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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]

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 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]

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]

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]

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]

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]

The idea of creating small spacecraft and constellations based on them, placed in low and medium earth orbits, has been formulated and widely used in the field of space systems development over the last two decades. To date, there are a large number of small satellites designed for remote sensing of the Earth. This paper describes the purpose and main technical characteristics of the small earth remote sensing satellite "AIST-2D". It also describes the progress of the flight tests and the results obtained with the few payloads. In addition, examples of the use of the received Earth remote sensing data to solve various problems are presented. Finally, conclusions are drawn on the interim results of the spacecraft's operation and the prospects for the development of the "AIST-2" platform are described. The small spacecraft "AIST-2D" created in the shortest possible time (less than three years) is an example of successful implementation of a project to create a modern multifunctional spacecraft for remote sensing of the Earth. The experience gained in the development, construction, experimental testing and operation of the "AIST-2D" spacecraft opens up opportunities for the creation of a series of spacecraft based on the "AIST-2" platform with minimal costs and short production times, which will allow us to count on the creation of highly efficient space monitoring systems in the foreseeable future. [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]

• 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]

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]

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 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]

• 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]

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]

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]

The article presents the results of a study of perforated plates effectiveness to ensure safety in case of emergencies on spacecraft and launch pads. The degree of shock wave attenuation by perforated plates with different permeability coefficients has been studied experimentally and numerically. Experiments were carried out with conical shock tube, which was used to produce spherical shock waves of varying intensity, and perforated plate with a permeability coefficient of 41% and. Shock wave reflection and attenuation coefficients were obtained. To study the qualitative picture of a shock wave interaction with perforated plate, a numerical model using Ansys Fluent was created and verified. The numerical model made it possible to qualitatively evaluate the breakup and coupling of a shock wave front after passing perforated plate. The degree of a shock wave attenuation by plates with permeability coefficients of 29% and 20.6% was numerically studied. It was shown that perforated plates can effectively reduce a shock wave overpressure. • Perforated plates effectively reduce shock wave overpressure. • Shock wave attenuation nonlinearly depends on permeability coefficient. • Perforated plates do not change shock wave positive phase duration. • Shock wave breakup and coupling does not depend on characteristic size of hole. [ABSTRACT FROM AUTHOR]

• A novel multi-stage temperature control system was proposed. • Control performance of the proposed system was experimentally verified. • The proposed digital-analog control scheme could precisely maintain the temperature. • Temperature fluctuation of sensitive parts was precisely controlled within ±0.1 ℃. The temperature control precision is crucial for the sensitive elements in a spacecraft. In the present study, a temperature control system to improve the performance of spacecraft precision instruments was designed and experimentally investigated. In the temperature control system, heat pipes, high emissivity coatings, and semiconductor coolers are applied to dissipate the internal heat and control the temperature of a sensitive element. Heat pipes are arranged on the sensitive element and the innermost shell, high radiation coatings are sprayed on each shell surface to transfer the heat to the outer layer, and semiconductor coolers are mounted on the outermost shell and dissipate heat to the external environment. The temperature of the sensitive element is compensated by an electric heating device to reduce the temperature fluctuations. A digital-analog control scheme was designed to maintain the temperature of the outer shell and the sensitive component of the temperature control system. The control accuracy of the temperature control system is within ±0.1 ℃ under different working conditions. When the frequency point is at 0.089928 Hz, the signal amplitude of temperature is taken as the maximum value of 5.05 mK; the amplitude spectral density is 0.006 ∼ 9 mK/Hz1/2 in the frequency range 0.001 ∼ 0.25 Hz. [ABSTRACT FROM AUTHOR]

The long-term changes in the transmission spectra of glass contaminated with oleamide were investigated under vacuum and atmospheric conditions. Oleamide as a contaminant material was deposited onto quartz glass substrates using a vacuum evaporation method. The transmission spectra of the glass contaminated with oleamide showed a decrease in transmittance in the ultraviolet region and transparency in the visible and near-infrared regions. During storage under vacuum, the transmittance in the visible and near-infrared regions decreased gradually over an extended period. This decrease was attributed to the formation of oleamide flakes several microns in size and the concomitant increase in light scattering. The formation of the flakes depended on the thickness of the oleamide, and flake formation was not observed in the case of a thin (0. 5 μ m) oleamide film. This flake formation is attributed to the evaporation of oleamide. The results suggest that the long-term degradation of optical data obtained from spacecraft in orbit should be checked on the ground. • Oleamide was deposited onto the glass in vacuum. • The transmittance of the contaminated glass decreased in the UV region. • The transmittance decreased gradually during storage under vacuum. • This behavior resulted from the formation of micron-sized flakes of oleamide. • This flake formation depended on the thickness of the oleamide. [ABSTRACT FROM AUTHOR]

• An ESO-based controller is proposed for the attitude control of spacecraft with unknown faults. • The event-triggered policy reduces the effects of constants on the final convergence regions. • The controller reduces the frequency of data transmission and actuators updates. An event-triggered control strategy based on extended state observer (ESO) is proposed for the attitude tracking problem of small plug-and-play spacecraft with uncertain inertia parameters, external disturbances, and actuator faults. A simplified controller is developed based on the angular velocity and the general disturbances estimated by the provided ESO using the information of the system inputs and the angular velocities. In the designed event-triggered sampling mechanism, a state-dependent event-triggered strategy determines the triggering instant of the controller to reduce the frequency of information transmission between the controller and the actuator. In comparison with the previous literature, this paper considers uncertain inertia parameters, external disturbances, and actuator faults as general disturbances estimated by ESO, especially for the actuator faults. The inputs of ESO are the error of the angular velocities, which can simplify the controller design. Moreover, the designed ESO can effectively attenuate the influence of measured noises generated by the gyroscopes. The proposed event-triggered policy balances the performance of event-triggering and the control stability performance, which reduces the final state convergence regions without increasing more triggering times compared to existing studies. Furthermore, the investigated policy achieves Zeno-free triggering. Numerical simulations verify theoretical results. [ABSTRACT FROM AUTHOR]

Successful human space exploration requires more products than can be taken as payload. There is a need, therefore, for in-space circular manufacturing. Requirements for this include limited resource inflow, from either Earth or other planets and the generation of minimal waste. The provision of nutritious food is a clear need for human survival on the Moon or Mars and is one of the most complex to solve. Demand in large quantities, constant and reliable provision of food requires the development of specialist agricultural technologies. Here, we first review the history of space farming over the past five decades. This survey assesses the technologies which have been tested under the harsh conditions of space, identifying which modern horticultural components are applicable for in-space plant growth. We then outline which plants have been grown and under what conditions, and speculate upon the types of plants that could be selected to best nourish astronauts. Current systems are focussed on experimentation and exploration, but do not yet provide turn-key solutions for efficient food production within a long-term space exploration scenario. With that take, this review aims to provide a perspective on how an engineered closed circular environmental life-support system (ECCLES) might be constructed. To exemplify the latter, nutrient auto accumulation by biofortification is proposed through the integration of space farming and space mining, which is uncharted on Earth. • Summing up the technological improvement of space horticulture over five decades. • Reporting plants that have been grown in space, how, and the key results obtained. • Identifying potential plants and technologies for future space farming. • Introducing the engineered closed circular environmental life-support system (ECCLES). [ABSTRACT FROM AUTHOR]

In view of the limitation of ground-based Tracking Telemetry and Command (TT&C) system in covering the geostationary satellite in space and time, the method of determining the orbit of the geostationary satellite by the LEO (Low Earth Orbit) multi-satellites network with small orbit inclination was proposed. According to the space environment and optical viewing conditions, the simulation data were screened to simulate the real observation scene. The precise orbit determination (POD) of geostationary satellite was calculated by using the optical angle measurement data and the numerical method. By comparing with the reference orbit, under the condition of platform's orbit accuracy of 5 m, measurement accuracy of 5-arcsecond, and 12 hours of observation, the POD accuracy of geostationary satellite by two LEO satellites can reach the order of kilometers, while the POD accuracy by four LEO satellites can reach the order of 100 meters. Therefore, the POD accuracy has been greatly improved with the increase of the number of LEO satellites. [ABSTRACT FROM AUTHOR]

Aiming at the thermal control challenges of the Chang'E−5 lunar module, which is a complex of lunar lander and lunar ascent vehicle, a design method was proposed based on optimized management of energy. An integrated and autonomous thermal management system for robotic lunar spacecraft was designed including a single-phase fluid loop, water sublimator and high temperature thermal protection shield. The single-phase fluid loop driven by a miniature pump served as the thermal bus to collect heat. The water sublimator was applied as a supplementary heat sink of panel radiators to dissipate heat load on the lunar surface. The lightweight thermal protection shield was applied to protect the equipment in confined space from high temperature plume of big-powered engine during powered descent and ascending. The combination heat sink of fixed panel radiators and water sublimator was managed to meet different heat dissipation requirements of the Chang'E−5 lunar module. The flight data showed that the proposed thermal control system achieved excellent performance. The electronic equipment onboard the Chang'E−5 lunar module was kept between −7.7 °C to 39.6 °C during the mission profile. An optimized goal was achieved that there was zero thermal power compensation for the electronic equipment, while the ratio of thermal load peak-to-trough was as high as 7:1. At the same time, an inter-layer heat transfer model of the heat shield was verified with a temperature difference of less than 3 °C between the prediction result and flight data. • Thermal control challenges of the Chang'E−5 lunar module were described. • Thermal management system for robotic lunar spacecraft was designed. • Thermal control system was designed based on optimized management of energy. • The thermal management system was verified by the operation on the lunar surface. [ABSTRACT FROM AUTHOR]

This paper proposes a new angular velocity observer for attitude tracking of spacecraft based on contraction analysis. The observer is designed in the inertial reference frame via estimating the inertial angular momentum to avoid the square term of the angular velocity when the spacecraft dynamics is expressed in the body frame. It employs a continuous angular-velocity dependent innovation term generated by means of a simple first-order linear filter, instead of a discontinuous attitude-dependent innovation term commonly used in angular velocity observer designs, resulting in a smooth behavior. The global exponential convergence is achieved. Moreover, when combined with the exponentially convergent tracking controller devised in this paper, it gives an overall system with exponential stability relying on a separation property. Finally, a switching function with hysteresis is introduced to stabilize the closest equilibrium in the configuration space, achieving the global exponential stability. Numerical simulations are included to illustrate the performance of the proposed observer in the closed loop, comparison with similar results, and robustness verification under inertia parameter uncertainties and noisy measurements. • A novel angular velocity observer with separation property is designed. • Global exponential convergence is shown for the proposed observer, relaying on contraction analysis. • Two attitude tracking controllers are developed in a controller-observer scheme, resulting in an overall convergent system in view of the contraction of hierarchical systems. • The separation property and robustness of the proposed controller-observer schemes are analyzed and illustrated through numerical simulations. [ABSTRACT FROM AUTHOR]

This paper is concerned with the agile rest-to-rest attitude maneuvering of spacecraft using pyramid-type single-gimbal control moment gyroscopes (SGCMGs). In this method, the gimbal rate command to the SGCMGs is determined by calculating the pair of optimal trajectories of the attitude and the gimbal angles without any "steering law". Thus, the proposed method does not suffer from the singularity problem causing the increase in control error and settling time. Before maneuvering, the pair of the initial optimal trajectories of the attitude and the gimbal angles is calculated by the minimum-time problem. Then, to compensate for errors in maneuvering caused by disturbances and/or modeling errors, the trajectories of the attitude and the gimbal angles are recalculated and updated frequently during the control. In this paper, optimization problems for generating and updating the trajectories are formulated, and specific control algorithms are provided. Finally, numerical simulations and ground experiments with an air floating table are performed to verify the feasibility and effectiveness of the proposed control method. For the rest-to-rest attitude maneuver, the proposed method can be adapted not only to the presence of modeling error in the inertia tensor, but also to the failure of one or two CMGs in the SGCMG system, without changing the control algorithm. These demonstrate the high robustness and practicality of the proposed method. [ABSTRACT FROM AUTHOR]

Soft landers are a common systems for planetary exploration and have been successfully landed on Moon, Mars and the comet Churyumov–Gerasimenko. In this study numerical simulations of the touchdown process of soft landers on the moon surface are presented. The focus is on the soil footpad interaction and the resulting loading on the primary strut of the landing gear. Main influence parameters like slope inclination, touchdown velocities and soil properties are varied to show each individual influence on the structural loading of the landing gear. The numerical simulations are done using the CEL-Method allowing for significant soil deformation at touchdown. [ABSTRACT FROM AUTHOR]

In space magnetic exploration missions, the measurements obtained from the magnetic sensors mounted on a spacecraft are often influenced by the spacecraft's own magnetic interference. Therefore, to ensure precise magnetic exploration, it is imperative to conduct a thorough magnetic characterization of the spacecraft. Conducting the magnetic characterization of spacecraft within a magnetic shielded room (MSR) can mitigate the influence of environmental noise on the characterization process, thereby minimizing uncertainty and enhancing modeling accuracy. However, the utilization of high-permeability materials (such as permalloy) in the MSR introduces a disparity in magnetic characteristics between the spacecraft inside and outside the MSR. According to the method of images, the magnetic field generated by the spacecraft inside the MSR is the superposition of the magnetic fields produced by the spacecraft and that produced by several images of itself. Therefore, this may lead to an ambiguity problem, where spacecrafts with different magnetic characteristics and their corresponding images can generate identical magnetic field distribution in the MSR. To address the aforementioned issues, this paper proposes a novel method for magnetic characterization that integrates both the magnetic field vector data within the MSR and the magnetic field gradient data outside the MSR generated by the spacecraft, allowing for the elimination of the ambiguity problem without introducing environmental magnetic noise interference. As a result, it obtains a unique model for precisely representing the spacecraft's magnetic characteristics in terms of the spherical harmonic coefficients. [ABSTRACT FROM AUTHOR]

An overview of the status and prospects of the search for life in the universe is presented. The search for life beyond the Earth is being carried out using a variety of techniques in three distinct realms of space. First is the study of habitable locations in the Solar System using spacecraft and robotic probes. Second is the detection of the global alteration of exoplanet atmospheres by metabolic processes of microbes, with targets typically Earth-like terrestrial planets within 20 or 30 parsecs. Third is a search for the signatures of technology from extraterrestrial civilizations, in the form of pulsed radio and optical signals, thermal footprints, or physical artifacts. Targets for this third search span a significant fraction of the galaxy. Rough estimates are made of the odds that each of these strategies will be successful and the timescale for that outcome. All the estimates assume that life will form given suitable conditions and chemical ingredients. The best near-term prospect for the detection of life is the detection of biomarkers in exoplanet atmospheres using the James Webb Space Telescope and a set of large, ground-based telescopes that will come online later this decade. Optimal chances of detection center on super-Earths orbiting red dwarf stars. The next best prospect will come from the search for relic traces of life in ancient Mars rocks returned to the Earth, early next decade. Subsequently, missions to several potentially habitable locations in the outer Solar System could detect biomarkers. The most exciting of these involves Titan, where biology would likely have a different biochemical basis than life on Earth. The strategy deemed least likely to succeed, having indeterminate odds of success, is the search for extraterrestrial intelligence, SETI, or more accurately referred to as the search for extraterrestrial technology. The technical capabilities to detect artificial radio and optical pulses has been increasing exponentially, but the parameter space of the search is enormous, and is it impossible to interpret the meaning of a null result. While SETI could succeed at any time, it's likely that life will first be detected on a relatively nearby exoplanet, using spectroscopic methods allied to high resolution imaging. The discovery, when it comes, will be one of the most important in scientific history. • The search for life beyond the Earth is being carried out using a variety of techniques in three distinct realms of space. • The best near-term prospect for the detection of life is the detection of biomarkers in exoplanet atmospheres. • The next best prospect will come from the search for relic traces of life in ancient Mars rocks returned to the Earth. • The strategy least likely to succeed, having indeterminate odds of success, is the search for extraterrestrial intelligence. • The discovery of life beyond Earth, when it comes, will be one of the most important in scientific history. [ABSTRACT FROM AUTHOR]

Remote sensing satellites require very accurate pointing to specific locations of interest with high resolution and small latency. Therefore, the space imaging systems attempted to achieve the possible highest ground resolution with minimum cost as possible. In this paper, design process of an electro-optical imaging payload is described considering the limited weight and small volume in addition to minimum power consumptions. The most promising prospects for high resolution imaging are connected with passive optical sub-systems, especially push broom one. For design a multiband Space remote sensing payload with 4°7′ field of view that operates in multi spectral bands and a panchromatic band, it requires the high speed Schmidt–cassegrain objective. The most promising prospects for high resolution imaging are connected with passive optical sub-systems, especially push broom one. The Schmidt-Cassegrain objective is a catadioptric objective that combines a Cassegrain reflector's optical path with a Schmidt corrector plate to make a compact objective that uses simple spherical surfaces only. Because of the limited weight of the camera (22.2 kg), the multispectral consists of only one single optical system large F-Number (F5) and provided with spectrum splitter. In addition, the key points in design of such types of cameras were listed as well as clarifying the optical design of such objective. [ABSTRACT FROM AUTHOR]

Studying solid combustion phenomena in microgravity environments can be complex, and this is furthered by many limitations and constraints in the available microgravity research platforms. Consequently, fire safety in spacecraft is also a complex subject. The main limitations found in the field are related to the microgravity quality, the duration of microgravity conditions, the rig capabilities in volume and size, time scales, length scales and the diagnostic systems, and these are therefore the focus in the current investigation. The laboratory capacity of ground-based platforms has remained somewhat stalled since 1990s, some drop towers have recently been upgraded to extend their performance. New space-based platforms have been or are being established and could extend the windows-of-opportunity to perform research. In addition, a discussion is provided on the implications of the fact that the phenomena studied in the experimental investigations and the type of material employed covers both programmatic and scientific needs. It is found that a handful of materials are most widely studied to quantify and characterise some of the phenomena, while some materials have been employed even in single experimental efforts. The current literature review provides a very comprehensive overview of previous experimental studies and the experimental methodologies utilised. Thus, this study can become an aid to planning for future studies. • The ground-based laboratory capacity for microgravity research has remained somewhat stalled since the 1990s. • New space-based laboratory capacity is extending the windows-of-opportunity for research. • The major constraints behind every research platform are experimental duration, sample scale and microgravity quality and levels. • Previous studies have investigated a range of diverse materials to comply with programmatic and scientific needs. • Diagnostic systems have improved the quantification of the phenomena behind solid combustion in microgravity. [ABSTRACT FROM AUTHOR]

This paper presents a Model Predictive Control (MPC) design for the test mass release phase of the LISA space mission. LISA is a gravitational wave observatory consisting of a triangular constellation of three spacecraft. The gravitational waves are detected by measuring the relative distance between free falling test masses by means of a laser interferometer. Each test mass is a cubic body located inside an electrostatic suspension that is initially locked by a clamp mechanism. Once the plungers are retracted, the test masses are released with high initial offsets and velocities. To detect the gravitational waves, each test mass must be accurately positioned at the cage centre and its attitude must be aligned with the local cage frame. However, the low actuation authority of the electrostatic suspension along with the critical initial conditions, make the attitude and translation control a difficult task. MPC is a suitable technique for this application because it can systematically account for command saturations, state constraints and can provide optimal (or sub-optimal) control inputs by solving an optimization problem online. In this paper, an MPC controller is designed and validated by means of Monte Carlo simulations, achieving satisfactory results. • Model Predictive Control is a suitable method for the test mass release control problem of the LISA mission. • Model Predictive Control can deal with the low actuation authority of the electrostatic suspensions. • Model Predictive Control can deal with the state constraints of the test mass inside the cage. • Model Predictive Control can manage the operating mode transitions from Wide Range to High Resolution. [ABSTRACT FROM AUTHOR]

The spacecraft in geosynchronous orbit interacts with the space plasma environment, which is prone to electrostatic discharge and affects the safe operation of the spacecraft. In order to realize simple, accurate and reliable electrostatic discharge detection in space environment, based on the energy spectrum characteristics of spacecraft electrostatic discharge electromagnetic radiation electromagnetic pulse and the electromagnetic pulse sensing principle of interdigital electrode, a centimeter-level two-dimensional planar ultra-small interdigital sensor with a size of only 4 cm × 3 cm × 0.01 cm is developed. The impedance matching circuit is established to greatly improve the transmission efficiency of electromagnetic pulse, which is 59.34 % higher than that before the matching circuit is added. Finally, a spacecraft vacuum electrostatic discharge simulation experiment platform is built to test the frequency response characteristics of the sensor. The results show that the frequency response characteristics of the electromagnetic pulse signal perceived by the sensor developed in this paper are consistent with the main frequency characteristics of the spacecraft electrostatic discharge radiation electromagnetic pulse energy spectrum, which can be used for spacecraft electrostatic discharge detection in space environment. [Display omitted] • The sensor is designed based on the ESD characteristics of spacecraft and the sensing principle of interdigital electrode. • By designing the impedance matching circuit, the transmission efficiency of the sensor is increased by 59.34 %. • A centimeter-level 2D ultra-small interdigital sensor with a size of only 4 cm × 3 cm × 0.01 cm was designed and developed. • The experimental results show that the developed sensor can be used for ESD detection of spacecraft in space environment. [ABSTRACT FROM AUTHOR]

The abnormal state of supporting bearings significantly affects the directional accuracy of spacecraft flywheel systems. A self-sensing triboelectric nanogenerator (TENG) offers a desirable route for on-orbit health monitoring and can potentially improve the intelligence level of spacecraft. Here, an extremely compact and lightweight TENG (CL-TENG) for the nonguided clearance of bearings is proposed and fabricated for the condition monitoring of a flywheel assembly. Using the bearing radial space allows the CL-TENG to perceive the revolution and whirling behavior of the cage; therefore, the output of the CL-TENG is simultaneously affected by the bearing speed and load. A test platform for capturing the kinematic of the cage is established to validate the effectiveness of the CL-TENG based on bearing skidding, cage rotation, and whirling instability. A negative exponential correlation between the output voltage and dynamic whirling clearance index is shown to be the basis of cage whirling sensing. The internal clearance and materials of the CL-TENG are optimized to enhance the output performance while preventing wear on the flexible electrodes. The application of the CL-TENG to an actual flywheel system in a simulated space environment and nonstationary operating conditions demonstrate its advantages in detecting the abnormal operating state of the cage. [Display omitted] • Extremely compact and lightweight TENG for bearing on spacecraft flywheel is proposed. • The best innovation of CL-TENG is the self-sensing of bearing cage whirling. • The CL-TENG ensures the integrity and durability of flywheel bearing. • The CL-TENG exhibits excellent performance under non-stationary working conditions. [ABSTRACT FROM AUTHOR]

Due to the high-heat flux issue caused by the increasing power consumption of electronic devices in spacecraft, there is an urgent need but still a significant challenge to develop high thermal conductivity elastomeric thermal interface materials (TIMs) for spacecraft. In this study, mesophase pitch-based carbon fibers (CFs) are added as fillers on the basis of conventional alumina/silicone rubber TIMs, and the CFs are oriented in the matrix by a strong magnetic field during the preparation process. Attributed to the phonon-transport highways composed of vertically-oriented CFs and the synergistic improvement based on CFs and alumina, the through-plane thermal conductivity of the final elastomeric composite with 20 vol% CFs is increased by nearly 14 times compared with the alumina/silicone rubber composite, and it also confirms to the vacuum outgassing performance standard for spacecraft. Moreover, with the help of a verification system, the actual heat transfer capability of the composite is evaluated through the ground test and on-orbit test, and the results suggest that it has a perfect interface heat transfer capability both on the ground and in space, which reveals that this composite can potentially be applied as elastomeric TIM for spacecraft thermal control. [Display omitted] [ABSTRACT FROM AUTHOR]

• A method to predict the FEM capability to represent the dynamic characteristics. • Method is well suited for base excitation problems such as a spacecraft. • Correlation is demonstrated using flightworthy spacecraft structure. • Correlation can be done as part of base excited qualification test. The accuracy of the finite element model is an important aspect, especially in the aerospace domain, where design margins are barely minimum. Hence, standard test-analysis model correlation methods apply to assess the quality of the analytical model. However, it is observed that these standard correlation techniques are inadequate to assure the model capability to predict the forced response characteristics of complex structures. The base force assurance criterion is found to be a good tool to overcome this inadequacy. During the dynamic test of the structure, the force measurement device can be used to measure the transmitted force to the base. Thus measured base forces, and the moments computed from the measured forces are vector correlated with the analytical force and moments to predict qualitatively the analytical model capability to represent the design critical peak acceleration responses and the global modes. The procedure is demonstrated using a set of experiments conducted on a real spacecraft under the sine excitations in two mutually perpendicular directions. The sensitivity of the base force assurance criterion to the inaccuracies that may be present in the finite element model and/or in the experimental results is evaluated using simplified Monte Carlo simulations. [ABSTRACT FROM AUTHOR]

Relying on contraction analysis, this paper addresses the global attitude tracking problem of a spacecraft when angular velocity measurements are corrupted by bias. A nonlinear observer with exponential convergence is designed firstly to estimate the bias in gyro sensors. Then an exponentially convergent attitude tracking controller with gyro bias correction is devised. Next, to remove the topological constraints of unit quaternions for global stability, a switching variable with hysteresis is incorporated in the control loop, enhancing the robustness in the presence of measurement noise and energy efficiency by preventing the unwinding phenomenon. Numeric simulations are shown to illustrate the performance and compare with other similar controllers in terms of tracking error, estimation error and energy efficiency, as well as the robustness to noisy measurements and time-varying bias in gyro sensors. • A nonlinear gyro bias observer with global exponential convergence is designed based on quaternion kinematics. • An attitude tracking controller with global exponential convergence is devised. • A global exponential unwinding-free version of the previous controller is developed. • Comparison of the proposed controller with three representative controllers reported in the literature is carried out through numerical simulations. [ABSTRACT FROM AUTHOR]

This paper is concerned with the transfer trajectory design among multiple planets using multiple flybys. The proposed trajectory design method consists of a semi-analytical tuning strategy and iterative patched-conic method. The tuning strategy refers to a method realizing the smooth connection between a hyperbolic trajectory near a planet and a trajectory obtained by Lambert's problem between the planet and another one. The proposed method iteratively solves Lambert's problem and the tuning strategy so that the arrival or departure velocity obtained by Lambert's problem coincides with the velocity at the corresponding point in the sphere of influence (SOI), assigned by the tuning strategy. Consequently, a smooth and continuous trajectory with multiple flybys can be obtained, where the SOI of each flyby planet and the flyby time are explicitly taken into account. Moreover, analytical conditions with respect to the orbital elements of the arrival and departure hyperbolas are provided to minimize velocity mismatch at the periapsis of a flyby planet. Through numerical simulation using the Cassini example, it is shown that the resultant trajectory by the proposed method is close to the optimal trajectory obtained by the n -body problem. Thus, a suitable approximate trajectory can be effectively generated without solving the n -body problem. • Interplanetary trajectories with multiple flybys are designed. • The proposed method consists of analytical tuning strategy and iterative patched-conic. • The SOI and flyby time for each flyby planet are explicitly taken into account. • Analytical conditions of the orbital elements are provided to minimize velocity mismatch. • A suitable approximate trajectory is effectively generated without solving the N-body problem. [ABSTRACT FROM AUTHOR]

The problem of attitude takeover control of spacecraft by using cellular satellites with limited communication, actuator faults and input saturation is investigated. In order to lighten the communication burden of cellular satellites, an event-triggered control strategy is adopted. The filtered attitude information needs to be transmitted only when the defined measurement error reaches the event-triggered threshold in this strategy. Then, to deal with the unknown inertia matrix, actuator faults, external disturbances and the errors caused by event-triggered scheme, fuzzy logic systems is introduced to estimate the uncertainties directly. Combining fuzzy logic control strategy and the event-triggered method, the first event-triggered adaptive fuzzy control law is developed. Then, torque saturation of cellular satellites is further considered in the second control law, where the upper bound of the uncertainties is estimated by fuzzy logic systems. The resulting closed-loop systems under the two control laws are guaranteed to be bounded. Finally, the effectiveness of two proposed control laws is verified by the numerical simulations. [ABSTRACT FROM AUTHOR]

Highly porous aluminum alloys and honeycomb structures may be utilized as single-use shock absorbers in landing gear and other elements of spacecraft thus ensuring safety of the landing process. This paper presents the results of studies that make it possible to estimate mechanical properties and shock-absorbing characteristics of highly porous aluminum alloys and honeycomb aluminum alloy structures, the initial average densities of said alloys and structures ranging from 330 to 2250 kg/m³. We list analogs that correspond with different stages found on stress-strain curves of highly porous aluminum alloys and honeycomb structures. The paper also provides the foundations of calculating the specific and effective energies of shock absorption. • The deformation diagrams of cellular aluminum are close to the idealized one. • Specific and effective impact absorption energy calculated by the deformation diagram. • Some honeycomb structures have a deformation diagram with hardening. • The presence of hardening worsens the shock-absorbing characteristics. • The deformation diagram depends on the porosity, matrix material and pore structure. [ABSTRACT FROM AUTHOR]

• First review of vapor compression refrigeration for space applications. • Summarizing more than 40 years of development. • Covering gas-liquid two-phase flow, thermal gravitational scaling and system design considerations. • Suggesting a general evaluation criterion to select cooling technologies for spacecraft. Using a vapor compression cycle for cooling in microgravity environments was already suggested in the 1970s to leverage the high coefficient of performance. Since then, only a few systems have operated in microgravity with scarce documentation of these flights. The lack of measured data and detailed documentation makes identifying the necessary next steps difficult for researchers entering the field of refrigeration in space. This paper provides a review of available literature for vapor compression systems flown in microgravity by outlining the history of vapor compression devices in space and presenting performance data. Moreover, gaps in the literature are highlighted and open questions are posed based on the reviewed material. Next steps of research are suggested to support and ultimately achieve reliable vapor compression refrigeration in space. Calculating equivalent masses for a fair comparison of different microgravity cooling technologies is proposed by capturing both energy consumption and used volume. [ABSTRACT FROM AUTHOR]

This paper presents a neural-aided Guidance, Navigation & Control algorithm for reconfiguration of distributed space systems. The guidance algorithm is based on Artificial Potential Fields (APF) in the Relative Orbital Elements (ROE) space. Since the relative orbit determination measurements are typically referred to the Cartesian metrics (e.g. range or range rate), a linear mapping between the set of ROE and the Cartesian coordinates expressed in the Local-Vertical-Local-Horizontal (LVLH) reference frame is derived. The navigation and control algorithms rely on the relative dynamics expressed in the same ROE set of coordinates. To cope with uncertainties and nonlinearities of the system, a Radial Basis Function Neural Network (RBFNN) is employed to reconstruct the perturbed dynamics. The Artificial Neural Network (ANN) is coupled with an adaptive Extended Kalman Filter for state estimation. A feedback control is designed to track the desired state, whose stability is analyzed using Lyapunov theory. The guidance, navigation and control algorithms are tested in a high-fidelity numerical orbit propagator. Moreover, the algorithm is tested in relevant Processor-In-the-Loop (PIL) simulations using a TI C2000-Delfino MCU F28379D. The results demonstrate the effectiveness of the algorithm for relative reconfiguration maneuvers involving relative distances ~ 10 2 m with limited fuel consumption and constrained available thrust (⩽ 1 mN). In particular along-track maneuvers, relative plane change and formation enlargement are analysed in the paper, showing the comparison between the proposed algorithm and the legacy one without the neural network. The benefit of implementing a neural network is particularly highlighted when the nonlinearities or unmodelled terms in the on-board dynamics become prominent. [ABSTRACT FROM AUTHOR]