Search

Modern space missions are increasingly venturing across cislunar space, requiring expansion of space awareness functions. Legacy space domain awareness (SDA) systems were not originally built to detect and track cislunar objects, and this could require acquisition of new sensor systems. There are numerous parameters, including sensing type, altitude, and number of platforms that could be varied for each system. One of the proposed options is a pole-sitter satellite located at 2.5 million km above either the North or South Pole. One key advantage to any “pole-sitter” is that it has a position well outside the ecliptic plane and offers a unique, in some cases orthogonal viewing geometry that here to fore have not been developed for operational deployment. Such orbits offer continuous watching and tracking of candidate objects with the advantages of efficiency and characterization. The efficiency results from not having to repeatedly reacquire the object due to blockage by Earth eclipse; and once detected and continuously tracked, any behavior of the object begins to reveal its characteristics, pattern of life, and potential intents. In this paper, the physics of the pole-sitter trajectory, the trade of fuel usage against altitude, and updates in required technology are discussed. All these elements point towards the feasibility of demonstrating a pole-sitter SDA capability in the near term. In addition, this paper devises a proposed prototype using small spacecraft in conjunction with ground-based sensors along with descriptions of current technology ready for deployment. [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]

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

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

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]

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

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

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

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

Autonomous optical navigation is one of the most promising techniques to estimate the position and velocity of a spacecraft in Earth orbit without the supports of Earth-based tracking stations. To improve the navigation performance, this paper presents a novel autonomous optical navigation method, where a star camera on the spacecraft is utilized to measure the line-of-sight (LOS) directions of a number of non-cooperative space targets, whose position vectors are supposed to be not precisely known in advance, and an improved Q-learning based extended Kalman filter (IQEKF) is developed to obtain the accurate motion state estimate of both the spacecraft and the space targets based on the LOS direction measurements. The main advantage of the presented method is that the LOS directions of the space targets can be acquired with high-accuracy by using the state-of-the-art star camera, such that the superior navigation accuracy is achievable. In addition, the whole motion state of the spacecraft, such as position, velocity, and attitude, can be obtained with the star camera, in the case that the space targets and the stars are observed simultaneously. The high performance of the presented autonomous navigation method is illustrated through a representative simulation of a medium Earth orbit (MEO) satellite. Furthermore, the simulation results indicate that the IQEKF yields more accurate solutions than the traditional navigation filtering algorithms. [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]

The paper deals with the problem of low-quality voltage characterizing the common bus of a spacecraft (SC) power-supply system (PSS) with distributed three-level hierarchical control. The hierarchical control consists in the droop control of power modules connected in parallel forming a common bus for to provide uniform current distribution, which provides an increase in the impedance of this bus to such a value that is incompatible with the requirements for the voltage quality of SC onboard equipment. To solve the problem, it is proposed to apply control with three degrees of freedom (3-DOF) to divide the processes of controlling local voltages of power modules and droop control into different speeds, neutralizing the effect of the droop control of modules exerted on the impedance of the SC PSS common bus and, as a consequence, on voltage quality. The operability of the proposed solution has been verified by a series of experiments with the use of a laboratory test bench. [ABSTRACT FROM AUTHOR]

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

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

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

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

Guest‐Editor Mark Garcia takes us on a spatial journey through some of the notions and precedents of spaceships and spacecraft, their fictive and architectural precursors and current conceptual preoccupations. The contexts these new posthuman architectures and posthumans will have to occupy, endure and travel through are mind‐boggling in their varied complexity. [ABSTRACT FROM AUTHOR]

Copyright of Lubrication Engineering (0254-0150) is the property of Editorial Office of LUBRICATION ENGINEERING and its content may not be copied or emailed to multiple sites or posted to a listserv without the copyright holder's express written permission. However, users may print, download, or email articles for individual use. This abstract may be abridged. No warranty is given about the accuracy of the copy. Users should refer to the original published version of the material for the full abstract. (Copyright applies to all Abstracts.)

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.

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.

Planetary protection is a guiding principle aiming to prevent microbial contamination of the solar system by spacecraft (forward contamination) and extraterrestrial contamination of the Earth (backward contamination). Bioburden reduction on spacecraft, including cruise and landing systems, is required to prevent microbial contamination from Earth during space exploration missions. Several sterilization methods are available; however, selecting appropriate methods is essential to eliminate a broad spectrum of microorganisms without damaging spacecraft components during manufacturing and assembly. Here, we compared the effects of different bioburden reduction techniques, including dry heat, UV light, isopropyl alcohol (IPA), hydrogen peroxide (H2O2), vaporized hydrogen peroxide (VHP), and oxygen and argon plasma on microorganisms with different resistance capacities. These microorganisms included Bacillus atrophaeus spores and Aspergillus niger spores, Deinococcus radiodurans, and Brevundimonas diminuta, all important microorganisms for considering planetary protection. Bacillus atrophaeus spores showed the highest resistance to dry heat but could be reliably sterilized (i.e., under detection limit) through extended time or increased temperature. Aspergillus niger spores and D. radiodurans were highly resistant to UV light. Seventy percent of IPA and 7.5% of H2O2 treatments effectively sterilized D. radiodurans and B. diminuta but showed no immediate bactericidal effect against B. atrophaeus spores. IPA immediately sterilized A. niger spores, but H2O2 did not. During VHP treatment under reduced pressure, viable B. atrophaeus spores and A. niger spores were quickly reduced by approximately two log orders. Oxygen plasma sterilized D. radiodurans but did not eliminate B. atrophaeus spores. In contrast, argon plasma sterilized B. atrophaeus but not D. radiodurans. Therefore, dry heat could be used for heat-resistant component bioburden reduction, and VHP or plasma for non-heat-resistant components in bulk bioburden reduction. Furthermore, IPA, H2O2, or UV could be used for additional surface bioburden reduction during assembly and testing. The systemic comparison of sterilization efficiencies under identical experimental conditions in this study provides basic criteria for determining which sterilization techniques should be selected during bioburden reduction for forward planetary protection. [ABSTRACT FROM AUTHOR]

An analytical solution of the problem of the movement of a spacecraft from the starting point to the final point in a certain time is given. First, the method of fast sine expansions is used. The space problem considered here is essentially nonlinear, which necessitates the use of trigonometric interpolation methods that surpass all known interpolations in accuracy and simplicity. In this case, the problem of calculating Fourier coefficients by integral formulas is replaced by the solution of an orthogonal interpolation system. In this regard, two cases are considered on the segment [0, a]: universal interpolation and trigonometric sine and cosine interpolations. A theorem on the rapid decrease of expansion coefficients is proved, and a compact formula for calculating the interpolation coefficients is obtained. A general theory of fast expansions is given. It is shown that, in this case, the Fourier coefficients decrease significantly faster with the growth of the ordinal number compared to the Fourier coefficients in the classical case. This property makes it possible to significantly reduce the number of terms taken into account in the Fourier series, significantly increase the accuracy of calculations, and reduce the amount of calculations on a computer. An analysis of the obtained solutions of the spacecraft motion problem is carried out, and their comparison with the exact solution of the test problem is proposed. An approximate solution by the method of fast expansions can be taken as an exact one, since the input data of the problem used from reference books have a higher error. [ABSTRACT FROM AUTHOR]

The regular quaternion equations of the orbital motion of a spacecraft (SC) proposed by us earlier in four-dimensional Kustaanheimo–Stiefel variables (KS variables) are considered. These equations use as a new independent variable a variable related to real time by a differential relation (Sundman time transformation) containing the distance to the center of gravity. Various new regular quaternion equations in these variables and equations in regular quaternion osculating elements (slowly changing variables) are also constructed, in which the half-generalized eccentric anomaly, widely used in celestial mechanics and space-flight mechanics, is used as a new independent variable. Keplerian energy and time are used as additional variables in these equations. These equations are used to construct quaternion equations and relations in variations of KS variables and their first derivatives and in variations of Keplerian energy and real time; the isochronous derivatives of the KS variables and of their first derivatives and the matrix of isochronous derivatives for the elliptical Keplerian motion of the spacecraft are found, which are necessary for solving the problems of predicting and correcting its orbital motion. The results of a comparative study of the accuracy of the numerical integration of the Newtonian equations of the spatial restricted three-body problem (the Earth, the Moon, and a spacecraft) in Cartesian coordinates and the regular quaternion equations of this problem in KS variables are presented, which show that the accuracy of the numerical integration of regular quaternion equations is much higher (by several orders) than the accuracy of numerical integration of the equations in Cartesian coordinates. This substantiates the expediency of using regular quaternion equations of the orbital motion of a spacecraft and the quaternion equations and relations in the variations constructed in the paper on their basis for the prediction and correction of the orbital motion of a spacecraft. [ABSTRACT FROM AUTHOR]

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

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

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

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

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