Aviation and spacecraft engineering. Рубрика в журнале - Siberian Aerospace Journal
Статья научная
One of the main systems of a spacecraft (SC) is a power supply system (PSS). The basis of a PSS are secondary power sources (SPSs), which use various methods of controlling and converting electricity, which leads to significant differences in their dynamic properties. On the onboard consumer side, the dynamics of the PSS is determined by the total internal resistance (impedance) of the SPSs. When conducting ground electrical tests of spacecraft EES (electrical engineering systems), due to the complexity of using power supply systems, test complexes are used, the basis of which is simulators of electrical characteristics of the PSS (PSSS). Modern PSSSs use a modular configuration principle, which makes it possible to produce PSSSs of different powers, but energy modules have fixed or adjustable impedance frequency response over a narrow frequency range, which leads to the limitation of the types of PSS being simulated. Equipping the PSSS with the property of regulating frequency response in a wide frequency range expands the functionality of the PSSS, as it allows simulating the dynamic properties of the PSS containing different types of SPSs. The purpose of the work is to study and carry out the comparative analysis of three methods for regulating the impedance frequency response (IFR) of the PSSS module. The methods for regulating the frequency response of a PSSS are being considered on the basis of its generalized functional diagram containing mathematical models: adder amplifier (AA), feedforward compensator (FC), power amplifier (PA), voltage divider (VD) and load (L). The article analyzes options for regulating the impedance frequency response of PSSS, and considers three methods for regulating the IFR: two with a passive FC and one with an active FC. The paper presents a simulation model in the MicroCap package of the electrical circuit of the PSSS module, and computational experiments have been carried out on each method of regulating the IFR of the PSSS. Based on the results of the study, a method for correcting and regulating the IFR of the PSSS is recommended, which makes it possible to separately regulate the low-frequency and mid-frequency regions of the IFR, which allows us to significantly simplify the configuration and provision of the IFR of the PSSS in accordance with the specified requirements.
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Concentration of carbon dioxide in products of combustion of GTE NK-16ST and NK-16-18ST
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This paper considers the design of two combustion chambers of a gas turbine engine running on natural gas. One combustion chamber has 32 burners, and the other has 136 nozzles located in two rows in the flame tube head. A major contributor to global warming is considered to be the significant emissions of greenhouse gases, primarily CO2, including those emitted by gas turbine engines and power plants. The reduction of carbon dioxide levels by developing a set of structural measures in the combustion chamber is one of the urgent tasks of engine construction which requires a solution in order to meet modern environmental requirements for gas turbine engines serving as blower drives for gas compressor units. The presented research is dedicated to the analysis of influence of changes in combustion chamber design on reduction of CO2 level in exhaust gases of gas turbine engine NK-16ST. Two modifications of the combustion chamber are considered. The first one was a serial combustion chamber with diffusion combustion; the second one was a modernized combustion chamber with a modified front device. Each of the chambers considered was tested as part of the engine. During the study, combustion products were sampled directly in the exhaust tower and their concentrations, including the CO2 content, were determined. As a result of this work, it was confirmed that there is a possibility to reduce the concentration of CO2 in the engine combustion products up to 20 % without affecting the engine parameters. This reduction in carbon dioxide content was made possible by reducing the completeness of fuel combustion in the combustion chamber. The obtained data on changes in CO2 concentration can be useful in selecting the most suitable mode of engine operation, and the presented approaches to combustion processes organization can be used by developers in designing combustion chambers of natural gas-fired gas turbine engines.
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Control and regulation equipment of electric power system for a prospective piloted transport system
Статья научная
The aim of this work is to consider solving complex of tasks focused on fulfilling the complicated tactical and technical requirements for regulation and monitoring equipment (RME) of electric power supply system (EPS) for a prospective spacecraft. These requirements are imposed due to the need to ensure high reliability of the equipment during operation under the influence of external factors (vacuum, vibro-impact loads, radiation, absence of convective cooling), as well as to achieve high mass-dimensional parameters of the equipment and its high functionality The complexity of problem solving lies in the need to ensure conflicting requirements – high levels of energy density, weight and size characteristics, reliability and durability. These problems fully apply to the RME of the EPS for a prospective piloted transport system (PPTS) which design example shows ways of solving abovementioned problems. The most rational way of solving these contradictions is to increase the specific energy indicators of the main components of the RME devices – power converters, which can be achieved by using modern power electronic elements, using new materials and semi-finished products, for example, printed circuit boards with a metal heat sink, as well as increasing the layout density design. Determining solution is to select an optimal structure of the power converter, which provides the best efficiency. An additional way to reduce the mass-dimensional indicators of the RME is the use of a digital control method, the collection of telemetric information, and the receiving and processing of commands. At the same time, on the contrary, to ensure the specified reliability of the equipment, it is necessary to use excess reservation at the element level – for power components, and the principles of majority reservation at the functional block level – for control and telemetry schemes. Using the example of RME, developed by CJSC “Orbita”, the main EPS parameters of a new generation spacecraft are shown and most important power supply subsystems are considered in the article: the solar energy control subsystem and the power storage subsystem, ways to build them for meeting specified requirements, taking into account the proposed solutions. As a result of this work, the optimal structures of power converters – the current regulator of the solar battery and the current regulator of the battery – were selected, the basic principles of power components reservation ensuring the operability of the equipment in case of a single failure of any component without loss of performance and deterioration of RME parameters as a whole are shown. Block-modular construction method is used for optimal layout and high reliability of the RME, it ensures uniform heat removal from electronic components, which is especially important in vacuum conditions, minimum dimensions and mass optimization of the RME, as well as high mechanical strength of the structure. The implemented principles of building the RME for PPTS using this approach will allow to increase the active lifetime (ALT) and reliability of the spacecraft with a simultaneous decrease in mass and dimension parameters.
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Статья научная
Modern liquid rocket engines of low thrust (LRELT) represent complex engineering struc-tures, which are subject to very high requirements in terms of efficiency, reliability, and cost-effectiveness. To confirm the characteristics of the developed designs, a comprehensive set of tests for prototype samples is required, allowing their operability to be verified under conditions close to real-life operation. As part of this work, a thermodynamic calculation of the LRELT chamber for fuel components such as liquid kerosene and gaseous oxygen was conducted. The injector calculation method used in this work is based on the application of similarity criteria. This allows the transition from small-scale injectors to those suitable for full-scale testing, including stand tests using the “hydroflush” method. For testing, a specialized test rig was created, allowing the testing of injectors manufactured using modern additive technologies, such as 3D printing from polymer materials. This not only reduces the cost of creating prototypes but also accelerates the testing process. The injector tests on the stand play a crucial role in verifying their operability. This testing method allows studying the behavior of injectors in condi-tions as close to operational as possible. In this study, injectors manufactured using additive technologies from polymer plastic were used. The use of such materials in the early stages of testing helped to reduce costs and time resources for producing prototype samples. During the tests, the injectors were subjected to liquid at a specified pressure differential, which allowed their operability and fuel distribution uniformity to be assessed. The results of the tests demonstrated a high degree of correlation between theoretical calculations and actual data. The injectors showed stable operation corresponding to the calculated characteristics, and also proved their suitability for further development stages. The use of additive technologies in the manu-facturing of the injectors confirmed its effectiveness, allowing the prototype production cycle to be short-ened and costs reduced. Moreover, the “hydroflush” method proved to be a reliable means of verifying and validating the working characteristics of the injectors, which is an important step toward their implementa-tion in real-world operations. Thus, the proposed methodology, which includes the use of similarity criteria and additive technologies, significantly simplifies the process of development and testing, improves accuracy, and brings the results closer to real operating conditions. This is especially important for increasing the reliability and quality of final products used in rocket and space technology, contributing to a reduction in operational risks.
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Design of a flexible spoke for a spacecraft umbrella antenna
Статья научная
Umbrella type antennas are often used in modern spacecraft. Their advantage is the possibility of compact placement during orbital insertion. At the same time, they must provide the necessary stability when deployed in space. Antenna stiffness mainly depends on the stiffness of the spokes, the design of which is a complex task of scientific research. Spacecraft antennas must provide functional performance and, at the same time, have a minimum mass. The cardinal direction of improvement of space antennas consists in application of new structural materials. Composites are characterized by high specific mechanical properties, which allow to create structures with a high degree of weight perfection. The problem is related to the presence of a large number of design parameters that affect the performance of composite structures in a complex way. Determining the optimal combination of these parameters for each structure and a particular design case leads to the need for a complex numerical experiment based on specialized algorithms, methods and programs. The aim of the study is to design a composite spoke for the umbrella antenna of a spacecraft, providing the required load-bearing capacity and maximum stiffness at a given mass limit of the structure. It involves the development of finite element models of the composite spoke of various designs, which would include the possibility of optimizing the design parameters by the criteria of strength, load-bearing capacity and stiffness. As a result of numerical experiment, the ways of increasing the bearing capacity and stiffness of the deployed spacecraft antenna are determined.
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Design of a multifunctional electric propulsion subsystem of the spacecraft
Статья научная
A common way to form an electric propulsion subsystem of the spacecraft is to create specialized equipment or to select the most suitable one from the ready-made ones. However, there are cases when the use of existing equipment is not optimal enough and leads to an unjustified increase of the subsystem mass. Therefore, the ques-tion of creating a minimum equipment set possibility from which it would be possible to form propulsion subsys-tems in optimal way is of interest. The set of tasks, variants of use and possible schemes of placing orbital cor-recting propulsion on the spacecraft are presented. The list of necessary propulsion subsystem elements is pre-sented as follows: a thruster block, a tank, a xenon feed unit, a power processing unit consisting of a power unit and switching units, the complete set of cables and pipelines, the software and mechanical devices for control of the thrust vector (as an option). The necessary capacity of propellant tanks for the tasks of correction and rais-ing of the satellite to GEO with a high-pulse Hall thruster is defined: for orbit correction tasks – up to100 kg, for orbit correction and raising to GEO tasks – up to200 kg. Necessary angle rates of mechanical devices for con-trol of the thrust vector are defined taking into account possible schemes of placing thrusters on the spacecraft. It is shown that in cases when it is required to apply two or more thrusters to increase overall thrust, it is more preferable in the weight aspect to apply a combination of power and switching units instead of monoblock type of power processing units, and advantage can reach tens of kilograms. Provided the listed set of functional units is created, the offered concept will make it easy to form propulsion subsystems of the spacecraft for solving a wide range of tasks. It will reduce the time and money spent on creation of propulsion subsystem for new space-crafts.
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Design of the fastening assembly of the guys on the power spokes reflector
Статья научная
Currently, global communication systems are developing towards mastering high frequency bands for organizing high-speed information transmission channels, which requires large-sized antenna systems with reflectors up to 50 meters. Most of the technical solutions used for assembling large-sized reflectors are based on technological volumetric templates that geometrically completely imitate the necessary reflective surface of the reflector. The mass of such templates increases in cubic dependence on the increase in the diameter of the reflector, which is why it becomes more and more laborious to use them when assembling large-sized antennas due to the increase in the dimensions and weight of the templates. The purpose of the study is to design the attachment point for guy wires on the power spoke of the re-flector for a "templateless" assembly. The spoke is a composite isogrid structure on which arms are fixed for attaching power units. The fastening unit is an assembly unit consisting of a bracket and clips and al-lows you to precisely adjust the necessary pull tension force to fix the cord in the working position without the use of one-piece fastening methods. The analytical approach and finite element analysis were adopted as research methods. Using an ana-lytical calculation, the maximum tensile force of the guys in the designed unit was determined, thereby set-ting the maximum load for its operation. The coefficient of friction between the cord and the clamp in each individual case is determined experimentally. After simplifying the design and construction scheme of the bracket, the analytical calculation was carried out for a three-dimensional rod frame. To confirm the results of the calculation, a finite element model of the bracket was built and its static analysis was carried out. For the developed model of the bracket, the maximum stresses were determined and their comparative analysis was carried out with the results obtained analytically. With the help of solid modeling, the mass and overall characteristics of the braces fastening unit are determined. The limiting ranges of tension forces and the materials used in the knot, as well as its strength characteristics, were determined. This unit can be used in the "templateless" method of assembling a reflector for a wide range of large antennas, it has high manufacturability and versatility.
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Determination of the digital controller’s characteristics of the switched-mode power converters
Статья научная
The development of spacecrafts equipment is on the way to digitalization. In particular, energy spacecraft conversion devices are being modernized by introducing digital automatic control systems instead of analog ones. This leads to an increase in the efficiency of the power supply system, but at the same time, there is a need to create methods to determine characteristics that will confirm with a high degree of accuracy and conformity of the manufactured sample with the technical requirements specified during its design. The article describes the features of functioning and methodology for determining digital control channel of a pulse voltage converter’s characteristics. The proposed approach is a toolkit for verifying the correct implementation of both the hardware parts of the control channel and the controller, which is a program code implemented on digital control devices. The technique is based on determining the degree of responses correspondence to typical external influences of a hardware-implemented control channel and its model. Based on the transfer functions of the IIR and FIR digital filters, using standard built-in models, the control channel of the pulse voltage converter corresponding to the tested hardware-implemented device is simulated in the package Matlab Simulink. The basic principles of building the software architecture experiment are described. A block diagram of the test complex has been developed, including sources of external influence, control channel, and a test management tool (in this case, a personal computer). An example of applying such a technique to verify the parameters of the developed PID controller is given. Operability and accuracy of the proposed method to determine characteristics of the control channel by reaction to a sequence of rectangular pulses, and by constructing the AFCL are experimentally shown. Application of this verification method to production conditions will allow a complete check of individual central control units (CCU) of energy-converting equipment with closed feedbacks even at the stage of devices development, which will eliminate errors in the implementation of regulators in control loops.
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Статья научная
The main function of a thermal control system (TCS) is to maintain the temperature at nodal points of a spacecraft in given ranges due to redistribution of thermal energy and the discharge of excess thermal energy into space. TCS may have a different design and principle of operation. One of the most common options is TCS using a liquid circuit (LC) and pumping coolant circulation. In the development of promising design-layout schemes for instrument compartments of nonhermetic formation spacecraft, it becomes necessary to state and solve new problems associated with the creation of computational and mathematical models of intermediate convective heat transfer in a fluid circuit. For systems of integral equations of a LC thermal model with fairly complex topographic boundaries and connections, the justification and use of the defining (equivalent) thermal resistance seems to be a compromise of counting implementation of a system that simulates a TCS with integration along the length of the LC. In this paper, for the computational model of the liquid circuit of the thermal control system, including the system of equations of two-dimensional thermal balance of the characteristic surfaces of a nonhermetic formation spacecraft, a method of calculating the determining thermal resistances was proposed and implemented. This method includes the calculation of the complex heat transfer coefficient and the local heat transfer coefficient to the heat carrier flow. The approach considered in this paper allows us to obtain a numerical solution for the distribution of heat flows and temperatures of liquid circuits with complex topographic boundaries and connections with minimal loss of accuracy. The determination of the local heat transfer coefficient makes it possible to take into account the influence of changes in the temperature of the coolant flow on the overall picture of convective heat exchange.
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Developing the laboratory test bench of fuel three-point measurement
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The development of digital technology allows continuous improvements in many areas. This paper reflects the development of a new fuel measurement method. To measure the fuel, the authors propose three fuel sensors and a computational element to simulate the position of the fuel level in space with further calculating the volume of fuel, to reduce errors due to the fuel meters operation. The main advantage of this system is the elimination of errors arising from the evolution of an aircraft, as well as its uneven movement. The paper demonstrates a phased development of a laboratory test bench to study the three-point method to measure fuel. In the course of the work, a vessel is assembled to simulate the fuel tank of the aircraft. The vessel is a glass container with submersible measuring sensors. Also, the research contains calculation of the bridge electrical circuit to compute a voltage value at each sensor. In the test, transformer fluid substitutes fuel, since it acted as a dielectric. The program code for the microcontroller is recorded. The proposed method has several advantages in comparison with traditional methods of measuring the fuel level; a mathematical model is presented, on the basis of which the level of fuel in the aircraft fuel tank is measured.
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Статья научная
The paper presents the developed methodology and designed a device for determining the Mach number during supersonic gas outflow. An analysis of various methods for determining the Mach number was carried out, including measuring the pressure at the flow boundary, the use of shock waves, and the use of optical methods. A comparison of the accuracy of the readings when using the considered methods was made. Based on the results obtained, a technique for high-precision determination of the Mach number has been developed, including a combination of several independent measurement methods. A device has been designed that implements this measurement technique, and the results of experimental tests in a wind tunnel have been reviewed, including instrument readings, graphs and tables confirming the accuracy and reliability of the data obtained. Their accuracy and reliability are analyzed. Using the analysis, it is possible to ensure the selection of the most rational method for determining the Mach number at the initial stage of designing aircraft, such as airplanes, missiles, fighters, and UAVs. Accurate knowledge of the Mach number allows engineers to optimize the aerodynamic characteristics of the aircraft, ensure flight safety, improve engine efficiency and overall air transport performance. In addition, the Mach number is the most important criterion of similarity when modeling in aerodynamic research, which makes the developed methodology and device relevant not only for the design of aircraft, but also for a wide range of scientific and engineering research in the field of aeronautical technology. It is emphasized that the presence of a reliable method for determining the Mach number can significantly reduce the time and resources spent on testing and improving aircraft, and also contributes to the development of innovative technologies in the field of aviation and astronautics.
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Статья научная
Today, we can distinguish a number of promising reusable launch vehicles “SV Wing” – a reusable cruise stage of a light-class launch vehicle; “Baikal-Angara” reusable booster of the first stage of the An-gara launch vehicle; “Soyuz-7” is a reusable two-stage medium-class launch vehicle; flight design tests of “Soyuz-7” are planned for 2025. To maintain the operational characteristics of aircraft, it is necessary to develop a maintenance system that ensures the specified reliability of flying vehicle assemblies. The pur-pose of this work is to develop a model for detecting malfunctions in the process of maintenance of units and systems of aircraft. Within the framework of this work, an algorithm has been developed, which is based on the method of statistical testing, which allows, at low computer time, to analyze the maintenance process in more detail, taking into account the duration of separate operations and their effectiveness. Da-ta on the duration and efficiency of separate operations can be obtained in the process of special tests of equipment by timing and analysis of service results. For modeling it is necessary to have the following ini-tial data: the law of distribution of the duration of separate operations; the effectiveness of troubleshooting during separate operations. The algorithm implements two types of maintenance: full and reduced. Re-duced maintenance provides for operations that are most effective in terms of the number of faults to be eliminated: adjustments, regulations, search for faulty elements. The developed model makes it possible to investigate the possibility of reducing the downtime for maintenance without a significant decrease in the quality of maintenance, namely: to assess the effectiveness of maintenance when it is carried out according to the full and reduced scheme; evaluate the effectiveness of maintenance when performing maintenance in a limited time; justify the most appropriate ways to improve the quality of service, provided that the down-time for maintenance is limited and predict the likelihood of detecting malfunctions during the maintenance process. The practical significance of the results of this work can be achieved in the aerospace industry, in particular, at the design stage (testing and operation) of a maintenance system for reusable elements of launch vehicles.
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Статья научная
This article presents a detailed description of an experimental setup for testing two-phase thermal control systems of spacecraft. The setup is a climatic chamber for simulating real operating conditions of elements in the subzero temperature range and includes three circuits: a coolant pumping circuit corresponding in parameters to the thermal control system under study, a cooling system circuit, and a thermal load simulator circuit. Electric heating elements are used as a thermal load simulator. Transparent inserts provide the ability to visually monitor the structure of a two-phase flow during boiling and estimate the volume content of the vapor phase. A testing methodology was developed for the installation under consideration, including a testing algorithm and a description of the software and hardware testing tools. The software and hardware testing tools include electronic measuring tools for the main parameters of the operating modes of the two-phase thermal control system and an automated testing process control system. The developed automated testing process control system provides the ability to monitor a wide range of thermal and physical parameters of the coolant at various points in the circuit. The control system is based on the use of programmable logic controllers. To automate the operation of the installation, the OWEN PLC200 controller is used in the stand – a monoblock controller with discrete and analog inputs/outputs, designed to control and manage the operating modes of small systems. The software part of the algorithm is based on the CODESYS environment. The results obtained in the work can be used in planning and implementing programs for ground-based experimental testing of two-phase thermal control systems, in developing test stands for conducting research on spacecraft systems.
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Development of interface module emulator architecture for spacecraft life support systems
Статья научная
The article gives an analysis of special characteristics of ground-based experimental evaluation of on-board radioelectronic equipment, taking the control unit of up-to date spacecraft on-board control complex as the test objective. The focus is the problem of providing testing procedures of the specific software employed in design and manufacture process. A solution of the problem is worked out on the basis of performance of a hardware-software complex which emulates interface modules for the computing module of control unit. According to the general operation algorithm of the control unit, the developed complex is regarded as a multi-user system. The main functional requirements for hardware-software emulator, regarded as the corresponding queuing system, are also defined. The results of the experiments with the computer module operation prompted the requirements for the emulator response time from the point of view of its operation stability in real strict-time mode. In order to ensure the required efficiency of operation, the emulated functions of the interface modules are classified according to the severity level of their execution determinacy. The results of experimental evaluation оf the service channel hardware design variants when applying multi-functional reconfigurable input-output digital devices allowed to develop a hardware-software emulator structural circuit based on operation parallelism of programmable integrated logic circuits and flexibility of software reconfiguration. The realization of emulated functions of selected classes within the available architecture was carried out using the corresponding hardware blocks and software module. The presented analysis of the emulator response limits was performed with the application of National Instruments technologies. The results of the developed hardwaresoftware emulator evaluation and practical application, as well as other possible ways of applying the proposed approach for tests of spacecraft on-board radio-electronic equipment and space system components were also analyzed.
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Development of the CYCLOPS CubeSat payload
Статья научная
The number of CubeSat satellites launched has been increasing over the past decades. These satellites have a number of advantages: short development time, low cost, possibility of modifications for certain scientific tasks and testing of technical solutions and new developments. This article describes payloads of a small spacecraft: CubeSat 3U CYCLOPS designed by D.F.Ustinov Baltic State Technical University "Voenmeh" within grant under Space-Pi program. The purpose of the study is to create, test and study the performance of payload modules of the vehicle, built using commercially available components, under space flight conditions. The text describes the structure of interaction between the payload and the OrbCraft-Pro 3U platform from Sputnix LLC. The process of development of payload control board is con-sidered. The control system software for mechatronic and multiaxis actuator modules with logging and error correction is described. In addition to the above-mentioned modules, the payload control system was also developed to carry out a series of experiments in the presence of a small spacecraft in orbit. The paper explains how the spacecraft communicates with the ground via special software Houston control applica-tion and Houston Telnet. The results describe tests performed on the mechanical components of the space-craft. Examples of telemetry packets received from on-board the spacecraft are given. The article also re-flects further plans for the project and the prospects of using the developed hardware for implementation in large-scale space systems and complexes. Also as part of the project students were able to gain engineering experience in the development of devices designed to work in space conditions.
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Development of the concept of a reusable liquid rocket engine with three-component fuel
Статья научная
The article considers a promising direction for the development of liquid-propellant rocket engines (LPRE) – the use of three-component propulsion systems. The interest in this topic is based on a number of advantages that can be obtained by using this LPRE concept, namely: saving the mass of the launch vehicle (LV) by using a denser hydrocarbon fuel at the initial launch site; high specific impulse values at high-altitude launch sites due to the use of a more efficient pair of fuel components (FC): liquid oxygen + liquid hydrogen; reducing the cost of removing the payload, due to the use of a single propulsion system for both launch sites. An analytical review of implemented three-component LPRE schemes developed in Russia and abroad has been conducted, and their main advantages and disadvantages have been highlighted. Based on a detailed study of a number of circuit solutions for liquid-propellant rocket engines running on three-component fuel, the concept of a two-mode single-chamber three-component engine made according to a closed circuit with afterburning of generator gas is proposed. The oxidizer is liquid oxygen, the fuel is RG-1 kerosene and liquid hydrogen. In the first mode, the engine runs on three components, the share of liquid hydrogen in the fuel mixture is 4% of the total consumption of components. In the second mode, the engine runs on FC liquid oxygen + liquid hydrogen. The results of a computational and analytical study of the optimal design parameters of the engine are presented. The aim of the study was to understand the qualitative picture of the influence of various fuel parameters on the thermodynamic properties of the combustion products of the fuel mixture and the engine efficiency. Based on the results of the study, the optimal percentage of fuel components was determined. A mathematical model for calculating a three-component LPRE has been developed. The results of calculation of energy coupling are presented. A comparative analysis of the mass characteristics of the designed propulsion system is carried out.
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Development of the heat panel of the small space apparatus for navigation support
Статья научная
To clarify the trajectory of the spacecraft in a given orbit, the parameter of unmodeled acceleration is taken into account. Today, in the design and manufacture of a spacecraft to meet the requirements of the technical specifications for the maximum allowable values of unmodeled accelerations during the operation of on-board equipment, it is necessary to take into account the effects of asymmetric heat fluxes from the panels of the spacecraft on the deviation of its center of mass from a given orbit. This article discusses the problem of the influence of asymmetric heat fluxes from the surfaces of the spacecraft emanating from the panels ± Z, + Y (deterministic and non-deterministic component) on the level of unmodeled accelerations, which significantly affects the trajectory of the spacecraft. In order to meet the requirements for the temperature control system in terms of ensuring efficient heat removal from the on-board equipment devices and its distribution over the surface of the instrument installation panel, it is necessary to significantly improve the technical characteristics of heat transfer and heat conduction processes in the spacecraft. The analysis of the current thermal control system in modern satellites is carried out and its shortcomings are revealed. A constructive option is proposed for creating an energy-intensive thermal panel, which allows more efficient heat removal from devices and distribution over the panel. The designed thermal panel is a flat sealed panel of a single complex design of aluminum alloy, made by the additive technology method. The dimensions of the thermal panel are limited by the structural dimensions of the working area of 3D printers. At the moment, the main dimensions reach 600-800 mm. An increase in the working area in the future will enable the installation of large-sized electronic equipment. A two-dimensional mathematical model for calculating heat transfer processes in the designed thermal panel is presented. For the calculation, specific average values are introduced that characterize the effective cross sections for the vapor channels and the wick in the longitudinal and transverse directions, physical parameters (porosity of the wick and its degree of liquid saturation).
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Статья научная
The article provides a brief description of the flight scheme of a prospective automatic spacecraft intended for the study of Mars and its satellites by remote and contact methods. At the near-Martian expedition site, it is planned to first bring the vehicle into orbit of the artificial satellite Deimos, and then landing on Phobos with the subsequent delivery of its soil to Earth. The main ballistic characteristics of the spacecraft flight conditions at all stages of the flight at launch after 2025 are given. The time frames for the five starting periods are considered - in 2026, 2028, 2030, 2033 and 2035. The launch of the spacecraft on the flight path to Mars is performed by a heavy class launcher. The article describes the design of the vehicle, propulsion systems of its modules and flight scheme at all stages – from launch from the Earth to landing on Phobos, and returning back to Earth. The article describes the propulsion systems of the main spacecraft units proposed for the mission implementation – the propulsion module, the flight landing platform and the return vehicle. The designs of these units are provided in the work. Flight schemes have been developed in accordance with their characteristics, which allows conducting remote study of Deimos, making a soft landing on the surface of Phobos, and then delivering samples of its soil to Earth. The project should be developed on the basis of the spacecraft launch from the Vostochny launch site by the Angara- A5 launch vehicle and the KVTK upper stage. An alternative variant of the construction of a spacecraft involves the use of a vehicle of a lighter class - perspective Soyuz-5 launch vehicle and Fregat-SBU upper stage. In this case, the engine module is excluded, and the flight and landing module is replaced by a heavier version with larger tanks. Both proposed options for constructing a spacecraft make it possible to implement the developed trajectory, while ensuring full-time operation of the target equipment and conducting a set of experiments during a given period of active spacecraft existence.
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Dynamic flow parameters in natural curvilinear coordinates for a current line in a rotating channel
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A special interest in the topic of mathematical analysis of the flow of heat transfer processes is determined by the scientific significance and practical application in the development, design and production of rocket and space vehicles and installations. Substantiation of the developed techniques and modeling of the data obtained during the experiment using 3D process technologies gives an advantage. The accuracy and reliability of the calculation results play a key role in ensuring the safety and reliability of rocket and space systems. Regular verification and verification of the results are also necessary to ensure a high degree of reliability and safety. The comprehensive analysis of the fluid flow in the inter-vane channel of the impeller of a low-flow centrifugal pump presented in the article, with the construction of the energy characteristics of the impeller, can be used to clarify the number of vanes. The developed calculation method consists of four parts: firstly, an expression is obtained to determine the projection of the pressure gradient on the longitudinal axis φ, secondly, an expression is obtained to determine the projection of the pressure gradient on the transverse axis ψ, thirdly, the derivative of the longitudinal velocity in the transverse direction is determined, and fourthly, the results are presented numerical and experimental visualization: the power balance, the dependence of the pressure and the coefficient of influence of a finite number of vanes on the flow rate of a low-flow centrifugal pump. Based on the results of theoretical research, an algorithm and a calculation program were developed that allows calculating local values. The considered approach is confirmed by verification of the results of mathematical modeling by graphical visualization of the flow and measurement of the power balance of a low-flow centrifugal pump. The obtained expressions for pressure gradient projections, determination of the derivative of the longitudinal velocity and experimental visualization play an important role in the calculation and analysis of the operation of centrifugal pumps. However, there is a need for further elaboration of the method to bring it to a form that allows calculating the three-dimensional flow of the working fluid in an arbitrary channel.
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Статья научная
The paper investigates areas of dynamically unstabilized flows characteristic of the elements of the flowing parts of turbopump units of liquid propellant rocket engines; studies of rectangular variable cross-section, cylindrical variable cross-section, rotational currents in cavities with fixed walls, fixed and rotat-ing walls. The specific elements include: delivery and discharge units, side cavities between the rotor and the stator, cavities of hydrodynamic seals and elements of the interblade channel of centrifugal pumps and gas turbines. Due to the specific features of the operating and design parameters, the initial components of dynami-cally unstabilized flows are predominant in the flow parts of the delivery units. These areas have a signifi-cant impact on the energy parameters of the assembly and affect the heat exchange processes and, as a result, the reliability of structural elements. In the specific elements of the feed systems, both laminar and turbulent modes of the flow of the working medium are realized. Using the methods of three-dimensional boundary layer theory, specific thicknesses of boundary layer such as thickness of dynamic boundary layer, displacement thickness and momentum loss thickness are determined. Dependences for determination of flow core velocities, necessary for evaluation of losses due to the length of specific sections, are obtained. Proper selection of friction laws and velocities profiles in the boundary layer and consideration of initial section is necessary for the purposes of reliable determina-tion of energy parameters. Obtained dependences consider velocity distribution profile in the boundary layer on specific sections of laminar and turbulent regimes cases.
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