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The article presents the results of computational studies of aerodynamic characteristics for unmanned lift-generating multi-rotor drones of various configurations. The distinctive features of rotors flow were characterized. The rotor interaction was evaluated. The computations were based on the nonlinear rotor blade vortex theory in a non-stationary arrangement. The combinations of four, eight (four coaxial) and fourteen two-bladed rotors at velocity V = 100, 150, 200 km/h were considered. Semi-empirical methods were employed to select the rotor angles of attack, rotation speed, blade installation angles and geometric parameters at the given take-off weight for each combination of rotors and flight airspeed. The computations showed that for a four-rotor lift-generating design (quad-rotor), two rotors installed downstream, depending on the velocity due to the mutual effect, have values of the thrust coefficients ≈10...20% less than those of the rotors located upstream. For a coaxial quad-copter, the effect of the upper front rotor on the upper rear rotor is similar to the effect of the front rotors on the rear ones in a four-rotor lift-generating design. The effect of the upper front rotor on the lower rear rotor does not vary in terms of the average thrust value, and variations are only local in nature. The interaction of other rotors is identical to that of the four-rotor version. A fourteen-rotor lift-generating multi-rotor drone has a complex flow pattern, which generates deviance in the thrust coefficients variation with respect to time. Depending on the mode and rotors location, the average rotor thrust coefficient can vary approximately twice. The computations showed that with the similar geometric parameters and kinematics characteristics, rotors thrust is substantially subject to variation, which causes destabilizing moments to a significant degree without additional control input. Thrust pulsations and, respectively, vibrations grow in intensity as the flight airspeed increases. Probably, the right choice of the rotor configuration and the automatic control system can counterbalance thrust surge by so-called "phasing", i.e. selecting an initial azimuth angle for each rotor.

In the design of multi-engine aircraft, one of the important issues is the interaction between the propellers and airframe configuration components, especially in take-off and go-around procedure modes. Modern propeller-driven aircraft concepts in the pulling configuration are characterized by a high disk loading and an increased number of propeller blades used to increase cruising speed and reduce excessive noise. The first problem arising due to high disk loading is the direct impact of forces by operating propellers (thrust, normal force) on fixed-wing stability, especially at angles of attack different from a zero value. The second one involves a high-energy level of the propeller slipstream, having a significant indirect impact on the aircraft’s aerodynamics, stability and controllability. This impact is primarily associated with the interaction of propellers slipstream with other aircraft’s configuration elements. The complexity of taking into account the slipstream-wing interaction and other airframe components stipulated the application of experimental methods to study the problems of propellers – airframe interaction while designing propeller-driven aircraft configurations. This article presents an analysis of the experimental studies results of the operating propellers- airframe interaction for a light twin-engine transport aircraft. The aerodynamic aircraft’s configuration is executed using the conventional pattern of a high-wing and the carrier-on deck type empennage. The high-lift wing device is a fixed-vane doubleslotted flap. The wind-tunnel tests of the model in the cruising, takeoff and landing configurations were carried out in TsAGI lowspeed wind-tunnel T-102. Measurement of forces and moments, acting on the model, was performed by means of an external sixcomponent wind-tunnel balance. Measurement of forces and moments, acting on the propeller, was conducted using strain gauge weighers installed inside the engine nacelles of power plant simulators. The simultaneous combined use of external and internal balances allowed researchers to determine the direct and indirect contribution of operating propellers to the model longitudinal aerodynamic characteristics under variation of loading factor B ranging from 0 to 2.

Introduction. Computer-aided design systems for microwave devices are an effective tool for assessing the backscattering characteristics of complex-shaped objects. However, these calculations are often associated with significant computational costs, especially at large values of the ratio of the characteristic dimensions of the object to the wavelength. The use of asymptotic methods in combination with the mesh coarsening of object partition can significantly reduce these costs. However, in each practical case, this leads to a deterioration in the accuracy of the estimates obtained, which is hard to predict.Aim. Comparative assessment of the results of modeling the scattering field in the CAD of microwave devices using various methods for calculating and detailing the object model in the decimeter and centimeter wavelength ranges.Materials and methods. The research object was an anti-tank guided missile FGM-148 Javelin. The scattering field of Altair FEKO microwave devices was modeled in CAD using the methods of moments and physical optics in the frequency range from 1 to 10 GHz and angles from 0 to 180°. A comparison of one-dimensional backscatter diagrams and radar images obtained using these methods was carried out.Results. For the class of objects under consideration, the method of physical optics provides acceptable accuracy at frequencies of 5 GHz and higher with a step of partitioning the model surface of the order of one centimeter and a total calculation duration of the order of several minutes (Intel Core i5-4460 PC / 3.2 GHz / 8 MB RAM). At lower frequencies, acceptable accuracy and a similar calculation duration are achieved when calculating by the method of moments and a partitioning step of about 20 cm. The possibility of using the Altair FEKO CAD system for modeling radar images of objects with a resolution of at least 20 cm is demonstrated.Conclusion. The results obtained complement the well-known studies in the field of comparative assessment of the time and accuracy characteristics of various methods for calculating the scattering field of objects in the CAD of microwave devices.

The development of space transport systems for the delivery of payloads and the study of the lunar surface is an important scientific and technical challenge. The article considers a near-lunar space tether system consisting of a station and a microsatellite. The station is considered as a rigid body having a cylindrical shape, and the microsatellite is considered as a spherical rigid body. The tether is considered as a weightless inextensible rod of variable length. The station moves in a near-lunar orbit, which is influenced by the Earth's gravity. The process of deployment of a radially directed near-lunar tether system is considered. The equations of motion of the space tether system are obtained using Newton's second law and the theorem on the change in the angular momentum. To release the tether and bring the orbital tether system to a working state, the article proposes to use the control program of tethers tension force, which ensures the deployment of the tether system to a position close to the vertical. A comparison of the motion of the tether system along the unperturbed lunar orbit and along the perturbed one, taking into account the gravitational influence of the Earth, is made. To substantiate the theoretical results, a numerical simulation was carried out, based on the results of which a conclusion was made about the influence of the Earth's gravity on the amplitude of oscillations of the microsatellite relative to the local vertical.

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

Objective. The process of designing a long-range aircraft, in particular at the preliminary design stage, involves comparing and analyzing a large number of design alternatives with specified performance criteria. At the same time, the problem of choosing the composition of rational design parameters is solved, which is an actual problem of the preliminary design stage of a long-range aircraft. Methods. The developed method allows determining the vector of parameters that provides rational characteristics for a given efficiency criterion. In this regard, the authors propose to introduce a global efficiency criterion - the takeoff weight of the long-range aircraft, and the specific criteria are the aerodynamic quality in cruise and the value of fuel efficiency. Results. In this formulation of the design problem, it is necessary to solve several interrelated problems, some of which are formalized, while others do not yet have a mathematical construct and software that allows automating the process. A distinctive feature of the proposed approach to finding a rational design solution is the use of statistical analysis methods in combination with methods of high-precision mathematical modeling, software-implemented in a single information environment using the Fortran V and C++ languages. Conclusion. The proposed method for selecting the composition of rational design parameters allows forming the appearance and providing the specified characteristics for its component elements at the early stages and stages of designing the long-range aircraft.

This article presents software tools designed for construction of heterogeneous models by software integration of logical models having their work methods set in the knowledge bases and engineering models created in a technical simulation environment. In order to build logical models we use our software «Software-and-mathematical model of the command-andmeasurement system’s onboard equipment». Engineering models are built in the graphical programming environment of the LabVIEW. The software functions are described for creation of graphical schemes of models, data structures and telemetry data packages design, creation of the command base and setting of the methods of the model’s elements work in the knowledge bases’ rules. We provide an example of a heterogeneous model simulating the logics of the spacecraft’s command-and-measurement system’s onboard equipment operation including the library of virtual devices for the receiving and transmitting path simulation and the knowledge base with the rules of the onboard system’s interaction during reception, transmission, execution and confirmation of commands. This example demonstrates the points of integration of logical simulators and virtual devices. The model allows to assess the parameters of the receiving and transmitting path and the influence of destabling factors on communication line’s quality, simulate generation and transformation of the signal coming from the source of information – ground segment – to the onboard equipment and simulate logical interaction of the systems if the necessary signal level is provided. Each element of the model can be transformed and processed by the subject area’s specialist depending on the specifics of the solved tasks.

In 2014, the Eurasian Economic Union was created. The states of the Eurasian Economic Union have a unique geographical and geo-economic position, high transit potential. Transport systems in the space of the EAEU member states can ensure the functioning of reliable intercontinental transport communications between the countries of Europe and Asia. Activities are underway to implement a coordinated transport policy of the EAEU member states. The development policy of the transport sector of the economies of the member states of the Eurasian Economic Union is based on the use of the existing integration potential, the synergistic effect of the joint use of rail, road air and water transport.

Introduction.Precision elements of the target equipment and sensitive elements of the stabilization and orientation system of the advanced spacecraft are considered in the framework of this research. A method and software for modeling the dynamic characteristics of these elements are developed and validated. At that, the processing data results from the experimental studies on active and passive oscillators are taken into account.Materials and Methods. It is shown how the method of weightlessness provides simulation of the conditions that as much as possible conform to the real-time use of advanced space vehicles, precision structural elements, target equipment and their drives. Schemes of the corresponding experimental facilities are presented. Mathematical modeling methods, techniques of machine mechanics and dynamics are applied. Basic parameters of the proposed design dynamics, which are governing parameters in terms of the implementation of the target functions of the spacecraft, are calculated. Rational versions of layout and approximate cycle patterns of the operation of advanced space vehicles are formed to reduce microperturbations from driving gear with rotating masses.Research Results. A simulation technique for the dynamic characteristics of the drives of advanced space vehicles considering the regular oscillator operation is developed and validated. A complex of methods is presented for solving the problems of identifying dynamic parameters of a mathematical model of an advanced spacecraft based on the processing data results obtained through the experimental testing of active and passive oscillators. Two types of vibration from flywheel engines are noted. The first type is according to the commands of the position control and stabilization control system. The second type is due to residual imbalance from the solar constant meter. It is shown how these vibrations affect the dynamic characteristics of the gyro mounting seats and of the multispectral scanner for hydrometeorological support of the spacecraft. The data obtained are meant to solve the problems of assurance of the dynamic accuracy of advanced space vehicles.Discussion and Conclusions. A technique for modeling the dynamic characteristics of advanced space vehicles when operating in the precision orientation mode is proposed. The solution is based on the results of theoretical and experimental studies presented in the paper, and it considers the operation of standard oscillators. The implementation of this method is brought to software and algorithmic support for assessing the dynamic characteristics of standard oscillators of an advanced space vehicle. Recommendations to reduce the effect of active oscillators are established. Initial data are selected to determine the dynamics of advanced space vehicles from the point of view of fulfilling their target functions. The layout and approximate cycle patterns of the operation of advanced space vehicles to identify the driving gear with rotating masses as sources of micro-perturbations are proposed.

Показано применение прецедентов функционирования имитационной модели для анализа результатов испытаний бортовой аппаратуры космического аппарата. Прецеденты содержат различные сценарии передачи команд управления и соответствующие им варианты изменения параметров телеметрии бортовых систем космического аппарата. Сценарии описываются правилами базы знаний и моделируются процедурами логического вывода. Разработанные структуры данных и программное обеспечение позволяют выполнять имитационные эксперименты, сохранять результаты в базе прецедентов и сопоставлять их с результатами испытаний бортовых систем как в процессе выполнения испытательных процедур, так и ретроспективно. Применение результатов имитационного моделирования расширяет возможности испытательного программного обеспечения и способствует повышению качества конструкторских решений., This article represents a method for intellectual analysis of the results of testing spacecraft onboard equipment on the bases of the precedents of the simulation model. A simulation model is founded on a knowledge base describing different peculiar properties of the onboard equipment’s function, settings of the reception-transmission tract, scenarios of control commands’ transmission and corresponding changes of the parameters of onboard devices’ telemetry. We have designed data structures and software that allows conducting simulation experiments, save them in the precedent base and compare the results of the simulation modelling with the results of the onboard system’s testing. This analysis is carried out both during tests and after they are finished. In the first case, an onboard systems’ command is sent to the object of testing and to the simulation model. The model contains methods of logical inference that forms a conflict set of rules, choose and complete the applied actions simulating the functions of onboard equipment at reception and execution of the given commands. The results of modelling are represented in telemetry that is compared with the telemetry received from the objects of testing. A designer is given a list of parameters that were changed in the process of the logical inference and the telemetry parameters of the object of testing. In the other case, the telemetry of the onboard system obtained from the test storage results is compared with the precedents of the simulation model from the data base. Precedents contain examples of execution of big variety of commands and sets of the rules of the knowledge base that were completed for their acquisition. If telemetry parameters coincide, software allows a step-by-step review of the tests thus making a comparison with the actions of the simulation model. Comparison of the simulation model precedents with the results of testing allows revealing special features of the onboard equipment function that may remain unnoticed when other methods of analysis are utilized. Intellectual methods of logical output for analysis of tests extend the capabilities of test software and provide better quality of designer solutions., №3(24) (2019)

Представлена технология выращивания двурядника тонколистного на примере ООО «Феличита», специализирующаяся на его промышленном выращивании в средней полосе России. В Московской области в открытом грунте семена высевают с мая до середины августа, в Тульской области – с конца апреля до конца августа с интервалом 7 суток. Продуктивность колеблется от 1,9 до 8,4 г/растение, урожайность – от 0,5 до 1,8 т/га за каждый цикл выращивания, общая урожайность превышает 35 т/га., The agrotechnology of the perennial wall-rocket is presented on the example of LLC Felichita, specializing in its industrial cultivation in the central strip of Russia. In the Moscow region in the open ground seeds are sown from May to mid-August, Tula region - from the end of April to the end of August by an interval of 7 days. The productivity varies from 1.9 to 8.4 g/plant, the yield is from 0.5 to 1.8 t / ha for each growing cycle, the total yield exceeds 35 t/ha., №1 (2019)

The analysis of possible approaches to the construction of control systems for guided projectile of caliber 120(122)-mm is carried out. Autonomous, non-autonomous and combined control systems are considered.

This paper presents a review of the references on solar sails, which implies that this subject is relevant and actively developed by various researchers from around the globe. It points out that an important role is played by various imperfections, which would affect the parameters of real solar sails compared to the ideal model.The differential equilibrium equations of the blades of a heliogyro solar sail considering small deformations are written at the beginning of the article. The blade is loaded by light pressure and centrifugal forces, and is fixed at one end in the rotation axis while the other end is free. The differential equation for the blade deflection is obtained. This equation is solved in a statement that the reflection coefficient is independent on the mechanical stress. The deflection function has a simple analytical formulation. The differential equation for the transverse deflection of the solar sail blade is also solved analytically in the formulation of the linear dependence of the reflection coefficient on the tensile stress. The results obtained in both derivations are compared at the maximum amount of deflection of the blade of heliogyro solar sail. It is proved that the maximum difference in the magnitude of the deflection is achieved at tip section. An analytical expression for the relative change in the maximum deflection of the sail blades is given. The chart of this dependency is given for the blade of perspective solar sail with a length of 1000m and made from PET material, which is considered in the linear-elastic formulation.To calculate the projection of the resultant vector of light pressure on the axis of rotation the hypothesis is introduced that the decrease of the reflection coefficient of the material exactly equals to the increase in the transmittance, and vice versa. Considering this hypothesis for PET material the proportionality factor is found, showing a decrease of the reflection coefficient with increasing tensile stress.The expression for the infinitesimal light pressure force on the blades of the element of blade of a solar sail in the projection on the axis of rotation is given. This expression is numerically integrated by two formulations for the ideal and non-ideal solar sail. The numerical value characterizing the reduction of the projection of the main vector of the light pressure force for the non-ideal solar sail is obtained. By analyzing the references it is shown that the obtained difference in the light pressure is substantial and can lead to a change in ballistic performance of solar sails.

The propellant to fill the fuel tanks of the spacecraft, upper stages, and space rockets on technical and ground-based launch sites before fueling should be prepared to ensure many of its parameters, including temperature, in appropriate condition. Preparation of fuel temperature is arranged through heating and cooling the rocket propellants (RP) in the tanks of fueling equipment. Processes of RP temperature preparation are the most energy-intensive and timeconsuming ones, which require that a choice of sustainable technologies and modes of cooling (heating) RP provided by the ground-based equipment has been made through modeling of the RP [1] temperature preparation processes at the stage of design and operation of the groundbased fueling equipment.The RP temperature preparation in the tanks of the ground-based systems can be provided through the heat-exchangers built-in the internal space and being external with respect to the tank in which antifreeze, air or liquid nitrogen may be used as the heat transfer media. The papers [1-12], which note a promising use of the liquid nitrogen to cool PR, present schematic diagrams and modeling systems for the RP temperature preparation in the fueling equipment of the ground-based systems.We consider the RP temperature preparation using heat exchangers to be placed directly in RP tanks. Feeding the liquid nitrogen into heat exchanger with the antifreeze provides the cooling mode of PR while a heated air fed there does that of heating. The paper gives the systems of equations and results of modeling the processes of RP temperature preparation, and its estimated efficiency.The systems of equations of cooling and heating RP are derived on the assumption that the heat exchange between the fuel and the antifreeze, as well as between the storage tank and the environment is quasi-stationary.The paper presents calculation results of the fuel temperature in the tank, and coolant temperature in the heat exchanger, as well the heat flows and the relative amounts of the liquid nitrogen used to cool the fuel RG-1 as compared with other cooling technologies.The RP temperature preparation process using a heat exchanger, placed directly in the tank of the filling system is applicable for any high-boiling RP and has some of the best performance characteristics.Modeling the heating and cooling RP processes in the internal tanks of the ground-based systems using the numerical solution of the equations presented can be applied when calculating the RP temperature preparation processes with estimating their effectiveness and time of heating and cooling operations of RP.

We consider the optimal control problem for spacecraft motion during correction of its position in an inertial coordinate system by means of control torques. Control torques arise from the acceleration of inertial flywheels of a strapdown inertial navigation system. We investigate optimal control, which ensures a smooth change in the spacecraft orientation. This smooth corrective motion is described as the motion along the shortest path in the configuration space of a special orthogonal group SO(3) . The shortest path coincides with the large circle arc of the unit hypersphere S3 . We also consider a control algorithm using the original procedure of nonlinear spherical interpolation of quaternions. Four inertial flywheels are used as the main executive bodies for orientation of the dynamic control loop of the strapdown inertial navigation system when solving the optimal control problem. Three flywheels are oriented along the axes of the spacecraft. The fourth flywheel is oriented along the bisector. The simulation results are presented. We consider examples for corrective motion in which the spacecraft has zero velocity and acceleration at the beginning and end of the maneuver. We give an animation of the corrective movement of the spacecraft. The proposed formalism can be extended to control the spacecraft motion at an initial angular velocity different from zero, as well as in the orbital coordinate system. © 2018 Udmurt State University. All rights reserved.

Molecular dynamic methods that use ReaxFF force field allow one to obtain sufficiently good results in simulating large multicomponent chemically reactive systems. Here is represented an algorithm of searching optimal parameters of molecular-dynamic force field ReaxFF for arbitrary chemical systems and its implementation. The method is based on the multidimensional technique of global minimum search suggested by R.G. Strongin. It has good scalability useful for running on distributed parallel computers.

The paper concerns modeling of an uneven flow and electromagnetic field parameters in the combustion chamber of the liquid rocket engine with a near-wall layer available.The research objective was to evaluate quantitatively influence of changing model chamber mode of the liquid rocket engine on the electro-physical characteristics of the hydrocarbon fuel combustion by-products.The main method of research was based on development of a final element model of the flowing path of the rocket engine chamber and its adaptation to the boundary conditions.The paper presents a developed two-dimensional non-stationary mathematical model of electro-physical processes in the liquid rocket engine chamber using hydrocarbon fuel. The model takes into consideration the features of a gas-dynamic contour of the engine chamber and property of thermo-gas-dynamic characteristics of the ionized products of combustion of hydrocarbonic fuel. Distributions of magnetic field intensity and electric conductivity received and analyzed taking into account a low-temperature near-wall layer. Special attention is paid to comparison of obtained calculation values of the electric current, which is taken out from intrachamber space of the engine with earlier published data of other authors.

The advanced scientific and methodical apparatus for calculating the aerodynamic and aerostatic characteristics of a bound aerostat is considered with the aim of substantiating the basic requirements for the mobile balloon radar complex detection of low-level targets.

This paper presents the results of numerical investigations of the interaction with the Mars surface of four supersonic jets of ExoMars landing platform propulsion system. The cases of impingement of supersonic jets on a curved surface are considered depending on the values of propulsion system thrust. According to the results of numerical studies are obtained the values of normal stresses on the surface of Mars at altitudes of 1.0, 0.5 and 0.3 meter to the surface of the landing. To define the occurring shear stresses Mohr-Coulomb theory was used. The maximum values of shear stresses were defined for the following types of soil of Mars: drift material, crusty to cloddy material, blocky material, sand and Mojave Mars simulant. The conducted evaluations showed, regardless of the propulsion system configuration, that when the final stage of the controlled landing of the ExoMars landing platform, the erosion of the Mars regolith would be insignificant. The estimates are consistent with the available data from previous Mars missions.

Temperature preparation of liquid propellant components (LPC) before fueling the tanks of rocket and space technology is the one of the operations performed by ground technological complexes on cosmodromes. Refrigeration of high-boiling LPC is needed to increase its density and to create cold reserve for compensation of heat flows existing during fueling and prelaunch operations of space rockets.The method and results of simulation of LPC refrigeration in the recuperative heat exchangers with heat carrier which is refrigerated by-turn with liquid nitrogen sparging. The refrigerating system consists of two tanks (for the chilled coolant and LPC), LPC and heat carrier circulation loops with heat exchanger and system of heat carrier refrigeration in its tank with bubbler. Application of intermediate heat carrier between LPC and liquid nitrogen allows to avoid LPC crystallization on cold surfaces of the heat exchanger.Simulation of such systems performance is necessary to determine its basic design and functional parameters ensuring effective refrigerating of liquid propellant components, time and the amount of liquid nitrogen spent on refrigeration operation. Creating a simulator is quite complicated because of the need to take into consideration many different heat exchange processes occurring in the system. Also, to determine the influence of various parameters on occurring processes it is necessary to take into consideration the dependence of all heat exchange parameters on each other: heat emission coefficients, heat transfer coefficients, heat flow amounts, etc.The paper offers an overview of 10 references to foreign and Russian publications on separate issues and processes occurring in liquids refrigerating, including LPC refrigeration with liquid nitrogen. Concluded the need to define the LPC refrigerating conditions to minimize cost of liquid nitrogen. The experimental data presented in these publications is conformed with the application of the simulator and its parts.In the proposed algorithm of calculations a simulator consists of three blocks with data and intermediate results exchange with each other. Blocks contain subsimulator of heat gain in the LPC and heat carrier circulation loops, cooling-heating subsimulator of LPC and the heat carrier in the heat exchanger and heat exchange subsimulator of LPC and heat carrier in its tanks. Functional dependence of thermophysical properties of LPC and heat carrier and its temperatures, as well as characteristics of the pumps are taken into consideration in calculations. Pump heat gain subsimulator includes hydrauliccirculation loop calculation and the calculation of the heat flow emitted by pump. Subsimulator of the heat exchanger is based on differential equations of static heat exchange, where the integration variable is length coordinate of heat exchanger. The output data of these subsimulators are input data in heat exchange subsimulator of LPC and heat carrier in its tanks, which is based on the differential equations of quasi-stationary heat exchange.Simulation results are interpolated functions of LPC, heat carrier and the metal shells temperatures. Analyzing these functions graphs one can determine the performance indicators of the refrigerating system: refrigerating time and the relative costs of liquid nitrogen (mass of liquid nitrogen spent, referred to the weight of the refrigerated LPC). For example in the case of refrigerating portions 13260 kg and 3300 kg of high-boiling oxidizer with various combinations of design and functional parameters the refrigeration time is 2.8 ... 78.3 hours, and the relative costs of nitrogen amounted to 0,178 ... 0,322.Concluded the possibility of applying the proposed method for LPC refrigeration simulation with ground equipment systems using liquid nitrogen and intermediate heat carrier to determine rational design and functional parameters of the system, which allow to achieve the lowest cost of liquid nitrogen.

The article presents mathematical modeling of the workflows in the combustion chamber of the ramjet (RJ) model, taking into account the geometric characteristics of the latter, parameters of air, and gas generation products. It provides formation of the computational domain with the airflow supply in the afterburner chamber at the angle of 45° to the axis of the latter, and at 45o angle to the RJ plane of symmetry. Calculations are made on the clusters «HyperBlade T-ForgeHB-20", using the software complex HyperFLOW 3D», developed at Bauman Moscow State Technical University and based on the solution of systems of nonstationary two- and three-dimensional differential equations for gas (k-ε turbulence model in the Eulerian coordinate system) and the dispersed phase particles (Lagrangian coordinate system) using the parallel computing and automatic spatial partitioning calculated region. The total number of active nodes of computational cells is ~ 2.2 million; a size of the elementary volume is 5.2 x 1.8 x 1.8 mm. The k-phase available in the gasification products is represented by the polifractional distribution of fuel particles based on the alloy of aluminum and boron, their interaction with atmospheric oxygen being modeled via the integral reaction considering the heat proportional to the residence time of particles of each fraction. For the first time the ignition and combustion of boron fuel polifraction particles in the air stream are computationally studied at various initial values of operating condition of RJ model. The article examines peculiarities of movement and combustion of single-fraction particles in the afterburner chamber, which are a part of the polifraction mixture of condensed particles of gasification products. It also constructs the trajectories and presents the mass concentrations of a particles stream of single-fractions, temperature fields of the gas phase of velocity and oxygen concentration, determines the complete combustion of the fuel particles along the length of boron afterburner, depending on the coefficient of excess oxidant. The possibility for organizing an effective workflow in the gasifier and RJ afterburner has been proved.

The article presents the calculation-and-theory-based research results to examine the possibility for using the jet thrust impulse to increase a penetration depth of high-velocity penetrator modules. Such devices can be used for studies of Earth surface layer composition, and in the nearest future for other Solar system bodies too. Research equipment (sensors and different instruments) is housed inside a metal body of the penetrator with a sharpened nose that decreases drag force in soil. It was assumed, that this penetrator is additionally equipped with the pulse jet engine, which is fired at a certain stage of penetrator motion into target.The penetrator is considered as a rigid body of variable mass, which is subjected to drag force and reactive force applied at the moment the engine fires. A drag force was represented with a binomial empirical law, and penetrator nose part was considered to be conical. The jet thrust force was supposed to be constant during its application time. It was in accordance with assumption that mass flow and flow rate of solid propellant combustion products were constant. The amount of propellant in the penetrator was characterized by Tsiolkovsky number Z, which specifies the ratio between the fuel mass and the penetrator structure mass with no fuel.The system of equations to describe the penetrator dynamics was given in dimensionless form using the values aligned with penetration of an equivalent inert penetrator as the time and penetration depth scales. Penetration dynamics of penetrator represented in this form allowed to eliminate the influence of penetrator initial mass and its cross-section diameter on the solution results. The lack of such dependency is convenient for comparing the calculation results since they hold for penetrators of various initial masses and cross-sections.To calculate the penetration a lunar regolith was taken as a soil material. Calculations were carried out for initial velocities of interaction between the penetrator and the soil within the range of 250 and 1500 m/s with Tsiolkovsky number from 0.1 to 0.5. Obtained results show that there are optimal times when the jet thrust engine "switches on" and operates, thus providing a maximum increase in penetration depth. This time optimum is due to two competitive factors associated with the reactive projectile penetration. On the one hand, there is an additional reactive force that contributes to penetration depth increase. On the other one, due to fuel combustion, the penetrator mass decreases, thereby leading to its reduced penetration capability.It was shown that a value of Tsiolkovsky number has a significant influence on the motion of penetrator using a jet engine. With raising Z, a penetration depth increases as well. At initial velocity of 500 m/s and optimal time parameters of reactive pulse, penetration depth increases almost by 40% for Z = 0,1, 90 % for Z = 0,25, and 2.5 times for Z = 0,5. As initial velocity of the penetrator grows, effectiveness of additional reactive acceleration significantly decreases. This is due to decreased relative portion of chemical energy of rocket propellant combustion as compared to the initial kinetic energy of penetrator with its reducing velocity.A conclusion based on research results was drawn up that a penetrator module under examination equipped with the pulse jet engine is an efficient facility for the significant increase of penetration depth in low-strength soil targets. It was emphasized that the maximum increase in penetration depth was reached when a running jet engine makes the penetrator to move in the target rather than at its prestart (before coming in contact with the target).

The purpose of this paper is to study and find the ways to improve the reliability of onboard laser ranging system (LRS) used to control the spacecraft rendezvous and descent. The onboard LRS can be implemented with optical-mechanical scanner and without it. The paper analyses the key factors, which influence on the reliability of both LRS. Reliability of LRS is pretty much defined by the reliability of the laser source and its radiation mode. Solid-state diode-pumped lasers are primarily used as a radiation source. The radiation mode, which is defined by requirements for measurement errors of range and speed affect their reliability. The basic assumption is that the resource of solid state lasers is determined by the number pulses of pumping diodes. The paper investigates the influence of radiation mode of solid-state laser on the reliability function when measuring a passive spacecraft rendezvous dosing velocity using a differential method. With the measurement error, respectively, 10 m for range and 0.6 m/s for velocity a reliability function of 0.99 has been achieved. Reducing the measurement error of velocity to 0.5 m/s either results in reduced reliability function

The paper considers a maneuver for deorbiting the large space debris using an active spacecraft connected with the debris by an elastic tether. Tether slacking during the maneuver can lead to the tether rupture, kinking, and winding on the descending object. Therefore it is important to prevent slacking. The objective of this work is to find the law of thrust force control of the active spacecraft to ensure a continuously strained tether during the maneuver.Using Lagrange formalism a mathematical model to describe the system plane motion is developed. This model considers the active spacecraft as a mass point, the space debris as a rigid body, and the tether as a weightless elastic rod. A thrust force is directed along the local horizon of the spacecraft. Linearization of nonlinear differential equation describing longitudinal oscillations of the tether length is performed. Its phase portrait is analyzed. An approximate expression describing the position of the center on the phase portrait is obtained. A time-optimal control with full feedback to ensure that the tether is in the strained state is found by solving the Bellman equation. To use the obtained optimal law it is necessary to set the measuring equipment on the spacecraft, which is capable of accurate measuring a distance to the space debris and its relative velocity. An alternative control law, which is simpler in terms of the practical implementation, is proposed. As an example, the descent from an orbit of nonfunctioning Soviet satellite Meteor-2 is considered. It is shown that both proposed laws provide continuous strain of the tether during deorbiting of the satellite. Moreover, slack does not occur even at the first period of oscillation of the tether length. It is shown that the use of the proposed control laws leads to slight increase of deorbiting time as compared to the case of using the constant thrust.The results can be used to develop the control systems of small spacecrafts designed for deorbiting of space debris.

When designing any spacecraft, the opening elements of the structure are applied in which deployment and locking units are used.There is a transporting step i.e. an insertion of the deployment unit in the folded state into orbit. During this step the deployment unit is under considerable vibratory loads, which are accepted by thrusts. Since the thrusts at this step work under increased wear conditions it is offered, considering a protuberant surface hardness of the micro-arc oxide (MAO) coating, to replace a constructional material of the thrusts with a larger density for the MAO-coated aluminium alloy.The step of deploying, at which required angle of deployment is provided by the rolling unit, is realized after the spacecraft has been placed into calculated orbit. This unit contains the target sleeve with the bearing, which provides reliable locking by means of the penetration in the inner cone of beam sector. To increase the product reliability and avoid a possibility for occurring the wear slot and burrs it is necessary to increase the hardness of the sector surface on which the target sleeve bearing rolls. The most rational solution is to form the local micro-arc oxide coating on the beam sector surface.Application of the micro-arc oxidation of aluminium alloys and research of the micro-arc oxidation specificities enable us:1) to decrease the typical deployment unit mass (by 0.221 kg or 0.85 %);2) to increase the rolling unit operation reliability;3) to design the schemes and select the conditions to form the micro-arc oxide coating on the parts of the typical deployment unit.The micro-arc oxide coating formed on the aluminium alloy is of significance for the aerospace industry that uses a diversity of aluminium alloys and prefers the structure mass decrease as much as possible. It should be appreciated that the process to form the micro-arc oxide coating is production friendly because it does not need the complex expensive equipment and is easily regulated by changing the electrolyte composition and the current conditions.

Ensuring the safety of the cargo separated from the aircraft is one of the most important criteria in the development of cargo transportation systems. The article shows that the experimental studies using wind tunnels are time- and cost-consuming. Therefore, at the early stages in designing an aerodynamic configuration of the aircraft the software systems of solid modeling and computational fluid dynamics are, usually, used. The article describes the shortcomings of such software systems because of the large-scale computational costs. The offered technique for calculating the aerodynamic characteristics of cargo separated from the aircraft uses two main points. Firstly, in calculations a strategy of reducing a size of the grid cells when approaching to the object surface was used. Secondly, in modeling the trajectories of the separated cargo in the neibourhood of carrier aircraft a multivariate interpolation was used for prossessing the results of computational experiments and their subsequent applications.The proposed technique allowed a significant reduction of the computational costs when determining the aerodynamic characteristics. This technique was used to evaluate a reliabiity of obtained results through their comparison with available experimental data.

Currently promising development direction of space propulsion engineering is to use, as spacecraft controls, low-thrust rocket engines (RDTM) on clean fuels, such as oxygen-methane. Modern RDTM are characterized by a lack regenerative cooling and pulse mode of operation, during which there is accumulation of heat energy to lead to the high thermal stress of RDTM structural elements. To get an idea about the thermal state of its elements, which further will reduce the number of fire tests is therefore necessary in the development phase of a new product. Accordingly, the aim of this work is the mathematical modeling and computational study of the thermal state of gaseous oxygen-methane propellant RDMT operating in pulse mode.In this paper we consider a model RDTM working on gaseous propellants oxygen-methane in pulse mode.To calculate the temperature field of the chamber wall of model RDMT under consideration is used the mathematical model of non-stationary heat conduction in a two-dimensional axisymmetric formulation that takes into account both the axial heat leakages and the nonstationary processes occurring inside the chamber during pulse operation of RDMT.As a result of numerical study of the thermal state of model RDMT, are obtained the temperature fields during engine operation based on convective, conductive, and radiative mechanisms of heat transfer from the combustion products to the wall.It is shown that the elements of flanges of combustion chamber of model RDMT act as heat sinks structural elements. Temperatures in the wall of the combustion chamber during the engine mode of operation are considered relatively low.Raised temperatures can also occur in the mixing head in the feeding area of the oxidant into the combustion chamber.During engine operation in the area forming the critical section, there is an intensive heating of a wall, which can result in its melting, which in turn will increase the minimum nozzle throat area and hence will change the operating conditions of engine.

The aim of research is to develop software system for solving gas dynamic problem in multiply connected integration domains of regular shape by high-performance computing system. Comparison of the various technologies of parallel computing has been done. The program complex is implemented using multithreaded parallel systems to organize both multi-core and massively parallel calculation. The comparison of numerical results with known model problems solutions has been done. Research of performance of different computing platforms has been done.

В статье рассмотрена и проанализирована существующая нормативно-техническая база по вопросам утилизации летательных аппаратов. Рассмотрены вопросы как непосредственно нормативно-технической базы процессов утилизации, так и нормативно-технической базы по утилизируемым материалам, а также проанализирована сфера требований к качеству продукции, полученной в результате утилизации., У статті розглянута і проаналізована існуюча нормативно-технічна база з питань утилізації літальних апаратів. Розглянуто питання як безпосередньо нормативно-технічної бази процесів утилізації, так і нормативно-технічної бази по утилізованим матеріалами, а також проаналізована сфера вимог до якості продукції, отриманої в результаті утилізації., The article is devoted to describing and analyzing the existing normative and technical documentation for recycling of aircraft and quality ensure of this process. The problems of directly normative and technical documentation of recycling processes and normative and technical documentation for recyclable materials are considered, as well as analyzed the scope of requirements for the quality of products obtained from recycling. As a result, it is concluded that currently recycling of aircraft elements is seen as an integral part of recycling is practically not separated; normative and technical documentation largely determined for recycling of metallic compounds, and often impose a requirement for quality of recyclable products, and not to the quality of products obtained from recycling and almost no normative and technical documentation for evaluating the quality of products obtained from recycling of aircraft elements.

Processes and phenomena are under consideration, which can be realized from the viewpoint of nonlinear dynamics and are typical for geosystems. Attention is focused on the main features and new possibilities given by the ideas and approaches of nonlinear dynamics for solving applied problems of geophysics. Premises for physical interpretation of spontaneous electromagnetic and seismoacoustic emission as a product of temporally uninterrupted transitional dynamic processes related to interaction of different subsystems of geo-medium as its metastable conditions and succession of such conditions are analyzed. An attempt has been undertaken to analyze lithosphere-atmosphere-ionosphere relations, methods of multiparametric aerospace monitoring of short-term precursors of strong seismic events. Its possibilities and limits are considered and the necessity of integrating passive monitoring systems based on the study of medium-generated emission signals spectra with active ones, oriented to system receptivity control. Baselessness of attempts to give reliable and accurate short-term forecast without studying physics of producing destruction sources and development of deformation process within the Earth’s crust during the final stage in a regime with intensification has been noticed.