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The Lucy mission is NASA’s 13th Discovery-class mission and the first mission to the Trojan asteroids. The spacecraft conducts flybys of 8 Trojan asteroids over the course of 12 years. A series of 3 Earth Gravity Assists are used to increase the aphelion of the spacecraft’s orbit and to target the final Trojan asteroid flyby. Over the course of 2 years the spacecraft conducts 4 flybys in the L4 swarm to explore 6 Trojan asteroids, which includes two small satellites. Near the end of the mission, Lucy flies past the near-equal size binary, Patroclus-Menoetius, in the L5 swarm. The concept of operations for the Trojan flybys invokes a standard timeline for spacecraft operations to allow a science sequence that is tailored to each Trojan asteroid. The concept of operations enables efficiency of observations and resiliency in the observing sequence to robustly meet the Lucy science requirements. [ABSTRACT FROM AUTHOR]

From 1 January 2010 to 24 August 2023, 566 orbital stages and 511 spacecraft with a radar cross section >1 m2 have re-entered without control the Earth's atmosphere. The total returned mass was 1650 metric tons, corresponding to a mean of 115 metric tons per year. 77% of the mass belonged to orbital stages, 23% to spacecraft. The uncontrolled re-entries of orbital stages are currently dominated by China, accounting for more than half of the decaying mass, while for spacecraft 2/3 of the mass belongs to American satellites. 60% of the re-entries occurred within 2 years of the launch. The ground casualty expectancy due to orbital stages was always predominant over that from spacecraft, by an average factor of nearly three. From 2010 to 2018, the total casualty probability remained substantially stable, with a mean annual value just over 1%. Since 2019, instead, the annual casualty probability of both spacecraft and orbital stages progressively increased, reaching a total value of around 3% in 2022 and 2023 (extrapolated). Even assuming a stable launch activity, in the coming years, when many of the recently launched spacecraft will start to re-enter, the casualty expectancy of orbital stages will remain basically the same, while that of spacecraft might progressively increase by a factor of 20. This would lead to an annual casualty probability of about 20%, even more in case of a further growth in launch activity, very likely based on current forecasts. The quick implementation of widespread and effective mitigation measures, like controlled de-orbiting and design for demise, is therefore necessary, to prevent the situation from deteriorating too much. • The statistics and ground casualty risk of uncontrolled re-entries of large spacecraft and orbital stages were revisited. • The uncontrolled re-entries of orbital stages are currently dominated by China, accounting for more than half of the decaying mass, while for spacecraft 2/3 of the mass belongs to American satellites. • Since 2019, the annual casualty probability of both spacecraft and orbital stages progressively increased, reaching a total value of around 3% in 2022. • 70% of the casualty expectancy is associated with the uncontrolled re-entries of orbital stages, 30% with the re-entries of satellites. • It is necessary to take effective mitigation measures as soon as possible, for instance requiring the generalized adoption of controlled de-orbiting for large orbital stages and satellites. [ABSTRACT FROM AUTHOR]

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

Spacecraft are subjected to a harsh and complex dynamic environment during launch and on-orbit operation, necessitating the use of vibration isolation techniques to protect the onboard components and the payload from harmful vibration. Moreover, the stringent requirements for modern spacecraft's payload stability and pointing accuracy have fueled research for higher-performance vibration isolation. Consequently, a number of vibration isolation systems based on active, passive, semi-active, and hybrid isolation methods have been proposed. This study employed a systematic literature review methodology to examine the most recent methods for vibration isolation in spacecraft, taking into account both whole-spacecraft vibration isolation (WSVI) and microvibration isolation (MVI) and primarily referencing works published in the literature between 2012 and 2022. The review highlighted current research trends in vibration isolation methods and implementation strategies reported in the literature, as well as the status of research on WSVI and MVI. The findings suggest that more research was conducted in the domain of MVI compared to WSVI, with passive techniques accounting for about 42 % of the articles reviewed. This could be attributed to the advances made in smart materials, metamaterials, and the quasi-zero-stiffness concept to address the shortcomings of traditional passive isolators. To this end, remarks on the limitations of the current studies and future prospects were highlighted. [ABSTRACT FROM AUTHOR]

Fundamental lateral mode frequency of a spacecraft is determined experimentally by performing a base excited vibration testing. For the above vibration testing, the spacecraft is mounted on the slip table of the shaker system. It is shown in this work that even if the test fixture is quite stiff, the fundamental lateral mode of the spacecraft can be influenced by the slip table system. The frequency obtained in the test is lower than the base fixed frequency. In this mode, even though the test fixture does not undergo bending, the slip table undergoes bending resulting in rotation of the test fixture. The stiffness of the bearing and the bending stiffness of the slip table are the causes of reduction in stiffness. A mathematical model that represents this behavior is developed for two typical shaker-slip table systems, one having a capacity of about 150 kN force and the other about 300 kN force. It is seen that the test system can cause a drop of about 3 Hz for a 3000 kg class of spacecraft, while using 150 kN shaker system. This is quite significant as their frequencies are in the range of 12–15 Hz. [ABSTRACT FROM AUTHOR]

To solve the non-uniformity of stress in space membrane structure and the lack of shear compliant border configuration design method, shear compliant borders are designed, optimized, and verified in terms of configuration. Firstly, an orthotropic model of the borders is built by combining Hill and Christensen-Lo composite material models. Secondly, a finite element form-finding method is put forward by establishing rectangular and cylindrical coordinates in different areas. The configuration of borders is obtained and the influence of the borders on the edge of the membrane is 0.23%, which means that the borders are compatible with the existing tensegrity systems, especially the tensioning components and the cable sleeves. Thirdly, simulation verifies that borders can cut the spread of shear stress and improve the stress uniformity in membrane structure. The maximum stress in the membrane effective area is decreased by 35.6% and the stress uniformity is improved by 30.5%. Finally, a membrane extension experiment is committed to compare the flatness of membrane surface under shear stress with and without shear compliant borders. The borders decrease the increment speed of flatness by 58.1%, which verifies the amelioration of stress uniformity. The shear compliant border configuration design method provides a reference for space membrane structure stress-uniform design. [ABSTRACT FROM AUTHOR]

Cameras are rapidly becoming the choice for on-board sensors towards space rendezvous due to their small form factor and inexpensive power, mass, and volume costs. When it comes to docking, however, they typically serve a secondary role, whereas the main work is done by active sensors such as lidar. This paper documents the development of a proposed AI-based (artificial intelligence) navigation algorithm intending to mature the use of on-board visible wavelength cameras as a main sensor for docking and on-orbit servicing (OOS), reducing the dependency on lidar and greatly reducing costs. Specifically, the use of AI enables the expansion of the relative navigation solution towards multiple classes of scenarios, e.g., in terms of targets or illumination conditions, which would otherwise have to be crafted on a case-by-case manner using classical image processing methods. Multiple convolutional neural network (CNN) backbone architectures are benchmarked on synthetically generated data of docking manoeuvres with the International Space Station (ISS), achieving position and attitude estimates close to 1 % range-normalised and 1 deg, respectively, an established rule of thumb for the navigation measurement accuracy during final approach. The integration of the solution with a physical prototype of the refuelling mechanism is validated in laboratory using a robotic arm to simulate a berthing procedure. • A novel AI-based solution for spacecraft docking and refuelling is presented. • First demonstration of deep learning for vision-based pose estimation in docking. • Various CNN backbones are benchmarked on photorealistic ISS refuelling datasets. • ResNet-101 reports average 1.19% error in position, 0.29 ∘ in attitude per trajectory. • Prototypes of the developed mechanism are validated in lab with 7-DOF manipulator. [ABSTRACT FROM AUTHOR]

The capabilities of the SIRIUS-S software environment for designing spacecraft motion control systems - modeling, synthesis of guidance and control laws, nonlinear analysis, computer simulation and graphical display of dynamic processes - are briefly represented. Simulation results on guidance and control algorithms are presented that demonstrate their effectiveness. [ABSTRACT FROM AUTHOR]

This article discusses an experimental study on the combustion of water vapor and aluminum powder for chemical micropropulsion. The study aims to address the constraints and safety requirements that have delayed the development of this propellant combination. The researchers measured the combustion pressure and reacted aluminum fraction under specific conditions, such as water vapor pressure under 100 kPa, aluminum powder diameter over 10 μm, and a combustion chamber smaller than 1 L. The results showed that the nitrogen gas promoted combustion by diffusing the powder, and controlling powder diffusion and ignition timing were important factors. This research provides valuable insights for the development of chemical micropropulsion systems. [Extracted from the article]

Thermal control is a critical functionality in space applications due to the narrow operation temperature range of the on-board systems, and, on the other hand, due to the harsh environment the spacecraft is subject to. Thermal control systems based on phase change materials have the main advantage that are passive and, if properly designed, are highly reliable and efficient. Some Phase Change Materials (PCMs) – paraffins – have other applications for spacecrafts, such as mechanical actuators, which convert temperature changes to mechanical work. In contrast to thermal control terrestrial applications, space applications have more constraints and must ensure compatibility with more systems. One of the most important constraints in space systems is the mass. The design and choice of thermal regulation systems often boils down to replacing heat dissipation radiators mass with lighter, PCM elements. The review is centered around the main application area of PCMs in space applications, discussing numerical and experimental studies on the design and multi-objective optimization of thermal control systems for spacecrafts. Other PCM applications in spacecraft technology are discussed such as micro-actuators, thermoelectric devices, propulsion systems and thermal protection for reentry vehicles. Not only PCM applications are discussed but also some specific issues which are highly important in the design and effective operation of PCM-based thermal control systems for spacecrafts. Thus, issues such as materials, performance enhancement, stability, compatibility and phase transition process under microgravity or hyper gravity were reviewed. • Phase change materials applications in space systems were reviewed. • Phase change materials applications other than heat storage were discussed. • Microgravity and hyper gravity effects on the phase transition process were discussed. • Thermal efficiency enhancement methods were reviewed. • Corrosion and stability issues were discussed. [ABSTRACT FROM AUTHOR]

The aim of the present research endeavor is to propose a framework to model, simulate and control the reorientation of an exploration-class spacecraft for space observation operations. Reorientation under boresight evasion requirements will be tackled by a virtual-potential theory which encodes a reorientation target as well as mandatory-direction-type and forbidden-direction-type attitudinal constraints. The mathematical model of the rotational motion of a spacecraft as well as the control fields are entirely written by coordinate-free Lie-group-type formulations. Numerical simulations complement the theoretical endeavor and illustrate the achieved progress. • Spacecraft reorientation under boresight evasion is a fundamental maneuver in space missions. • Reorientation by virtual potential is capable of tackling mandatory and forbidden attitudinal constraints. • Coordinate-free modeling and control are advantageous in modern spacecraft maneuvering. [ABSTRACT FROM AUTHOR]

Here we describe the novel, multi-point Comet Interceptor mission. It is dedicated to the exploration of a little-processed long-period comet, possibly entering the inner Solar System for the first time, or to encounter an interstellar object originating at another star. The objectives of the mission are to address the following questions: What are the surface composition, shape, morphology, and structure of the target object? What is the composition of the gas and dust in the coma, its connection to the nucleus, and the nature of its interaction with the solar wind? The mission was proposed to the European Space Agency in 2018, and formally adopted by the agency in June 2022, for launch in 2029 together with the Ariel mission. Comet Interceptor will take advantage of the opportunity presented by ESA's F-Class call for fast, flexible, low-cost missions to which it was proposed. The call required a launch to a halo orbit around the Sun-Earth L2 point. The mission can take advantage of this placement to wait for the discovery of a suitable comet reachable with its minimum Δ V capability of 600 ms − 1 . Comet Interceptor will be unique in encountering and studying, at a nominal closest approach distance of 1000 km, a comet that represents a near-pristine sample of material from the formation of the Solar System. It will also add a capability that no previous cometary mission has had, which is to deploy two sub-probes – B1, provided by the Japanese space agency, JAXA, and B2 – that will follow different trajectories through the coma. While the main probe passes at a nominal 1000 km distance, probes B1 and B2 will follow different chords through the coma at distances of 850 km and 400 km, respectively. The result will be unique, simultaneous, spatially resolved information of the 3-dimensional properties of the target comet and its interaction with the space environment. We present the mission's science background leading to these objectives, as well as an overview of the scientific instruments, mission design, and schedule. [ABSTRACT FROM AUTHOR]

The Lucy spacecraft successfully performed the first of two Earth Gravity Assist maneuvers on October 16th 2022, flying 360 km above the Earth's surface at 11:04 UT. The flyby was essential for the Lucy mission design, but also provided a wealth of data for scientific, calibration, and public engagement purposes. The Earth and Moon provided excellent calibration targets, being large, bright, and well-characterized, though instrument saturation was sometimes an issue, as the instruments are designed for operation 5 AU from the sun. Calibration data of the Earth and/or Moon were taken with all Lucy instruments, improving knowledge of instrument alignment, stray light characteristics, and sensitivity to resolved targets. In addition, Lucy obtained scientifically valuable thermal emission spectra of the Moon, and extensive images of the DART mission impact into the Didymos system, from a unique geometry, 20 days before the Earth flyby. [ABSTRACT FROM AUTHOR]

After the space robot captures the out-of-control spacecraft, they form a combined spacecraft. For the parameter identification problem of combined spacecraft with complex angular rate measurement noise, deep learning could provide a new solution, due to its ability to extract robust features from noisy data. However, the deep neural network (DNN) has substantial computing resource requirements, which restricts its application in spacecraft systems. To solve the above problems, an inertia matrix identification approach combining a structure optimization strategy of DNN is proposed to identify the inertia matrix in the presence of complex measurement noise using the DNN with optimized structure. The attitude dynamics of combined spacecraft are firstly derived, and the identification problem considering complex measurement noise is described. Based on the dynamics model, the construction process of DNN for accurately identifying the inertia matrix is proposed. A structure optimization strategy of DNN is then designed to optimize the structure of DNN for increasing computing efficiency. This strategy provides a whole set of solutions including the structure selection of the initial DNN, the redundant neurons pruning process, and the end-of-optimization condition. By combining the proposed construction process of DNN, the structure of the DNN is finally optimized by iterating the structured pruning and retraining step. The performance of the proposed DNN with optimized network structure is verified on a testing set with complex measurement noise. The results show that the proposed approach can accurately identify the inertia matrix of combined spacecraft, and meanwhile greatly improve the computing efficiency of DNN. [ABSTRACT FROM AUTHOR]

At the European Space Agency (ESA), Modeling and Simulation (M&S) plays a fundamental role during the lifetime of a spacecraft, being used from the design phase to the testing and during operations in space. M&S tools embed general physics-based models and disciplines characterized by configurable parameters which have to be calibrated in order to mimic the behavior of the actual flying spacecraft. However, their calibration requires a large number of simulations which are unfeasible to be obtained through computationally expensive high-fidelity simulation models. Thus, the inability to calibrate the high-fidelity simulation models poses limitations for the use of M&S tools during spacecraft operations. In this light, the present work proposes the use of a surrogate model-based approach for the calibration of simulation models of spacecraft. The approach integrates a computationally inexpensive deep-learning-based surrogate model, which mimics the high-fidelity simulation model without requiring the same computational burden, and a metaheuristic optimization algorithm, to identify the optimal values of the simulation model configurable parameters. This enhances the capabilities of M&S tools and allows their use in operations. The approach's effectiveness is shown by its application to real flying Earth observation satellite data and simulation models. [ABSTRACT FROM AUTHOR]

In this article, the attitude tracking problem for rigid spacecraft subject to external disturbance, inertia uncertainties, and actuator faults is investigated. A novel continuous control strategy is proposed to guarantee that the attitude tracking errors converge with desired performance before an appointed time. In addition, the proposed control law can enable the closed‐loop system to be unwinding‐free. For the purpose, a modified appointed‐time performance function (APF) is used to construct a disturbance observer‐based adaptive attitude control law based on the special orthogonal group (SO(3)). With the aid of the modified APFs, desired performance metrics, including transient and steady‐state performance, on the attitude tracking error can be specified in advance without the requirement for initial values of the tracking system to avoid singularity. By resorting to error transformation and backstepping technique, the control design process is simplified and a few control gains are involved in the proposed control law. Rigorous stability analysis based on the Lyapunov theory is provided to prove the appointed‐time stability of the attitude tracking error as well as the boundedness of all signals in the closed‐loop system despite external disturbance, inertia uncertainties, and actuator faults. Finally, numerical simulations are conducted to support the analysis. [ABSTRACT FROM AUTHOR]

This study introduces a data-driven surrogating modeling approach for determining thermal parameters in undetermined thermal mathematical models (TMM) used in spacecraft systems. Maintaining appropriate temperature levels for onboard components is crucial for the success and safety of space missions. However, accurately determining thermal parameters in the undetermined TMM presents a challenge. To address this, we propose an approach that combines TMM correlation with test data and data-driven surrogate modeling. Our approach minimizes an objective function that measures the deviation between predicted temperatures from a surrogate model, incorporating thermal parameters, and steady state temperatures obtained from thermal vacuum balance test results. This optimization is performed using a genetic algorithm while imposing constraints to ensure the physical validity of the thermal parameters and correlation criteria related to temperatures. We demonstrate the effectiveness and validity of our approach by applying the obtained thermal parameters to the TMM. The resulting calculated temperatures exhibit excellent agreement with the test results. This methodology offers a user-friendly and reliable way to determine thermal parameters in undetermined TMM within spacecraft systems, ultimately enhancing temperature control and mission success. • Efficient thermal parameter determination in spacecraft TMMs through data-driven surrogate models. • TMM correlations achieved by minimizing temperature deviation with surrogate models and test data. • Reliable TMM correlations achieved by considering realistic constraints, overcoming thermal parameter non-uniqueness. • Reasonable selection of thermal parameters, enhancing reliability and reducing user knowledge requirements. [ABSTRACT FROM AUTHOR]

This paper investigates the fault-tolerant attitude control problem for a spacecraft with actuator faults using an event-triggered-based concurrent-learning control (CLC) approach. A concurrent-learning controller, updating itself through previous control signal values to generate the current control signal, is studied in this paper. The concurrent-learning controller is designed under an event-triggered policy to compensate for the actuator's data transmission constraints caused by bandwidth limit. In addition, a robust fault observer is designed to facilitate the concurrent-learning controller in compensating for the actuator faults. The key feature of the given approach is in the simplicity of control structure to stabilize spacecraft attitude considering actuator faults, model uncertainty, actuators' dynamics effect, and data transmission; simulation results show the proposed approach's efficiency. [ABSTRACT FROM AUTHOR]

AbstractTo address the challenges of spacecraft design involving over and under-designing, uncertainty thermal analysis offers a practical solution. This method typically requires extensive simulations under varying conditions, leading to high computational costs, primarily due to the complex ray tracing calculations for solar radiation external heat flux in different scenarios. This study introduces an solar radiation external heat flux expansion (EHFE) formula that simplifies these calculations. By applying ray tracing only once and using matrix operation for other conditions, it significantly reduces the need for complex calculations. Additionally, this approach effectively eliminates errors caused by the cut-off factor. An example with a specific spacecraft demonstrates that this new method, while enhancing computational efficiency, maintains the accuracy of traditional uncertainty-based calculations for solar radiation external heat flux. [ABSTRACT FROM AUTHOR]

The paper discusses priority scientific tasks related to the study of small bodies in the Solar System, identifies the most promising objects for investigation from a spacecraft on a flyby trajectory and the sample return mission, and develops proposals for the preliminary composition of scientific instruments for remote asteroid research methods. A long-term and phased Russian scientific program for studying small Solar System bodies using spacecraft with electric propulsion has been proposed. The project is designed in such a way as to explore the largest number of scientifically interesting asteroids using a smaller number of spacecraft. A design concept for a small spacecraft to investigate near-Earth asteroids on a flyby trajectory and a main spacecraft for studying metallic asteroids in the Main Belt and sample sample return has been developed. A ballistic analysis of the flyby of five near-Earth asteroids and three metallic asteroids in the Main Belt is presented, as well as a ballistic analysis of the sample return mission from a Main Belt asteroid. The option of sample return using the nuclear tug Zevs is also considered. [ABSTRACT FROM AUTHOR]

A reconfigurable metasurface constitutes an important block of future adaptive and smart nanophotonic applications, such as adaptive cooling in spacecraft. In this paper, we introduce a new modeling approach for the fast design of tunable and reconfigurable metasurface structures using a convolutional deep learning network. The metasurface structure is modeled as a multilayer image tensor to model material properties as image maps. We avoid the dimensionality mismatch problem using the operating wavelength as an input to the network. As a case study, we model the response of a reconfigurable absorber that employs the phase transition of vanadium dioxide in the mid-infrared spectrum. The feed-forward model is used as a surrogate model and is subsequently employed within a pattern search optimization process to design a passive adaptive cooling surface leveraging the phase transition of vanadium dioxide. The results indicate that our model delivers an accurate prediction of the metasurface response using a relatively small training dataset. The proposed patterned vanadium dioxide metasurface achieved a 28% saving in coating thickness compared to the literature while maintaining reasonable emissivity contrast at 0.43. Moreover, our design approach was able to overcome the non-uniqueness problem by generating multiple patterns that satisfy the design objectives. The proposed adaptive metasurface can potentially serve as a core block for passive spacecraft cooling applications. We also believe that our design approach can be extended to cover a wider range of applications. [ABSTRACT FROM AUTHOR]

With the increasing demand ol spacecraft lor 3D-printing metal materials, spacecraft propose higher requirements for reliability of 3D-printing metal materials. Based on the analysis results of service environment in spacecraft, the characteristics of 3D-printing materials and the application lunctions in spacecraft, this paper established the properties evaluation technical index system, designed the technical route of scheme and delined the properties index evaluation method Irom the perspective ol 3D-printing metal raw materials, typical components and products. Finally, a technical scheme lor reliability evaluation ol 3D-printing metal materials for spacecraft was lormed, which provided technical support for the sale and reliable application ol 3D-printing materials in the aerospace lield. [ABSTRACT FROM AUTHOR]

This paper investigates the attitude control of rigid spacecraft with uncertainties and actuator faults. An immersion and invariance sliding mode control with time-delay estimation (IISMCTDE) is proposed to cope with uncertainties, disturbances, and actuator faults. First, time-delay estimation is utilized to reconstruct the total disturbances including uncertainties, disturbances, and actuator faults. Then, based on the time-delay estimation, an immersion and invariance sliding mode control (IISMC) scheme is developed to achieve accurate attitude tracking from the reconstructed knowledge. The proposed IISMCTDE can not only achieve finite-time convergence but also be simple and easy to operate. Simulations are provided to demonstrate the effectiveness of the proposed approach. [ABSTRACT FROM AUTHOR]

To realize rapid, safe and optimal avoidance for spacecraft in complex threat situations and incomplete-information conditions, and considering the problems in existing hierarchical decision-making modes, a learning-based spacecraft reactive anti-hostile-rendezvous (AHR) maneuver control framework is proposed in this paper. First, a fundamental AHR action planner based on the interfered fluid dynamical system (IFDS) is proposed, in which the avoidance strategies for hostile and non-adversarial threats are appropriately integrated. An avoidance strategy matched with the IFDS is further designed by specifically considering high-intensity confrontation situations where a threat maintains persistent pursuit. Then, the "state-action" solution methods based on deep reinforcement learning (DRL) and multi-agent-DRL (MADRL) methods for orbital avoidance maneuvers are proposed to control the avoidance time and directions by optimizing the IFDS parameters. In the offline training stage, the DRL method will first simulate the interactive processes between the spacecraft and non-adversarial threats so that the spacecraft can gradually adapt to complex space environments and generate the relative primary avoidance maneuver strategies. On this basis, for hostile threats, the MADRL-based solution method of orbit game strategies is further established, in which the spacecraft and hostile threats are constructed as two types of opposite agents to simulate the game processes so that the strategy levels can be bilaterally evolved and the AHR success rates can be improved. After training, the corresponding neural networks are extracted from the agents in the DRL-based and MADRL-based methods for reactive AHR maneuver control. Moreover, a progressive agent training mechanism matched with the above DRL-based and MADRL-based solution methods is proposed, which can reduce the invalid interactions in initial training episodes, improve the training efficiency, and enhance the generalization ability of orbital game strategies. Finally, the effectiveness of the proposed framework is demonstrated by simulations. [ABSTRACT FROM AUTHOR]

This paper describes the method of determining reasonable test levels for product of spacecraft with big size, low mass and high test levels when the force of the shaker facilities is limited. The major step is to compute the test envelope from both Newton's law and Ampere's force aspects, the minimum of which is defined as the final safety control envelope for flight spacecraft hardware. By using this method, the shaker facilities can provide maximum force for spacecraft experiencing effective assessment and the interruptions can be avoid during the process of test operation. [ABSTRACT FROM AUTHOR]

This paper reports the development of a 10Ah lithium-ion pouch battery cell using stainless-steel laminated film as the casing material for JAXA's SLIM lunar lander. By selecting the pouch format, the mass ratio of the case and electrode terminals is significantly reduced, which is superior to conventional space batteries using prismatic metal cases. A high specific energy of approximately 130Wh/kg was achieved. In addition, unlike standard batteries using aluminumlaminated film, the battery developed during this study deforms little when subjected to a vacuum. It can thus be mounted on a spacecraft with a simple, lightweight bracket, contributing to higher specific energy as a module. This paper reports the results of environment tests for the SLIM mission and the basic electrical characteristics of the battery. [ABSTRACT FROM AUTHOR]

Thermal analysis of spacecraft systems is a critical process for mission operations. However, knowing the temperature distribution of entire systems is not easy due to the uncertainty of the thermal mathematical model (TMM) and limited temperature sensors. This paper proposes a temperature estimation method using physics-informed machine learning (PIML). The PIML-based thermal analysis allows us to estimate the actual temperature distribution by seamlessly bridging the limited observations and the TMM. To evaluate the estimation accuracy of the proposed method, we conducted a numerical experiment using a pseudo small satellite model consisting of 100 nodes. The proposed method was applied to three different model error cases and was found to improve temperature estimation accuracy in all cases. In addition, the impact of the number of temperature sensors and their placement on estimation accuracy was investigated. • A data-driven thermal analysis that can bridge the uncertain model and limited measurements is proposed. • The physics-informed machine learning was applied to the thermal mathematical model. • A numerical experiment was conducted using a pseudo small satellite model. • The effect of the number and placements of sensors on the estimation accuracy for three model error scenarios was discussed. [ABSTRACT FROM AUTHOR]

• A novel I&I adaptive scheme is presented. • Immersion and invariance based adaptive terminal sliding mode control is proposed. • Immersion and invariance based adaptive terminal sliding mode control with saturation compensation is presented. This work investigates attitude control of spacecraft with uncertainties and input saturation. An immersion and invariance based adaptive robust control (IIARC) is proposed to improve tracking performance of spacecraft. The proposed approach consists of an immersion and invariance (I&I) estimator, a robust adaptive law, and a terminal sliding mode control (TSMC). The continuous fast TSMC is designed to stabilize the system and obtain fast convergence rate. To estimate the disturbances and uncertainties, I&I methodology is introduced into the proposed approach. The robust adaptive law is applied such that the prior information of disturbances is not required. IIARC with saturation compensation (IIARCSC) is presented to reduce the effect of input saturation on the attitude tracking of spacecraft. It is proved that the developed approach could guarantee the finite-time convergence (FTC). Extensive comparative simulations demonstrate the effectiveness of the presented scheme. [ABSTRACT FROM AUTHOR]

Large increases in the number of low earth orbit satellites are projected in the coming decades [L. Schulz, K.-H. Glassmeier, Adv. Space Res. 67, 1002–1025 (2021)] with perhaps 50,000 additional satellites in orbit by 2030 [GAO, Large constellations of satellites: Mitigating environmental and other effects (2022)]. When spent rocket bodies and defunct satellites reenter the atmosphere, they produce metal vapors that condense into aerosol particles that descend into the stratosphere. So far, models of spacecraft reentry have focused on understanding the hazard presented by objects that survive to the surface rather than on the fate of the metals that vaporize. Here, we show that metals that vaporized during spacecraft reentries can be clearly measured in stratospheric sulfuric acid particles. Over 20 elements from reentry were detected and were present in ratios consistent with alloys used in spacecraft. The mass of lithium, aluminum, copper, and lead from the reentry of spacecraft was found to exceed the cosmic dust influx of those metals. About 10% of stratospheric sulfuric acid particles larger than 120 nm in diameter contain aluminum and other elements from spacecraft reentry. Planned increases in the number of low earth orbit satellites within the next few decades could cause up to half of stratospheric sulfuric acid particles to contain metals from reentry. The influence of this level of metallic content on the properties of stratospheric aerosol is unknown. [ABSTRACT FROM AUTHOR]

Solar-hydrogen energy for spacecraft, space interplanetary stations, space bases on the planets of the solar system must reliably provide energy: spacecraft during peak periods (repairs at a space facility, energy-consuming experiments, cleaning of facilities, regeneration of waste to extract oxygen and hydrogen, etc..d.) and bases on planets, especially where the night lasts for a long time. The onboard hydrogen module of the solar-hydrogen plant includes: a container with water; distiller; deionizer; water lines with adjustable shut-off valves; electrolyzer; cryogenic side tanks with hydrogen and oxygen; fuel cell, hydrogen safety sensor systems based on leakage and concentration sensors; level gauges; cryogenic and gas pipelines of hydrogen and oxygen; onboard cryo-refrigerator; automatic control system for the hydrogen module. The article presents the results of experimental studies of static, dynamic and mass-dimensional characteristics of power conversion equipment (hereinafter - PCE) in the power supply system (hereinafter - PSS) of the spacecraft (hereinafter - spacecraft) with a digital control system. The purpose of experimental research is to verify the theoretical results obtained earlier in the design of digital PCE control systems of high-voltage PSS spacecraft, including the study of static and dynamic characteristics. We developed and manufactured an experimental PCE sample with an output power of 2500W in this research. It includes: four voltage stabilization modules and the 100V load supply output bus filter. Each voltage stabilization module includes a battery energy conversion channel, a solar battery energy conversion channel and a digital control system unit. The power conversion channel of the accumulator battery is designed and made on the basis of a bridge volt-booster inverter-transformer circuit. Voltage control in energy converter of accumulator battery is based on phase shift in control pulses of transistors in adjustable inverter rack related to unregulated one. The power conversion channel of the solar battery is designed and made on the basis of a two-phase circuit of a direct voltage converter of an increasing type with switching throttles. The digital control system of the voltage stabilization module is designed on the basis of analogues in the modern electronic component base in the Space class and consists of four identical channels. Each channel of the control system controls its own voltage stabilization module. We obtained the following results during our experimental studies: some pulsations of the output voltage of the energy conversion equipment during operation on an active load, transient processes of the output voltage of the energy conversion equipment during pulsations of the load current without changing the mode of operation, transient processes of the output voltage of the energy conversion equipment when changing the modes of its operation and weight and size characteristics for different values of the output power. The obtained results of experimental studies confirm the previously obtained theoretical results, and also showed that the use of a distributed digital control system will increase the specific energy characteristics of the newly developed PSS spacecraft while ensuring the given static and dynamic characteristics of output voltage quality. The results of the work are designed to create a new on-board power conversion equipment for power supply systems of promising spacecraft. The area of application of the project results is space instrumentation. [ABSTRACT FROM AUTHOR]

In the context of nonlinear systems, variational Bayesian filters play an important and alternative role to sampling‐based Monte Carlo methods. Although they have been used in conjunction with derivatives of Kalman filters, their behavior, both in terms of convergence, stability, and a solution without a stepwise approximation procedure, has not been well explored. In this paper, we propose a novel, stochastically stable, stepwise approximation‐free variational Bayesian Kalman filter (hence, termed as VBKF) for nonlinear systems. The proposed filter, unlike other posterior distribution estimation methods, is very flexible with the number of parameters that should be known a priori, offers an excellent convergence rate, stochastic stability, and more specifically in its ability to handle non‐Gaussian, unknown, inaccurate and drifting noise models. Furthermore, one of the key advantages of the proposed solution is that it is stepwise approximation‐free, and as such, less computationally demanding. We also prove that the proposed filter is stochastically stable using the Posterior Cramér–Rao Lower Bound. In addition to the stability, by using Lyapunov function, we show that the estimation errors are exponentially bounded by the mean square error under certain conditions. In addition to this, we also demonstrate the practical applicability of the proposed approach by developing a novel filter for estimating the attitude of spacecrafts using star cameras and gyro‐sensors. We then evaluate the efficacy of the novel filter for attitude estimation, and thus that of the VBKF, using a set of detailed simulations mimicking different initialization errors that are common in attitude estimation. Our results show that the proposed approach offers an excellent convergence rate within conventional 3σ$$ 3\sigma $$ bounds, better root mean square error when compared against several of the common, and state‐of‐the‐art methods. [ABSTRACT FROM AUTHOR]

This paper conceptualizes the main principles of comprehensive software verification for an onboard spacecraft control system. An optimal comprehensive verification strategy for onboard software is selected by rigorously stating and solving the corresponding optimization problem. Software verification methods with functional correctness indicators are proposed. [ABSTRACT FROM AUTHOR]

We have been conducting hypervelocity impact tests and numerical simulations to develop a Japanese spacecraft design guideline for the protection of satellites from a certain degree of M/OD impacts. A working group is formed by members of JAXA, spacecraft manufacturers, experts, and researchers in the field of hypervelocity impacts to investigate the effects of M/OD impacts on satellite critical parts and bumpers by hypervelocity impact (HVI) tests and analysis. The knowledge acquired is now being reflected in the spacecraft design guidelines. We investigated the effect of hypervelocity impacts of micrometeoroids and small-scale space debris on space structures by comparing numerical simulation results obtained using the AUTODYN-2D hydrocode with the results of experiments using a two-stage light gas gun since 2005. We will consider the case of a satellite hydrazine fuel tank and the consequences of a hypervelocity impact from space debris. The final purpose of this study is to better understand the mechanisms of detonation of hydrazine vapor during a hypervelocity impact on a low-pressure reservoir. As part of this effort and an initial study, the response of a Ti–6Al–4V plate which is used tank material to 6 km/s high-velocity impacts is shown and discussed. It was found that the ballistic limit of the titanium tank material for both curved and flat plates was generally predictable using the existing ballistic limit equation. The effect of curvature on the impact scar size of flat (t = 1.2 mm) and curved plates was examined, and no significant difference was found. • We introduced our achievements and plans overview related protection for small debris and meteoroid related space in Japan. • The ballistic limit of the titanium tank material of spacecraft was predictable using the existing ballistic limit equation. • The effect of curvature on the size of the hiper-velocity impact scar was examined, but no significant difference was found. [ABSTRACT FROM AUTHOR]

We propose an analysis of the recovery strategies for the science mode of the Laser Interfer-ometer Space Antenna (LISA) mission after a meteoroid impacts the spacecraft. The mission consists of a three-spacecraft constellation traveling in a heliocentric orbit, detecting gravitational waves through laser interferometry. To this end, each spacecraft must travel in a free-fall condition in order to reject any possible disturbance and noise affecting the control loop. Nevertheless, if one of the three satellites crosses a meteoroids stream, the collisions can produce attitude perturbations that must be compensated by the control loop. Indeed, in this latter case, the interferometer laser links can be lost. Unfortunately, the link recovery is accomplished through a quite time-consuming re-acquisition maneuver, implying a significant reduction of the science mode time. For this reason, we propose different strategies for a fast recovery of the nominal attitude. The strategies are supported and traded-off by means of extensive simulations, including a Monte Carlo campaign and a worst-case analysis. • Different GNC recovery strategies for LISA mission are identified and compared. • An attitude guidance and control system is designed for handling the attitude perturbation due to micrometeoroid impacts. • Different navigation filters, based on sensors measurements, are designed and compared. • Monte Carlo and worst-case simulations are presented to verify the effectiveness of the proposed GNC strategies. [ABSTRACT FROM AUTHOR]

This study examines geomagnetic storm conditions for attitude estimation of a spacecraft having magnetometers with various measurement noise levels at low Earth orbit. The geomagnetic field models are introduced by taking the external field effects into account. It is important to discuss the limitations of the models and the measurements that are being used. Therefore, the external effects are evaluated under different noise levels of magnetometer to determine and discuss the suppression level caused by the noise levels in a general framework. External magnetic field effects are examined by the ratio to noise under various magnetometer standard deviations up to 300 nT. An analysis is carried out for attitude estimation using Kalman-type filter in order to determine what noise level is acceptable on a particular sensor with a specified attitude requirement, magnetic field model, and space weather conditions. This study indicates that external magnetic fields are vital for establishing attitude processes based on magnetometers with varying noise on which the spacecraft relies for accuracy. [ABSTRACT FROM AUTHOR]

The attitude tracking control problem of rigid spacecraft subjected to external disturbance and prescribed performance is studied. A fast fixed‐time stability theorem is firstly developed. A continuous disturbance observer is then designed with its estimation error can be driven into a tiny neighborhood containing zero within fixed time. Moreover, the spacecraft attitude tracking error constrained by prescribed performance function is transformed into an unconstrained dynamics. A disturbance observer‐based prescribed performance tracking controller is designed to stabilize the unconstrained dynamics within fixed time. It also means that the attitude tracking errors meet the prescribed performance as the stabilization of the unconstrained dynamics. Spacecraft simulation examples are finally given to validate the effectiveness of the proposed controller. [ABSTRACT FROM AUTHOR]

This work presents the results of complex experimental and computational studies of the shock-wave and thermodynamic parameters of the collision of a particle with the wall of the pressurized compartment of a spacecraft, depending on the physical and mechanical characteristics of the impactor material and the impact velocity. The results of calculations of the parameters of the wall breakdown and air waves in the pressurized compartment of the spacecraft are compared with the data of experimental measurements. [ABSTRACT FROM AUTHOR]

Recent advancements in space technology and reduced launching cost led companies, defence and government organisations to turn their attention to low Earth orbit (LEO) and very low Earth orbit (VLEO) satellites, for they offer significant advantages over other types of spacecraft and present an attractive solution for observation, communication and other tasks. However, keeping satellites in LEO and VLEO presents a unique set of challenges, in addition to those typically associated with exposure to space environment such as damage from space debris, thermal fluctuations, radiation and thermal management in vacuum. The structural and functional elements of LEO and especially VLEO satellites are significantly affected by residual atmosphere and, in particular, atomic oxygen (AO). At VLEO, the remaining atmosphere is dense enough to create significant drag and quicky de-orbit satellites; thus, thrusters are needed to keep them on a stable orbit. Atomic oxygen-induced material erosion is another key challenge to overcome during the design phase of LEO and VLEO spacecraft. This review covered the corrosion interactions between the satellites and the low orbit environment, and how it can be minimised through the use of carbon-based nanomaterials and their composites. The review also discussed key mechanisms and challenges underpinning material design and fabrication, and it outlined the current research in this area. [ABSTRACT FROM AUTHOR]

A new method is proposed for determining the electron density in rarefied plasma, based on simultaneous measurements of the Interball-2 satellite potential using IESP-2 (electric field instrument) and KM-7 (electron temperature sensor) probe devices. This makes it possible to estimate the photoelectron current density based on a procedure proposed earlier by the authors of this study. The electron concentration was determined only for the positive potential of the spacecraft. The balance equations for the satellite and the probe between the currents of the surrounding plasma electrons and photoelectrons emitted by the illuminated surface were compiled. In the magnetosphere, to bring the probe potential to the potential of the surrounding plasma, a bias current is directed into the probe, which was taken into account in the current balance equation for the probe. The electron energy used in the calculations was kTe = 1 eV. We analyzed data from ~350 orbits in the auroral region of the magnetosphere at altitudes of 2–3 RE from October 1996 to March 1998 during the period of low solar activity at the beginning of the 23rd cycle. Examples of the calculated electron density are given, which is in the range of 1–30 cm–3. [ABSTRACT FROM AUTHOR]

The significance of reliable energy storage systems in spacecraft applications cannot be overstated, since they play a vital role in ensuring continuous power supply and prolonged mission durations. This research deals with the modeling of a hybrid multi-mission radioisotope thermoelectric generator (MMRTG)-lithium-ion (Li-ion) battery integrated energy storage system for spacecraft applications to combine the RTGs' long lifespan and reliability benefits alongside the Li-ion battery's rechargeability and high energy density to achieve a single energy unit. The investigation's main problem was exploring a power unit that improves the limitations of MMRTG and Li-ion batteries to achieve a highly efficient and reliable power supply for autonomous systems, such as a spacecraft. The proposed hybrid system comprises a 110 W/32 V RTG and a 3.6 V/43 Ah Li-ion battery connected to a DC motor through power converters. Results demonstrate the potential of the adopted hybrid energy system in improving the efficiency, reliability, and mission duration of spacecraft missions. The assessment of the hybrid energy system under various load conditions shows that the highest power peak of 3500 W was achieved at a load resistance of 1 Ω. Furthermore, the results show that the hybrid energy system output voltage at temperatures of 253 °K and 293 °K are relatively equal. However, the power cycle was wider and required a long time before dropping. [ABSTRACT FROM AUTHOR]

Currently, managing a group of satellites or robot manipulators requires coordinating their motion and work in a cooperative way to complete complex tasks. The attitude motion coordination and synchronization problems are challenging since attitude motion evolves in non-Euclidean spaces. Moreover, the equation of motions of the rigid body are highly nonlinear. This paper studies the attitude synchronization problem of a group of fully actuated rigid bodies over a directed communication topology. To design the synchronization control law, we exploit the cascade structure of the rigid body's kinematic and dynamic models. First, we propose a kinematic control law that induces attitude synchronization. As a second step, an angular velocity-tracking control law is designed for the dynamic subsystem. We use the exponential coordinates of rotation to describe the body's attitude. Such coordinates are a natural and minimal parametrization of rotation matrices which almost describe every rotation on the Special Orthogonal group S O (3) . We provide simulation results to show the performance of the proposed synchronization controller. [ABSTRACT FROM AUTHOR]

The aim of this paper is to provide information on the radial dependence of turbulence quantities as input for ab initio numerical cosmic-ray modulation models. We present an evaluation and analysis of 48-s resolution Voyager 1 data for the normal component of the magnetic field. These data were preprocessed to fill in data gaps and filter out spurious peaks. The effectiveness of this process is illustrated for a sample data set, for which we show that the underlying spectrum is indeed Kolmogorov. We use a second-order structure function to find correlation times, and also the spectral index and spectral level in the inertial range from expressions derived by Huang et al. (2010). The usefulness of the second-order structure function for data with significant gaps is illustrated by using synthetic data. Gaps up to 68% cause changes of less than 5% in the quantities that we calculate. We show that beyond about 20 au there are significant changes in the behaviour of the magnetic field magnitude and the inertial range spectral index, neither of which has a reasonable explanation. We find that the ratio δ B N / B N remains almost constant out to 20 au, that the correlation times is consistent with results from other studies, and that the spectral index in the inertial range is consistent with the Kolmogorov value. A new result is that we could infer that the break between the energy- and the inertial range has a radial dependence that is consistent with that of the correlation scale. [ABSTRACT FROM AUTHOR]

Summary: This paper investigates the attitude tracking control problem of rigid spacecraft with inertia parameter identification. Based on the relative attitude and angular velocity error dynamics, a basic adaptive backstepping based attitude tracking control scheme is firstly designed such that asymptotic attitude tracking can be achieved. However, the parameter identification error cannot decay to zero if the persistent excitation (PE) condition is not satisfied. To solve this issue, a filtering‐based concurrent learning adaptive backstepping control scheme is then proposed, by incorporating torque filtering technique with concurrent learning technique. A more mild rank condition, which consists of some collectable historical data, is provided to guarantee the convergence of parameter identification error. In addition, a valid data collection algorithm is given. It should be mentioned that a distinctive feature of the proposed filtering‐based concurrent learning adaptive control scheme is that the convergence rate of attitude tracking errors can be improved from asymptotical to exponential. Finally, simulation results are provided to illustrate the effectiveness of the proposed control schemes. [ABSTRACT FROM AUTHOR]

• A novel linear model-based controller is designed for the CEACS attitude regulation. • The Takagi-Sugeno fuzzy model with a parallel distribution approach has been introduced. • The Fuzzy-MPC controller achieves the desired CEACS attitude pointing accuracy. • The proposed controller effectively handles the CEACS actuator constraints. • The CEACS attitude actuation consumes a lower on-board power. A Combined Energy and Attitude Control System (CEACS) is a dual system in which flywheels are used as energy storage and attitude control devices. This work is a progress on CEACS for small satellites to improve the attitude accuracy. In this maiden work, the Fuzzy–MPC controller is introduced to regulate the CEACS attitude with both linear and nonlinear initial angles, and in the presence of actuator constraints. The nonlinear attitude model is transformed into a set of linear subsystems using the Takagi–Sugeno fuzzy modeling approach. The subsystem is designed with a MPC controller and a total control action is summed up using the parallel distributed compensation approach. The numerical results have been analyzed in terms of attitude pointing accuracies, actuator constraints, and energy consumed by the actuators. A performance comparison between the Fuzzy–MPC and Fuzzy–LQR controllers has also been done. The results validate that the Fuzzy-MPC controller achieves the desired CEACS attitude pointing accuracy of 0. 0010 ° with a zero steady-state error and keeps the control torques within the actuator constraints. On the other hand, the Fuzzy–LQR controller gives some measurable steady-state error and achieves the CEACS attitude pointing accuracy of 0. 0012 °. Therefore, the Fuzzy–MPC based CEACS architecture not only achieves the desired pointing accuracies but also produces ten times smaller control torques than the Fuzzy–LQR controller, which result in a lower onboard power consumption. [ABSTRACT FROM AUTHOR]

Thermal analysis of spacecraft involves complex heat flux outside space and heat transfer calculation of equipment inside space, which is the basic guarantee of its long-term stable operation. In this paper, the core module of the Chinese Space Station is taken as an example to conduct hybrid modeling of its complex radiation and heat dissipation network. The external heat flux received by spacecraft is affected by many factors, such as its attitude, orbit parameters, and solar orientation. First, based on the defined reference coordinate system, the whole station is divided into eight quadrants. Then, an intelligent occlusion algorithm is proposed to solve the complex occlusion relationship according to the space quadrant method and fuzzy logic, and then the external heat flux is obtained. Finally, the dynamic response of the station under different conditions is studied by combining traditional mechanism modeling and an intelligent algorithm, and the maximum heat dissipation capacity was obtained. The results show that the temperature control valve opening will seriously affect the main flow of the loop, and the maximum difference can be 0.1 kg/s. In addition, the external heat flux changes periodically, so that the flow of each branch and the valve opening also changes periodically. In the sunlight area, the radiator receiving the external heat flux is about 77,000–90,000 W, and the maximum heat dissipation capacity is 12,000 W. Those values are 28,000 W and 22,000 W, respectively, in the shadowed area. Moreover, the hybrid modeling method proposed in this paper has an important reference value for spacecraft thermal analysis and research in other fields. [ABSTRACT FROM AUTHOR]

This paper deals with the nonfragile fuzzy tracking control technique of a spacecraft subject to external disturbances and control gain uncertainties. First, the Takagi–Sugeno (T–S) fuzzy model of the tracking error dynamics is newly derived. Unlike the existing studies, we assume that the controller output is norm bounded and its gain matrix is perturbed because of the aging of the actuator. Next, a fuzzy disturbance observer (DOB) estimating the external disturbance is constructed to enhance the robustness of the control system. Then, a fuzzy controller feeding back the estimated disturbance is designed to attenuate external disturbances. In addition, we establish the linear matrix inequality (LMI) conditions, by which both the H ∞ disturbance attenuating performance and the asymptotic stabilization are guaranteed. From the solution to the LMI conditions, we obtain the controller and DOB gain matrices. Finally, a simulation example validates the disturbance estimation and tracking performance of the proposed method. [ABSTRACT FROM AUTHOR]

Fungi biofilms have been found growing on spacecraft surfaces such as windows, piping, cables, etc. The contamination of these surfaces with fungi, although undesirable, is highly difficult to avoid. While several biofilm forming species, including Penicillium rubens, have been identified in spacecraft, the effect of microgravity on fungal biofilm formation is unknown. This study sent seven material surfaces (Stainless Steel 316, Aluminum Alloy, Titanium Alloy, Carbon Fiber, Quartz, Silicone, and Nanograss) inoculated with spores of P. rubens to the International Space Station and allowed biofilms to form for 10, 15, and 20 days to understand the effects of microgravity on biofilm morphology and growth. In general, microgravity did not induce changes in the shape of biofilms, nor did it affect growth in terms of biomass, thickness, and surface area coverage. However, microgravity increased or decreased biofilm formation in some cases, and this was incubation-time- and material-dependent. Nanograss was the material with significantly less biofilm formation, both in microgravity and on Earth, and it could potentially be interfering with hyphal adhesion and/or spore germination. Additionally, a decrease in biofilm formation at 20 days, potentially due to nutrient depletion, was seen in some space and Earth samples and was material-dependent. [ABSTRACT FROM AUTHOR]

The concept of solar-hydrogen systems for spacecraft, orbital stations, lunar and Martian bases is currently receiving a new impetus. The supply of solar energy to energy receivers aboard space vehicles is limited. The number of everyday tasks and energy-intensive experiments on board space objects is growing with the development of astronautics. To perform energy-intensive work and experiments, energy consumption exceeds the incoming solar fluxes. The modern solar-hydrogen system ensures the reception, conversion and accumulation of excess incoming energy in the form of chemical energy - hydrogen. Cryogenics makes it possible to miniaturize the solar-hydrogen system. The safety of the cryogenic solar-hydrogen energy system is ensured by hydrogen concentration and leakage sensors. The article proposes a comprehensive solution for miniaturization of Power Conversion Unit (PCU) of Energy Power System (EPS) of spacecrafts (SC), which consists in a joint solution of four key problems of miniaturization of power devices: energy, structural, design and technological, system. Miniaturization of the solar-hydrogen energy system (SHES) is achieved by installing onboard hydrogen and oxygen microcryogenic refrigerators, as well as hydrogen and oxygen cryogenic tanks, water tank, electrolyzer and hydrogen fuel cells (FC). The accumulation of chemical hydrogen energy on board the orbiting spacecraft ensures reliable operation when entering the shadow. Storage of hydrogen in cryogenic form significantly reduces the volume required. A cryocooler based on the Stirling cycle provides the process of liquefying hydrogen after the electrolysis of distilled water. In addition, cryogenic temperatures of 20.2 K can be used to thermostat precision instruments placed on board the spacecraft, For example, to ensure the operation of the SQUID. A 3D - model of a voltage stabilization module (VSM), on the basis of which EEC can be produced with different output power and redundancy depth. We give an example of a complete structural scheme of the EEC, which allows implementing all the fulfillment of all the tasks, assigned to the EEC of the PCU SC. [ABSTRACT FROM AUTHOR]

This paper considers the idea of reducing the debris of near-Earth space by releasing the spent parts of a spacecraft on orbits touching the conditional boundary of the Earth's atmosphere. We optimize the spacecraft transfer trajectory from a circular reference orbit of an artificial Earth satellite to a target elliptical orbit in a modified pulse problem statement. The convergence of Newton's method is improved by introducing a series of auxiliary coordinate systems at each point of applying an impulse action. The derivatives in the transversality conditions are calculated using a special numerical-analytical differentiation technique. [ABSTRACT FROM AUTHOR]