Aviation and spacecraft engineering. Рубрика в журнале - Siberian Aerospace Journal
Real-time carrier rocket mission control using space relay system
Статья научная
One of the purposes of the multifunctional space relay system “Luch” is telemetry provision for the timely control of launches from the “Vostochniy” spaceport. Launch vehicles and upper stage units have special high-speed and lowspeed relay user equipment for telemetry relaying using the relay system “Luch”. This article is about special programs for processing such kind of telemetry at the Mission Control Centre (MCC) of the federal state unitary enterprise "Central Research Institute for Machine Building" (TsNIIMASH) for mission launch control. These programs are the part of telemetry processing software-hardware system of the MCC. Their purpose is real-time reception, processing and depicting results of processing telemetry data from high-speed and low-speed relay user equipment. This article contains description of the telemetry structure, transmission scheme and description of telemetry reception and processing approach based on the specific characteristics of such kind of telemetry. It contains information about tasks, solved by the MCC telemetry complex for giving timely, objective and correct information about a launch process as well. Created programs, processing algorithms and representation forms of the results of telemetry processing successfully provided missions control of the launches of the spacecraft “Kanopus-V-IK”, “Meteor-M”№2-1 in 2017, “Kanopus-V” №3, 4 in 2018. We propose to use developed programs for telemetry
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Статья научная
This article presents some of the main parameters of metal mesh that affect the radio-reflective properties of reflectors and the technology for attaining these parameters by controlling the tension capacity of the mesh sheet at various stages of the radio-reflective surface manufacturing. The article also includes a brief overview of the developed methods (applied at the RESHETNEV JSC) of measuring and controlling the tension capacity of spacecraft reflectors with mesh shields, along with the analysis of the shortcomings revealed in the application of these methods. The authors introduce a new method of the tension capacity control based on the local deformation of the metal mesh by pressure of the air drawn through its surface. There are the results of development and testing of the metod, and of a model device for controlling the mesh tension capacity. The purpose of the research is to determine the operability of the device and the ways of its possible application in controlling the tension of the mesh-made reflectors ready for operation. The developed method and the model device for its application allow in-process control of the mesh tension at any spatial position of the mesh sheet. There is an analysis of possible application prospects of the method in rocket and space industry for manufacturing radio-reflecting surfaces of spacecraft antennas. According to the results of the research, the ratio of the mesh tension capacity to the pressure of the air drawn through the sheet was determined. The need for further improvement of the designed device aimed at increasing the measurement data accuracy was also pointed out.. In case the laboratory testing of the upgraded device for controlling the mesh tension capacity proves successful, further tests will be carried out in the process of sheet cutting and in the reflector shield manufacture.
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Статья научная
The article provides an overview of current methods for measuring and controlling the tension force of the material of the spacecraft metal-mesh reflectors. The purpose of the research: to obtain the most accu-rate results of measuring the mesh tension. Based on the review, a comparative analysis of the advantages and disadvantages of each of the considered methods was carried out, when measuring the tension force on mesh for large-sized transformable spacecraft reflectors. In the article considered the methods of force measurement: photo method, the operation principle of which consists in the sequential photographing of a mesh on reflector frame certain zones; a method based on pattern recognition, similar to the photo method, but using an information and measurement system with a given information processing algorithm; a meth-od based on local membrane deformation, related to the contact type, in which the mesh tension measure-ment is defined as a reaction from the impact of physical force on the mesh surface; a method based on the influence of sound waves impact, which is based on the influence of sound waves impact on the mesh sur-face. The experimental part of the research described in the article includes the measurement of the mesh tension by the resonance method, as the most optimal method according to the analysis. The prospects of possible use of the resonance method in the aerospace industry of radio-reflecting surfaces of spacecraft antennas are proposed and analyzed. According to the results of the conducted research, the dependence of tension force of mesh is established on the sound vibrations frequency affecting the mesh, at which reso-nance occurs.
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Research of the motion parameters of a spaceplane entry of into the atmosphere
Статья научная
After cessation of operation of the International Space Station in 2028, the Russian Federation plans to develop a national orbital station project. The Russian Space Station will differ from its predecessor in a greater practical aspect. One of the tasks assigned to the station will be the launch and management of a group of small satellites for remote sensing of the Earth, as well as the interaction and maintenance of prospective satellite groups. Due to the limited maneuverability of the orbital station and the potential for a malfunctioning device to be at a significant distance from it, the use of an autonomous spaceplane is proposed to increase the transportation and technical capabilities of the station. The research presents two aerodynamic designs of the spaceplane, and one of them is chosen based on the results of the aerodynamic and weight analysis. The spaceplane configuration and algorithms for its operation on the orbit and descent to the atmosphere are also presented. The goal of the research is to compare the trajectory parameters during the descent of the spacecraft from different descent orbits. For this purpose, a task was specified to determine the dependence of the area of the descent corridor on the initial parameters. The area of the descent corridor is determined by the boundary conditions, which depend on the operational parameters of the spaceplane. A computational program is written to solve differential equations of flight dynamics of a spaceplane by Euler's method in general and by Runge-Kutta method in a computational case. The results of the research are presented as the dependence of the area of the descent corridor on the altitude of descent. Graphical representations of the primary parameters of the spaceplane descent for the computational case are also provided.
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Results of monitoring the radiation environment in medium circular orbit
Статья научная
Problem definition – these data will form the basis for the development of technical solutions that will minimize mass, time and financial costs while ensuring the radiation resistance of on-board equipment and the spacecraft as a whole. Goal – the experimental dose control complex measures the level of absorbed ionizing space radiation doses in the sensitive element, assesses the radiation effects influence on the spacecraft, determines spacecraft’s residual radiation resource and refines impact models of the ionizing space radiation, located on an experimental spacecraft “Skif-D”, which was launched into orbit H=8070 km and inclination 90°. Results – flight experiment demonstrated high convergence of the comparative analysis’ results of the experimentally obtained impact levels in orbit of the operation of the “Skif-D” spacecraft with the impact model stated in the Russian Federation Scientific and Technical Documentation (OST134-1044-2007 amend.1 (2017) “Methods of the calculation of radiating conditions on-board of spacecrafts and specification of requirements for resistance of radio-electronic equipment of spacecrafts to the action of the charged particles from the space of natural origin”); Practical value – successful modernization of the ICDRM integral accumulated dose sensors in terms of their miniaturization and transition to a digital output (flight qualification of the sensors was obtained); the prospects of the concept of monitoring the integral accumulated radiation dose using semiconductor detectors with individual mass protection; experimental confirmation of a higher radiation exposure in the range of typical protections for ECB equal to 0.5–3 g/cm², on a 8000 km circular orbit compared to the GEO and 1500 km circular orbit.
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Статья научная
The aim of the work is the calculation and experimental substantiation of the expediency of using (both on Earth and on the surface of other planets) active-reactive type penetrator projectiles (SPART) for solving a number of scientific problems related to the formation of wells in the ground and the delivery of payloads to a certain depth. Research methods: various launch schemes (options for organizing the functioning process) of SPART are considered. The depth of penetration of an active-reactive type penetrator projectile into loam is calculated for the case when SPART is fired from a ballistic launcher located in such a way that the projectile exit velocity is equal to the velocity of its entry into the ground, and the thrust of the propulsion system is twice as great as the static resistance of the soil. From a variety of options, three SPART design schemes are selected depending on the combustion rate of the fuel used to ensure normal operation of the engine. As a result of the conducted calculation and experimental studies to determine the depth of penetration into loam of 152.4 mm penetrator projectiles 4.6 m long, launched from an artillery mount using the same powder charge weighing 18 kg, it was found that from the moment the engine is turned off until it comes to a complete stop, , which is more than twice the penetration depth of the same penetrator projectile if it moved in the soil only by inertia. Conclusion: the results presented in the article can be useful for researchers, graduate students and engineers involved in the creation and operation of aviation and rocket and space technology, and can also be useful for students of technical universities studying in the relevant specialties.
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Simulation of the mode of electron-beam welding of a thin-wall structure from AD31 alloy
Статья научная
The article contains the results obtained by the authors in the study of the possibility of using electron beam welding (EBW) for thin-walled structures made of AD31 aluminum alloy. Today, EBW of similar de-signs are not used due to the lack of technology. Currently, other technologies for connecting similar struc-tures are used in production, but they have a high cost, the reason for which is due to the high percentage of defects. The method of using EBW proposed by the authors will significantly improve the quality of the joint in thin-walled structures and the reproducibility of the technological mode of welding products. The authors have developed a technological solution to the presented problem, based on many years of experi-ence in the use of models of thermal processes, accompanied by electron beam welding. As a subject of research, modeling of the parameters of electron-beam welding of thin-walled pipes for waveguide paths made of aluminum alloy AD31 is proposed. The article presents the results of mathematical modeling of technological parameters during heating of an aluminum alloy by energy sources equivalent to an electron beam during EBW. The analysis and evaluation of the simulation results was carried out using the optimal-ity criterion developed by the authors. The calculations performed by the authors are based on functional using mathematical models of metal heating by a complex heat source consisting of moving instantaneous point and linear energy sources. The article presents the results of calculations for a plate with a thickness of 0.12 cm, which corresponds to prototypes used in the manufacture of waveguide paths. As a result, by changing such values as: beam current and welding speed, the temperature distribution on the surface of the product during the EBW process was obtained, which showed the applicability of modeling for develop-ing a new technological process.
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Small satellites for sea surface sounding
Статья научная
The paper presents a method of processing signals of radar sensing of the underlying surface using il-lumination from existing spacecraft (navigation, communication) and a constellation of small receiving spacecraft using synthetic aperture antennas. Methods and Results. Small spacecraft have many advantages over large satellites. Thus, they are rela-tively inexpensive to build, take minimal time from design to launch, can be easily modified to solve a particu-lar problem, and create less radio interference. The approach under consideration consists in redistribution of tasks to be solved between the constellation of satellites in orbit. Both regular high-orbit communication satellites and low-orbit satellite communication systems, as well as navigation satellites are represented as transmitter carriers (underlying surface illumination). These space systems use the necessary broadband sig-nal. Receivers of reflected signals are placed onboard small spacecrafts, and one of the tasks of the system is to perform research experiments, including on-line monitoring of fast-moving atmospheric cyclones. The work applies the method of sea surface radar imaging based on reflected signal models. The main results of the research are as follows: (1) possibility to use as a probing signal both a pulse and a broadband signal with a priori unknown modulation law, (2) acceptable resolution, (3) possibility to significantly reduce the system cost as compared to the existing space radars of sea surface survey. Conclusions. As a result of using a multi-position radar system, which uses small receiving antennas with synthetic aperture and sea surface illumination from operating spacecraft, it is possible to move to a qualitatively new level of solving problems of sea surface remote sensing with spatial resolution up to 1 meter, regardless of illumination and cloud cover presence.
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Solar thermal propulsion systems with various high-temperature power sources
Статья научная
The paper provides an overview of space thermal propulsion (STP) systems using concentrated solar energy as the main source of power. The paper considers solar thermal rocket engines of various configurations including those with afterburning of hydrogen heated in the “concentrator – absorber” system (CAS) with various oxidizers. Together with hydrogen the oxidizers form high-energy fuel compositions with a high value of ratio of components mass flow-rates which allows reducing the dimension of the CAS. The extreme dependences of the engine thrust on the specific impulse are shown for various values of the hydrogen heating temperature and the oxidizer-to-fuel ratio. The coefficients of the regression dependencies for the efficiency of a two-stage absorber and an absorber with the maximum non-isothermal heating having the highest possible energy efficiency are presented. The algorithms for calculating the main design parameters of the STP system as a part of a spacecraft (SC) are given, taking into account the ballistic parameters of the multi-turn transfer trajectory with multiple active segments applied to the STP systems having an energy-efficient non-isothermal CAS. The engine configurations with thermal heat accumulation and possible afterburning of heated hydrogen are also considered. Thermal accumulation allows accumulating energy in the solar-absorber during passive movement in the illuminated portions of the transfer orbits regardless of the lighting conditions of the apsidal orbit portions where the engine is turned on. Suitable heat-accumulating phase transition materials (HAM) such as the eutectic alloy of boron and silicon as well as refractory beryllium oxide are selected for different phases of the interorbital transfer to the geostationary Earth orbit (GEO). The main characteristics of different configurations of the STP systems in the problem of placing a spacecraft (SC) into high-energy GEO orbits are shown. A model of the SCSTP system operation is given taking into account ballistic parameters and the possibility of accumulating thermal energy. It is shown that the oxidizer-to-fuel ratio in STP systems with thermal energy storage (TES) increases with the decrease of the interorbital transfer time. The STP configurations with a two-stage TES showing a large energy-mass efficiency at moderate values of the solar concentrator accuracy parameter are considered.
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Spacecraft motion in a low circular orbit in establishing intersatellite link
Статья научная
The article investigates the problem of inter-satellite linking in the constellation of spacecraft in a low circular orbit. A specific problem of establishing intersatellite link (IL) in that orbit – cross-pointing of the antennae – is also studied. To support cross-tracking, it is important to place spacecraft (SC) in the orbital plane so that they are constantly in the zone of mutual visibility. The line-of-sight range is analyzed both in one orbital plane and between adjacent planes. IL is treated in terms of the orbital constellation (OC) ballistic formation. Several typical modes of motion of SC with IL in adjacent planes are determined – parallel, orthogonal, oncoming. The parameter values of IL antenna pointing are also assessed. The obtained results of OC formation and antenna pointing parameters’ calculations may be relevant for establishing a modified system.
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Specialized LED assembly for out-atmospheric solar simulator
Статья научная
Traditional solar simulators for thermal vacuum tests of spacecraft are based on gas-discharge lamps. Due to the characteristics of such lamps, they can only be installed outside the thermal vacuum chamber. High-efficiency LEDs can be installed directly in the thermal vacuum chamber, which can significantly improve the luminous and operational characteristics of solar simulators. Obtaining a spectrum close to the spectrum of the extraterrestrial Sun (AM0) is one of the primary and most difficult tasks in ensuring that the luminous characteristics of the solar simulator meet the requirements. The article considers a pre-viously proposed model of a combined emitter consisting of halogen lamps and assemblies of high-performance LEDs of various wavelengths. We have proposed a method for determining the spectral match for AM0 solar simulators and determined the requirements for LED assemblies intended for use in the combined emitter. Simulation with a sample of the most suitable commercially available LED assembly, at the nominal power level of halogen lamps, showed a good spectral match, which deteriorates significantly with decreasing lamp power. At the same time, many programs and methods of thermal vacuum tests re-quire simulation of different irradiance levels. Taking this into account, the authors developed an experi-mental LED assembly. Simulation of the combined emitter with this LED assembly showed the best results. The required spectral match is maintained at various irradiance levels. The achieved characteristics of the developed LED assembly are not limiting and can be improved by further optimization.
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Статья научная
This paper discusses and estimates the effect of some design parameters on the value of axial thrust appearing during functioning of the core component of a spacecraft’s (SC) thermal control subsystem – electric pump unit (EPU). The major causes of axial forces in centrifugal pumps of in-line arrangement are described and analysed. Design parameters having an effect of axial thrust value are: impeller position relatively to EPU diffuser (position was chosen based on dimension chain calculation), presence and size of discharging holes in the impeller, number and shape of impeller vanes (numbers of 14 & 16 were considered). EPU impellers with different number and shape of vanes were designed and manufactured. A series of experiments was carried out in order to research the effects of all aforementioned parameters: measurements of head vs flow curves and axial thrust values at given flow values. Each parameter’s contribution in the value of axial thrust appearing during EPU functioning is evaluated. Vibration measurements were obtained and analysed for electric motor DBE 63-25-6.3 fitted with different impellers. In this study, a DLP additive process was used for impellers manufacturing, which significantly sped up the tests. Obtained results will extend knowledge of processes taking place in EPU impellers, enable choice of the aforementioned parameters at design phase so to minimise axial thrust appearing during functioning of a centrifugal EPU of a spacecraft’s thermal control subsystem. Outcomes of this study are capable of improving SC reliability at all phases of its life because EPU axial thrust causes its premature loss of operability.
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Study of the payload extraction trajectory heavy class carrier rocket
Статья научная
As the weight and complexity of the payload that needs to be launched into orbit increases, the relevance of rational trajectory selection to ensure maximum efficiency and minimum costs for delivering the payload to a given orbit increases. Rational choice of the trajectory of a heavy-class launch vehicle has a number of important practical applications. Firstly, it allows you to increase the payload capacity of the launch vehicle and reduce the cost of delivering payload to the target orbit. This is especially important in the context of the development of the space industry, when more and more companies and organizations are showing interest in launching their own satellites and other spacecraft in conditions of fierce economic competition. Choosing a rational trajectory for launching a payload into orbit will significantly reduce the cost of launches and make them available to a wider range of potential customers. Secondly, the choice of launch vehicle trajectory parameters is important for ensuring safety and minimizing risks during spacecraft launches. Thanks to the rational choice of trajectory, it is possible to reduce adverse impacts on the environment and eliminate the possibility of emergency situations associated with loss of control over the flight of the launch vehicle. Rational selection of launch vehicle trajectory parameters is a complex task that requires comprehensive research and consideration of various factors, such as aerodynamic parameters of the atmosphere, mass and characteristics of the payload (spacecraft), engine operating parameters, characteristics of the target orbit, features of the launch of the launch vehicle and many other factors. A more thorough and systematic study of the influence of these parameters will significantly improve the efficiency and reliability of launching spacecraft into orbit. Thus, the choice of rational parameters for the launch vehicle trajectory is a relevant and important topic for scientific research. Increasing the rocket's payload capacity, reducing the cost of delivering a spacecraft to a given orbit, and ensuring launch safety are tasks that depend on the chosen shape and parameters of the rocket's trajectory. Such research has important practical significance and can become the basis for the development of new technologies and methods in the space industry. The purpose of the study is to study and select rational parameters for the trajectory of a heavy-class launch vehicle when launching a payload. The main task is to determine the flight path parameters that will allow achieving maximum efficiency and accuracy in delivering the payload to a given orbit. To achieve the goal of the study, the analysis of various factors influencing the launch parameters of the spacecraft is required, such as structural and aerodynamic characteristics of the rocket, the influence of aerodynamic factors and the Earth’s gravitational field on the flight path. Taking these factors into account, numerical calculations were carried out on the basis of a system of differential equations of motion using a computer program created in the MAPLE software package. Based on the calculations, modeling of the shape and parameters of the launch vehicle flight path was carried out. Research results. During the study, the rational parameters of the trajectory of a heavy-class launch vehicle were selected. The calculations were carried out using numerical modeling of the parameters of payload launch trajectories, and the analysis of the resulting trajectories was carried out. Minimizing the rocket's flight time was identified as the main criterion for the rational choice of a trajectory, which allows increasing launch efficiency and saving energy resources. An increase in payload mass and minimization of fuel consumption were adopted as additional criteria. Conclusion. The procedure for choosing rational parameters for the trajectory of a heavy-class launch vehicle proposed in this work will improve the delivery accuracy and reliability of spacecraft launches at the stage of ballistic analysis when designing rockets. The results of the study have practical significance for the development of future heavy-lift launch vehicle missions and improving the efficiency of space launches.
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Статья научная
Due to the constant tightening of flight safety requirements in the country and abroad, with the constant growth of air traffic, more and more requirements are imposed on the reliability, non-failure operation of air navigation systems and methods for updating their air navigation databases. Besides, there arises the problem of the relevance of the databases used in accordance with the AIRAC cycle, since the risk of emer-gency situations or disasters increases in the case of using non-updated aeronautical information in the flight management system (FMS), flight management computer system (FMCS), satellite navigation system (SNS), ground proximity warning system (GPWS). The paper proposes to consider the issues of improving aircraft navigation systems and updating databases using FMS-type computing systems. Russian aircraft use FMCS-95-1V, onboard ground proximity early warning systems (GPWS) and onboard satellite naviga-tion systems operating with an orbital satellite constellation (GPS, Glonass). All of them are equipped with aeronautical databases, which, in accordance with the AIRAC cycle, are updated on the ground by engi-neering and technical personnel every 28 days. The frequency of updating depends on the receipt of chang-es in navigation data for the operation of these systems. The paper considers the issues of operational characteristics analysis, methods of data transmission to onboard aircraft systems, development of an aer-onautical data transmission system, development of a remote transmission control system, as well as the development of data transmission algorithms, theoretical and experimental justification of the choice of a transmission system model. The use of the considered complex leads to a qualitatively new level of efficien-cy, reliability of updating air navigation databases in the FMS, SNS, GPWS, FMCS, which will affect the increase in flight safety, as well as the regularity of flights in the absence of aircraft downtime according to the criterion of operational updating databases under the AIRAC cycle.
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Studying static stability of a model rocket
Статья научная
Relevance. When designing flying models of rockets, one of the difficult tasks is to ensure the static stability of the rocket in flight along a given trajectory. Static stability refers to the ability of a model to return to an equilibrium position disturbed by external forces (wind, model asymmetry, etc.). In this case, the model must be stabilized in terms of the angle between the longitudinal axis of the model and the direction of flight (velocity vector), that is to maintain a zero angle of attack. The condition for ensuring the static stability of the rocket model is the location of its center of gravity ahead of the center of pressure. In this case, when the angle of attack is different from zero, the aerodynamic forces will create a stabilizing moment, which will return the model to a zero angle of attack. The purpose of the study is to develop and compare methods for determining the position of the center of pressure of a rocket and determining its static stability. The research considers a simplified method, an analytical calculation, a graphical method, and various practical approaches that can be used in rocket modeling. As research methods, an analytical approach, a graphical method and finite element modeling in the SolidWorks Flow Simulation program were used. In addition, a number of approximate calculation methods were considered. The study analyzes the capabilities of the considered methods for determining the static stability of a model rocket and the error of their application. To confirm the results of the calculation, a computer experiment was carried out in the form of blowing a finite element model of a rocket using the SolidWorks Flow Simulation program. The results of computer simulation confirmed the reliability of the proposed methods for determining the static stability of a model rocket. Research results. Simplified methods for determining the static stability of a rocket are the simplest and most reliable when designing model rockets. It is advisable to use it for launching demonstration rockets with an allowable misalignment error of 15% or more. Analytical methods are useful for designing sports models of rockets with high flight requirements, for example, for international competitions. Conclusion. The proposed method for ensuring the static stability of a model rocket makes it possible to simplify the design process of both demonstration and sports models of rockets for reliable demonstration launches.
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Testing of spacecraft orientation and stabilization systems using starry sky simulators
Статья научная
The paper investigates the need to create a method of simulating the starry sky for testing spacecraft and conducting tests of orientation and stabilization systems in laboratory conditions. Modern space exploration and, as a consequence, the complexity of technical requirements for flight support facilities are constantly increasing, respectively, the requirements for ensuring the accuracy of determining the position and orientation of the spacecraft are increasing. The history of the development of astroorientation devices and, in particular, stellar sensors is given. The modern stage of development of stellar sensors came with the advent of matrix radiation receivers: charged coupled device (CCD) and complementary metal-oxide semiconductor (CMOS) video matrices. Such stellar sensors are no longer tied to individual, predefined stars, but determine their orientation from images of groups of stars visible in the field of view of the device. Examples are given for their field of application, namely, determining the orientation of the sensor, pointing some device mounted on a spacecraft, and others. Modern requirements for astrogation are given. The basic principles of ground-based testing of the spacecraft orientation and stabilization system using starry sky simulators are considered. This is a stage of development and autonomous tests on a hardware and software stand of semi-natural modeling. To date, the ISS JSC enterprise has a complex modeling stand for conducting these types of spacecraft tests, using methods of both mathematical and semi-natural modeling, which includes various simulators of the starry sky. The development of these simulators has a long history, a comparative table of previously used simulators is given. The structures of both past and modern simulators of the starry sky are shown. The conclusions state the need to create a method that will simulate the rotation of the spacecraft at speeds up to 15–30 °/s. This method will allow testing the orientation and stabilization system of modern spacecraft.
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The aircraft hydraulic system units and pipelines heat exchange parameters study
Статья научная
The paper offers a method of mathematical modelling of aircraft hydraulic system thermal state. The given mathematical model presents a system of partial differential equations for carbon-fiber composite thermal insulation together with ordinary differential equations for hydraulic system components that describe their heat exchange with the ambient air and close-located surfaces. To solve the direct thermal state problem for hydraulic system components, i. e., to solve a stiff ordinary differential equation system, a Rosenbrock-type second order approximation numerical scheme for non-autonomous systems was applied. A solution of a partial differential equation system in Monte-Carlo method based on a probabilistic representation of the solution as a functional expectation of the diffusion process was also used. The inverse problem of the hydraulic system elements’ thermal state was solved applying a composition of the steepest descent method, Newton method and quasi-Newton method of Broydon-Fletcher-Goldfarb-Shanno. A mathematical model of the thermal state of a hydraulic system unit operating in an unpressurized aircraft compartment has been also developed, and the confidence intervals of each of the required model coefficients have been estimated using 2 1 α χ distribution at confidence probability = 0.95.
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The choice of the energy parameters of an oxygen-hydrogen propellant expander cycle rocket engine
Статья научная
In liquid-propellant rocket engines (LRE), made according to a gas-free scheme, the turbine of the turbopump unit (TPU) is driven by heated fuel in the coolant system of the combustion chamber (CC). The absence of a gas generator greatly increases the reliability of the LRE and provides a number of advantages over other engine schemes. At the moment, the existing oxygen-hydrogen gasless liquid-propellant rocket engines do not meet mod-ern tactical and technical requirements for the level of thrust and pressure in the (CC) engine. Therefore, it is necessary to study ways to increase the energy parameters of the liquid-propellant rocket engine and identify promising engine schemes. This article proposes schematic solutions for an oxygen-hydrogen LRE, provides an analysis of the influence of various factors on the power parameters of the engine, as well as recommendations for the design of gasless LRE. A mathematical model for calculating the main energy and geometric parameters of the engine has been developed. Prospective pneumohydraulic schemes of an oxygen-hydrogen gasless liquid-propellant rocket engine are proposed, depending on the tactical and technical requirements.
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Статья научная
The article proposes the concept of an experiment for conducting flight testing of various samples of photo voltaic converters. The purpose of the experiment is to study the behavior of new types of solar cells in outer space. The research will be carried out by testing samples on board the spacecraft. The experiment will be carried out by a device that collects data on the electrical properties of solar cells. The information re-ceived will be presented in the form of a voltage characteristic. During the experiment, its dependence on external parameters will also be studied. In particular, the correlation of the current-voltage characteristic from the values of temperature and illumination of solar cells will be investigated. Based on the data ob-tained, the efficiency of photo voltaic converters will be determined. Their degradation as a result of expo-sure to cosmic ionizing radiation will also be studied. The authors are tasked with designing and develop-ing an experimental installation that will be a payload module of a small CubeSat-class spacecraft. Based on the results of the work, the appearance of the flight testing experiment was developed, the re-quirements for the payload module were determined and a project for its creation was proposed. At this stage, the circuit design and software implementation of the module itself are under development. In the course of the work, the main requirements that this module provides to the main systems of the spacecraft were also formulated. To carry out the mission of the experiment, it is planned to integrate the payload module on the plat-form of the ReshUCube-2 form factor 3U. This satellite will be equipped with equipment enabling techno-logical experiments.
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The definite questions of simulation of transformable space structures dynamics
Статья научная
This paper describes large transformable space structures with various configuration in the folded transport position and in the open working one. As an example, simulation of transformable space structures dynamics is shown for the antenna circuit foldable load-bearing frame with diameter of 5 m. For investigation of the foldable frame deployment dynamics, a design scheme presented by a system of rigid bodies connected with each other by hinges is accepted as it is simple, but at the same time it considers features of the structure well enough. For performing stress analysis of the foldable frame elements during deployment, the frame shape at the certain time point of deployment, when relative velocities of adjacent elements are ultimate, is chosen. As a results of calculation using MSC.Adams software, positions, velocities and accelerations of the centres of mass of the foldable frame elements as well as the angular velocities and accelerations of the elements for each time step of the deployment are obtained. To perform stress analysis of the foldable load-bearing frame, finite element model of the frame is developed using MSC.Patran/Nastran software. As a results of investigation of stressed and deformed states of antenna circuit foldable frame elements both without taking into account damping and with consideration of damping, stresses arising in the foldable frame elements at the certain time points during deployment are found.
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