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

This paper analyzes the design of an alpha-decay sail. While previous studies have been conducted on the feasibility of alpha-decay sails, none have investigated the effects of different backing materials and backing material thicknesses on propulsion. This paper conducts simulations using the particle simulation toolkit Geant4 to determine the thrust produced from alpha-decay sails constructed of 238Pu and a backing layer made of aluminum, beryllium, lithium, or polyethylene. The results demonstrate that alpha-decay sails can be a very useful form of in-space propulsion, having the potential to dramatically decrease travel time beyond distances of 100 AU. [ABSTRACT FROM AUTHOR]

With the increasing frequency of human space activities, the orbital space environment is deteriorating. It is of great practical significance to enhance the strength and toughness of spacecraft structures. High strength & toughness bio-inspired helicoidal composite metastructures with mid-plane symmetry characteristics were designed and a corresponding hot-pressing preparation process was developed. The carbon fibre reinforced polymer metastructures with cross-ply, quasi-isotropic and 5°, 10° and 20° helicoidal lay-ups were characterized by quasi-static indentation performance tests, and the damage modes and the failure mechanisms were analyzed. The load-displacement curves, peak forces, failure displacements, stiffnesses and energy absorptions, were used as the mechanical property measures, with the thicknesses of the structures including 37 and 73 layers. The results show that, compared with the traditional lay-up method, the symmetric helicoidal lay-up can effectively reduce the interlayer stress and significantly improve the quasi-static indentation performances of the metastructures. Especially with a helicoidal angle of 10°, the metastructures have excellent performance enhancement in terms of peak loads and energy absorptions. The research results not only provide a theoretical support for the design and fabrication of high-performance composite metastructures in the aerospace field, but also lay a practical foundation for their practical application. [ABSTRACT FROM AUTHOR]

The results of creating a map of the circumpolar regions of the Moon are described. The maps are limited to latitudes +/–55° to show the location of the Luna-25 landing. The article also provides a brief overview of space missions to the Moon. [ABSTRACT FROM AUTHOR]

During eclipse periods, the spacecraft relies on electricity, which its solar arrays produce and store in batteries. Forecasting a solar array's power output employed during space missions is of significant importance. Varying space environments and satellite loads, which are characterized by significant randomness and uncertainty, affect the generated power of the spacecraft's solar array. These challenges pose difficulties in power prediction. To address these issues and achieve a more accurate estimation of the solar array's generated power during a space mission, this study develops a metabolic model termed TDGM(1, 1, r) that incorporates an enhanced accumulating fractional-order, optimizing it within the discrete grey TDGM(1,1) model's framework with three parameters. The optimization model's objective function is defined as the mean absolute percentage error (MAPE) within the modeling context. In order to minimize MAPE, the differential equation's order and accumulation number are determined using a particle swarm approach. The TDGM(1, 1, r) demonstrates superior forecasting performance in comparison to the classical GM(1,1) and Grey-Markov models. These findings indi-cate the superiority of TDGM(1,1,r) over GM(1,1) and Grey-Markov, with improvements of 84.2% and 81.2% for MAPE (from 1.83% to 0.29% and from 1.54% to 0.29%). The metabolic TDGM(1,1,r) employing the particle swarm algorithm (PSO) is better suited for short-term predictions. Finally, relevant suggestions for future development of the prediction model are proposed. [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]

This paper investigates the attitude control of rigid spacecraft in the presence of uncertainties, disturbances, and actuator faults. In order to effectively address these challenges and improve the performance of the system, a novel actor-critic neural-network-based fractional-order sliding mode control (ACNNFOSMC) has been developed for spacecraft. The integration of actor-critic neural network, fractional-order theory, and sliding mode control enables dual functionality: the actor-critic neural network serves to approximate the aggregate of uncertain parameters, disturbances, and actuator faults, thereby facilitating their compensation, while the fractional-order sliding mode control mechanism significantly improves the system's tracking precision and overall robustness against uncertainties. Theoretical analyses are presented to analyze the stability of the proposed control framework. Thorough examination via simulation experiments affirms the effectiveness and control precision of attitude of our proposed control strategy, even in complex operational scenarios. [ABSTRACT FROM AUTHOR]

The development of the Drag-Free and Attitude Control System is crucial for space-based gravitational wave detectors. The Drag-Free and Attitude Control System includes for spacecraft attitude control, drag-free control, moving optical subassemblies pointing control, and test masses electrostatic control during multiple phases of the mission lifetime. An accurate model of the spacecraft dynamics model is not only a precondition for the control design of Drag-Free and Attitude Control System, but also can serve as a valuable reference to assess the noise budget of gravitational wave detectors. This paper presents a generic, nonlinear mathematical model that describes the multi-body dynamics for the gravitational wave detectors. The model fully considers the couplings between the bodies involved, with a particular focus on the impact of the moving optical subassemblies on the spacecraft attitude. Additionally, a corresponding reduced model for the control design of the science mode is divided. Both models are validated numerically by Monte Carlo simulations and motion analysis, respectively. • Introduce a spacecraft dynamics model that fully describes the motion coupling. • Simscape simulation has validated the accuracy of the dynamics model. • The spacecraft dynamics model is applicable to multiple phases of motion. • Monte Carlo simulations have validated the applicability to different parameters. [ABSTRACT FROM AUTHOR]

Hydrazine's carcinogenic and toxic nature makes its use in space propulsion expensive and necessitates careful launch campaign planning, which may cause delays. One already established alternative is to use energetic salt-based ionic solution propellants. However, the costs remain high because these salts are costly and highly regulated energetic materials. Additionally, propellants based on these energetic salts produce much higher temperatures during decomposition and combustion, as is the case for hydrazine. Consequently, expensive refractory metal-based alloys and high-temperature oxidation-resistant catalytic reactors must be used, which further drives up the costs. Alternatively, hydrogen peroxide may be used, though only in applications with a low specific impulse requirement. This study explores an alternative approach: using nitromethane, a commonly available laboratory solvent. NMP-001 is a nitromethane-based monopropellant that contains additives that allow for low-pressure combustion and reduce sensitivity to mechanical stimuli. Multiple combustion chamber configurations with varying characteristic lengths were tested in rocket combustion experiments. With the 11.5 m L ∗ -configuration, a 15.1 bar low-pressure limit could be determined. The 7.3 and 5.0 m L ∗ -configurations yielded a 16.2 and 28.3 bar low-pressure combustion limits, respectively. These outcomes are significantly below the 30 bar limit reported in the literature. Despite using only a simple heat-sink combustion chamber with a non-sophisticated injector element and a non-optimized propellant formulation, combustion was 83 to 95 % efficient with a C* ranged between 1133 to 1302 m/s. These figures are higher than for hydrogen peroxide and are comparable to the theoretical performance of state-of-the-art green liquid monopropellants. • Existing alternative ionic propellants are costly and highly regulated due to their instability and energetic properties. • Nitromethane offers a cost-effective solution with comparable performance, leading to the development of NMP-001, a nitromethane-based monopropellant with reduced sensitivity to mechanical stimuli. • Tests showed a lower low-pressure combustion limit of 15 ba. This is significantly lower than the limits reported in the literature. • The measured C*-performance for NMP-001 is better than hydrazine and comparable to LMP-103S. [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]

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]

The heat dissipation capacity prediction and cooling adjustment of the space station are rarely reported and are currently very urgent. Therefore, this paper does the following work. First, a twin system equivalent to the actual thermal system is built based on the digital twin technology, and the net heat dissipation capacity that can reflect the energy storage characteristics of the radiator and the total heat that can reflect the load changes in the cabin are identified. Then, the optimum operating state of the thermal system is planned by using the changing characteristics of the two, to predict the maximum heat dissipation capacity and heat dissipation margin of the space station. Finally, the cooling adjustment method is given. The results show that the heat dissipation capacity of the shadow area of the space station is about 16kW, which is about 1.5 times the sunlit area. The series control reduces the time difference between hot and cold source from 677 s to about 77 s. The radiator absorptivity is only increased by 0.15, and its accumulated heat dissipation will be reduced by 7.48 kWh in one orbital period. The radiator emissivity decreased by 0.05, and the accumulated heat dissipation decreased by 1.6 kWh. When the solar azimuth is 20° and 160°, the heat dissipation capacity ratio of the whole module in the sunlit zone is about 0.9 of a single orbital period. It is worth mentioning that the idea of heat dissipation capacity prediction and cooling adjustment proposed in this paper has important reference significance for both fluid network and non-fluid network structures. [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]

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]

Ground experimentation is an important method to verify the performance and quality of spacecraft structures. How to simulate the force and heat fields that a spacecraft may be subjected to during flight is a challenge for experiment design. In this paper, the static strength experimental method of spacecraft in the combined force and heat fields is studied. Furthermore, the coupling cases of the spacecraft during the flight process are simulated using an airbag and a resistive heater. Finally, it can be concluded that the accuracy of the proposed experimental method as well as the structural strength of the spacecraft can meet technical requirements during flight. [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]

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]

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]

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]

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]

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]

Space debris is an urgent problem of our time. This paper considers the idea of reducing near-Earth space debris by releasing the spent additional fuel tank (AFT) and the booster's central block (CB) into the Earth's atmosphere. The spacecraft transfer from a reference circular orbit of an artificial Earth satellite to a target elliptical orbit is optimized. The transition maneuvers are carried out using a booster with a high limited-thrust engine and the AFT. The second zonal harmonic of the Earth's gravitational field is taken into account. The optimal control problem is solved based on Pontryagin's maximum principle. Bulky derivatives are calculated using a specially developed numerical-analytical differentiation technique. The Pontryagin extremals obtained below are the next step in implementing the problem hierarchy methodology. [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]

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]

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]

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]

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]

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]

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]

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]

• 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]

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]

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]

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]

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]