Search

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 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.

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.

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.

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.

An executive precision universal mechatronic module is considered, which allows performing precision rotational motion of the output kinematic link. The universal mechatronic module is designed to create the kinematics of mechatronic guidance systems and positioning of objects of aerospace technology. The key features of the executive precision universal mechatronic module are high positioning accuracy, low weight and overall dimensions. The aim of the work is to create a finite element model and analyze dynamic loads to assess the resistance of the structure to external vibration loads. The first section describes the creation of a finite element model taking into account the radial and axial stiffness of the supporting structural elements. The second section describes performing static analysis to create a preloaded state, and performing dynamic analysis. The influence of sinusoidal and random vibrations is investigated, the natural frequencies of the universal mechatronic module design are determined.

Design models and algorithms have been developed to estimate the spatial direction of the reorientation of spacecraft for optoelectronic observation of the Earths surface during object shooting, taking into account the relative movement of the surveillance objects in the field of view of the spacecraft. Modeling is simplified due to the choice of such coordinate systems in which the spacecraft is stationary and the surveillance objects move relative to it. The direction of rotation of the spacecraft with a given angular velocity is found at the point of the upcoming intersection of the optical axis of the observation equipment on the Earths surface with the surveillance object moving in the selected coordinate system. The models and algorithms were developed in order to expand the capabilities of Samara Universitys existing software package for evaluating the target indicators and design characteristics of observation spacecraft. Based on the developed re-targeting models, algorithms have been developed for selecting the sequence of shooting observation objects according to the criterion of their maximum number on the orbit of the spacecraft flight and counting the captured objects to assess the performance of survey. An example of the implementation of one of the graphic windows of the software with the developed models and algorithms is given, in which the process of re-targeting the spacecraft is visualized, which is one of the ways to verify the adequacy of the developed models and algorithms.

This article characterizes the spacecraft as a means of performing various tasks, including by precision equipment. Emphasis is placed on vibration as one of the influencing factors and on the need for a low level of vibration on the spacecraft. Among the side effects of the electric pump unit of the thermal control system, the urgent task is to reduce vibration activity when developing a new model for advanced spacecraft. An integrated approach to solving the problem is indicated with the consideration of electric pump unit as a system and the selection of several directions to achieve the root-mean square value of vibration acceleration of 0,01g and below. Possible methods for reducing vibration activity in each of the directions and implemented solutions in the developed sample are described in detail. On several versions of electric pump unit for experimental determination of the level of root-mean square value vibration acceleration at a frequency of 5–1000 Hz. According to the results of the experiment, the contribution to the reduction of vibroactivity in several directions is estimated. It is concluded that only an integrated approach makes it possible to achieve a very low level of vibration activity of electric pump unit.

The aim of the work was to create a software package for modeling the solution of target problems by multi-satellite constellations. The software consists of modules for the dynamics of the center of mass of the satellites, remote sensing of the Earth, satellite communications (including routing of messages over inter-satellite communications), space monitoring and others. The main feature of the implementation was the mathematical and computational optimization of calculations, which made it possible to carry them out on a conventional computer without involving additional computational power. In the course of the work, corrections to the Everhart method were proposed to improve the trajectory integration step. Parallel implementation also helped to significantly speed up the calculations. To illustrate the work, examples of solving two different problems are given: calculating the optimal orbit for the task of monitoring outer space and solving the problem of determining the minimum number of satellites needed to scan the border of the Russian Federation once every half an hour. A feature of the soft is the work with the constellation as an integral object at the level of setting tasks and the subsequent analytics of its solution, which is presented in the form of information on target tasks, subgroups of the constellation, spacecraft and the entire constellation.

Background. In modern scientific publications of foreign scientists, the concept of controlling an electromechanical engine-flywheel unit with high-precision feedback on the speed of rotation of the rotor introduced into it is described. Examples of the implementation of this concept in domestic spacecraft do not meet the modern qualitative requirements for the management of highly dynamic vehicles. The purpose of the study is to create a new spacecraft control system based on controlling the rotation speed of the rotor of the electromechanical engine block-flywheel. Materials and methods. Practical tests of a typical number of domestic flywheel engines have been carried out using signal modeling approaches of a complex of testing equipment developed in “VNIIEM Corporation” JC. Results. Based on the tests results, a new prototype of the flywheel engine was created, on the basis of which the main technical characteristics were obtained and the restrictive requirements for the orientation and stabilization of the spacecraft necessary to create a control function for the rotation speed of the rotor of the electromechanical block of the future control system were determined. Conclusions. The created spacecraft control system made it possible to find the optimal target function for controlling the rotation speed of the rotor of the electromechanical unit of the flywheel engine type, which provided an increase in the efficiency indicator for a particular indicator of the accuracy of working out the required speed by more than 12% compared with the presented domestic spacecraft.

The article discusses the key role of the use of automation in the stands, as well as the ignition system of the stand of low-thrust engines and the features of its operation as part of the stand. The ignition system scheme was developed, assembled and worked out, on the basis of which the ignition of the components of the propulsion system mixture will be carried out. The ignition system was tested by directly connecting the power supply to the circuit, which during the tests showed its suitability for use due to the long and continuous operation of the system. The power was turned on by controlling relay switches based on the Arduino MEGA 2560. Also, an algorithm was developed for the operation of the automation system for switching on, off and checking the entire system before testing the propulsion system on the stand of low-thrust engines to ensure automation and safety of using the stand. This algorithm includes control of two channels of the oxidizer and fuel pipeline, emergency power off, opening and closing of cut-off valves, taking readings and checking pressure gauges and flow meters, as well as shutting down the system after testing low-thrust engines using an Arduino MEGA controller.

In recent years, the number of space objects located in near-Earth outer space, especially in the near operational zone, has increased significantly due to the build-up of space groupings, including dual-use (for example, Starlink) and the remnants of their vital activity (space debris). This factor increases the importance of the task of recognizing and classifying space objects by type in the shortest possible time and entering them into the main catalog of space objects. The developed methodology allows automated analysis of optical images of space objects using software to solve the problem of their recognition and classification by type using a convolutional neural network. The purpose of the study is to increase the efficiency of processing and analysis of optical images of spacecraft. The experimental results of the study confirm the achievement of the research goal. The developed methodology contributes to the development of software and hardware for image processing and can be used in calculations and data preparation for information support of interested officials. For the first time, a training set for a convolutional neural network has been prepared using real optical images of spacecraft obtained in the visible range.

The classical problem of optimal control of the attitude maneuver of a spacecraft as a rigid body of arbitrary dynamic configuration under arbitrary boundary conditions for the angular position and angular velocity of a spacecraft without restriction on the control vector function and with a fixed transition time is considered. As a criterion of optimality, the functional of the energy spent on the rotation of a spacecraft is used. Within the bounds of the Poinsot concept describing arbitrary angular motion of a rigid body in terms of generalized conical motion, a modification of the problem of optimal control of the angular motion of a spacecraft is carried out and its trajectory is given in this class of motions. At the same time, the generality of the original problem is practically not violated, since the known exact solutions to the classical problem of optimal angular motion of a dynamically symmetric spacecraft in cases of plane rotation or regular precession and similar solutions of the modified problem completely coincide; in other cases, in numerical calculations of the classical and modified problems, the discrepancy between the values of the optimization functional is no more than a few percent, including spacecraft rotations at large angles. Therefore, the proposed solution of the modified problem can be used as quasi-optimal with respect to the classical problem. Explicit expressions for the quaternion of the orientation and the vector of angular velocity of a spacecraft are given, a formula for the vector of the control moment of a spacecraft is obtained based on the solution of the inverse problem of the dynamics of a rigid body. The quasi-optimal algorithm for optimal rotation of a spacecraft is given. Numerical examples showing the proximity of solutions to the classical and modified problems of optimal reorientation of a spacecraft are given.

The article presents the theoretical foundations of automated design of on-board equipment layout in spacecraft compartments in terms of topology. The purpose of the work is to create a flexible arrangement model that would allow solving the problem for any configuration of the structure, taking into account the engineering and operational restrictions imposed on the objects located in a limited space. The technology is formalized according to the methods of the foundations of topology and in this work is demonstrated for a two-dimensional case by comparing it with a reference design-layout scheme. The functional, installation, dimensional, thermal, vibroacoustic requirements are taken into account with the mass-centering restrictions of the layout. In the course of the study, a new approach to describing the design process of assembling on-board equipment was obtained, which can be fully digitized and then integrated into known computer-aided design systems. It has been verified using the example of the power supply system of the Yantar-2k spacecraft.

The article presents the results of the analysis of heating of printed circuit conductors of printed circuit boards mounted on a metal base, on-board device of spacecraft when pulsed current flows through them. Thermal transient and impulse characteristics of a system consisting of a printed conductor and a printed circuit board are considered. The analysis of the dependence of the temperature rise rate of printed conductors and the duration of transient processes on the linear dimensions of the system elements is given. It is noted in the article that the time of transient processes most strongly depends on the thickness of the layer of insulating materials. All the results were obtained in the process of modeling transients in the ANSYS 2019 R1 Transient Thermal module. Practical recommendations on the consideration of thermal transients in printed conductors in the design of printed circuit boards are given.

The authors describe an algorithm that allows calculating the parameters of maneuvers performed on several turns by the low-thrust engine, which ensure the flight of the active spacecraft to the specified vicinity of the target space object. The movement takes place in the vicinity of a circular orbit. Linearized equations of motion are used in solving the problem. The influence of the non-centrality of the gravitational field and the atmosphere is not taken into account. The determination of maneuver parameters takes place in three stages. At the first and third stages, the parameters of the pulse transition and the transition performed by the low-thrust engine, are determined analytically. At the second stage, the distribution of maneuvering between turns, which provides a solution to the meeting problem, is carried out by iterating over one variable. This method of solving the problem provides simplicity and high reliability of determining the parameters of maneuvers, which allows it to be used on board the spacecraft. The paper investigates the dependence of the total characteristic speed of solving the meeting problem on the number of turns of the flight and the magnitude of the engine thrust.

A design feature of the construction of electric heaters for non-pressurized spacecraft using a heating element in the form of a tape of a material with high resistivity, located on a flat base (electrical insulating substrate) in the form of a snake with the organization of the necessary gaps between the tapes, is revealed. A mathematical model has been developed for calculating the main design parameters of electric heaters: power, mass, heated surface area. The nomenclature of efficiency indicators used as optimization criteria has been formed: power utilization factor, filling factor of an electrically insulating substrate tape, surface density, energy-mass efficiency, thermal efficiency of heating, target efficiency. The methodological principles for ranking the requirements for the parameters of electric heaters are determined using the selected optimization criteria: the preferred use of tape materials with high resistivity and minimum thickness. A method is formulated for step-by-step calculation of the parameters of electric heaters in the presence of restrictions on the maximum heat flux density determined from the conditions for the use of electric heaters. General recommendations have been developed for choosing the parameters of electric heaters: use increased voltage, minimize the thickness of the heating element tape and the width of the gap between the tapes, determine the minimum value of the tape length for the maximum heat flux density at a given voltage rating.

The gear assembly of the rotational motion drive of the spacecraft system is considered. Such gear assemblies are one of the main elements of spacecraft systems with the use of large-sized structures that are folded in the transport position. Due to the development of computer technology, it becomes possible to create more accurate simulation models of such structures. The aim of the work is to create a simulation model that allows to analyze emerging loads on a wave transmission, which is part of the gear assembly of the spacecraft drive. Section one discusses the process of creating a finite element model of the drive. One of the features of creating a finite element model is a method of taking into account the preliminary stress-strain state in the structural elements of the wave transmission. Section two analyzes the data obtained during the calculation. The obtained equivalent stresses on the structural parts, contact stresses in the gear engagement are investigated, the torsional rigidity of the gearbox elements is determined, various tooth positions are considered during movement. The influence of loads applied to the output link of the gearbox on the surface of the gears is investigated. The obtained simulation model allows us to consider similar gear assemblies in which wave transmissions are used.

This paper presents an analysis of various designs of transformable membrane reflector antennas for spacecrafts, considers the dependence of the surface accuracy on the operating frequency of the antenna and the requirements imposed on deviations of the antenna surface from the paraboloid. A classification of membrane transformable antennas is presented depending on the presence or absence of supporting elements and stiffeners. Structural schemes for opening membrane antennas are described, from the classical umbrella opening scheme to original schemes for winding ribs on a central hub, tension supporting elements, winding paraboloid segments on a central hub, and folding the reflector surface using origami methods. The design of an offset membrane transformable antenna is considered. Modern materials for the manufacture of reflectors of membrane antennas, such as polymer composite materials with triaxial weaving of carbon fabric with a silicone matrix, polyamide and polyimide films with and without reinforcement, are considered. Materials with shape memory for opening membrane antennas are considered. Calculations of structures for natural vibration frequencies, calculations for bending of folds are described. A description of ground-based experimental testing, in particular, for the disclosure of various types of membrane antennas, was carried out.

The paper presents the results of computational study of sections of spatial truss structures. They can form large supporting structures on spacecraft to place onboard equipment. The calculations were carried out by the finite element method in FEMAP with Nastran. The chosen section, the object of study, is a set of straight rods rigidly connected at the nodes in such a way that the cross section of the truss structure is a triangle. Calculation models, procedures for calculation and results analysis are presented. The purpose of the calculation is to determine how the relative position of the rods in the structural scheme affects the mechanical properties of the structure. The main criterion for strength estimating was the magnitude of the stresses derived by load. Stiffness was determined by the value of the first natural frequency. The sequential addition of rods and varying their connection allow modifying the structural schemes of truss structures. Next, the mechanical properties of the structure which effected by made modifications were evaluated again. Thus, six structural schemes of sections of the truss structure, obtained from the results of the study, have been developed. At the same time, the mass of the section, its shape and dimensions, the material and shape of the rods section as well as the initial and boundary conditions, remained unchanged. Based on the results of the analysis, someone can notice that each structure has unique mechanical characteristics. Thus, the paper gives recommendations for choosing a specific structural scheme of the section, depending on the required operating conditions and acceptable manufacturing technology. So, the criteria for choosing one of the above various sections of a truss structure or the principle of its building can be the complexity of manufacturing, maximum stiffness and strength, or minimum displacement. Structural schemes of sections with the highest strength and capability have been selected for use as part of spatial truss rods of spacecraft.

Background. The object of research is existing and prospective spacecraft. The subject of the study is the control of the technical condition and diagnostics of the onboard equipment of spacecraft. The aim of the work is to develop a model of an intelligent monitoring and diagnostics system, the use of which will increase the reliability of determining the technical condition of the onboard equipment of spacecraft at the stage of their testing and application. Materials and methods. When developing a model of an intelligent system for monitoring and diagnosing the onboard equipment of spacecraft, the main attention is paid to a method based on machine learning, which allows you to build predictive models using both retrospective and current data received in real time from measuring equipment and obtain patterns that are not obvious at first glance. Results. An approach is proposed to increase the reliability of determining the technical condition through the use of an intelligent system for monitoring and diagnosing the onboard equipment of spacecraft. The intellectualization of this system is determined by the presence of mathematical, algorithmic, software and instrumental components of intellectual support for decision makers in the management of spacecraft. Conclusions. The developed model of an intelligent system for monitoring and diagnosing the onboard equipment of spacecraft makes it possible to expand the capabilities of the spacecraft by implementing the functions of correcting the algorithms of the onboard and special software based on the results of machine learning and neural network processing of television measurements, contributing to a more reliable determination of the technical condition of the onboard equipment. The proposed model is proposed to be used at the stages of testing and application of spacecraft to increase the reliability of determining the technical condition of on-board equipment and the efficiency of taking measures to prevent the occurrence of emergency situations.

The place of electric rocket thrusters in a spacecraft propulsion system is pointed out; the tasks for such propulsion system are described. The modern classification of electric thrusters is given, in accordance of that one there are four classes, dissimilar due to a nature of processes that impart a kinetic energy to a propellant: electrothermal, electrostatic, pulsed and magnetoplasmodynamic. In order to illustrate an electric rocket thrusters sate of art, examples different classes electric rocket thrusters for modern spacecrafts are presented; the technical characteristics, such a thrust, specific impulse, powered energy and efficiency are systematized. As a general electric rocket thrusters for spacecraft propulsion system were chosen electrostatic thrusters: stationary plasma (Hall thrusters) and ion thusters. For the different classes of electric rocket thrusters the common characteristics are defined as a set of: specific impulse (that describes thruster propellant efficiency), thrust cost (determines electric power consumption per unit thrust) and propellant type (combines thruster’s technical and economical characteristics). A necessity is stated to take into account the cost parameters of an electric rocket thruster design and a manufacturing during a propulsion system developing for a next generation spacecraft. The tasks are determined for the further research in the area of a techno-economic efficiency of electric rocket thrusters exploiting as a part of a spacecraft’s propulsion system.

It is shown that the basis for the creation of an «artificial» mass which is not related to either the volume or the density of the material contained in an artificially inert object, is the requirement to satisfy Newton's second law. Of interest is the possibility of creating conditions for the emergence of an artificial moment of inertia. Since the principle of equivalence between inertial and gravitational mass does not apply to artificial mass, a device with artificial mass can be incomparably lighter than its mechanical counterpart with the same moment of inertia. This quality makes such a device particularly attractive for use in astronautics as light flywheels for orienting spacecraft. The purpose of the study is to analytically synthesize the artificial moment of inertia and establish the parameters that determine it. If an artificial electric flywheel is placed in a «black box» with the shaft outside, then no experiments can determine whether the artificial or «natural» flywheel is inside. The main advantages of an artificial or electromagnetic flywheel over a «natural» one are incomparably lower weight and the possibility of electrically controlling its moment of inertia over a wide range by changing the magnetic field (excitation) and capacitance.

The issues of ensuring the functioning of an additional energy-receiving channel for the onboard power supply system of a low-orbit spacecraft are considered. To compensate for the shortage of electricity on board the spacecraft, various options for remote power supply from the system of space power stations transmitting energy to the solar battery by laser radiation during periods of its forced inactivity are possible. Long shadow-sunlight periods of time are a distinctive feature of functioning of low-orbit spacecraft. During these periods the solar battery is idle, and, in addition, even in the daylight portion of the spacecraft's orbit, the energy output from the spacecraft may be reduced to almost zero during the operation of some special systems that require special modes of spacecraft orientation. Reception of energy from a CES with energy flux density higher than that of the solar flux may lead to overheating of the solar battery panel of conventional design, a decrease in its efficiency and even to its failure. Therefore, the analysis of thermal modes of reception and conversion of laser radiation energy which affect the energy efficiency of the solar battery is an important aspect of remote power supply. Relationships of a mathematical model for estimating the available power supply from the solar battery operating in the mode of optimizing control of its power are proposed. The results of approbation of the model for maintaining the safe thermal mode of the solar battery panels during remote power supply of the spacecraft are presented.

The goal of the work is to research the case of isotopy of spacecraft compartments in the problem of on-board equipment location. A definition of this engineering topological property is given in the article. The most frequently used spacecraft compartment hull shapes and various compositions of their casings are considered. Also, general strategies of taking decisions that determine the choice of modeling technologies and performance factors are discussed. A method of comparing various layouts on the strength of the object location performance factors was developed. The problem of on-board equipment accommodation is solved. Generalized mathematical models of performance factors are presented. Design models and diagrams are constructed. The results obtained make it possible to upgrade algorithms of accommodation.

An algorithm as well as a program for calculating mechanical loads in spacecraft structure are developed in the frame of its preliminary strength assessment. The algorithm was developed with the aim of calculating loads of several design cases simultaneously. The main object for study of this paper was an algorithm for calculating internal forces for a bending beam. As part of a preliminary assessment, such an algorithm can be used to calculate the responses, which are appeared in the spacecraft structure during its ground and flight operations. In order to provide an effective use of developed algorithm, a computer program has been written. This program provides an automated calculation of loads in the spacecraft structure with the ability to output data in a format which convenient for its further use. Such parameters as geometrical data, mass properties of the spacecraft, its test and launch conditions, qualification factors and safety coefficients are initial data to start the calculation. To verify the results obtained with the program, they were compared with the results of the analytical solution of a similar problem. From the check it was revealed, that they are coincide.

The paper reports a brief description of spacecraft operating conditions, the main reasons of heating and thermal gradient appearance and need to reassign the thermal energy. Active thermal control systems and their advantages are considered, spacecraft for which the use of this type of thermal control systems is a priority. The electric pumping unit is pointed as a key unit of active thermal control systems. The electric pump unit is considered from the as the electromechanical system, its key elements are pointed. A description of the preferred pump types is reported and the types of active thermal control systems are briefly discussed. The foreign and domestic operating experience of spacecraft electric pumping units, the features of their designs are considered, the most common types of key elements are determined. Based on the results of the review, it is concluded that the most relevant layout of the electric pump unit is a centrifugal electrical pump with a brushless DC motor and hydrodynamic bearings. It is also indicated that the electric pump unit is a product with a long lifetime, which complicates the task of monitoring the technical condition in order to prevent failure.

The work is devoted to the protection of a spacecraft from the influence of unacceptable internal vibration sources. The urgency of reducing the vibration activity on board the spacecraft to improve the accuracy of the target equipment is indicated. A particular problem of vibration protection of the spacecraft platform from a vibration source – an electric pump unit of a liquid thermal control system – is being solved. The basic requirements for electric pump unit vibration protection have been determined. Possible ways to reduce the level of vibration excited by the electric pump unit on the surface of the spacecraft fixation are considered. Particular attention is paid to such vibration protection methods as damping and vibration isolation, implemented by installing special vibration protection devices between the source (electric pump unit) and the object (spacecraft) – vibration isolators and vibration dampers. The principles of operation of vibration dampers and vibration isolators, the most common materials for vibration dampers are described. Examples of constructive solutions for linear single-axial vibration isolators are considered, recommendations for the use of promising products are developed. Particularemphasis is placed on the use of metal rubber as a material for vibration isolators. With regard to a specific design of electric pump unit, a diagram of the spatial structure of vibration isolation is proposed. Formulas for calculation are given in detail, a mathematical model of the vibration isolation system is developed. The procedure for calculating the parameters of the system has been formed. Based on the model, the maximum possible level of vibration suppression in the mid-frequency region was determined. Minimum required number of operable pixels was identified for monitoring the water surface with sufficient accuracy and reliability.

The main issues associated with the development of two-phase mechanically pumped loops (2-MPL) for thermal control systems of spacecraft with large heat dissipation were formulated back in the early 80s. They have undeniable advantages over single-phase loops with mechanical pumping and two-phase capillary pumped loops at power more than 6 kW and heat transfer distance more than 10 meters. Intensive research and development of such systems started in the USA together with European, Canadian and Japanese specialists due to plans to build new high-power spacecraft and the Space Station Freedom project. In the 90's, S. P. Korolev Rocket and Space Corporation Energia (Russia) was developing a 2-MPL for the Russian segment of the International Space Station with the capacity of 20...30 kW. For this purpose, leading research organizations of the former Soviet Union were involved. In the last two decades, interest in two-phase heat transfer loops has significantly increased because of high-power stationary communications satellites and autonomous spacecraft for Lunar and Martian missions. The paper presents a retrospective review of worldwide developments of 2-MPLs for thermal control systems of spacecraft with large heat dissipation from the early 80's to the present. The participation of scientists and engineers of the Ukrainian National Aerospace University "KhAI" and the Center of Technical Physics is considered. The main directions of research, development results, and scientific and technical problems on the way to the practical implementation of such system are considered. Despite a large amount of research and development work done, there were no practically implemented projects of spacecraft with the high-power thermal control system until recent days. The first powerful stationary satellite with the 2-MPL was SES-17 satellite on the NEOSAT platform by Thales Alenia Space - France. The satellite was successfully launched into space on October 24, 2021 by onboard Ariane 5 launcher operated by Arianespace from the Europe’s Spaceport in Kourou, French Guiana.

The article proposes a refined method for calculating the width of the conductors of printed circuit boards on a metal base for the onboard devices of spacecraft, depending on the current flowing. The constructed refined mathematical model of the process of conductive heat exchange between the conductors and the metal base is described. The results of calculations of various, most common, locations of layers of printed circuit boards are presented. An analysis was carried out and a refined methodology was developed based on the results obtained. It allows you to easily (without complicate calculations) calculate the necessary values of the width of conductors. This technique is based on graphical methods, but allows you to perform technical calculations with sufficient accuracy. Accuracy is achieved by using special formulas that simplify the determination of the value of a physical quantity on a logarithmic scale. The disadvantages of the proposed method are indicated.

The basis for ensuring a high level of reliability and safety in dynamic operations is a comprehensive experimental development of products in conditions as close to the real conditions of operation as possible with the use of mathematical and simulation modeling. The technical task of conducting dynamic operations in orbit is analyzed, which in turn boils down to two basic: bringing the initial state of the spacecraft grouping into a given area of space and performing maneuvers in a given area of space during the active life. These management tasks and terminal requirements are considered to be initial in relation to the stochastic and minimax approaches that are offered for use in the construction of working algorithms. The method of developing a software-modeling complex of the decisive task of developing the means of conducting dynamic operations of the spacecraft has been proposed. The application is illustrated by examples.

In order to further improve the methods of protecting the onboard spacecraft equipment from the effects of electrostatic discharges, based on a comprehensive study of the problem of the occurrence of electrostatic discharges, the authors carried out the studies described in this article to analyze the functioning of the onboard spacecraft equipment under the influence of electrostatic discharges. In the course of the research, the methodology for assessing the resistance of the onboard equipment of the spacecraft to the effects of electrostatic discharges was refined, based on the generalization of the calculation and analytical work. The analysis of the resistance of the onboard equipment of the spacecraft was carried out taking into account the main mechanisms of the effect of electrostatic discharges on the onboard space network. Taking into account the need to assess the resistance of the onboard space network equipment to electrostatic discharges outside real operating conditions, the authors assess the resistance of the onboard space network equipment on the basis of mathematical and physical modeling. As a result of the study, an algorithm was developed for assessing the resistance of the onboard equipment of a spacecraft to the effects of electrostatic discharges, which is described in detail in this article.

This article presents the methods of complex computational and analytical assessment of ensuring electromagnetic compatibility of technical means (electrical, electronic and radio-electronic means of launch vehicles and spacecraft). These methods allow us to conduct a comprehensive computational and analytical assessment of the electromagnetic compatibility of the technical means of rocket and space technology, which in its essence allows us to assess the mutual influence of the technical means of rocket and space technology already at the design stage of rocket and space technology objects. The analysis of the results of the assessment of the electromagnetic compatibility of the technical means of rocket and space technology allows us to develop organizational and technical measures aimed at ensuring the electromagnetic compatibility of the technical means of rocket and space technology at all stages of the life cycle of objects of rocket and space technology. To assess the electromagnetic compatibility of the technical means of rocket and space technology, an integrated approach was chosen due to the large dimension of the initial data and the variety of possible variants of interference penetration.

Solar arrays supply electrical power to spacecraft equipment and also provide charging of electrochemical batteries used in the shadow sections of the orbit. Photovoltaic converters of light energy of solar radiation form the basis of the solar battery. Their principle of operation is based on the phenomenon of the photoelectric effect. The article provides an overview of the development of solar cells and the classification of modern designs of solar cells for spacecraft. The review considers stationary and deployable solar batteries used both on the first spacecraft and on space stations. The classification of solar cell designs is made taking into account their characteristic features. These features are the rigidity of the supporting structure, the method of placement in the starting position and the method of orientation towards the light source. The classification covered rigid panel solar arrays, flexible substrate solar panels, inflatable solar arrays, self-expanding solar arrays, and solar concentrator panels. In each design group of this classification, corresponding examples of solar cells are presented. The presented review and classification makes it possible to track trends in the development of solar array designs for spacecraft.

The important problem for rocket-space, aviation and transport equipment fields is designing high-resource electric pumps. Electric pumps are active energy conversion devices to provide the thermal control system operation. Electric pumps failure leads to fault of the interfaced equipment, system failures and significant economic damage accidents. The most unfavorable factor for electric pump reliability is vibration. That the reason of relevance, practicability and importance to research the electric pumps vibration activity, the reasons for the appearance of vibrations and ways to control them. The paper provides the main sources of electric pump of the spacecraft vibrations as mechanical, hydrodynamic and electrical. This paper presents the recommendations of foreign and domestic authors to control each vibration source. To solve the problem of reducing the electric pump vibration activity the research of the technical level of existing developments, results, main technical solutions and technical solutions trends was produced. The main directions of further electric pump vibration activity researches leading to decreasing of vibration activity were defined on this paper.

The purpose of the study is an analytical description of the section of the ballistic trajectory corresponding to the normal fall of the spacecraft on the surface of an atmosphereless planet. In this case, the motion of a normally falling body is characterized by an increasing acceleration of gravity. The problem of the speed, time and acceleration of the normal fall of a body on the planet's surface in the absence of an atmosphere is reduced to solving a second-order differential equation, which is solved by the standard method. A feature of the solution is the formal use of the tabular integral at an intermediate stage. It turned out, however, that his formula is unreliable, namely, the derivative of the right-hand side is not equal to the integrand. It follows from this that the possible existing solutions to this problem, based on the use of the indicated tabular integral, are incorrect. The article presents the correction of this tabular integral, which is an incidental result of the study. In this work, the time equation of motion of a body normally falling on the surface of the planet in the absence of an atmosphere, as well as the time equations of its speed and acceleration are obtained. The results obtained can be useful in calculating passive gravity assist during interplanetary flights and calculating the sheer fall of small celestial bodies and spent structural element s of spacecraft.

An important stage in the development of temperature control interface modules for measuring instruments, implemented according to the main-modular principle of construction on the basis of a central instrument module and used as part of a measuring system on spacecraft for various purposes, is to ensure their noise-immune operation when the measuring system is exposed to a combination of various types of interference which determine the electromagnetic environment on the spacecraft. The article is devoted to the analysis of the characteristics of various types of interference affecting the measuring system, identifying the ways of their penetration into the temperature control interface module and determining the influence of the characteristics of these interference on the choice of measures to ensure the required noise immunity of the temperature control interface module as part of the measuring device. It is shown that the paths of interference penetration into the interface temperature control module depend on the frequency and time characteristics of the interference affecting the measuring system. Measures to mitigate these interference must be determined after assessing the danger of their penetration into the interface temperature control module for each path separately in the entire frequency range of their impact. Based on the results of such assessments, a set of reasonable measures should be determined to ensure the noise immunity of the interface temperature control module as part of the measuring device, implemented in combination at the design levels of the measuring system, the measuring device and the interface temperature control module.

The article discusses the features of the radiation of ultrafast point clusters of charged particles moving uniformly near a gyrotropic medium interface. It is shown that some types of electromagnetic radiation – transient and/or Cherenkov radiation – have the characteristics of superchiral fields. Therefore, they can be effectively used to study chiral structures (for example, to detect circular dichroism, the frequency characteristics of refractive indices), various materials, including biomaterials. Ultrafast (relativistic) particles can serve as a «tool» not only for studying the structure of various materials. They can also be used as «generators» of quasiparticles that determine the «dynamic» properties of the materials under study, as well as the features of their interaction with radiation of various nature and response to external influences. In this paper, some types of circularly polarized EM waves propagating in optically active (magnetoactive, naturally active, gyrotropic, and chiral) media are considered. Using the generalized reciprocity theorem for media characterized by the Hermitian permittivity tensor, we consider the transient and Cherenkov radiation excited by a uniformly moving bunch of charged particles when it crosses (or moves along) the interface of media, one of which is an optically active gyrotropic medium. It is shown that the superchiral electromagnetic fields of the transition and Cherenkov radiation of dipoles can serve as a source of chiral collective excitations in magnetoactive and naturally active media. The investigated mechanisms of interaction of electromagnetic radiation with chiral materials (structures and media) are one of the possible physical approaches to solving the problem of the chiral purity of the biosphere and to elucidate the factor of deracemization of the organic primeval environment. A new hypothesis is presented, suggesting that ultra-high-speed clumps of charged particles of cosmic origin can cause deracemization of the prebiosphere.

The article considers a spacecraft for remote sensing of the Earth with high-resolution or ultra-high-resolution optical-electronic equipment. During the shooting process, the recorded image constantly moves through the photodetector matrix at a non-constant and/or excessive velocity, which is not suitable for this photodetector. The purpose of the article is to synthesize a method for the control of the orientation and stabilization of the remote sensing spacecraft, which will provide a strictly specified velocity of the image motion on the photodetector. It is proposed to find such a law of motion (functional dependences of the angular rate of the remote sensing spacecraft on time), which will allow, when applied in the control loop, to compensate for the image motion velocities that are unsuitable for this photodetector. The method used consists in time differentiation of the fundamental equation of space photogrammetry in the guiding cosines, as well as in differentiation of the matrix of guiding cosines. This provides a transition between the guiding cosines in the space of images and the space of objects. The result obtained in the article is the derived equation of space photogrammetry in kinematic form, as well as the functional dependences of angular rate on time. In the present article, a mathematical model of scanning images of the Earth’s landscapes with the help of remote sensing spacecraft is compiled. The obtained functional dependences can be applied in the development of on-board algorithms for controlling the orientation and stabilization of the remote sensing spacecraft. When implementing orientation and stabilization control in the on-board computer based on the obtained functional dependencies, a strictly specified speed of image movement in the focal plane of the on-board shooting equipment can be provided, and, consequently, the quality of the scanned image is improved by improving the function of transmitting the modulation of the kinematic “smudge” (blurring) of the image.

In the quaternion formulation, the problem of mathematical modeling of the spacecraft movement in an elliptical orbit was considered. Control is an acceleration vector from jet thrust. Control modulus is constant. The control is directed orthogonally to the plane of the spacecraft orbit. The quaternion differential equation of an orbital coordinate system orientation was used to describe spacecraft movement. An approximate analytical solution of the quaternion differential equation of the orbital coordinate system orientation in the form of an expansion system of linearly independent basis functions was constructed. The method of pointwise collocation was used to find unknown quaternion coefficients of this decomposition. The above decomposition was simplified taking into account the well-known solution of the orientation equation of the orbital coordinate system for the case when the spacecraft orbit is circular. With respect to the desired coefficients, a system of linear algebraic equations is obtained in which the components of the stiffness matrix and the column of free terms are quaternions. A program in Python was written to conduct numerical simulations of the spacecraft movement. A comparison of calculations for analytical formulas obtained in the paper and the numerical solution of the Cauchy problem by the Runge – Kutta method of the 4th order accuracy was done. Error tables have been obtained for determining the orientation of the orbital coordinate system for cases when basic functions are polynomials and trigonometric functions. Examples of numerical solution of the problem are 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 describing changes in the components of error quaternion in determinig the orientation of the orbital coordinate system are constructed. The analysis of the received results is given. The features and regularities of the spacecraft movement on an elliptical orbit are established.

The optimal attitude maneuver control problem without control constraints is studied in the quaternion statement for an axially symmetric spacecraft as a rigid body under arbitrary boundary conditions on angular position and angular velocity of a spacecraft. The performance criterion is given by a functional combining the time and energy used for the attitude maneuver. Using substitutions of variables, the original problem is simplified (in terms of dynamic Euler equations) to the optimal slew problem for a rigid body with a spherical mass distribution. The simplified problem contains one additional scalar differential equation. In the class of generalized conical motions, the traditional optimal slew problem is modified to obtain analytical solutions for motion equations. The solutions contain arbitrary constants and two arbitrary scalar functions (generalized conical motion parameters). The proposed approach fits well with the Poinsot concept that any arbitrary angular motion of a rigid body around a fixed point can be considered as some generalized conical motion of a rigid body. Moreover, for the cases of analytic solvability of the classical problem of optimal reversal spherically symmetric spacecraft, when restrictions are imposed on the edge conditions of the problem (plane Euler turn, conical motion) solutions of the classical and modified tasks are completely the same. An optimization problem is formulated and solved with respect to these functions, the second derivatives of which serve as controls in the optimization problem. The resulting analytical solution of the modified problem can be treated as an approximate (quasioptimal) solution of the traditional optimal slew problem under arbitrary boundary conditions. The quasioptimal algorithm of the optimal turn of a spacecraft is given. Numerical example showing the closeness of the solutions of the traditional and modified optimal slew problems for an axially symmetric spacecraft is given.

The article provides an overview of designs of reflector antennas for spacecraft with solidstate precision dimensionally stable reflectors. The requirements for modern reflector antennas of spacecraft intended for the transmission of high-frequency radio signals are formulated. The features of using composite materials in manufacture of antenna reflectors in the context of their reflective capabilities are considered. The classification of reflector antennas is given. Features of the placement of antennas on a spacecraft are outlined. Multiple designs of reflectors of modern antennas are considered. Various ways of ensuring accuracy of shape and rigidity of reflectors are analyzed. Designs of antennas with a contoured service area and dichroic antennas are described. Antenna assemblies which are placed on a spacecraft and their advantages are considered. An overview of the engineering analyzes carried out during the antenna design phase is provided. The types of tests used to check the performance of spacecraft antennas are considered.

The topical problem of choosing the standard size of the unified space platform developed by the joint-stock company «Academician M. F. Reshetnev» Information Satellite Systems» for geostationary communication satellites has been formulated. Previously, the task of assessing the effectiveness of using the developed unified platforms for creating a new spacecraft was solved to a greater extent empirically, by generalizing the groundwork for existing developments and assessing the possibility and feasibility of using the existing groundwork in the future (development continuity). However, a methodological basis for solving this kind of problem has not yet been developed. As part of the study, a model for assessing the efficiency of a spacecraft based on a unified space platform was developed. A method for evaluating the effectiveness of the use of unified space platforms with an excess resource (in terms of mass and energy consumption) for the payload and a method for evaluating the effectiveness of using a modernized unified space platform, which makes it possible to meet increased resource requirements (in terms of weight and energy consumption) for the payload, have been developed. With the help of the developed methods, the assessment of the range of effective use of a unified space platform for geostationary communication satellites developed by JSC ISS was carried out, both without revision and with the necessary modernization for a specific project (payload).

The paper presents a method for minimizing the torque value of the multi-time deploymentfolding hinges. The objects of research were the hinges assembly of multi-time deployment-folding mechanical devices of solar array, which are used as part of the platforms of spacecraft. A computational analysis of the forces and moments that act in the hinges in the process of their opening and folding is carried out. The minimization of the excessive torque value without changing the design and layout of the hinge is possible by changing the shape of the dependence of the spring drive force on the angle of rotation of the hinge. The possibility of using constant force springs as part of the hinge is considered, the existing design algorithms are improved and a program for calculating the geometric and mechanical characteristics of constant force springs is developed. Experimental data were obtained for measuring the forces of springs of various configurations, showing the dependence of the magnitude of the force on the magnitude of the spring deformation and confirming the efficiency of the proposed calculation algorithm. A prototype of a hinge assembly with a constant force spring drive was developed and manufactured. The possibility of minimizing the excess torque without the need to change the design of the hinge while maintaining its technical characteristics was experimentally confirmed.

Due to the increasing complexity of radio electronic equipment, the problem of its reliability becomes much more acute. At the same time, the task of evaluating on-board equipment at the design and production stages becomes very urgent. An important issue is also the reliable determination of numerical values of its reliability indicators. Ground-based experimental testing of samples of on-board radio-electronic facilities, devices, assemblies and units facilitates the solution of this problem. Inthe course of this testing, various tests conducted. It been shown that the timeliest information about the quality and reliability of the equipment can be obtained from autonomous tests. In the article, the choice of an autonomous test facility has made. Amicroprocessor-based temperature controller chosen as an autonomous test object. The expert evaluation method used to select the object. Analysis of specifications carried out. The scope of autonomous tests was determined. An algorithm for autonomous testing of on-board radio electronic equipment of spacecraft was developed. It has properties of discreteness, determinacy, limb and mass. The development of the autonomous test algorithm takes more account of the specifics of the technology used to manufacture the product under test and the test sequence. One of the options for mathematical modelling of the process of selecting test equipment for autonomous tests been considered.

The paper discusses the issues of designing a control moment gyroscope electric drive with strict requirements in terms of the accuracy of ensuring a given rotation rate of the gyro motor suspension. A brief description of the control moment gyroscope electric drive applied currently is presented and the issues of improving the electric drive characteristics are discussed. As a solution, an electric drive is proposed which operates in the mode of feedback loop using angle sensors located on the axes of the gyroscope suspension and the engine rotor. The paper describes the arrangement of the control moment gyroscopes on advanced spacecraft for Earth remote sensing and presents the analytic expressions needed to calculate the control moments that affect the spacecraft. The moments are in the projection to the coordinate system brought into coincidence with the spacecraft. The paper compares spacecraft angular velocity stabilization errors for the cases of using the conventional scheme of control moment gyroscope electric drive and the newly developed one. The presented results can be used for developing control moment gyroscope electric drives to be mounted on spacecraft of different purpose with strict requirements on ensuring operation at specified rotational velocities.