Aviation and spacecraft engineering. Рубрика в журнале - Siberian Aerospace Journal
Inter-satellite optical communication link
Scientific article
A two-level system of data transmission in the optical range is considered between a low-orbit spacecraft located in a sun-synchronous orbit and a repeater satellite located in a geostationary orbit. This topic is rather relevant due to the fact that the rapid development of remote sensing satellites resulted in the increase of the amount of transmitted information, which in consequence introduced new requirements for communication systems. The increase of data transmission rate and severization of requirements for communication systems contributed to the development of one of the most promising areas of space communications, based on the information transmission via a laser channel, due to a high energy concentration and a much higher carrier frequency. The prospects for the application of optical communication systems are designated by lower power consumption, dimensional specifications and the mass of the transceiver equipment of the optical range (compared to radiofrequency range systems). The article describes the solution of application of optical communication link between a low-orbit spacecraft and a repeater satellite. The main factors that contribute to the attenuation in the process of signal propagation along the route are presented and analyzed. A model of a communication channel between a low-orbit spacecraft and a repeater satellite is provided for a visual image. Two different approaches of mutual guidance and tracking of laser terminals are described for using beacons and without ones. EDRS foreign system is considered as an analogue. The estimation of the main parameters of the communication link is given. The communication system considered in the article will allow for greater carrier capacity of the data transmission in the optical range between the low-orbit spacecraft and repeater satellite. The application of this system will allow solving problems, including in the interests of any departments and structures of the Ministry of Defense of the Russian Federation, for which the rate of obtaining information is one of the basic requirements for a satellite communication system. The tasks of precise targeting of receiving and transmitting devices arising as a result of narrow beam patterns can be solved with current technical means.
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The relevance of the study is due to the need to improve the reliability and durability of steam turbines, which are widely used in the energy industry. One of the critical factors affecting the performance of turbines is the occurrence and development of cracks in the blades, which can lead to their destruction and emergency situations. This type of defects can significantly change the dynamic characteristics of a structure, reducing its life and increasing the likelihood of failure. Therefore, analyzing the effect of cracks on the vibration parameters and strength of blades is an important task for predicting their reliability and developing diagnostic methods. This article examines the effect of cracks on the dynamic and strength characteristics of steam turbine blades. The object of the study is a working wheel made of 304 stainless steel. The finite element model in the ANSYS Workbench environment was used for the analysis. The natural frequencies and vibration mode of the blades at different angular speeds of rotation, as well as the effect of cracks of various lengths on the dynamic characteristics of the structure, are investigated. The analysis results show that the presence of a defect leads to a decrease in natural oscillation frequencies, especially for low-frequency forms of bending vibrations. In addition, the effect of crack growth on the durability of the blades and the entire working wheel has been studied. It has been found that increasing the crack length significantly reduces the life of the blades, and the durability of the working wheel decreases more slowly due to the interaction of the blades with each other. The obtained results can be used in the development of methods for diagnosing and predicting the life of turbomachines, as well as to optimize their design in order to increase operational reliability.
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High-precision and reliable inspection of thermal deformations is necessary in terms of simulating the effects of space in the ground-based experimental processing of antennas and mirror systems of spacecrafts. Inspection of objects up to 1.5 m in size is considered in the paper. In practice, it can reach sizes up to 10 m. Requirements for thermal deformation are in range of 10–200 micrometers. The deformable surface is rough (Ra » λoptic). The measurement error, however, should not exceed ± 1 micron. The electronic speckle pattern interferometry (ESPI) method is the most suitable for solving this problem. The method allows to inspection objects with a randomly inhomogeneous surface. The method assumes that it is necessary to calculate the wave phase values from the recorded picture by the digital matrix. It is the phase that contains information about the deformation, and the spatial phase shift method is used to calculate it. One of the measuring systems based on this method is the measuring system PulsESPI (Carl Zeiss Optotechnik GmbH production, Germany). It has a high sensitivity which is about 50 nm. However, this measuring system is designed for single measurements. In this regard, an additional software module for processing and visualization the result of a series of several hundred measurements has been developed. The experimental test bench with a test object has been developed to research the metrological characteristics of the PulsESPI system in accordance with thermal deformations measurements (multiple determinations). The PulsESPI system and the Renishaw XL-80 interferometer introduced into register of measuring instrumentation of Russian Federation were located on different sides of the object 1.5 m in size. As a result of measuring the surface displacement measured by the Renishaw XL-80 interferometer and its corresponding point from the PulsESPI system deformation map are compared. Three types of tests were carried out at the developed bench. The root-mean-square deviation of single measurements was no more than ± 0.2 μm. Error was no more than ± 1 μm when the series of measurements was conducted in which a total strain of 200 μm was obtained. The results obtained suggest the possibility of using this system for high-precision inspection of thermal deformations of large objects.
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The object of the study is a cylindrical mesh shell without a regular structure covering, made of carbon fiber composite material. A distinctive feature of this class of structures is the intersection of families of annular and spiral ribs. Mesh shells are used as power elements of spacecraft, therefore, when designing them, the main re-quirements are a reduction in the mass of the structure, high strength and stiffness characteristics. The re-duction of the shell mass is achieved by varying and selecting the structural and geometric parameters of the ribs. The article considers a set of mesh cylindrical structures of fixed mass. The authors have developed and presented an algorithm for calculating the number of elements of a regular rib structure and calculating the values of geometric parameters of elements of rib families. Two approaches to the formation of rib sys-tems are considered: by changing the heights or thicknesses of the rib structure. A macro has been developed for modeling parametric discrete models of such rib structures in the An-sys Mechanical APD software package. When constructing discrete grid models, a one-dimensional two-node finite element BEAM4 was used. The model was rigidly attached at the nodes along the lower edge, a load was applied to the nodes of the upper edge. Two types of loading were considered. The “non-flight” mode was determined by the axial loading of the shell evenly distributed along the upper edge. Flight mode – additionally took into account the applied moment. The displacement and deformation fields were calculated numerically in the ANSYS finite element package. The article presents the results of a study of the effect of the density of the rib structure on the stability of mesh shells under static axial loading, natural frequencies and waveforms. It is shown that with increasing density, the rib structure becomes thinner. At the same time, the critical load decreases, the values of the natural frequencies of the shells decrease, and the number of waves in the forms increases. The influence of approaches to the modeling of rib structures on the results of numerical calculations is noted.
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We have considered the issues of ensuring the resistance of high-voltage solar battery (SB) of spacecraft to the effects of secondary arc discharges. Research in this area has been going on for more than 50 years, but the answer to all the questions has not yet been found. First of all, this is due to the complexity of the electrophysical processes occurring on the surface of the spacecraft in space and in laboratory conditions. The second reason is the random nature of secondary vacuum arc discharges, which requires the use of special test methods to confirm the effectiveness and reliability of selected design and technological solutions. Tests in conditions close to full-scale conditions do not allow us to solve this problem. We have given a retrospective review of publications on the physical features of secondary arcs arising on SB of spacecraft, the mechanisms of their initiation, experimental research and testing methods. We paid considerable attention to the issues of the occurrence of secondary arc discharges SB of the spacecraft in the conditions of ionospheric plasma and plasma generated by electric propulsion thrusters. We have shown that despite the large amount of accumulated data and knowledge, the transition from low-voltage SB to high-voltage SB remains a difficult scientific and technical problem, which requires additional research to solve. In addition, it is already necessary to start training personnel who possess a wide range of knowledge and are able to work on this topic. To do this, it seems advisable to organize sectoral research, as well as the allocation of targeted funds for the training of highly qualified specialists and their independent research. This approach will make it possible to solve the problem of creating high-voltage SB in the shortest possible time and prepare personnel for the development of this technology.
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Scientific article
In this work, an analytical determination of the local heat transfer coefficient in the planes of rotation of gas turbines is carried out using an affine-like model for the distribution of temperature and dynamic spatial boundary layers with a convective component (at Pr < 1). The method of analytical study used in the work led to results close to the experimental values. The problem of determining the thickness of the energy loss is solved using the integral relation of the energy equation of the temperature spatial boundary layer, which makes it possible to integrate the necessary curvature over the surface. The law of heat transfer of the turbulent boundary layer for the rotational motion of the flow and motion according to the law of “solid body” is expressed. Equations are obtained for determining the local heat transfer coefficient by the Stanton criterion for various external flow laws for a power-law velocity distribution in the boundary layer according to the affine-like model of the temperature boundary layer. Heat transfer coefficients correlate with sufficient accuracy with experimental data and dependencies published by other authors: J. M. Owen, L. A. Dorfman, I. V. Shevchuk. The deviation of the results obtained from the dependence of the model with a convective component and with affinity-like profiles do not have statistically significant differences. The obtained results of the study and their comparison with the results of other authors showed that they are suitable for engineering calculations and analysis of the impact of local heat transfer coefficients on high-temperature units of a turbopump unit.
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Low-energy Earth – Moon – Earth flight trajectory design using optimization procedures
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The problem of designing a low-energy Earth Moon Earth spacecraft flight trajectory using optimization procedures is considered. The proposed approach combines rough and ready analytical methods with numerical population-based optimization techniques, that results in significant reduction in computational time compared to existing methods that require boundary value problem solution and numerical integration of differential equations. The proposed approach to design a spacecraft flight scheme utilizes spheres of influence method, which involves segmenting the trajectory into several sections. Each section is represented as an orbit defined by a conic section. The first segment of the trajectory is a geocentric orbit of the spacecraft flight to the Moon. The second segment of the trajectory is a lunar orbit of spacecraft flight within the Moon sphere of influence. The last segment represents is the trajectory of the spacecraft leaving the Moon and returning to Earth along a geocentric orbit. To ensure a passive lunar flyby and subsequent return to Earth without using additional impulsive maneuvers, the parameters of each trajectory must be determined by the initial conditions. To do this the optimization problem was formulated aimed at determining the trajectory initial parameters. The cost function is the criteria for minimizing the spacecraft’s closest approach distance to the Moon and the total flight time. By varying the weight coefficients in the cost function, various trajectory configurations can be formulated. The result of the optimization problem solutions is the initial parameters of the flight trajectory to the Moon from Earth orbit were selected, ensuring the spacecraft’s entry into the Moon sphere of influence and enabling its return to Earth without impulsive maneuvers. The results show the fundamental applicability of the proposed approach to designing lunar missions using a genetic algorithm.
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A mirror geometry control system for the Millimetron Observatory is being created to work as part of the on-board complex of scientific equipment. The system is designed to monitor the quality of the space telescope’s mirror system and use the data received as feedback signals for pre-setting and tuning the telescope’s optical system in outer space. The goal of the system is estimation of the multidimensional vector of unknown parameters of the telescope’s mirror system by indirect measurements obtained as a result of the measurement of the telescope by 3D scanning. A mathematical model has been created, numerically describing the process of pre-measurement of the mirror system of the Millimetron Observatory using optical control marks on the surface of the mirror system. The linear mathematical model allows to link the actual indirect measurements of the mirror system with the unknown biases of its parameters, determining the shape of the telescope. A formula has been developed for the optimal reverse problem solver in the process of pre-measurement of the mirror system. The method of measuring the components of the telescope as part of its pre-setting is described. The measurement of control marks is based on a onboard 3D scanner embedded in the design of the mirror system control system. The error analysis was carried out using the optimal solver, and a covariance matrix was obtained for the error vector of estimated parameter.
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In the conditions of continuous financing of the programs of the Ministry of defense of the Russian Federation, the question of finding the most effective ways to modernize weapons and military (special) equipment, the developments in which are maximum and the processes of their improvement can take no more than a few years, is particularly acute. Such products, in particular, include aviation artillery weapons (AAO), the prospects for the use of which remain for the entire period of the army's existence with conventional weapons. The main factor influencing the quality of the AAO functioning is considered to be the thermophysical loading of a small-caliber artillery barrel (hereinafter referred to as the barrel) during firing. The problem of increasing the accuracy of determining the temperature field of the barrel is again updated by tightening the conditions for striking targets. Issues closely related to the intensification of AAO application regimes have come to the fore. These are issues of heating, cooling, thermal strength, wear, barrel survivability, issues of safety and firing efficiency. Despite the methodological evidence of analytical and numerical approaches to formalizing heat transfer in the wellbore, their practical implementation is rather complicated. The physical and mathematical meaning of this reason is as follows: possible instability of solutions; manifestation of oscillations in areas of large gradients; simultaneous presence in the solution regions of supersonic, sonic and subsonic zones; the existence of laminar, turbulent flows and other non-linear formations; non-triviality of setting boundary conditions; the presence of thermal resistance of surfaces, etc. However, the practical needs of ensuring safety and increasing the efficiency of fire operation of AAO dictate the need to obtain a close approximation of the problem under consideration to its possibly existing exact analytical solution. The aim of the work is to improve the mathematical apparatus that simulates the temperature field of the shaft based on a combination of heat transfer methods and mathematical physics. By verifying the reliability of the developed mathematical model (hereinafter referred to as the model, if from the context of the presentation of the material it is clear that we are talking about the proposed tools), the facts of the absence of methodological errors in the formation of the constituent blocks of the model and the increase in the accuracy of determining the thermal loading of the wellbore by 9.4 % were established. Based on the accents of the stated problem, the directions for improving the model are argued.
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Mathematical modeling of the thermal condition of pressurized aircraft compartments
Scientific article
A method for determining the thermal state of the instrument pressurized compartments of the aircraft, based on the use of a mathematical model of the thermal state of the compartments, has been developed. The mathematical model of the air-conditioning compartment is represented by a system of equations of honeycomb thermally insulated skin, ordinary differential equations of convective heat transfer of the inner surface of the thermal insulation of the skin and compartment structures, on-board equipment, air and enthalpy transfer from the air conditioning system. The radiant exchange coefficient in the model is determined by the Monte Carlo method. With parametric identification of the parameters of compartments and the air conditioning system, methods developed for solving the direct and inverse problem of heat transfer and determining confidence intervals for parametric identification estimates. Confidence intervals for estimating the coefficients of the nonlinear mathematical model of the thermal state of the compartment are determined by the method of projecting the joint confidence region of estimates onto the coordinate axes of the coefficient space. The research is carried out in accordance with the Airworthiness Standards. The required characteristics of the air conditioning system and the thickness of the honeycomb thermal insulation of the compartment were obtained.
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Measurement of temperature distribution using a three-wire system of sensors based on thermistors
Scientific article
Improving the reliability and increasing the avionics resource is associated with possibility of continuous control of temperature fields of printed circuit boards. This problem can be solved only with the use of a large number of temperature sensors. It raises the problem of connecting the measuring elements and recording equipment. Several methods with their own advantages and disadvantages are proposed. One of the implemented and patented methods is using a set of resistive diode sensors installed in series on a threewire line. The temperature sensors are pairs of counter - connected diodes with a sequential survey when applying sawtooth voltage. The system is simple and easy to implement, but its main drawback is the method of determining the temperature by measuring the amplitude of the total reverse currents of diode pairs. It determines the large measurement errors, especially in the temperature range less than 20°C. The article deals with a similar design of a three-wire circuit, but with a fundamentally different approach to temperature measurement. The temperature sensor here is not diode pairs, but thermistors with a well-known dependence of resistance on temperature and high accuracy, and diode pairs record only the moment of coincidence of the sawtooth voltage with the voltage on the thermistors. This approach allows using mathematical methods of signal processing to accurately determine the voltage drop on the thermistor, and this ensures the accuracy of the resistance/temperature and the expansion of the temperature range. Given the fact that thermistors are increasingly used to measure temperature, simplifying their inclusion in a large number will allow to register the temperature field of electronic units, which is extremely important for spacecraft. The proposed version of a three-wire circuit for connecting temperature sensors at several points was tested experimentally, including at negative temperatures.
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Based on the previously compiled physical and mathematical model of the vortex ejector workflow, its solution was made. The solution of this model made it possible to compile two calculation methods: a method for calculating the optimal geometry of a vortex ejector for given thermodynamic characteristics and a method for calculating the thermodynamic characteristics of a vortex ejector with known geometric dimensions [1; 2]. Vortex ejectors are used in many areas of aerospace engineering. The compiled development of a method for calculating the thermodynamic characteristics of a vortex ejector with known geometric dimensions will make it possible to use vortex ejectors more widely in the aerospace industry. The calculation method is based on the concepts of tangential stresses arising in a viscous medium when two flows moving at different speeds interact. The mechanism of kinetic energy transfer from a high-energy gas to a low-energy one is shown.
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When designing rocket engines, the problem of providing the specified basic design parameters is solved. In connection with the increase in requirements for products of rocket and space technology, the requirements for ensuring the energy efficiency of all its constituent elements are also increasing. As a rule, the task of increasing the energy characteristics of a rocket engine is carried out by increasing the pressure in the combustion chamber and the rotational speed of the turbopump shaft. An increase in the rotational speed of the shaft of a turbopump unit requires the provision of a cavitation-free operation of the pump with the absence of cavitation breakdown. This problem can be solved in various ways: by constructive improvement of the pump or by increasing the pressure parameter at the pump inlet. However, too much increase in inlet pressure is not possible, as this will increase the thickness of the walls of the rocket's fuel tanks and a corresponding increase in the mass of the entire rocket. Turning on the screw, although it does not guarantee cavitation-free operation at any inlet pressure, is the most preferred method. The geometry of the bore part of both the screw prepump and the pump blades is designed to ensure non-cavitational operation. When designing, at the stage of experimental testing of pump modes, it is pos-sible to use the methods of computational fluid dynamics (Computational Fluid Dynamics, CFD). These methods are used in various areas of general engineering and have proven themselves well. However, the rocket motor pump has a high pressure drop with relatively small dimensions. The question arises of adapt-ing CFD methods to modeling cavitation tests. This work is aimed at deriving a function approximating the TPU test data set with a view to its further adaptation for CFD methods.
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This article formulates the topical problem of formalizing the methodological principles of the search for a compromise between properties repeatability and modification (novelty) for the new models of space technology – communication satellites. The main methodological principles of development continuity are the unification of products and their component parts, means of technological equipment and technological processes, which allows to reduce their diversity and nomenclature. The order of the nomenclature of the articles and their component parts is achieved by developing parametric and type-size series with rationally chosen intervals between the adjacent members of the series according to a complex criterion, a link to the target product performance with the cost of creating the product. The research develops a project model and defines criteria for selecting the size type of a universal space platform. In designing a new spacecraft on the basis of a unified space platform (USP), there is a need to refine it to meet the resource requirements of the new payload (mass and energy consumption). The article sets out the methodological principles for assessing the range of effective applications of the universal space platform for the two most extreme cases of resource requirements (by mass and energy consumption) of the payload: the resources of the platform are excessive or insufficient. Methodological principles have been developed to form a number of unified space platforms. Using the methodological principles for the formation of a series of unified space platforms, the effectiveness ranges of USP were evaluated and the completeness of a number of geostationary communication satellites developed by “ISS” was assessed.
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Methodological principles of space vehicle design for the maximum energy supply of the payload
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The design of spacecraft at the initial stages is carried out in the presence of uncertainties in terms of parameters and conditions. The determination of design parameters is performed step by step: determination of the nominal values of design parameters, normalization of resource reserves (mass, volume, energy consumption) according to the design parameters to parry uncertainties, designing spacecraft for marginal resources. The operation of spacecraft with an electrical load switched on includes several stages: launching into the target orbit, putting into regular operation, regular operation for the intended purpose, decommissioning from the intended use in case of emergencies. The power supply system is designed to provide uninterrupted autonomous power supply to the onboard equipment in all modes and at all stages during the period of active existence of spacecraft, taking into account the presence of shadow zones of the orbit from the Earth and the Moon. In this article, the methodological principles for designing spacecraft for the maximum power supply of payload in the presence of uncertainties in parameters and conditions are developed. The mathematical models for calculating the parameters of the energy balance of spacecraft have been developed for various options for realizing the power of the session load, depending on the level of illumination of the orbit and the period of operation of spacecraft. The effectiveness of using the methodological principles of designing spacecraft for the maximum power supply of the payload, depending on the level of illumination of the orbit and the period of operation of spacecraft, has been evaluated. A technique has been developed for rationing reserves by spacecraft energy resources to parry uncertainties in terms of parameters and conditions, as well as the principles of its application when designing spacecraft for maximum payload power supply.
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Methodology for assessing reliability of stand-bed systems in testing liquid throat engines
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In the process of design processing of low thrust rocket engines great attention is paid to special bench test methodologies, technical use of benches, simulation measurements of the physical conditions of outer space, as well as the use of diagnostic studies and equipment for various physical studies and measurements. The efficiency of ground (bench) testing is ensured by simulating the conditions of full-scale tests and taking into account the influence of all operational factors affecting the reliability of the assessment of reliability indicators during design testing in ground conditions. A special place in the issues of achieving test efficiency is occupied by the requirements to ensure the accuracy and reliability of test results. A significant amount of testing during the development of engines should be carried out under the required vacuum conditions on test benches equipped with pressure chambers with vacuum systems. As a result of failures of some elements of a complex bench system, the quality of functioning deteriorates and the probability of successful performance of the functions that determine the output effect of the system decreases. Therefore, the task of evaluating the reliability of the systems of the stand for firing tests of rocket engines is reduced to elucidating the effect of element failures on the quality of operation and the output effect of each system. When testing, the given conditions must unambiguously determine the technical characteristics of the test stand, including the pressure chamber and vacuum equipment. Tests must be carried out with a sufficient degree of certainty. When assessing the dynamic characteristics in pulsed modes, significant errors are introduced by inertial forces. Methods for ensuring the dynamic similarity of the characteristics of the engine supply systems with fuel components on the stand and as part of the propulsion system of the spacecraft, including the correspondence of the hydraulic, inertial and wave characteristics of the mains, are considered. An analysis of the errors in the test results was carried out. The tasks of the methodology for calculating instrumental errors are formulated. An assessment of the frequency characteristics of bench hydraulic lines was carried out. Recommendations have been developed to improve the accuracy of measuring parameters during bench firing tests of low-thrust rocket engines.
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This paper considers the methodology for calculating the de-weighting system of spacecraft elements for ground tests, taking into account the deployment options, de-weighting conditions, types and options of de-weighting systems. An example of calculation for a 3-section solar battery without a beam with incomplete de-weighting and with minimization of moments in the hinges is given. Genetic algorithms are used as an algorithm for determining the parameters of the de-weighting system, which allows obtaining the minimum moments in the hinges. The moments and forces acting in the system were checked by plotting diagrams in the expanded state. In addition, a check for compliance with the specified distance, based on design constraints, between the points of application of the weighting forces was made.
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Methodology of determination of balancing weights mounting places inside spacecraft compartments
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The paper presents a methodology for determining the mounting places of balancing weights inside spacecraft compartments, based on the use of a combination of methods of analytical and computational geometry, mathematical programming and computer graphics. The use of balancing weights is necessary to ensure the required position of the center of gravity of the compartment and the product as a whole. When using the methodology, the problems of ensuring a minimum mass of balancing weights and reducing the labor intensity of developing options for their installation are solved to speed up the preparation and approval of design documentation. The balancing weight placement zone is considered as a set of spatial regions free from compartment structural elements and other component parts. To minimize the overall mass of the balancing weights by determining their placement locations on a coordinate grid, the balancing weight placement problem is proposed to be represented as a linear programming problem. For testing, a conical compartment of a product with a spherical bottom was used as an example. It was determined that the balancing weight placement zone should be located near the junction of the bottom and the hull shell. The configuration of the placement zone was identified, taking into account the surrounding structural elements. The coordinates for placing the balancing weights were determined, and their masses were selected. Testing has showed the performance of the proposed methodology and the algorithm based on it. Effective use of the methodology is possible with the availability of a specialized calculation software package.
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Modeling a methane-hydrogen flame using a vortex burner
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The article devoted to the development of a mathematical model and the calculation of a methane-hydrogen flame generated by a vortex burner. The combustion products of hydrocarbon fuels, used in energy and transportation facilities, are the main source of greenhouse gas emissions, leading to increased ambient temperatures and global climate change. Therefore, there has been a recent focus on reducing carbon dioxide emissions from aircraft gas turbine engines and industrial gas turbines. The use of methane-hydrogen fuel can significantly reduce CO2 emissions, but it also leads to changes in combustion modes. There is an increase in flame temperature and propagation speed, which can lead to increased NOx emissions and burnout of installation elements. Therefore, when designing combustion devices and chambers, it is important to study the various combustion modes of methane-hydrogen flames in detail. Computational fluid dynamics methods are widely used to solve these problems, but mathematical models of combustion for methane-hydrogen fuels in relation to vortex flames are still not fully developed. To optimize the design and operation of burner devices, it is necessary to conduct complex mathematical modeling of aerodynamic, heat, and mass transfer processes and combustion. This article describes models for these processes, which were justified and selected based on previous research by the authors for different types of flames. It also presents a mathematical model for calculating swirling methane-hydrogen flames using the vortex-resolving large-eddy simulation (LES) model to describe turbulence. Additionally, the article discusses FGM combustion models with a kinetic reaction mechanism developed at the Institute of Chemical Kinetics and Combustion SB RAS, as well as a discrete ordinate radiation transfer model. A comparison of the calculation results with experimental data obtained by the German Aerospace Research and Technology Center (DLR) showed that the selected mathematical models of turbulent aerodynamics, heat and mass transfer, and chemical reaction processes, as well as the calculation algorithms, make it possible to simulate, with sufficient accuracy for engineering practice, the combustion of methane-hydrogen mixtures in swirling flows formed by vortex burners, which are widely used in the combustion chambers of gas turbines. The computational resources required for such calculations are reasonably acceptable when using available cluster systems.
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The main requirements for LRE gas generators are high stability of operation, ease of workflow management, as well as high efficiency of the generator gas. A particularly difficult task is to ensure the sustainability of the workflow. In addition to the probability of transverse vibrations in the in-chamber volume, due to the presence of additional volumes of various configurations and lengths attached to the reaction chamber, acoustic vibrations of complex longitudinal modes may occur. Most of the existing methods of testing a gas generator are criterion-empirical in nature and are based on the processing of experimental results, which does not always provide the required accuracy of calculating dynamic and thermal characteristics. The need for experimental and theoretical refinement of the calculation methods of thermodynamic processes of gas generators is an urgent task that will significantly reduce the material and time costs for preliminary design, testing and fine-tuning of modern models of engines and power plants of aircraft. Therefore, the calculation and analysis of the LRE gas generator is an important stage in the design and development of modern engine designs. Using the finite element method of the SOLID WORKS software package, a model of a two-zone gas generator for supercharging fuel tanks of the LRE was built. A study was conducted on modeling the workflow in a gas generator, visualization of thermodynamic processes in the product was built, numerical characteristics were obtained. The method of autonomous bench (firing) tests of fuel tank supercharger gas generators, the method of verification of numerical methods are considered.
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