillations in pressure and heat release rate. Self-excited transverse instabilities in a rectangle multi-injector rocket combustor powered by n-C 12 H 26 and O 2 , are investigated experimentally and numerically in this paper. Transverse combustion instabilities are captured during repeatable experiments, while dynamic characteristics in the combustion chamber and driving mechanisms of instabilities are analyzed via numerical simulation based on stress-blended eddy simulation. Dynamic mode decomposition analysis reveals coupling behaviors between the transverse mode in the combustion chamber and longitudinal mode in the oxidizer post, resulting in propellant mass flow rate oscillations. Substantial evidence demonstrates that the intensities of heat release rate near the end walls are higher than that detected around the chamber center due to its intense coupling to the local first transverse (1W) acoustic mode. Flow field animations indicate strong deflection of oxygen posts and liquid particles by a transverse moving gas mixture leading to flame interactions and collision between adjacent share injectors. The propellant injected into the chamber does not burn immediately and is transported to the combustor end wall. Propellant mixing performance is significantly enhanced during interactions between acoustic pressure and end wall, causing a sudden heat release near 1W pressure anti-node, thereby strengthening pressure oscillations. Finally, a combustion instability driving mechanism associated with propellant mass flow rate oscillations, and interactions between unsteady acoustic pressure and heat release rate is proposed. The insights obtained from this work can aid comprehension of transverse combustion instabilities encountered by rocket engines. • Both liquid fuel and gaseous oxygen are highly deflected by transverse moving gas mixture when transverse combustion instability appearing. • Large flame interactions and collision structures appear inside the combustion chamber. • Combustion instability is driven by pulsated heat release rate near the end walls and periodic supply of propellant mass flow rate. [ABSTRACT FROM AUTHOR]

The Proxima Mission of ESA Astronaut Thomas Pesquet began with the launch of the 49 Soyuz (49 S) from the Baikonur Cosmodrome on 17 November 2016 (GMT322) at 20:20 GMT. Following a nominal 34-orbit rendezvous, the 49 S docked with the Rassvet module on the International Space Station on 19 November 2016 (GMT324). Along with ESA crewmember Pesquet were Russian crewmember Oleg Novitsky and US crewmember Peggy Whitson. The Proxima mission concluded on 02 June 2017 (GMT153) after Pesquet spent 196 days in low earth orbit. This paper gives an overview of the Proxima mission and the forty ESA objectives conducted during this time ranging from human physiology to material science, from robotics demonstrations to educational outreach activities. On-board and ground activities are described, crew-time usage is assessed, and scientific results are compiled. Five years after the completion of Proxima, we review the results of this research. • Research overview of Proxima, Thomas Pesquet's 2016–2017 space station mission. • Forty studies: Human physiology, biology, physics, material science, and tech demos. • 207 h of crew-time spent on ESA research during 196-day mission. • Space affects brain structure/function, headaches, bone, cartilage, and immune system. • Methods for identifying and mitigating biological contaminants in spaceflight. [ABSTRACT FROM AUTHOR]

The heterogeneity of the small body population complicates the prediction of small body properties before the spacecraft's arrival. In the context of autonomous small body exploration, it is crucial to develop algorithms that estimate the small body characteristics before orbit insertion and close proximity operations. This paper develops a vision-based estimation of the small-body rotational state (i.e., the center of rotation and rotation axis direction) during the approach phase. In this mission phase, the spacecraft observes the rotating celestial body and tracks features in images. As feature tracks are the projection of the landmarks' circular movement, the possible rotation axes are computed. Then, the rotation axis solution is chosen among the possible candidates by exploiting feature motion and a heuristic approach. Finally, the center of rotation is estimated from the center of brightness. The algorithm is tested on more than 800 test cases with two different asteroids (i.e., Bennu and Itokawa), three different lighting conditions, and more than 100 different rotation axis orientations. Each test case is composed of about 250 synthetic images of the asteroid which are used to track features and determine the rotational state. Results show that the error between the true rotation axis and its estimation is below 10 ° for 80% of the considered test cases, implying that the proposed algorithm is a suitable method for autonomous small body characterization. • Small bodies have heterogeneous characteristics complexifying mission design and operations. • The small body properties are hard to predict before the encounter. • Vision-based techniques enables autonomous estimation of the small-body rotational state. • The proposed method estimates the small-body rotational state from images. • The proposed method is tested on numerical simulations with different observational geometries. [ABSTRACT FROM AUTHOR]

The Japan Aerospace Exploration Agency had planned a sample return from the Jupiter Trojans using a solar power sail in its Oversize Kite-craft for Exploration and AstroNautics in the Outer Solar system (OKEANOS) mission. Solar power sail is an extended concept of solar sail consisting of a sail membrane covered with flexible solar arrays. OKEANOS is capable of generating a large amount of power sufficient to drive high-specific-impulse ion engines in the outer solar system, thus making it possible to perform sample return from the Jupiter Trojans. Various innovative technologies are needed to realize such a challenging mission in the distant space. Specifically, OKEANOS deploys a small lander to ensure the safety and save propellant of the mother spacecraft. The lander collects samples of the asteroid, and separates them into different groups, namely, those for in-situ analysis and those for return. After completing the in-situ analysis, the lander lifts off the asteroid, docks with the mother spacecraft, and delivers the samples to the reentry capsule mounted on the mother spacecraft. The reentry capsule is designed so that it can manage the reentry speed of 15 km/s. Although OKEANOS was proposed as a Japanese large-class mission and advanced to the final selection, it was not selected because of a cost issue. However, its sample return concept is inherited by small- and massive-body explorations coming in the next generations. In this paper, we report the state-of-the-art technologies that enable sample returns from the outer solar system, which were developed through the OKEANOS mission study. • OKEANOS is a solar power sail that can perform a round trip to Jupiter Trojans. • A multi-spacecraft system is used to explore the Trojan asteroid. • A small lander collects surface and subsurface samples to bring back to Earth. • The lander and the mother spacecraft perform autonomous rendezvous and docking. • The reentry capsule was designed so as to manage the reentry speed of 15 km/s. [ABSTRACT FROM AUTHOR]

Slosh dynamics in a liquid propellant tank during a spacecraft maneuver can have an effect on the system stability and potentially lead to critical problems for the mission operation. In particular, the fuel sloshing problem has drawn significant attention in aerospace applications where a large portion of the gross weight is comprised of liquid fuel. Recently, many active control methods have been introduced as alternatives to passive approaches. However, the majority of these methods basically require the full-state condition. To guarantee the availability of the full-state condition, this paper proposes a cascade nonlinear state observer based on the sliding mode theory for estimating the unmeasurable slosh variables in a planar model. Additionally, for large amplitude estimation, a complete nonlinear mathematical model with a pendulum analogy is formulated without a linearization process. The stability analysis is theoretically proven by Lyapunov method. The effectiveness is demonstrated through simulations, and the performance comparison with a different observer method is also presented in the results. [Display omitted] • Observer design of cascade structure for slosh-variable estimation. • Complete nonlinear equation with pendulum analogy for slosh dynamics. • Full-state availability and large amplitude estimation are achieved. • The stability and small bound convergence are theoretically proven. • The effectiveness and performance comparison by numerical simulations. [ABSTRACT FROM AUTHOR]

At present, detonation principles are widely used for propulsion purposes. Stable initiation of detonation and its sustainable propagation are the key factors for safe operation of detonation chambers. In this paper, the evolution of the cellular structure of detonation waves under the condition of non-uniform initiation is investigated within the frames of two-dimensional simulations in a stoichiometric mixture of hydrogen with air. The dependence on cellular structure on the heterogeneity of initiation sources, chemical kinetic mechanism is studied. The simulation results are compared with experimental data. It is demonstrated, that in case of detonation decay after switching off initiators the flow structure essentially depends on heterogeneity of initiator, while in achieving the self-sustaining propagation mode the wave structure becomes independent on initiation conditions and "forgets" initial heterogeneity. • Two-dimensional simulations of the cellular structure of a detonation wave. • The dependence on cellular structure on the heterogeneity of initiation sources, chemical kinetic mechanism is studied • On achieving the self-sustaining propagation mode the wave structure becomes independent on initiation conditions and "forgets" initial heterogeneity • Comparison of simulation results with experimental data demonstrates a good agreement. [ABSTRACT FROM AUTHOR]