Aviation and spacecraft engineering. Рубрика в журнале - Siberian Aerospace Journal
Scientific article
The article presents the designs of several variants of combustion chambers for the NK-16ST engine. The combustion chamber designs are made within the framework of known concepts of lowemission fuel combustion to reduce nitrogen oxide emissions. A serial combustion chamber with a stepped air supply along the length of the flame tube is considered. A combustion chamber with a reduced volume of the flame tube and a twozone combustion chamber, as well as a combustion chamber with preliminary preparation of the fuel-air mixture in the burner device are considered. The features of burner devices of these combustion chambers are considered. For a serial combustion chamber, the burner device contains a nozzle, a swirler and a diffuser nozzle attachment. For a shortened combustion chamber, a confuser nozzle attachment is used. A two-zone combustion chamber contains two types of burner devices: diffusion for the duty zone and with preliminary mixing for the main zone. The combustion chamber with preliminary preparation of the fuel-air mixture contains a dual-circuit burner device with hollow blades and a central nozzle, which allows for regulation of fuel consumption along the circuits to ensure the required level of toxic emissions in the corresponding engine operating modes. The paper also provides a description of the bench equipment for purging burner devices, on which tests were conducted to determine their flow characteristics. The parameter of throughput, which is determined for each burner device, is derived. A description is given of the test bench equipment for purging combustion chambers, on which tests were carried out to determine pressure losses on the walls of combustion chamber fire tubes and total pressure losses in the combustion chamber under various inlet conditions for air velocity. A calculation and experimental method for determining the excess air coefficient at the outlet of the burner devices is presented. It is determined that in the combustion chambers under consideration, there is a tendency for the fuel-air mixture to become leaner at the outlet of the burner device. This ensures a reduction in nitrogen oxide emissions. Regulation of fuel consumption in the corresponding combustion chamber zones allows for an optimal level of emissions in various engine operating modes.
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The choice of the energy parameters of an oxygen-hydrogen propellant expander cycle rocket engine
Scientific article
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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Scientific article
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
Scientific article
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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The evolution of the concept of a “small” aircraft and the development of its design
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This study analyzes the design features of small, aircraft-type unmanned aerial vehicles. The first stage of the study examined the chronology of the evolution of the concept of “small”. It was shown that this concept should be understood as a combination of two factors: small size and low flight speeds. Because of this, the boundary layer flow occurs in a laminar-turbulent transition regime. Due to boundary layer instability, a “laminar bubble” forms above the wing, reducing the lift properties of the vehicle. Reynolds numbers in the range of Re = 105…106 were proposed as an aerodynamic criterion for smallness. At speeds of 70–150 km/h, the mean aerodynamic chord size is in the range of baverage = 0.1–0.9 m. The second stage analyzed the influence of the “square-cube” law on the energy efficiency of the vehicle during the miniaturization of its design. According to this law, the specific wing loading and the (weight)1.5/power ratio are reduced by the magnitude of the reduction factor for the linear dimensions of the aircraft. This is interpreted in the article as a reduction in the efficiency of the aircraft, which converts the kinetic energy of the airflow into lift. When transporting a unit of weight over a unit of distance, a small aircraft consumes several times more fuel than a full-size aircraft. A potential way to increase energy efficiency is to reduce or eliminate the trim losses of lift. This reduction is achieved by reducing the area of the rear-mounted horizontal tail, and the elimination of this loss is achieved by replacing the standard aerodynamic configuration with alternatives: canard or tailless configurations. These configurations do not use a rear-mounted horizontal tail, which generates negative lift, requiring additional energy to compensate. At the 3rd stage of the study, recommendations were formulated for the developers of the considered class of devices: 1) small-sized devices have a naturally low aerodynamic and energy efficiency, increasing which to the level of full-size devices is practically impossible; 2) improving aerodynamics is possible through the use of special profiles, morphing wing technology and boundary layer control methods; 3) increasing the duration and range of flight is possible due to the reduction of balancing losses of lift, using a beam tail in a normal configuration or using alternative aerodynamic configurations; 4) the article presents a block diagram of the iterative process for determining the wing area, the value of which will satisfy the conditions of the balance of forces and moments for various values of the ratio of the wing and horizontal tail areas.
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The feature of raising the “Express-AMU3” and “Express-AMU7” satellites into geostationary orbit
Scientific article
At present, in order to increase the launch mass, raising satellites into geostationary orbit by their own propulsion subsystem is widely used. Oly the JSC “Academician M. F. Reshetnev “Information Satellite Systems” applied such a scheme for several satellites of their own design - "Express-AM5", "Express-AM6", "Express-80" and "Express-103". Along with this, some diversity of approaches to the implementa-tion of this operation can be noted. In particular, the orbit raising of the above satellites was carried out using the onboard propulsion subsystem based on SPT-100 plasma thrusters. The operation was carried out by one or two thrusters. The use of two thrusters of the "Express-80" and"Express-103" satellites was due to the desire to keep within a reasonable amount of no more than six months with a significant increase in the output mass. Nevertheless, the duration of the orbit raising of about 150 days, which took place dur-ing the raising of satellite data, is also excessively long. It is evidemnt that it can be reduced, other things being equal, only by increasing the available thrust of the thrusters. This can be achieved both by increas-ing the thrust of individual units, and by increasing the number of simultaneously used thrusters. Therefore, for the new Express-AMU3 and Express-AMU7 satellites (with dimensions similar to the Express-80 and Express-103 satellites), for which a paired launch was also assumed, both of these methods were used. For orbit raising, two SPT-100V thrusters and, additionally, an SPT-140D type thruster were used. The total thrust of a cluster of thrusters made it possible to count on a significant reduction in the duration of orbit raising in comparison with the Express-80 and Express-103 satellites. The SPT-140 thruster developed by JSC "Experimental Design Bureau FAKEL" was used in Russia for the first time. For its power supply, the CCS-140D control and conversion device was specially created at the JSC "Design Bureau Polyus". The use of a combination of three thrusters made it possible to significantly reduce the duration of the opera-tion of raising into geostationary orbit.
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Scientific article
The burning of fuel in the combustion chamber of a gas turbine engine (GTE) is accompanied by formation of toxic substances. The most dangerous among them are carbon oxides that have a detrimental effect on humans and environment. In this regard the article is solving the urgent problem of determining the optimal method of gaseous fuel supplying in GTE combustion chamber to ensure low carbon-oxide emissions. The paper presents the design features of injectors that work with a separate supply of air and fuel. Natural gas is used as fuel. One of the considered injectors provides jet fuel supply by means of a perforated spray, and another one provides twisted fuel supply by means of a swirler built into the fuel channel. The main geometric parameters of the injectors are given as well, such as the size of the swirler, the number of blades, and the diameter of the output nozzle. In this regard the quality of air-fuel mixture preparation in a swirl jet in the outlet of the burner with two types of injector is defined. It is found that the best quality of mixing is ensured by the injector with jet spray. The design of a heat pipe simulator, in which the tested nozzle is placed, is considered. The design of a stand installation designed for testing injectors in a heat pipe simulator, as well as the modes under which these tests were carried out, are presented. The results were obtained in a heat pipe simulator with installed jet injectors and injectors with a swirling fuel jet. An analysis was conducted, which resulted in conclusions about the effectiveness of using jet injectors. According to the conducted research, the parameters of the injector with a swirling fuel jet are characterized by the presence of high values of CO levels in the combustion products, which is explained by the extremely low quality of mixing fuel with air and, consequently, low efficiency of fuel combustion. Jet fuel injection has low CO values, which indicates good quality of mixing fuel with air and high efficiency of a combustion process. As a result, we have received recommendations on setting the selected type of injectors in a full-size combustion chamber.
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Scientific article
In this paper, the issues of ensuring the resistance of the onboard equipment of spacecraft to the effects of ionizing radiation from outer space, which significantly limits the period of active existence of the space-craft, are considered. The paper describes the methodology for ensuring radiation resistance, developed by the specialists of JSC “ISS”. The result of the work done is to ensure the guaranteed performance of the target function by spacecraft with long period of active lifetime of 15 or more years. Among the outer space factors affecting the spacecraft, ionizing radiation of outer space is the main factor limiting the period of active existence. Exposure to energetic particles of ionizing radiation from outer space causes degradation of the electronic component base, which leads to failures and malfunctions of on-board equipment and degradation of its functional surfaces. Ensuring the radiation resistance of a spacecraft (SC) is a complex task, one of the stages of which is to determine the radiation resistance of components that complete the on-board equipment. As a result of accumulated experience in conducting radiation tests and analysis of the results, specialists of JSC “ISS” developed a methodology that allows to guarantee the radiation resistance of the spacecraft under conditions of tight production deadlines and optimized costs.
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Scientific article
A method of synthesizing a digital controller for a solar energy conversion channel in a power supply system of a spacecraft is presented. The method is based on the initial functional diagram of the pulse converter and the method of switching discontinuous functions. In accordance with the technique, which is formally presented in the form of eight consecutively executed items, a block diagram of the shunt converter has been developed in the basis of switching functions, which is taken as an example for testing the technique. The shunt converter is one of the three energy conversion channels in modern power supply systems of a spacecraft. The block diagram showed that all nonlinearity of the system can be reduced to nonlinearities of two multiplication links and nonlinearity of a pulse-width modulator. The possibility and acceptability of joint linearization of each of the specified nonlinear multipliers with a pulse-width modulator at the selected operating point is shown. A linearized block diagram of the control object was obtained, after which the transformation and simplification of the block diagram to a convenient form for calculation was carried out. Using the transfer functions of the linearized block diagram, the logarithmic frequency characteristics were calculated analytically and the results of their comparison with the frequency characteristics obtained experimentally on a simulation model, which confirmed their identity in the working frequency domain, were presented. At the same time, the specified simulation model of a shunt pulse converter, built in the Simulink package of the Matlab design environment, took into account all the mentioned nonlinearities of the real converter. According to the obtained logarithmic characteristics, a classical synthesis of the analogue prototype correcting section was produced. The transition from the analog correcting section of the prototype to the implementation of the digital correcting section is shown. Simulation modeling of a closed-loop power supply system with a synthesized analog controller, in its mode of operation from a solar battery, confirmed the correctness of the methodology and the achievement of the goals. The results of the work are intended to create a new onboard energy conversion equipment for power supply systems of high-potential spacecrafts. The scope of application of the project results is space instrumentation.
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The method of the disk friction determining of low mass flow centrifugal pumps
Scientific article
Low mass flow centrifugal pumps are currently widely used in the energy supply system of liquid rocket engines, the engines of correction, docks, consisting of on-Board power sources on-Board sources power supply system of fuel components in the in gas generator systems for inflating fuel tanks, and in temperature control systems of aircraft and spacecraft. When designing low mass flow centrifugal pumps for aerospace purposes, methods for calculating and optimizing the flow rate are often used corresponding to the design methods of full-size centrifugal pumps, which limits the mode and design potential of pumps and affects their energy characteristics and reliability. Reliability requirements often lead to the need to reserve units and fuel-supply systems. Despite the large amount of research works, the issues of reliable design of low mass flow centrifugal pumps with high energy and operational parameters for spacecraft and aircraft remains an urgent task. The article analyses the operational parameters of low mass flow centrifugal pumps used in aircraft and spacecraft power systems. Taking into account working fluid used and the temperature range, it was found that a laminar rotational flow with Reynolds number characteristic Re 103 3105 is realized in the lateral cavity between the impeller and the pump housing. The determination of power losses on disk friction of the impeller technique is developed taking into account design features and the applied schemes. Equations for determining the disk friction coefficients are consistent with the dependencies obtained by other authors. The obtained equations for the laminar rotational flow made it possible to determine the dependences for the resistance moment and the disk friction power of the impeller determining of a low mass flow centrifugal pump.
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The possibility of using methane-hydrogen fuel in converted gas turbine engines for power plants
Scientific article
Taking into account the fact that recently the topic of using methane-hydrogen mixtures as a fuel for gas turbine engines used in power plants has been actively developed, it is necessary to have engineering methods for calculating the fuel system and combustion chamber of engines operating on such fuel. The paper proposes the methodology that allows performing such calculations. A gas turbine unit (GTU) based on a converted aircraft engine NK-16ST was taken for the calculation. The calculation according to this method is carried out in three stages. At the first stage the composition is selected and the thermophysical characteristics of the gas under consideration are determined. At the second stage the fuel system is calculated, the consumption characteristics of the engine fuel system and the combustion chamber system are built. The consumption characteristics built for natural gas and for methane-hydrogen mixture are compared. The analysis makes it possible to develop recommendations for optimizing the design of the fuel supply equipment and fuel nozzles in terms of changing the volume of internal channels. At the third stage the combustion chamber is calculated and recommendations about the need to change the flame tube head or redistribute air along the flame tube length are made. The volumetric heat intensity parameter is used to estimate the sufficiency of the available volume of the flame tube for operation on methane-hydrogen mixture and to determine the gas average temperature in the combustion zone of the combustion chamber. The possibility of operation of the NK-16ST gas turbine unit on a methane-hydrogen mixture was confirmed on the basis of the results of the work performed. It was also concluded that in order to supply large volumes of methane-hydrogen mixture in comparison with natural gas, it is required to increase the size of fuel pipelines, metering and control units and fuel nozzles.
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Scientific article
A mathematical model describing the dynamics of a resonant converter with a T-shaped oscillatory circuit in the mode of power supply from a current source, which is a simplified model of a solar battery, is analyzed. It is shown that increasing the output filter capacity to improve the quality of transients is an ineffective measure. As an alternative measure, it is proposed to supplement the output circuit of the converter with an additional low-pass filter. A linearized equivalent circuit of the upgraded converter has been developed. It has been found that the quality of the transients of the input and output circuits of the converter is practically independent of the inertial properties of the converting T-shaped oscillatory circuit due to the fact that the time constant of the input and output filters is much greater than the constant of the rise time of the circuit current. It is shown that the device in question is a current-to-voltage converter not only in statics, but also in dynamics, and the equivalent circuit of the input circuit of the converter, without taking into account the input filtering capacity, is described by an electrical circuit that is dual to the output circuit of the converter. The developed equivalent circuit makes it possible to increase the speed of computer simulation of transient processes in the converter tenfold. The adequacy of the proposed approach is confirmed by modeling the transients of a resonant transducer in the Matlab 2021b software package. It has been found that the use of an additional low-frequency output filter significantly reduces the time of transients, reduces the amplitude of the input voltage transition, and also improves the mass-dimensional characteristics of the converter.
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Scientific article
Today, the scope of application of electric propulsion systems for orbit correction and spacecraft’s attitude control is rapidly expanding due to their high efficiency compared to liquid jet systems. The main elements of electric jet systems are plasma or ion thrusters. To ensure power supply of such thrusters, complex electronic power processing systems – power processing units (PPU) – are used. These units are capable to operate for a long time (up to 15 years or more) in a high vacuum environment and generate sufficiently high accelerating voltages – from 300 V and higher. PPU’s comprise various EEE-parts, mainly in the case design. As a rule, the technology of their production is such that air or nitrogen is initially located inside the housing at atmospheric pressure. During the operation of the unit, the non-absolute hermeticity causes pressure decrease inside EEE housings. Due to high voltages applied, this can lead to electrical breakdowns between current-carrying ele-ments inside the parts, their failure with the subsequent failure of the functional blocks of the unit. The paper considers the physical principles of the breakdown occurrence inside EEE-parts cases. The results of non-hermiticity measurements of several types of HV EEE-parts are presented. The dynamics of the pressure drop to the values dangerous from the point of view of breakdown event and the relevant occurrence duration are esti-mated. It is shown that duration of being exposed to the pressure-dangerous conditions can be as long as space-craft service lifetime. It can make difficult to use packaged gas-filled EEE-parts at the level of units intended to operate in non-pressurized compartments of spacecraft. Recommendations are provided for selecting the design of EEE parts with an operating voltage of about 300 V or more, as well as circuit solutions used to develop high-voltage equipment intended to operate in vacuum environment.
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Scientific article
The rapid development of electric-powered aircraft (with electric and hybrid power plants, including unmanned ones) is determined by the following advantages: reduction of harmful effects on the environment, reduction of noise exposure, reduction of maintenance costs. However, the development of such aircraft is hindered, in addition to organizational and legal difficulties, also by technical ones, the key of which are: low power-to-weight ratio of the aircraft (compared to piston and jet aircraft), electrification, underdeveloped infrastructure of airports and airfields. One of the options for solving the problem of increasing the power and energy efficiency of an unmanned aerial vehicle (UAV) with electric traction is the installation on board of a chemical-kinetic energy storage (rotating battery packs), with the possibility of using static electrification to recharge batteries. This system allows you to implement a continuous consumption / recovery cycle during the flight. The design was developed and a simplified assessment of the energy and technical characteristics of a promising UAV based on the selection of static electricity was given. The principle of operation of the static electricity extraction system based on rotating battery packs is shown, the main advantages over traditional systems of energy storage / recovery of aircraft are given. The results of the calculation and research of a prototype (a promising unmanned aerial vehicle) are presented, showing the versatility of the design, high operational and technical characteristics, as well as recommendations for the practical application of the developed UAV design. The results obtained are the initial stage of promising work to improve unmanned aerial vehicles and electric-powered aircraft systems. The low weight and size characteristics, as well as the unusual principle of operation of the static electricity extraction system based on a rotating battery packs allows it to be used as an autonomous power source for unmanned aerial vehicles, which allows in the future to extend this principle of energy consumption for any electric-powered aircraft. This work is an important stage in the development of highly efficient power plants operating on unconventional principles of energy production, which, on the one hand, expands the field of research of storage/ recovery/ consumption systems in modern aircraft construction, on the other hand, gives a powerful impetus to the development of engineering and scientific thought in the field of unmanned aircraft construction.
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Three-component aerodynamic load cells
Scientific article
The article examines the effect of flow on models studied in wind tunnels. To determine the force effect of the flow on the model under study, a more accurate and reliable method of directly measuring forces and moments using aerodynamic strain gauge balances is proposed. When solving a plane problem for a symmetrical model at zero slip angle, a design of three-component scales is proposed that measures the lift force, the drag force and the pitching moment. To eliminate the interaction between the supporting devices and the model, which causes disturbances in the flow near the model, the scales are located outside the model and the working part of the wind tunnel. The components of the aerodynamic force and moment acting on the model are measured using resistance strain gauges, which convert the deformation of the elastic element into a change in electrical resistance, which is measured by an instrument connected to an appropriate measuring circuit. The choice of strain gauges as weight elements is due to their very small size and weight, the ability to measure very small relative deformations of elastic elements, low inertia, which makes it possible to measure not only static but also dynamic loads, and the possibility of remote measurements. To compensate for the influence of various sources of errors, increase sensitivity and ensure greater measurement accuracy, the strain gauges are connected via a bridge circuit and included in all four arms of the bridge. Deformation of the horizontal measuring beam causes a change in resistance not only in the strain gauges that measure the pitching moment, but also in the strain gauges designed to measure the lift force. Since the design of the scales does not allow for electrical separation of these components, the influence of the pitching moment on the magnitude of the lift force is determined during the calibration process and is assessed using a special influence graph constructed from the results of calibration data. In strain gauge measurements, the output values of forces and moments acting on the model under test are obtained in the form of corresponding readings from a device that measures electrical signals proportional to the applied forces. To convert instrument data into values of forces and moments, a joint calibration of scales and instruments is carried out in order to obtain calibration coefficients. Additional components of aerodynamic forces and moments created by the holder are determined by purging it in the presence of the model. Calculated dependencies for determining the components of the aerodynamic impact are given. The values of the coefficients of aerodynamic forces and moments are given in the flow coordinate system. The pledge has been given.
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Scientific article
Ensuring high reliability of unique high-critical products at the design stage is an actual task that the aerospace industry faces. For high reliability indicators, at the design stage, it is necessary to ensure the basic property of the product – its strength, with a high probability of non-destruction. It is provided by introducing the corresponding coefficients – «safety coefficient» and «margin of safety» into the strength calculations. The necessity in these coefficients is based on the spread of values of external loading factors: magnitude of forces, combination of forces, kind of actions, place of connection, etc. In this case, the safety coefficient is related to external factors. The margin of safety refers to internal factors: the spread of the mechanical characteristics of the product material, the spread of the geometric dimensions of the product, etc. To determine, with a given probability, the safety coefficient and margin of safety, it is necessary to know their dependence on the combination of spread of external and internal factors. The purpose of this work is to determine the mathematical connection between the internal factors of the spread and the safety coefficient, external factors of the spread and the margin of safety, the combination of these factors and the probability of non-destruction of structures. In this work the values of internal and external factors, which affect the strength and probability of non-destruction of the product and have the boundaries of the spread of their values, using the tools of probability theories, were characterized as random variables, the values of which are determined by the distribution density, expected value and variance. I this work there was found a high dependence of the product strength on the spread of its geometric characteristics and tools were defined to determine the total spread of the values of the main strength characteristics of the product with a given probability of non-destruction. The practical significance of the results of this work can be achieved in the aerospace industry, in particular, at the design stage of unique high-critical products.
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To the question of forecasting the technical condition of low-thrust liquid rocket engines
Scientific article
In the rapidly developing space and rocket industry, spacecrafts are being equipped with low-thrust liquid rocket engines. Нigh requirements are imposed on the reliability, efficiency and economy of fuel use for this type of rocket engine. To ensure monitoring of the characteristics of spacecrafts, a functional diagnostic system is used, which includes telemetry and analytical data processing. Telemetry performs the functions of receiving and transmitting information. Information processing is carried out in computer centers located on the spacecraft and the Earth. The most promising computing tool capable of predicting time series and classifying a large amount of interconnected data is considered an artificial neural network. In this regard, the subject of research in the work is data processing methods based on an artificial neural network. The purpose of the work is to develop a method for forecasting the technical condition of low-thrust liquid rocket engines using an artificial neural network. The relevance of research on the use of a neural network in the system of functional diagnostics of low-thrust liquid rocket engines for spacecraft is explained in the introduction. In the main part, an analysis of many telemetric data of the rocket engine is carried out and their strength in the forecast of the main diagnostic parameters is determined. It is proposed to use traction, specific impulse, and temperature of the structure as diagnostic parameters. The prognostic capabilities of the neural network were investigated and a schematic diagram of a method for predicting the technical condition of a low-thrust liquid rocket engine was developed. In the developed method, at the first stage, the neural network performs the approximation of the function and extrapolates the time series of telemetric data; the second stage determines the probable class of the technical condition of the engine. The conclusion outlines a plan for further experimental research in the study area and provides recommendations on the development and improvement of algorithms for functioning of artificial neural networks as part of the functional diagnostics system of the spacecraft. Due to the generalized nature of the methodological schemes, the results of the work can be applied to any type of rocket engines and used at all enterprises of the rocket and space industry of the corresponding profile.
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Scientific article
Liquid gas generators (LGG) are additional firing units in the power system of liquid rocket engines (LPRE). The LGG ensure the operation of the power units of the turbopump unit (TPU) of the engine by feeding combustion products (CP) to the turbine drive. The main criteria for the efficiency of the generator gas is the complex (RT)gg and the thermodynamic properties of the mixture, depending on temperature, pressure, the degree of excess of the oxidizer and the enthalpy of the fuel, attributed to the conditions of supply to the nozzles of the GG. Changing the parameters of the generator gas leads to a change in the turbine power parameters due to its effect on the adiabatic operation of the Lad turbine. Depending on the engine circuit under consideration, CP GG can perform work in other units and elements of the engine, as well as influence many parameters of the LPRE. Among the main ones can be noted: the power of the booster gas turbine of the booster turbopump unit (BTPU) in the case of the selection of the generator gas after the GG or turbogas after the main turbine; the temperature of heating the refrigerant in the heat exchanger introduced in the GG; specific impulse of a liquid rocket propulsion system (LRPS), depending on the quantity and properties of the turbogas entering the exhaust pipe of the engine (for the engine circuit without afterburning the generator gas); mixing in the combustion chamber (CC) due to afterburning of turbogas entering the engine chamber after the turbine (for the engine circuit with afterburning of generator gas); parameters of the firing wall of the engine in the case of using a high-temperature gas curtain by blowing generator gas into the supersonic part of the nozzle. For many pairs of fuel during combustion in GG, the nonequilibrium of combustion products is characteristic (especially in hydrocarbon fuels).Due to the fact that the combustion products (CP) during the combustion of an oxygen-hydrogen mixture, due to the simplicity of the reaction, have time to form while staying in the GG (i.e., the time of chemical equilibrium of the CP is less than or equal to the time of stay in the GG), their thermodynamic parameters can be reliably determined using programs that simulate chemical equilibrium reactions. In this article, the issue of obtaining reliable results of thermodynamic calculations of generator gas at low and high coefficients of oxidant excess is investigated. Verification of parameters obtained in the programs “Astra” and “Rocket Propulsion Analysis” with calculated values was carried out. The most suitable program for performing engineering calculations and modeling the thermodynamics of liquid gas generators has been determined.
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Virtual imitation of a fueling panel for modern civil airplane
Scientific article
The article considers the design and development of a computer model of the fueling monitoring and control panel of the Sukhoi SuperJet 100 civil aircraft. The presented model is included in a simulator program for training technical specialists in aircraft maintenance skills under conditions of limited access to real or hardware-simulated equipment. In the process of designing of the presented simulation model, the refueling panel and the SSJ-100 aircraft refueling system sufficient components were considered and selected for further software implementation. The selection of the necessary components for the model was carried out using the decomposition method of the real system. First, the functional elements of the refueling panel itself were selected, after which the refueling system was disassembled into components that allow simulating the operation of the simulated panel. To implement the simulation model, software classes of objects and interactions between them were described. Software algorithms are implemented in the Unity environment using the C# language. The created program uses a three-dimensional graphic component and compiled for launching on a web browser. Software components have also been developed that allow studying the functions of the fueling panel both independently and in the mode of control of knowledge of the elements and algorithms for working with the fueling panel. The developed model is used as part of a practical simulator at the Reshetnev Siberian State University, and can be functionally expanded in the future.
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Scientific article
Recently, there has been an increase of interest in satellite orbit raising using electric propulsion subsys-tems. Theoretic analyses and practical experience demonstrate that while orbit raising to the geostationary orbit (GEO) via a transfer orbit is feasible, it requires a certain amount of time due to the thrust of onboard electric thrusters being low (40-300 mN) and thus incomparable with that of propulsion systems of liquid propellant thrusters (22-400 N). Due to low thrust, orbit raising by electric thrusters is time-consuming. However, the associated increase in mass to GEO may counterbalance the long duration of satellite commis-sioning. Calculations demonstrate a potential added satellite mass on GEO of up to several hundred kilo-grams with orbit raising duration of about 6 months. In particular, with satellite mass not exceeding 2500 kg, coupled launch is possible using existing launch vehicles. ISS took into consideration the positive results ob-tained with Express-AM5, and Express-AM6 satellites to design the Express-80 and Express-103 with orbit raising in mind. Such approach allowed for a coupled launch on the Proton-M carrier rocked with a Breeze-M upper stage, and a twofold launch cost saving. To increase thrust during orbit raising and decrease its du-ration, coupled thruster operation in high thrust mode was implemented. The resulting total mass on GEO increase constituted over 700 kilograms with maneuver duration of up to 158 days. This allows performing coupled launches of heavier satellites with orbit raising by means of electric propulsion in a feasible timeframe.
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