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The pervasiveness of deep space radiation remains a confounding factor for the transit of humans through our solar system. Spacecraft shielding both protects astronauts but also contributes to absorbed dose through galactic cosmic ray interactions that produce secondary particles. The resultant biological effects drop to a minimum for aluminum shielding around 20 g/cm2 but increase with additional shielding. The present work evaluates for the first time, the impact of secondary pions on central nervous system functionality. The fractional pion dose emanating from thicker shielded spacecraft regions could contribute up to 10% of the total absorbed radiation dose. New results from the Paul Scherrer Institute have revealed that low dose exposures to 150 MeV positive and negative pions, akin to a Mars mission, result in significant, long-lasting cognitive impairments. These surprising findings emphasize the need to carefully evaluate shielding configurations to optimize safe exposure limits for astronauts during deep space travel.

Members of the family Trichomeriaceae, belonging to the Chaetothyriales order and the Ascomycota phylum, are known for their capability to inhabit hostile environments characterized by extreme temperatures, oligotrophic conditions, drought, or presence of toxic compounds. The genus Knufia encompasses many polyextremophilic species. In this report, the genomic and morphological features of the strain FJI-L2-BK-P2 presented, which was isolated from the Mars 2020 mission spacecraft assembly facility located at the Jet Propulsion Laboratory in Pasadena, California. The identification is based on sequence alignment for marker genes, multi-locus sequence analysis, and whole genome sequence phylogeny. The morphological features were studied using a diverse range of microscopic techniques (bright field, phase contrast, differential interference contrast and scanning electron microscopy). The phylogenetic marker genes of the strain FJI-L2-BK-P2 exhibited highest similarities with type strain of Knufia obscura (CBS 148926T) that was isolated from the gas tank of a car in Italy. To validate the species identity, whole genomes of both strains (FJI-L2-BK-P2 and CBS 148926T) were sequenced, annotated, and strain FJI-L2-BK-P2 was confirmed as K. obscura. The morphological analysis and description of the genomic characteristics of K. obscura FJI-L2-BK-P2 may contribute to refining the taxonomy of Knufia species. Key morphological features are reported in this K. obscura strain, resembling microsclerotia and chlamydospore-like propagules. These features known to be characteristic features in black fungi which could potentially facilitate their adaptation to harsh environments.

Background: The commercialization of space travel will soon lead to many more people living and working in unique built environments similar to the International Space Station, which is a specialized closed environment that contains its own indoor microbiome. Unintended microbial growth can occur in these environments as in buildings on Earth from elevated moisture, such as from a temporary ventilation system failure. This growth can drive negative health outcomes and degrade building materials. We need a predictive approach for modeling microbial growth in these critical indoor spaces. Results: Here, we demonstrate that even short exposures to varying elevated relative humidity can facilitate rapid microbial growth and microbial community composition changes in dust from spacecraft. We modeled fungal growth in dust from the International Space Station using the time-of-wetness framework with activation and deactivation limited growth occurring at 85% and 100% relative humidity, respectively. Fungal concentrations ranged from an average of 4.4 × 106 spore equivalents per milligram of dust in original dust with no exposure to relative humidity to up to 2.1 × 1010 when exposed to 100% relative humidity for 2 weeks. As relative humidity and time-elevated increased, fungal diversity was significantly reduced for both alpha (Q < 0.05) and beta (R2 = 0.307, P = 0.001) diversity metrics. Bacteria were unable to be modeled using the time-of-wetness framework. However, bacterial communities did change based on constant relative humidity incubations for both beta (R2 = 0.22, P = 0.001) and alpha diversity decreasing with increasing moisture starting at 85% relative humidity (Q < 0.05). Conclusion: Our results demonstrate that moisture conditions can be used to develop and predict changes in fungal growth and composition onboard human-occupied spacecraft. This predictive model can be expanded upon to include other spacecraft environmental factors such as microgravity, elevated carbon dioxide conditions, and radiation exposure. Understanding microbial growth in spacecraft can help better protect astronaut health, fortify spacecraft integrity, and promote planetary protection as human activity increases in low-Earth orbit, the moon, Mars, and beyond. 5L5VMn4841emXVRzhjbC6H Video Abstract [ABSTRACT FROM AUTHOR]

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

This paper investigates the distributed formation control of a group of leader-following spacecraft with bounded actuation and limited communication ranges. In particular, connectivity-preserving and collision-avoidance controllers are proposed for the leader with constant or time-varying velocity, respectively. The communication graph between the spacecraft is modeled via a distance-induced proximity graph. By designing a virtual proxy for each spacecraft, the spacecraft–proxy couplings address the actuator saturation constraints. The inter-proxy dynamics incorporated with a bounded artificial potential function fulfill the coordination of all proxies. In addition, the bounded potential function can simultaneously tackle connectivity preservation and collision avoidance problems. The distributed formation controllers are proposed for multiple spacecraft with constant or time-varying velocities relative to the leader. A sliding mode control approach and the proxies' dynamics are used in the design of a distributed cooperative controller for spacecraft to address the cooperative problem between the followers and the leader. Numerical simulations confirm the effectiveness of the anti-saturation distributed connectivity preservation controller. [ABSTRACT FROM AUTHOR]

The design of the space hatch door mechanisms is crucial in the aerospace field, impacting not only durability and reliability but also the life safety of astronauts during space missions. This review extensively researches vehicle doors and hatches in civil and military systems across various environments, including land, sea, deep sea, aviation, aerospace, and extreme conditions. Specially, it focuses on the structural design of hatches and related mechanisms in civil aviation and military aerospace environments, such as opening and closing mechanisms, release mechanisms, locking mechanisms, sealing mechanisms, and the ergonomic design of door structures. The review highlights the integration of bionic design principles with hatch mechanisms to explore future solutions. By systematically examining these aeras, this review addresses the lack of comprehensive studies in previous reviews, which often overlook the interconnectivity and applicability of hatch mechanisms across different fields. The absence of such holistic reviews has led to fragmented knowledge and missed opportunities for cross-industry innovation. This review aims to fill these gaps by providing a wide range of design solutions and offering insights that can enhance the development of more reliable, efficient, and safe hatch mechanisms in aerospace and other high-stakes environments. [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 cold spray deposition technique has been used to produce a new class of solid lubricant coatings using powder feedstocks of the metal disulfides WS2 or MoS2, either pure or mixed with Cu and Ni metal powders. Friction and cycle lives were obtained using ball-on-flat reciprocating tribometry of coated 304 SS flats in dry nitrogen and vacuum at higher Hertzian contact stresses (Smax = 1386 MPa (201 ksi)). The measured friction and thickness of the coatings were much lower than for previous studies (COF = 0.03 ± 0.01 and ≤1 µm, respectively), which is due to their high metal disulfide:metal ratios. Cu-containing metal sulfide coatings exhibited somewhat higher cycle lifetimes than the pure metal sulfide coatings, even though the Cu content was only ~1 wt%. Profiling of wear tracks for coatings tested to 3000 cycles (i.e., pre-failure) yielded specific wear rates in the range 3–7 × 10−6 mm3N−1m−1, similar to other solid lubricant coatings. When compared to other coating techniques, the cold spray method represents a niche that has heretofore been vacant. In particular, it will be useful in many precision ball-bearing applications that require higher throughput and lower costs than sputter-deposited MoS2-based coatings. [ABSTRACT FROM AUTHOR]

Abstract Background The commercialization of space travel will soon lead to many more people living and working in unique built environments similar to the International Space Station, which is a specialized closed environment that contains its own indoor microbiome. Unintended microbial growth can occur in these environments as in buildings on Earth from elevated moisture, such as from a temporary ventilation system failure. This growth can drive negative health outcomes and degrade building materials. We need a predictive approach for modeling microbial growth in these critical indoor spaces. Results Here, we demonstrate that even short exposures to varying elevated relative humidity can facilitate rapid microbial growth and microbial community composition changes in dust from spacecraft. We modeled fungal growth in dust from the International Space Station using the time-of-wetness framework with activation and deactivation limited growth occurring at 85% and 100% relative humidity, respectively. Fungal concentrations ranged from an average of 4.4 × 106 spore equivalents per milligram of dust in original dust with no exposure to relative humidity to up to 2.1 × 1010 when exposed to 100% relative humidity for 2 weeks. As relative humidity and time-elevated increased, fungal diversity was significantly reduced for both alpha (Q

Due to the rapid development of the space industry, ever higher demands are being made for the optimization and improvement of spacecraft thermal management systems. Thermal control technology has become one of the key bottlenecks that restrict the level of spacecraft design. In this paper, the thermal management technologies (TMTs) for spacecraft electronics are reviewed according to the different heat transfer processes, including heat acquisition, heat transport, and heat rejection. The researches on efficient heat acquisition include the utilization of high thermal conductance materials, the development of novel package structure based on micro-/nano-electromechanical system (MEMS/NEMS) technologies, and advanced near-junction microfluidic cooling techniques. For the heat transport process, various heat pipes and mechanical pumped fluid loops (MPFLs) are widely implemented to transport heat from heat generation components to the ultimate heat sinks. The heat pipes are divided into two categories based on their structure layout, i.e., separated heat pipes and unseparated heat pipes. The merits and demerits of these heat pipes and MPFLs (including the single-phase MPFL and the two-phase MPFL) are discussed and summarized respectively. In terms of the heat rejection for spacecraft, thermal radiators are normally the sole option due to the unique space environment. To meet the requirements of large heat dissipation power and fluctuated thermal environment, research efforts on the radiators mainly focus on the development of deployable radiators, variable emissivity radiators, and the combination with other techniques. Due to the fluctuated characteristics of the heat power of internal electronics and the outer thermal environment, the phase change materials (PCMs) exhibit great advantages in this scenario and have attracted a lot of research attention. This review aims to serve as a reference guide for the development of thermal management system in the future spacecraft.

In the quaternion formulation, the problem of mathematical modeling of the spacecraft motion in an elliptical orbit is considered. The control is a modulo-limited acceleration vector from the jet thrust, directed orthogonally to the plane of the spacecraft orbit. The motion of the spacecraft center of mass is described using the quaternion differential equation of the orientation of the orbital coordinate system. An approximate analytical solution of the quaternion differential equation of the orientation of the orbital coordinate system is constructed in the form of a uniformly suitable asymptotic expansion by degrees of eccentricity of the spacecraft orbit (small parameter). To eliminate the secular terms in this expansion, the renormalization method was applied. Taking into account the known solution of the equation of orientation of the orbital coordinate system for the case when the spacecraft orbit is circular, allowed to simplify the form of the above expansion. The nonlinear oscillation frequencies of each component of the desired quaternion were found. Analytical transformations were performed using the SymPy symbolic algebra package. To carry out numerical simulation of the spacecraft motion, a program was written in Python. Calculations based on analytical formulas obtained in the paper (in the absence of secular terms) and previously obtained results in the presence of secular terms are compared. An example of modeling the controlled motion of a spacecraft is given for the case when the initial orientation of the orbital coordinate system corresponds to the orientation of the orbit of one of the satellites of the GLONASS orbital grouping. Graphs were built to show error in the module (and components) of the quaternion describing the orientation of the orbital coordinate system. It is shown that the elimination of secular terms using the renormalization method made it possible to reduce the error in determining this module with an increase in the number of spacecraft revolutions around the Earth. The analysis of the obtained approximate analytical solution is carried out. The features and regularities of the spacecraft motion in an elliptical orbit are established.

The Advanced Space-based Solar Observatory (ASO-S) marked China’s first comprehensive solar mission in space. Drawing upon the previous reports covering 2018‒2020 and 2020‒2022, we present here an update on the ASO-S made from 2022 to 2024. In August 2022, ASO-S completed its Phase D study and was successfully launched on October 9, 2022. The commissioning phase was carried out and concluded within the first nine months following the launch. The data and associated analysis software have been opened to the community and the research on the early ASO-S data has been well developed. We anticipate also the achievements in data research pertaining to ASO-S in the near future.

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]

The article presents the results of a study on the basis of which a set of mathematical models was developed for ballistic analysis of a mission to send a small spacecraft to the satellite of Jupiter, Callisto and its landing on the surface of the satellite. In this mission, it is proposed to use a gravity assist maneuver around the Earth and an aerodynamic maneuver near Jupiter to reduce the cost of the working fluid of the spacecraft. The minimum required thrust of the engines and the duration of the soft landing maneuver of a spacecraft with a given mass on the satellite are estimated. The optimal launch date was found for the possibility of launching a spacecraft using a medium-lift Soyuz-2 launch vehicle. The simulation of the movement was carried out numerically, in the Mathcad mathematical package, all the dependence diagrams necessary for the analysis of the movement were constructed.

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 introduces a medium voltage direct-current (MVDC) system for large spacecraft megawatt-scale (MW) power and propulsion systems intended for interplanetary transport, including missions to the Moon and Mars. The proposed MVDC system includes: (i) A nuclear electric propulsion (NEP) that powers a permanent magnet (PM) generator whose output is rectified and connected to the MVDC bus. (ii) A solar photovoltaic (PV) source that is interfaced to the MVDC bus using a unidirectional boost DC-DC converter. (iii) A backup battery energy storage system (BESS) that connects to the MVDC bus using a bidirectional DC-DC boost converter. (iv) A dual active bridge (DAB) converter that controls the power to the spacecraft's electric thruster. The NEP serves as the main power source for the spacecraft's electric thruster, while the solar PV and BESS are intended to provide power for the payload and spacecraft's low-voltage power system. The paper will (i) provide a review of the spacecraft MVDC power and prolusion system highlighting state-of-the-art main components, (ii) address the control of boost converters for the PV and BESS sources and the DAB converter for the thruster, and (iii) propose an uncertainty and disturbance estimator (UDE) concept based on current control algorithms to mitigate MVDC instability due to unpredictable factors and external disruptions. The proposed UDE can actively estimate and compensate for the system disturbance and uncertainty in real time, and thus, both the system tracking performance and robustness can be improved. Simulation studies have been conducted to substantiate the efficacy of the proposed schemes. [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]

Presented method allows one to calculate the of maneuvers performed on several turns using a low-thrust engine. These maneuvers ensure the flight of an active spacecraft within a given area of the target space object. The flight is carried out in the vicinity of a circular orbit. Simplified mathematical models of motion are used to solve this problem. The influence of the non-centrality of the gravitational field and atmosphere is not taken into account in the calculations. The process of determining the parameters of the maneuvers is divided into several stages: in the first and third stages, the parameters of the impulse transfer and the transfer carried out by the low-thrust engine are calculated using analytical methods. In the second stage, the distribution of maneuvering between turns, ensuring a successful solution to the meeting problem, is determined by changing one variable. This method is characterized by its simplicity and high reliability in determining the parameters of maneuvers, which makes it applicable on board a spacecraft. As part of the study, an analysis of the dependence of the total characteristic velocity of solving the meeting problem on the amount of engine thrust was also carried out. The maneuver parameters can be refined using an iterative procedure to take into account the main disturbances.

The localization of electromagnetic radiation leakage through cabin gaps is a critical and challenging aspect of electromagnetic compatibility design for spacecraft. This paper proposes a localization method based on synthetic aperture interferometric passive radiometry imaging. Electromagnetic radiation signals are measured at a certain distance from the spacecraft surface to form visibility samples. These visibility samples and the modified brightness temperature of the spacecraft surface electromagnetic radiation leakage constitute a pair for Fourier transform. The electromagnetic leakage location image of spacecraft surface is obtained through inverse Fourier transform. A sparse sampling method based on particle swarm optimization is proposed to improve testing efficiency. The impacts of various factors, including positional parameters, positioning accuracy of the test antenna, and scanning parameters on the imaging results are analyzed. Experiments are conducted on a cabin with 1 m $\times 1$ m $\times 1$ m having 51 holes on one surface. The algorithm proposed in this article is validated to effectively image and locate electromagnetic leakage points at different frequencies, with an absolute positioning error not exceeding 16mm. In addition, the effectiveness of the sparse sampling method is confirmed. Hence, it can accurately locate the position of electromagnetic leakage by saving 83.3% of the test time, with a misjudgment rate of 5.9%.

In order to find a simpler method to simulate the charging of spacecraft surface and further analyze the characteristics of the charging reaction of low-orbit spacecraft surface materials, this paper uses COMSOL to perform finite element analysis on the charging of low-orbit spacecraft surface materials at an altitude of 300–400 km. First, the surface charging potential is established using the particle transport equation. By calculating the thermal diffusion movement of particles, the adsorption, secondary electron emission and backscattering effects of the material surface are evaluated to simulate the charging potential of the spacecraft surface. The results show that COMSOL based on the plasma module can replace SPIS to simulate the charged properties of spacecraft surface materials, indicating that different physical properties of the material surface are attached with different magnitudes of potential and small potential differences; In addition, based on the particle transport equation, this paper obtains the surface charging results of the Kapton connector between the solar panel and the spacecraft body, which is prone to charging. This research content provides a simplified method and new direction for the simulation study of spacecraft surface charging.

This paper proposes a generalized control for multiphase multilevel neutral-point-clamped (MPML-NPC) converter in a large spacecraft electric propulsion (EP) system. The spacecraft EP scheme considered in this paper has been proposed by NASA for long-term crewed missions and includes a heat source that acts as a prime mover to a fixed speed permanent magnet generator (PMG) whose output is rectified via a MPML-NPC converter before connecting to a high-power electric thruster. The proposed generalized control is based on space vector modulation (SVM) developed for MPML-NPC converters with any number of phases and levels. The paper will discuss mathematical modelling of multiphase systems and derive the equations that govern the proposed control. The generalized control generates optimized switching commands based on a space vector diagram (SVD). The analytical SVM models are verified through simulations for 3-phase, 6-phase, and 9-phase MPML-NPC converters. Further, the simulations are validated through test results from a scale-down laboratory prototype MPML-NPC converter.

A linear time-varying model predictive control (LTV-MPC) method based on $SE(3)$ is proposed, to enhance the control precision and reduce energy consumption of the spacecraft. First, left-invariant principle of Lie group, Lie algebra $SO(3)$ , and other differential geometry theories are applied to extend LTV-MPC to $SE(3)$ . Then, considering the obstacle avoidance problem of the spacecraft in space, a suitable optimization function is selected to ensure smooth tracking of the desired working trajectory. By controlling the incremental output with the fastest convergence rate, the controller enables effective trajectory tracking. Finally, the effectiveness and applicability of the controller are verified by simulation experiments. These simulations validate the controller’s ability to achieve constraint satisfaction, optimize transient processes, and enhance control accuracy. The results provide compelling evidence of the controller’s potential for real-world spacecraft control applications.

The impact sensitivity assessment of spacecraft is to obtain the probability of spacecraft encountering the OD/M (orbital debris or meteoroid), which is a prerequisite for survivability assessment of on-orbit spacecraft. An impact sensitivity assessment method of spacecraft based on virtual exterior wall was proposed to improve the computational efficiency. This method eliminates determination of the outermost surface elements of the spacecraft before generating the debris rays, which are assumed to originate from a non-concave virtual wall that completely wraps the spacecraft. The DistMesh method was adopted for the generating of the virtual wall to ensure its mesh quality. The influences of the sizes, mesh densities, shapes of the virtual wall on the efficiency and accuracy were considered to obtain the best combination of the size and mesh density of the wall and spacecraft. The results of this method were compared with those of S3DE (Survivability of Spacecraft in Space Debris Environment), BUMPER, MDPANTO, ESABASE2/Debris to verify the feasibility of the method. The PCHIP (Piecewise Cubic Hermite Interpolating Polynomial) was used to fit the size vs. flux relationship of the space debris to acquire the impact probability of OD/M with arbitrary size on the spacecraft.

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]

Controlled vacuum environments as in space applications represent a challenge for the lubrication of tribological components. In addition to common space lubricant requirements like, e.g., low evaporation, a broad operational temperature range and a high stability during operation, long-term-storage (LTS) properties have gained increasing attention recently. The term addresses the time-dependent stability of a lubricant under static conditions, which can mean chemical degradation processes such as oxidation on the one hand, but also the physical separation of oil and thickener in heterogeneous lubricants like greases. Due to the extended storage periods of lubricated components on-ground but also during a space mission for several years, it has to be ensured that a lubricant is still functional after LTS. This article depicts the development of a space lubricant grease with LTS properties. Firstly, LTS requirements and methods for their assessment are discussed. In the following, a systematic approach towards the design of a grease formulation compatible with LTS is described. Finally, the manufacturing of prototype formulations and their broad characterization by means of LTS behaviour, outgassing, and tribological performance is presented. [ABSTRACT FROM AUTHOR]

The aim of this research paper is to propose a framework to model, simulate and control the motion of a small spacecraft in the proximity of a space station. In particular, rendezvous in the presence of physical obstacles is tackled by a virtual potential theory within a modern manifold calculus setting and simulated numerically. The roto-translational motion of a spacecraft as well as the control fields are entirely formulated through a coordinate-free Lie group-type formalism. Likewise, the proposed control strategies are expressed in a coordinate-free setting through structured control fields. Several numerical simulations guide the reader through an evaluation of the most convenient control strategy among those devised in the present work. [ABSTRACT FROM AUTHOR]

This paper studies the navigation technology of the On-Orbit Servicing Spacecraft (OSS) and proposes an overall solution for the OSS navigation system. The composition of the OSS satellite navigation simulation system and simulation supporting environment are studied. Then a new SSUKF navigation filter is proposed based on the hyper spherical distributed feature sampling point transform algorithm (SSUT) and the non-scanning Kalman filter (UKF). The numerical simulation of the OSS system is studied based on MATLAB/RTW. Finally, the SSUKF algorithm and the conventional UKF algorithm are simulated digitally. The effectiveness and advanced nature of the hybrid Kalman filter applied to spacecraft autonomous navigation are verified. [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]

Permanent magnet synchronous motor (PMSM) servo system is the control and execution mechanism of spacecraft, and its precise modeling is of great practical significance for improving the performance of the spacecraft. A comprehensive model considering elasticity, backlash, and friction was established based on the structural composition and transmission characteristics of spacecraft. Firstly, an elastic-dead zone model was established based on the transmission characteristics, with two mass elastic and gear backlash dead zones as the main features. Then, an improved LuGre friction model was established that considers the effects of vibration and environmental temperature, and the three models were organically combined. Finally, a method for identifying nonlinear parameters of the system using sinusoidal sweep and the least squares method was proposed, and the accuracy of the constructed model was verified. The experiment shows that the error of the simulation results is consistent with the experimental results. This method can precisely obtain the servo system model, providing theoretical support for control method research. [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]

This paper presents a simplified method for launch vehicle coupled loads analysis (CLA) with spacecraft which is normally conducted in the process of the spacecraft development. The method allows spacecraft developers to calculate coupled vibration response of their own without performing transient analysis for the coupled whole system model. A new method, which allows us to obtain coupled vibration response by convolving non-loaded acceleration time history at the spacecraft boundary with corresponding impulse response of combined apparent mass matrices in frequency domain, has been proposed. The proposed method functions well in any CLA cases including lift-off forced events where loads accompany quasi-static component in nature. We propose a novel approach to escape the error caused in the process of convolution for the quasi-static event by making use of the method to convolve only dynamic acceleration response extracted from the original acceleration by inertia relief technique. Finally, we demonstrate the validity of the proposed method through the analysis using practical launch vehicle and spacecraft models.

It is known that today the problem of microvibration in spacecraft structures is relevant and widely researched. On- orbit operation of modern spacecrafts such as geodetic satellites, Earth observation satellites or space telescopes is associated with the use of onboard equipment sensitive to micro- vibration. In most cases, the conditions for the normal functioning of precision spacecraft equipment are provided by the operation of electromechanical devices, which in turn are sources of vibration. This article attempts to highlight the main approaches to the microvibration analysis that are known today. In this regard, the analysis of both native and foreign sources, in which the central problem is the analysis of microvibration, is carried out. This paper takes formulation of the microvibration problem and its role in the creation of modern spacecraft process. It also gives a literature review on methods and approaches to the microvibration analysis of spacecraft. Based on the analysis of publications on a similar topic, here one can find the tasks to solve in the process of the spacecraft microvibration analysis. The advantages, disadvantages and limitations of existing approaches and methods are revealed.

In order to preclude the impact of atmospheric factors on the quality of received data, state of the art astronomy research projects require spacecraft mounted telescopes. However, such designs entail the issue of sustaining the accuracy of telescope boresight attitude relative to the object of observation for the required period of time. This paper is a summary of design principles for an attitude determination and control system of a space telescope mounted rigidly onboard a spacecraft. The paper formulates requirements for attitude accuracy, and proposes a viable equipment configuration as well as a three- stage attitude control algorithm. The first stage of the algorithm ensures primary boresight pointing towards the target object prior to the object being captured in the telescope’s FoV based on star tracker and angular rate sensor data. The second stage of the algorithm ensures boresight positioning relative to the target object direction up to an error value allowing for the required image quality, based on telescope readings. The third stage of the algorithm ensures telescope boresight holding relative to the target object direction with a required accuracy for the duration of exposure. Interferences from operating satellite equipment are reviewed, and the principles of reducing their impact on sighting accuracy are formulated. Low inclination, low eccentricity geostationary orbit is recommended as the best operating environment for the space telescope’s scientific instrumentation, as well as for providing a continuous data interface between spacecraft and ground segment.

With the advent of expanded standards of transfer of information flows (DVBS2X type) from Earth remote sensing spacecraft (remote sensing), there was an opportunity to switch its parameters for transfer of the increased (reduced) data stream to VHSR operating time. For example, switching of modulation and coefficient of interference-immune LDPC (Low Density Parity Check codes) of coding. It allows to reduce power costs for data transmission and to raise the general information flow of data from remote Sensing Satellites (RSS) to Ground Stations (GS) several times. Estimations of theoretically possible transitions to the next transmission mode of data (transition to the level of 0.5-1.5 dB) are made that gives a total gain in increase in a flow of transmitted data can reach 2.0-2.5 times with the possible number of switchings up to 10. In practice, considering the fluctuation nature of the distribution ogtht the signal level during its transmission to Ground Stations, switching is provided in smaller (3-5) number of transitions (at achievement of a power stock 3-5 dB) that leads to a gain in data transmission by 1.5-2.0 times.

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.

Simplified models and algorithms have been developed to estimate the current value of the angle between the normal to the plane of the solar panel and the sun vector at the time of object survey and during the reorientation of the Earth remote sensing spacecraft, taking into account the relative movement of observation objects in the field of view of the spacecraft. The models differ from the existing ones in that they do not require knowledge of the spacecraft control programs at pitch, roll and yaw angles at each moment of time of the process of its targeted functioning and re-targeting, Based on the obtained models, a software module that does not require large computing resources used in the simulation software package simulating targeted functioning of the space surveillance system has been developed, taking into account the long-term orbital motion of satellites, changes in the parameters of the orbit, the sun vector and other factors. The adequacy of the developed models was verified by visualizing the orbital motion of the spacecraft and its turns during re-targeting, as well as by comparing the simulation results with the results of analytical calculation for particular cases of the initial data. Examples of the implementation of software windows are given.

An analysis and classification of methods for limiting technogenic pollution of the near-Earth space according to selected criteria are presented. A brief description of existing and prospective methods for limiting technogenic pollution of the near-Earth space is given, indicating the level of technological readiness and the scope of application of a particular method.

A digital model is described for estimating the drag force of the Earth's upper atmosphere acting on a low-orbit spacecraft. Unlike the classical model of aerodynamic force, the digital model is the result of a computer algorithm. This algorithm calculates the field of aerodynamic force values, taking into account changes in the flight altitude, aerodynamic coefficient and the area of the spacecraft midsection over time. The calculations also use a digital model of the dynamic density of the Earth's upper atmosphere. The digital model of the drag force of the Earth's upper atmosphere is useful in the design of low-orbit spacecraft with a low-thrust jet propulsion system (thrust order – millinewtons). On the basis of the calculated values of the aerodynamic force, it is possible to form requirements for the geometry of the spacecraft, attitude modes and arrangement of low-thrust jet engines (for example, electric jet engines). With the accumulation of big data in the form of tabular functions of aerodynamic force from the space of various design and ballistic parameters of low-orbit spacecraft, it is possible to create a machine model (ML-model). It will make forecasts of the values of the design characteristics for the low-orbit spacecraft being developed according to a limited set of initial requirements.

The extreme temperature fluctuations and the vacuum of the space environment make growing plants in outer space challenging. To simulate the temperature fluctuations and vacuum conditions associated with space environments, dry tomato seeds were placed in a thermal cycle simulator and vacuum simulator chamber of space systems, respectively. A Bradford method was used to determine the total protein content of each group of seeds. Sodium dodecyl-sulfate polyacrylamide gel electrophoresis was used to separate proteins. The seed of the thermal cycle group had the highest protein content (26 to 31 mg/ml), followed by control seeds (8-10 mg/ml) and the vacuum seeds (4-5.6 mg/ml). The molecular weights of the peptides ranged from 8 to 42 kDa. The intensity of the protein bands was significantly different in the thermal cycle group from the other two groups, and vacuum group had the lowest intensity. Water and oil released from seeds in the vacuum environment resulted in a reduction of protein content. In the thermal cycle group, the total protein content and the intensity of the bands were significantly higher than those treated with the control group, which can be attributed to the degradation of storage proteins involved in seed germination in the control group.

The design of the space hatch door mechanisms is crucial in the aerospace field, impacting not only durability and reliability but also the life safety of astronauts during space missions. This review extensively researches vehicle doors and hatches in civil and military systems across various environments, including land, sea, deep sea, aviation, aerospace, and extreme conditions. Specially, it focuses on the structural design of hatches and related mechanisms in civil aviation and military aerospace environments, such as opening and closing mechanisms, release mechanisms, locking mechanisms, sealing mechanisms, and the ergonomic design of door structures. The review highlights the integration of bionic design principles with hatch mechanisms to explore future solutions. By systematically examining these aeras, this review addresses the lack of comprehensive studies in previous reviews, which often overlook the interconnectivity and applicability of hatch mechanisms across different fields. The absence of such holistic reviews has led to fragmented knowledge and missed opportunities for cross-industry innovation. This review aims to fill these gaps by providing a wide range of design solutions and offering insights that can enhance the development of more reliable, efficient, and safe hatch mechanisms in aerospace and other high-stakes environments.