Gorbovskoi V. S., Kazhan A. V., Kazhan V. G., Shenkin A. V. Numerical studies of nozzle thrust characteristics of supersonic civil aircraft by computational gas dynamics method. Aerospace MAI Journal, 2019, vol. 26, no 4, pp. 7-16.

One of the most urgent problem while developing a new generation supersonic civil aircraft is ecologic requirements ensuring, including the community noise level near the airport. It requires developing and studying new technical solutions ensuring both low nozzle thrust losses level at all flight modes and reduction of jet flow velocity to decrease its noise level at the take-off/landing modes. One of the possible trends for this problem solving is mixer-ejector type nozzle application on the supersonic civil aircraft. Its operation principle consists in the fact that at the take­off mode with sound absorption, the hot jet is split into smaller jets by the multi-lobe nozzle. The increased surface area of the ruffled jet intensifies its mixing with atmospheric air, and reduces the length of the mixing layer initial section. The mixed jet velocity in the nozzle outlet section reduces, and thus the effect of acoustic suppression is achieved. Mixing zone shielding by the tail part elements of the airframe allows additional enhancing of acoustic suppression. At the flight modes without acoustic suppression the mixer- ejector type nozzle transforms into conventional supersonic nozzle with much higher thrust characteristics.

To reduce time and financial costs at the preliminary design stages, it is expedient to employ computational methods, ensuring high level of confidence. Modern software for fluid numerical modelling are applicable for solving a wide class of problems. However, refinement of flow numerical modelling technique is necessary for solving concrete problem. In the presented work, rational parameters for computing and computational grids were selected.

Modern Computational Fluid Dynamics (CFD) software allows solving a wide class of problems. However, refinement of flow numerical modelling technique is necessary for solving concrete problem. In the presented work, rational parameters for computing and computational grids were selected to study physics of the flow and obtain integral characteristics of the nozzle, such as mixer-ejector nozzle, at the take-off, landing, transonic and supersonic flight regimes. This method is employed to predict the nozzle thrust losses with ANSYS CFX commercial CFD code of Reynolds- averaged Navier-Stokes equation numerical solution. The numerical study of losses in mixer-ejector nozzle with active system of acoustic suppression at the take­off and landing modes are performed, and obtained results are validated by the experimental data. The accuracy of validation does not exceed 0.5% of the ideal thrust losses at all flight modes.

Murav’ev V. I., Bakhmatov P. V., Grigor’ev V. V. Specific defects forming features while aircraft bulky titanium structures assembling. Aerospace MAI Journal, 2019, vol. 26, no 4, pp. 17-27.

This article presents the results of the study of specific defects forming while VT20 and VT23 titanium alloys electron-beam welding. It was established that the presence of capillary-condensed moisture, resided in the defects of the edges’ surface, impacts dominantly on the submicropores formation. Other conditions electron-beam welding conditions, which may lead to specific defects forming, were revealed. These conditions may include:

  • Improper assembling and preparation of the abutting edges for welding;

  • Electron-beam welding modes;

  • A solid-phase joint formation prior to the front of the molten bath;

  • Oscillatory processes of the electron beam (~0.5 mm), which may lead to uneven melting (due to insufficient temperature of the edges’ overall melting) over the grains boundaries with submicropores forming (less than 0.00025 mm), which cannot be detected by modern X-ray machines;

  • Hydrodynamic collapse of the crater leading to the root defect generation as peak-shaped formations.

It was revealed by radiographic control and scanning microscopy that defects in the form of dark stripes represented the chains of submicropores projected onto each other. It was established also that specific defects formed while electron-beam welding impacts significantly on the strength properties of welded joints, as well as on their destruction stadiality. The performed studies allowed make a conclusion on the necessity of monitoring such basic factors as the surface quality of the abutting edges for welding; electron beam focusing conditions, its power and oscillatory processes; and hydrodynamic instability in the weld penetration channel.

Moshkov P. A. Problems of civil aircraft design with regard to cabin noise requirements. Aerospace MAI Journal, 2019, vol. 26, no 4, pp. 28-41.

The presented work is devoted to the problem of modern aircraft design with classical power plant layout, i.e. two turbofan engines on pylons under the wing, with account for the cabin noise requirements. The objective of the work consists in developing the list of scientific research and development activities, which execution is necessary for an aircraft design by the specified parameters of acoustic comfort.

The article considers the problem of noise level normalization in the aircraft cabin and cockpit. The main sources noise in the cabin were determined based on SSJ-100 aircraft testing. To minimize their sound pressure levels in the cabin a list of works while civil aircraft design was developed.

Determining relative contribution of various sources to the total sound pressure level along the cabin length, measured with the A-weighted scale of a standard noise level meter, is necessary for the right selection of methods and means for its reduction. The main sources of noise in the cabin and cockpit are the systems for air conditioning and ventilation, as well as pressure pulsation fields in the boundary layer on the aircraft fuselage surface.

Noise from the engines vibrational impact does not appear to be significant while evaluating total noise level in dBA. Acoustic radiation of the power plant, such as ventilator and jet noise, does not affect total levels of sound pressure weighted by A scale of a standard noise level meter in the cabin and cock pit at the cruise flight mode. The sound of aircraft avionics is not a significant source. But it can be said in general that placement of aircraft equipment systems aggregates should be executed with account for their acoustic characteristics.

The noise level they create in the cabin should be 10-15 dBA lower than the calculated sound pressure level in the cabin of the aircraft under development, determined at the control point of the cabin as the energy sum of noises from air conditioning system and turbulent boundary layer.

The results of this work can be used in the design of modern civil aircraft, with regard for the requirements to acoustic comfort.

The cabin noise problems of civil aircraft was considered. It was shown, based on the SSJ-100 flight tests that the dominant sources of noise in the cabin were the turbulent boundary layer and air conditioning system. The main directions of scientific and research activities, necessary for the aircraft design according to the specified parameters of acoustic comfort were formulated for these two main sources. Basic methods for noise reduction in the cabin were considered.

Valitova N. L., Kostin V. A. On probabilistic methods application to solving aircraft strength inverse coefficient problems. Aerospace MAI Journal, 2019, vol. 26, no 4, pp. 42-50.

Solving problems of static strength, fatigue resistance, and aeroelasticity can be performed in both deterministic and probabilistic formulation. Deterministic approach for aircraft strength computing is adopted as the basic one both in this country and abroad. Aircraft safety requirements increasing leads to the necessity of considering probabilistic safety criteria and development of normative standards for them.

The article deals with solving the inverse strength problems in a probabilistic setting in a general form. In the most general case, the elements of the “output”, as well as parameters of the structure under study, characterized by a certain operator, are stochastic. It is assumed that the probabilistic measure of the “output” is known and can be defined in the form of theoretical distribution law. In this case, the inverse strength problem in probabilistic setting is reduced to either determining the probabilistic measure of parameters of the “input” (at the determined parameters of the “object”), or to determining the probabilistic measure of the “object” parameters. It is assumed initially, that the problems under consideration are quasi-static, and unique deterministic dependence between the “input” and the “output” is known.

Examples of linear transformations for random variables are given when determining probability characteristics of load restoration and identification of structures for the two models, namely a beam and a thin-walled Odinokov’s structure.

Further, the article presents methods for analyzing static systems with random parameters. The real structural elements parameters randomness is being caused by the external environment disturbing effects, unavoidable technological production errors etc. It manifests in the form of cracks, starved spots, initial irregularities and other factors, which may affect the structure behavior in various ways. In particular, destruction may be associated with a large number of dislocations and stresses redistributions. This allows expecting non-linear manifestations in the structure material behavior in the form of hysteresis loops, leading in general case to non-Gaussian distribution of random values.

When considering static systems hereafter, an internal random value (e.g. crack) is being interpreted as an additional random impact at the deterministic system input. This affects the system behavior and leads to natural mixing of random output processes while their transformation in the system, i.e. the effect of natural formation of mixture of distributions.

The examples of determining the probability density for the potential energy dissipation of the rod deformation at random thermal effects, as well as functions of the mixture density in the presence of the internal defect in the beam were considered.

The obtained material can be recommended for developing a base of standards on mixtures’ references necessary for the purposes of structures diagnostics.

Amir'yants G. A., Malyutin V. A., Soudakov V. G., Chedrik A. V. On strength and aeroelastic characteristics of a large-scale model of an airplane wing section. Aerospace MAI Journal, 2019, vol. 26, no 4, pp. 51-65.

The article presents the computational and experimental results of aeroelasticity issues studies accompanying design and testing in wind-tunnel of a large-scale model of a passenger aircraft-demonstrator wing element the 7-th European framework program AFLoNext. The goal of the project consists in developing advanced flow control technologies for new aircraft configurations to achieve a quality leap in improving their aerodynamic performance.

Design, manufacture and assembly of a large-scale model, which serves for visual presentation of typical phenomena of flow separation in the fixation area of the wing with engine with high degree of bypass, were performed. However, such engines application on arrowhead wings causes undesired phenomenon of flow separation on the wing at low speeds and high angles of attack, which may lead to deterioration of the aircraft overall aerodynamic characteristics. To avoid these phenomena, the two newest types of technologies for active flow control are studied within the framework of the project. The pipe tests of the model were performed on the aerodynamic balance of the ADT-101 TSAGI pipe.

Based on the developed demonstrator CAD-model, detailed mathematical model of a demonstrator was built to compute the strength and safety of the pipe tests. Preliminary calculations of the structure stress- strain state indicated the need to strengthen the attachment area of the caisson spar to the beam of the supporting device. Comparison of natural frequencies and shapes of the first tones of mathematical model oscillations with the results of ground frequency tests was performed prior to testing. The difference between experimental and computed natural frequencies of the first oscillation tones did not exceed 10%.

Analysis of the structure behavior in the flow revealed the most loaded elements, in which minimum safety margin was η = 3, which corresponds to the ADT-101 TSAGI requirements. To control the nacelle and slat oscillations at the start-ups, computation of overloads limit values on nacelle and slat for understated strength margin of η = 2 with reference of the “stall” phenomena and turbulence was performed.

Critical flutter and divergence speeds were determined for ensuring safety of the demonstrator mathematical model tests performance in the pipe. The obtained values were out of the bounds of the velocities realized during the tests.

High measurements accuracy of the wing flow control systems efficiency was ensured by a comparative analysis of the local angles of attack of the structure under the impact of the ADT flow.

Kolyshev E. S., Krapivko A. V. Experimental methods for determining dynamic characteristics of aircraft landing gear. Aerospace MAI Journal, 2019, vol. 26, no 4, pp. 66-80.

The article describes methods and algorithms for determining the fundamental eigen modes of landing gear, such as torsion, lateral and longitudinal bending of support, according to the amplitude-phase frequency characteristics measured at characteristic points of the structure. Resonant frequencies, shapes and decrements of vibrations are determined using transfer functions (dynamic compliance and dynamic stiffness). A typical accelerometers arrangement of a system for oscillations registering and arrangement of vibration exciters are given. The described methods for obtaining dynamic characteristics were developed based on the long experience in landing gears GVT of various aircraft.

The novelty in landing gear GVT is marked:

  1. Moveable carriages with vibration exciter mounted on them, which are equipped with special connecting devices for attaching rods to the axis of wheels. The rods are equipped with forces sensors transmitted to the structure, in order to eliminate the excitation system effect.

  2. The GVT is performed for the landing gear both in a free state and at various vertical loads on supports created from action of the aircraft mass by hydraulic lifts.

  3. The applied shock method application on landing gear to obtain amplitude-phase frequency characteristics at the selected points of structure according to the results of response functions processing. This method allows giving an operational evaluation of the landing gear resonant characteristics and speed up the ground frequency testing procedure.

  4. The GVT results processing is performed using transfer functions of dynamic compliance and dynamic stiffness of landing gear strut for bending and torsion and their cross links.

  5. To determine hydraulic lifts effect on landing gear dynamic characteristics, the GVT in a free state is performed in cases when the aircraft is installed on the standard hydraulic lifts and when the aircraft is installed on pneumatic supports.

Parakhin G. A., Rumyantsev А. V., Pankov B. B., Katashova M. I. Low-current cathode designing for small stationary plasma thruster. Aerospace MAI Journal, 2019, vol. 26, no 4, pp. 81-89.

At present, the interest of spacecraft producers to low-power electric propulsions and propulsion installations on their basis is growing. The above mentioned fact imparts topicality to the task of expanding the family of cathodes for such thrusters towards decreasing discharge current maintained by the cathode.

It is well known, that effective cathode of the electric propulsion does not require any additional heat source in a steady-state operation, and thermoemitter operating temperature maintaining is ensured by the ion current on its surface. This article describes two complementary trends of works aimed at such cathode designing.

The first trend consists in the cathode thermal scheme optimization and thermal losses reduction. Some of design solutions, related to this field of work, were employed in the cathode experimental design and demonstrated their efficiency. On the other hand, the optimized design appeared to be sensitive to the smallest changes in the thermal scheme and, thus, needed a retrofit.

The second trend is a development and application of new thermal emissive materials with a lower operating temperature. The article presents the results of the works which have been in progress with some intermittences since 2013. The article demonstrates the results of Barium oxide-based thermoemitter samples developed and tested at EDB Fakel. The issues of thermoemitter manufacturing procedure; raw materials (powders) purity and dispersity; sintering temperature, and tool set, developed in the course of the works, are tackled.

As the result of handling of work, the authors came to a conclusion that for a higher efficiency of the new cathode design being developed it is necessary to consolidate the results of works in both trends. Further additional measures for the design optimization are planned.

Gol'berg F. D., Gurevich O. S., Zuev S. A., Petukhov A. . The onboard mathematical model application to control gas turbine engine with extra combustion chamber. Aerospace MAI Journal, 2019, vol. 26, no 4, pp. 90-97.

Modern gas turbine engines control is performed by the parameters accessible for measuring, which for the most part characterize indirectly the engine critical parameters such as thrust value R, specific fuel consumption CR, as well as parameters, affecting directly operational safety and reliability, such as gas temperature  in the combustion chamber (CC), stall margin (ΔSm) etc.

Employing the all-modes self-identified thermo­gas-dynamic model of the above said engine in modern digital automatic control systems (ACS) offer scopes for new opportunities of substantial control quality enhancing. This model allows computing with high precision the engine critical parameters in real-time scale, and realize the engine control directly by these parameters.

The article presents the results of studying such methods for controlling the fuel consumption GFE into extra combustion chamber, and nozzle throat area FT of the multi-mode engine.

The scheme of structural and algorithmic construction of such system is introduced.

Implementation of the three control programs, such as thrust changing RΣ depending on throttle position, and minimum  and maximum  values limiting of the air-to-fuel ratio αECC in the extra combustion chamber is being accomplished by affecting the fuel consumption (GFE).

Ensuring the minimum possible value of the specific total fuel consumption C = (GFM + GFA )/RΣ) , as well as restriction of fan stall margin, are implemented by affecting nozzle throat area by the extremal controller.

The effectiveness evaluation of the control methods under consideration was brought about by the integrated mathematical models “Engine – ACS – Onboard Mathematical Model” employed in CIAM.

It was shown, that direct engine thrust control by the impact on fuel consumption into the extra combustion chamber allowed ensuring the thrust value invariance to the engine components degradation while in operation.

The impact on the nozzle throat area herewith minimizes specific fuel consumption and limits the fan stall margin.

Baryshnikov S. I., Kostyuchenkov A. N., Zelentsov A. A., Lukovnikov A. V. Effectiveness estimation of turbo-compound scheme application on purpose of indicators increasing of aircraft piston diesel engine of 300 H.P.. Aerospace MAI Journal, 2019, vol. 26, no 4, pp. 98-107.

The goal of the presented work consisted in improvement of the engine basic indicators — specific power and effective specific fuel consumption (ESFC). This goal achieving is possible though three methods, based on a heat balance equation, namely, effective power increasing, as well as heat emission decreasing into cooling system and exhaust energy utilization. Effective power increase seems to be a conservative method that ensures relatively low performance increase, and is the main research trendHeat removal limitation to the cooling system was actively studied in 90-s, and currently considered unworkable. Thus, the best way to increase the engine indicators radically is the exhaust gases energy utilization.

There are many ways realization, including mechanical and electric compounding, the Renkin cycle application, thermoelectrical generators. However, the most efficient way from the niewpoint of specific parameters is mechanical compound.

Historically, turbo-compounding is a logical continuation of turbocharging. Turbo-compound engines are the pinnacle of aviation piston engines. VD-4K and Napier Nomad engines represent the examples of such engines, demonstrating at that time the unsurpassed fuel efficiency levels.

A six-cylinder boxer four-stroke turbocharged CI water-cooled engine was selected for the purpose of this study. The key factor for the diesel engine selection was the high air to fuel ratio, which was about two times higher than this for the gasoline engine. Owing to this, other things being equal the compound turbine will ensure twice as much power.

In this work, identification of the basic engine was being performed with the AVL BOOST software. The Patton, Nitschke, Heywood friction model, allowing determine friction losses based on the engine arrangement; Vibe combustion model, and Woschni 1978 heat exchange model were employed. Based on the obtained model a turbo-compound modification was developed. Optimization of basic parameters, such as charge pressure, pressure drop on both power and compressor turbines, gas distribution phases and ignition advance angle.

Based on the obtained results, a comparison of three variants of the engine, such as basic one; with the Garret turbine, which roughly corresponds to domestic prospective turbines; and the one with reference turbine was performed.

As a whole, the achieved results fit theoretical estimations with high degree of precision, with the exception of the exhaust gases temperature: contrary to the initial expectations, the temperature decreased. However, this result fits the pattern, established in other authors’ works.

The results of the comparison revealed that the power increment in the turbo-compound engine could achieve 10%, and ESFC reduction could achieve 11%.

Kiselev F. D. Fracture diagnostics and operational workability evaluation of working turbine blades of aircraft engine. Aerospace MAI Journal, 2019, vol. 26, no 4, pp. 108-122.

The topmost constituent part of the study on determining the cause parts of destruction of the aircraft in operation is fracture diagnostics employing the methods of physics-of-metals analysis of the fracture structure, material structure and composition determining, defect detection control, mechanical properties characterization, parts strength and survivability analysis.

Diagnostics of aircraft turbine blades operational fractures was performed, factors contributing to destruction were revealed, and causes of blades destruction were established. The article considers operational damageability specifics, on frequent occasions differing from the test bench ones, the systematization results of loading types, fracture mechanism, and operational fractures of gas turbine engine blades.

Methodical aspects were developed and new techniques were elaborated for fracture diagnostics were developed. The article systematizes external, fractographic and metallographic signs of diagnostics characteristic to anomalous (abnormal) modes of the engine functioning and a blade fracture at normal aircraft engine functioning (operating parameters did not outrun the operational limitations). The suggested classification allows determining blades fractures while operative diagnostics with account for joint action of static, vibration and thermal stresses in the blade material. It helps identifying blades fractures by the operational fractures types and revealing thermo­loading factors, determining the fracture mechanism, outlining it from all set of mechanical and thermal loadings acting on the blade.

The article presents the results of experimental studies of cyclic crack resistance of the blade made of VZHL12U (equiaxial crystallization) and ZHS26, ZHS32 (directional crystallization and single-crystal version correspondingly) alloys. Characterization of the blades material resistance to fatigue destruction with kinetic diagrams plotting (dependence of the crack growth rate on the stress intensity factor) was performed at the temperature of 850°C with samples loading on the vibro-bench. Eigen oscillations frequencies of the samples were of 70-120 Hz. Pulsating stretching scheme with the frequency of 50 Hz was used as well. The values of the cycle asymmetry coefficient in both cases were 0.15 and 0.35.

According to the results of high-temperature test and fatigue crack growth rate measuring on the samples from the above said alloys, kinematic diagrams of fatigue destruction, i.e. dependence of fatigue crack growth rate on stresses intensity coefficient values were plotted.

Based on the conducted fractographic studies and their results comparison with experimentally obtained ones and schematic kinetic diagram of fatigue destruction the schemes are developed; fractographically illustrated stages of fatigue crack growth and various fracture micromechanisms at different sites of the kinetic diagram of fatigue fracture in the material of the samples and blades.

The results of the work can be applied for developing more advanced modifications of turbine blades of high reliability.

Ezrokhi Y. A., Kadzharduzov P. A. Working process mathematical modelling of aircraft gas turbine engine in condition of elements icing of its air-gas channel. Aerospace MAI Journal, 2019, vol. 26, no 4, pp. 123-133.

The article presents general approaches to of aviation gas turbine engine operation modelling in icing conditions.

Component-level engine model is considered, in which the parameters, determining each component operation mode, represent a set of independent variables. These variables values are computed as the result of solving a system of nonlinear equations that determine conditions for the engine system components concurrent operation and its control laws Airflow continuity with account for its bleed and leaks, compressor and turbine power balance for the shaft of each engine are related to the concurrent work conditions, while fuel feeding conditions to the main combustion chamber and afterburner, as well as conditions, determining position of the nozzle actuator inlet guide vanes are related to the control laws.

It is assumed, that the ice formation in air-gas channel of this or that compressor stage, which leads to its airflow capacity reduction due to reduction of its conditional cylinder area of the inlet cross-section. The losses level the of inlet total pressure increase in the compression duct in consequence of inevitably occurring deterioration of compressor elements flow-around due to icing. Quantitative values of these impacts are determined from the engine gas-flow channel sizes, rate of ice growth, as well as the results of well-known generalizations on the unevenness effect of gas-flow channel on the total pressure losses in it.

Ice accretion rate may be set as data of engine testing results in icing conditions, or as a variable allowing evaluating its effect on the main engine performance parameters (thrust, rotation frequency, fuel consumption etc.). The other way to identify the ice accretion rate is solving of complicated thermodynamic problem of ice accretion on this of that part of engine duct surfaces.

The possibilities of the developed mathematical model were demonstrated based on data of test results of the ALF502R turbofan engine tested in ice crystal conditions in NASA Glenn Research Center. Good calculated and tests results matching herewith was demonstrated, which indicates the principal and proved approaches of turbofan operation modeling under the influence of this external factor.

Varsegov V. L., Abdullah B. N. Gas dynamic optimization of wedge-shape vaned diffuser of a centrifugal compressor of small-sized turbojet engines based on numerical modelling. Aerospace MAI Journal, 2019, vol. 26, no 4, pp. 134-143.

A competitive small-sized turbojet engine development under modern conditions of aviation engines building requires high efficiency values of parts with high degree of pressure ratio. Centrifugal compressors find extensive application while developing small-sized gas turbine engines employed for unmanned aerial vehicles and gas turbine power plants.

To ensure high efficiency and compressor pressure ratio, a numerical gas-dynamic calculation is performed with Ansys Workbench (Fluid flow CFX) program, which allows studying the air flow in the diffuser channels.

The presented article considers the flow in a wedge­shaped diffuser and optimize geometry optimization of the wedge-shaped diffusers blades of a centrifugal compressor, as well as geometry impact on the total pressure loss coefficient ξ, and the coefficient of static pressure recovery in the diffuser Cp at different entry angles α3l .

The main task of the calculation consists in determining the optimal shape of the wedge-shaped diffuser blades, insuring required parameters and characteristics of the diffuser, with an uninterrupted flow and a minimum of energy loss at given input angles.

The article presents also the results of the compressor stage numerical study, i.e. joint operation of the impeller with a diffuser to assess the quality of the geometry and operation of the diffuser to increase the compressor efficiency.

In the presented work, the calculation model is built with the SolidWorks program, and then, using the Turbo Grid program, the computational grid was applied. The flow simulation was performed using the SST turbulent viscosity model.

Nadiradze A. B., Frolova Y. L., Zuyev Y. V. Conical plume model calibration of the stationary plasma thruster by the thruster integral parameters. Aerospace MAI Journal, 2019, vol. 26, no 4, pp. 144-155.

The article presents the analysis of possible reasons for divergence of parameters measured under laboratory conditions and realized in space, based on application of multi-fractional conical model of the stationary plasma thruster jet. Three possible methods for the jet model calibration by the thruster integral parameters, such as discharge current, flow-rate and engine thrust were considered. The study of measuring conditions impact on the jet integral parameters was conducted. The need for calibration is stipulated by the fact that jet measured parameters may incorporate essential errors associated with the effect of experiment conditions and vacuum chamber walls. Calibration coefficients, linking measured and integral parameters of the jet, such as total ion current, flow-rate by ions and the jet axial impulse, are being introduced to minimize errors. Inasmuch as the jet integral parameters are being measured with high precision, the thruster jet model accuracy may be significantly increased after calibration.

The calibration methods regarded in the article allow obtain concurrence either by current density or by the flow-rate, or by the thrust (axial pulse). Jet calibration by the ion current and ions flow-rate gives the undervalued thrust value. Calibration by the thrust gives the jet parameters estimation for the worst case (overvalued parameters by the ion current and flow rate) necessary for analyzing the jet impact on a spacecraft. However, it is impossible to obtain the exact concurrence for parameters due to the effect of jet «disintegration» caused by interaction between the accelerated ions and neutral particles. Besides, the particles of residual atmosphere in vacuum chamber may affect the processes of jet formation in the acceleration channel of the thruster. To obtain more accurate jet model, it is necessary to account for the above-mentioned factors, and to use more complicate correction methods.

Marchukov E. Y., Vovk M. Y., Kulalaev V. V. Technical appearance analysis of energy systems by mathematical statistics techniques. Aerospace MAI Journal, 2019, vol. 26, no 4, pp. 156-165.

Aerospace industry development is impossible without implementation of up-to-date samples of high-efficiency new generation energy systems (ES). The term “technical appearance” implies the aggregate of parametric, structural and technological solutions, reflecting most substantial specifics of the system appearance [5]. It is well-known that designing and production of new technology, inclusive of ES in aerospace industry, leads to the necessity of taking compromise optimal or rational engineering and technological decisions. Besides, designer always faces the requirement for conformity of technical appearance forecast of the ES being developed to its real-life prototype. An engineering approach based on statistical analog technique for decision-making while developing new technology may be of help for the appointed tasks solution and meeting the above said requirements [10, 15]. This technique foundation consists in the fact, that deep analysis and synthesis of static structural and energy data of the ES, selected analogs and prototypes according to the parameters of technical requirements to the design according to [15-17] are performed while prospective equipment development. The article regards the energy system (ES) in general form as a mechanical machine for input energy conversion into useful work. Methodological basics of the new generation ES optimal appearance forecasting by mathematical statistics techniques [15-24]. The article demonstrates that development and introduction of the special statistical criterion, integrating all operational parameters in the form of multi-parametrical function, is urgent for solving scientific and engineering problems of new ESs development with specified properties of enhanced effectiveness. This criterion may be named forecast criterion. The introduced special forecast criterion is based on ES statistical analog data fields processing (already created and successfully operated) by mathematical statistics techniques [15-17]. The criterion of the analytical form analysis by independent parameters-arguments leads to formulation and solution of the extreme problem of a multi-parameter function optimizing by known mathematical methods [18, 20, 24], where obtained optimal parameters determine the forecast of the newly created ES optimal technical appearance. Algorithm for compiling and special forecast criterion computing in general is presented. To demonstrate the legitimacy of the criterion introduction, an example of computing the forecast of the ES technical appearance in general is given. The scientific results of the article may be used to develop a comprehensive software product for modeling technical optimal concept of a new generation ES with increased output energy operational parameters and optimal mass-dimensional (volumetric) characteristics.

Kartas S. S., Panchenko V. I., Aleksandrov Y. B. Geometric parameters effect of ejector with curvilinear section of mixing chamber on its characteristic. Aerospace MAI Journal, 2019, vol. 26, no 4, pp. 166-173.

Ejector is the simplest device without moving parts for liquids, gas, and other media moving. Power transfer from one stream to the other proceeds by their turbulent mixing. Very often, injector is employed to maintain continuous airflow in a duct, or a premise, thus performing a fan role. It is used also for jet engines testing. The exhaust stream flowing from the jet nozzle draws in the air from the shaft into the ejector, ensuring thereby the premise ventilation and engine cooling.

Over the past 60 years, plenty of studies has been performed on ejectors as a part of jet engines, which purpose consisted in increasing engine thrust, and reducing fuel consumption, jet noise and output temperature.

In modern conditions, these devices are used in various fields, such as aircraft and machine building, firefighting equipment, and as pumps, compressors, and mixers at oil tank farms.

In general, the described ejector structures include straight-line mixing chambers. Employing a curvilinear section of mixing chamber, which allows improve the ejector parameters, may be suggested as an option of such ejectors. An option of the ejector of this kind consists of a high-pressure flow nozzle, a low-pressure flow nozzle, mixing chamber, and diffusor. With this, the initial section of the mixing chamber is curvilinear.

The disadvantage of this ejector is certain difficulties in manufacturing curvilinear surfaces of nozzles and initial section of the mixing chamber. The advantage of this ejector consists in average velocity reduction of the active jet at the mixing chamber inlet, and, as a consequence, mixing losses reduction.

The article presents the results of numerical calculation of the  characteristics of curvilinear ejectors with F1/F2 = 1 geometric parameter (elbows and bends) at relative sizes of R/a = 1; 2. These results revealed that with the same ejection coefficients, the relative pressure drop is greater for a curvilinear ejector with a relative radius of R/a = 2. The numerical calculation was performed in a stationary setting using the Fluent program and the k-e RNG turbulent viscosity model. Based on preliminary calculations and the grid independence analysis of the obtained results, the grid models were selected.

Volkov S. S. Assessment techniques for psychophysiological state of special purpose systems operators. Aerospace MAI Journal, 2019, vol. 26, no 4, pp. 174-183.

The article deals with assessing techniques for psychophysiological state (PPhS) of a flight crew, cosmonauts, test pilots and other representatives of the aerospace industry. An approach, involving gas discharge visualization method in conjunction with fuzzy logic system for psychophysiological state monitoring is being offered for consideration. Prospectives of automation system for psychophysiological state assessment techniques implementation in the interests of aerospace comples are demonstrated.

The purpose of the work consists in demonstrating the increase of the PPhS assessment quality of special purpose systems operators of the aerospace industry. Special purpose systems operators are both civil and military aviation flight crew, cosmonauts, test pilots, and specialists dealing with robotic systems.

This work novelty lies in the intelligent tools application for operators’ PPhS determining. The interest to this method application is caused by the fact that human ability to perform professional duties is characterized by his psychophysiological state. Psychophysiological state monitoring of operators of special purpose systems (SPS) of aerospace industry allows increasing efficiency of their decisions and raise their readiness to perform special duties. Eventually, the ability to perform special duties unconditionally may and must be controlled and monitored to enhance readiness to perform the assigned task during the periods of flying vehicles flights and testing.

In this respect, the necessity for performing control of SPS operators of aerospace industry at the stage of their preparation for flights and tests performing, as well as during special assignments performing with automation tools application is imminent. It would allow assess with certain fidelity their readiness to perform the assigned tasks during flights and tests, and point out to particular official the necessity to pay attention to this or that pilot, cosmonaut or technician. However, such control implementation is not possible without methodological tools and means for assessing flight crews, cosmonauts and other aerospace industry prepresentatives fitness for their functional assignment.

As the result of the studies, an algorithm of the decision-making support system with fuzzy logic system for automated assesment system of PPhS operators was developed.The fuzzy logic system operation is based on the Mamdani algorithm.

The PPhS assessment techniques implementation, described in the article, in the aerospace industry will allow monitoring the health of the flight crew, cosmonauts, test pilots and operators of robotic systems, as well as reducing the risk of injury and mortality factor while equipment operation.

Boyarskii G. G., Sorokin A. E., Khaustov A. I. Experimental pressure-flow characteristics determining of micropumps for orbital station biotechnical system. Aerospace MAI Journal, 2019, vol. 26, no 4, pp. 184-190.

While conducting research at the space stations, great attention is paid to revealing the weightlessness effect on the cells, which allows the results transferring to the other objects and models in various areas of biology and medicine. For such studies performing, the authors suggest to apply a biotechnical system for cell culture (BTS CC) in conditions of spaceflight, which main element is a micropump, meeting the following requirements:

– to possess minimum size: diameter of not more than 10 mm, and length of not more than 50 mm,

– to ensure a liquid supply with viscosity of 1 cSt from zero to 0.1 liters per minute,

– to ensure pressure of up to 3 J/kg.

The existing techniques for axial pumps design do not allow correctly determine the micropump geometric size and its pressure-flow characteristic, since with a pump size reduction compared to the full- size pump, relative size of gaps and roughness increase, which changes significantly redistribution of the velocities fields and volume leakages, as well as disk and friction losses. A micropump designing with such specifics requires new structural and designing concepts.

Based on the full-size pump designing experience and with account for the BTS CC pump operation specifics, a new micropump of 6.5 mm diameter and 45 mm length was developed. Its control block allow changing rotation speed and the electric motor and impeller of the micropump by setting the current frequency and value, varying hereby the pump delivery and pressure.

Any pump characteristic is its head dependence H on delivery Q at various rotation frequencies of the pump shaft, i.e. H = f (Q, n). Thus, to determine the micropump pressure-flow characteristics, experimental studies are necessary to examine the effect of geometric size and mode parameters on its characteristics.

The main difficulties in the pressure-flow characteristics determining of micro-pumps, i.e. the dependence of the pump head on its supply and shaft speed, is their small size, commensurable with the sensors size.

Analysis of publications related to the study of fluid micro-flows in micro-pumps revealed that they use tracers were employed for this purpose, which introduction disrupts the micro-pump operation. Thus, to determine micro -pumps characteristics, a test

bench was designed and manufactured. It includes non­inertial micro-sensors (for the pressure drop-head registration and measurement). The flow rate was measured by weight, with account for the liquid evaporability. The micropump pressure-flow

characteristics are modeled by changing hydraulic resistance at the pump outlet by varying the flow section of the throttle. The measurements were repeated for different speeds of the impeller shaft from 2000 to 20,000 rpm.

The results of the tests revealed that the micropump pressure-flow characteristic represent a falling dependence typical for the full-sized axial pumps. However, stratification of dynamic characteristics is being observed at various impeller rotation frequencies. Thus, for the range of n1 > 8000 rpm the pressure-flow characteristic goes higher, than for n2 < 8000 rpm. The obtained pressure-flow characteristic of the developed micro-pump allows estimating the effect of the micropump micro-sizes on its efficiency.

Kirsanov A. P. Stealthy movement of aerial object along rectilinear paths in the onboard doppler radar station detection zone. Aerospace MAI Journal, 2019, vol. 26, no 4, pp. 191-199.

Onboard radar stations operating in the pulse–Doppler mode show the characteristic feature in the detection zone. This feature consists in the fact that in every point of the detection zone the aircraft has a sector of directions moving along wich it not detected by the onboard Doppler radar. This sector is called the sector of invisible motion directions of the aircraft. Due to these features, there are stealthy paths allowing an aircraft stays non-detected by Doppler radar station, such as radar station of an airborne early warning aircraft, while moving along them. The majority of stealthy trajectories is curvilinear with variable curvature. The article deals with the study of the rectilinear paths of the aircraft stealthy movement in the onboard Doppler radar station detection zone. It was established that any aerial object position relative to the early warning aircraft might be the start of the rectilinear stealthy path at the appropriate selection of direction of movement. An equation to determine the stealthy movement duration along the rectilinear path depending on the aircraft initial position and its direction of movement was obtained. Areas in the detection zone of the pulse-Doppler radar station to which the aerial object may enter, moving along the rectilinear stealth paths, were plotted. Their shapes and sizes depending on the aerial object position and motion parameters relative to the radar station were studied. Conditions of the unlimited time duration of movement along the stealthy paths, and conditions of the rectilinear stealthy paths for the aerial object outgoing to the onboard Doppler radar station location were found.

Bezzametnov O. N., Mitryaikin V. I., Khaliulin V. I. Low-speed impact testing of various composites. Aerospace MAI Journal, 2019, vol. 26, no 4, pp. 216-229.

The purpose of the study consists in technique development for detecting impact damages character of composites with various nature of reinforcing material and interlacement type. A series of experiments on the presence of internal defects after impact damages inflicting was conducted while this work performing. The samples based fourteen fabric types were selected as the subject of the study, including fiberglass cloth, hybrid materials, Kevlar® and high molecular polythene. Temperature mode was developed, and technology for plates manufacturing by the compression molding technique was worked out.

The experiment technique was being developed with regard for the international Standards recommendations for damage resistance testing while the falling load impact. Evaluation of criteria on impact resistance was performed within the energy range of 10, 20 and 30 J. Initially the dent depth was determined with digital detecting head. The internal damages areas were being estimated by the semi-automated ultrasonic NDT complex with phased array. This technology allows obtaining scanning results in the form of projections onto three planes, namely C-scan (top view), S-scan (end view) and B-scan (side view). To analyze the damages areas of samples after the impact, the C-scan, depicting the scanned area below the sensor, was registered. The layer-by-layer studying of the samples damages character was performed by the X-ray computer tomography method. This method allows visualize the sample internal structure by processing shadow projections obtained while the object X-raying.

The obtained results allow determine optimum characteristics of the composite material pack content while developing the structure with the set requirements to the impact resistance. The nose part elements and high lift devices of an aircraft, helicopter blades and transmission shafts, moving parts of jet engines may be among these structures.

Based on these works results graphs of the damages areas dependence on the impact energy of each material were plotted. The less damage area was demonstrated by the fiberglass samples, while the greatest one belonged to the fabrics of hybrid content. To evaluate the impact resistance criteria the energy of the damage initiation, maximum load of impact and absorbed energy were registered. Maximum value of the damage initiation energy was demonstrated by the samples from hybrid fabric material, and the least one by the fiberglass samples. This criterion reflects the limiting value of the impact energy which a material can sustain without being damaged.

Savel’eva L. V., Vendin I. O. Cutting conditions effect on tool front surface wear rate while workpieces machining. Aerospace MAI Journal, 2019, vol. 26, no 4, pp. 209-215.

The article tackles the issue of determining the degree of various cutting modes effect (cutting speed, cutting thickness, cutting width, feed, cutting depth, temperature, front angle, vibration) on the front surface wear of the cutting tool.

The authors describe the nature of cutting modes effect on the front surface wear of the tool, and suggest recommendations on optimal cutting modes, which ensure maximum life span of the tool.

The article consists of three main sections: introduction, the bulk section, conclusions.

The introduction considers causes of the tool wear. As a rule, cutting tools wear occurs under the impact of molecular adhesion forces of the treated metal surface with the cutting tool, or under abrasive action of solid particles existed in the structure of the machined material.

The main section regards the tool wear process over the front surface. It analyzes an experimental dependence of the cutting speed impact on the tool wear intensity. As the result of the analysis conclusion was made that the wear increased with the cutting speed increase. According to professor A.M. Danielyan’s studies, with the cutting speed, feed and cutting depth 20% increase the cutter surface wears out correspondingly 3.5, 1.7 and 1.05 times faster. This research data demonstrates that the largest effect may be achieved not by the cutting speed increase, but by the cut width and thickness increase. The effect of the cut thickness and feed on the wear intensity of the cutting tool is analyzed. With large cut thickness (more than 0.5 mm), a misgrowth of significant height is formed, eliminating the contact of the rear surface with the cutting surface. Only the front surface of the tool thereby wears out. With the cut thickness reduction, the wear occurs on both back and front surfaces simultaneously. At very small cut thickness (less than 0., the misgrouth is of rather insignificant height, and the wear occurs only on the back surface. With feed increase, the cut thickness increases either, and, thus, the wear on the front surface increases. The experimental dependence of the cut depth impact on the tool wear intensity is analyzed. As the result, the optimal cutting depth is determined, at which the front surface wear is minimal. The experimental dependence of the tool temperature influence on the tool wear intensity is analyzed. The optimal tool temperature, at which the wear of the front surface is minimal, is determined. The effect of the tool front angle and vibrations on tool wear is analyzed.

Recommendations on selection of optimal cutting modes, ensuring maximum tool life are presented in conclusions.

Ushakov I. V., Simonov Y. V. Experimental detection of micro-destructions viscosity in central and boundary areas of brittle samples while loading on the substrate by vickers pyramid. Aerospace MAI Journal, 2019, vol. 26, no 4, pp. 230-239.

The main purpose of the work consists in developing the earlier proposed technique for viscosity detection of micro-fracture of thin brittle amorphous nano-crystalline samples.

The regularities of deformation and fracture under local loading of solid thin samples of nano-crystalline material by Vickers pyramid are determined experimentally. The main studies were performed on amorphous metallic alloy Co71,66Si17,09B4,73Fe3,38Cr3,14, converted into the nano-crystalline state by controlled isothermal annealing.

The dependency of the symmetry of micro-patterns of destruction from the load value and a distance to the sample boundary was established. It is established that with the load growth occurrence of symmetry elements starts to be observed in the initially asymmetric fracture patterns. Statistical analysis of symmetric cracks, as well as the distances between them, allows find the micro-destruction viscosity of the material. At a certain optimal load, the probability of symmetrical micro-patterns formation is maximal. A further load increase leads to the symmetry reduction, and, accordingly, to the decrease of micro-destruction viscosity calculation accuracy.

For the first time, a technique for determining the minimum allowable distance to the boundary of a thin sample, on which the micro-destruction viscosity determining was possible, was proposed. It was established that the optimal load value while determining the micro-fracture viscosity near the sample boundary coincides with the value of such for the central areas.

For the first time, mechanical testing modes, which allow obtain symmetrical and analyzable micro­patterns of destruction were determined. These conditions include the following: using the optimal load on the indenter; accounting for the allowable distance between the adjacent loading areas and a distance from the loaded area to the sample boundaries. Based on the experimental results analysis, algorithms for to determining the optimal load on the indenter and the allowable distance to the sample boundary have been developed. The obtained results allowed improve the earlier proposed technique for micro-fracture viscosity detection by local loading of thin, hard and brittle samples.

Sedel'nikov A. V., Belousova D. A., Orlov D. I., Filippov A. S. Assessment of temperature shock impact on orbital motion dynamics of a spacecraft for technological purposes. Aerospace MAI Journal, 2019, vol. 26, no 4, pp. 200-208.

The main objective of the work is assessing the of temperature shock impact on the orbital motion dynamics of the spacecraft for technological purposes.

The problem consists in the uncertainty of center of mass displacement due to the impact of temperature shock and, thus, the motion control error. This problem is particularly relevant for the spacecraft for technological purposes, and products sensitive to the experimental conditions.

The importance of assessing the impact of temperature shock is determined by the need to ensure the spacecraft functioning with the specified parameters of motion, as well as maintaining controllability of the spacecraft in the presence of orbital eclipse periods.

Analysis of the studies by the scientists from various countries reveals that control of a small spacecraft with no large elastic elements in the design-layout scheme often reduces to the target values active control of the angular velocity of its rotation.

In this case, the orbital eclipse periods are not highlighted separately, and no changes in spacecraft movement control law are made while its immersing in and out of Earth shadow.

The article deals with the issues related to the temperature shock impact on the orbital motion change of a spacecraft for technological purposes, and modeling the scale and nature of the effect.

The temperature shock impact assessment is based on the 3D modeling of the processes occurring at the spacecraft entering and exiting the orbital eclipse period.

For a small “Return— MKA” type spacecraft the three-fold excess of admissible micro-accelerations was obtained.

As the result of the conducted study, a conclusion was made that control algorithms development, levelling the temperature shock from the viewpoint of occurring micro-accelerations compensation, was required for successful implementation of gravity- sensitive processes onboard the spacecraft for technological purposes with the orbital eclipse period.

A three-dimensional heat conduction problem was solved to determine the target parameters of control algorithms. The following simplifying assumptions were introduced to solve the problem:

– the elastic element model was a frame structure;

– the elastic element was rigidly fixed in the small spacecraft body;

– the elastic element properties satisfied the conditions of homogeneity;

– the heat flow was uniform;

– operating temperature range was −170... + 110 °C;

– the properties of the elastic element material were considered constant throughout the operating temperature range;

– orientation changing of normal to the elastic element surface due to its own oscillations was neglected.

Ivanov P. I., Kurinnyi S. M., Krivorotov M. M. Asymmetry in the parachute canopy filling process. Aerospace MAI Journal, 2019, vol. 26, no 3, pp. 7-16.

The main purpose of the work consists in studying dynamics and specificity of filling the large area parachutes of the main class employed for rescuing re­entry spacecraft as well as large weight cargoes airdrop of civil and military hardware. The problematic issues here are these associated with the occurrence of large aerodynamic load values while parachute dynamic filling, which may lead to premature loss of its strength. The issues of long delay in the filling process, which increases the path and height loss and is very dangerous while low-height airdrop, are of no less importance. The article tackles the issues associated with the filling process deviation from the rated value, such as asymmetry occurring while the parachute canopy filling.

The dependence between the filling time and aerodynamic load on the parachute, i.e. maximum drag force value, was established experimentally. The article demonstrates that with the parachute filling time increasing the aerodynamic loads on the parachute and overloads on the object decreased, while the filling path increased.

The relationship between the edge contour of the canopy inlet orifice shaping, filling time and aerodynamic loads on the parachute was established. One of the possible causes of both deceleration and intensive canopy filling dynamics, consisting in substantial asymmetry of the shaping process of the edge contour of the parachute canopy inlet orifice, was revealed.

The authors introduced the notion of the canopy contour shaping asymmetry coefficient at the intensive dynamics of the canopy filling process, as an effective tool for studying the processes of canopy edge shaping processes and their quantitative evaluation.

Setting the rated boundary value for the asymmetry coefficient it is possible to make judgments on the tendency of the canopy shaping by the degree of distance from this boundary. Thus, it will show the propensity of the specified parachute for the asymmetric filling and the ensuing negative consequences, associated with intensive dynamics of the filling process and load-carrying capacity loss. Practically, the asymmetry coefficient represents the square root of the ratio of impact pulses from the air- velocity pressure (which form local pressure drops along the carrying surface) for the canopy with asymmetry, and a canopy being filled symmetrically, under the same initial conditions on speed.

The larger the coefficient of asymmetry, the larger the dome is predisposed to asymmetric filling, the more shock impulses differ. In this association, the probability of the canopy and shrouds destruction increases in the local loading area from the pressure drop at the loads being measured by the strain sensor in the parachute thimble, which are substantially lower than its load-bearing capacity.

Novogorodtsev E. V. Numerical study of total pressure in the air intake with sharp edges applying eddy-resolving sbes-method. Aerospace MAI Journal, 2019, vol. 26, no 3, pp. 17-31.

The values of the total pressure oscillations intensity root mean square parameter ε in the channel of isolated air intake with sharp edges were determined as applied to industrial aerodynamics problems based on numerical solution of Navier-Stockes system of equations. Numerical solutions of Navier-Stockes system of equations were obtained using eddy­resolving Stress Bkended Eddy Simulation (SBES) approach employing ANSYS CFX solver. Simulation of the 3D flowing of the viscous compressible gas around and inside the object was performed employing spatial regular multi-shell grid. The procedure of computational grid generation was being performed in manual mode employing ICEM CFD software.

To evaluate fidelity of the computational study based on SBES method application, comparison of the obtained values of the root-mean square parameter of pulsations intensity with experimental data was performed. The data processing procedure herewith was conducted in concordance with the standard experimental technique approved in TsAGI.

Numerical simulation results are presented in the form of plots of parameter e values in the engine section as a function of the specific reduced air flow q(λen) through the engine cross section. The air intake duct throttling was modelled by cross-clamping of the auxiliary duct in the form Laval nozzle. The auxiliary duct wall profile in the longitudinal section herewith was constructed using the Vitoshinsky formula.

The article performed a comparison of total pressure oscillations obtained while computational study in monitored points of the metering cross­section with oscillograms obtained while experimental study according to readings of the total pressure pulsations sensors, installed on the model at the same points of the reference cross-section.

The parameter ε values obtained in the framework of this work in the engine cross-section for the air intake and engine synchronization mode in all regarded range of of the incoming flow Mach numbers M = 0-1.8 (at zero angles of attack and sideslip) are in good agreement with the experimental data. Maximum discrepancy between computational and experimental results was Δε max = 1% in absolute units of the ε parameter.

The ε parameter values were obtained for both the air intake configuration without a boundary layer control system, and the one with a boundary layer control system.

Bragazin V. F., Gusarova N. A., Dement’ev A. A., Skvortsov E. B., Chernavskih Y. N. On practicality of deflectable thrust vector application for civil aircraft. Aerospace MAI Journal, 2019, vol. 26, no 3, pp. 32-42.

The study focuses on the engine deflectable thrust vector (DTV) application on the civil aircraft to improve its controllability, as well as take-off and cruising-flight characteristics.

Thrust vector deflection is achieved through the movable nozzles. Three options of the engines location in the aircraft layout, namely, on the pylons under the wing, as well as on the pylons of the fuselage nose and tail parts were considered. Esteems of the DVT application as an additional element to the aerodynamic control elements were obtained.

The DVT application as an additional balancing element of pitch and/or yow control leads to the possible reduction of the horizontal tail (HT) and/or vertical tail (VT). Thus, for the aircraft layout with the engines under the wing, the HT area reduction may be of 11%, and VT area reduction of 8%. For the aircraft layout with the engines in the fuselage tail part, the VT area reduction may be of 13–20%. The DVT application along with the aircraft aerodynamic control elements allows increase the effectiveness of the lateral, pitch and yow control, as well as reduce the aircraft response time to the steady-state overload.

The aircraft cruising aerodynamic quality changing depending on the engines position on the aircraft and thrust vector deflection was considered. The largest increase in maximum quality was realized with the engines location in the front part of the fuselage and upward thrust vector deflection. It was revealed, that aerodynamic quality increases about 2% within the angles range of 0° to ±10°. According to the preliminary estimates, the aggregate impact of several factors may ensure the fuel consumption reduction in the cruising flight by approximately 3–4%.

While studying the takeoff trajectory, it was found that the largest trajectory slope angle at the safe takeoff speed was possible with the DVT engines application in the taili part of the fuselage.

According to the preliminary data, the DVT application bears a potential to improve a civil aircraft operational characteristics. The DVT significant useful effects are the possibility of aircraft control dynamics improvement and flight safety enhancement at the takeoff/landing and climbing modes.

Levin V. I., Karasev D. Y., Sitnikov M. S. Aircraft break wheels designing using 1D thermodynamic models. Aerospace MAI Journal, 2019, vol. 26, no 3, pp. 43-61.

The OEM, EASA and ICAO requirements to aircraft systems and equipment force manufacturers to conduct more verification calculations and tests to confirm the announced characteristics, as well as analysis of various modes of operation. Currently, there are already new methods of design, as well as automation of calculations and tests. Thus, it is necessary to develop both theoretical and practical basis for their implementation.

The objectives of this work consist in determining a convenient method for thermal processes computing in the the aircraft wheel structure, as well as describing a method for developing a 1D model for the wheel thermodynamic calculations, performing computations by this model, and comparing the obtained results with the results of test modes.

The article provides a summary of the research and work conducted at the enterprise of the brake wheels manufacturing company. The approach to computing the thermal energy distribution dynamics over the friction disk volume and the wheel structure while braking process is being substantiated. The adequate accuracy while using the reduced model of the disks temperature computing is demonstrated. The article presents the processes and methodology issues of developing architecture and parameterization of the wheel structure model for computing the points of the monitored temperature. The model additionally accounts for the convective thermal exchange with the pneumatic network of the air cooling from the brake wheel. Speed, direction and successive air heating are also being accounted for. The results of computing and testing at three test modes are presented. The adequate accuracy of the computational results compared to the testing data is being determined.

Eventually, all declared goals were achieved. A convenient method for thermodynamics computing of the wheel based on the 1D model was determined. Virtual testing was performed on both a model and a test bench. Analysis of the results allows stating the expediency of the 1D models while brake wheels designing.

Virtual tests were performed on the developed and validated model, which allowed determine more optimal modes of the test bench equipment application. This, in its turn, allowed the time reduction of the field tests and the number of test launches.

Currently, a set of documentation has been developed to justify changes in the regulations for the design and conduct of accelerated life tests of the wheels. The prospects for the used computing method development for solving the related tasks of the break wheels design.

Boldyrev A. V., Pavel’chuk M. V., Sinel’nikova R. N. Enhancement of the fuselage structure topological optimization technique in the large cutout zone. Aerospace MAI Journal, 2019, vol. 26, no 3, pp. 62-71.

Topological optimization techniques play an important role while selecting a structural layout of aggregates for a flying vehicle of minimal mass. The goal of the presented work consists in increasing weigh efficiency of the aircraft structure in the stresses concentration zones. The article proposes a of topological optimization method for edging of the cutout for the hatch in the fuselage, based on the full- stress concept with regard for the functional limitations on the generalized hull skin displacements at the cutout contour.

For the design object synthesis, a method, based on Komarov’s mathematical model of a deformable solid body with variable density is being applied. An artificial material with variable density and rigidity, called a “filler", in which the strength and elastic properties linearly depend on density, is being employed.

Finite element models, integrating the manifold of the load-bearing elements of the structure and continuous medium of variable density are being developed while topological design. Earlier, such combined model was employed in [25, 26]. The material distribution in the filler allows revealing theoretically optimal structure and, using the strategy [8], developing the structural layout closest to the theoretical solution from the viewpoint of its stressed operation. The topological optimization process is based on stage-by-stage substitution of the filler by structural elements, realizing the technical decisions being taken

The article presents a numerical example of the fuselage compartment design with rectangular cutout, demonstrating the operability of the suggested technique. Conventional layout with well-known prototypes technical solutions is adopted as an initial structure. The topological optimization resulted in obtaining new technical solution allowing 16,7% reduction in the mass of the strengthening members of the cutout relative to the initial structure. The parts of the internal panel are shifted inward the fuselage from its theoretical contour and duplicate the hull skin at the cutout portion. The internal panel is fixed to the hull skin by the longitudinal and sloped walls, reinforced and ordinary bulkheads. The manifold of stressed elements forms closed and hollow contours in the cutout corners, enhancing the structure rigidity in the hatch cutout zone in radial and longitudinal directions.

Mamedov I. E., Sharifova B. A. UAV functioning mode optimization while seawater sampling. Aerospace MAI Journal, 2019, vol. 26, no 3, pp. 72-79.

Water is a necessary factor for the humankind survival. For this reason, the quality of water resources should be protected. Thus, it is necessary to organize permanent monitoring of water resources. Industrial and agricultural wastes are the main sources representing danger for water basins. Water quality of rivers and lakes may be evaluated by monitoring such indices as quantity of dissolved oxygen, pH., temperature, and electric conductance. Low concentration of oxygen dissolved in the water, undesirable temperature and abnormal salt content lead to water quality degradation. The article is dedicated to the issues of UAV application for the seawater salinity and conductance determining. The UAV application for this purpose allows increasing space-time resolution of the results of the studies being performed. The task of forming the UAV empirical model in water sampling mode was formulated. Electric conductance sensors while corresponding UAV flight altitude control are being immersed into the water and taken out after conduction measuring. Thermal sensors are applied herewith, installed on the other UAV flying 30-40 meters higher than the first one. Temperature survey is performed to reveal undercurrents of the incoming external water, which temperature and salinity differ greatly from those of the basic water body. The studies employing heuristic procedure of collating the values of the searched indicator, computed by different representations in the form of one graphics data, and checking the obtained results by the data represented by the other graphics data were performed. The article suggests an empirical model of the UAV, employed for the water quality studying. The empirical model of the UAV in the mode of sampling for the samples analysis is presented as well. Specific issues of realizing the suggested empirical algorithm for the empirical model development were considered. Indirect validation of the developed empirical model demonstrated close agreement of experimental and modelled dependencies character obtained based on heuristic algorithm of the UAV functioning in the water quality studying mode.

Abdulin R. R., Podshibnev V. A., Samsonovich S. L. Determining planetary gearing optimal gear ratio allowing minimize its outer diameter at the specified load torque. Aerospace MAI Journal, 2019, vol. 26, no 3, pp. 80-90.

Mass and size parameters reduction is one of actual issues of aircraft electromechanical drives design. It concerns especially mechanical transmissions employed in drive systems of mechanical transmission. Harmonic and planetary gears are most compact. They allow obtaining large gear ratio for a single stage. Their application as the output stage of a multi-stage reduction gearbox of an electro-mechanical drive, as a unit transmitting the largest moment, allows mass and size parameters reduction of a drive system.

The goal of this article consists in determining the optimal values of gear ratios at which the outer diameter of planetary transmissions has its minimum size for the specified load moment.

It was demonstrated, that the main parameter affecting the outer diameter of planetary transmissions for the specified load moment was the carrier radius. For a single-row planetary transmission this radius was expressed through the gear tooth module value, the number of teeth of the central sun-gear and gear ratio between the sun-gear and satellites. The article presents substantiation of the above said parameters selection. Minimum acceptable carrier radius was found. It was established, that optimal gear ratio value of the single­row planetary transmission equaled four.

The carrier radius planetary gear with double-row planets was expressed by gear tooth module and two gear ratios, namely between the central sun-gear of the planet gear and first-row satellites, and between planet gear of the second row and the crown-wheel. The dependence of the carrier radius on these gear ratios, which is represented by a surface with «ravine», was plotted. A unified optimal gear ratio value was not obtained for the planetary transmission with double­row satellites was not found. However, a set of quasi­optimal values do exist. The “ravine” direction, along which the quasi-optimal values were located, was determined. The optimal relationship of gear ratios between the central sun-gear and the first-row satellites, and between the second-row satellites and the crown wheel was derived. This relationship allows ensure minimum outer diameter of the planetary transmission with double-row satellites. An example of the minimum outer diameter of the planetary gear with double-row satellites computing is given.

The obtained optimal gear ratio values expand the knowledge on planetary transmissions and allow minimize overall dimensions of aircraft drive systems while developing multi-stage reduction gearboxes for electromechanical drives with output planetary transmission.

Nikolaev E. I., Nedelko D. V., Shuvalov V. A., Yugai P. V. External airbags application onboard a helicopter. Aerospace MAI Journal, 2019, vol. 26, no 3, pp. 91-101.

The subject of the presented article is an energy absorption system in the form of external airbags, fixed under a helicopter fuselage. The external airbags are meant for reducing the risk of injury of the passengers and helicopter damage in case of a crash landing.

The study of the external airbags impact while crash landing was performed by the finite elements method. The airbags mathematical model, accepted in the computations, assumes gas simulation by the thermodynamic parameters (pressure, temperature) averaged by the airbags volume. The article presents the airbags initial characteristics for the case of the gaseous nitrogen application. Gas leakage from the airbags is determined by the area of the vent hole and the value of relative pressure for initiation of the gas outflow from the vent hole. The initial pressures values and the holes areas were selected by the condition of overloads minimizing and the strength of airbags material ensuring.

The purpose of this work consists in analyzing the helicopter fuselage loading with the external airbag, and identifying the time dependencies of main thermodynamic parameters of the gas work. The study of a helicopter collision encompasses the moment of time of the airbags contact with the ground to the moment of the fuselage gaining a stable position on the ground. The process visualization of the helicopter fuselage spatial position changing so far as the airbag crimping is demonstrated. Velocities and overloads in the helicopter fuselage center of mass are presented according to the results of computations. The obtained dependencies of pressure, temperature and mass flow rate may be employed for technical requirements forming to the external airbags and gas generating elements structures. Computational results considered in the article allows drawing inference on the possibility of the external airbags application for the helicopter energy shock absorbing and increasing the rate of passengers and a crew survivability. The presented values of loads acting on the fuselage from the airbags side may be employed for the detailed designing of the airbags fixing to the fuselage. The conclusion presents the issues which may become a further development of the research topic.

Chernovolov R. A., Garifullin M. F., Kozlov S. I. Validation of designing and manufacturing procedures of aircraft dynamically similar models with polymer composite materials application. Aerospace MAI Journal, 2019, vol. 26, no 3, pp. 102-112.

Drained dynamically scaled models have been designed for studying unsteady aerodynamic characteristics in wind tunnels. At present, such models testing is of the greatest interest both from the viewpoint of their application for studying safety of the prospective aircraft from the flutter and buffeting, and for verification of calculated aerodynamics with account for the structure elasticity.

The article presents an algorithm for design parameters selecting of a dynamically scaled model and its tuning by test results. The proposed procedure for implementing this algorithm is demonstrated on a simple example (a beam of constant cross section, reinforced by layers of a polymer composite material). Issues of technology for design and manufacturing of a typical element of the dynamically scaled aircraft model applying polymer composite materials are considered. Frequency tests conducting technique is presented, as well as the results of computational and experimental studies of the shapes and frequencies of natural oscillations with account for the additional loads placement. Computed shapes and frequencies of natural oscillations obtained by the finite element method using several successively condensed grids are given. The research findings comparison indicates that calculated values of the cross-section bending stiffness obtained using theoretical relationships and characteristics of the material, accounting for epy specifics of dynamically similar model manufacturing technology, are close enough to those obtained by the experiments at static loading and resonant tests conducting. Setting-up such model does not require special efforts. It allows considering, that the accepted calculating and design technique ensures obtaining required characteristics of the dynamically similar model.

Matiukhin L. M. The fuel molar weight impact on filling, and indicator indices of a piston combustion engine. Aerospace MAI Journal, 2019, vol. 26, no 3, pp. 113-123.

The problems arising while improvement of any type of the internal combustion engine (ICE), such as reciprocating, rotary-piston, gas turbine or jet engines, are common for all of them.

The notions of the volumetric efficiency (nv) and residual gases (γr) traditionally used in the theory of piston internal-combustion engines do not allow characterize the air-fuel mixture composition, which defines the all power, economic and ecological indices of the engines. All the above-mentioned coefficients are applied only while the reciprocating ICE design. With this, the main indicator of pistons filling, namely volumetric efficiency, characterizes not so much the cylinders’ filling as its downgrade due to the presence of hydraulic resistances and incoming charge warming up. The essential drawback of all known equations for the volumetric efficiency determination is ignoring the impact of the fuel type, excess-air coefficient and recirculation’s degree on the cylinders filling. The general-technical concepts of (volume) fractions are far more informative. The aggregate of air-fuel mixture fractions determines its composition and thermodynamic characteristics values. The incoming charge (air) fraction allows unambiguous judgment on the degree of filling the whole cylinder volume, i.e. on the existing reserves of filling. Using the air or mixture volumetric fraction as the main filling indicator while piston ICE cycle computing allows accounting for the fuel molar weight and recirculation impact on the engine indices. As the result of the analysis, in order to account for the fuel impact on the filling the so-called “displacement coefficient” was proposed. Power and economic indices of the engine depend on this coefficient value. The value of this coefficient determines the degree of qualitative power regulation efficiency. Together with the recirculation degree, this coefficient determines the value of stoichiometric relationships and, thus, affects the indicator and effective indices of the engine.

As the sum of the fractions equals to the one, there is no necessity with the suggested approach in separate determining the fraction of the residue gases, since this fraction is equal to the difference between the one and the incoming charge fraction. The suggested approach is of prime importance while analyzing operating cycles of the engines operating on gaseous fuels, and on hydrogen in particular. As a result, the structure of the main calculation dependencies is simplified, and their analysis becomes more clearly evident and easy- to-understand. The possibility of the computing results visualization facilitates their analysis and is a great advantage of the suggested approach in terms of didactics.

Employing the ICE computation as a base of the air-fuel mixture fractions in modern applied programs might have led to the labor intensity reduction and execution time cutting due to the number of variables reduction.

Zubrilin I. A., Didenko A. A., Dmitriev D. N., Gurakov N. I., Hernandez M. M. Combustion process effect on the swirled flow structure behind a burner of the gas turbine engine combustion chamber. Aerospace MAI Journal, 2019, vol. 26, no 3, pp. 124-136.

The article presents the results of computational and experimental study of the swirling flow structure of a swirling jet behind the burner unit of an industrial gas turbine installation. The burner unit being studied in this work is intended for burning poor pre-prepared mixtures. The burner consists of an axial vane swirler with hollow blades through which the main part of the fuel enters, and a “central body”, functioning as a stabilizer with a pilot flame. Natural gas is employed as a fuel. The studies were performed by applied methods of computational gas dynamics and experimental methods. Experimental velocity measurements were performed with a laser Doppler particle velocity meter LAD-056S. Combustion products composition measurements were performed by sampling with subsequent chromatographic analysis. Experimental studies were conducted under the following conditions:

- The inlet temperature Тк = 330 К;

-  Differential pressure ΔP* ≈ 3,3%;

-  Reynolds number at the burner outlet Re ≈ 12000;

-  The proportion of fuel consumption in the standby zone is 11.5% of the total fuel consumption;

-  The excess-air factor for the case of mixing fuel without combustion was α = 2.08, and for the case without combustion α = 1.8.

The flow and combustion processes modelling was performed in three-dimensional unsteady formulation using Large Eddy Simulation (LES) method. Combustion processes were being described with the Flamelet Generated Manifold model. The GRI 3.0 mechanism was selected as the kinetic mechanism of chemical reactions. As a result, a comparison of time- averaged velocity fields and turbulence characteristics was being performed for the case of fuel combustion and without combustion. The obtained simulation results are well agreed with the experimental data on the flow velocity, its fluctuation components, as well as chemical composition. Thus, the employed approach may be applied for calculation study of the combustion processes of the gaseous fuel in swirling flows. An exception is carbon monoxide, which needs to be modeled using approaches accounting for non­equilibrium chemical combustion processes, such as a network of ideal reactors. The flow structure behind the burner was studied in detail, and the characteristics of the recirculation mixing zone were obtained. It was shown, that the fuel supply does not significantly affect the flow structure. It was found, that the combustion process changes the shape of the reverse streams, increasing it in diameter. Mass flow while combustion is significantly lower than in the so-called “cold” case. Due to the air-fuel mixture low consumption through the recirculation mixing zone for the given burner unit, the combustion process characteristics are mainly affected by the interaction between the recirculation mixing zone and the main flow. Pressure fluctuations associated with the vortex core precession, detected while cold purges, were not found during combustion.

Grigor'ev V. A., Zagrebel'nyi A. O., Kalabuhov D. S. Updating parametric gas turbine engine model with free turbine for helicopters. Aerospace MAI Journal, 2019, vol. 26, no 3, pp. 137-143.

A priori estimation of an aircraft engine mass takes on an important role while its creation, especially at the initial designing stage, when conceptual basics of the engine are being established. At this stage, when the design working out of the engine is not done yet, its weight estimation together with fuel economy indicators allows making valid selection of the engine working process parameters values. The presented work refines the parametric model of a gas turbine engine with the free turbine (GTE FT), used in the problem of the helicopter engine working process parameters optimization at the conceptual design stage. With this, while performing parametric studies the design mass of the power plant should be estimated according to the GTE parameters, though, up to now these dependencies are not studied quite well. Thus, the estimation of the engine mass dependencies on its parameters is being performed at present based on the generalized statistic data on the already accomplished structures or parametric mass models, since there is no more precise information at this stage. In fairness, it should be noted that they are all related to the aircraft engines. A rather smaller number of works is oriented of the mass estimation of the helicopter GTE FT. This is primarily due to the fact, that these engines belong to the class of the small-size and have thereupon a number of specifics.

At the same time, as new versions of gas turbine engines appear the periodical refinement the parametric model coefficients values is required. he article considers the mass model of the gas turbine engine with free turbine for several options for the reduction gear mass accounting for, namely, both as a part of the engine, and the power plant. The authors suggest representing the coefficients used in the above said GTE FT models in the form of dependencies on the working process parameters. It allowed perform parametric studies and obtain predictive solutions corresponding to the achieved current design level of gas turbine engines.

Mil’kovskii A. G., Atamasov V. D., Kolbasin I. V., Ustinov A. N., Kalinina A. M. New phenomena in the space experiment on creating an artificial solar eclipse while the spaceships “APOLLO”-”SOYUZ” joint flight. Aerospace MAI Journal, 2019, vol. 26, no 3, pp. 144-151.

The presence of gas-and-dust plasma atmosphere is discovered in every spacecraft, which is confirmed by many domestic and foreign researchers. Due to the medium mixing under the impact of parameters gradients, the radionuclides of plasma atmosphere formed with the intensive impact of gamma and neutron radiation of the reactor would migrate to the outboard space area, surrounding protected part of the spacecraft structure and instrument bay with electronic equipment. These elements would be exposed to radiation due to the induced radiation. In this case, the deterioration of the spacecraft radiation protection against the onboard reactor occurs, which would lead to fluences excess of radiation fluxes on the instrument bay and sensitive structural elements relative to the acceptable levels. Formation of the flows of the eigen external atmosphere (EEA) substance irradiated by the reactor from the operating reactor into the area of the instrument bay and back is stipulated by the presence of parameters gradient of the EEA substance between the specified areas. These parameters are the volume plasma potential and, correspondingly, concentration of charges, pressure and temperature of the gas-and- dust plasma medium. This plasma migration got physical substantiations, published in many scientific works on nuclear physics, performed under I.V. Kurchatov guidance, which attaches authenticity and meaningfulness to the outlined concept, as well as determines the necessity to developing measures for the spacecraft extra radiation protection.

In 1975, an international experiment was conducted in the outer space under the “EPAS” program, during which the artificial Eclipse of the Sun and the solar corona was photographed during the Apollo and Soyuz spaceships joint flight. The spacecraft EEA was repeatedly registered while this experiment. We employed the said photos to analyze the properties of the spacecraft outboard atmosphere. It allowed comprehending the similar processes in the atmosphere of the spacecraft with nuclear reactor.

The physical phenomenon of the “identic luminosity” was recorded by the experimental method in conditions of the space flight under the EPAS program. This phenomenon is a confirmation of the induced radiation phenomenon from the EEA area being under the direct impact of the radiation source due to the various processes of the radiant energy transfer between the particles of the atmospheric environment, varying in weight, shape, chemical content etc., to the shadowed area, protected from direct radiation of the nuclear source, into the atmosphere area. The “identic luminosity” of atmospheric matter can only be explained by the fact that the energy losses while the radiation migration between the described areas are minute. This phenomenon is reliably rendered on all published EEA photos employing high-sensitivity photo film. Such film employing was predetermined by the weak luminosities of the phenomena studied in the experiment such as solar corona and the spacecraft Apollo EEA. They are approximately millions of times smaller weaker than the Sun radiation. Thus, they are being detected only during its full eclipse. This was artificially created in the “Apollo”-“Soyuz” spaceships joint flight (EPAS).

It is necessary to add justification for the necessity for measures to clean the spacecraft outboard space from the EEA caused not by the induced radiation phenomenon only, but also by other non-traditional processes that lead to disturbances in the spacecraft onboard systems functioning.

Lepeshinskii I. A., Tsipenko A. V., Reshetnikov V. A., Kucherov N. A., Sya S. . Joint measurement of gas-dynamic parameters of two-phase highly concentrated flows by laser-optical and probe methods. Aerospace MAI Journal, 2019, vol. 26, no 3, pp. 152-160.

The article considers the problems of joint application of the laser-optical technique for measuring parameters of the two-phase highly concentrated gas-drop flow. Each technique does not allow measuring all necessary parameters. The probe method allows adequate measuring of the local values of the phase flow rates and determine concentration, while measuring phase velocities and drops dispersivity requires suggestion of various hypotheses, requiring experimental verification.

Laser methods allow measure the drops velocities and their sizes in the two-phase flow. However, earlier they could not be applied for studying the flows with large concentration of dispersed phase, as well as determining the gas phase parameters in the two-phase flow. The laser engineering evolution resulted in developing lasers with high spatial and temporal definition, allowing their operation in the area of high concentration of the condensed phase. Combining these two techniques for the two-phase flow study allows go ahead in the area of measuring the parameters, which were either impossible to be measured, or determined with significant error. Particularly, to measure the gas phase velocity and improve measurement accuracy.

Laser-optical methods and Probe methods have long been employed to measure two-phase flow parameters. They are the ones of the few, by which local phase flow rate can be measured. However, their application arouses a number of problems. This is isokinetic problem while sampling and the impact elasticity coefficient selection. Certain design improvements and the probe technique application in compilation with PIV-method allows solving these problems and determining all parameters of the two- phase flow at high concentrations.

The probe represents a cylindrical channel employed in two modes: sampling and measuring the stagnation pressure of a two-phase flow. The problem of isokinetic sampling and selecting the elastic coefficients values of the impact of drops, determining the kinetic energy transfer in the two-phase flow during its braking (the stagnation pressure measurement), were analyzed. To ensure isokineticity, a structural solution was proposed for the probe, which ensures significant error reduction. Application of laser with high temporal and spatial resolution for measuring (PIV-system) allowed determine the drops velocity in a highly concentrated two-phase flow, and, based on the joint measurement with a probe, the coefficient of impact elasticity. The proposed techniques allowed measuring for the first time all the necessary parameters of the two-phase flow. Particularly, we managed to measure the gas phase velocities, and to perform a qualitative comparison with the flow rate of the gas phase at the two-phase flow outlet from the nozzles of the engine combustion chamber mixer.

Katashova M. I., Parakhin G. A., Rumyantsev А. V. Multiple mode cathode-compensator developing for the stationary plasma thruster. Aerospace MAI Journal, 2019, vol. 26, no 3, pp. 161-166.

There is a need today in creating a highly efficient multi-mode stationary plasma thruster capable of both inter-orbital transfer and spacecraft position keeping in a set point. A multi-mode cathode-compensator capable of operating at a discharge current up to 15 A is needed for this purpose. The cathode operates on the principle of a gas-electric source of electrons based on a hollow cathode, and it is the most thermally and energy intensive element of the thruster. The K-3/15 cathode structure was designed and studied experimentally on the possibility of flame operation in at least two modes within the discharge current ranges fr om 3 to 5 A and from to 15 A at the experimental design bureau “Fakel” base. The main purpose of the К-3/15 tests was verifying the cathode operability at various start-up powers, propellant flow rates and discharge currents to determine optimal start-up modes. In the process of stand-alone testing, it was determined that the optimal start-up mode for the cathode is a start lies within (160±5) sec at the heating power of 130-139 W and at the cathode flow rate from 30 to 0.60 mg/s. A special attention was paid to determining the current-voltage and voltage-flow rate characteristics in the discharge current range from 3 to 15 A at propellant flow rates to the cathode in the range from 0.30 to 0.60 mg/s. A comparative analysis of the main characteristics of the КН-3В cathode and К-3/15 cathode was performed as well. It was revealed, that compared to the KH-3B cathode the cathode K-3/14 current effectiveness value would manifest itself at the high-current modes (above 10 A), wh ere this parameter value was three times lower. It was determined that the K-3/15 cathode ensured the multi-mode operation with respect to the discharge current and had much higher resource parametrics compared to the KH-3B cathode. It is being forecasted, that parameter changing of the thermo­emitter from mono-crystal lanthanum hexaboride will allow three times increase of the flame operation.

Artyushenko V. M., Kucherov B. A. Analysing the system of restrictions on spacecraft control means application, accounted for while their scheduling. Aerospace MAI Journal, 2019, vol. 26, no 3, pp. 178-189.

A number of tasks of various resources scheduling should be solved to ensure spacecrafts mission control. One of such tasks is tracking, telemetry and command (TT&C) ground stations scheduling. That task is performed under strict resource restrictions. These restrictions include both restrictions on the resource being scheduled and temporal restrictions being imposed on the operativeness of the ground stations distribution plan developing. To ensure operative and qualitative TT&C, accounting for all these restrictions is required.

The restrictions on employing ground stations include the ones on applying separate ground stations as well as restrictions on various ones simultaneous employing. Restrictions stipulated by mission control centers capabilities to perform communication sessions with spacecraft are also a part of the restrictions on TT&C ground stations application.

The restrictions on employing a separate ground station include radio-visibility zones, a set of ground stations network for each spacecraft, a set of service operations to be done for ground station (during which it cannot be used to perform communication sessions with spacecraft) and a set of operation modes supported by each ground station. The restrictions on simultaneous application of different ground stations include ones caused by electromagnetic compatibility and restrictions caused by necessity of employing same resources. The restrictions caused by electromagnetic compatibility can be defined through the sets of two communication sessions characteristics, which cannot be performed simultaneously. These definitions can be used to identify conflict situations while TT&C ground stations scheduling. The resources which simultaneous application may be limited can be sharable or non- sharable. Demands for such resources can be associated with ground stations or their models. It will allow, in is turn, identify conflict situations while ground stations scheduling. Another restriction, which should be regarded while identifying conflict situations during ground stations scheduling, is the maximal number of communication sessions, which each mission control center can perform concurrently. The presented restrictions can be considered as the system of resource restrictions to be accounted for while TT&C ground stations scheduling. The proposed mathematical task formulation of accounting for the system of restrictions can be employed in future development of methodical support for ground stations scheduling.

Maron A. I., Maron M. A., Lipatnikov A. Y. Defining the number of employees for project realization of ground-based radio engineering flight support means upgrade. Aerospace MAI Journal, 2019, vol. 26, no 3, pp. 190-200.

The study relevance is stipulated by the fact that at present the number of projects for ground-based flight support radio engineering means (REFSM) is increasing. The REFSM upgrade represents a project. Such project is associated with a large number of works to be performed. Thus, just one division of the St. Petersburg Center for the of Air Traffic Organization performs technical operation of retranslation stations equipment in the area from Priozersk to Nizhni Novgorod. It is required defining the number of employees for the project completion in the specified time. It should be noted herewith that the same employees ensure operative runability restoring of equipment. The error-free running time of modern REFSM means is tens of thousands hours. It is ensured by both redundancy and technical servicing. A the same time, the defects causing the unit transfer from the operation condition to the fault operable state occur more frequently than the defects leading to inoperability. Such defects require operative elimination since they increase the failure occurrence probability. This problem has not been resolved up to now. Classical methods for queuing systems computing are based on computing probabilities of the system being in various states. They are practically inapplicable due to the dimensionality of the problem under consideration. Simulation methods describe special cases only. They do not guarantee the solution of the problem without analytically found initial approximations to the required number of personnel. The presented article solves the problem by the mean dynamic method. It presents the program for performing computations of the required number of employees in MathCAD Prime. The example of the number of employees computation is given. The proposed method gives practically exact results when the number of units to be upgraded is a couple of dozen or more. In case they are less in number, the obtained number of employees should be refined by simulation. The values obtained by the proposed method herewith will be the initial approximations. The materials of the article are of practical value for the managers of the flight support and communication REFSM services while the upgrading projects planning.

Ied K. . Developing a technique for hazardous situations warning system design while piloting errors occurrence. Aerospace MAI Journal, 2019, vol. 26, no 3, pp. 201-209.

Studying the accident rate of sports aircraft indicates a large number of accidents associated with control loss etc., due to piloting errors and piloting at unacceptable speeds, altitudes and overloads. The current situation requires a flight test methodology developing and specifying airworthiness standards for aerobatic aircraft to improve flight safety.

To define the safe altitude of the maneuver commence, it is also necessary to identify the probabilistic characteristics of piloting errors. Obtaining a functional relationship, based on studying altitude changes in the presence of piloting errors with the regard to the probability of these errors, will allow determine the safe altitude of the maneuver commence with a specified degree of probability.

A mathematical model was developed for studying the impact of pilot’s errors on the changes of trajectory parameters when performing maneuvers on an aircraft.

As a rule, control system of a light sports aircraft is characterized by the extreme simplicity, and is not supplemented with the capability of automated control (autopilot system). Thus, a task arises to develop a warning system, which is not based on automated control (automatic withdrawal from the dangerous altitude), but produces a warning signal only. It requires developing a technique for the warning system developing, which level should be associated directly the probability of the emergency occurrence to prevent this situation transfer to catastrophic one.

The article suggests this problem solving by the technique, according to which it is necessary to supplement the aircraft system with a unit, which would receive velocity and altitude parameters and compare them with the preset values of the acceptable velocities. This is important for warning the pilot on a possible situation to withdraw straightway from the maneuver being performed.

Korobeinikova E. S. Evolvement of quality management systems effectiveness assessment mechanism in aerospace industry. Aerospace MAI Journal, 2019, vol. 26, no 3, pp. 210-219.

Two significant disadvantages are inherent to the procedures of aerospace industry suppliers’ quality management systems (QMS) certification for compliance with whether the universal standard ISO 9001:2015 “Quality Management Systems – Requirements” or industry-specific AS/EN 9100:2016 “Quality management systems – Requirements for aviation, space and defense organizations” have two significant disadvantages. These disadvantages do not let the interested parties (primarily, customer companies and the State) to obtain maximum value added fr om external audits.

Firstly, only the inference on the compliance / non-compliance of QMS with the requirements of the declared standard is the result of certification, without quantitative estimation of the QMS maturity level of the monitored enterprise. Secondly, within the audit the QMS effectiveness is assessed in terms of achieving the results determined by each particular enterprise, whereas, there are quite specific indicators in the aviation industry, characterizing the effectiveness of the implemented systems and the competitiveness of the enterprise.

The aim of the article was to develop recommendations for improving the methodology of the QMS effectiveness assessing. Two trends of improvement were proposed, namely, creating a mechanism for quantitative assessment of the QMS effectiveness level, based on the AS9101 Standard for effectiveness assessing of separate processes, as well as detecting competitiveness rates of the enterprises critical to the specified industry (and, accordingly, clarifying the term “competiveness” for an aviation enterprise).

The first is the development of a mechanism for quantitative assessment of the QMS effectiveness level. The mechanism is based on the one used for assessment of the individual processes effectiveness in the standard AS 9101. The second direction is determining the competitiveness indicators that are critical for organizations of the aerospace industry (and, accordingly, clarifying the term “competitiveness” for aviation enterprises).

A quantitative assessment of the system effectiveness can be performed using the QMS assessment matrix (based on the PEM – process evaluation matrix – used in AS 9101). It is proposed to mark one of its axis with the level of the planned results of the activities

It is proposed to mark the level of planned performance results achievement on one of the matrix axes, and the level of implementation of the QMS standard requirements on the other. The final quantitative assessment of the QMS effectiveness is a score fr om one to four, obtained at the intersection of grades on both axes.

The planned performance results herewith, indicated on the second axis of the QMS assessment matrix, are computed as a complex indicator of the enterprise competitiveness.

This indicator will be computed by the formula:

where αi is the weight of the indicator i, determined by experts;

ci is the parametric index of the parameter i, computed by the differential method (the values of relative indicators determined by the industry are assumed as the base). Individual and group indicators, evaluated while computing the complex indicator, can be derived from the definition of the aerospace enterprise competitiveness specified by the author. Thus, the competitiveness is the ability of an enterprise to meet the consumer needs in terms of the competitive production. This means the qualitative production, corresponding to the consumers’ expectations on acquisition costs operation. It implies also the servicing quality, and related products and services in the necessary quantity and within the required terms, as well as demonstrating to the parties concerned (both direct customers and integrators of various levels, primes) the steady development in conditions of changing external medium, characterized by the costs cutting and profit rising. It should demonstrate also, the effective management, flexibility and ability to optimize their activities, including implementation of new management technologies, peculiar to the industry, namely increase labor productivity, maintain labor, scientific potential and cooperation expressed in the number of customers and partners increasing

С = f (C ; P; R; P; V; V; K; Q; N; m),

Cp — product competitiveness;

P — profit;

R — profitability;

PT — labor productivity;

Vp — the volume of production;

Vr — sales volume;

K — human resources;

Qcoop — an indicator of cooperation activity (increase in customers, suppliers and partners while maintaining the existing ones);

N — scientific and technical potential (includes such indicators as growth in new technologies applicaton (including IT technologies), the volume of in-house development, R&D costs);

M — effective management (increase in use of new management technologies - for example, risk management, lean production and others).

Thus, due to the new methodology application, the QMS effectiveness esteems and the set of competitiveness indicators while QMS analysis of the existing aerospace industry enterprises, the audit emphasis are shifting from the system correspondence to the Standards requirements to the system effectiveness in terms of achieving specific indicators, important to the customers of the aviation industry. Besides, the audits results a cquire quantitative character and allow comparing various suppliers.

Liu L. ., Shi J. ., Bao H. . A metal-composite joint and its mechanical performance. Aerospace MAI Journal, 2019, vol. 26, no 3, pp. 220-227.

A jointing technique, which can be employed in metal-composite joints and may enhance the ability to non-admission of joints disbond, is proposed in this article. This type of joints will contain a certain number of thin pins running though the substrates in the overlap region of the metal-composite adhesive bonded joints. There is adhesive on the surface of the pins and thus, the pins are bonded together with the substrates. And thus, the pins running through the joint plates not only arrest the cracks in the adhesive layer of the bonded joints, also transfer some load between the metallic and composite components. Comparative test results show that the proposed joint method can increase the strength, the failure strain of the metal-composite joints comparing with the traditional adhesive joints, moreover, the joint method can decrease the suddenness of the joint significantly and therefore, improve the damage tolerance performance of the bonded joints. Secondly, the effects of the number and arrangement of the pins on the mechanical performance of the joint will be analyzed in accordance to the test results also. And finally, an optimized method which can improve the load capacity and fracture toughness of the joints will be obtained.

Nasonov F. A., Gavrilov G. A., Babaitsev A. V., Nazyrova O. R. Target modification of constructional epoxy-carbon plastics as a materials science approach to the effect of mechanical joints orifices on bearing capacity. Aerospace MAI Journal, 2019, vol. 26, no 3, pp. 228-242.

The materials science approach to polymer matrices physic-mechanical properties management requires the assessment of modifying additives impact on technological and main operational properties of compositions. Works on studying and intercomparing the main technological properties of the initial epoxy composition and the one modified by technological Zinc Stearate (ZC) technological addition were conducted by viscosimetry and thermo-analytic methods. The developed kinetic model of the compositions hardening process revealed the trifling impact of the composition modification on the hardening process. Pilot samples from the plastics filled with carbon long-fibered fillers (impregnating under pressure and autoclave molding) were fabricated, and their non-destructive control and standard samples testing were performed for mechanical properties measuring.

Estimation by the computer tomography method revealed the stability augmentation of material structure along the edge of the orifice contour after machining for carbon plastics modified by ZC within the interval of 0.1-2% of mass. Thermal effects measuring of machining processes with various tools were performed by IR-thermography method combined with recording function at the specified intervals. The dependence of thermal effects from the modifier concentration was established. The article demonstrates that while this parameter measuring as an integral characteristic, temperatures reduction (temperatures maximums) is observed at the modifier content in matrix samples of 0.1–0.3% by weight, and at the content of 0.2–0.5% by weight in the carbon plastic samples (depending on the applied tool).

Podguiko N. A., Marakhtanov M. K., Khokhlov Y. A. Magnetron discharge application prospects as an electrons emitter in cathode-compensator for electric propulsion thrusters. Aerospace MAI Journal, 2019, vol. 26, no 3, pp. 167-177.

The subject of the presented article consists in assessing the prospects of magnetron discharge application as an electrons emitter for electric propulsion thruster cathode-compensator. This theme relevance is associated with the development of new stationary plasma thrusters (SPT) for the spacecraft operating on iodine, as well as low-orbit spacecraft employing outboard air as a working substance.

The paper assesses the energy aspect of magnetron cathode-neutralizer application for modern stationary thrusters. The highest operating voltages of the prospective dual-mode SPTs are 500-800 V. If a ten percent sacrifice of the propulsion system efficiency is possible with the view of increasing the service life and chemical resistance of the cathode-neutralizer, then the operating voltage of the magnetron cathode should be reduced to 120-180 V.

The article proposes a mathematical model of a magnetron discharge, on which basis a theoretical estimation of the magnetron minimum operation voltage and its dependence on the secondary ion- electron emission coefficient is presented. For a magnetron discharge with a copper cathode in the argon atmosphere, the minimum operating voltage equaled to 126 V. Besides, the minimum magnetic flux necessary for the discharge existence was computed.

An experimental study of plasma-forming gas pressure impact on the operating voltage value of the magnetron discharge was conducted for several options of the cathode material-working gas combination. These combinations were copper - argon, stannum - argon, stannum - argon-air mixture and aluminum - argon-air mixture. Minimum discharge voltage of 160-170 V was obtained when operating on an argon- air mixture and employing an aluminum cathode.

The performed studies allowed making the following inferences and recommendations:

  1. Cathode design should ensure optimal values of both the magnetic flux above the cathode surface and working gas pressure in the discharge area for the effective operation (minimum voltage).

  2. One of the ways to the electron cost in the magnetron cathode is the optimal.

Anisimov K. S., Kazhan E. V., Kursakov I. A., Lysenkov A. V., Podaruev V. Y., Savel’ev A. A. Aircraft layout design employing high-precision methods of computational aerodynamics and optimization. Aerospace MAI Journal, 2019, vol. 26, no 2, pp. 7-19.

Nacelle shape and engine position optimization was performed for Blended Wing Body aircraft (BWB). Aerodynamic characteristic computing method, used in the optimization procedure, is based on numerical calculations of the Reynolds-averaged Navier-Stokes equations. The EWT-TsAGI software, used for the flow computation, is based on the finite volume method of the second approximation order for all variables and includes monotonic modified Godunov scheme. The engine is simulated by the “active disks” method. Computations were performed on multi­block structured meshes with hexahedral cells. The power plant was designed with account for the initial requirements to the aircraft formulated in the AGILE project.

The developed optimization procedure consists of the two steps. At the first step, the isolated nacelle for the high bypass ratio engine is being developed and optimized for the cruise regime. Geometry of axially symmetric nozzle is described by the 11 parameters Parametric geometry of the inlet is specified by 7 control geometric parameters: 6 parameters specify the axially symmetric inlet, and one parameter (incidence angle) is employed for the air intake 3D design. The engine effective thrust is an objective function of optimization at the specified engine flow-rate constrains. To find the optimum solution, the Efficient Global Optimization method, based of simulation models, is used. It was shown, that SEGOMOE optimization method decreases the number of computed geometries.

At the second step, installation angles and the engines position over the airframe are optimized. A total of nine parameters is varied. The objective function is the effective thrust of the total layout (thrust minus layout drag) with the specified lift force constraint. An automatic structural mesh rebuilding is realized for the effective optimization procedure. The EGO based optimization algorithms require the initial points set calculating for the simulation model creation. It is shown, employing the large set of initial points (DOE) is more effective for the optimization process parallelization. Aerodynamic characteristics of the final layout with optimally installed engines were calculated. The main source of aerodynamic losses for the obtained configuration at the cruise flight’s Mach number of 0.85 is the compression shocks occurring due to the interference of the airframe with engine nacelle and between the neighboring engine nacelles. The subsequent studies should pay special attention to the aerodynamic interaction of the airframe and engine nacelles.

The described procedure was performed in the context of the third generation multidisciplinary optimization techniques, developed within the AGILE project. During the project, the new technologies were implemented for the novel aircraft configurations, selected as test cases for the AGILE technologies application.

Galkin N. A., Kondratenko A. N., Gaponenko O. V., Chiryukin E. V., Sviridova E. S. Methodical approach to aggregating computing of spacecraft manufacturing labor intensity. Aerospace MAI Journal, 2019, vol. 26, no 2, pp. 20-33.

For the purpose of the aerospace industry (AI) enterprises readiness to the implementation of State and commercial programs, it is necessary to perform an assessment of the production capabilities loading with regard to the labor costs for development efforts (DE) and spacecraft (SC) production.

The set task was being solved by the product capabilities conformity evaluation of the aerospace equipment (AE) head manufacturer with the federal target and government programs determining the required nomenclature and number of products, as well as the due dates of their production.

The spacecraft production is of a unit character with irregular repetition in the course of the years of production, where the products after the flight development tests (FDT) of the SC No 1 may have changes in the composition of the onboard equipment and design. The SC of manned programs production is individual and depends on the crew list and mission objectives.

Nowadays, based on the experience of the previous works and the prospective trends of development, engineers worked upon a number of unified space platforms (USP), which can significantly reduce the labor intensity of the SC manufacture. Development of the unified space platforms significantly reduces the volume and design cycles. In connection to the tried- and-true structural elements application the share of testing per one product set, which allows reduce the number of manufactured experimental installations.

The algorithm of SC manufacturing labor cost determining describes the sequence of labors costs computing of classification groups, containing tactical and technical characteristics of the products. The initial data on the actual and planned labor intensity of the SC production at the manufacturing enterprises were the products, both being manufactured and under development.

The first article of the stock-produced item manufactured for the flight development tests (FDT), at both single and several SC launch is assumed as a calculated labor intensity. The labor intensity calculation does not account for labor costs for the product manufacturing for performing inspection­sampling and periodical test.

The algorithm for the aggregating assessment of the SC production labor intensity is based on the layout solutions classification (constructive-technological schemes) of various types of SC. This algorithm has successfully proved itself within the framework of the “The SC Investments” research effort (RE) implementation, significantly increasing the accuracy of the loading prediction per product.

Calculation by the proposed algorithm is determined by a sufficient degree of technical solutions study at the stages of technical, draft and working projects, when analogous products, novelty factors or structural complexity of a new product can be determined.

Based on the obtained calculations, it is possible to evaluate and analyze the loading of the production capabilities of the main enterprise, specializing in the SC manufacturing. This will ensure the authenticity, completeness and estimation efficiency of the similar enterprises potential production.

Further development of this aggregating calculation algorithm of the DE and SV production labor intensity within the framework of assessing the feasibility measures of strategic plans for the technological development of the AI, the authors see in its automation. Besides, a coefficient characterizing technical level and industrial organization at the main manufacturing enterprises of the AE should be added to the algorithm. The proposed algorithm for the labor costs of SC production calculating was used by the center of integrated planning specialists of NPO “Technomash” in assessing the feasibility of the Russian Federal Space Program policy and the tasks of the Defense Procurement and Acquisition in 2017–2018, which confirmed its practical significance.

Calculated evaluation of labor costs for the SV production are recommended for employing as a basis for conducting technical and economic analysis, comparing alternative projects and developing perspective plans and programs. This labor input intensity algorithm will increase the accuracy of the enterprise predicted loading, resulting in the balance of the production program.

Kargaev M. V. Stresses computing in the main rotor blade based on the nonlinear loading model under static wind impact. Aerospace MAI Journal, 2019, vol. 26, no 2, pp. 34-42.

Wind is an important factor collateral to the helicopters operation. Due to a number of aeroelastic characteristics specifics, the non-rotating helicopter blades are sensitive enough to the wind impact. With this, the level of loads, acting on the blade, is commeasurable with the loads acting in flight. Traditionally, with high wind speeds mooring is employed to ensure the blades safety in parking position. It represents a flexible wire rope, which one end is fixed to the blade mooring unit, located as a rule at the blade tip section, and the other end is attached to the mooring node at the fuselage, or chassis of the helicopter. It represents a flexible wire rope, which one end is fixed to the blade-mooring unit, located as a rule at the blade tip section, and the other end is attached to the mooring node at the fuselage, or chassis of the helicopter.

The non-rotating main rotor blade according to its characteristics relates to flexible rods with deflections within the elastic deformations of the material commensurable with their length. This stipulates the necessity to consider the problem of the moored blade wind loading in a nonlinear formulation.

In this article, the parameters of the stress-strain state of the blade required for the mooring efficiency analysis are obtained based on a nonlinear model, which accounts for both geometric and aerodynamic nonlinearities. Computational algorithm for the initial nonlinear equation solution of the blade loading, developed based on the V.V. Petrov’s method of successive perturbation of parameters of was realized. The static loading is being considered as a process, developing at monotonous increasing of the loading parameter. The interval of load changing via its step- by-step application with small increments is split by steps, and for each step the linearized boundary value problem is being solved.

The blade deformed state, obtained in this manner at the current step, is assumed as the initial state for the next loading step. For error correction at each loading step, an iterative process is used, which allows performing calculations with a given accuracy.

The mooring effectiveness analysis was realized based on the computations performed for the moored and non-moored main rotor blades of the Mi-8 helicopter. The article presents the dependencies of critical gliding angles and limiting, under the strength condition, wind velocities values corresponding to them.

The article presents the dependencies of critical gliding angles and corresponding to them limiting, under the strength condition, wind velocities values. It also presents the dependencies of limiting velocities at the condition of a swaying absence condition on the characteristic section installation angle for the modes of blowing from both front and rear edges. The optimum installation angle, at which the range of safe wind speeds for the main rotor as a whole was the largest, was determined. This allows recommending to set the angle of the total step equal to the optimum one while a helicopter parking.

Alekseev V. V., Bobrov A. N., Kalugin K. S. Study of complex strength characteristics of gas turbine odels fabricated by additive methods. Aerospace MAI Journal, 2019, vol. 26, no 2, pp. 43-50.

Recently, the studies related to the additive technologies application in various industries, including aviation and space-rocket mechanical engineering, are considered promising. An indisputable advantage of additive technologies is minimization, and, in some cases, complete elimination of the need for parts machining, which significantly reduces both the time consumption and the finished part cost.

There are several basic 3D-printing methods, differing in the source material and technology of the parts formation. Recently, the parts production by selective laser sintering of metal polymer compositions powders (SLM-printing) has become topical.

The SLM-printing technology consists in layer-by layer deposition and sintering of powder on a special substrate. However, application of the selective laser powders sintering method is associated with problems of the porosity formation and a decrease in the strength of the parts produced. Thus, the issue of practical application for parts of the space-rocket and aviation equipment, created by the 3D-printing, still remains open.

To substantiate the possibility of 3D-printing application in turbines production for laboratory test benches on compressed air, the strength calculation of the turbine from PLA-plastic printed on the 3D printer were performed. The tests were performed to confirm the calculations results.

When developing a turbine 3D-model the rotor wheel geometry was selected, based on the prototype, which was used in the turbine structure employed in the laboratory test bench installation at the BMSTU for the laboratory works for studying the energy characteristics of active turbines.

Besides the external loads, the gas turbines rotor wheels load-bearing capacity is affected by loading conditions, such as gas temperature. However, the gas turbines employed in laboratory work benches on the compressed air are operating, as a rule, at low operating temperature of 30-50°C. Thus, the temperature stresses may be neglected while strength calculations of the turbine disk.

A 3D-model of the turbine under test was built with the Autodesk Inventor program. A finite-element model containing about 4.15 million elements was built for the above said model. Its strength analysis was performed with the Autodesk Simulation Mechanical 2019 module. The mesh thickening was reduced to the base of one blade only, since the load distribution is symmetrical. It can be seen from the safety factor distribution fields that minimum safety factor corresponds to the root sections of the blades, and it is no less than 3.3.

While theoretical calculations the modified safety factor n1, accounting for the effect of the part material porosity (for the case of its manufacture by 3D­prototyping) through coefficient k, was 3.28.

For tests performing, an axial active supersonic gas turbine was manufactured from PLA-plastic according to the SLM-printing technology.

For tests performing, a test bench, consisting of an electric motor, a voltage regulator, a tachometer, a video camera, as well as a turbine under study was assembled.

The methodology of the experiment conducting is as follows: the turbine is fixed on the motor shaft by the keyed and glue joints. When the motor is connected to the mains (220 VAC), the shaft and the turbine begin rotating. The rotational speed is changed by a voltage regulator connected to the motor circuit, and can aquire values from 0 to 24000 rpm, which corresponds to the voltage range in the motor network from 0 to 220 V. The data on the motor rotational speed are read from the digital optical tachometer. The experiment is being shot by the video camera.

The strength calculations of the axial supersonic gas turbine fabricated from the PLA-plastic by the SLM-printing additive technology revealed that the safety factor in operation conditions of laboratory test benches with compressed air was higher than the maximum allowable one for the considered unit.

As a confirmation for calculations, the turbine rotational speed during the test reached 24,000 revolutions per minute, which is the maximum possible value for the engine used in the tests. With this, visible defects were not detected in the turbine itself.

On the assumption of the performed studies it was established that the turbine manufactured using additive technologies can be employed for the laboratory text benches operating on compressed air.

Pronin M. A., Ryabykina R. V., Smyslov V. I. Experimental study of the aircraft forced vibrations while the engine blade break-away. Aerospace MAI Journal, 2019, vol. 26, no 2, pp. 51-60.

The presented article is a generalization of works relating to the ground reproduction of the force impacts on the aircraft structure, on the part of the engine with imbalance in case of the blade loss.

While ground testing the engine rotor does not rotate, and rotating force is formed by the fixedly installed vibration exciters. The immediate purpose of the experiment consists in frequency characteristics measuring, which associate the aircraft vibrations with the excitation force from the engine rotor imbalance. These characteristics are necessary for the computational dynamic scheme correction of the structure employed in loads computing in flight, possibly prolonged, while the blade break-away over the water surface. These computations are used for the aircraft safety evaluation while the blade loss.

The article presents the testing technique and facilities. The estimates of the modelling method applicability and its trustworthiness are given for the first time. The text is supplemented by the examples of real data of the tests.

The quantitative confirmation for the case of the ground experiment is given in the applicability esteems of the rotating inertial force reproduction by the harmonic forces stationary in space. At the same time, it was noted that the loads calculation while flight fluctuations, with a high level of the engine overloading, can not be based on either use of only relative acceleration of the blade, or the approximate theory of the gyroscope.

The circumstance of the experiment performing while the compulsory routine tests prior to its first flight was considered separately as practically the only possible for the experiment under consideration. The domestic tests on the aircraft with the engine blade loss modelling performed for the first time revealed the feasibility and possibility of their realization in conditions of dire time deficit prior to the first flight.

The presented details and features of the technique allow apply them in the future in the practice of such tests by the design bureau itself.

The main result is substantiation and practical confirmation of the possibility of reproducing on the ground the forced oscillations of an airplane after the blade loss, and while the mandatory regular modal tests.

Avdeev A. V., Katorgin B. I., Metel'nikov A. A. Energy characteristics computing technique for mobile multifunctional laser power plants based on fiber lasers. Aerospace MAI Journal, 2019, vol. 26, no 2, pp. 61-69.

Multifunctional Laser Power Plant (MLPP) should simultaneously solve the tasks of energy generation (Power Supply System (PSS)), radiation conversion and transmission (Laser System (LS)), and heat removal (Thermal Mode Supporting System (TMSS)). Meanwhile, the above said tasks are duly elaborated in modern projects. Thus, it is necessary to develop the MLPP design methodology, which accounts for the above listed subsystems interaction.

The article presents the developed technique for parameters analysis of the LS, TMSS and PSS subsystems of a multifunctional laser power plant, and results of its approbation while solving the task of space debris removal.

Computing was performed for the initial data Xtask based on the analysis presented in [1–5, 8]:

  1. acting on the Space Debris Fragment (SDF) with the orbit of HSDF = 1000 km by the ΔhSDF value required to its descent to [50; 900] km;

  2. the FSD velocity change per one pulse ΔFpulse of [0,1; 1,6] m/s;

  3. the impact distances range of RySDF [10; 150] km;

  4. the height difference of the SDF and spacecraft (SC) orbits of Horb [0; 150] km;

  5. relative FSD and SC closing-in velocity of Vrel [10,8; 12] km/s.

The following requirements to the MLPP operation mode (Υmode) were obtained for the initial data presented above: the energy density of [2,5⋅104; 2,5⋅105] J/m2 at the SDF; pulse duration of [2,7⋅10-9; 2,7⋅10-7] s; FSD exposure time of [2; 28] s; pulse frequency of [1; 1250] Hz.

The requirements to the sub-systems performance for this mode are as follows:

  1. LS (XLS): the output aperture dimensions of [0,5; 3] m; M2 and λ LS are assumed equal to 1 for calculations simplification; efficiency is [0.31, 0.59]; the laser pulse energy of [3⋅105] J; the threshold pulse power for one channel of 4,2⋅106 W; the beam strength of fiber of [0,01; 0,08] J.

  2. Requirement to the PSS generated energy is NPSS = [0,87; 5,7⋅108] W.

  3. The energy removed by TMSS is NTMSS = = [0,5; 4,5⋅108] W.

As a result, the inference cam be made that the data obtained while the technique application allow perform the MLPP parameters analysis for selecting the types of PSS, TMSS and their parameters, necessary for the MLPP required operation mode. Besides, this technique allows determining the limitations imposed by the PSS and TMSS subsystems on the LS pulse energy. The presented technique may be employed for the integrated assessment of the subsystems parameters and recommendations development of the MLPP application.

Vetrov V. V., Morozov V. V., Kostyanoi E. M., Os’kin A. S., Fedorov A. C. Caliber air-intake device for a flying vehicle with rocket-ramjet engine. Aerospace MAI Journal, 2019, vol. 26, no 2, pp. 70-80.

The work is devoted to the caliber air-intake device development for an aircraft with a rocket-ramjet engine moving in the dense layers of the atmosphere.

Analysis of the trends in the near-range aircraft with active start development demonstrates that one of the main directions of their improvement is the flight range increase The mass-size characteristics of the aircraft herewith remain at the same level, which does not allow employ the extensional development trends. Under these conditions, an important place is ranked by the trend related to the rational onboard energy utilization, within which framework the already classical solution are employed. However, the potential of these solutions is currently close to its limit.

In this regard, special attention is paid to propulsion systems (PS), which energy capabilities can be improved through the atmospheric air employing, and to a rocket-ramjet engine (RRE) in particular.

One of the key elements that largely determines the rocket-ramjet engine efficiency in total is the air-intake device (AID).

The proposed work novelty lies in the fact that the guided artillery shell (GAS) with its specific layout and functional features is considered as the object of study, and the search for a reasonable compromise between the requirements for the propulsion system and the shell as a whole is performed.

The problem of the AID rational configuration is being solved complexly based on the combination of numerical modelling methods and wind tunnel tests.

The initial variant of the twelve-nozzles caliber AID was developed for the pilot studies.

The works aimed at obtaining the throttle characteristics were performed.

One of the key features of the AID initial version was low efficiency of the boundary layer drainage system, which negatively affected its characteristics. In this regard, the initial model was modified to the second and later to the third option, characterized by an increased area of drain channels.

A positive result, manifested in an increase in the coefficient of the total pressure restoration by 14-20%, and the coefficient of air consumption by 11-27% for the third option, allowed form priorities for the subsequent AID configuration with a modified boundary layer discharge system and boxlike nozzles.

This solution allowed maintaining the aft location of the caliber non-regulated AID and the power plant with moderate total pressure losses and more stable air intake operation.

The performed studies allowed soundly obtain the most rational option of the caliber four-nozzle non­regulated AID for aft located RRE, integrated into the GAS structure. According to the preliminary estimates, this solution ensures provides a flight range increase by 25% compared to the GAS, equipped with the solid engine and bottom gas generator.

Osipov I. V., Remchukov S. S. Small-size gas turbine engine with free turbine and heat recovery system heat exchanger within the 200 HP power class. Aerospace MAI Journal, 2019, vol. 26, no 2, pp. 81-90.

The article presents a preliminary study of a small- size gas turbine engine (SGTE) of the 200 HP power class with a free turbine (FT) and a heat exchanger (HE) of the engine exhaust heat regeneration system. The presented engine is being developed primarily for unmanned aerial vehicles of various types and purposes (helicopters and airplanes).

The engine is available in two versions, namely, without a heat exchanger of the heat regeneration system, for the aircraft with short range and flight duration, and with a heat exchanger for the aircraft with long flight duration.

Characteristics calculations were performed for both the TSEr-200 engine with complex heat regeneration cycle and for the TSE-200 engine without heat regeneration [5].

Computational studies on sizes and type of the recuperative heat exchanger, rational for the given problem, were performed while the TSEr-200 engine development. A bundle of tubes was employed to determine basic dimensions of the heat exchanger matrix, on the assumption of the preliminary computation convenience (as the most worked out) [6].

The design arrangement of the heat exchanger and gas genera The structural layout of the heat exchanger and gas generator was developed based on the primary matrix computations.tor was developed based on the primary matrix computations. The heat exchanger includes 12 separate modules interconnected by the common manifold. Each matrix module is placed in individual casing.

Computational studies of various plate matrix types, as the most technologically worked-out at present and less expensive, were performed after the general layout developing. These computational studies were performed with the Ansys software package [11] using existing techniques for gas dynamic flows computing [12-15]. The computation results revealed significant hydraulic losses in the place of the flow turning inside the heat exchange matrix. Analysis of the results led to the necessity of studying the one- pass scheme of the coolant movement.

Computational studies of the heat exchanger option with the one-pass flow scheme revealed that total hydraulic losses for coolants did not exceed 3%. However, the layout of the heat exchanger with the engine was changed to organize the return of the air, preheated in the heat exchanger, to the combustion chamber. A distinctive feature of the proposed layout of the heat exchanger with SGTE is that the heat exchanger consists of 8 unified blocks, arranged in a circle among the three manifolds: the front one and two rear ones. All manifolds are cast and they are bearing elements of the engine.

For further work on the heat exchanger of the TSEr-200 engine, an option of the matrix with the “Frenkel packing” type plates of a single-pass scheme was adopted.

To confirm the feasibility of the heat exchanger project for the TSEr-200 engine, a matrix of the demonstration version of the heat exchanger with the “Frenkel packing” type heat exchange surface was developed. The module will be tested on the CIAM universal test bench as a part of the demo small gas turbine unit with the 4 kW capacity.

Ezrokhi Y. A., Khoreva E. A. Studying criterion parameters of the total pressure input non-uniformity impact on the thrust of a turbojet engine. Aerospace MAI Journal, 2019, vol. 26, no 2, pp. 91-98.

The presence of the total pressure non-uniformity may affect the basic engine parameters, and, in the first place, its gas-dynamic stability margin, as well as thrust-economic characteristics. Circumferential non­uniformity of the total pressure and its non-stationary component greatly affect the engine gas-dynamic stability. As for the engine thrust, the radial and circumferential effects are close enough, and non­stationary component does not affect the engine thrust at all. It allows employ one-dimensional approaches while this phenomenon modelling, and consider the impact of both stationary components of non–uniformity of the total pressure (both circumferential and radial) from the single methodological positions

In case of a non-uniform input flow, the flight-thrust decrease occurs for to the several reasons. Reduction of the general level of the total pressure along the engine passage, which leads to the pressure drop reduction in the jet nozzle pressure difference and, correspondingly, the decrease of the engine specific thrust may be assigned to the first cause. Besides, due to the general level of the input pressure reduction, physical air consumption reduction through the engine occurs as well.

The second reason of flight thrust reduction is associated with additional total pressure losses due to the “wash-out” of areas with various level of the total pressure in compression elements. It leads to the additional losses of the total pressure in compressor stages, which reduces the aircraft engine thrust to an even greater degree.

The authors suggested and justified criterion parameter Er for correct estimation of the thrust- economic parameters of the engine, operating in conditions of non-uniform input field of the total pressure. To the contrary of the W parameter, this parameter reflects additionally the relative values of the area, occupied by the zones with various total pressure values, being conditional indicator of the reduced pressure “concentration” per unit of the input area.

On a calculation example of the one-shaft turbojet with sufficiently conservative level of the design parameters the effect of the total pressure non–uniformity on its key parameters, such as thrust and gas-dynamic stability margin of the compression system was considered. This kind of engine selection is explained by the fact that to the contrary of the bypass jet engine, considered in the previous articles, the non-uniform field at the turbojet compressor inlet is considered as known, and its impact on a single compressor would be determinant for the whole turbo jet engine.

The performed calculation estimations revealed that the decrease in the engine thrust δR due to the non-uniform field of the total pressure at the inlet was completely defined by value of this parameter (dependence between δR and Er is almost linear), and also by the engine operating mode, such its shaft rotation frequency.

Zuev A. A., Nazarov V. P., Arngol’d A. A. Determining local heat transfer coefficient by a model of temperature boundary layer in gas turbine cavity of rotation. Aerospace MAI Journal, 2019, vol. 26, no 2, pp. 99-115.

Accounting for heat transfer specifics in flow­through parts of turbo-pump assemblies of liquid rocket engines (LRE) is a topical task. Currently, accounting for the specifics of the flow with heat transfer while realizing both potential and vortex rotary flow in the flow-through parts is implemented generally by the following methods: employing empirical equations, numerical and analytical methods for solving partial differential equations [1].

High temperatures of the working fluid lead to thermal deformations of components, including the turbine disks [18]. When designing the flow-through parts of the LRE turbo-pump units and assemblies, it is necessary to account for the temperature change of the working fluid flow along the working channel, since the viscosity parameter is a function of temperature and determines the flow regime and, as a result, losses, particularly disk friction and hydrodynamic losses in the flow-through part. The LRE turbo-pump energy parameters modelling is a topical scientific and technical task. The issues of the workflow parameters optimization, and the propulsion system mathematical model were reviewed in the V.A. Grigoriev’s treatise [19], where analysis of the models was performed, and merits and demerits for various design stages were disclosed.

A model for dynamic and thermal spatial boundary layers distribution with convective component for the combustion products turbulent flow in the LRE gas turbines rotation cavities is proposed. For combustion products, the Prandtl number is less then unity (Pr < 1), and dynamic boundary layer thickness is less than the thermal boundary layer one. It was assumed, that the temperature change and thickness of energy loss within the dynamic boundary layer border occurs due to the dynamic velocity transfer, and beyond the border – due to thermal conductivity only. This assumption complies well with the inferences of many authors [20, 21, 24]. Thermal resistance manifests itself over the entire thermal boundary layer thickness. Thermal resistance exists within the dynamic boundary layer borders due to the turbulent heat transfer, and beyond the border – due to thermal conductivity [24]. The distribution model of the dynamic and thermal spatial boundary layers with convective component is necessary for analytical determination of the local heat transfer coefficient in the LRE turbines rotation cavities.

The main objects of research, where the potential and vortex rotational flow is realized, are the flow­through components of LRE gas turbines such as inlet and outlet devices, as well as cavities between the stator and the working wheel [20].

An integral relation for the thermal spatial boundary layer energy equation, allowing integration over the surface of any shape, which is necessary for determining the thickness of energy loss, was obtained. The expressions for determining the energy loss thickness for thermal spatial boundary layer are necessary to determine the local heat transfer coefficients for the typical flow cases with account for the heat exchange.

Expressions for determining the local heat transfer coefficient in the Stanton number form for the straight linear uniform flow, rotational flow according to the rigid body law, and rotational flow of the free vortex of a power profile distribution for dynamic and thermal boundary layers parameters in case of Pr < 1 were obtained analytically.

Local heat transfer coefficient in the Stanton number form for straight linear uniform turbulent flow is

where m — is the turbulization degree of spatial boundary layer dynamic velocity profile,

– is the dynamic and thermal boundary layers ratio of the thickness, λ — is the coefficient of thermal conductivity,

 – the laminar sublayer coefficient of turbulent velocity distribution profile (obtained considering the two-layer turbulence model with a viscous laminar sublayer), Re — the Reynolds number.

Local heat transfer coefficient in the Stanton number form for rotational flow according to the rigid body law is

where ε — is the angle tangent of the bottom streamlines bevel, J — is the relative characteristic thickness.

Local heat transfer coefficient in the Stanton number form for rotational flow of a free vortex is

Analytical expressions for heat transfer coefficients agree well with the experimental data and dependencies of other authors [7–10].

The obtained analytical expressions well agree with the data of other authors and are necessary for engineering calculations while designing the LRE flow-through parts of turbo-pumps.

Baklanov A. V. Experimental study of the flame tube temperature state of a gas turbine engine multi-nozzle combustion chamber. Aerospace MAI Journal, 2019, vol. 26, no 2, pp. 116-125.

The flame tube walls cooling is one of the important components while organizing processes in the gas turbine combustion chamber. The combustion chamber operation reliability and engine endurance as a whole depend on the effective flame tube walls cooling. Convective-film cooling is one of the most widespread cooling systems. It includes the air film forming, which does not allow the hot gas interaction with metal and drawing heat from the backside of the wall due to the convection. The article presents the results of the studies on the flame tube walls temperature determining of the gas turbine engine operating on the gaseous fuel.

The article presents the combustion chamber structure of the converted aviation gas turbine engine serving as the gas pumping unit supercharger drive. The combustion chamber walls preparation and its testing as a part of a gas turbine engine were performed. The article presents the results on the flame tube walls temperature for the two operation modes of the gas turbine installation corresponding to 16 and 18 MW. The analysis of the obtained results allowed revealing that with the gas turbine installation power increase from 16 to 18 MW the temperature state of the wall did not drastically change. The walls temperature at the considered modes does not exceed 800°С, which indicates the flame tube sufficient cooling. However, the temperature distribution in various cross-sections was not of the similar nature. In some cross-sections maximum compared to the other cross-sections temperature was observed. It can be explained by the fact that the air passed through the conduit is split upon the hole flanging forming a vortex flow. As a result, the film-cooling loses its effectiveness, and the wall temperature behind the hole increases. The film-cooling effectiveness was determined at various sections on the flame tube walls. A technique for the wall temperature computing was developed, and comparison of computational and experimental results was performed.

Semenova A. S., Zubko A. I. Studying technical condition of the interrotor bearing with the SP180-M vibratory-diagnostic test bench after passing life tests. Aerospace MAI Journal, 2019, vol. 26, no 2, pp. 126-138.

The presented article deals with the studying of roller bearing after accelerated life test for the resource of 2000 hours.

To analyze the 5АВ1002926Р4 bearing vibration state a cpmprehensive analysis was being performed, including spectral analysis, RMS analysis in low-, medium- and high-frequency ranges, analysis of a pick-factor in low- and high-frequency ranges, and analysis of a “raw” signal of records.

The obtained test results allowed evaluate the bearing technical condition and transfer to further life tests with the test bench at “CIAM named after P.I. Baranov”.

It is well-known, that machines and mechanisms reliability depends essentially on their bearing assemblies working capacity. It is especially important for aviation engines as their bearing assemblies are one of the most responsible units often limiting an engine resource.

A reliable estimate of roller bearings technical condition, applied in gas turbine engines presents a problem at the aircraft building enterprise while both manufacturing and incoming inspection and fault detection. It concerns especially the indecomposable bearings since their technical condition estimation system currently in force is based mainly on the subjective methods such as checks on ease of rotation, or noise. Thus, the instrumental control methods implementation allowing not only estimate, but also forecast the working capacity during the operational process with more fidelity, is of current interest.

One of such instrumental methods is the quality monitoring of bearings vibration characteristics (a method of vibration diagnostics), operating with the specified loadings and frequencies of rotation. For vibrations measuring the vibrational converters, i.e. seismometers or accelerometer are used.

Methods of bearings vibrations measurement at control test benches are defined by the Standards [4, 5, 6]. The bearings condition is defined through the analysis of vibration signals [7].

Currently, various test benches, installations and diagnostics complexes, realizing this technique, have been developed, and being manufactured. One of them is the SP-180M test bench for roller bearings incoming inspection, being produced by LLC “Diamekh”. The test bench is meant for experimental studies for technical condition evaluation of separate bearings by vibration diagnostics method. These are the bearings of the first category (new), and bearings of the second category (being reinstalled), being installed in the engine while assembling.

The roller bearings, depending on the structure specifics of the product, where they are employed (parameters of inertia, stiffness and damping) may generate vibration of various intensity at various frequencies.

The vibration sensors mounting location and their characteristics significantly affects measuring results.

Thus, the SP-180 test bench has the single-type fixing of bearings, and fixed position of vibration sensors

The vibration signal amplitude, generated while interaction of working surfaces and external and internal rings of the bearing will depend on the rotational frequency of the test bench. Thus, its operating frequencies have the specified values.

Krylov A. A., Moskaev V. A. A technique for fluoroscopic control and analysis of technical condition of aircraft structural elements with honeycomb filler. Aerospace MAI Journal, 2019, vol. 26, no 2, pp. 139-146.

Application of various non-destructive testing (NDT) methods and means in conditions of operation is an effective method for sustaining the required reliability of aerotechnics. The structures with honeycomb filler from aluminum, steel and titanium alloys are employed in the modern aircraft airframes elements. Currently, x-ray method is the most effective one for such structures inspection. The article covers the non-destruction inspection technique performing of the aircraft structural elements with the honeycomb filler, and estimation of the images obtained by the fractal analysis.

The proposed technique consists of three main blocks:

1.   The block forming initial data, restrictions and assumptions:

a)   Variable parameters of the fluoroscopic installation (“Norka” X-ray TV unit);

b)   Invariable parameters characterizing design specifics of aircraft or control object (CO).

2.   A block of the fluoroscopic control methodology of aircraft design elements with honeycomb filler:

a)   A model for images base formation with account for the fluoroscopic installation parameters adjustment:

-     The CO X-raying schemes elaboration;

-    forming the images base when changing the anode voltage value at the emitter and the distance from the emitter to the CO. The best picture of the element with a honeycomb core was obtained in the framework of the experiment at U = 50 kV; F = 90 cm (F is a focal length, U is the anode voltage);

b)     A model for the image quality assessing:

-    Expert evaluation of the images database, with the concordance coefficient calculation [3];

c)     The CO fault detection performing:

-    Parameters adjustment of the “Norka” X-ray TV unit according to the image quality assessment model;

-    The CO fault detection according to the X-raying scheme;

-    The fault detection results decoding and analysis by fractal analysis.

3.   Recommendations formation on fault detection and repair of aircraft structural elements with honeycomb filler.

Fractal dimensions of the honeycomb filler without defects and the one with defects (the presence of moisture and geometry violation of the honeycomb filler structure boundaries) were obtained applying FracLab software.

The result of fractal dimension computing was obtained using the FracLab program by the direct geometric method of counting the cells of the honeycomb filler structure without defect and the one with defect.

The graph deviation of the structure with a defect from the linear dependence, characterizing the self­similarity of the structure under study, is twice as large as on the graph without a defect. It indicates the boundaries structure violation of the honeycomb filler. In addition, the graph with a defect in the double logarithmic coordinates has a kink, characterizing transition between different types of the structure (liquid presence in the honeycomb filler).

The additional information on the state of the system under study can be extracted by determining the self-similarity ranges limits.

Thus, employing the fluoroscopic control technique will allow performing the fault detection inspection of the aircraft structural elements with the honeycomb filler based on fractal analysis, as well as analyzing the obtained images base, and trace the dynamics of the honeycomb filler parameters changes, and defects of its internal structure, while the aircraft operation. However, it should be noted that the fractal analysis may be employed in the long term for automated parameters adjustment of the “Norka” X- ray TV unit, and the images base decoding without an operator.

Levochkin P. S., Martirosov D. S., Kamenskii S. S., Kozlov A. A., Borovik I. N., Belyaeva N. V., Rumyantsev D. S. Liquid rocket engines functional diagnostics system in real-time mode. Aerospace MAI Journal, 2019, vol. 26, no 2, pp. 147-154.

The hardware-software complex of the functional diagnostics system of the liquid jet engines operation during fire tests was developed. The system analyzes data in the real time mode. It deals with troubleshooting of units, structural elements or loops of a liquid rocket engine and determines the time instant of their occurrence.

Theoretical studies of the processes occurring in a rocket engine have been conducted since the 1930s. Differential equations reflect the dependencies between the engine parameters. The developed system employs the linearized equations of dynamics allowing accelerate computing and obtain numerical results in the real-time mode.

Each engine and each of its units are described by mathematical equations, on which basis the parameters values are calculated.

At each stationary mode, the averaged values of the operating engine measured parameters computed employing a mathematical model are compared.

If a calculated value deviates from the actual one, then there is a considerable probability of a defect presence in a unit, or in the entire engine. Functional diagnostics is based on this principle.

Modern measuring systems and high-speed computing systems are employed to diagnose engines in real-time mode.

The system consists of a hardware-software complex, an information system and a database, a telemetry signal emulator and an operator’s automated workplace.

The LRE functional diagnostics system solves the following tasks:

1. Increases the safety of the LRE fire tests conducting;

2. Determines the the engine functioning correctness in all stationary modes specified by the test profile;

3. Detects and localizes the malfunctions disrupting the proper functioning;

4. Identifies the engine “weak points”, such as elements or loops prone to structural or manufacturing failures.

5. Confirms the engine reliability before prior to its employing as a part of the launch vehicle.

The results of the emergency protection system and functional diagnostics system operation were compared. The proposed system has always found a failure before the emergency protection system did.

Petukhov V. G., Zhou R. . Computing the perturbed impulse trajectory of transferring between the near-earth and near-lunar orbits by the continuation method. Aerospace MAI Journal, 2019, vol. 26, no 2, pp. 155-165.

The problem of computing a two-impulse flight between circular near-Earth and near-lunar orbits with specified altitudes and inclinations over a specified time is considered. A mathematical model of motion, accounting for the Earth, Moon and Sun attractive forces as point masses and the second zonal harmonic of the Earth gravity potential at all spacecraft movement sections is used. The first velocity impulse is formed at the initial near-Earth orbit, and puts the spacecraft on the lunar flight trajectory. At the Moon passage instant at the minimum distance the second impulse is formed putting the spacecraft on the near-lunar orbit.

A numerical method for calculating two-impulse transfer between the circular orbits of the Earth and the Moon for a fixed time with account for the main perturbing accelerations has been developed. The method consists of the procedure for calculating the guess values, using the method of point-like spheres of impact, and the procedure for solving the boundary value problem for calculating the perturbed flight trajectory using the continuation method for reducing the boundary value problem to the Cauchy problem.

The advantage of the developed method is the procedure automation for selecting the initial guess values for solving the boundary value problem, and the computational stability of the solving process of the boundary value problem itself. The method revealed its efficiency and computational stability when calculating a series of transfers to a polar circular low lunar orbit of an artificial lunar satellite for various start dates and flight durations. The developed method may be applied for the design-ballistic analysis and operational planning of prospective lunar missions.

The article presents the numerical examples of trajectories computing for the flights between the low near-Earth and near-lunar orbits. Computing of the series of such trajectories allowed calculate the optimal start date and optimal flight duration, as well as dependencies of the required velocity impulses and longitude of the ascending node of the near-lunar orbit on start date and flight duration.

Nikolaeva E. A., Starinova O. L. Application of a heavy spacecraft with low-thrust engines for asteroid deviation from a dangerous trajectory. Aerospace MAI Journal, 2019, vol. 26, no 2, pp. 166-174.

The problem of asteroid danger for the Earth has long enough attracted the attention of scientists and society. Studying the traces of the space originated catastrophes on surface of the Earth and celestial bodies, as well as observing asteroids in the near-Earth space reveal the seriousness of asteroid hazard for the Earth civilization and the necessity of developing measures for its prevention.

The studies related to the issues of asteroid hazard encompass several trends.

Above all, detecting dangerous asteroids approaching the Earth (AAE) and their orbits determining. Currently, there are several national programs for optical observation of such bodies (NASA, LINEAR, ESA). It is assumed that these programs allowed detect great majority of such bodies with the size order of a kilometer or more. A whole number of such studies and projects envisage the countermeasures against these outlanders by their changing orbits or their destruction into small splinters, burning down in the atmosphere.

The urgency of the asteroid danger overcoming is beyond doubt at present, and the developing measures for its prevention should be one of the most important tasks to be solved by the humankind in the 21st century.

The goal of the presented work consists in developing a mathematical model, simulation and effectiveness analysis of the Earth protection systems to overcome the asteroid danger by the gravitational tractor.

To achieve the set goal, the following tasks were solved:

1)   Studying parameters asteroids approaching the Earth;

2)   Developing mathematical models of the joint motion of asteroid and all the bodies involved in the process of deviation from the dangerous trajectory (Sun, Earth, spacecraft, asteroid);

3)   Developing a software package, ensuring simulation and visualization of the proposed method of the asteroid danger counteracting;

4)   Analyzing the simulation results of the proposed method of the asteroid danger counteracting.

The main results obtained in the work are as follows:

-     a mathematical model of the motion of bodies, with perturbations from the gravitational tractor acting on them: a variable mass asteroid, spacecraft, the Earth and the Sun, with account for the gravity of all bodies;

-     based on a a mathematical model of the bodies motion system, the software package “Simulation of the Earth protection systems functioning to overcome the asteroid hazard” for the asteroid trajectory simulation by the selected method of the asteroid danger overcoming in heliocentric coordinate system was developed;

-     simulation of the potentially hazardous bodies deviation method (asteroid deviation by the gravitational tractor) for the 99942 Apophis asteroid was performed with the developed software complex “Simulation of the Earth protection systems functioning to overcome the asteroid hazard”;

-   the simulation resulted in obtaining the flight trajectories of all the bodies of the system under consideration (the Earth, the Sun, asteroid and a spacecraft) and heliocentric movement parameters;

-   the efficiency analysis of the selected method was performed.

Ermakov V. Y. Studying the effect of the beam aerial drive control algorithm on its vibration activity onboard a spacecraft. Aerospace MAI Journal, 2019, vol. 26, no 2, pp. 175-181.

Modern space vehicles (SV), as a rule, include bearing-out structures of slight rigidity. These are solar batteries, antenna-feeder devices, elements of thermal conditioning systems. Actuators and special purpose units, as well as units of technological and support systems are being placed inside the SV hull. SVs are exposed to vibrations from the external and internal perturbance sources both on Earth and in orbit. The feature of the SV loading in orbit is low-force spectrum of perturbances up to tens of Newtons with frequencies from fractions of hertz to hundreds of kilohertz. Vibrations may have deleterious effect upon both orientation and stabilization accuracy, and movement dynamics including various types of orbital maneuvering. These perturbances might be created, for example, by operation of the narrow-beam aerial (NBA) drive, which leads to occurrence of elastic vibrations of the structure and mounting faces of the precise equipment. While the observation session onboard an SV, mechanical disturbances, stipulated by operation of aggregates with non-balanced masses, may occur. This may affect both the orientation accuracy of the SV itself and equipment elements which may degrade the quality of the registered information, and introduce significant error to the SV angular position measurements, obtained by the orientation and stabilization control system. This, in turn, may make the SV mission target task performance impossible. To reduce these perturbances an algorithm for the NBA drive operation for the “Spectr-R” type SV was developed. Dynamic analysis of data obtained for the suggested algorithm and conventional was performed. Positive results of the suggested algorithm, tested on the “Spectr-R” type SV are demonstrated.

Silaev M. Y., Es’kova E. A., Gerus D. S., Remshev E. Y. Acoustic emission method application while determining mechanical characteristics of the brnicrsi-2,5-0,6-0,7 wire for elastic elements production. Aerospace MAI Journal, 2019, vol. 26, no 2, pp. 182-192.

A great number of electromechanical systems, an important part of which represents an elastic element from bronze, are applied in aerospace technology.

Severe requirements are placed to the physico- mechanical characteristics of these parts. The existing standard methods for mechanical properties determining are not sufficient for such products.

Acoustic emission method is one of the promising methods to solve this problem.

Acoustic emission is radiation of mechanical waves by the material, caused by local dynamic rearrangement of its structures This method is non destructive.

Beryllium bronze is used as a rule in special products. This project studies a cheaper substitute for Nickel-chromium-siliceous bronze.

Besides, mechanical tensile tests of the wire with parameters registration of acoustic emission were being conducted. Bronze was subjected to various heat treatment to select the optimal mode.

As the result of this work, the microstructure of the samples was studied for various thermal treatment modes. It was revealed that the acoustic emission parameters were the figures of strength and plasticity.

The strength and plastic characteristics are related to the grain size by the dependence proposed by Hall- Petch. This dependence modernization allowed adopt the stress at maximum value of the pulse amplitude up to the yield point achieving as the stress corresponding to the dislocations motion start.

The possibility of determining the microplastic deformation starting of wire samples by AE method was established. Based on the obtained regularities, it was revealed that the number of signals is a characteristic of strength, while the amplitude is a characteristic of plasticity. The Hall-Petch dependence modernization may allow developing a technique for operational control of microstructure in the release of special products.

Kovalev A. A., Zinova V. V. A tool-blank state monitoring while cutting process using kalman filter. Aerospace MAI Journal, 2019, vol. 26, no 2, pp. 193-204.

The article discusses the issue of the cutting process monitoring possibility using the acoustic emission method by processing the input signal using the Kalman filter. A filter was selected to solve the problem. The inference was drawn on the possibility of monitoring the gradual wear-out and chipping of the cutting edge by Kalman filter.

The article consists of three main parts: introduction, the main part, and conclusions.

The introduction considers the problems occurring while automating the technological process of blank parts machining. With this, a part of events is deterministic, while the other part is random. Thus, to ensure the required quality level in the process of automation the cutting zone continuous monitoring is required. It will allow making changes directly while blank parts machining technological processes executing.

The main part of the article presents operation principles of the monitoring systems, based on the

system harmonic oscillations analysis. Various filtering algorithms were considered in particular.

The Kalman filter was chosen as the object of study as one of the most common algorithms in the theory of automatic control. The goals were set and the tasks were formulated. Criteria are being set, which the desired filter should meet for continuous for the cutting area monitoring. The main approaches to solving filtering problems are being considered and compared with the Kalman filter. The inference is being drawn that this filter is the most suitable for solving the set problem. Measurements are being performed, the results, processed by the three Kalman filters versions are being analysed, and one of them, best meeting all the necessary requirements is being selected.

The conclusions formulated the possibilities for Kalman filter application for continuous monitoring of the tool blank state in the cutting process and gave recommendations to the future work, and filter coefficient selecting in particular.

Khryashchev I. I., Danilov D. V., Logunov A. V. Developing a sparingly doped high-temperature nickel alloy for gas turbine blades. Aerospace MAI Journal, 2019, vol. 26, no 2, pp. 205-218.

Development of mono-crystal high-temperature nickel alloys for gas turbine blades and vanes is one of the leading trends ensuring enhancement of parameters, efficiency and reliability of modern gas turbines.

Currently, one of the most widely used alloys for turbine blades manufacturing is the second-generation domestic ZhS32 alloy with Re content of about 4.5%. The goal of this work consists in the alloy creation with the equivalent level of heat resistance, but with no expensive elements, such as rhenium and ruthenium.

Besides, determination of the optimum heat treatment mode based on experimental works in production is a costly method.

Computing diffusive activity of doping elements may allow decrease development costs and optimize the regime for realizing the total potential of the alloy, embedded while it’s designing.

Analysis of nickel high-temperature alloys was performed while this work execution, and an optimal scheme of doping process to achieve maximum heat resistance was selected. With application of the computer aided method for high-temperature alloys optimization a new sparingly doped alloy for gas turbine blades, meant for operating at the temperatures up to 1050°C. The alloy is distinguished by high structural stability and economical use of doping elements. The new sparingly doped alloy relates to the first generation. With this, it complies with the third generation GS32 alloy by the level of heat resistance at 1000°C.

In the course of the works, development of nickel- based heat resistant alloys has been analyzed and an optimum alloying system has been selected to achieve the maximum heat resistance of the alloy. With the use of computerized optimization method of heat resistant alloys, a new lean alloy has been developed for gas turbine blades intended for operation at temperatures to 1050°C. The alloy exhibits high structural stability and efficient use of alloying elements. A new lean alloy is the first-generation alloy but its heat resistance at 1000°C corresponds to that of the third-generation alloy ZhS32.

A unique techniques for determining the diffusion coefficient of doping elements, and, based on the obtained data, for determining an optimal duration of the thermal treatment, were developed.

The microstructural studies of a new sparingly doped SLZhS32 alloy were conducted; a thermal treatment mode was tested with account for the diffusion processes kinetics; the samples were fabricated and strength tests were conducted.

The developed new sparingly doped alloy can be widely used for gas turbine blades manufacturing, ensuring the cost reduction without deterioration of the alloy operational properties.

Aydemir T. ., Golubeva N. D., Shershneva I. N., Kydralieva K. A., Dzhardimalieva G. I. Formation, structure and magnetic properties of nanocomposites obtained by Fe(III)Co(II) cocrystallized complexes thermal decomposition. Aerospace MAI Journal, 2019, vol. 26, no 2, pp. 219-228.

Considerable interest in d-elements nanoparticles well as the possibility of creating magnetic carrier with is stipulated by their magnetic properties specifics, as high information recording density on their basis.

Magnetic particles are widely used in biomedicine, and ferrous oxides (magnetite and maghemite), possessing high biocompatibility, play exceptionally significant role. Iron- and cobalt-containing particles are characterized by high values of coercive force and magnetic susceptibility. For example, for magnetite Fe3O4, the saturation magnetization (δs, Ms) is 92 eme⋅g-1, and for γ-Fe2O3-74 eme⋅g-1, the coercive force magnitude for anisotropic nanoparticles of the latter ranges from 200 to 400 Oe.

The structure and properties of metal-containing nanocomposites obtained while thermal transformations of Fe (III) Co (Il)-acrylate complexes were studied in this article.

It was shown that thermal transformations of the complexes under study included the stages of dehydration, solid phase polymerization and decarboxylation of the forming metal polymer. The solid phase product of the complexes thermal transformation are metal-containing nanoparticles, stabilized by carbonized polymer matrix. The crystalline nanostructured phases are Fe3O4, CoFe2O4 and CoO. The average crystallite size is 10 nm. Magnetic properties of the obtained nanocomposites also were studied. Hysteresis loops measured at temperatures below 200 K are open and displaced to a negative field. The coercive force and residual magnetization are 0.18 T and 15.5 mT, respectively.

An original approach consisting in combining nano-size metal particles synthesis with its stabilizing polymeric shell in situ was developed. The approach is based on metal containing monomers homo- and copolymerization in the solid phase with subsequent controlled thermolysis of the formed metall-polymers.

Accordingly, matrix-stabilized metal oxide nanoparticles were obtained by the method of polymer-mediated synthesis. In the nanocomposite obtained at 643 K and conversion of Δ m = 42%, the crystalline phase contains nanoparticles of ferromagnetic oxides Fe3O4 and CoFe2O4, and CoO antiferromagnetic nanoparticles. The nanocomposite microstructure includes polycrystalline agglomerates with sizes of 30 nm, consisting of individual nanocrystallites with an average size of 10 nm. The magnetic properties of the obtained products depend on the nature of the components, the temperature and the magnitude of the applied magnetic field. The coercive force and residual magnetization at room temperature are 0.18 T and 15.5 mT, respectively. The strong dependence of the magnetic characteristics on the phase composition, temperature, and magnetic field suggests that nanocomposites of this type are of interest for the sensor materials production for aerospace and biomedical applications.

Antipov V. V., Prokudin O. A., Lur'e S. A., Serebrennikova N. Y., Solyaev Y. O. Sial interlaminar strength estimation based on the results of the samples’ three-point bending tests. Aerospace MAI Journal, 2019, vol. 26, no 2, pp. 229-238.

Laminated aluminum-glass plastics (GLARE, SIAL) are promising structural materials for application while aircraft structural elements manufacturing These composite materials represent layered panels formed by thin layers of fiberglass and aluminum alloy. Compared with metals, SIALs possess increased specific strength, long-term strength and fire resistance. Studying the dependence of SIALs mechanical properties on the parameters of their reinforcement is an important task, which solution is necessary for the structures’ design and strength computation. One of the important characteristic, determining the SIALs structural properties, as well as the other composite materials, is interlaminar strength.

The samples testing on the three-point bending by the “short beam” technique is one of the simplest techniques for determining the interlayer strength of composite materials. This method is widely used in composite structures research and development, since it does not require the application of complex experimental equipment and strain gauges. At the same time, the interlayer strength is an important parameter from the designing viewpoint, as it is used in formulating the strength criteria of composite materials The interlaminar cracks occurrence may lead to a decrease in the bearing capacity of structural elements, and further to their destruction, for example, by the local buckling mechanism.

However, such a simple method as testing on three-point bending holds certain disadvantages associated primarily with the fact that during such tests a complicated stress state is realized in the samples, that is, not only the interlaminar shear stresses occur, but the also tensile / compressive stresses arise as well, leading to errors in determining the materials characteristics. Besides the above mentioned errors associated with non-uniform tensed-state of the samples, the complexity, occurring while samples testing on the interlaminar shear, consists in the fact that the interlaminar strength being determined while testing proves to be not a constant of the material, but it depends on the distance between the supports. This problem is known both for conventional composite materials and for metal-polymer composites. It is explained by a decrease in the tangential stresses actually acting in short samples (according to the standards the samples relative elongation shoul be of 5 to 10), compared to the classic beams models. These models assume the constant value of the shearing force, and, correspondingly, constant values of tangential forces (up to sign) along the sample length. Thus, application of the traditional relation for estimating transversal shear stresses acting in a beam, according to the formula 3 P / (4 b h), leads to the increase in the apparent interlayer strength of the material. Besides, the sample length impact on the results of the tests on the interlaminar strength is explained by:

1)    Stress concentration nearby the supports;

2)   Statistical dependence of strength on the sample size;

3)   The interlaminar cracks occurrence not on the neutral axis of the sample and

4)   Special dependence of interlayer strength on the parameters of fracture mechanics.

The article proposes a scheme for SIAL testing on the interlaminar shear strength by the short beam technique. These tests employ the samples with the large number of layers and unidirectional reinforcement scheme, which allows reduce the error of experiments while employing the standard equipment. The samples apparent interlaminar strength, depending on the distance between the supports, was determined by the results of the tests. Based on the calculations, the accordance of the obtained experimental data and theoretical estimates is demonstrated. The calculated SIAL interlayer strength value was of ~ 60 MPa, which corresponds to the typical interlayer strength of polymer composites. However, while testing the destruction was being realized at the contact boundary of metal and composite layers, which allows affirm that the found interlayer strength value is a characteristic of the metal / composite contact.

Shved Y. V. Determining technique for optimal rigging angle and aspect ratio of the soft wing with sling support. Aerospace MAI Journal, 2019, vol. 26, no 1, pp. 7-18.

While developing paragliders and gliding parachutes many issues on the optimal selection of the airfoil, its relative thickness and twist over the span, the law of the wing arc distribution and its shape in the sweep, length and slinging arise. Selection criteria for of some of these parameters may be transferred practically without changing the methods, rather explicitly elaborated for the historically earlier appeared aerial vehicles with balancing by the payload weight (hang-gliders). However, the paraglider, also related to the flying vehicles balanced by the load, has some specifics, since it employs momentless carrying shell.

The parameters estimates of the aerial vehicles with the soft wing and sling support with various working-out degree are presented in [5-19]. However, the issue of working-out the simple and vivid analytical technique for obtaining optimal characteristics of the above said aerial vehicles, which does not employ iteration approximating and general empirical assumptions, still remains open. The article is devoted to the study of some aspects of this technique.

author proposes to perform the calculation in the following sequence:

  1. It is assumed, that in the assigned flight mode, the wing has the required angle of attack. Aerodynamic coefficients of the airfoil Cxp and Cya for the specified mode are being elected.

  2. Based of the obtained coefficients, the glading angle is calculated according to the expression proposed in the article. Then, with account for the obtained gliding angle, the gliding speed is calculated using the following expression.

  3. After selecting several options of the wing profiles and aspect ratio the comparative calculation of the flight quality is performed. With too small values of the wing lift coefficient, the main contribution to the resistance is brought by the air-dropped cargo and slings. If the Cya is too large, the inductive resistance becomes prevalent. Consequently, for each wing aspect ratio, the system slings and cargo type it is possible to determine the optimum carrying capacity of the designed wing profile. Conversely, it is possible to determine the optimal aspect ratio with given the remaining design characteristics.

  4. After the final selection of the profile, by the center of pressure on the wing MAC (middle aerodynamic chord) is determined. Further, with account for the obtained coordinates of the center of pressure on the MAC, the coordinate of the wing suspension relative to the load center of gravity is determined by the proposed formula.

The article demonstrates also the independence of the of self-balancing wings angle of attack from the thrust magnitude. This conclusion is based on the fact, that for the angle of the slant of the slings relative to the center of the pressure of the MAC in the horizontal flight mode under thrust and in the gliding mode, identical equations were obtained.

In [1] the algorithm for static parameters calculation of the motor flight vehicle with a soft wing is presented. In the presented article it was expanded for the gliding descent mode.

Brutyan M. A., Potapchik A. V., Razdobarin A. M., Slitinskaya A. Y. Jet-type vortex generators impact on take-offand landing characteristics of a wing with slats. Aerospace MAI Journal, 2019, vol. 26, no 1, pp. 19-26.

To increase maximum lifting force coefficient of the aircraft wing with fixed geometry, it is reasonable to use the flow control concept. For this purpose, the new way of flow control about a wing with deflected slat, suggested by authors, is being studied experimentally and numerically. A number of slanting holes (along the flow and longwise a wingspan), through which the air jets are blown-out, is made to create vortex cores at the nose section of the upper surface of the wing’s main part, which opens while the slat root section moving-out. The pilot experimental studies of the new method of the wing with slat flow- around at the take-off and landing modes were performed on a model of a modern long-range aircraft with mechanized wing with moving-out slats and flaps.

The slats are made along the wingspan with a gap along the motor-nacelle pylon. The aircraft model testing while the landing state of the high-lift device with jet-type vortex generators and without them were performed with ADT T-106 TsAGI, equipped with aerodynamic scales. Slats and flaps were in landing state; with corresponding deviation angles of δsl = 24° and δfl = 36°. Weight measurements of aerodynamic characteristics were performed at the Mach number of the incident flow М = 0.15. It corresponds to the Reynolds number value of Re = 3.1⋅106 at the pressure pumping up to 5 atm in the working section of the tube. The angle of attack was being changed from 4 to 26°.

Numerical simulations of jet-type vortex generators impact on the wing flow-around pattern in a take-off and landing configuration were performed. Numerical calculations were performed to compare the experiment and the expanded range of the studied parameters. The well-known ANSYS CFX software based on the numerical solution of averaged Navier-Stokes equations for the compressible perfect gas with two-parameter SST turbulence model was used. The flow was considered turbulent starting from leading edge. The surface of the model was assumed adiabatic; the viscosity-temperature relation was determined by Sutherland’s law with the constant C = 110.4 K. The number of computational nodes used for the flow-around modelling with streams increased approximately up to 68 million.

The performed studies of passive technique for streams forming by the air blow-by from low the wing underside to its upside at the numbers of Re = 3.1⋅106 and М =0.15 revealed the possibility of the maximum lifting force coefficient increase.

Artamonov B. L., Shydakov V. I. Algorithm of transient flight modes performance by convertiplane. Aerospace MAI Journal, 2019, vol. 26, no 1, pp. 27-40.

The article considers the Project Zero convertiplane implemented according to the structure with two rotary screws positioned in the fixed wing. The screws are driven by electric motors powered by batteries, and controlled by a common and cyclic step. Electric transmission of the Project Zero convertiplane allows smooth change of the propeller rotations while transient flight mode performing with minimum required power.

The article analyzes control laws of screws, which allow performing transient flight modes from helicopter to aircraft without losing altitude at minimal engine power consumption. The described algorithm uses the results of experimental studies of the convertiplane body model in the t-1 MAI wind tunnel by th angle of attack at various rotation angles of the screws axes of rotation relative to the fuselage longitudinal datum line. This allowed reduce the problem to a system of transcendental equations of the convertiplane motion, which was solved numerically by successive approximations method. The aerodynamic characteristics of the propellers located in the ring fairings are being computed based on the disk vortex theory.

It is shown that while the convertiplane transition from hover mode to flight mode the screw control laws are of a rather complex character, and may be realized only by employing automation. The obtained convertiplane control laws at the transient flight mode are effective from the energetic viewpoint. The power consumption in the transient process endpoint is three times less than in the hover mode, which allows further convertiplane flight speed increase.

Gubernatorov K. N., Kiselev M. A., Moroshkin Y. V., Chekin A. Y. Studying elements reliability impact on the aircraft functional systems architecture. Aerospace MAI Journal, 2019, vol. 26, no 1, pp. 41-50.

The reliability of the more electric aircraft and its systems must not be less than the reliability of the conventional aircraft and systems to meet the required safety level. The level of the system reliability is specified in the part 25 or 23 of FAR. Power system of the more electric aircraft is a very important system due to the approach of ensuring the dragging and operating systems, such as control system and landing gear system. The weight-size parameters and the reliability of the more electric aircraft power system are opposite and depend of the power and energy system architecture.

This article demonstrates an approach to the architecture design of the more electric aircraft power system, that follows modern trends and ensures the required safety level and minimum volume and mass using state-of-the-art technologies, such as permanent-magnet generator and power electronics.

The current reliability level of power supply system elements (generators, rectifiers) cannot provide an extremely improbable event of the functional failure of the power generation system. Thus, the power supply system designers are forced install emergency (alternative) power sources such as batteries, a RAT, and auxiliary power unit, providing power to important systems to complete the flight and perform a safe landeing. These systems for example represent to an engine and an aircraft control system. The emergency (alternative) power supplies and the associated cables and switching system possess a considerable mass and volume. For example, the modern aircraft such as Boeing-787 and Airbus-350 have a very complicated power system to meet the required level of reliability. So these systems employ additional power converters, batteries, ram-air turbines and complicated distribution system. All of these have mass and occupy the aircraft volume.

Here is another example. The MC-21 emergency energy system weight is about 85% of the main energy system weight.

Hence, we can conclude that in order to meet the safety requirements, the power supply system designers should install almost one more power generation system onboard.

It is worth adding, that besides generation function the emergency power sources perform some other functions such main engines on-board starting, voltage ripples smoothing in the DC power systems with batteries and other. However, these functions are not taken into account in the presented article. The main attention is paid to the electric power supply system architecture developing, which meets the safety requirements, and contains minimum set of components to reduce weight-size parameters at large.

Shustrov T. L. Simulation as a substantiation of the trace contaminants removal system selection. Aerospace MAI Journal, 2019, vol. 26, no 1, pp. 51-63.

The article is dedicated to one of the most important problems while preparing any potential long-term or interplanetary space missions, namely the inefficiency of the life support subsystems employed at the habitable spacecraft. The article focuses mainly on the trace contaminants removal system (TCRS) being an important element of the space object life support system. It purifies the atmosphere of an object from any contaminants, and keeps it at the predetermined chemical balance.

The main hazards requiring permanent system regeneration and its keeping at the maximal possible technical level are as follows:

  • Atypical habitability conditions at the space object;

  • The crew impact (chemicals secretion as a metabolism result), as well as the spacecraft itself (chemicals emission as a result of degradation of coverings, used for internal plating, ) on the artificial living space.

The artificial atmosphere of any isolated airtight object is affected by its inhabitants, which could lead to the sensitive equipment failures, destructive emergencies, and deaths among the crewmembers. The presented article suggests employing simulation model as an attempt to improve the design and production of the future trace contaminants removal systems. The model allows computing the resulting amount of trace contaminants formed by any number of potential sources. The model structure provides the designer with maximum flexibility while the process regulation, which might help while creating individual configuration of the trace contaminants removal systems with account for the space mission scenario.

The article presents mathematical/technical description, structure, and examples of the simulations results. Most subprocesses are at the final stage of testing. The simulation results correspond to the telemetry data from the space station. In the future, after the final testing the authors plan to create the “artificial helper” for the model that will perform automatic selection of the trace contaminants removal system based on the results obtained after the simulation.

Gaponenko O. V., Gavrin D. S., Sviridova E. S. Structure analysis of the strategic plans of the space-rocket industry development by method of space functional and industrial technologies R&D classification. Aerospace MAI Journal, 2019, vol. 26, no 1, pp. 64-81.

The task of a subject for study classification arises while information analysis support of strategic programs for space-rocket industry technological development and managerial decision making on a sectorial level. In this case, it is an aggregate of scheduled measures, namely research and development work (R&D) on the cosmonautics and aerospace industry technological development.

The existing R&D classification in Federal target and Government programs (FTP) “Military-industrial complex development of the Russian Federation” does not fUlly reflect the structure of program activities, i.e. an aggregate R&D technological development R&D, and is applicable only to industrial technologies. In the Federal Space Program (FSP) the R&D is classified according to the target purpose of finished products. The R&D classification employing is not applicable in other FTP and vice versa. In the authors’ opinion, classification according to technological trends is the most efficient.

In domestic practice of analytical studies associated with the space activities technological R&D are subdivided into the intrinsic cosmonautics technologies (the space functional technologies), and industrial technologies for the space engineering development (the space industrial technologies).

There are also the third system-wide studies in the programs of cosmonautics and rocket building development, besides the functional and industrial technologies. These include complex system analytical research.

The forecasting of the space technology development without accounting for the capabilities of aerospace industry risks turning into vain dreams and fiction, and vice versa, the development of industrial production with no strategic targets in the form of promising space technologies may lead (and already leads) to creation of inefficient and economically unviable production structures.

The same technology, depending on the stage of the product life cycle of aerospace technology, can be attributed both to the target technology and to the of industrial production technology.

The unified R&D classification system of aerospace functional and aerospace manufacturing technologies and system-wide research effort is advisable. There is a necessity of a unified classifier for the cosmonautics development strategic programs (FSP, state programs “Development of the MIC”, strategic programs and plans of other governments) in parts of R&D sections.

The article proposes a unified classifier of space- rocket and manufacturing technologies. It is based on the classification features of technologies used by NASA in the technological road maps of 2015.

The classifier was realized by the authors in the form of an object-relational database on PostgreSQL. The database is switched as an external data source to Excel, and further the analytical capabilities of the free Excel table mechanisms are used.

A comparative analysis of R&D technologies performed by NASA, the European Space Agency and State Space Corporation “Roscosmos” within the framework of long-term strategic programs of space activity was performed using the developed classifier. The classifier allows also compare the same technological trend in different programs.

Besides the number of works the developed classifier allows analyzing their financing, starting/ ending dates and starting/ending level of technological readiness by technological trends.

The classifier allows reveal the technological development trends, to which most attention is paid in the states participants of the space activities, and vice versa which are related to unessential, and their studies are not financed by strategic programs. The structural specifics of each of the considered programs of technological development can be analyzed.

Practical implementation of techniques, associated with program events classification forming and scientific-methodological support of the strategic programs of national space-rocket industry development (including application of the classifier suggested by the authors) with subsequent analysis of the obtained classes will contribute to the managerial decisions effectiveness in Russian space-rocket industry, and eventually in rational implementation of the State budgetary funds allotted for this purpose.

Nikitin S. O., Makeev P. V. A project of the “Synchropter” type high-speed helicopter with pushing air propeller. Aerospace MAI Journal, 2019, vol. 26, no 1, pp. 82-95.

Due to the helicopters ability to perform vertical take-off and landing, as well as effective operation while hover mode, they became indispensable practically in all regions of the world. With that, the requirements for the helicopters flight performance enhancement become ever more acute, primarily concerning the increase in speed and range.

Currently, a number of rotary-winged aircraft structures of vertical take-off and landing, realizing increase in speed and flight range, are under development and in some cases at the stage of testing and batch production in leading world countries. There is a number of concepts and technical solutions, mainly in the field of aerodynamics, allowing increase a helicopter cruising speed. In this regard, the exploratory research and these projects implementation development are highly relevant.

The presented work is devoted to creation of a project of a perspective passenger high-speed aircraft with vertical take-off and landing based on a helicopter with intermeshing rotors and a pushing air propeller.

The project employs a set of the following technical solutions:

-    The blades rotational speed reduction (from 220 to 180 mps) as the flight speed increase; special arrow­shaped tips setting on the blades to reduce to zero the probability of a wave crisis on the advancing blades with flight speed increasing;

-    Balancing the unbalanced lateral tilting moments on the two rotors of a “synchropter” scheme, rotating in opposite directions;

-     Application of rotors with elastic torsion sleeves;

-    Application of a system of the blades individual control to prevent the flow disruption on the retreating blades;

-    The aircraft fuselage layout with account for the specifics of the scheme with low frontal resistance at near-zero angles of attack;

-    Application of a propulsion propeller with maximum efficiency in operating conditions.

The capabilities of modern computer-aided design technologies were demonstrated while the project developing. The main emphasis is made on the aircraft dynamic designing with implementation of modern tendencies of the high-speed helicopters development. The main limitations and possible ways for the helicopter speed increase implementation were considered. The article presents the computational results of aerodynamic characteristics with account for the decisions made.

The developed project has the following characteristics: the take-off weight of 6500 kg, payload mass of 1000 kg, maximum speed of 420 km / h, static ceiling of 4700 m, dynamic ceiling of 5600 m, and flight range of 1228 km.

The obtained results indicate the achievement of indicators close to the modern world level, demonstrated on similar developed helicopters.

The developed project has prospects for further flight performance improvement by improving the‘ aerodynamic characteristics of the fuselage, propellers, as well as exploiting more fully the capabilities of the individual blade control system.

Erkov A. P. Buckling of stepped beams. Aerospace MAI Journal, 2019, vol. 26, no 1, pp. 96-110.

The article discusses the problems of stability of two types of beams of variable stiffness: with a stepped change in cross section with two zones and with a step change in section with three zones. Simply supported boundary conditions at two ends are considered, as well as with embedding at one end and with a free second end. Beams of isotropic material and of the laminated composites are discussed.

To study the stability of beams of variable stiffness, the Ritz method was used. Beams with the ratio of the maximum and minimum flexural rigidity in the zones does not exceed 8 are considered, since in practice the ratio greater than 8, as a rule, is not applied. Analytical expressions for determining the critical force are obtained. The calculation results and their verification are given.

The results of analytical calculations were compared with the results obtained by the finite element method (MSC.Nastran / MSC.Patran). Based on a comparative analysis, graphs of the error of analytical solutions (relative to the solution obtained by the finite element method) were constructed. To minimize the error of analytical equations, a correction factor was introduced.

The study showed that the equations applicable for calculating the critical force of isotropic beams are also applicable to composite beams. Correction factors obtained for isotropic beams are also applicable to composite beams.

In addition to assessing the accuracy of analytical equations for the critical force, the influence of local effects in the area of the junction of zones with different flexural rigidity is investigated. In practice, the Bernoulli hypothesis does not work in the junction area of the zones, which has some influence on the magnitude of the critical force.

Results of investigation:

- Analytical equations were obtained for determining the critical force for two types of beams of variable stiffness with two types of boundary conditions;

- The accuracy of analytical equations was investigated. A correction factor was introduced, which allows to obtain a more accurate result for the critical force;

- The technique can be applied to other types of beams of variable stiffness and other boundary conditions not considered in this paper;

- The resulting analytical expressions are easy to automate. For this suit, for example, Microsoft Excel can be used.

Baklanov A. V. The impact of the of fuel supplying method to the combustion chamber on carbon oxides formation in combustion products of the gas turbine engine. Aerospace MAI Journal, 2019, vol. 26, no 1, pp. 111-125.

The fuel burning in the combustion chamber of a gas turbine engine (GTE) is attended by toxic substances formation. Carbon oxides, having deleterious effect on human and environment, are of particular danger. In this regard, the article solves the actual problem of determining the optimal method of gaseous fuel supplying to the GTE combustion chamber to ensure low emission of carbon oxide.

The article considers the burner with two types of injectors, differing by the gas spray method. The first injector is a centrifugal gas injector (CGI), and the second one is a jet injector (JI).

A technique of target feeding of a jet, formed by the injector in the burner unit was developed.

The fire tests of nozzles were performed. While the tests performing, it was revealed that during the burner operation with the fuel feeding by the CGI, the flame front was being stabilized along the walls of the burner nozzle extension with visible hollow red colored core. Behind the main flame, the reddish “tail” which length corresponded to the length of the main flame was observed. This indicates that the fuel has no time to burn out in the primary zone, and flame front is stretching out.

In this regard, the quality determination of air-fuel mixture preparation in the swirled jet at the outlet of burners with two types of nozzles was performed. It was established, that the nozzle with the jet-like fuel atomization ensured the best mixing quality. The engine throttle characteristics were determined, and carbon oxides concentration in the combustion products measuring was performed by the results of the experiments. The results demonstrated that with the power increase the carbon oxide concentration level in the combustion products decreases. The 25% from the initial variant decrease in concentration was observed herewith for the combustion chamber with JI, which corresponds to the 28775-90 State Standard.

Khramin R. V., Slobodskoi D. A., Lebedev M. V., Sobul’ A. V. The test bench development improvement of the gas turbine engine due to the application of the new method for axial force determining impact on the radial-thrust bearing. Aerospace MAI Journal, 2019, vol. 26, no 1, pp. 126-133.

A radial-thrust bearing of rotor supports is one of the most critical elements of aviation gas-turbine engine, as its failure leads to the engine destruction. To ensure the required reliability of such bearings, experimental studies allowing increase the accuracy of design models employed for the bearing life determination under engine operating conditions were performed. One of the main factors affecting the endurance of radial-thrust bearings is the axial load.

The current quantitative method of axial load determination during engine tests employs the technological supports with dynamometric rings. Qualitative methods of axial load determination based on vibration sensor readings do not allow correct determining of the axial load.

This article presents the method used to measure the axial force applied to radial-thrust bearing. The method is based on dynamic strain gauging of bearing rings. Strain gauges are installed into special slots in the bearing rings. The slot width should be maximum possible but not exceeding the distance between the adjacent rolling elements. The slot depth should comply with the requirements for admissible deformation of raceways and sensitivity of the strain gauges.

The strain gauges readings are taken in the values of relative strain (mm/mm). For ease of use, these values are converted into stresses values (kgf/mm2) by multiplying them by the elasticity modulus of the bearing ring material.

To determine the dependency of the strain gauge readings on the axial load, calibration on a special installation is performed. During calibration, the strain gauges measure the variable stresses in the slot. The amplitude of variable stresses with flicker frequency of the rolling elements is proportional to the axial load, and is a key parameter. To determine it, the signal from the strain gauge, at any given moment, is represented as a Fourier series, and spectrum of the signal amplitude-frequency response is formed. This spectrum is being used to determine the amplitude on flicker frequency of the rolling element. Based on the test results, the calibration factor is determined which characterizes the dependency of axial load on the amplitude of the strain gauge reading signal. Then, by the measured dynamic stresses recalculation, the axial load applied to the bearing is determined.

The accuracy of axial load measurement by dynamic strain gauging of bearing rings does not exceed ±1% of the reference load. The above­described method has been applied during engine tests together with the current method with temporary supports and dynamometric rings.

Based on the test results, the accuracy of axial load determination has been increased and the number of the required engine tests has been reduced.

Gogaev G. P., Nemtsev D. V. The study of flight conditions impact on high-pressure turbine disk damaging of the highly maneuverable aircraft. Aerospace MAI Journal, 2019, vol. 26, no 1, pp. 134-142.

The increase in the GTE life cycle cost brings to the forefront the problem of the full safe use of the aviation engines lifetime, which can be achieved by the transition to operation on a technical condition. This transition is possible with the sufficient product testability ensuring obtaining the objective information required for the reliable technical condition estimating.

The crucial problem herewith consists in methods and algorithms developing for estimation the lifetime depletion, accounting for loading specifics of each engine.

Excessive conservatism is inherent to the currently employed methods for lifetime cycle depreciation control due to the lack of actual operation conditions record keeping. Premature engines exclusion fr om operation occurs thereby, which is unfavorable and has an adverse effect on supporting the required combat readiness level of the aircraft fleet.

Thus, the trend of control techniques improvement, analysis of loading and GTE lifetime deprecation control, fully accounting for the operation specifics of each engine is relevant enough.

The purpose of this work consists in studying the impact of flight conditions on the high-pressure turbine (HPT) disc damaging of highly maneuverable aircrafts.

The main contribution to the parts damage accumulation of the highly maneuverable aircraft engine is made by the damages, caused by intermittent operation modes (the low-cyclic fatigue mechanism), and operation at the maximum set modes (the mechanism of long-term strength depletion).

As the service experience of the 4th generation engines being a part of highly maneuverable aircraft of the task aircraft fleet shows, the contribution of a static component to the overall damage of the basic engine parts is significantly less than the cyclic one. Thus, the estimation of the residual engine life is made, as a rule, based only on accounting for the cyclic damages of its basic parts.

The main idea of the 4th generation engine life deprecation accounting for consists in comparing the actual value of the technical condition parameter (the accumulated damage) of the engine basic parts during the operation with its maximum permissible value, accumulated while the endurance tests, with subsequent determination of the residual resource of the engine basic parts according to this comparison.

Currently the number of cycles before the failure (Npi) and the single damage (Пi) for each cycle type are is detemined at the extreme loads (engine power rating, speed, and flight altitude) for the given GTE operation range.

However, the performed analysis of the highly maneuverable aircraft operation belonged to the task aircraft fleet, revealed that about 80% of the operation was performed at subsonic speeds and heights up to 10 km (without participation in combat operations), at which the basic part load was much lower than its maximum value. Thus, the existing methodology application leads to the excessive conservatism of the accumulated damage calculation.

To assess the effect of flight conditions on the single damage of the main parts, a complex of calculations for HPT disk of the 4th generation engine were performed. The obtained results demonstrated that the single damage of all cycle types of the HPT disk significantly depends on the flight conditions. Thus, the single damage of the loading cycles in the zone, wh ere 80% of operation time is performed in default of combat operations participation, is on average 25% below the values at the maximum loads for all cycle types.

In the context of the HPT disk of the 4th generation engine, the article shows that the existing technique for the lifetime deprecation monitoring by low-cycle fatigue of the 4th generation GTE basic parts includes assumptions leading to the accuracy reduction of determining the accumulated damage and the residual life of the engine and its main parts. This, in turn, leads to an early removal of a serviceable engine, and the life cycle cost increasing.

To avoid the excessive conservatism of the currently used technique, it is necessary to accumulate the cyclic damage of the engine basic parts with account for real flight conditions.

Tkachenko A. Y., Filinov E. P. Gas turbine unit efficiency upgrading for gas-turbine locomotive of a new generation. Aerospace MAI Journal, 2019, vol. 26, no 1, pp. 143-151.

Up to now, at least half of the railways are not electrified. Thus, it is necessary to employ heat engines to set a locomotive into motion. Employing a gas turbine unit (GTU) is one of the possible options. The GTU power is transferred to the generator, and electric motors set the locomotive into motion. It is worth mentioning that in the future aircraft engines of the civil aviation with worked-out lifetime, and updated for the railway application may be installed on a gas-turbine locomotive. Such an approach would significantly reduce the transportation cost value and gas-turbine locomotives implementation to the national economy.

This work was performed in several stages:

– Mathematical models verification used while performing design calculations and GTD operational characteristics computing to increase their identity to the mathematical models employed by PJSC Kuznetsov;

– Studying the number of stages of a low-pressure compressor (LP) effect on the of a gas turbine unit performance employed as a part of the gas-turbine locomotive;

– Proposals development on improving the units’ joint operation to reduce the air consumption through the gas turbine unit.

One of the ways to improve the operation efficiency of gas turbines for application as a part of the gas turbine locomotive consists in the air flow reduction through the unit, which would allow reduce the total pressure losses in the suction tract due to more rational operation conditions of the air filters. The possibility of air consumption reduction through the engine in condition of preserving the effective power of the gas turbine unit by eliminating one and two stages of the low pressure compressor will be discussed further.

The following main scientific results were obtained as a result of the study:

  1. Mathematical models verification used while performing design calculations and GTD operational characteristics computing to increase their identity to the mathematical models employed by PJSC Kuznetsov. Comparison of the results of GTU climatic characteristics computing, based on the initial gas generator, with data obtained at the PJSC Kuznetsov allows talking about the identity of mathematical models of thermo-gas-dynamic computation, performed by the PJSC Kuznetsov, and ACTPA mathematical models;

  2. A study of the low-pressure compressor number of stages impact on the operational characteristics of the GTU employed as a part of the gas-turbine locomotive. Based the obtained results, a conclusion can be made on the inexpediency of changing the number of stages of the low-pressure compressor without refinements (changing the joint operation conditions of the GTU units by throughput efficiency correction of nozzles assembly);

  3. Proposals on improving the joint operation conditions of the units to the effect of air consumption reduction through the GTU, and the most rational options of nozzles assembly of the low-pressure turbine and a free turbine were elaborated.

Dukhopel'nikov D. V., Vorob'ev E. V. Technique justification for erosion profile determining of the accelerating electrode of ions gas-discharge source. Aerospace MAI Journal, 2019, vol. 26, no 1, pp. 152-157.

Determining erosion rate of electric rocket engines elements and other gas-discharge devices is the most important stage of their design and testing. The simplest method to determine the surface erosion rate under the ion bombarding may be employing of optically contrasted multilayer coatings pre-applied to the surface under study. The pattern of alternating optically contrasted bands occurs while sputtering these coatings by the non-uniform ion beams. The boundaries between these bands are the lines of equal erosion depth.

The surface slope angle in the erosion zone while a massive material sputtering by a non-uniform ion beam is determined by the equation


where Ma is the atomic mass of the material, ρ is its density, Seff is effective sputtering rate, j is the ion current density, t is the ion beam exposure time, and q is the ion charge.

While selecting a multilayer coating structure computation of separate layers thickness δi is performed on the assumption of the required band width and the surface slope angle in the erosion zone

The layers thickness herewith should be selected so that the bands widths on the image repeatedly exceeded the registration resolution of the equipment employed for the sputtered patterns photo-registration. Thus, to obtain accurate results using the represented technique, the correct surface slopes angles a; determining is required.

At the same time, while sputtering multilayer coatings, different points of the layer, lying in depth, begin sputtering at different time moments, in contrast the massive material. Thus, the necessity occurred to confirm the correctness of application of the expressions, obtained for the massive material, to the layers thicknesses computing of the multilayer coating.

This article is dedicated to the analytical proof of the expressions usage appropriateness to calculate the erosion slope angle and the layers thickness in the depth of the multilayer coating. It shows that these expressions can be used for any layer of any material located in the depth of the multilayer coating of arbitrary structure.

Volkov S. S. Psychophysiological condition assessment of an operator of the ground complex’s ergative system. Aerospace MAI Journal, 2019, vol. 26, no 1, pp. 158-165.

The article considers an automated system for psychophysiological condition (PPhC) assessment of a flight crew, spacemen, test pilots and other representatives of the airspace industry. The PPhC operation is based on the gas discharge visualization (GDV) method.

The purpose of the work consists in demonstrating the effectiveness and necessity of the psychophysiological state monitoring of the ergative system operators. The ergative system operators are the flight crew of both military and civil aviation; astronauts; test pilots; robotic systems specialists.

This work novelty consists in the GDV method application in a new area. The interest to this method application is caused by the fact that operators are working in special conditions of professional activities. In this regard, they suffer fatigue, overtiredness, undersleeping, performance decrement, stress etc. The PPhC neglecting may lead to tragic aftermath. Thus, the authors suggest developing prospective automated system for operators’ psychophysiological condition estimation, which would allow monitor operators’ readiness to perform their service duties while their professional activities.

During the survey, the snapshots of ten fingers are made with the filter, and another ten without it. The obtained images are being separated into sectors. Further, the mathematical apparatus described in the article is applied to them. The stressed background and normalized glow area, necessary for the psychophysiological state determining, are being computed. After obtaining the information on the operator’s PPhC the official takes a decision on the given person’s readiness to perform his service duties.

The results of the studies allowed developing an algorithm for the software operation of the operatots PPhC estimation system. Neural network technologies are supposed to be the basis of this work. They will improve and expedite the information processing process.

The automated PPhS estimation system, described in the article, introduction into the aerospace industry, will allow monitoring the health of the flight crew, cosmonauts, test pilots and robotic complexes operators, as well as reduce the risk of injury and death while equipment operation.

Anan’ev A. V., Filatov S. V., Petrenko S. P., Rybalko A. G. Experimental approbation of free-falling uncontrolled containers application, employing short-range unmanned aerial vehicles. Aerospace MAI Journal, 2019, vol. 26, no 1, pp. 166-173.

Suppression of enemy’s air defense systems by employing small size striking unmanned aerial vehicles (UAV) to reduce the risk of the piloted aircraft fire damaging is a topical task. The world practice of the small-sized UAV application for striking with free- falling uncontrolled containers (FFUC) is a premise for their application. The majority of scientific publications, describing the UAV striking application, are based, in general, on mass media information, combat effectiveness estimation simplified to its lowest limit, expert esteems of the UAV combat effectiveness without their transformation into qualitative estimations. Highly in-depth academic studies are known also. However, they are based on the probabilistic apparatus, which application is impossible due to the lack of probabilistic laws and random values parameters required for the calculations.

Thus, by this time, the full-fledged computational ballistic algorithms for the small size striking UAVs cannot be realized in practice. With account for the above said, practical approbation of the UAV striking application as the most crucial stage of the aviation complexes lifetime is of first and foremost interest.

Thus, the first and most valid method for the UAV striking capabilities estimation is performing experimental ballistic tests. Their results can be employed for such UAVs efficiency estimation in striking variant, and forming tabulated data on FFUC hitting accuracy, parameters spread, according to which firing tables will be composed.

To reach the purpose set in this work, the following problems were defined and solved:

- The target environment was created for refinement of the FFUC practical application employing UAV;

- Estimation of FFUC with UAV application in striking embodiment on the land objects with application of the simplified deflection measuring technique and estimates of arguments of the FFUC dispersion was performed;

- Statistical data on experimental UAV application in striking mode while hitting ground objects and the enemy’s manpower, for subsequent determination s of FFUC dispersion were collected and processed.

A target, simulating the command center of the medium range surface-to-air missile system battery was employed while testing.

Systematized data on the FFUC dropping were obtained according to the results of the work. They can be utilized for mathematical support developing for the command post of the short range UAVs in striking configuration while developing aiming algorithms.

The obtained results confirm the hitting effectiveness of the FFUC equipped with ammunition of “tactical grenades” type of the enemy manpower and vulnerable (light armored) ground objects.

By results of the obtained statistical data and preliminary calculations, the accuracy of the FFUC application was from 8 m to 10 m.

Legaev V. P., Generalov L. K., Galkovskii O. A. An analytical review of existing hypotheses about the physics of friction. Aerospace MAI Journal, 2019, vol. 26, no 1, pp. 174-181.

Assigned the task to determine the laws of change in the coefficient of friction and determine the factors affecting it as part of the research work. The purpose of this work is to improve the performance parameters of precision machines.

Physics of the external friction process has the next form. When the contacting solids are shifted, the external friction force increases due to deformation of these solids, this phenomenon is called preliminary displacement. Static friction force Fs is the force of friction, corresponds the highest value of the preliminary displacement. One of the contacting solids moves irreversibly (slides) across the surface of another solid after a static friction force has been achieved, in

Relation of external friction force F to the movement x this case the external force is equal to the kinetic friction force Fk [2].

Friction interaction occurs in certain parts of the nominal contact only. Friction interaction is the third solid. The complete complex of frictional links forms a frictional interaction, which is discrete [1].

The preliminary displacement is caused by the redistribution of the contact irregularities in the support surface [1].

The total static friction force is the boundary point, under which preliminary displacement friction passes to kinetic friction force.

Kinetic friction is the friction of two solids that are in motion relative to each other [4].

Kinetic friction has a dual molecular-mechanical nature. It is caused by volumetric deformation of the material and overcoming intermolecular bonds

where Ffm - is the molecular component of the friction force; Ffd - is the mechanical (deformation) component of the friction force [2].

If the adhesion bond is less strong than the underlying layers, then there is a positive gradient of mechanical properties at depth:

Under normal friction process, the positive gradient rule is always come true.

The contact is always discrete and the area external friction depends on the applied load at external friction. Contact surface is continuous and independent on the applied load at internal friction.

The coefficient of friction depends on three factors almost equally: combination of materials; construction of friction pair; operating mode [1].

To execute the rule of positive gradient must be present lubrication film in the friction contact, or oxide film, soft components film [1].

The growth of the film slows down with increasing its thickness [1].

The growth of the film reduces the coefficient of friction to a known limit. Very thick films increase the coefficient of friction [1].

The relative sliding of two solids produces heat in a thin surface layer. The temperature rising can lead to local softening and melting of the material. The temperature field leads to a change in the mechanical properties of the material in a thin surface layer. The intensity of the heat flow depends on the friction work and the size of the area on which it is generated [1].

Important constructive characteristics of the friction units is the coefficient of mutual overlap, proposed by A.V. Chichinadze,

where Aa1 - the nominal friction area of the first element; Aa2 - the nominal friction area of the second element; Aa2 ≤Aa1 .

Wear products have a great impact on the strength and coefficient of friction [2].

Friction and wear characteristics and mechanical properties of friction pairs materials are in various nonlinear functional dependencies. At the same time, these dependencies can significantly change from the friction mode and from the thermal mode of friction pairs.

The construction of the friction unit significantly affects the force and coefficient of friction. In this regard, the nominal Aa, contour Ac and actual areas of friction Ar, the coefficient of mutual overlap Kov, the shape and size of the contact elements, their stiffness and elasticity is among the main parameters determining friction.

More rigid elements of surfaces intrusion into softer counterbody due to waviness, roughness, heterogeneity of mechanical properties and duality of molecular-mechanical nature of friction.

Accordingly, the speed ν of the indenter determines the friction force. At the same time, an increase in the load on the separately selected indenter leads to an increase in the friction force. However, the support reaction force N affects the area of the actual contact Ar in the actual operating conditions of the friction pair. The actual contact area depends on the load. Increasing the area of actual contact reduces specific pressure. Thus, the dependence of the friction force on the relative velocity of the friction pair and the load is not linear and differs for different materials.


  1. In the research of the friction of polymeric and metallic materials should be used adhesion- deformation theory of friction, which includes the definition of the molecular and mechanical components of the friction forces.

  2. The thermal and mechanical properties of materials should be determined by the known friction force of the friction pair.

  3. A positive gradient rule should be observed and lubrication films, oxide films or films of a soft component in the friction contact should be provided.

  4. It is necessary to determine the area of the contacting surfaces at the given micronutrients and friction forces.

  5. It is necessary to take into account the shape and size of the friction unit and the coefficient of mutual overlap.

Mudrov A. P., Faizov M. R. A spherical simulator motion study. Aerospace MAI Journal, 2019, vol. 26, no 1, pp. 182-191.

The article presents a spherical mechanism allowing perform spatial movements along a sphere. A 3D model of the mechanism was developed with the SolidWorks software. The model allows synthesize and examine the mechanism structure. Computing of the angular displacement, speed and acceleration of the connecting rod from the center point of the link and the slide of the mechanism itself was performed. The center point calculation was performed with account for the small and large thickness of the mechanism links Calculations were made for all angles between the links, which were employed for calculation of the spherical mechanism with two degrees of freedom. Based of the obtained mathematical model, computing of the moment of inertia from a given crank motion was performed. The motion parameters along the coordinate axes were determined, which would allow application of the direction cosines formulas. Additional angles calculation used when creating a mathematical model for the moment of inertia were obtained from the spatial sphere around the mechanism. The instant rotation moment of the mechanism was obtained. Using to the obtained data, a certain movement of the mechanism and the time interval of its movement were set, which are reflected by the obtained plots. These plots were obtained for comparing the two methods. The obtained plots reflect the movement of the connecting rod itself, and the slide mechanism. In addition, using Maple, the computation of motion with the moment of inertia of the mechanism itself, with a specified various masses, but with certain geometrical parameters of the mechanism links, was verified.

Rebrov S. G., Yanchur S. V., Drondin A. V., Zernov O. D. Developing the concept of solar energy units robotic assembly in orbit. Aerospace MAI Journal, 2019, vol. 26, no 1, pp. 201-211.

The recent foreign experience in the spacecraft development, including lengthy sup-porting structures developing, and giant space telescopes and solar energy systems con-struction indicates that further development in this field of engineering is impossible without transferring the structures manufacturing directly into the space.

As applied to power systems, this is motivated by the low packing efficiency (measured in percent of occupied volume) of the power system elements inside the launch vehicle, and, correspondingly, the small value of the “Stowed Volumetric Power”, or “Stowed Volume Power Density”, or “Stowed Volume Efficiency” (measured in kW/m3) parameter of the head fairing. This practically excludes the other ways of increasing the power systems total power, other than forming by independent delivery of the power systems parts by way of several launches. The latter leads to a multiple increase of the projects costs, which is not often acceptable.

The article proposes a solution to the describedabove said problem in the form of a concept of a robotized assembly of solar power arrays in space, which is based on the application of the Solar Arrays Assembly Machine (SAAM).

SAAM is a robot with which a solar cell of a large area is being assembled by attaching the mounting panels to each other. The mounting panel can be a honeycomb of high stiffness, allowing the SAAM to move along its plane. When moving, the SAAM “clings” to the reference holes made on the mounting plates in advance. SAAM has four telescopic supporting rods for moving around the mounting plates and two mounting arms for fixing the panels.

The concept demonstrates the scheme of the SAAM application. determines The optimal route for the SAAM movement and the order of the solar array assembly are determined. The scope of its possible application has been determined: for assembling a wing of a solar array with an area of less than 64 m2, the target (competitive) mass of the SAAM is of linear dependence on the area of the solar array. When assembling solar arrays with an area of more than 64 m2, traditional deployment systems cannot be employed. So the SAAM does not have competitive alternatives implemented.

The basic SAAM size are determined. A layout was made allowing develop the basic technological operations and algorithms of moving and assembling. The system weight and size parameters depend on the materials used, electromechanical assemblies, SAAM batteries, and will be refined further further work. The time of the solar array forming depends on the speed of SAAM electromechanical units and manipulators operation. But this is not a limiting factor, since the modular structure of the system should allow the SAAM to recharge from the assembled segments at several stages of the assembly.

Kovalev A. A., Konovalov D. P. Workpiece thermal deformations simulaiton occurring while holes drilling process. Aerospace MAI Journal, 2019, vol. 26, no 1, pp. 201-211.

The article tackles the issue of determining the error caused by the workpiece thermal deformations occurring in the holes drilling process in a part, being the main part of the unbraked wheel of the aircraft landing gear and is called the “Drum”.

The article describes the mechanism of these errors occurrence. A method for the treatment process simulation was developed, and proposed an algorithm for estimating the error in the workpiece size occurred due the thermal deformation while drilling.

The article consists of three main parts, namely introduction, body part and conclusions.

The introduction considers the mechanism of errors occurrence due to thermal deformations of the workpiece, which in turn presents one of the total machining error components. It presents the cases when this error may significantly affect the total machining error. Thus, it is relevant that this error component is estimated.

The basic part presents a method for computing the temperature in the cutting zone for further machining process simulation. It describes the object of simulation, i.e. the operation of drilling a through hole of 13.5 mm diameter with the tolerance range of 120 μm in a workpiece from the ML12 magnesium alloy with the cutting modes recommended by the cutting tools manufacturer, namely, the cutting speed of 264 m/min and feeding of 0.35mm/rev. An algorithm for the size error estimation is presented as a block diagram. The step-by-step description of the hole drilling simulation process is presented on the example of this operation. As a result, the temperature distribution, equivalent von Mises stresses, and displacements caused by thermal deformations over the part volume were obtained. Based on the diagram of displacements, caused by thermal deformations, the error was 191 μm at the specified cutting modes and machining conditions, which appeared greater than the tolerance range by the size of the hole.

The conclusions note that the cutting parameters recommended by the cutting tool manufacturer do not always provide the required machining accuracy. It was concluded that the required accuracy was not achieved for a specific hole drilling operation. The ways leading to the error reduction due to changes in cutting parameters, as well application of the other types of cutting fluid are presented in the conclusion.

Verchenko A. V., Kurskaya I. A., Chigrinets E. G., Maksimov D. V., Geiko Y. S. Water-jet cutting process optimization of work pieces from aircraft materials. Aerospace MAI Journal, 2019, vol. 26, no 1, pp. 212-229.

Single and small batch production is predominant in parts production for aerospace industry. Stamped blanks and castings manufacturing with small batches is not cost-effective due to the high cost of tooling. Thus, forgings or billet plates made of thick plates that are close to the part’s shape are used in increasing frequency as work pieces.

One of the most up-to-date and promising method of cutting and obtaining finished parts is the method of water-jet cutting. It ensures wide ranges of processed material thickness, the ability to cut almost any material, high performance, obtaining high quality cutting surface, the ability to process complex geometry. All this makes this method of processing the most popular in conditions of modern aircraft building, shipbuilding, etc. The absence of thermal impacts on the material, low cutting force, the erosional nature of the destruction do not contribute to the development of internal stresses in the cut zone.

The process of water-jet cutting is complex, poorly understood, which result is affected by many technological parameters such as cutting pressure, nozzle feed, grain, hardness, abrasive consumption, distance from nozzle to the surface being processed, physical and mechanical characteristics of the material being processed. The design complexity of the cutting technological process consists in selection of optimal cutting conditions, which will ensure the specified quality of the part surface layer at the minimum cost. The production technologist faces the difficulty of determining not only the cutting surface hardness, but the size of the smooth and wavy cut zone as well.

goal of the work consists in improving the efficiency of the waterjet cutting process by optimizing processing modes based on the development of an adequate theoretical model for the formation of surface roughness at different depths of the cut section.

To achieve this goal the following tasks were solved:

  1. Theoretical and experimental studies of the cut surface roughness profile formation depending on the processing parameters;

  2. Theoretical studies of the wavy cut zone formation depending on the technological parameters of the process;

  3. Development of methods for predicting the quality of the cut surface;

  4. Development of methods for optimizing the process of water-jet cutting.

The paper presents the results of theoretical and experimental studies of the surface roughness profile formation while water jet cutting of various materials, such as 30HGSA steel, hardened 30HGSA steel, D16T aluminum alloy, fiberglass-titanium composite material. A theoretical model for the roughness profile formation of the cut surface was obtained, which shows the dependence of roughness on the main technological parameters of the process (nozzle feed, particle radius, mixture pressure, etc.) depending on the depth measurement of the cut surface roughness. It reflects thereby the distribution of the ratio of the smooth and wavy cut zone. Statistical processing of the studies results was performed using MathCad. The experimental studies result was the obtained dependencies of the number of the particles’ useful encounters with a material on the magnitude of the nozzle feed, abrasive consumption, and section depth. One and two-factor regression equations describe the effect of abrasive consumption, nozzle feed, thickness of the material being processed, section depth on the cut surface roughness.

A two-factor regression model for the formation of the roughness profile of the cut from the nozzle feed rate and the roughness measurement depth while polymeric composite materials (PCM) processing of the fiberglass-titanium type was obtained. The material layering and shagging while cutting were not detected, the cut quality was high. To assess the water impact while cutting fiberglass-based PCM, an analysis was performed using differential scanning calorimetry, which resulted in the conclusion that the waterjet cutting technology can be used for PCM processing.

Based on the theoretical and experimental studies results, a for designing and optimizing the technological processes of water-jet cutting technique has been developed, with account for the specified cut surface roughness ensuring and obtaining the minimum cutting costs.

The optimization of the technological process of water-jet cutting of the “Bracket” part of the Mi-28 helicopter was performed, which resulted in a 2.5 times reduction in labor intensity, a cost cut of 843.51 rubles, which allowed the company to save 1286 rubles while each part production. The technique for the water-jet cutting technological process optimization application was undergone industrial tests at the Rostverol plant.

According to the technical requirements for rotor blade manufacturing, as well as the results obtained by the authors, the possibility of hydro-abrasive cutting application for removing the technological allowance in the basis part of Mi-28 helicopter main rotor spar as an alternative to the rough milling was demonstrated. Application of cutting feed within the 160-240 mm / min range min reduces the labor intensity by 80% with the required quality indicators.

At present, measures for the suggested technology introduction into batch production are under development at the PJSC Rosvertol Blade plant.

Zhemerdeev O. V., Kondratenko A. N. Methods for determining technical potential state of the enterprises based on a modified model of factors of production. Aerospace MAI Journal, 2019, vol. 26, no 1, pp. 230-235.

The core indicators, characterizing the enterprise technological capacity are technical level of production (TL), identified by the technical level of the leading elements of the fixed productive assets (FPA), and wear-out (W). The existing equipment classification is expanded with account for clean zones and premises (CZ&P). Each FPA element group in the classification is associated with the TL (l), adopting values fr om 1 to 7. Accounting for CZ&P is especially relevant while determining the TL of an instrument making enterprises, production of electronic components and optical elements, as well as some assembling industries of machine building.

Basic coefficient of the technical level of production at the enterprise is defined as a weighted average value (l). Weighting factors calculation is performed employing gross book (replacement) value of the group elements adduced to the prices of the current year, using deflator indexes of the Ministry of Economic Development “Fixed Investments”. The calculating formula is based on the effect of labor efforts decrease with the technical level (l) growth, and weighting factor considers the accomplished capital investments has been made. The TL coefficient for particular production method (technological lim it) is defined similarly.

The transitive coefficient of production TL is an extra tool for monitoring and prediction of the technical level of production. Its calculation is similar to the of the calculation of the basic coefficient technical level of production. With this, while weighting coefficients computing, besides the gross book (replacement) value of the group elements the real wear-out of FPA elements is considered. To determine the real wear-put of the elements it is most preferable to employ the probability models approach based on lognormal distribution. The TL transitive coefficient presents interest for the basic TL coefficient trend forming due to the FPA elements disposal. Actual wear-out (W) is determined as a weighted average value of actual wear of FPA elements.

The developed method is based on an accessible input data, and the proposed variables of technological capacity are “tied” with the capital investments.

Nochovnaya N. A., Nikitin Y. Y., Savushkin A. N. Exploring the properties changes of the titanium alloy blades surface after chemical cleaning from carbonaceous impurities. Aerospace MAI Journal, 2019, vol. 26, no 1, pp. 236-243.

It is important to understand how a cleaning technology can change the physico-chemical properties of the material being cleaned after removing carbonaceous impurities from the compressor parts surface of a gas turbine engine. In continuation of the previous work, the creep of VT20 titanium alloy samples was examined, and one of the selected chemical technologies that remove carbonaceous impurities was tested on compressor blades with subsequent determination of some surface properties.

To evaluate the creep of VT20 titanium alloy characteristics, the standard flat samples, some of which were coated with carbonaceous impurities that simulate exploitation, were fabricated. Two titanium compressor blades of a gas turbine engine were used in the research work: blade 1 (small) after operation with a small amount of contaminants on the surface, and blade 2 (large), on which carbonaceous impurities, imitating operation, were coated.

The creep tests results proved that the impurities removal by cleaning solution No. 1, alkaline and acid solutions (“loosening + etching”), and HDL 202 did not reduce the time of the samples destruction and degraded their plasticity, compared to the original samples.

Allowing for the results of the previous work, 9. cleaning solution No. 1 was selected for testing the of carbonaceous impurities removal from the surface of the blades. The results of blades processing revealed 10 that the surface was completely cleaned. In in X-ray microanalysis spectrograms the elements such as sulfur, oxygen and carbon, indicative of the presence of carbonaceous impurities, are missing. The values of surface roughness and micro-hardness did not sustain significant changes. Processing in the indicated solution leads to activity (potential) increase of the of blade No. 1 surface. The lower values of the blade No. 2 surface potential were observed (about 10%) compared to the initial state.

Kolesnikov A. V., Mikhailov I. V. Superplastic forming of aerospace facilities’ parts and multilayer structures from vt20 titanium alloy. Aerospace MAI Journal, 2019, vol. 26, no 1, pp. 244-250.

The structures from titanium allows are increasingly employed in aerospace structures. Labor intensity may be significantly reduced while titanium parts manufacturing by application of superplastic forming process (SPF) and combined process of SPF and diffusion welding (SPF/DW). Superplasticity manifests itself in alloys with a fine-grained structure under certain strain-rate conditions and maintaining a constant temperature during the formation process. Maintaining a constant strain rate in the process of shaping is ensured by a continuous change in the forming pressure over time. The computing of the plot of the forming pressure change with time is rather labor consuming. For this problem solving and the process visualization, modeling with the MSC “Marc” program was performed.

By the example of forming a cellular panel from VT20 titanium alloy, the possibility of manufacturing parts by the SPF method is demonstrated. The simulation result allowed obtain relative deformations distribution, which analysis revealed that maximum relative deformations constituted 97%. This is quite acceptable, and there will be no destruction while the forming process. The simulation results allowed also develop the control program according to which the cellular panel was produced by the superplastic forming press.

The article considers the form shaping modeling of multilayer wedge-shaped panels with transverse and longitudinal corrugation set. It follows from the relative deformation distribution analysis that maximum relative deformations in the structure constituted 126.6%, which is acceptable. The forming of the wedge-shape three-layer panels was performed by the SPF/DW method according to the computed plot of forming pressure change with time.

After the superplastic forming process, there are no both corrugation forming and springing effect, which eliminates the finishing work.

Thus, the SPF and SPF/DW technologies and modeling the process of production to obtain the forming parameters allow significantly enhance the production possibilities while producing complex parts from titanium alloys.

Klimov V. G., Nikitin V. I., Nikitin K. V., Zhatkin S. S., Kogteva A. V. Wear-resistant composites application in repair and modification technology of the gtd rotor blades. Aerospace MAI Journal, 2019, vol. 26, no 1, pp. 251-266.

The of production and operation costs of gas turbine engines employed in aviation, oil and gas or energy industries constitutes a significant portion of costs reducing the net marginal profit of operator- enterprises. These costs reduction is a natural desire of any holding. Against this background, the ability to maintain the resource of the gas turbine engine at the lowest cost to itself remains the main criterion of the competitiveness of the producer in the market.

It should be kept in mind, that the operation costs of gas turbine engines through their lifetime cycle often exceed their original cost. To be precise, the effective repair technologies often stops the loss risks in future orders.

A distinctive feature of domestic aviation gas turbine building is a low assigned and overhaul period of the engines operated according to the first performance strategy.

Often the causes of understated life cycles are the imperfections of the structures that occur at the development stage. Consequently, the presence of the extremely expensive parts and units with a relatively short lifetime requires their permanent replacement or recovery. These parts are the rotor blades, and the turbine stator. Many factors can lead to their failure, starting from structure changing due to the uneven temperature fields, to the loss of geometry due to burn­out or mechanical damage. The last one is the factor, most frequently occurred in the products.

From the viewpoint of repairing technologies, the turbine blades recovery is the most cost-effective among all other engine parts. The cost of the engine hot section (turbine) producing exceeds the cost of a cold section (compressor) by average of 400-700%. However, the repair complexity remains the main obstacle in its implementation.

This article proposes to employ the high-temperature nickel powders of the VPR type as wear-resistant surfacing materials applied by laser action. The structure formation peculiarity of the described materials is revealed. It is manifested at high cooling rates in the form of natural composites formation with dispersion eutectic hardening along the boundary of the dendritic framework. This structure has a non­directional arrangement of strengthening phases that increases the wear-resistant characteristics of the resulting composite.

The original method of restorative surfacing is described. It allows repairing and modifying rotor blades of gas turbine engines (GTE), with increasing the wear-resistant characteristics of the part contact surfaces. Based on the conducted comparative studies, including analysis on a scanning electron microscope; measurement of micro-hardness and the coefficient of the materials linear expansion; testing of abrasion resistance of cladding and their fatigue strength, the possibility of VPR type materials application of as an alternative to classical wear-resistant composites with mechanical admixture of various carbides was proved. It is shown that under conditions of pulsed laser action at high cooling rates, the average hardness and overall resistance to abrasive wear of certain VPr alloys grow due to the formation of a finely dispersed stable eutectic structure close to the initial powder material. The positive performance characteristics of alloys of VPr 11-40N and VPr 27 grades were obtained, which allows employ them when rebuilding the GTE rotor blades.

Kuznetsov E. N., Lunin V. Y., Panyushkin A. V., Chernyshev I. L. Boundaries of non-separation flow-around of bodies of rotation, with the nose part in the form of Riabouchinsky half-cavity. Aerospace MAI Journal, 2018, vol. 25, no 4, pp. 7-15.

Bodies that are optimal at the so-called low critical Mach number M*, at which at least one sonic point on the body flown-over surface occurs, were studied theoretically in Ref. [1]. It was confirmed that M* achieves its maximum value and, consequently, the wave drag minimum value occurred for the body identical to the Riabouchinsky finite cavity in the classical theory of incompressible fluids. It was experimentally studied in Ref. [12], which demonstrated that in the transonic velocities range the Riabouchinsky half- cavity had the smallest drag among the bodies of rotation with the same aspect ratio  λ=L/D=0.87(where L is the nose part length and D is the diameter of its mid-section). This conclusion is incorrect for aspect ratios λ>2 due to the friction impact the drag as it follows fr om Ref. [24]. The absence of turbulent boundary layer separation from the side surface of the body of rotation under study at zero angle of attack in the range of Mach numbers 0.8≤M≤0.95 was demonstrated in Ref [17].

The main objective of this work is determination of angles of attack αsep at which turbulent boundary layer separation from the side surface of the studied body occurs. The study was performed with NUMECA FINE/Open software based on Reynolds Averaged Navier-Stokes equations (RANS). The solution of the problem was performed in the framework of fully turbulent flow model without accounting for laminar-turbulent transition using Spalart-Allmaras (SA) and k-ω SST turbulence models. To determine the boundaries of the non-separated flow-around computation was performed in stationary problem setting at various angles of attack. With that, the flow separation indicator was the presence of the zone on the model surface wh ere the friction coefficient Cf < 0. The results obtained with two turbulence models are close to each other, and the difference between the two separation angles does not exceed 1°.

The results of the study obtained for αsep for the nose part with aspect ratio of are as follows:

αsep=15° for М=0.5, αsep=9° for М=0.65,

αsep=12° for М=0.8, αsep=13° for М=0.85,

αsep=5° for М=0.9, αsep=11° for М=0.95.

Computing results for the longer nose part with aspect ratio are:

αsep=20° for М=0.5, αsep=13° for М=0.7, αsep=21° for М=0.9, αsep=18° for М=0.95.

The angles of attack αsep which realize turbulent boundary layer separation from the side surface of the investigated body at free-stream Mach numbers 0.5≤M≤0.95 were obtained. Separation zones location is shown for various models and modes.

Bragin N. N., Kovalev V. E., Skomorokhov S. I., Slitinskaya A. Y. On evaluation of buffeting of a swept wing with high aspect ratio at transonic speeds. Aerospace MAI Journal, 2018, vol. 25, no 4, pp. 16-27.

The article presents to the development of a technique for buffet initialization boundary evaluation, occurring on a swept height aspect ratio wing at increasing angle of attack during cruising flight modes. The lift coefficient value of the buffet onset is one of the limitations that should be accounted for while designing the win aerodynamic layout of a subsonic aircraft. According to the norms, the margin between the cruise flight mode and the CLbuff value or the lift coefficient of the buffet onset should be at least 30%. Thus, knowing the CLbuff value through the entire operational speed range is a prerequisite for an aerodynamic wing configuration design beginning from the preliminary design stage. The problem of determining the CLbuff has become of special urgency at the transonic speeds due to the substantial aspect ratio increase of the (by 15–20%) of the long-range aircraft wings due to the composite materials application in load-carrying structure.

Despite the successes in CFD aerodynamics gained over the last years, non-stationary separation flow modes are studied, basically, using experimental tools, including wind tunnel tests of airplanes high scale models. Though the cost of such studies is high, they ensure required reliability of the obtained results. It is worth mentioning, that the time consuming computations on multiprocessor computers are costly as well. With this, the high accuracy and reliability of the obtained results are not guaranteed. Preparing mathematical model and building-up computational meshes with hundreds of millions of nodes are commensurable with costs of developing and manufacturing scaled models for tests in the wind tunnel. Thus, numerical methods do not always prove to be less labor consuming and costly than the experimental ones. Despite the fact that computer facilities and software develop rapidly, and the situation gradually changes, experimental methods remain as before the basic tool while performing the studies of complex flows.

The article presents the analysis of typical features of the wing flow-around at the angles of attack corresponding to the start of the buffet mode. The technology of application of the program for computing transonic flow-around based on the full potential method for the buffet initialization computing is demonstrated. Computational results comparison with the data of experimental studies obtained for the model of the wing with fuselage in the wind tunnel is presented.

Danilenko N. V., Kirenchev A. G. Vortex formation of gravity flows. Aerospace MAI Journal, 2018, vol. 25, no 4, pp. 28-36.

Vortex formation analysis of the air medium as a gas turbine engine (GTE) propellant allows extracting one of its specifics, namely the gravity character of the technogenic vortices formation. The fudamentals of this vortex formation are being subordinated to the natural vortexes formation of the Earth atmospheric environment. The theory of technogenic vortices of the GTE operating on the ground, the same as the theory of natural atmospheric vortexes is being in the state of its development. It confirms the state of the issues of the working process principle and classification of both technogenic and natural vortices stated in scientific and technical literature and textbooks on the theory of gas turbine engines and metrology. The most informative is the database on meteorological studies of natural vortices (tornados, cyclones, circulations and atmospheric fronts). Thus, due to the technogenic and natural gravity vortex forming similarity, the gist of technogenic vortices' work process should be searched for in the gist of the cyclonic type vortices of the environment. The work process study herewith of the cyclonic type vortices (tornados and others) may be the basis for creating a theory of natural vortex forming.

The problem is set to study the work process of small-sized technogenic vortices with their subsequent adaptation to the work process of natural vortex forming.

The above said problem should be solved relying on the basic equations of gas dynamics (gas flow energy preserving and other) with subsequent yield to the methodology and essence of the gravity type vortices' work process. The most accessible to learning the gravity vortex formation and its vorticity is the energy conservation equation, including its components in the form of internal and kinetic energy of a gas flow, kinetic energy of the environment angular rotation, and heat exchange elements in the form of external mechanic work and heat. Hence, extracting the master unit of vortex formation (angular rotation energy) allows establish functional dependence of the vortex formation under study from the sources of energy capable of generating gravity vortices of various types.

The article presents the methodology of studying, and analysis of the problems of work process cognition of vortex movement of the Earth's ambiences. Classification of vortices according to the gist of their work process. The article indicates the way of splitting the vortex formation into the vortices according to the gist of the work process included into its classification, and further cognition of their physical entity and exploration resulting in vortex characteristics, their consequences and application areas.

Yudintsev V. V. Rotating space debris objects net capture dynamics. Aerospace MAI Journal, 2018, vol. 25, no 4, pp. 37-48.

By now, several methods for near Earth orbits active cleaning from large-size space debris were suggested. The most difficult stage of such mission is the stage of space debris capture. Capturing method selection and subsequent orbital transportation of space debris depends on its type and angular motion. Rockets' orbital stages may rotate with high angular velocity, which aggravates their capture by manipulators and other means. One of prospective techniques of such object capture is application of a net connected with the space tug by a tether. The object capture by a net can be performed by the net separation with a certain relative speed in relation to the space tug and space debris, or by the net unrolling on the trajectory of the space debris object relative to the space tug. Elastic properties of the net and tether allow reduce the load acting on the space tug while an object capturing process and control the value of this impact.

The paper presents discrete mathematical model of the net movement as a system of material points' elastic interaction, as well as these components interaction with the space debris surface. The possibility of capturing an orbiter type object, rotating with significant angular velocity was demonstrated through this model. The article demonstrates that capturing the object, rotating with angular speed of 5 degrees per second, requires the speed of the net relative to the space debris from 2 to 5 m/s. To capture an object, rotating with angular speed of 30 degrees per second, the net speed should be no less than 10 m/s.

Bolsunovskii A. L., Bondarev A. V., Gurevich B. I., Skvortsov E. B., Chanov M. N., Shalashov V. V., Shelekhova A. S. Development and analysis of civil aircraft concepts employing integration principles. Aerospace MAI Journal, 2018, vol. 25, no 4, pp. 49-63.

The search for the technologies allowing significantly improve operational characteristics of the prospective civil aircraft was the goal of the work.

The study of innovative technologies, including those, providing the airframe and engine integration, was performed applying methods of alternatives analysis, based on the factorial analysis, and experiments planning assuming performing a series of computing experiments with subsequent comparison of their results.

Three possible innovations trends were considered:

– application of a turbojet engine with the increased bypass ratio for the fuel consumption and noise-at-terrain reduction;

– application of the so-called distributed power plant with the separated gas generator and the fan connected by mechanical transmission;

– airframe and power plant elements integration for obtaining useful effects in aerodynamics and structure, as well as and obtaining new operational properties.

According to these independent principles a number of the long-haul aircraft possible configurations, differing by various combinations of the bypass ratio, the turbojet schemes and technologies of elements integration into the “airframe + engine” system was developed. The number of possible strategies of the integral aircraft, including the base option, corresponds to the number of binomial coefficient of the three factors 23 = 8 according to performing the full-factorial experiment.

A multidisciplinary expert assessment of aircraft configuration options was performed, which turned into the basis for the most effective concepts selection.

Comparison of possible characteristics revealed that some options of airframe and engine integration under consideration had potential for considerable of fuel consumption reduction compared to the conventional long-haul aircraft configuration. The results of the study allow recommend two strategies for further studies. These strategies also possess potential for additional improvement of the other operational characteristics, such as noise-at-terrain and operating safety. Other configurations possess a number of useful elements that can find application while critical technologies development and reduction of technical risks.

Egoshin S. F. Impact evaluation of multi-propeller wing blow-over system on the stol aircraft characteristics. Aerospace MAI Journal, 2018, vol. 25, no 4, pp. 64-76.

The article undertook the attempt to obtain an analytical solution to compute the take-off length of an aircraft equipped with the multi-propeller wing blow-over installation, and estimate the benefits of such engineering decision through the transport operation evaluation of this aircraft.

The main difficulty of this problem lies in the fact that the wing and propeller interaction is an extremely complex and insufficiently studied task. Currently, only approximate semi-empirical formulas for calculating the aerodynamic characteristics of the wing at small relative diameters of the blowing airscrew exist. The exact calculation of the wing flow-around in this case is possible only for strictly specified, nonparametric configurations using the finite-difference method.

In addition, the current complicated situation in the sphere of local air transport in Russia (reduction of airlines and operating airfields) requires the search and evaluation of effective technical solutions for a prospective aircraft of local airlines. One of the ways is envisaged as developing a short takeoff and landing (STOL) aircraft capable of carrying out transport operations in conditions of an underdeveloped airfield infrastructure. It is considered that equipping such a STOL airplane with a multi-propeller electrically powered blow-over system will be an effective solution to the problem. However, the above said complex aerodynamic task does not allow a quick search for the optimal characteristics of this aircraft.

The developed mathematical model, under certain constraints, allows obtain an evaluative analytical solution for the take-off run length of such STOL aircraft, reveal the specifics of parameters interaction and evaluate possible advantages and disadvantages of the aircraft. Within the framework of the model, it was demonstrated that the maximum possible power consumption from the main engines is the optimal value of the corresponding parameter of the electric power plant. The amount of this power consumption determines the blown part of the wing area through the relationship with the critical rotation speed of the auxiliary propellers.

As for performance of a transport STOL aircraft based on L-410, it was shown that a blow-over system based on conventional electrotechnical materials can reduce the take-off run by 30% (up to 300 m), while reducing the payload by 1520% at flight ranges up to 300 km or up to 50% when flying to the maximum range. At the same time, if the electric power plant is designed based on high-temperature superconductors (HTSC), the payload reduction will be much less: negligible at flight distances up to 300 km or about 25% with flight to the maximum range. This allows conclude that the HTSC technology application for such STOL aircraft creation is rather promising.

Lopatin A. A., Nikolaeva D. V., Gabdullina R. A. Experimental data generalization on heat transfer in cooling system with axial sectional finning in conditions of free convection. Aerospace MAI Journal, 2018, vol. 25, no 4, pp. 77-85.

At present, power electronic components with high heat release have been widely applied in various fields of modern industry. The main problem the developers of the element base are facing consists in creating cooling and thermal stabilization systems capable of removing heat fluxes of high density, while working in a wide range of ambient temperatures. When creating such systems, it is necessary, alongside with the thermal ones, to account for the mass-size characteristics of the device as a whole. Thus, much depends on the heat transfer intensification method selection.

Quite enough attention is paid in modern literature to the issues of radio electronic equipment thermally loaded elements, as evidenced by a significant number of articles and monographs on this topic. Heat release is one of the main causes of the unstable operation of radio electronic equipment. Among the basic factors exerting maximum destructive effect, the increased temperature of the elements is one of the main ones. Thus, the devices operation in the field of the radio electronic equipment is closely associated with heat removal from the thermally loaded components. Depending on the structure and shape of the cooled components, various solutions are employed for organizing continuous heat removal.

Certain problems of large heat fluxes removing in the elements of industrial electronic equipment were considered in [1-3]. The correct choice of the cooling system type ensures trouble-free operation of all cooled components of the device.

A considerable amount of publications in modern scientific publications, both in Russia and abroad, attest to the considerable interest in the issues related to the heat transfer intensification for surfaces of various shapes and sizes as applied to cooling systems for electronic equipment. The issues of heat transfer intensification are set forth in [4-10]. In particular, the criterion equations of various authors for the Nusselt number computing for natural convection are presented in [2, 9]. Experimental studies of the convective heat transfer intensification in rectangular dissected channels and in channels with discrete turbulators were described in [1, 10]. In the studied dissected channels, a process of rational convective heat transfer intensification was implemented, reliably controlled by changing the values of the basic dimensionless geometric parameters. The generalizing dependences for discrete-rough channels were obtained in [7] for free convection conditions, and flow modes and mechanisms of intensification were studied. In [11-14], the authors experimentally studied one of the parameters characterizing the cooling systems both qualitatively and quantitatively, namely, the thermal resistance.

The fins application as a method of heat transfer intensifying leads to the increase in the heat transfer coefficient value by a factor of tens. This method of intensification implies a wide variety of vatious types of fins, such as: longitudinal, transverse, rolling, spiral and many others [15,16]. In [15] the analysis of the expediency of employing different types of fins from the viewpoint of the coolant aggregate state is presented. The optimal edges number selection is presented in [16]. The heat transfer intensification of the systems with a cut-off fin is also considered in [17-20].

The purpose of this work consists in studying the efficiency of the split finning under conditions of natural convection. A test bench was developed for performing the experiment on the study of heat transfer. While the experiments on heat transfer near the cut edge under conditions of natural convection, criterion dependencies were obtained.

Relying on the analysis of literature sources and accounting for the results obtained while experimental studies, the authors established that from the viewpoint of the of axial split finning practical applicability, there are a number of specifics, associated primarily with the fact that the “petals”, obtained as a result of dissection of the heat exchange surface, can be considered as independent ribs. The studies of heat transfer intensification under conditions of natural convection with the cut ribs application were conducted and presented as a result of the work. While the experiment, the effectiveness of the use of split finning is demonstrated, and the most optimal geometric parameters of the working area were revealed. The process of heat transfer was visualized. The boundary layer thickness near the cut edge was computed. Criterion dependencies for heat transfer computing of the systems with axial cut fins were obtained. The prospect of this study is of experimental data verification by numerical modeling programs.

Zichenkov M. C., Ishmuratov F. Z., Kuznetsov A. G. Studying the gyroscopic forces and structural damping joint impact on the wing flutter of the aeroelastic euram model. Aerospace MAI Journal, 2018, vol. 25, no 4, pp. 86-95.

The article deals with the structural damping role analysis while studying the gyroscopic forces impact on the flutter speed. The algorithm for accounting for gyroscopic forces in polynomial Ritz method while computing the aircraft aeroelasticity dynamic characteristics, developed earlier by the authors, was employed. The algorithm was realized in the KC-M software developed in TSAGI and validated while solving aeroelasticity problems in many practical applications.

The computations were performed on the example of the wing of the well-known aeroelastic research model of the four-engine long-distance aircraft EuRAM (European Research Aeroelastic Model), developed and studied in the framework of the European project 3AS (Active Aeroelastic Aircraft Structure). The model characteristic feature is the flutter form occurrence associated with the lateral vibrations of external engines. This form is affected by the gyroscopic forces due to the engines rotating rotors.

The flutter characteristics analysis at various levels of structural damping (characterized by logarithmic vibrations decrement δ ) revealed that the vibrations tones interaction character with account for gyroscopic effect was not principally changed. It was found herewith, that the gyroscopic forces impact on the speed of the considered flutter form might be of different sign depending on the level of the structural damping.

For example, at δ = 0.02 the flutter speed increases by 11.5%, with the maximum value of the engine rotor kinematic momentum (scaled to the model). While increasing the structural damping value, in the beginning, the gyroscopic forces' impact on the flutter speed decreases, it does not practically exist at δ = 0.04, and with further increase of the decrement the impact changes its sign, and the flutter speed decreases. The flutter absence was marked at δ > 0.046 in the range of small rotor rate speed, but while the rate speed increase the flutter may occur. Its speed decreases at that (about 10%) with the rate speed increase. This indicates the importance of accounting for the dynamics of rotor systems while the aircraft aeroelastic phenomena analysis.

The obtained the results were confirmed also by the finite element computing method in NASTRAN system using Rotordynamics module (accounting for the rotor systems dynamics). The computing results on gyroscopic forces impact on aeroelasticity characteristics at various structural damping values performed with KC-M software accord well with computations performed with NASTRAN software.

It was noted that while experimental validation of the gyroscopic impact on the flutter speed of the model in the wind tunnel various results might be obtained depending on the structural damping level. Thus, the detailed computational and experimental analysis of the model dynamic characteristics is required while such tests preparing and running.

Ezrokhi Y. A., Kalenskii S. M., Morzeeva T. A., Khoreva E. A. Analysis of a concept of the distributed power plant with mechanical fans drive while integration with a “flying wing” type flying vehicle. Aerospace MAI Journal, 2018, vol. 25, no 4, pp. 96-109.

The article presents analysis of a concept of the distributed power plant (DDP) while its integration with a “flying wing” type flying vehicle.

A modified airframe model of a prospective long-range aircraft (LRA) of PJSC “Tupolev” development with two power plants integrated into the tail-end was selected as a flying vehicle.

Those power plants represent a bypass turbojet engine where two taken-out fan modules are driven by mechanical transmission from fan turbine of this turbojet. The choice in favor of a mechanical way of power transfer for the aircraft of 2030 level is based on the results earlier performed studies on the engines of new schemes in CIAM named after P.I.Baranov.

The results obtained while numerical modeling of the flow on the upper surface of an LRA airframe were also employed. This modeling revealed that for a long-range flight the mean values of the full pressure's losses prior to the fans differed greatly and depended monotonously on the flow deceleration level in the air intake. According to the calculations, the average value the full pressure restoration coefficients was correspondingly ~0,923 for the first fan module, ~0,952 for the bypass turbojet and ~0,958 for the second fan module.

Refining of the earlier developed model of the distributed power plant was performed to evaluate the impact of the conditions at the inlet of each of fan modules. The performed of mathematical model refinement allowed implementing independent selection of parameters, dimensionality and gear-ratios of reducing gear for DPP fan modules drive, as well as performing independent regulation of output devices of these modules.

The article considers separately the impact of the two main factors on the engine thrust, namely, the fall of the full presure level at the inlet, and its proper heterogeneities.

Calculations revealed that for the earlier selected DPP option while its integration into the flying vehicle under consideration, regulation of nozzles of the turbojet bypass loop and fan modules was required at the cruising mode. With this, gas temperature increase prior to the turbine by ~70 К was required.

Three different variants of the engine which allow excluding the above said regulation were investigated while this work.

The first variant is a version with fans equal by dimensionality and pressure ratio at the designed cruising mode.

The second variant is a version with the first fan module with the pressure ratio increased by 5% relative to the BTJEs fan at the cruising mode.

The third variant is a version with first fan module air consumption decreased by 50% at the cruising mode.

Parametric studies performed employing the develop methodology allowed selecting the degree of bypassing and the degree of pressure increase in the fan optimal by the specific fuel consumption at the cruising mode for each DPP option. The dimensionality of fan modules and main DPP units was refined with account for various losses levels at the inlet.

Analysis of effects associated with the presence of a non-uniform field of the full pressure and leading to its average level decrease at the fan inlet revealed that impact of the presence of non-uniformity might be from 15 to 30% of total impact on the engine thrust.

At the same time, while the power plant parameters selection at the cruising mode with account for the degraded coefficients of the full pressure preservation prior to the fan, the fall of the thrust level due to the proper non-uniformity might be ~2,53% at the given mode. This should be accounted for while selecting an optimal DPP appearance of the configuration under consideration.

Panov S. Y., Kovalev A. V., Aisin A. K., Achekin A. A. Aircraft air intakes location impact on vortex formation intensity. Aerospace MAI Journal, 2018, vol. 25, no 4, pp. 110-119.

Kalugin K. S., Sukhov A. V. Methane application specifics as a fuel for liquid rocket engines. Aerospace MAI Journal, 2018, vol. 25, no 4, pp. 120-132.

At present, a significant part of the research aimed at increasing the energy and mass characteristics of liquid-propellant rocket engines (LPRE) is being performed in the field of new materials application and processing technologies. Other studies are aimed at modernizing the principle of organizing the work process. However, new fuels application is a more long-term and quickly realizable prospect notwithstanding the research costliness in the field of LPRE. Methane may appear one of the propellants, which application may become a new stage in rocket and space industry development. The article considers the historical process of methane formation as a liquid rocket fuel component since it was firstly mentioned by Tsiolkovsky in his works (“Exploration of the World Spaces with Jet Devices”, “Space Rocket”, “Jet Airplane”, “Achieving the Stratosphere”, etc.). A comparative analysis of methane with kerosene was performed in view of the similarity of the work process organization in the LPRE combustion chamber, as well as close hydrocarbon structure. A component close to methane, currently in use in rocket engines, is hydrogen due to the cryogenic nature of both components, which creates difficulties at the design stage of valves, pipelines and gas lines, as well as the organization of the work process in the combustion chamber. Additionally, analysis of the most common fuels based on kerosene, methane and hydrogen was performed. This is especially interesting, since methane fuel pair of oxygen-methane occupies an intermediate position between “oxygen-kerosene” and “oxygen-hydrogen” pairs with respect to the specific impulse and fuel mixture density. The analysis was performed based on physical-chemical, energy, operational, environmental, economic and some other characteristics. This allowed identify the main advantages and disadvantages of methane application as a LPRE fuel and determine its prospects both in Russian and foreign rocket and space industries.

A brief analysis of liquid rocket engines on methane, created or projected in NPO Energomash by V.P. Glushko, KBHA them. S.A. Kosberg, KB Himmash them. A.M. Isaev, the Research Center. M.V. Keldysh, and also to the American firm SpaceX.

Finally, it was concluded that the methane LPRE could replace oxygen-kerosene engines in the near future, since the fuel pair oxygen-methane outperformed the oxygen-kerosene pair by its energy, environmental and economic indicators. Interplanetary flights can become a special field of methane application, since a large amount of methane, the main element of natural gas, can be found almost everywhere in the solar system: on Mars, Titan, Jupiter and many other planets and satellites, which will allow refueling rockets in flight, significantly increasing them the flight range.

Kuz'michev V. S., Omar H. H., Tkachenko A. Y. Effectiveness improving technique for gas turbine engines of ground application by heat regeneration. Aerospace MAI Journal, 2018, vol. 25, no 4, pp. 133-141.

The requirements for the ground based gas turbine installations efficiency improvement are constantly increasing.

Heat conversion into work in gas turbine engines, operating by the Brayton cycle, is attended by significant losses, which depend on the cycle parameters and reach up to 60-70% or more. At present, high-tech aircraft engines and their modifications are widely used as ground-based gas turbine plants, making provision for the gas turbines efficiency improvement based on the of combined thermodynamic cycles application.

The article considers the schemes of gas turbine units (GTU) for ground application with combined cycles, allowing improve their efficiency. One of the ways for the gas turbine units cycle improving is heat regeneration of the exhaust gases by installing a heat exchanger at the turbine outlet where a part of the heat is transferred into the air behind the compressor. However, relative bulkiness and substantial weight of the heat exchanger (even of plate type) do not allow at present active application of this scheme in aviation, but it is widely employed in ground applications.

In the case of ground based gas turbine unit, heat exchangers are located in the exhaust chamber or tower behind a power turbine. Thermal ratio of the most widely used tubular heat exchangers is  θ = 0.8-0.9, and plate- type heat exchangers are characterized by the thermal ratio of θ = 0.5-0.8.

It is obvious that the main parameters of the thermodynamic cycle of gas turbines unite, such as the gas temperature T*g and the compressor pressure ratio (π* ), as well as the parameters determining the working process of additional units (heat regenerators, steam turbine, etc.) of the combined installation play an important role in its efficiency improving. Comprehensive optimization of these cycle parameters is the main goal of the gas-turbine combined unit thermodynamic design.

Computer models of a gas turbine unit with combined thermodynamic cycles developed in the ASTRA SAE-system allowed solve the problems of nonlinear multi-criteria optimization of their operating parameters, determine the most rational schemes depending on the intended purpose and operating conditions of the gas turbine unit.

Russian Turbofan engine TRDDF RD-33 was selected as the basis for studying the heat regeneration impact on efficiency effectiveness. Its low pressure compressor was cut off to eliminate the bypass duct while converting it into ground based installation.

The following variation values of the cycle basic thermodynamic parameters were selected (π*kΣ = 15, 30, 45, 60 и T*g = 1200, 1500, 1800, 2100 K). The GTE module without exhaust gases heat regeneration and a GTE with exhaust gases heat regeneration were developed employing ASTRA computer program. The paper presents some results of the study on GTE efficiency improvement.

Ogloblin D. V., Gorelov A. D., Voroshilin A. P., Zueva K. S. Automated testing system for technical diagnostics of spacecraft power supply systems. Aerospace MAI Journal, 2018, vol. 25, no 4, pp. 142-151.

One of the urgent tasks at present consists in time reduction of flying vehicle launch preparation through introduction of new technologies, equipment and various kinds of tests based on computational experiments.

One of the most expensive types of aircraft is a spacecraft for various missions and tasks in the near-earth space and interplanetary flights. The main system of a spacecraft is the power supply system. Stringent requirements on external impact stability, and operability maintenance in emergency situations, since its failure results in the spacecraft loss. The power supply system preparation and testing is a major part of the testing program and it is performed employing rather labor-intensive methods.

One of the most interesting objects for the test complex optimization are the spacecraft for astronomical observations, as they present a set of a large number of complex technical devices for various purposes and control systems. Such aircraft requires a special approach to ensuring the quality of various electrical systems tests, as well as control and monitoring systems.

It is important to ensure uninterrupted power supply of the onboard service and scientific equipment for the timely data obtaining during the mission. Thus, the task of rapid and high-quality electrical tests performing of such spacecraft is of paramount importance.

The following core systems are being subjected to comprehensive electrical testing: the onboard radio telemetry system, onboard control system, propulsion system, the solar panels orientation system, power system, electrification control system.

Besides scientific equipment, these systems form the basis of almost any spacecraft. Due to the large number of systems subjected to electrical checks, the issue of the electrical tests time reduction, while preserving their quality (guaranteed reliability level of the systems) is relevant. It is necessary to determine the level of reliability and the number of tests based on the system model.

This above said problem can be solved by optimizing the measurement devices' number and functional characteristics, as well as application of automated measurement systems (AIC) for processing a large number of parameters (with account for specifics of electrical tests). This solution allows optimize the testing process, while reducing herewith the number of employed measuring equipment and the test program cost.

The upgraded AIC version allows combine system data into one subsystem of the power bus monitor. It also ensures centralized output of the test data to the Central computer, reducing the number of additional jobs, and simplifying the operators work.

The final layout option has a more optimized build structure. The structural diagram of the modernized automated testing system for electrical testing is presented.

Application of the reliability computation model allowed estimate the number of necessary checks, and the proposed system ensured the electrical tests duration reduction by 30%.

Implementation this complex of has increased the voltage, current and resistance parameters measurements accuracy by 0.01%.

This set allows define the parameters of voltage, as well as the leakage current and the signal occurrence at a specified time instant. This allows localize and fix the problem with the power supply during the test.

The software has a flexible customizable interface that allows quickly respond to changes and emergencies.

The proposed complex can be employed for spacecraft testing and, first of all, telecommunication satellites for various purposes based on of the “Navigator” platform. As examples of products using this platform, we can cite the spacecraft of the Arctic family, Electro and Spectrum.

Sedel'nikov A. V., Puzin Y. Y., Filippov A. S., Khnyreva E. S. Soft hardware efficiency estimation for a small spacecraft rotation angular velocity provision and monitoring. Aerospace MAI Journal, 2018, vol. 25, no 4, pp. 152-162.

The article presents the efficiency estimation of an AIST type small spacecraft rotation angular velocity provision and monitoring, employing the improved modification of the software hardware compared to the basic model, undergone flight tests as a part of the flight and trial samples of the “AIST” small spacecraft.

The article deals with magnetometer sensors application as a measuring tool for estimating the rotational motion parameters for various types of spacecraft.

The gist of the article is narrowed down to the fact that for there is no problem of mounting place selection for magnetometers and magnetic actuators (electromagnets) installation on the large-weight spacecraft and space stations. At the same time, for the small spacecraft this problem is topical in view of the impact of onboard scientific and supporting equipment together with magnetic devices of orientation hardware, while magnetic moment work out, on magnetometer sensors measuring data.

The article analyses in detail operation results of a number of devices, containing magnetometer sensors, as a part of various spacecraft types, such as “Foton” No 2,“AIST”, “AIST-2D”. The analysis results revealed that for “Foton” No 2 weighting 6546 kg, almost all deviations in magnetometer sensors measurements could be explained by the measurements instrumental error, while for the “AIST” series spacecraft weighting up to 50 kg the sensors readings demonstrated significant discord.

The authors conclude that the problem of magnetometer sensors' measurement data correctness for small spacecraft is specific due to the essentiality of the impact of scientific and supporting equipment operation, which arises from the dense layout of the payload in the inner space of a small spacecraft.

One of the ways of this problem solving may be the software hardware development, accounting for this impact for each normal operations mode of the equipment.

As an example of such solution, the article presents the software hardware installed on the “AIST-2D” small spacecraft. In the case of the “AIST-2D” small spacecraft, the discrepancies in estimating the angular velocities by two different magnetometer sensors were much lower than for “AIST” small spacecraft. This was achieved by the improvement of software hardware.

Donskov A. V. Analysis of modern evaluation and modeling methods of contingencies occurrence risks onboard a spacecraft. Aerospace MAI Journal, 2018, vol. 25, no 4, pp. 163-169.

The purpose of the article consists in analysis of modern evaluation and modelling methods of contingencies occurrence onboard a spacecraft and substantiation of methods selection for their subsequent application in the process of operative flight control. The process of the aircraft flight control while contingencies occurrence onboard a spacecraft (time limitation for the decision making on contingencies parrying and their type, high dynamics of the processes flow, multicriteriality of the spacecraft current state, the presence of sources of uncertainty) was studied. The inference was drawn, that all considered methods of the onboard contingencies risks evaluation and modelling were not exhaustive. Depending of the current situation, any of the considered methods and on account of the problems being solved could be employed for the contingency risk occurrence onboard a spacecraft evaluation and modelling. The accumulated experience of flight control in manned astronautics revealed that the most interest was provoked by those methods of contingencies risks evaluation and modelling, which reflect the ways of their evolution and aftermath. The selection of the contingencies occurrence risks onboard a spacecraft methods (logic-linguistic and theory of fuzzy sets) is being substantiated by the fact that it allows develop scenarios of the contingencies occurrence onboard a spacecraft and prepare initial data for the decision making on contingencies parrying in case of uncertainty. Methods of contingencies occurrence risks onboard a spacecraft considered in the article may be implemented as tools both in the systems for decision making support on contingencies onboard a spacecraft parrying and in expert systems.

Kovalev A. A., Tischenko L. A., Antipin M. A., Shakhovtsev M. M. Oxide film homogeneity provision on the surface of silicon monocrystal substrates while their thermal oxidation process. Aerospace MAI Journal, 2018, vol. 25, no 4, pp. 170-177.

The article deals with the impact of the thermal oxidation process technological parameters, such as the work process temperature and the value of oxidizer reagents consumption, on the oxide film uniformity on the silicon substrate surface. It evaluates the heat-treatment furnace preheating temperature range by modelling in ANSYS with the specified boundary conditions. The homogeneity characteristic, being computed by Min-Max method, was used for homogeneity estimation. The type of silicon wafers under consideration satisfied the number of parameters, such as d = 100 mm, thickness 400 µm, the polished silicon emissivity ε = 0.89, silicon thermal conductivity λ = 31.2 W /(mK), silicon density ρ = 2.33 g / cm3. The preheating temperature range for an atmospheric thermal furnace with a silicon carbide boat is 550–700 °C. Increase in the oxide film distribution homogeneity over the silicon substrates surface due to the uniform distribution of the temperature field was observed as the result of simulation. As a consequence, chemical reactions were close in the intensity of the oxidation processes flow. Experimental validation of the homogeneity increase was obtained due to the thermal furnace boat preheating.

Optimal values of oxidant reagents (O2, H2O) consumption at temperatures of 1100, 1000 and 850 °С in the two-phase heating furnace (heating – dry oxidation, work process – wet oxidation) were obtained experimentally. These values allow producing wafers with high-quality silicon oxide (U < 1%).

The article gives technological recommendations on high-quality oxide film provision on the surface of monocrystal silicon wafers by furnace preheating and maintaining temperature between the thermal oxidizing processes, as well as oxidant reagents consumption values selection, allowing producing wafers with high-quality silicon oxide.

Na L. ., Zhefeng Y. ., Yi F. . Shock absorber using inward-folding composite tube and its application to a crew seat: numerical simulation. Aerospace MAI Journal, 2018, vol. 25, no 4, pp. 178-188.

This paper presents an innovative energy absorber consisting of an inward folding of composite tube, which is cut axially and turned into the inner of the itself. There is no excess of composite fragments after the composites destruction, and the debris will fill in the inner part of tube to increase the energy absorption. The impact energy is absorbed mainly by the fibers fractures, as well as delamination and friction between composite tube and the cylinder wall of the cap. Impact tests were performed to study the energy absorption performance. To study the shock absorber effect on the shock-resistance of the helicopter crew seat, a four-degree-of-freedom nonlinear biodynamic model corresponding to 50th-percentile male occupant was developed. The simulation results revealed a good shock-absorber shock-resistance performance.

Zhuravlev S. V., Zechikhin B. S., Ivanov N. S., Nekrasova Y. Y. Analytical technique for magnetic field calculation in active zone of electric motors with superconducting excitation and armature windings. Aerospace MAI Journal, 2018, vol. 25, no 4, pp. 189-201.

Creation of systems with electric motors application is complicated by the restricted potential for conventional electromechanical transducers characteristics upgrading, namely due to their low specific and voluminous power. In this regard, Russian and foreign scientists are working on creating apparatuses based on high-temperature superconductors (HTS), which, as the studies demonstrate, are capable of ensuring higher values of specific power. In particular, developments of superconductor electric motors for prospective all-electric aircraft, sea vessels electric propulsion and wind-driven power plants are renowned. The article presents the problem solution of computing magnetic fields distribution in the active zone and parameters of excitation and armature windings of electric motor based on high temperature superconductors and ferromagnetic yoke of rotor and stator. The above said problem can be reduced to computing of the magnetic field, created by periodical system of current coils, placed between the two cylindrical ferromagnetic areas under the following conditions: the ferromagnetic sections permeability is assumed infinite, and the motor is considered long enough. The current coils systems herewith may be both of various external and internal radiuses, and of equal ones. As a result, a technique based on Poisson and Laplace equations solution relative to vector magnetic potential was developed. The active zone complicated area of a motor herewith was represented as a set of simple homogeneous partial areas according to the harmonic analysis method. The obtained formulas bear general character and can be employed for the computing parameters of motors of various structural schemes. The proposed technique accounts for the superconductor critical parameters, including the transport current dependence on the external magnetic field; the number of pairs of poles, high-order harmonics impact; the number of slots per pole and q phase. This technique can be applied to the ring armature winding. The solution presented in the article allows determine the HTS motor basic parameters such as quiescent E.M.F., inductive resistance of the armature winding phase, power, weight, etc. HTS generator computing was performed using the obtained formulas. The size of the motor active zone was obtained in accordance with the above mentioned computations. Finite elements analysis of magnetic fields distribution was performed to verify the analytical calculations results and their correction. The obtained results testify the high accuracy of the developed technique.

Bazhenov N. G., Filina O. A., Ozerova E. Y. Uniaxial gyrostabilizer application for the gyroscopic stabilization system in self-contained control systems. Aerospace MAI Journal, 2018, vol. 25, no 4, pp. 202.

The proposed unit relates to the field of navigation technology. This type of gyroscopic devices is the simplest and low cost compared to a rotary gyroscope. Single-axis and dual-axis gyroscopes, widely used for movement direction determining, are wellknown, by now, both in civil and military spheres. The main disadvantage of the now-employed gyroscopes is a low speed of the oriented direction determining, reaching up to several minutes. The proposed gyrocompass allows significantly reduce the above said disadvantage and at the same time dramatically improve the accuracy in the “North-South” direction determining. It aims at improving accuracy and speed in the “North-South” orientation determining. This type of gyroscope is much simpler and low cost at commensurable accuracy compared to a rotary gyroscope. The article presents a kinematic scheme of the HS, which allows implement a twin-axis perturbation control method. With this, it allows achieving the required control dynamic characteristics; stabilization accuracy, and the kinematic moment values by selecting the appropriate transfer coefficients Its main property consists in the ability to hold fixed direction of the axis of rotation in space in the absence of the external forces impact on it. This gyroscopic device structure consists of a gyroscope with a rotor of a “brick" shape; communication on angular deviations through amplifiers; torque sensors; and standard stabilizing motors. Thus, the whole complexity of the device consists in the gyro rotor manufacturing, and float chambers in particular. The proposed unit operates in the following way. At the effect of the moment on one of the axes, the rotor gyroscopic moment appears. Pulsating signals along the perpendicular axis appear as well. Thus, there are two types of signals, which can be employed to stabilize the object containing the above said unit is installed. These findings are supported by the presented equations, where  are expressions for the deviations on the two axes. The gyrostabilizer motion was considered in the mathematical model under condition that the values of the moment of inertia on the axes α and β , with account for the presence of additional inserts, would be expressed in the form:  . Thus, the equations were composed with account for these expressions.

The said unit allows find application in the aircraft automatic devices systems.

Grachev N. N. Stability provision of electromechanical transducer characteristics in conditions of flight in the upper atmosphere. Aerospace MAI Journal, 2018, vol. 25, no 4, pp. 207-215.

Scientific results obtained in the work are devoted to the issues of measurement reliability and accuracy of small, slowly changing parameters of low-density gas flows effecting an aircraft in the upper atmosphere. The main trend of the research consists in measuring the aerodynamic forces effecting a flying vehicle in a depleted gaseous environment Measured electric signals coming from primary transducers are often of a low level, sometimes reaching the values less than the noise and interference levels by several orders of magnitude. The most acceptable way of such processes measuring while reaching the specified metrological characteristics is electrochemical modulation (interruption) of the incident gas flow by choppers. It allows obtain the required large gain of the AC signal; easily separate the modulated beam signal from the non-modulated background signal, even if the beam density is lower than this of the background, as well as employ the signal accumulation, which increases the signal-to-noise ratio at the transducer output.

To determine the optimum frequency of electromechanical conversion and ensure the specified accuracy and stability of metrological characteristics, the author proposed employing the method of probabilistic investigation of the stability of the transducer output characteristic. The study of the characteristic's stability is based on the method of probabilistic stability research, ensuring the account for the random character of structural and electro-physical parameters deviations under the impact destabilizing factors and in conditions of mass production. The underlying method of moments allows obtaining the required accuracy with a small amount of computation.

To analyze the stability of the membrane-capacitor transducer transfer characteristic and sel ect the optimum frequency of the flow interruption, programs for computing the nominal values of the frequency response (AFC) and phase-frequency characteristics (PFC) of the transducer, the sensitivity of the response frequency and phase response to the specified parameters were developed. The program for computing the mathematical expectations and dispersion of manufacturing tolerances and temperature coefficients was developed based on the analysis of frequency dependencies of the transducer parameters coefficients effect on its transfer characteristic. Besides, the dependencies of the transducer transfer characteristic fr om temperature at the flow chopper operating frequency, as well as from the temperature change of the transducer membrane were computed.

Based on the measuring transducer transfer function representation in the form of polynomials ratio and analysis of its absolute sensitivity functions to the manufacturing tolerance, structural and electro-physical parameters, as well as destabilizing factors computing dependencies of the transducer transfer function for its stability evaluation and selection of optimal electromechanical conversion frequency were obtained.

Sychev Y. A., Kuznetsov P. A., Zimin R. Y., Soloveva Y. A. Problems of current and voltage high-order harmonics compensation in conditions of distributed generation. Aerospace MAI Journal, 2018, vol. 25, no 4, pp. 216-210.

The main task of this article consists in improving parallel power active filters operation, with account for the networks with distributed generation specifics, and their operation processes modeling.

Micro-power grids, and the features of their control algorithms, have recently gained considerable attention in a wide range of research community. While the potential for increasing the efficiency, reliability and adaptability of the local grid is the most important motivation for their development, micro-energy systems, in turn, may be implemented to meet the growing demand for electric energy in numerous applications. Compared to the high-power energy systems, micro-energy systems can depend on non-inertial generators, such as photovoltaic batteries arrays, which are connected through inverters. Despite the lack of inertia and other micro-power grids properties, which cause certain difficulties in control, micro-energy systems are controlled better through new control laws, such as, those depending on distributed computations, rather than on centralized processors.

Thus, the distortions of the shapes of the current and voltage waveforms introduced by the high-order harmonic components from the nonlinear load can be supplemented by distortions from the sources of distributed generation and the units for their synchronization with the grid. Wind turbines are the most common renewable energy sources. So the parallel power active filters application is considered in the article for the purpose of compensation of the high-order harmonics, generated in association with their operation features. Currently, there are three most common types of wind generators:

1. Induction generators directly connected to the grid. It is an old concept with a mechanical transmission between the turbine and the generator. The generator operates in a rather narrow range of speeds (just above the synchronous one), the gearbox provides a more or less constant speed of the generator at highly differing wind speeds. The generator requires a considerable amount of reactive power, so often a capacitor bank is connected to it.

2. Synchronous generators with permanent magnets. Usually they are delivered in the kit with rectifiers and inverters. The generator is connected directly to the turbine, and it rotates at a low frequency. The generator is being excitated by magnets, and it is not regulated. The rectifier converts the generator voltage/current into DC. At constant voltage there is a capacitor (for smoothing pulsations and as a certain energy buffer). At the DC voltage side the bulk capacitor is present (for tripples smoothing and as an energy buffer). The DC voltage/current is converted thereafter into AC voltage/current by the inverter with 50 Hz frequency and specified characteristics. Everything is determined by the inverter control system. In fact, the rectifier-inverter is an original DC insertion (the like are employed at the borders between the countries, or for energy transmission over large distances).

3. Double fed induction generators. It is just an induction machine with a phase rotor. The stator of the machine is connected directly to the grid, and the rotor is connected via a rectifier-inverter.

It depends, on many respects, on the concrete case, how separate wind generators are combined into wind farms. However, interference is generated in any case, and an effective compensation system is necessary at each stage.

The parallel active filter circuit and developed mathematical model for the conditions of distributed generation and combined power supply were proposed. They allow efficiently compensate for various harmonic interferences in the grids with distributed generation due to the revealed dependence of the efficiency indices of high-order harmonics correction by the parallel active filter on the value of the supply grid internal resistance and the load node parameters.

Tereshkin V. M. Theoretical justification of the possibility of reducing vibrations of electromagnetic origin in a five-phase alternating current machine in comparison with a three-phase machine. Aerospace MAI Journal, 2018, vol. 25, no 4, pp. 229-239.

Multiphase electric machines with an odd number of phases can become an alternative to three-phase machines in areas where a stable rotation speed within one revolution of the shaft is required, as well as other areas requiring highly reliable electric drives with low noise and vibration, for example in special ventilation systems and complexes .

The problem of formation of the resulting current of 3-phase and 5-phase windings, which are fed from the bridge converter, is considered in the work. A comparative analysis of the resulting currents is given from the point of view of the harmonic composition. In solving the problem, a classical approach was used with the use of the vector method.

It was found that the components of the phase currents of the 5-phase symmetrical winding form the 1st and 11th harmonics of the resulting forward current and the 9th harmonic of the reverse sequence.

The third and seventh harmonics of the phase currents do not form a rotating field; their temporal alternation does not coincide with the spatial alternation of phases. The 5th harmonic of the resulting current is absent; in the phase-current spectrum of a 5-phase symmetrical winding, the 5th harmonic component is not contained.

The components of the phase currents of the 3-phase symmetrical winding form the 1st and 7th harmonics of the resulting forward current, as well as the 5th and 11th harmonic of the reverse sequence.

The 3rd and 9th harmonics of the resulting current are absent, because in the phase-current spectrum of the 3-phase symmetrical winding, the 3rd and 9th harmonic components are not contained.

By harmonic composition, the resultant current of the 5-phase winding takes precedence over the resultant current by a 3-phase winding. This allows us to assume that within the period of the first harmonic (fundamental rotation frequency) in a 5-phase winding, the vibrations of electromagnetic origin will be less than for the three phase windings.

Experimental studies of prototypes of 5-phase and 3-phase synchronous machines made using identical magnetic systems have shown that the level of mechanical vibrations of a 5-phase machine is lower than that of a 3-phase machine.

Rebrov S. G., Yanchur S. V., Faustov A. V., Filin S. V. Laminated lithium-ion cells with high specific characteristics. Aerospace MAI Journal, 2018, vol. 25, no 4, pp. 240-248.

Lithium-ion batteries are widely employed for space applications due to a number of advantages compared to other electrochemical systems, especially due to their high specific energy and volume values. Laminated lithium-ion cells possess maximum specific characteristics among lithium-ion cells of other types (cylindrical and prismatic). Besides, they allow use available space more effectively. Despite skeptical attitude towards the laminated Lithium-ion batteries, however, there is information on their successful application in space conditions.

Keldysh Research Center is developing the space oriented Lithium-ion batteries with improved operational and specific characteristics. For example, Keldysh Research Center has already developed and tested the Lithium-ion battery for application in outer space conditions. The objective of the studies being performed is creation of the laminated Lithium-ion battery with improved specific characteristics, large calendar and cyclic life and a possibility to function in conditions of outer space. The task at this stage of the research work consists in searching for cathode and anode compositions allowing achieve maximum specific and volumetric parameters, as well as study electrical characteristics of these batteries under normal climatic conditions.

The studies conducted by Keldysh Research Center with cooperation of the Research Center of Applied Acoustics aiming at searching for cathode and anode composition allowed obtain maximum specific and volumetric parameters for laminated lithium-ion cells (often also called polymer Li-ion by mistake).

Two types of Lithiated oxide of Nickel-Manganese-Cobalt (NMC) and three types of Lithiated oxide of Nickel-Cobalt-Aluminum (NCA) were used as active cathode materials. Six different types of battery-grade carbons were used as active anode materials. N-methylpyrrolidone was used as a solvent for cathode slurries manufacturing, and water was used for anode manufacturing. As the studies revealed, the brand NCA A801-COA was the best choice as an active cathode material and carbon black brand AGP-2A was the best choice as an active anode material. The mass fraction of the active cathode material in the cathode mass was improved up to 93%; the fraction of the active anode material was improved up to 95%. Specific characteristics increase while manufacturing laminated batteries of high capacity. It indicates the batteries manufacturing flexibility while transferring from small-sized laboratory cells to large-sized experimental cells.

The obtained optimal electrodes compositions for Li-ion cells with rated capacity in the range of 1.6-15.3 A·h allow achieving the specific and volumetric values at the level of 230-268 W·h/kg and 520-560 W·h/lrespectively. The following characteristics were demonstrated herewith: for charge-discharge currents of 0.2C-0.2C the cyclability was about 1200 cycles with 97.75% efficiency; for charge-discharge currents of 0.2C-0.1C it was about 980 cycles with 87.8% efficiency, and for charge-discharge currents of 0.5C-1C the cyclability was about 460 cycles with efficiency of 89.1%.

Rygina M. E., Petrikova E. A., Teresov A. D., Ivanov Y. F. Studying the possibility of hypereutectic silumin surface layer structure and properties modification by intense pulsed electron beam. Aerospace MAI Journal, 2018, vol. 25, no 4, pp. 248-256.

Silumins of the hypereutectic composition in the cast state are characterized by a high level of porosity, the presence of large silicon inclusions and intermetallides, which significantly reduces the range of this material application in industry. To eliminate these drawbacks, samples of hypereutectic silumin (Al- (20-22) wt% Si) were irradiated in vacuum with an intense pulsed electron beam in the surface layer melting mode. Irradiation of the surface of the silumin samples was performed by an intense pulsed electron beam (“SOLO” facility, HCEI SB RAS). The irradiation was performed in a residual argon atmosphere at a pressure of 0.02 Pa with the following parameters: 18 keV; 40 J/cm2; 200 µs; 0.3 c-1; 20 pulses. The selected mode, as shown by the results of modeling the temperature field formed in the surface layer of silumin, results in the surface layer melting of the material up to 70 μm thickness. Investigations of the elemental and phase composition, the state of silumin defective substructure in the initial state, and after irradiation with an intense pulsed electron beam were performed using scanning electron microscopy (SEM-515 Philips) and transmission electron microscopy (JEM 2100F), X-ray diffraction (XRD 6000, imaging copper-filtered radiation of Cu-K 1, monochromator CM-3121). The samples microhardness was being determined with the PMT-3 device with an indentor load of 0.1 N. The wear-out parameter and friction coefficient were being identified on a TRIBOtechnic tribometer. The results of the studies performed revealed that the high-speed melting and subsequent high-speed crystallization were led to a nonporous surface layer forming of up to 100 μm thickness with the structure of cellular crystallization free of primary inclusions of silicon and intermetallides.

The size of the cells of high-speed crystallization formed by a solid solution based on aluminum was 0.4-0.6 μm. The cells were separated by interlayers enriched with silicon, copper, nickel and iron atoms. The transverse size of the interlayers was up to 100 nm. It was revealed that the nonporous surface layer formation with a multiphase submicro- nanocrystal structure was accompanied by an increase in the silumin microhardness by 4.5 times, and wear resistance by 1.2 times compared to the cast state.

Parkhaev E. S., Semenchikov N. V. Wings aerodynamic optimization technique for small-sized unmanned aerial vehicles. Aerospace MAI Journal, 2018, vol. 25, no 3, pp. 7-16.

The article suggests combined technique for wings aerodynamic optimization of mini unmanned aerial vehicles (MUAV), which flight modes correspond to critical Reynolds numbers within the range order of 105–106. According to this technique, non-viscous flow-around, flow without separation and aerodynamic characteristics of the finite span wing are being computed in the beginning. The wing planform shape, wing aspect ratio and other geometrics are assumed known and specified. Computation is performed by reliable panel technique. Then the wing profiles shape optimization is performed with account for laminar-turbulent transition, and separation phenomena.

The following assumptions were assumed while wings optimization algorithm developing: the flow-around parameters computation employing 3D analysis model is non-viscous and non-separable. Viscous separated flow-around computing is performed in the contest of 2D-problem of viscous-invicid interaction. Integral aerodynamic characteristics over the wing span are being computed by the technique of lifting line theory using nonlinear section lift data. The suggested technique came from the supposition that aerodynamic characteristics of an isolated wing profile can be extrapolated on the wing. It associates with the fact that the MUAVs wings have, as a rule, a large aspect ratio (AR> 3), and hypothesis of flat sections is applicable for such kind of wings.

The article presents the results of numerical optimization on maximum quality criterion for rectangular wings planform, aspect ratios AR = 5 and AR = 10, at Re = 200 000, as well as arrow-type wing employing the suggested technique.

It was demonstrated that, the moment coefficient constraint allows increase the wing lift-drag ratio, reducing the share of resistance associated with laminar-turbulent transition occurrence and local flow separation formation. At the same time, while optimization in the absence of the moment coefficient constraint each successive quality improvement occurs due to the moment coefficient and wing middle surface curvature increase. The Cya(a) distribution herewith deviates from the initial one.

Khmelnitskii Y. A., Salina M. S., Kataev Y. A. A spacecraft solar batteries panels strength calculation. Aerospace MAI Journal, 2018, vol. 25, no 3, pp. 17-24.

The solar battery panels can be divisible by construction into the following types:

– Solar batteries with panels in the form of a frame with the stretched net-like fabric (net-like fabric panel);

– Solar batteries with panels in the form of a frame with orthogonally stretched strings (a string panel);

– Solar batteries with panels in the form of a frame with the stretched flexible film (a film panel);

– Solar batteries in the form of three-layer panels with a honeycomb core (a panel with a honeycomb core);

– Solar batteries with three-layer panels of integral construction (an integral panel).

The structures analysis of various panels reveals that at present all world firms employ generally three-layer panels with honeycomb core.

The structure of such panel consists of carbon fiber-reinforced plastic encasement and a metal honeycomb core.

Pursuing a goal of developing the rigid and light panel, recommendations on selection of carbon fiber-reinforced plastic, honeycomb core, adhesive film and dielectric film are issued based on experiments.

It allowed create lightweight rigid design structure of a solar panel. It was necessary herewith to perform strength, rigidness calculations and vibrations under effect while transportation and operation.

The stress-strain state of panels, forms and natural frequencies were being defined. Calculations were performed by a finite element method in MSC/Nastran.

CQAD4 sheathing element was selected for encasement and honeycomb cores modelling. The CQAD4 element accounts for all internal forcing factors and the encasement geometry, since it perceives membranous, shear, transversal and flexural loadings.

Calculations reveal that tension, occurring in the elements of the offered light-weight structure, have considerable safety margin, and high rigidity at which the maximal shifts do not exceed 0.05 mm, while oscillation frequencies change in within range of 16-91 Hz. The three-layer panel specific mass herewith is only 1.27 kg/m2. The structure opens possibilities for further improvement.

Smirnov A. V., Egoshin S. F. Energy balance analysis of prospective regional turbo-electric aircraft. Aerospace MAI Journal, 2018, vol. 25, no 3, pp. 25-33.

The presented work deals with studying the possibility and practicality of high-temperature superconductors (HTSC) application while regional aircraft with hybrid electric power plant (the flying weight of up to 50 tons) developing.

The analysis was performed with mathematical model based on collating the power plant required and disposable power while a cruise flight. The basic energy and mass equations, characterizing hybrid power plants of various structures, including the structure with HTSC application, were derived.

It was revealed, that the turbo-electric aircraft is inferior to the aircraft with conventional power plant in the disposable power value. With application of conventional electrical materials, such as copper, this penalty is significant at any flying weight, and such hybrid aircraft developing is inexpedient. With HTSC technologies application this penalty is significantly lower, but it persists at any flying weight.

It can be explained by the presence of additional units in the power plant, which weight is much higher than the weight of the reducing gear, which they replace. The specific mass characteristics of the units based on conventional electric materials are significantly lower, than for HTSC units, which explains the difference in their application effectiveness. The efficiency change of power transfer herewith is insignificant.

At the same time, it was demonstrated in the framework of the model that the trend of the turbo-electric aircraft upgrading was application of installations and units (both gas turbine engines and electric motors) with the most advantageous specific energy-mass characteristics. With this, as it follows from the derived equations, the power plant should include minimum possible number of electric motors based on HTSC technologies.

It was confirmed in the framework of the constructed mathematical model that if the development of superconductor technologies allows develop HTSC-motors with specific characteristics at the level of 20 kW/kg, then the turbo-electric aircraft disposable power would attain the disposable power values of aircraft with classic power plant. It will ensure unconditional possibility for energy effective regional hybrid aircraft creation.

Kargaev M. V., Mironenko L. A. Bending stresses computation in a helicopter unmoored rotor blade blown about by the wind flow. Aerospace MAI Journal, 2018, vol. 25, no 3, pp. 34-43.

Ensuring the main rotor blade strength remains as before one of the main problems, the designer faces while developing a helicopter. Heretofore, at the main rotor blade design in a part of static strength ensuring, the designers confined themselves to its computing under the impact of a force of its own weight. While a helicopter operation thereat, the damages of the main and tail rotor units occur after the storm wind impact.

For the main rotor blades, the following situations are possible: the blade spar bending with residual deformation occurrence down to its destruction; corrugation occurrence at the tail sections; the blade contact with ground or the helicopter tail beam. The above said phenomena prove to be possible due to the small inherent rigidity of the rotor blades, which makes them rather sensitive to the wind loading. The designers should take measures on the wind-flow impact protection ensuring while developing rotary wing aircraft.

According to the 29.675 b item of the AC 29-2C recommendation circular, which gives procedures for determining compliance with the requirements of the AP-29 airworthiness standards, when designing the carrier system, it is necessary to avoid overloading the stops and blades in conditions of wind gusts in the parking lot, or from the rotary wing aircraft's main rotor, taxiing nearby.

The article presents a method for computing flexural stresses in an unmoored blade of a helicopter, blown over by a wind flow. It consists in determining the positions of the elastic axis points of the idealized blade model.

These positions fully determine the shape of deformations and, hence, the magnitude of flexural stresses acting in the blade. The initial equation of the blade bending by a wind loading in a linear setting by the Galerkin method is reduced to an equation relative to an unknown deformation coefficient. This coefficient is determined under the condition of neglecting the additional aerodynamic loadings stipulated by the blade elastic deformations, and with their accounting for. The load increase factor was determined from comparing the obtained relations comparison, on which basis the solution allowing avoid the direct integration of the initial equation was obtained.

The equations are presented in a form convenient for numerical determination of the elastic axis points positions of the blade, slope angles and bending moments (stresses). Computation results for the rotor blades of the Mi-8 helicopter are presented. It was shown, that accounting for elasticity introduces significant changes in the bending moments (stresses) distribution along the angle of the blade azimuthal position, which determines the direction of its blow over.

Lopatin A. A., Gabdullina R. A., Terentev A. A., Eremeeva C. F., Biktagirova A. R. Analysis and characteristics of prospective thermoelectric generators in aircraft electric power supply systems. Aerospace MAI Journal, 2018, vol. 25, no 3, pp. 44-54.

The goal of the presented work consists in developing a technique for computing a part of the aircraft engine casing realized as a thermoelectric generator. The thermoelectric generator (TEG) application onboard an aircraft allows discard the mechanical electric current generator, operating on account of energy extraction from the aircraft engine rotor. At present, a great number of thermoelectric materials, prospective for practical application have been studied. One of prospective trends in this matter is application of housing elements as a basis for TEGs design. The aircraft power plant is undoubtedly the most thermally burdened component. The temperature field along the engine path herewith is characterized by a significant gradient.

Since the thermoelectric gadgets' computing is accompanied by certain difficulties associated with electric and thermal parameters dependencies, the authors developed the technique for computing a housing element, represented in the form of thermoelectric generator of a cylindrical shape. The article presents computation results, performed according to the developed technique, which allow determine and evaluate the value of power output, as well as TEG electric parameters and boundary temperatures of housing walls at the design stage.

The electrical power of the thermo-generator module depends on the flow rate, which cools one side of the housing: a small increase in its speed up to 40 m/s, the power output increases up to 1 kW. It can be seen that under similar conditions with flow rate growth from 50 m/s the power output increases only by 550 W. A similar situation is observed for the case when a TEG is made of of bismuth telluride. Characteristic presented in the article allows determine what engine operation mode would be the most optimal for the TEG effective implementation onboard an aircraft.

To study characteristics and parameters of thermoelectric generator the test bench was employed. The following parameters were measured while the experiments: the resultant current and voltage in thermoelectric modules connected in series (each module is a 64 thermocouples per module, connected in series cased in an insulating ceramic housing), hot and cold junctures temperatures, speed and temperature of the hot and cooling flows.

The paper presents numerical and graphical results of analytical and experimental studies, on which basis the inference can be drawn on the perspective of practical implementation of thermoelectric modules as aircraft engines components. The prospect of TEGs application in high-temperature aircraft and spacecraft power plants is determined by the necessity to obtain powerful enough electric power source onboard with modest weight and size characteristics and high reliability.

Miodushevskii P. V., Legovich Y. S. Development of prospective multipurpose convertiplane. Aerospace MAI Journal, 2018, vol. 25, no 3, pp. 55-63.

Domestic and foreign helicopter building development in the last century opened prospects for convertiplanes application as transportation means for carrying cargoes of considerable weight over the vast territory in conditions of deficiency of the advanced airfield net. Convertiplane is capable of performing vertical take-off and vertical landing. However, convertiplane can ensure significant fuel or electric energy saving while horizontal flight compared to the helicopter of the same weight. Aviation history knows two successful practically realized convertiplane projects. The first project was Canadair CL-84. The second project was the US V-22 Osprey military transport convertiplane. Aero-mechanical schemes of both Canadair CL-84 and V-22 have significant disadvantages. The presented work offers an original convertiplane aero-mechanical scheme, eliminating these disadvantages.

The article lays out the results of studying characteristics of the developed multi-purpose convertiplane, possessing conceptually new aero-mechanical scheme. Various options of the multi-purpose convertiplane, such as ten seats passenger plane, special plane for rescue missions and ambulance, light unmanned convertiplane with high flight duration are considered. Technical characteristics of convertiplane were determined based on the developed technique of preliminary design employing computations of aero-mechanics, aerodynamics, structural strength, weights and centring, as well as comparing the results with the well-known calculation methods.

The results of the studies revealed that among all realization options the offered multi-purpose convertiplane configuration allows achieve higher characteristics, than those of conventional aerial vehicles.The article demonstrates that the existing technical state-of-the-art level allows developing a light multi-purpose convertiplane.

Convertiplane gains its significant advantages through the new turbo-electric power plant, where the last achievements of developing light and powerful electric generators and motors with high power to weight ratio values is employed.

Huang S. ., Kostin V. A., Laptevа E. Y. Application of the sensitivity analysis method for the solution of the inverse creep problem of a wingbox structure on the basis of super-element model. Aerospace MAI Journal, 2018, vol. 25, no 3, pp. 64-72.

The research paper considers the problem of isochronic curves recovery of thin-walled design structures creep referred to deformations measured within the process of full-scale live experiment. It is known that as time passes creep deformations can appear in the construction. They are graphically usually represented as «deformation-time» curves measured by standardized sample testing. However, it has been found out that the deformation curves obtained during testing procedures of the construction differ fr om standard samples due to various reasons: power-based, technological, thermodynamical, etc. The article presents an approach to the corresponding curves construction, based on the processing of the results of the aircraft construction strength experiment.

Setting up the problem for general thin-walled constructions in mathematical terms, we obtain the necessity to optimize the objective functional in the form of the squared residual error of the corresponded theoretical and experimental deformations to the minimum. Working out the solution of the optimization problem is carried out iteratively using the sensitivity matrix, which is the derivative of the deformation function vector along the vector of elastic parameters variables. As the required parameters which control the stress-strained state (SSS) of the structure we choose the secant elastic modulus of the material. To solve a direct problem of the stress-strained state value determination the finite element method (FEM) in the form of a super-element model is used. This makes it possible to reduce the number of diverse required parameters at sufficient accuracy.

Due to the lack of data from the physical experiment, we obtain the numerical deformation values, using the FEM. This is done by solving a direct problem, wh ere measure of inaccuracy typical for strain and load application gauging is introduced. A mathematical calculation has been made for a four-stiffener wingbox operating under the mechanical and temperature load. Figures of the first and second stiffeners show the change of values of the theoretically obtained deformations in case of iterations in the direction of the corresponding experimental values. Isochronic creep curves have been constructed. The application of the sensitivity function has made it possible to purposefully organize the iteration process in the search for elastic parameters and to construct creep curves for the structural elements. The results of the research can be useful for further development of methods of identifying and improving of thin-walled structures according to the testing data, in case of creeping process as well.

Filinov E. P., Avdeev S. V., Krasil'nikov S. A. Correlation-regressive model for small-sized aircraft gas turbine engines mass computation. Aerospace MAI Journal, 2018, vol. 25, no 3, pp. 73-81.

The article suggests several new correlation-regressive models for the aircraft small-sized gas-turbine engines mass computation at the stage of their conceptual design.

A database of the main data and thermodynamical parameters for mass computation of dual-flow turbojet engines was formed. The database includes 92 small-sized turbojets with the thrust less than 50 kN. Equations allowing compute the engine mass at the initial stage of design were derived by correlation-regressive method based on the accumulated statistics.

Model No. 1 uses the total air flow through the engine as an input parameter. The approximating function coefficients were determined based on 88 turbofan engines. The relative standard deviation value for this model was 25.5%.

Model No. 2 uses engine thrust as an input parameter. The approximating function coefficients were determined based on 92 turbofan engines. The the relative root-mean-square deviation value for this model was 18.6%.

The mass model No. 3 uses three input parameters: engine thrust, overall pressure ratio, by-pass ratio. This model involved 77 turbofan engines. The relative root-mean-square deviation value of this model was 13.4%.

The fourth model uses the total air flow, overall pressure ratio, gas temperature in front of the turbine, bypass ratio for calculating the mass.

Statistical coefficients for this model were determined based on 57 turbofan engines. The relative root-mean-square deviation value for this model was 10.1%.

The Kuzmichev mass model depends on five parameters of the gas turbine engine: Mдв = f (m,πкΣ,Gв,T*г, πв) . The total number of engines used in the statistics was 52. The relative root-mean-square deviation value of this model was 13.5%.

Based on the results obtained, we can draw the following conclusions: at the stage of the gas turbine engine conceptual design, the most preferrable models are model No. 4 and Kuzmichev's model. Models No. 1, No. 2 and No. 3, are most preferable for preliminary estimation of the mass of the propulsion system while an aircraft design.

Fokin D. B., Selivanov O. D., Ezrokhi Y. A. The studies on optimal shape forming of a turbo-ramjet engine as a part of a high-speed aircraft power plant. Aerospace MAI Journal, 2018, vol. 25, no 3, pp. 82-96.

Recently, the great attention is payed in many countries to the studies aimed at flight cruising speed increase of aircraft of various purposes. The projects aimed at considering the issues creating both passenger (Aerion AS2, QueSST, Sky-lon) and military (SR-72) high-speed aircraft are in full swing abroad.

The similar studies on building-up the flight speed of military planes are carried out in Russia too. Thereupon, the possibility of developing prospective Russian fighter-interceptor based on MiG-31 aircraft, which speed should substantially increased, presents undoubted interest. The same applies to an attack high-speed aircraft of the type of the Soviet T-4 scout bomber, or the US XB-70 “Valkyrie” strategic bomber with maximum flight speed, corresponding Mach number no less than M = 3.

The article presents the results of the study on the power plant optimal shape based on the turbo-ramjet engine with tandem configuration of the high-speed aircraft contours with cruising speed of Mcr = 4.

To solve the stated problem, the software complex consisting of mathematical models of the combined engine, including gas-turbine and direct-flow circuits, supersonic air intake and a full-range jet nozzle, as well as the technique for the aircraft performance characteristic computing. The developed program complex allowed evaluate the efficiency of such combined power plant application as a part of an aircraft with increased cruise speed.

The presented results demonstrated with high obviousness that the effort aimed at the power plant optimal shape formation is most expedient to perform in accordance to the procedure of optimization studies performing, which includes the task setting, the initial data preparation, parametric studies, post-optimization analysis and issuing recommendations.

Parametric optimization with seven parameters and three criteria with goal functions of subsonic and supersonic flight ranges at the optimal altitude, as well as required length of runway for the aborted-continued takeoff, was performed employing the above said approach. The optimization results revealed that the possibility of improving an high-speed aircraft performance relative to the conditionally preliminary basic variant.

Three aircraft options with the highest attractiveness level were selected out of the obtained twenty Pareto-optimal options by the “fuzzy sets” tool. Further final selection of the most expedient one out of these options always up to the development engineer and associated with taking a number of trade-off decisions.

Orlov M. Y., Anisimov V. M., Kolomzarov O. V. Design refinement of combustion chamber of gas turbine engine with toroid recirculation zone. Aerospace MAI Journal, 2018, vol. 25, no 3, pp. 97-106.

The design of a serial auxiliary power unit was employed as a prototype while developing a new engine. The schematic solution of inlet unit and centrifugal compressor was preserved in the new design, while the engine turbine underwent changes right from the start, since it became radial instead of axial. It required the changes of the combustion chamber design. After studying a number of possible schemes, a decision was made to choose the straight-flow combustion chamber of a ring-type of, which had substantial reserves for minimization on size with relative simplicity of its technological design. The specific feature of this particular combustion chamber is its diagonal positioning relative to the engine axis. A number of problems associated with the lack of experimental and calculation data arise while organizing a working process in the combustion chamber of this type.

The goal of the study is design refinement of the considered combustion chamber structure to optimize the workflow of the annular combustion chamber with the offset zone of a toroid type.

At the first stage, the design refinement of the flame tube structure was performed to organize a vortex structure in the primary zone by changing diameters and a number of clamping apertures and addition of a «springboard» of the internal rim of the flame tube. At the second stage the design refinement of the seat of flame in the primary combustion zone was performed. The atomizer was substituted by the spray injector, and vane swirlers were added to the duct between the deflector and the flame tube wall. The third stage was devoted to the necessary temperature field forming at the combustion chamber outlet. For this purpose the works shaping-up the necessary jets penetration depth, the number and location of shift apertures were performed.

The outcome of the activities consists in obtaining acceptable combustion chamber design of the engine being developed, in which the authors succeeded achieving the flame stabilization in the primary combustion zone, temperature field distribution inside the chamber, excluding its burn-through, and temperature filed irregularity reduction at the outlet.

Finogenov S. L., Kolomentsev A. I. Solar thermal rocket engine with beryllium-oxide phase-transition latent heat energy storage and hydrogen afterburning. Aerospace MAI Journal, 2018, vol. 25, no 3, pp. 107-115.

The article considers solar thermal propulsion (STP) with thermal energy storage (TES) containing high-temperature phase-transition material – beryllium oxide possessing high latent heat of phase transition “fusion-crystallization”. High melting temperature allows obtain the engine specific impulse at a level of 9000 m/s.

Joint optimization of the basic relevant parameters, such as masses ratio of solar mirror concentrator and TES in combination with mirror accuracy parameter was performed. It was demonstrated that that the ratio of TES energy capacity to solar radiation receiver thermal power, or ratio TES energy capacity to solar concentrator area in conjunction with optimal selection of accuracy parameter of the mirror can be accepted as an optimizing parameter. Maximum mass of a spacecraft being placed into geostationary orbit with time limitation of inter-orbital transfer from 30 to 90 days was selected as optimization criterion. Optimization was performed out by Gauss-Seidel method.

The optimization results revealed that optimal ratio of TES energy capacity and light detector power was 22-24 MJ/kW, which corresponds to the optimal ratio TES energy capacity to the concentrator area of 6-7 MJ/m2 at rational mirror accuracy parameter of 0.25 degrees. The STP characteristics with TES are presented and analyzed. The article shows that for relatively small flight time of 3040 days optimal values of excess oxidant ratio corresponding to payload mass maximum. The higher value of excess oxidant ratio corresponds herewith to the lower value of the flight time.

Dependences of the TES energy capacity and the concentrator diameter from excess oxidant ratio for a wide interval of flight duration are presented. Expedient areas of heated hydrogen afterburning application for various inter orbital flight duration were determined. The article shows that afterburning is expedient for the time of putting to geostationary orbit of 30 to 45 days. The corresponding excess oxidant ratio changes herewith from 0.3 to 0.1. For the flight above 50 days, the monopropellant hydrogen STP is expedient. Compared to alternative inter-orbit transportation means, employing the combination of small and large thrust engines combination, the gain is about 450 kg under the one and the same inter-orbital transportation time of 60 days.

Remchukov S. S., Danilov M. A., Chistov K. A. Computer aided design and computing of a plate-type heat exchanger for small-size gas turbine engine. Aerospace MAI Journal, 2018, vol. 25, no 3, pp. 116-123.

The article presents computational complex allowing perform computer aided design and calculation of a compact heat exchanger for a small-size gas turbine engine.

The coomputational complex includes a number of blocks based on open commercial programs. The blocks are united by the common software algorithm, developed at Central Institute of Aviation Motors (CIAM).

The input data is changed at each iteration to obtain the required parameters, namely, the regeneration degree and hydraulic resistance.

Computer aided design and calculation include the steps of the initial data entering into the parametric model, checking compliance with the restrictions, automatic model building, meshed models preparation, working medium flowing calculation and computational results output. The initial data is set with account for limitations, such as overall size restrictions and material outlet depth. The possibility of obtaining better thermohydraulic characteristics depending on the model geometry should be accounted for as well.

Automatic building of models is performed according to the set parameters.

At the next stage, the built models are loaded to the ICEM CFD program, and meshes building is performed.

The obtained grid models are used for calculation in Ansys CFX software. Full pressures and temperatures of air and gas at the inlet, as well as the flow rate of gas and air at the outlet are set as boundary conditions. The employed turbulence model is Shear Stress Transport model.

After calculation termination, the resulting file, containing all significant exchanger computational parameters, is formed in the form of a table.

In case of the obtained parameters discrepancy with the claimed requirements, the parameters correction is performed with subsequent repetition of the considered algorithm.

Automation of the design and computing algorithm allows employing it together with CAD complexes for multi-criteria optimization.

The developed computing complex allows obtaining the optimal heat exchanger configuration for a specific task within the specified limits. The calculating complex was being employed in CIAM for the heat exchanger envelope updating, which led to the regeneration degree increase from 62% to 76%, when total hydraulic losses decreased to 1,27% with requirements and restrictions compliance. The genetic algorithm was used as an optimization method.

Gulienko A. I., Schurovskiy Y. M. Experimental study of gte lubricating system oil-air mixture properties. Aerospace MAI Journal, 2018, vol. 25, no 3, pp. 124-133.

Lubricating system of the gas turbine engine rotor supports in many ways determines its operation reliability. The experience in developing aircraft GTEs lubricating systems accumulated by many manufacturers is associated with predominant employing of empirical and experimental techniques, and is not practically covered in literature. As a rule, the lubrication systems characteristics are being determined while operating with pure oil, and their recalculation onto the two-phase mixture may lead to overpower of the driving pumps motors and, as consequence, to their weight increase.

The article presents the analysis of the properties of two-phase mixtures pumped through the unit of the aircraft GTE lubrication system based on experimental data. This data was obtained at CIAM with the test bench for semi-natural simulation of lubrication system with oil chamber imitation, and installation for fine-dispersed oil-air mixture forming, where the mixture is formed by the air entrainment effect.

Using the results of oil-air mixture flow visualization, the article shows that in the area of the GTE lubrication systems operating modes the mixture may be considered as one-component homogenous media, possessing the properties of elastic continuum with homogenous sound velocity.

While air entering the exhausting duct the two-component flow of oil-air mixture and air-oil bubbles, clogging the pipe cross section and move relative to the oil-air mixture at low speed is formed.

Characteristics of a discharge gear pump, pumping the oil-air mixture, are affected not only by air and oil properties, but also by the structure formed by the pump throughput capacity.

It has been shown that in GTE lubrication systems a mode of emptying the exhaust gear pump inlet branch may occur with the possible realization of the stratified flow structure, as well as a dynamic locking mode in which a pulsating flow is formed with density waves forming and a polyharmonic fluctuations excitation in the system. Based on the experimental data, the air-oil mixture flow modes map was compiled.

The paper presents the relationships by which give possibility to calculate the thermo-physical properties of the two-phase mixture pumped in the tracts of the GTE lubrication systems. This approach showed good agreement of calculations with experiments in the lubrication system static and transient operation modes.

Semenova A. S., Gogaev G. P. Evaluation of destructive rotation frequency of turbo-machine disks applying deformation criterion with LS-DYNA software. Aerospace MAI Journal, 2018, vol. 25, no 3, pp. 134-142.

Turbine disk is the main part of the aircraft engine, since its failure may lead to either emergency or catastrophic situation. According to NTD the GTE load-bearing capacity is being evaluated by the destructive rotation frequency margin, applying the limit equilibrium theory, at destruction along meridian section fr om tangential stress, and at destruction along some cylindrical partially meridian section fr om tangential stress.

Factors affecting the disk load-bearing capacity are the meridian section shape, scheme of destruction (along meridian, cylindrical or mixed sections), the presence of stress concentrators, and the material properties. Allowance for these factors effect on the disk load-bearing capacity while applying the lim it equilibrium theory is not practical.

Destruction of most metals is the result of damages accumulation. Two main mechanism of damages such as voluminous damage (pores growth and merge) and shear damage (cracks growth and merge) are discerned. A model of damages accumulation based on shear damage, i.e. destruction criterion on maximum accumulated plastic deformation, can be employed for numerical determination of the destructive rotation frequency of the turbo-machine disks from nickel alloys.

The plastic flow theory can be employed to determine the disk lim it rotation frequency. A modified version of the classical flow theory with isotropic hardening makes allows implement an arbitrary stress-strain dependence given in the form of strain diagrams.

Several series of calculated overspeed test were performed. The effect of the following factors on the calculated destructive frequency was being studied:

– loading speed;

– the finite elements mesh size.

The computational studies results revealed that the finite element size and mesh computing time did not practically affect the convergence of computation and experiment.

The computational studies results revealed that the finite element size and mesh computing time did not practically affect the convergence of computation and experiment. However, the smaller the grid, the more accurately the cracks development on the disk can be traced.

The obtained computation results were validated based on the results of the overspeed test performed with the low-pressure turbine disk of AL41F-1C engine at the Central Institute of Aviation Motors (CIAM) stand.

Kolychev A. V., Kernozhitsky V. A., Levikhin A. A. Cooling system of gas turbine engine turbine blades made of heat-resisting alloys and conductive ceramics. Aerospace MAI Journal, 2018, vol. 25, no 3, pp. 143-150.

The article deals with thermionic cooling system (TCS) of turbine blades (TB) and other hot elements (HE) of aircraft gas turbine engines (GTD), which consists in coating them with a layer (thermionic- protecting layer (TPL)) from heat-proof and heat-resisting material, but with a low electronic work function (EWF). When the TBs and HEs heating, electrons start leaving their surface, taking with them 2-10 MW/m2 of thermal energy in exponential-like temperature dependence. It will allow increase significantly the GTE efficiency due to the temperature increase of the working gas prior to the turbine and extra thermionic transformation, as well as increase the GTE reliability and lifespan.

The thermionic cooling technique under development can be employed in aircraft building while creating power gas turbine installations-converters for the spacecraft of increased power capacity and prolonged active lifespan. It can be implemented also while developing commercial systems of putting a payload and tourists into orbit, including a spacecraft based on the reusable first stage of an aircraft type with GTE, or transport aircraft with thermionic GTE. Besides, the technology under development will be called-up for the fuel-and-power sector and shipbuilding while power plants developing, and in oil and gas sector for gas pumping units developing etc.

The TCS realization will allow increase the temperature of the working gas prior to the turbine without increasing the quantity of the air tapped off the compressor, or increase the resource of the most thermally stressed elements of the gas turbine parts, the efficiency increase, thermal stresses reduction in blades due to the thermionic sensitivity to the temperature. It will ensure continuous diagnostics of the turbine state and other high-temperature elements in real-time mode based on electrical engineering parameters, depending on the number of thermo-emission electrons perceived by the anode, and modernize gas turbine installations and GTEs produced in Russia with their resource enhancing due to the extra cooling and without their serious reconstruction.

Donskov A. V., Mishurova N. V., Solov'ev S. V. Automated system for space vehicle status monitoring. Aerospace MAI Journal, 2018, vol. 25, no 3, pp. 151-160.

The article considers the issue of a space vehicles status monitoring automated systems developing.

The goal of the work consists in analysis and application of the tools for flight control systems of conventional spacecraft improving.

The technique of a manned spacecraft current state status monitoring with account for affecting factors (an aircraft orbital movement parameters, structural specifics of an aircraft and ground-based control loop), as well as a throughput of radio communication circuit for telemetric information transmitting, sensor equipment capabilities, onboard measuring instrumentation and computing means were studied.

The conclusion was drawn, that the tasks of controlling processes automation while spacecraft flight control are not exhaustive.

Depending on the designation of an individual spacecraft or spacecraft orbital group not only the tasks and their aggregate set can change, but specific and independent assignments may arise as well.

With account for the current flight control practice in manned astronautics the approach at large to a space vehicles' on-board equipment status monitoring automated systems developing was formed. Automation of problems solving on telemetric information displaying and analysing coupled with information support of a specialist of the group of analysis allows increase the quality level of the managing group functioning. It is achieved through detecting an abnormal situation, potentially translating into emergencies, as well as operational provision of flight control operators with information over a wide range of the problems being solved.

The significance of the spacecraft status monitoring automated systems developing is being proved by the fact that it allows minimize the human factor in the process of a spacecraft control, increase information accessibility and ease-off the burden of analysis group specialist while performing routine operations.

The considered approach to the spacecraft status monitoring automated systems developing can be applied to both the process of of existing manned space vehicles flight control process improvement, and prospective manned spacecraft under development.

Danilenko N. V., Kirenchev A. G. Work process of the earth environments vortex formation. Aerospace MAI Journal, 2018, vol. 25, no 3, pp. 161-170.

The present-day science state-of-the-art allowed ensuring qualitative transfer to many branches of not only scientific but to technologic and other types of human activities. New knowledge aroused at the junction of the well-known scientific and technological trends. Though in certain trends of modern science development some so-called “blind-spots” still exist. The theory of vortex formation is an example of such modern science state-of-the-art. Currently the specialists of this scientific trend cannot establish the physical entity of atmospheric gas dynamic specifics of the vortices under the aircraft air intakes, well as the gist of their work process. The closer analogue of such vortices are the atmospheric whirlwinds, which working processes are associated with the Earth daily rotation. However, the capabilities of modern science do not allow establish the work process of the above said problem gas dynamic phenomena. The scientists in the USA and many other countries declare openly that the do not understand tornado – a small-sized vortex of a cyclonic type. In such circumstances, the scientists are compelled to give definitions to whirlwinds, tornadoes and cyclones by the facts of their physical manifestations in the field of visual perception. Such definitions do not contain the boundary conditions, work process elements, and limit their experimental modeling possibilities. The scientists face a great problem of exploring the work process of the Earth environments vortex forming. One of the main tasks of the Earth environments vortex forming research and its product is establishing the vortex characteristics, their corollary and application areas.

The article discloses the work process of the Earth environments vortex formation. It gives the definition of vortex formation, and specifies the product of vortex formation, including vortex field, tornadoes and air intakes vortices. The work process of vortex formation was established. The article presents the Earth vortex filed characteristics and their corollary.

The Coriolis force role in the process of vortex formation of natural and man-made vortexes was revealed. The results of experimental modeling of vortex formation under the air intake with account for the Coriolis force action are presented.

Razoumny Y. N., Samusenko O. E., Nguyen N. Q. Optimal options analysis of two-tier satellite systems for near-earth space spherical layer continuous coverage. Aerospace MAI Journal, 2018, vol. 25, no 3, pp. 171-181.

Nowadays, there is a wide nomenclature of practically new significant tasks of monitoring vast near-Earth space areas by space systems, associated with the space debris problems, spacecraft technical maintenance in orbit etc. All tasks of such kind in an abstract formulation can be interpreted in the form of mathematical problem on optimization of the satellite constellations orbital construction for continuous coverage of specified spherical layers of near-Earth space. However, still there is no theoretical apparatus for effectively solving this problem.

The article formulates for the first time the optimization problems of the two-tier satellite constellations orbital construction for near-Earth spherical layer continuous coverage by the criterion of the characteristic velocity minimum total costs on the system creation. Each tier of such a system is formed in circular orbits with the same altitude and inclination values for all satellites. The satellites of each tier are oriented herewith in such a way that observation cone, formed by the onboard equipment of the satellites in the upper tier are directed downward towards the Earth, while in the upper tier – towards the opposite side.

Decomposition of this problem and its reduction to the traditional problem of selection in the delta-systems class of one-tier orbital constellations and their optimization by the total characteristic velocity minimum was performed in this work. The authors suggest methodological approach to this problem solving; discuss the obtained numerical results and the suggestion on application of the obtained optimal options of the two-tier satellite systems for solving various practical tasks. The two-tier orbital structure in many cases has no advantage over the traditional, single-tiered option. However, under certain conditions the two-tier orbital construction appears after all more preferential.

Kovalev A. A., Tischenko L. A., Shakhovtsev M. M., Gorbatovskaya T. A., Vlasov E. Y. The study of silicon substrates pre-treatment technological parameters effect on their surfaces contact angle. Aerospace MAI Journal, 2018, vol. 25, no 3, pp. 182-189.

The presented article deals with the study of technological parameters (temperature and processing time in hexamethyldisilazane (or HMDS) vapours and dehidration) effect on contact angle of silicon substrates pre-processing, including oxidized ones, to evaluate their hydrophobicity. Contact angle measurements were being performed by method suggested by Bickerman. Those angles values were being obtained indirectly due to the known volume and diameter of a water drop. For non-oxidized silicon substrates technological parameters effect on contact angle consists in the following: the 1.5 degrees increase with 30 seconds increase in time of processig by HMDS vapours, 1.2 degrees increase with 120 seconds increase of dehydration time, 0.6 degrees decrease with 45 degrees increase of processing temperature. For oxidized silicon substrates technological parameters effect on contact angle consists in the following: the 2 degrees increase with 30 seconds increase of processing by HMDS vapours, 0,2 degrees increase with 120 seconds increase of dehydration time, 1 degree decrease with 45 degrees increase of processing temperature. Experimental data analysis was performed by Yates analysis, i.e. full fraction analysis. Based on the obtained results the inference was drawn that increasing time of substrates processing in HMDS takes the strongest effect on their contact angle change. Besides, on substrates temperature increase the contact angle decreases irrespectively to the oxide film presence or absence on their surface. The latter, probably, is associated with the fact that hexamethyldisilazane evaporates from the substrate surface, since their maximum heating temperature was close to the HMDS boiling temperature while this study.

Grachev N. N. Quality evaluation of aircraft electronic instrumentation assembling based on registration and analysis of mechanical joints electromagnetic emission. Aerospace MAI Journal, 2018, vol. 25, no 3, pp. 190-202.

The scientific significance of the results outlined in the work consists in studying mechanisms of contact radio interference originating, occurring due to the magnetic field effect of radio transmitters, located onboard an aircraft, on current-conducting mechanical contacts of the structures with non-linear variable resistance, as well as electromagnetic interference generation. The scientific results obtained from the studies demonstrated mechanisms of forming the current, induced by EMF, passing through the aircraft structural elements, which leads to formation of a mechanism of secondary electromagnetic field radiation, interacting with the primary irradiating magnetic field of the radio transmitter. There is a possibility to control the structural elements assembly quality by registering and analyzing the spectral composition of the electromagnetic radiation of the mechanically connected structural elements. Performing complex diagnostics, based on measuring the spectral content of the whole product, and placing antennae around the product under study allows performing reliable estimations of the assembly quality of both separate mechanic components and the entire structure.

The studies performed in this work can be applied to the development and study of a contactless express-method for assessing the structures assembly and erection quality. This method is based on the registration and analysis of artificially generated contact interference under the impact of mechanical vibrations and a high-frequency harmonic electrical signal on the aircraft structures' elements, forming phase-amplitude-modulated oscillation circuits, which can be recorded by either spectrum analyzer or a FAM receiver, or AM oscillations. With this, the levels of their spectral components are measured at a change of mechanical impacts frequency in the range determined by the operating conditions. The measured level of the spectral components of the emitted amplitude-modulated oscillations is compared with the level of the spectral components of the signal emitted by the reference block with given mechanical parameters and normalized level of contact interference.

The main result of these studies allows fruitfully employ the contact interference formation, considered as undesirable phenomenon in the field of electromagnetic compatibility, for estimating the mechanical qualities of the structures (their assembly quality) of various aircraft equipment and units, including assembly and erection quality (especially associated with fixture elements tightening force).

Balkovoy N. N. Analysis of application specifics of a reaction wheel with intrinsic disturbing moments compensation. Aerospace MAI Journal, 2018, vol. 25, no 3, pp. 203-211.

Inertial electromechanical actuators in the form of reaction wheels (RW) found widespread occurrence as actuators of spacecraft attitude control system. The RWs task consists in forming a dynamic controlling moment proportionally to a control signal. This task is reduced to the RW acceleration control.

The article suggests the RW classification, describes advantages and disadvantages peculiar to each of the types. Amid all varieties of the units, reactions wheels with intrinsic disturbing moments' compensation (RWMC) are outlined as one of the most prospective types. The presented work is devoted to the study of these units application possibilities and comparing them with classical ones, where control is performed only by the electromagnetic moment.

To study dynamic and accuracy characteristics of a spacecraft equipped with the RWMC under study, its mathematical model was developed. Analysis of the RWMC dynamic moment development transient was performed. It revealed that transfer functions of compensating and basic (electromagnetic moment control) loops may be represented with high accuracy by the aperiodic link. The time constants of these links were also obtained while the RWMC experimental testing.

The model of the controlled rotational motion accounts for the RWMC static and dynamic imbalances values, as well as the number of RWMC nonlinearities, such as saturation, associated with attaining the limiting angular speed by the rotor and the dead zone while the rotor passes the zero angular speed (for a model without disturbances compensation) etc. Modeling of the spacecraft control system operation in stabilization mode in conditions of ideal measuring of angular position and angular speed was performed to study the effect of the unit specifics on the control system operation.

The spacecraft attitude control system with RWCM was compared to classical RW. In both cases, the control system loop was closed by the PID-regulator, since external disturbances, affecting the stabilization static error value, impact the spacecraft together with disturbing moments.

The simulation results showed that RWCMs has higher accuracy and dynamic characteristics compared to the classical RW. This type of units appears more preferable for developing precise spacecraft attitude control systems, since it allows reduce the “dead zone” of control, as well as oscillation in stabilization transient, especially in the area of near-zero angular rotation speeds of the RW rotor.

Tereshkin V. M. Determining resultant current harmonic composition of an electric motor symmetric four-phase winding. Aerospace MAI Journal, 2018, vol. 25, no 3, pp. 212-219.

Modern power electronics and microprocessor technology state-of-the-art allows develop DC-AC converter with any number of phases in a wide power range.

Realization of a multiphase motor (m > 3) based on the magnetic system of a 3-phase motor is also practically a feasible task with certain modernization of the winding scheme.

As an illustration the article presents a schematic diagram of the four-phase winding, its vector representation, as well as four-phase converter control algorithm while vector pulse-width modulation realization.

The electric drive based on a multiphase motor may display certain advantages in compared to the traditional electric drive based on a three-phase motor and find application wherein the higher requirements are placed on vibrations. The cause of vibrations of electromagnetic origin may be the high-order harmonics of the resulting current, which creates an m.m.f. in the air gap.

Preliminary studies revealed that symmetrical 4-phase winding had the worst figures of the spectral composition of m.m.f., compared to the 5- and 7-phase windings. However, the traction electric drive of the “Granit” electric locomotive was just realized based on the 4-phase asynchronous motor. That is the electric drive based on multiphase motor is already an alternative to the electric drive based on the three-phase motor. It imposes the necessity for comprehensive comparative analysis of multiphase windings and control algorithms for converters to which multi-phase windings are being connected.

The article considers an approach based on classical vector method. With its application harmonic analysis of a resultant current of the symmetrical 4-phase winding. The analysis revealed the phase currents' 1, 5, and 9 harmonics formed the resulting currents of positive-sequence, and the phase currents' 3, 7 and 11 harmonics formed the resulting currents of the negative sequence. Accounting for the fact, that the 1, 3 and 5 harmonics are commensurable in magnitude, significant electromagnetic ripples are theoretically possible within the first harmonic period.

The approach based on the classical vector method considered in the paper can be used to analyze the harmonic composition of the resulting current of multiphase windings with any number of phases. This makes the approach universal for the comparative analysis of multiphase windings on the harmonic composition of the resulting current.

Antipov V. V., Nochovnaya N. A., Kochetkov A. S., Panin P. V., Dzunovich D. A. Effect of casting parameters on shaped castings quality of a new high-temperature TiAl based alloy. Aerospace MAI Journal, 2018, vol. 25, no 3, pp. 220-228.

The results of original research on cast structure and properties of a new high-temperature intermetallic gamma alloy Ti-45.5Al-2V-1Nb-1.5Zr(Cr)-Gd-B developed in VIAM [patent RU 2606368] have been discussed. A solidification temperature interval has been determined for the new alloy: solidus temperature 1471°C, liquidus temperature of 1528°C. The pouring gate system has been designed with the help of ProCast software taking into account centrifugal casting technique which provides both full mould filling with molten metal and absence of metallurgical defects in low pressure turbine blade castings. The research was focused on the effect of temperature and duration parameters of centrifugal casting on macro- and microstructures of shaped castings obtained in induction skull ALD Leicomelt 5 furnace. The X-ray spectral microanalysis has revealed that the samples matrix consists of alternating γ -TiAl and α2-Ti3Al lamellae; there are areas with lower aluminum content and higher content of vanadium and zirconium/chromium; also excess phases enriched with gadolinium and oxygen have been found (complex gadolinium oxides). Microstructure analysis after hot isostatic pressing has shown that plate-like morphology of structure doesn't change: alternating lamellae of γ and α2 phases are gathered into colonies within prior β(α)  grains with small amounts of β phase along grain boundaries (the plates possess similar geometrical orientation within each lamellae colony). It has been shown that structure homogeneity of castings strongly depends on pre-heating temperature of casting moulds. As the experiment has revealed the optimal pre-heating temperature of casting moulds for the new alloy falls in the interval 750850°C. The research results have given the opportunity to develop casting and heat treatment processes which allowed to obtain defect-free shaped castings of turbine blades for aviation jet engines.

Lapaev A. V., Ryashin N. S., Fomin V. M., Shikalov V. S. Properties of aluminum coatings of cold gas-dynamic spraying at corrosion damage zones of 1163RDTV alloy products. Aerospace MAI Journal, 2018, vol. 25, no 3, pp. 229-239.

Cold gas-dynamic spraying is a method for coating process, creation of 3D objects and new materials from powder metals, alloys, composites and powder mixtures. The method was developed based on cognominal physical phenomenon, discovered at the Institute of Theoretical and Applied Mechanics named after S.A. Khristianovich of Siberian branch of the Russian academy of sciences in the early 1980s. Nozzle assembly and a heater are fixed as a part of the cold gas-dynamic spraying test bench based on the industrial robot KR 16-2 in dust-noise proof chamber. While spraying the powder particles are accelerated by the gas flow to the velocities of 400-1200 m/s and form the coating without melting. In a number of works of domestic and foreign researchers the possibility of metallic objects recovery by this method is demonstrated, whereby the study of coatings and materials obtained by this method presents an undoubted scientific and practical interest.

The presented article studies the properties of aluminum coatings formed by the cold gas-dynamic spraying method at corrosion damage zones of the substrates from 1163RDTV structural alloy.

At the first stage of work corrosion damages in the form of surface corrosion of the plates from the 1163RDT alloy were simulated. Then they were recovered by the cold gas-dynamic spraying coatings from ASD-1 aluminum powder. The average measured size of the ASD-1 powder particles was 27 mcm.

Experimental dependencies of porosity and micro-hardness of these coatings and oxygen content in them from deceleration temperature while spraying were obtained. These dependencies allowed sel ect the better coating process mode for performance characteristics recovery of structural elements with corrosion damage.

During the experiments of the second stage the samples recovered by the cold gas-dynamic spraying coatings from the 1163RDTV alloy were tested on tensile strength while static loading. Experimental deformation and fatigue endurance curves were obtained. Due to the low porosity and micro-hardness of the cold gas-dynamic spraying coatings, applied at T0 = 200°C, the samples with corrosion zones recovered by these coatings were selected for static and fatigue stretching tests. The obtained experimental results analysis revealed that with the considered coating process mode the full static hardness characteristics recovery did not occur. Nonetheless, an A1 recovery by the cold gas-dynamic spraying coating from 1163RDTV alloy increases the sample static hardness characteristics in the elastic region of the deformation curve. The fatigue tests revealed the effect of the stress concentrator on fatigue strength, which should be accounted while cold gas-dynamic spraying application for recovering corroded structural elements.

At the final stage of the work, a coating fr om ASD-1 was formed on the TU-154 stringer fragment (an alloy of B95 series). It demonstrates the ability of applying these coatings on the fuselage frame elements.

The results of the presented work demonstrate the high potential of the cold gas-dynamic spraying method in solving the problems of aircraft construction elements recovery and repair.

Pashko A. D., Belichuk A. A. Development of anti guided missiles active protection system for aircraft and assessment of its application prospects. Aerospace MAI Journal, 2018, vol. 25, no 2, pp. 25-36.

At present, designers of almost all foreign countries (USA, UK, France, Germany, Israel, Japan, China and South Africa) decide upon thermal imaging tracking coordinator, employing matrix photo-detecting unit while the type of homing head selection for hew types on missiles. Its modern element base is intrinsically the basis of the fifth generation infrared homing heads. The main advantages of guided missiles of “air-to-air” class equipped with homing head containing matrix photo-detecting unit consist in the presence of significant field of vision, ensuring target patter recognition and its identification, capability of automatic aiming employing and high jamming immunity. All this requires aircraft protection means modernization.

Modern aircraft are being equipped with on-board defense systems, designed to protect an aircraft of various classes and purposes from hitting by aircraft rockets, antiaircraft rocket systems through detecting hazard occurrence and counteracting the attacking means. Onboard defense system “President-S”, “Talisman”, electronic countermeasures equipment of Su-30MKI and aircraft protection system “MANTA” are most up-to-date systems.

The results of performed analysis of modern aviation guided missiles and means of protection from high-accuracy weapons allow conclude that the existing onboard defense systems do not ensure enough level of protection. Namely, they ensure only a passive protection by creating interference action on missiles homing heads, which is inefficient with account for digital signal processing and jamming protection of the guided missiles. Modern heat flares are effective only for protection from the missiles' with single-element photo-detecting unit. Due to target image detection capabilities of modern homing heads with matrix photo-detecting units, the heat flares application is inappropriate. From all the above said, a topical problem of upgrading the onboard defense systems by developing new ways of an aircraft protection from guided missiles follows.

Improving the aircraft protection is possible by active impact on guided missile by protective ammunition included into active protection system, leading to its hitting, self-destruction or mishit.

The goal of the study is enhancing the aircraft protection from guided missiles of “air-to-air” type.

Thus, the developed active protection system is capable of ensuring in automatic mode all aspect detection and tracking of a guided missile, its destruction at a safe distance from the aircraft, in close interaction with the other aircraft systems.

Sinitsin A. P., Goza D. A., Rumyantsev А. V. Thermal calculations of liquid low thruster on pollution-safe fuel. Aerospace MAI Journal, 2018, vol. 25, no 2, pp. 109-116.

The article presents the results of development and application of the thermal model of a stationary liquid thruster on alternative mono-fuel. It allows calculate the thermal field, and determine internal and external conductive and radiation thermal fluxes, temperature variation gradients speeds in stationary and dynamic modes operating modes of the engine, and calculate heat emission in combustion chamber with subsequent recommendations on upgrading the engine thermal scheme and its reliability.

The purpose of the above said thermal calculation consisted in determining the thermal state parameters and characteristics of the low thrust rocket engine on alternative fuel. The thermal calculations using mathematical model developed and presented in this document solved the following problems: developing the engine thermal model, its verification by the thermal test results, calculation substantiation of the solutions,directed to temperature reduction of propellant delivery valve and capillary delivery tube.

The three-dimensional engine thermal model was built with SolidWorks Flow Simulation 2014, which employs the finite volume method ( a numerical method for integrating the systems of partial differential equations. In heat calculations, the boundary conditions were set identical to the conditions for thermal vacuum tests, which imitated the outer space in full-scale operating conditions.

The experimental data of the engine thermo-vacuum tests, obtained with the development design office Fakel test-bench, were used for the calculation thermal model verification. Verification of the thermalmodel consisted in heaters power selection from the condition of compliance of temperatures in the controlled points and measured ones.

Recommendations on thermal scheme optimization and constructional materials selection were developed according to the thermal calculation results.

Recommendations were also given on optimal structure selection of low-thrust liquid engine on alternative fuel for valve temperature reduction and power consumption reduction while thermocatalytic pack heating-up to +400 °C.

Several design options were considered, and recommendations were given on heat sink application and its impact on the thermal condition of the product, and the effect of the rack material on the thermal condition of the product. According to the results of thermal calculation of the engine structure in functioning mode recommendations are given on substitution of the engine structural elements (heater) and mounting blocks materials not answering the thermal criteria (working values the engine structural units temperature should not exceed qualification value of the temperature).

Malenkov A. A. Design solutions selection while developing a system of unmanned flying vehicles in conditions of multi-target uncertainty. Aerospace MAI Journal, 2018, vol. 25, no 2, pp. 7-15.

The article is devoted to the design solutions selection while developing a system of unmanned aerial vehicles in conditions of uncertainty. The presented article such system is assumed as a party of cruise missiles (CM) targeted at hitting an enemy naval ship grouping.

Besides solving the problem of cruise missiles optimal distribution over the target assignments this work solves the problem of ensuring stability at large. Here, the stability means achieving the probability of hitting the targets, no less than the specified one, for all possible values of uncontrolled factors.

By stability in the article is meant the achievement of the probability of defeat of target tasks not lower than given for all possible values of uncontrollable factors. Thus, the problem is set as:

where d is the vector of design parameters, E(ω) is the distribution function, and P is the probability of failure.

The distribution function E(ω) is constructed with engagement of statistical synthesis operations. A regularity criterion was adopted as a criterion of stability:

where Κ¡  is the Lipschitz constant in the i-th row of the statistical sample of the N volume, Κ¡pos is the specified value of the Lipschitz constant.

To ensure stable design solution, the contracting mapping is necessary, i.e. the Lipschitz constant should be less than one. With this, the less the Lipschitz constant value, the higher the degree of the design solution stability.

At each step of the statistical sample, two variants of design parameters are set. They are necessary for stability condition calculatiщn. The model values of the Lipschitz constant are restored in the class of trigonometric polynomials:

The problem of CM system optimal ranging is being solved at the already obtained stable vector of the design solution (the set of design parameters) yust.

The presented work solved the problem of CM system of optimal ranging, which maintains six target problems. The initial thrust-to-weight ratio and the wing area are assumed as design parameters. The target’s required payload mass, coordinates, speed and course are assumed as uncontrolled parameters.

Three nominal sizes of CMs were considered in the framework of the set problem:

Depending on the uncontrolled factors values, two variants of the cruise missiles optimal ranging were solved, and two distribution functions Ε(ω) were constructed. It is shown that the probability of the system performing the target task appeared to be the same and equals to Ρ – 0,9.

Further, the problem of a design solution selection stable to uncontrolled factors was solved. The stability conditions gave the following design parameters:

Thus, a cruise missile with such parameters solves all the target problems with uncontrolled factors given in the work, i.e. the cruise missile system includes cruise missiles of the same type, and the probability of accomplishing the target problem by the system is 0.9.

Aslanov V. S., Yudintsev V. V. Docking with space debris employing the unfolding flexible beam-strap. Aerospace MAI Journal, 2018, vol. 25, no 2, pp. 16-24.

For risk reduction of the uncontrolled growth of space debris number at the near Earth orbits, it is necessary to remove the most dangerous objects, such as the worked-out orbital stages of rocket carriers and non-operating spacecraft, which may be the sources of space debris. The most complicated stage of passive uncontrolled object's removal from an orbit is its capturing. Selection of capturing technique is determined by the type of space debris and its angular motion. For example, very often the orbital stages are being purposely spin-up around their transverse axis by jet nozzles to guarantee avoiding collision with the detached payload. It complicates capturing the object of such type by the space tug to remove them from the orbit. While employing manipulators or classical mechanisms of a beam-cone type for capturing, when the engine nozzle of a rocket carrier plays the role of the cone, significant overloads may occur in capturing units.

The presented article proposes employ for docking the expanded beam (strip) with aspect ratio. As in the classical docking technique, the nozzle of an orbital stage is being employed. The docking scheme being suggested allows reduce impact forces occurring while docking with rotating objects.

The docking assembly model was developed to study the effectiveness of the suggested scheme. The flexible beam was modeled by the system of solid bodies (beams) connected by cylindrical hinges. To imitate the bending stiffness a torsional spring was being installed in every hinge. The system model was developed using MSC.ADAMS CAE software. The system model was developed in MSC.ADAMS CAE software.

The process of docking with rotating orbital stage, using the three beams variants of large, medium and low stiffness, was analyzed through the developed model. While docking process, the reaction force value in the hinge, connecting the beam with the space tug hull, maximum tug angular velocity and the success of entire docking operation were controlled. The results of modeling confirmed the impact loads reduction while docking with reduction of the beam bending stiffness. The flexible beam will allow employ greater closing-in velocities with uncontrolled rotating objects of space debris to increase the successful docking probability. The beam elastic properties herewith allow reduce the effect of disturbance forces on the space tug while the beam contact with the docking surface of a docking port (nozzle).

Kargaev M. V., Mironenko L. A. Static stability of a helicopter main rotor flexible blade at the parking affected by wind. Aerospace MAI Journal, 2018, vol. 25, no 2, pp. 43-51.

The article is dedicated to the topical issue of helicopter design, namely static stability of an unmoored main rotor blade of a parked helicopter under the wind impact.

The article considers a general case when the wind velocity is directed at an angle to the helicopter longitudinal axis. Velocities and angles of attack of the blocked main rotor's sections are being determined. The authors used experimental data on straight and oblique wings blow-down, as well as circular blow-down of the NACA 23012 profile while determining aerodynamic loads in the blade section, blown by wind flow.

The aerodynamic load, acting in the blade section, is a function of the blade curve, and changes according to the blade's rotation azimuth. Thus, while considering the issue of the blade static stability, the problems on determining the most insecure direction and maximum allowed wind speed of the unmoored main rotor blade under specified position of parked helicopter is solved.

The article considers the blade bending in the plane of least rigidity. The blade torsional deformations are not accounted for while loads determining. It is considered, that the helicopter has main rotor of a common type with hinge mount blades.

Firstly, the solution for the homogenous blade with constant stiffness and aerodynamic characteristics was obtained. The design equation determining the value of wind flow critical velocity in various azimuthal positions was derived. It was established the main rotor blade's stability loss under the wind impact was possible only with oblique blow-down with negative sideslip angles, i.e. when the blade tip position was directed against the wind flow. The wind flow critical velocity minimum value and its corresponding direction were determined. The authors suggest employing the wind coefficient of the blade as a generalized parameter characterizing the blade tendency to the stability loss under wind impact.

Further, the solution for the blade with inhomogeneous parameters was obtained. The value of wind flow critical velocities was obtained by two methods, such as method of straightforward iteration, as well as a method, employing the wind coefficient of the blade.

The article presents the result of the wind flow critical speed computation, performed for MI-8 helicopter main rotor blades, blown-down from the front and back edges.

Vyatlev P. A., Sergeev D. V., Sysoev V. K. Holes formation mechanism while laser perforation of metallized thermal vacuum blanket films. Aerospace MAI Journal, 2018, vol. 25, no 2, pp. 37-42.

Perforation of thermal vacuum blanket (TVB) films is performed to ensure vacuum and protection from electrostatic charges effect.

Method of film materials mechanical perforation is the most widely spread for TVB films perforation. With this kind of processing It is impossible to achieve high productivity and perforation accuracy.

Laser perforation of thin materials is one of the high-efficiency technologies for processing materials, and has a number of advantages, such as increasing productivity and perforation accuracy. This method allows quick adjustment of both the diameter, and perforation step.

Fiber repetitively-pulsed laser with the wave-length of 1,062 microns was selected as laser light source. Dot cutting along the hole outline was selected as a cutting scheme.

The process of fiber laser emission action on metalized polyamide films is accompanied by bushy flame in the operation area. The reduction of laser light power and processing speed herewith results in disappearance of bright light emission and significant increase of thermal influence area width up to 300 microns.

From our viewpoint, daisy chain of the following physical effects could serve as such mechanism:

– evaporation of aluminum coating;

– ionization of its vapors;

– impact of this plasma, combined with light power, on polymer, leading to the hole cutting.

One of the evidences of such hole formation mechanism is performed physical-chemical analysis of the obtained holes' edge. The holes edge was studied by electron microscope of JEOL JSM-5910LV series together with INCAENERGY analytic system. The major results of these measurements revealed the carbon content increase in the holes edge area, while oxygen and aluminum content reduced more than three times. Thus, it can be expected that physical process of holes formation with laser perforation of metallized TVB films takes place under combined action of light power and plasma of evaporated aluminum surface layer on polymer base of the film.

Khmelnitskii Y. A., Salina M. S., Kataev Y. A. Spacecraft solar batteries dynamic analysis. Aerospace MAI Journal, 2018, vol. 25, no 2, pp. 52-60.

At present, the extensive studies of outer space are carried out to obtain scientific, economic and military results.

The solar battery is an important element of a spacecraft since it ensures functioning of its equipment.

The solar battery should have high rigidity at maximum loading factor. The structure rigidity exerts a certain effect on oscillatory process and frequency characteristics while a spacecraft maneuvering. It determines also deformations of a solar battery while its transportation to a specified orbit.

Insufficient rigidity reduces the solar battery efficiency.

The dynamic analysis of solar battery envisages determination of natural oscillations shape and frequency, and a time of the oscillatory process termination.

From these positions, comparison of the two spacecraft “Spectr-R” and 14F150 is being considered.

The finite element models were developed for these occurring while the spacecraft turn along the longitudinal axis were determined.

The inherent characteristics of a solar battery structure were being determined by the finite element method employing “NASSTRAN” software.

To determine values of inherent dynamic characteristics of a solar battery panel a series of simulations of the product dynamics were performed with parameters variation of its mathematical model.

These parameters were determined by elastic and dissipative properties of the solar battery panel.

Comparison of stiffness coefficients values and inertial links damping for these types of spacecraft revealed that the solar panels impact on the dynamic characteristics of these spacecraft was practically the same.

The transient time was of 1000 seconds, which exceeded the admissible values. For the solar battery in the considered configuration, the first mode frequency should be of the order of 0.45 Hz with damping factor of the order of 0.1.

In the considered configuration of the panels, their rigidity characteristics should be 16 times, and dissipative characteristics −3 times greater.

Nedelko D. V., Safiullin A. F. Finite element method application for determining water landing parameters of airplanes and helicopters of various types. Aerospace MAI Journal, 2018, vol. 25, no 2, pp. 61-72.

The problem of safety ensuring while an aircraft forced water landing is topical due to periodic incidents while the flights over the water. According to the European Aviation Safety Agency information, the helicopters that have passed the certification procedure topple when performing water landing. This fact indicates that the process of aircraft dynamic contact water surface is insufficiently studied, and the need to account for the aircraft spatial position parameters while water landing.

Confirmation of compliance with requirements of the airworthiness standards (AP, FAR, JAR) while emergency landing and subsequent sailing of the aircraft is based on the results of model tests, which determine the aircraft behavior, the structure loading, its possible destruction and conditions of the most favorable water landing. The hydrodynamic characteristics of models of aircraft fuselages of ground and water basing in the water landing mode, and helicopter equipment with the system of emergency splashdown are studied. A method allowing study such processes at the stage of preliminary design is the finite element method application. However, validity of the results obtained in this way should be verified based on experimental data to enable further practical application of the experience gained.

The article presents the verification results of finite element models of simple geometric bodies (inclined plate, cylinder). These are simplified models that duplicate the shapes of the amphibious aircraft float, the fuselage of the ground airplane and the helicopter ballonet. Verification was perormed employing the concept of Euler-Lagrangian interaction using the generalized “structure-to-fluid” communication simulation algorithm. For the inclined plate, the lifting force coefficients were determined for various deadrise angles and trim at its gliding on incomplete width. A graphic dependence comparing the experimental and computed values was plotted. The changing of overload at the center of gravity was demonstrated for the gliding cylinder, and comparison was performed with experimental data and approximate analytical theory. In all cases satisfactory convergence of the results was obtained.

A helicopter mathematical model with a system of emergency water landing was developed to compute the depth of the ballonet sinking, which determines the level of hydrostatic and hydrodynamic loads. The general case of driving a helicopter to the approaching slope of the wave was simulated with the presence of the initial slip at the moment of contact with the water surface. Based on the graphical dependence of the ballonet transoms movements, a technique for the computed immersion depths determining was formulated. The visualization of the helicopter position change while the water landing process is demonstrated. Based on the developed finite-element model, the other parameters of water landing of a helicopter with emergency splashdown system (overloading in the helicopter center of mass, loads on the fuselage bottom, etc.) can be determined.

The article shows, that a similar approach can be employed to simulate the process of various types of aircraft water landing, including amphibious and ground ones.

Baklanov A. V. Controlling fuel combustion process by burner design change in gas turbine engine combustion chamber. Aerospace MAI Journal, 2018, vol. 25, no 2, pp. 73-85.

Fuel burning in gas turbine engine combustion chamber entails toxic agents formation. Among them, nitrogen oxides and carbonic oxides, which prove deleterious effect upon a human and environment, present the special hazard. In this regard, the article solves the topical problem on upgrading the existing combustion chamber by changing the design of its burner.

At the first stage of the research, several types of burners, differing by nozzle extension geometry, were studied. The studies consisted in determining toxic agents' emissions concentration in the flame formed by the burner.

According to the results of the studies the inference was drawn that the most acceptable burner was the burner with convergent head piece, since it ensures minimum content of nitrogen and carbonic oxides in combustion products. The decision was made on continuing studies of both types of burners, namely, original with diffuser extension and the burner with convergent head-piece, which demonstrated minimum emission of toxic agents.

It was found that the residence time of the burner with converging nozzle extension in the reverse currents zone was 0.15 ms, and 0.025 ms for the burner with convergent head-piece, which is six times less. Testing results were colligated in the form of mathematical dependence of CO and NO from swirl parameter Sg, which characterizes the degree of the nozzle head-piece opening-out.

During the next stage, the studies on determining the throughput capacity of the burners, as well as the quality of air-fuel mixture preparation at their outlet were performed.

According to the results of the studies, it was revealed that due to the high velocity pressure there is no significant jet spreading behind the burner with convergent head-piece. The jet herewith has the high ejection capability and forms narrow flow core, in which intensive fuel and air mixing occurs. The burner with diffusion extension forms a wide concentration field and its low level, which is explained by volumetric recirculation zone.

The combustion chambers tests hereafter on determining thermal field   and obtaining hydraulic characteristics were performed. The measurements showed that at the outlet of the burner with convergent head-piece in the vicinity of thermocouple No 4 the temperature increase was observed compared to the burner variant with diffusion extension. But both cameras ensure temperature field regulated by general requirements.

While next stage the tests of the engines with the combustion chambers under study were performed. The tests data confirmed the reliability of air-fuel mixture ignition during the engine starting. They confirm also correspondence of NK-16ST throttle characteristic to the chambers with both convergent head-piece and diffusion extension in the burner.

The obtained data allowed conclude that employing the burner with convergent head-piece allowed reduce emission of nitrogen oxides by 20% and carbonic oxides by 75%. The main characteristics of the combustion chamber can be affected by changes in the design of the nozzle extension in the burner.

Mileshin V. I., Semenkin V. G. Computational study of reynolds number effect on the typical first stage of a high-pressure compressor. Aerospace MAI Journal, 2018, vol. 25, no 2, pp. 86-98.

At present methods of blade machines characteristics are widely used by many scientists all over the world. However, the applied methods of problem setting while flow modeling suppose the boundary layer to be fully turbulent in all regions, and do not reflect transient effects actually in the flow in effect. For the flows with low Reynolds numbers the problem setting with no account for laminar-turbulent transition might lead to significant disagreement between experimental and computational results.

The article presents the results of the computational study of Reynolds number effect on the first stage of high pressure K-8B compressor with the low aspect ratio of the rotor wheel blades (RW) (0.729). The stage has the following key geometry and gas-dynamic parameters:

The values of corrected specific mass flow rate through the stage are related to the values at the design point. The compressor stage regulation allows vary the setting angles of the inlet control assembly (ICA) and distributor, though at the rotor rotation frequencies under consideration (100% and 95%) zero angles were set. The ICA row, RW row and distributor row contain 46, 35 and 76 blades respectively. The gaps at the periphery and hub of guiding devices were assumed as 0.4 mm and 0.6 mm correspondingly in the stage model. The rotor row gap was assumed as 0.5 mm. The value of the total temperature at the input boundary condition is 288.15 K. For Reynolds number decrease modeling the values of total pressure were assumed as Pin = P0, Pin = 0,72P0, Pin = 0,29P0, Pin = 0,21P0, where P0 = = 101325 Pa is the standard atmosphere. The values of static pressure at the periphery were fixed on the outlet boundary condition.

Simulation of 3D viscous flow in blade channel of the stage was performed with ANSYS CFX SOLVER MANAGER in the setting of 3D averaged Navier-Stokes equations (3D RANS). The computational mesh was created with integrated automatic mesh generator ANSYS TURBOGRID and contains 3643432 elements. The solution for the setting with fully turbulent flow was obtained by Menter SST turbulence model. The calculations accounting for laminar-turbulent transition were also performed. For this purpose the Menter SST turbulence model supplemented with γ − Reθ transition model by Langtry and Menter was applied. For solutionconcordance, “stage” or in other words “Mixing planes” option was used at the rotor-stator interfaces.

According to the calculation results the stage characteristics degradation between maximum and minimum Reynolds numbers was as follows: adiabatic efficiency η*ad (4%), pressure ratio ( π* ) at the points of max η*ad (2.8%), corrected specific air flow rate (1.52%) at rotor rotation frequency n = 100%, and ∆ max η*ad = 5%, ∆π* = 4.3%, ∆Gcor= 2.3% for n = 95%. Thus, the shift of characteristics corresponding to lower Reynolds numbers occurs to the area of reduced flow of η*ad and π* . The transitional model addition affects these differences as follows: ∆ max η*ad= 3.9%, ∆π* =2.2%, ∆Gcor= 1.6% for n = 100% and ∆ max η*ad =3.7%, ∆π*=2%, ∆Gcor= 1.6% for n = 95%.

Comparing to the experimental results, obtained for n = 95%, application of transitional model of turbulence increases significantly the accuracy of the numerical study. Namely, deviations between experimental data and calculations with transitional model by values of max η*ad pressure ratio at the points max η*ad is less than 1%, while for standard SST model these deviations are of about 2% for maximum Re number, and 3.5% for minimum Re.

Comparing the fields relative to Mach numbers for two models (SST and SST γ − Reθ ), the basic difference in the flow while laminar-turbulent transition modeling consists in qualitatively true modeling of the processes occurring in the boundary layer. In this case, laminar boundary layer near the front edge of the blades, laminar separation and attachment really exist. Turbulization at the rotor wheel blades occurs at the shock wave location, after which the boundary layer already has turbulent structure for the most part with preservation of a very thin laminar layer. Besides, the changes in flow through the radial clearance in the rotor wheel are being present. For γ − Reθ “bubble” flow-over while Re number reduction slightly reduces its size. The separation near the back edge herewith becomes more intensive.

Ezrokhi Y. A., Khoreva E. A. Estimation of inlet airflow non-uniformity effect on turbofan thrust. Aerospace MAI Journal, 2018, vol. 25, no 2, pp. 99-108.

The article considers methodical approaches to developing mathematical model using “parallel compressors” method, intended for estimation of inlet flow non-uniformity effect on aircraft engine basic parameters. On the example of a two-shaft turbo-jet engine calculation at two characteristic cruise modes the results of calculated estimation, where the base value of σst and averaged value of σav stayed invariable, were presented. Parametric calculations herewith were performed for each selected relative value of the reduced pressure area.

It was demonstrated that degree of full pressure inlet non-uniformity effect on the engine thrust at the two considered modes differs significantly. Thus, if at subsonic mode this estimation could be reduced to accounting only for the effect of reduction of the averaged value of the total pressure at the inlet, while at supersonic cruise mode such reduction use might lead to considerable errors. With invariable values of pressure recovery factor at the engine entry, corresponding to the flight speed for the typical air intake, external compression σst and averaged value σav, the flow non uniformity factor Δσnu affects mainly the thrust. The degree of this parameter effect herewith depends significantly on the difference of sst and sav.

The obtained results of calculated estimations of temperature field non-uniformity at the engine inlet effect revealed that the dependence of relative thrust reduction only at the cost of relative heating was similar for the two considered modes (transonic maneuvering and supersonic flight). At the transonic mode herewith, corresponding to higher values of the reduced rotation frequency of both compressor stages, the thrust decay occurs less intensely due to relatively smaller decrease of air flow rate through the engine with reduced rotation frequency decrease due to air temperature rise at the inlet. As for the difference between the values of total thrust decay , which does characterize the effect of the input temperature field non-uniformity, with the increase of relative heating at the transonic mode it rises more intensively. It is explained by the fact that at this mode due to the less difference between air consumptions in air-gas channels of «parallel compressors» (more «dence» location of pressure downstream brunches) the speeds difference and, consequently, static pressures between the flows is much greater, than at the supersonic flight mode, which stipulates the higher losses level while these flows interaction.

Smirnov P. E., Khartov S. A., Kashulin A. P. Experimental study of radiofrequency cathode-neutralizer. Aerospace MAI Journal, 2018, vol. 25, no 2, pp. 117-124.

High specific impulse and low mass-flow rate of ion thrusters (IT) make them increasingly popular choice as a spacecraft propulsion system. Recent missions demonstrate the efficiency of these thrusters in such missions as orbit correction and exploration of Solar system. Moreover, there are many developing ideas of creating spacecraft with IT for wider spectrum of missions. However, IT needs to have a longer operation time, due to the small thrust (about several mN).

As a rule, such thrusters failure occurs due to the destruction of Ion Optics electrodes or failure of electron source. IT needs electron sources as a main cathode (for plasma producing), and as a cathode-neutralizer (to neutralize potentials of ion beam). Hollow cathodes are most used devices for Ion propulsion applications, due to low gas consumption and high electron current density.

Application of lanthanum hexaboride or tungsten with BaO impregnating as an emitter material, leads to the necessity of strict sustenance of hollow cathodes operational parameters. Interaction of emitter material with a small quantity of poison gas leads to its surface contamination and, as a consequence, to decreasing of the recoverable current even down to zero. It leads to more requirements to the gas purity, and hollow cathode handling prior to its placement in space. Moreover, to ensure effective operation, the emitter should be heated up to 0.6-0.8 of its melting temperature by the external heater, which, in turn, causes the emitter material evaporation (life span reduction), power consumption increase and longer cathode start-up procedure.

The problems of high reliability of traditional electron sources for ion thrusters led the authors to the idea apply them as cathode with plasma high-frequency discharge. In such device, plasma is generated and sustained by radiofrequency induction discharge. The absence of “loaded” (high temperature, powerful flows of charged particles) electrodes eliminates all problems of the cathode long-term operation provision. As with hollow cathode, the bulk plasma volume acts as an electron emitter, which allows generate high electron currents. The article describes the scheme of the prototype of this device, and the results of its experimental development. Currents generated by the high-frequency cathode were achieving up to 1.7 A at the input power of 120 W. Effectiveness evaluation of the high-frequency cathode is presented.

Abdulov R. N., Asadov H. G. Optimization of unmanned aerial vehicles detection in conditions of signal-to-noise ratio variation. Aerospace MAI Journal, 2018, vol. 25, no 2, pp. 125-131.

The problem of illegal unmanned aerial vehicles of various types detection and identification consists in their low flight heights, small sizes and high maneuverability. The presented article analyses the interrelated optimal selection of detection probability figures in the UAV-Radar system, corresponding to the minimum value of the signal-to-noise ratio at the output of the radar receiving part, i.e. the worst conditions of the UAV detection. The authors suggest a new setting of the problem, associated with several pulses detection at the radar input while the signal-to-noise ratio changing dynamically. The article considers the situation when the detection probability grows with time, and the integral of the sum of detection probability and false alarm probability is equal to a certain constant. In conditions of dynamically changing signal-to-noise ratio with account for the accepted condition of constancy of the integral of the sum while preserving the mutually inverse by nature character changing of detection probability and the alarm probability the problem of optimal interrelation above said probabilities values calculation the is being set. The optimization criterion was formulated in the form of the integral of the well-known expression, determining the interrelation between the signal-to-noise ratio minimum probabilities and false alarm. The gist of the formulated optimization problem consists in finding such probability dependence of false alarm from the detection probability in the series of operations of radar detection with growing detection probability, at which the minimum of the integrated value of minimum signal to noise ratios is reached, ensuring detection of point objects at each radaroperation. Based on the performed analysis the authors obtained the functional relationship of the false alarm probability from the detection probability for scenario, when a pinpoint target in the course of radar detection with growing detection probability is being detected at minimum achievable figure of integrated signal-to-noise ratio at the radar receiver input.

Mamedov I. E. Photometric informational method for unmanned aerial vehicles localization. Aerospace MAI Journal, 2018, vol. 25, no 2, pp. 132-138.

One of major factors affecting the successful UAV performing reconnaissance tasks is the possibility of its coordinates exact localization. The UAV position estimation in such systems is usually performed employing such features as reference points, margins or other images informative elements. Application of such characteristic as reciprocal information for this purpose is also possible. The major shortage of these methods consist in complexity of this function computation in real time scale. The presented article suggests the method, generalizing the main features of localization techniques based on reciprocal information calculation. In contrast to the well-known solutions, localization with the suggested technique is performed based on both information characteristics and optical illuminance characteristics of the analyzed images of various formats. The resemblance of real scene herewith with geo-referenced image is computed by subtracting them from the information characteristics, and for accuracy and reliability of the obtained result, the localization is performed based on multi-format geo-referenced images of the object. The localization problem is solved with this method as a problem of minimization of difference of the total volumes of information, obtained from the real object and reference image in the mode of studying the multi-format frames while meeting some additional condition, specified on total illuminance of the studied and compared images. As applied to the considered problem of the UAV localization, the obtained solution ensures maximum difference of estimations of information volume in the ground scene under study and geo-referenced image. The author concluded that the optimal selection should be considered as such a desired functional dependence, which differs to the greatest extent from the calculated function characterizing the studied extreme localization mode.

Spirin A. I. Flight data analysis as an operational decisions making basis of the long-term operating orbital stations usage manual. Aerospace MAI Journal, 2018, vol. 25, no 2, pp. 139-151.

Space mission control is an integral part of the control process. It allows obtain a fair presentation on the actual state and functioning of constituent parts of a spacecraft (SC), the degree of tasks implementation and its reaction to control actions.

As a rule, two tasks are solved while controlling. The first one consists in predicting the SC and crew abilities to perform the current flight tasks based on current data, and the second one of no less importance is to detect timely a failure onboard a SC and take measures to its elimination at short notice.

The analysis of the SC onboard systems state adds to the control, but this process is more complicated and it is aimed at revealing cause-and-effect relations of the control parameters both with each other and with external conditions. This analysis is performed for predicting the onboard systems state over the planned flight stages to reveal undesired tendencies in control parameters behavior, as well as for analyzing and revealing the causes of divergences and failures of the onboard system operation. The analysis of the onboard systems states is performed as a rule out of the bounds of a SC operative control loop.

For long-term orbital stations' (LTOS) the analysis of the onboard system state is particularly urgent due to the necessity of ensuring long-term operation in conditions of known restrictions on their structure changing. The flight data generalization and their analysis allow reveal the causes of divergences of the onboard systems states, elaborate recommendations on their elimination of reducing their negative effect, as well as elaborate operational decisions on optimization of the onboard systems operation modes, rational resources consumption, ensuring thereby long-term and effective operation of the LTOS.

The article presents methodological approaches employed for the of onboard systems state analysis with account for collateral data. The operational decisions examples, implemented based on the International Space Station flight data analysis, are considered for the events such as:

– parry the negative impact of the jets of orientation engines of transportation vehicles on solar batteries panels (SP);

– reduce fuel consumption during the SP effectiveness evaluation;

– improving the heat transfer of radiators of the thermal mode provision system during the «solar orbits» periods.

Kyaw Z. L., Moung H. O. Development of wind velocity estimation method using the airspeed. Aerospace MAI Journal, 2018, vol. 25, no 2, pp. 152-159.

The method suggested in this paper provides estimates for the three projections of wind velocity in earths normal coordinate system using satellite navigation systems (SNS) data, as well as on-board barometric airspeed measurements. The wind speed and its direction are assumed constant for a flight leg of 50-60 s duration. This means, that for the given time interval projections of the wind velocity values on the axis of the normal earth coordinate system are constant. Further, the object and observation models are presented, as well as the identification algorithm accuracy characteristics, obtained from the simulation data processing. The airspeed measuring error effect on the wind velocity estimation is also under discussion. The results, showing the accuracy of wind velocity estimation depending on the constant velocity measurement errors, are presented.

The analysis shows that horizontal projections of wind velocities are estimated with high accuracy (relative errors of 13%), but a certain time interval to obtain the proper degree of identifiability is necessary. After this, the accuracy of estimating the horizontal projections of wind velocities remains at a decent level, and does not depend heavily on the increase of the speed measurement error. The wind vertical projection estimation herewith leaves something to be desired. It makes 3040% even at zero flight speed error, and increases considerably with an increase of speed measuring error. Thus, we may conclude that the suggested method can ensure the good accuracy for estimating the wind velocities along the horizontal coordinate axes, and it is not applicable for estimating the vertical component of wind velocity.

Lebedeva N. V., Solov'ev S. V. Intelligent systems application while spacecraft flight operational control. Aerospace MAI Journal, 2018, vol. 25, no 2, pp. 152-159.

To perform automation of a spacecraft state control, it is necessary to define the scope of tasks, which are most dangerous from the viewpoint of their accuracy of estimate. Intelligent systems application whileoperational flight control does not assume the complete waiving from human in the control loop. It should complement his activities by in-depth and rapid evaluation of a vast amount of information, and help to elaborate the correct reaction to the current state of a spacecraft.

While operational efficiency computation, the nominal time t and its technological delay ∆t , spent for evaluation, is assumed as the main control criterion. This technological delay is associated with the time of data receiving from the spacecraft. The spacecraft normal operation evaluation is important as the main reference point for monitoring of its state changing.

While various operations execution the type of commands issued to the onboard systems to ensure the operation execution, capability of their issuing, as well as the ways of technical evaluation of the state of their execution are accounted for. For control automation, it is necessary also to account for the pre-planned possibilities of organized (nonrandom) effecting affecting its state. From the analysis viewpoint, the flight operation execution switches the spacecraft to a new state. Evaluation of the flight operation and the new state of the spacecraft is the purpose of the flight operation controlling.

The monitoring process includes also performing diagnostics of the spacecraft state. More than one point of its state herewith is determined for the current time (interval). Intelligent system application allows employ all previous diagnostic results and represents the dynamics of the development of the process of changing the technical characteristics of the spacecraft in the past, which can be used to the forecast systematic correcting and increasing its validity.

Operation of the intelligent system in real time mode will allow increase the response rate to anomalies occurrence and their development in time with accurate fixation of the drift of data deviation development. An essential advantage of such systems can be the immunity to accidental failures, such as information loss, as well as the determination of non-obvious changes, which might become a forerunner of failures.

Pigalova E. A., Abramova A. A., Kurnikov N. A. Plasma welding application prospects while airplanes of mig brand production as one of the methods to reduce welding deformations. Aerospace MAI Journal, 2018, vol. 25, no 2, pp. 172-183.

Welding is a complex technological process followed by occurrence of internal residual tensions and deformations of a welded structure.

While producing aircraft It is essential to reduce residual tensions and deformations, since:

– the structure's deformations affect an aircraft external aerodynamic contour reducing its aerodynamic characteristics;

– residual tensions sum up with tensions from external loads on the structure, leading to its destruction;

– residual tensions form volumetric stressed state in separate metal volumes, which complicate plastic deformation of a metal and contributes to its transition to brittle state, leading to local destruction of a structure.

This work is devoted to experimental research on automatic plasma welding application instead of argon-arc welding as one the methods for welding deformations reduction while aircraft structures fabrication.

Plasma welding is the welding performed by directed flow of plasma arc. The plasma arc is characterized by the high temperature (up to 30,000°C), and a wide range of its processing properties. It has much in common with argon-arc welding technology.

The main features that distinguish the plasma arc from the conventional one are:

  • a higher temperature;

  • a smaller arc diameter;

  • cylindrical arc shape (unlike the usual conical shape);

  • the arc pressure on a metal is 6-10 times higher conventional one;

  • the ability to hold the arc at low currents (0.2-30 A).

Conclusion: the plasma arc is more concentrated, powerful and universal source of heating in compared to the conventional one.

The conducted pilot studies consist in comparing parameters of the samples welded by both automatic argon-arc and automatic plasma welding. Based on the performed work, the following conclusions were drawn:

  • the sizes of the weld seam (the width of heat-affected area, the weld seam width, the samples bending angles) made by automatic argon-arc welding exceeded about 1.16 times the sizes of the weld seam made by automatic plasma welding;

  • the width of heat-affected area obtained while automatic argon-arc welding exceeded about 1.2 times the one obtained while automatic plasma welding;

  • the bending angles of the samples with automatic plasma welding are 2-3 times less than with argon-arc welding.

Based on the above said studies at NAZ “Sokol” the decision was made to implement automatic plasma welding. A new installation for sheets butt-joint automatic argon-arc and plasma welding was developed.

The installation consists of:

  • bedplate;

  • beams with clamping push buttons and mechanism for converging these beams;

  • carriages with plasma gun for automatic plasma welding and a burner for automatic argon-arc welding;

  • a mechanism for carriage move along three coordinates: along and transversely to the weld seam axis, as well as up/down;

  • supporting devices for sheet billets.

The interface of control panel software is intuitive and provides the following functionality:

  1. User identification.

  2. Identification of the installation readiness for welding.

  3. Welding programs database (DB) creating and editing.

  4. The ability of welding the parts of various thickness.

  5. Selection the already worked-out and saved welding programs.

  6. Control of welding parameters.

  7. Logging of the welding process.

The effect of implementing the plasma welding instead argon-arc:

1) Higher labor productivity in view of the higher welding speed (by 3-5 times).

2) Time consumption reduction for products leveling after welding (by 50-70%) due to minimal residual deformations in the weld seam due to more concentrated heating source.

3) Time consumption reduction for welding modes testing (by 50-70%) due to the the stored base of welding programs.

Golovnin S. M. Risk of problem solution skills loss by civil aviation pilots in uncertainty conditions. Aerospace MAI Journal, 2018, vol. 25, no 2, pp. 184-190.

Modern air transportation system is characterized by the great dependence on human, all its elements safe functioning determine the very same “human factor” playing a big role in management and stability of the entire system. In the course of time and aviation industry development, the role of the human factor in aviation accidents is being varied considerably. If for old aircraft, which were difficultly controlled and unreliable the human factor share was 5–7%, in the middle of the middle of the last century it was about 50%, and at present of the human factor is about 80% with the uptrend.

To reduce the risk of an aviation event, the Concept of Crew Resources Managing (CRM), based on the provisions of the human factor, is being actively implemented in modern civil aviation. This is a system of measures aimed at enhancing flight safety and effectiveness by the right implementation of human, technical and information resources, as well improving interaction within the crew, and the crew with the personnel of the other CRM components. CRM is an of practical implementation of the human factor principles.

The human factor as the cause of aviation event implies the human inability to react (interfere with) timely to an evolving or created emergency situation to avoid or minimize of this event aftermath.

One of the most important characteristics of a person is the response time of his reaction. In general, the response time is the time that passes from the moment of the an irritant occurrence to the motional response ending. In civil aviation, the ability to respond to irritants (signals, air traffic controllers' commands, aircraft cabin situations) is instilled in the early stages of training in flight schools. However, the practical development of reactions to events undoubtedly plays an important role in the development of the reaction rate under real flight conditions.

For this purpose, training programs for cadets include tasks for training with a list of events, which are practiced on simulators and imply the occurrence of the cadet's correct response to avoid an emergency situation development.

However, while delivering classes with cadets who are commissioned for a new type of aircraft after flight school graduation, it was noted that in the case of a series of one-type trainings, cadets began foresee a situation that wouldl be set by the instructor and developed while training process.

Thus, the effect of “suddenness” vanishes, and after all, failures or other predicaments, which may occur in flight, cannot be predicted in real flight conditions.

This regularity and foreseeing the possible scenario of situation development is able to abate significantly the pilots skill to respond and solve the unexpected problems and reduce the need for analysis and correct decision-making regarding a particular situation. As a consequence, the pilot's main skill “to fly a few seconds ahead of the aircraft” will be blurred and will subsequently be left without development, which will affect the further safe aircraft operation.

The following experiment was conducted to simulate alike situation with the cadets on the simulator. Ten cadets underwent a total of twenty training sessions, of which in 10 training sessions they knew that a simulated collision with the bird (imitation of broken glass) would be planned, and in 10 other cases the task for training did not indicated the planned collision with a bird.

The results were being recorded as follows: the training number and the number of people who could not properly perform the procedures while collision with a bird (imitation of broken glass) were recorded.

Thanks to uncertainty conditions modeling in virtual space (training device, simulator), these skills can be developed on the ground, preparing the pilot for action in almost any situation, and it does not matter whether the situation is caused by a person, a vehicle, or environment. Skills of action in the face of uncertainty will help the pilots in any case to make right decision and eliminate the problem in time.

Ismagilov F. R., Zarembo I. V., Kalii V. A., Vavilov V. E., Miniyarov A. K. Specifics of permanent magnet synchronous motor development for fuel pump of perspective flying vehicles. Aerospace MAI Journal, 2018, vol. 25, no 2, pp. 191-202.

Electric motors are one of the main actuating element ensuring aircraft systems functioning. Traditionally, they are employed in fuel pumps, oil pumping pumps, hydraulic stations, automation systems, as fans drive, and wing-flap systems. The variety of problems solved by electric motors on board the aircraft, makes them almost one of the main consumers of electric power.

Currently, several types of electric motors are employed in aircraft fuel pumps, such as DC motors with brush-collector unit, induction motors, inductor and reactive motors, permanent magnet synchronous motors (PMSM) with direct start, and brushless direct current motors (BLDCM). All the listed motors have problems related to energy efficiency and mass and size indicators.

Thus, the main promising motor version for employing in aviation fuel pumps at this stage is the PMSM. A number of scientific and practical works are devoted to the development of the PMSM for aerospace systems. In particular, the specifics of field simulation of the PMSM for aircraft air-conditioning systems and general approaches to PMSM development for aerospace applications are considered. The works are devoted to the study of the PMSM magnetic systems and solving the problems of creating a PMSM control system development. The design features herewith of PMSM for fuel pumps are not disclosed in the literature. Although this type of motors has a number of distinctive features, such as working conditions in the field of low negative temperatures, working capacity at low voltage, employing of graphite bearings, etc. All these specifics do not allow employ the results of the works to the full extent.

Thus, the purpose of this article consists in analyzing the design features of the PMSM for fuel pump by developing and examining the PMSM for fuel pump with concrete geometry with account for real operating conditions and evaluating the prospects for the development of the PMSM for fuel pump.

Nadaraia T. G., Selivanov A. I., Shestakov I. Y., Fadeev A. A., Vinogradov K. N. Hybrid energy storage device in power supply system for prospective spacecraft. Aerospace MAI Journal, 2018, vol. 25, no 2, pp. 202-209.

The article presents the improved version of spacecraft power supply system by rational selection of the hybrid power plant basic elements. Power supply system is the most important onboard system from the viewpoint of energy supply and reliability. Failure of this system entails failure of the whole spacecraft.

The main types of power plants, such as a combination of solar and chemical batteries, installations based on various physical phenomena, and electrodynamic tether systems, as well as nuclear ones are known.

Rational selection of the power-plant basic elements to solve specific problems allows improve technical, mass-and-size and cost characteristics of a spacecraft in total.

The improvement of the power supply system energy efficiency is achieved by special schematic architecture and joint application of chemical and kinetic energy storage devices. The hybrid energy storage device will allow maintain the required energy supply of the onboard equipment and compensate peak energy consumption onboard a spacecraft. This energy storage device includes ionistors. Ionistors serve to compensate fast transients while the installation start-up in orbit. Compensation of the occurring kinetic moment is realized by installing two energy storage devices operating in antiphase. Application of contactless, magnetic, high-temperature super-semiconductor suspension in the flywheel allows significantly reduce mechanical losses and increase the storage time of the stored kinetic energy.

The principle of the above said installation operation in both energy storing mode and energy return to the system to consumers' mode is described. The hybrid energy storage device operation in the process of energy return takes place with rotation speed changing, which leads to the necessity of solving the problem of obtaining the AC of stable frequency at the output. This problem is being solved directly by rotating converter or a specialized inverter. Smoothing the peak loads on the battery by ionistors and the lack of brush gear increase the lifespan of the hybrid energy storage device.

Indicative computations show that application of the hybrid energy storage device allow improve mass-and-size characteristics of the power supply system by 24%. The suggested approach will be employed in further activities associated with enhancing the energy-mass perfection of the spacecraft power supply system.

Galkin V. I., Galkin E. V., Paltievich A. R., Preobrazhenskii E. V., Borunova T. V. Analyzing technological schemes of production of “FRAME SEGMENT” type parts. Aerospace MAI Journal, 2018, vol. 25, no 2, pp. 210-220.

The article considers methods for frame segment obtaining from the B95 alloy by isothermal forging. This method allows obtaining forgings with minimum allowance for machining and requires value of punching force. Isothermal forging can be a more productive alternative to the now employed cutting operation with NC machine tool. The above said alloy is a certified material for aircraft industry and has a high specific strength. One of the B95 specifics consists in rather narrow deformation temperature region. On the one hand, herewith the forging temperature should be selected as maximum to reduce the required force, and on the other hand, the deformation heating-up may lead to overburning, i.e. irreparable damage of the material, characterized by drastic mechanical properties deterioration. To solve this problem, the authors propose to reduce the deformation loading of the material, which can be ensured by controlling the stress-and-strain state and heating temperature of a workpiece while forging.

The stress-and-strain state of temperature fields analysis was performed with engineering software complex Deform, based on finite element method. Deform software found wide application for the analysis of metals pressure shaping. It allows reduce the design period of the process and cost price, as well as increase the quality of production.

In the presented work several options of isothermal forging of a frame forged piece made of B95 allow were studied with finite element method. While modeling, the initial temperature of the process was being varied, and forging tools of various geometry were employed, as well as the auxiliary operations number. Workpieces of various cross-sections, such as circular, square and rectangular ones were used. The initial workpiece position in the stamp was accounted for. For all cases under consideration, the deformation ratio exceeds the permissible value of 60%, and the process temperature was non-uniformly distributed over the forging cross-section. In a number of cases the conditions that could lead to metal burn-out were observed. It was found, that the most rational scheme is the scheme of isothermal forging, in which a rectilinear pressed rod was used as a billet. Its cross-section area was equal to the section area of the frame forging, and the length of the shelves was 3 mm shorter. This scheme application allows produce forging with equivalent strains of no more than 60%, and allowable deformation heating, which does not lead to the of B95 alloy burnout.

Antipov V. V., Dobryansky V. N., Korolenko V. A., Lur'e S. A., Serebrennikova N. Y., Solyaev Y. O. Evaluation of layered aluminum-fiberglass plastic effective mechanical characteristics in conditions of uniaxial tensile. Aerospace MAI Journal, 2018, vol. 25, no 2, pp. 221-229.

The article presents the results of laminated aluminum-fiberglass composite material, formed by thin layers of aluminum alloy and fiberglass, mechanical characteristics modeling. A modified analytical model of layered material accounting for the presence of metal elastic-plastic layers in the composite structure with bilinear defining relationships is being employed for calculations. For the case of uniaxial tensile, the layer-by-layer analysis of the composite strength is being performed with account for residual tensions formed while the material fabrication. The Tsai-Hill strength criterion was used for fiberglass layers. The moment of yielding commence in metal layers is being determined by Mises criterion. The calculation results determined effective strength characteristics, yield stress and strength limit of composites in conditions of uniaxial tensile. The good agreement of calculation results and experimental data within the 90% of accuracy limits was shown.

The effective Young's modulus of the material in the calculations was 51.5 GPa (49 GPa in the experiment). The apparent yield stress of the composite, associated with the appearance of plasticity in the layers of aluminum, was 230 MPa, which in fact coincides with the experiment. The composite ultimate strength in calculation was 540 MPa (585 MPa in the experiment). In fact, it follows fr om the calculations that the yield stress of metal-polymer composite is determined by aluminum layers yield stress, while the strength limit is determined by the strength lim it of fiberglass layers oriented in the direction of load action. The proposed model allows evaluate the effect of residual tensions on the material mechanical strength characteristics. The results of calculations established that the residual tensions might lead to the composite mechanical properties degradation.

Kalugina M. S., Remshev E. Y., Danilin G. A., Vorob'eva G. A., Telnov A. K. A method of light alloys reinforcing by aero-thermoacoustic treatment for aerospace industry. Aerospace MAI Journal, 2018, vol. 25, no 2, pp. 230-239.

The article studies the possibility of developing technological basics of higher mechanical properties of aluminum casting alloys ensuring, and wrought aluminum alloy while employing aero-thermoacoustic treatment (ATAT).

The share of aluminum allows employed in aviation industry is high. Thus, both casting and wrought alloys find application in aerospace industry. Casting aluminum alloys are used for containers and tanks production. In machine building such casting aluminum alloys as silumin are widely spread.

Aluminum wrought alloys present great interest, due to their higher mechanical properties. They are used for aircraft hulls manufacturing. The above said alloys are employed for manufacturing prefabricated shells of aircraft hulls, representing rigid encasements of rather rigid sheet material, which should resist normal and tangent forces and carry all types of loads.

ATAT employing enables increasing the strength of silumins about 1.4 times, practically with preserving elasticity at the initial level or its slight reduction. Significant holding time reduction was observed as well.

The article studies ways of increasing strength characteristics of extra-high tensile wrought aluminum alloy without significant loss of plastic properties of the material.

The article studies ways of increasing strength characteristics of high-strength wrought aluminum alloy without significant loss of plastic properties of the material.

The ATAT effect on the structure and properties of aluminum casting alloys was revealed, which could be associated with the process of micro-plastic deformation and partial recrystallization while treatment, with diffusion processes acceleration, which ensures grinding of solid solution grains. The redistribution and reduction of macro and microstrains in the material significantly affects its properties.

Aslanov V. S., Yudintsev V. V. Parameters selection of space debris removal system with elastic elements by cable towing. Aerospace MAI Journal, 2018, vol. 25, no 1, pp. 7-17.

There are more than 1500 large artificial objects on the near-earth orbits, while only 7% of then are active spacecraft. The remaining objects are space debris. The greatest hazard is presented by the large space debris, such as non-functioning satellites, final stages of rocket carriers, staying on the orbit. Their destruction can lead to grave aftermath, since collision of such object with other the objects and fragments may lead to significant increase of the number of small debris, which, in its turn, can lead to impossibility of safe employing of some near-earth orbits. The space debris removal is one of topical problems, which humanity will have to solve in the nearest future.

A method of space debris removal, and transportation system parameters are determined in many ways by the properties of the garbage being removed. Objects capture and removal by tether systems is one of the prospective methods of large objects, such as non-functional satellites of rocket stages, removal from orbit. The removal of a non-functioning spacecraft with flexible elements herewith is a more complicated task, since the possibility of oscillations of elastic structural elements, such as solar batteries panels should be accounted for, which may lead to their destruction and greater clogging of near earth space.

The article considers cable transportation of a large-sized object of space debris with elastic elements, such as solar batteries. The goal of the work consists in studying the mutual effect of tether oscillations and oscillations of flexible elements while transportation active phase. The article presents the developed mathematical model of the system, consisting of space tug and towed space debris with flexible elements. It considers the simplest case when only a constant thrust force effects the tug. No other forces and moments (such as gravitational) are accounted for.

The transported space debris should not be destroyed while towing, and its attached elements (solar batteries) should not tear away. Otherwise, it may lead to greater clogging of space. To analyze the possibility of destruction and selection of such system parameters that will exclude the space debris structure destruction, mathematical model was developed. By dint of this model, the analytical expressions allowing select the tether rigidity depending on parameters of space debris and mass of the tug were obtained. The article demonstrates the existence of critical tether rigidity, that should be avoided while transportation system parameters forming. Direct numerical integrating of the initial equations of the motion substantiated all analytical and numerical results presented in the article.

Guryanov A. I., Kalinina K. L. Studying an atomizer for rain imitation while aircraft engines certification. Aerospace MAI Journal, 2018, vol. 25, no 1, pp. 18-27.

The purpose of the study is creating an atomizer for aircraft engines testing while ingress of rain, as well as checking the sample for conformity to the standard requirements for testing facilities adopted for aircraft engines' certification.

The review of fluid spraying problems and methods, which formed the grounds for further selection of the liquid spraying scheme, was performed in this research work. The paper presents the description of the technique for the pursuance of the pilot studies of the atomizer with determination of the parameters such as flow coefficient; water distribution irregularity ratio; rooted angle of a drip stream and drops distribution over the diameter with computation of the average median diameter. It presents also the scheme of installation for complex study of a water drip stream characteristics.

Experimental studies of atomizer prototype models were performed according to the above said technique for the purpose of increase integral parameters of the efficiency, as well as compliance check of range of drops diameters from 0.5·10-3 to 7·10-3 m, and the value of average median diameter of 2.66·10-3.

The tests allowed revealing the relationship between the geometric characteristics of atomizers and drip flows being obtained. Development of the most suitable prototype of atomizer allowed obtain the drops within the certification range with average median diameter of 2,656·10-3 m.

The results of the work are as follows: the problems of rain imitation were analyzed, the technique for the atomizer testing was developed, and the atomizer design was offered and substantiated. Experimental studies of parameters of the above said atomizer were performed design, and its conformity to certification requirements was confirmed.

Shorr B. F., Buyukli T. V., Shorstov V. A., Bortnikov A. D., Sal'nikov A. V., Frolov V. N., Serebryakov N. N. Developing calculation method for forced vibrations of turbomachines of a blisk type blades. Aerospace MAI Journal, 2018, vol. 25, no 1, pp. .

The subject of the article “Developing calculation method for forced vibrations of turbomachines of a blisk-type blades” by Shorr B.F., Buyukli T.V., Shorstov V.A., Bortnikov A.D., Salnikov A.V., Frolov V.N. Serebryakov N.N. is the blades of a blisk-type rotor wheels.

The research topic is the effect of amplitude-dependent damping in the material of blades on amplitude of the steady-state resonant vibrations.

The goal of the work is definition of the non-stationary components of the aerodynamic forces and resonant stresses amplitudes in the blades at steady-state vibrations.

The article employs the following assumptions: only the steady-state vibrations amplitudes are being computed. Aeroelastic phenomena relating to blade deformation (both oscillations' excitation and damping) are neglected, i. e. gas exciting forces are defined according to the geometry of air-gas channel elements at a specified operating mode regardless of blade vibrations. Mechanical damping in blades material is amplitude-dependent; i. e. blade behaves as a physically heterogeneous body in the sense of energy dissipation, which heterogeneity depends on variable tensions distribution at each form of vibrations. Damping properties are verified by dynamic tests of samples at various excitation levels and frequencies.

The methodology of the work includes a sequential computational study, which consisting of initial normal modes analysis with definition of the operating mode with possible resonances. It also accounts for of the non-stationary components of the aerodynamic forces definition by solving the Navier-Stokes equation at the operating mode of interest, transferring these components to the nodes to the mechanical finite element model of the blade. Finally, the extraction of the harmonic components of the force, and solving the problem of steady-state vibrations of the blade with amplitude-dependent damping.

Calculations revealed that employing of the constant decrement of oscillations might lead to incorrect results. The difference between calculated amplitudes of the vibratory stresses in the considered example was 25%.

Conclusions were drawn on the method structure, as well as that the considered example of calculating the rotor wheel forced vibrations at resonance with the 13th harmonic of the flow circumferential irregularity shows the utility of accounting for the dependence of the energy dissipation factor in the material on the vibratory stresses amplitude.

Il'inkov A. V., Gabdrakhmanov R. R., Takmovtsev V. V., Shchukin A. V. Effect of centrifugal mass forces on heat transfer when airflow of concave surface with transverse projections. Aerospace MAI Journal, 2018, vol. 25, no 1, pp. 39-48.

The article presents the results of a pilot study of average heat transfer on a concave surface model with regard to the cooling systems of the leading edge of a gas turbine engines turbine blade with spanwise semi-cylindrical ribs in turbulent flow. Relative curvature parameter was being varied by variation of the momentum thickness. Heat transfer has been studied employing a gradient method based on Fourier-Newton law. A test section was a plane channel of 140 × 100 mm consisting of a straight section and a 90° bend. The concave surface of the channel and the object under consideration had a radius of curvature Rw = 500 mm.

The visualization results revealed that when an undisturbed fluid flowed past the first spanwise rib, the reattachment length behind this rib depended on the surface curvature parameter. The latter is the ratio of the momentum thickness to the surface curvature radius. The increase in this parameter fr om 1.38 · 10-3 up to 2.5 · 10-3 resulted in the average of 1.6 times reduction in the reattachment length.

This result derived fro m flow visualization has been satisfactorily confirmed by the distribution of local heat transfer coefficients between the ribs. The reattachment length characterized by the peak heat transfer reduced approximately by 1.4 times. No effect of centrifugal body forces on heat transfer in the flow around the second and third ribs has been observed.

It has been shown that in the case of combined effect of centrifugal body forces and spanwise ribs on heat transfer, these factors do not meet the additivity concept of individual effects due to their mutual coupling. In the considered case, the effect of streamwise curvature of the concave surface was observed only behind the first spanwise rib wh ere the momentum thickness was large. This effect was suppressed further downst ream byboundary layer breakup caused by spanwise ribs. The contribution of centrifugal forces to heat transfer enhancement at a given surface curvature radius can grow if the rib height is decreased while the streamwise rib pitch remains constant.

Marchukov E. Y., Polyakov K. S., Kulalaev V. V., Petrienko V. G. Computation of magnetic liquid flow in annular channel of magnetic-fluid seal of a shaft with high-speed wall. Aerospace MAI Journal, 2018, vol. 25, no 1, pp. 49-56.

Purposes and objectives of the article consist in the following: formulating hydrodynamic boundary problem of computation of magnetic-fluid seal (MFS) parameters, which belong to the group of noncontact slot seals operating as a hydraulic lock. While developing MFS the annular packets of conducting magnetic brushes were used as magnetic field concentrators instead of teeth. A magnetic fluid resides between the bristles of these brushes in a narrow annular channel. Such a seal gives the minimum friction between the interfaced parts. Numerous calculation methods for the abovementioned hydrodynamic boundary problems developed historically almost independently from each other. General principles for creating numerical methods acceptable for all hydrodynamic boundary problems in general were slated. The aggregate of these concepts and methods allows eventually reduce the algorithm for solving complex hydrodynamics boundary problems to algorithms for solving simple problems of standard structure. An integral relations method employed in this work was developed based on conservation laws and eventually reduced to ordinary differential equation solving. At the domain boundary herewith the boundary conditions are specified both at the rigid wall and the boundaries through which the flow inflows and outflows. Additionally, adhesion conditions are specified. The presented article formulates the new boundary conditions of tracking-concatenation of viscous incompressible flow for internal flows in narrow annular channels. It gives mathematical formulation of the boundary problem for viscous incompressible magnetic flow with possible internal backflows, which detection experimentally is impossible. The boundary problem was set and algorithm for computation of viscous magnetic liquid flow field in the annulus with movable walls of the magnetic-fluid seal (MFS) by the structured method with the exact fulfillment of the boundary and initial conditions was presented.

The article shows that application of mathematical apparatus for solving the boundary problems by the structured method allows calculate in total parameters of the magnetic liquid flow: heat flows, coefficients of friction, heat transfer and distribution of these parameters through the radial clearance of annulus, revealing the areas of potential backflows.

The results of this work may be useful while developing and computing new type of magnetic-fluid seals (MFS) for high-speed shafts of structures and units for various industrial purposes.

Ezrokhi Y. A., Kalenskii S. M., Morzeeva T. A., Khoreva E. A. Accounting for the effect of the border layer at the inlet to the fans while integrating the distributed power plant and a flying vehicle. Aerospace MAI Journal, 2018, vol. 25, no 1, pp. 57-66.

The article presents the analysis of a distributed power plant concept for perspective long haul passenger aircraft, which is intended for ensuring more deep integration of a power plant and a flying vehicle, as well as enhancing its fuel efficiency.

While employing an aircraft engine of such kind, separate modules of a power plant may be installed both in the engine nacelle and inside an airplane fuselage, made according to a “flying wing” scheme.

A portion of a boundary layer, formed at the surface of an aircraft, gets into the inlet plane of fan modules, located at the top surface of the fuselage.

The variant of a submerged engine inside an aircraft assumes the presence of a rather long curvilinear intake channel, in which local separations and vortexes inevitably occur. It leads to additional losses of full pressure at the engine inlet.

The article considers separately the effect of two main factors on the engine thrust, namely, the drop of overall level of the total pressure at the engine inlet and its non-uniformity.

To evaluate the effect of the above said components, the results of preliminary work out of the distributed power plant parameters, obtained at CIAM, named for Baranov, in the activities progress on the engines' schemes of new types, were applied.

Calculations were performed employing the first level model of an aircraft gas turbine engine.

Parametrical studies performed using the developed technique allowed select an optimal degree of double-flowness on specific fuel consumption at course speed, and the degree of pressure increase in the fan. The fan modules' and main engine components dimensionality was redetermined with account for various losses levels at the inlet.

The effect of engine parameters changing on the its mass estimation value was performed with the developed modular technique, based on the idea of impeller machine mass proportionality to compression specific work and corrected specific air consumption. The modular technique coefficients characterizing the weight fraction of the turbojet modules were determined based on estimations obtained for detailed element-by-element mathematical model of mass, in the activities progress on the engines' schemes of new types, at CIAM, named for Baranov.

The obtained results of the parametrical studies make it clear that on deterioration of the factor of total pressure preservation at the inlet by 2%, minimum specific fuel consumption at a cruising mode would be achieved in the distributed power plant with double-flowness reduced by 3%, and the total pressure increase degree in the fan reduced by 0.6%. At the same time specific fuel consumption increases on 6-7 % of percent. The specific fuel consumption herewith is increases by 6-7%.

The power plant weight, without account for the weight of the remote fan modules transmission drive may increase by approximately 4-5 %.

Analysis of the effects associated with the presence of non-uniform total pressure field, resulting in its averaged level reduction at the fan inlet, revealed that the effect of non-uniformity presence itself might be of 15 to 30% of the total effect on the engine thrust. It should be accounted for selection of the distributed power plant shape of the configuration under consideration.

Pisarenko V. N. Testability management while an object operation. Aerospace MAI Journal, 2018, vol. 25, no 1, pp. 67-75.

Foreign-made aircraft (AC) ingress to domestic civil aviation airlines revealed a number of significant challenges, including the testability provision (abbreviated Tst). We will denote testability management hereafter by FTsbl symbol with Tstblt subscript. An ill-considered implementation of foreign-made components in aviation transportation system of Russia without comprehensive accounting for operation and maintenance factors leads to above-level downtime of cost intensive aerotechnics, and upset of calculated value of an aircraft testability. At present, revealing assessments and factors of testability management, gains special topicality and requires comprehensive analysis. Many scientists in Russia, including V.S. Shapkin, N. Gipich, .G. Evdokimov, A. Stepanov, V. Viktorova and abroad, including Douglas, T. Ross, studied testability as the means of equipment failure-free operation provision through its whole life cycle. However, the studies of testability provision while operation are insufficient. The testability management system is being reduced to compliance with the State Standard 27518-87 “Products diagnosis”, i. e. to totality of coordinating activities on management state, as a part of general enterprise management. These activities are not oriented with respect to testability while operation. They are fulfilled without adequate theoretical development on substantiating the required acceptable testability level of object under operation and control action. It does not achieve the desired goal since functional dependencies of testability management, controlled parameters and acceptable limits of testability parameters variation of controlled products are not substantiated theoretically.

The objective of this article consists in studying the possibility of testability management while operation and developing mathematical model of testability management of an object on the example of testability management of aerotechnics.

The article describes the testability as a function of the monitored object under operation. It presents description of testability computation models and algorithms. Based on the theory of optimal processes and Pontryagin's maximum principle the mathematical model of the function test was studied. A mathematical model of an operated object testability management on the example of aerotechnics. This model is based on measuring indices and parameters of operation, processing of the obtained data, analyzing and developing control action on the operated object.

A mathematical model of controlled object under operation testability on the example of aviation technology, based on the measurement of parameters and operating parameters, the processing of this data analysis and generation of control action on the object of exploitation. An approach to testability management of an object under operation was deduced.

Patrikeev S. A. Capabilities of onboard innovation measuring systems while ground and flight tests. Aerospace MAI Journal, 2018, vol. 25, no 1, pp. 76-83.

The main modern aviation development trends are based on the fact that aircraft qualities are defined not only by carrier characteristics but also by onboard equipment complex capabilities.

High rates of airborne equipment development and creation intrinsic to recent years have come into contradiction with long service life of airframe and engine. Resolution of this conflict supposes abruption of an aircraft common life cycle as an aggregate of an aircraft and its equipment and shift to logically interrelated separate cycles of aircraft and onboard equipment complex development.

The problem discussed herein consists in optimum employment of cost and time resources in aircraft flight and engineering test practice.

Particularly, the flight and engineering tests are described, which essence consists in giving the answer to the question on how the flight task was realized with the accuracy not worse than the specified one.

Parametric expected uncertainty within the problem formulated has some specific distinctions from situations discussed within statistical decision theory.

First of all, the values of the parameters, which define the hypotheses under checking. These parameters, H0 and H1, are not defined (a priori) within the sets of their values Ω0 and Ω1, responsible for the system state (H0 – the system complies with the requirements, H1 – it does not), and are defined in the sense of the system state (“YES” – H0, “NO” – H1).

Secondly, inasmuch as on the assumption of employing information methods for optimization of surveillance planning stage at the interval of an aircraft's ground tests the situation, when the probability in the context of alpha and beta errors is required, is inadmissible. The decision making in this case will turn out to be unobtainable due to the lack of information in the sample of observations.

Substantiated information and cost approach, general formulation and the ways of resolving the problem of surveillance of ground measuring complex means while performing aircraft flight and engineering tests, ensures the effectiveness of flight tests with existing test pattern and requirement for minimum consumption of all kinds of resources.

Proved relationship and interpretation of the results open a possibility of obtaining analytical expression of informational measures necessary within the framework of the problem discussed and formulation of the task for ground measuring system equipment observation plan optimization.

While application of this method, the effectiveness of proposed models was about 9 –15 % of augmentation in terms of economic indicators, and instruments and general structures controllability by 15 – 20 %. Thus, general effectiveness of the proposed model equals to about 20%, which allows for attributing it to qualitatively new flight controllability structures.

Dong Z. . Analysis of dynamics and motion control of low-orbital space tether system. Aerospace MAI Journal, 2018, vol. 25, no 1, pp. 84-91.

The paper analyzes the dynamics of a low-orbital space tether system (STS), consisting of a main and a small space vehicles, and a tether connecting them. Under consideration are the stages of deploying, free motion and stabilizing on a low and nearly circular orbit (170-180 km). The tether escapement is performed fr om the main spacecraft by the mechanism operating only on braking action, according to the feedback principle of measuring the tether length and escapement velocity. The tether length after deploying termination is several tens of kilometers.

The study of the STS motion on a low orbit becomes more complicated due to the necessity of considering the atmospheric drag acting on all system elements including the tether. It was demonstrated, that at the end of the STS deploying in a position close to vertical, unavoidable system oscillations relative to vertical occurred, caused by joint affecting of gravitational and aerodynamic forces (aerogradient effect).

The author suggests a nominal deploying program of the low-orbital STS at the position near to vertical. The proposed STS deploying program, compared to the known programs, accounts for the effect of the aerodynamic force acting on the end-bodies and the tether. The program law elaboration is realized by a simplified model with inextensible tether, and written in the orbital moving coordinate system. To verify the effectiveness of the suggested program the STS mathematical model with distributed parameters, wh ere the tether is represented as an aggregate of material points was elaborated and applied. Numerical simulation of the deploying process revealed that the suggested nominal program of the STS deployment allows decrease the amplitude of aerogradient oscillations of tether relative to the vertical by several times.

Simulation of the stages of free motion and stabilization was performed on the model with distributed parameters. When the orbital height of the system's center of mass decreases to a certain value, the low-orbital STS will switch to the stabilization motion in a given range of orbital height (170-180 km). Stabilization of the system orbital motion is realized by a correcting thruster, located on the main spacecraft. Employing the correctiing thruster ensures the flight stabilization of the low-orbital STS in the given range of orbital height. At the stage of STS motion stabilization, restrictions, imposed on the tethers angle deviations from the vertical are executed.

Repin A. I., Kashkina T. I. Specifics of application of minimax operations for aircraft lateral movement control. Aerospace MAI Journal, 2018, vol. 25, no 1, pp. 92-98.

The safety of aircraft landing on approaching the landing strip in difficult weather conditions is associated not only with the need to create light and strong devices, but also, mainly, the search for new principles (methods and tools) for building control systems, since the aircraft landing is the most laborious process and largely unsolved problem to date.

Safety upgrading is achieved by control automating while approaching the landing strip and aircraft landing. It is obvious that the use of standard methods for modeling, analyzing and managing of complex multi-level systems becomes less possible with complexity increasing. In this situation, fuzzy control methods are the most applicable to such complex technological processes as the control of aircraft landing.

Aircraft control systems based on the principles of fuzzy logic, allow increase the course stability of the aircraft. In such situations, energy consumption is reduced and the response time of the system is increased simultaneously. Besides, it is possible to make the system as a whole more stable to the effect of disturbing factors compared to the traditional aircraft automatic control systems.

Practice shows that the operator, in conditions of good meteorological visibility range, satisfactorily lands an aircraft without the help of a program control system and a trajectory control system.

In the case of poor meteorological visibility, with the lack of visual contact with the runway strip, radio technical, optoelectronic and inertial navigation systems are employed for aircraft landing. They are used in the control system as sensors of primary information for the automatic control system (ACS). Such systems are termed course-glissade systems. They determine the position of an aircraft on the course and on the glide path.

But, even with modern control systems provision equipped with computerized hardware and software systems, which functionality is largely determined by software, applied diagnostic models, information processing algorithms, etc., the final decision-making is delegated to the human, which is a consequence of the insufficient effectiveness of diagnostic models, reflecting real ACS and the environment.

Thus, the structure schemes of similar systems in the following stages should include the links with fuzzy transfer function WN(p) instead of links with functions Wo(p ) or Wa(p ) . To this effect, it is rational to implement the of the operator's behavior in such a situation as the basis for the fuzzy controller synthesis. In this situation, namely the methods of fuzzy control are the most applicable to such complex technological processes that will allow reduce by 10 times the duration of the longitudinal and horizontal movements' transients. The pilot in this case operates as a controller for the state of the control system.

Thus, the task consisted in developing models and algorithms for the design of control systems based on the methods of the theory of fuzzy-multiple apparatus.

A program in the C++ programming language was created to reproduce the min and max operations in on-board systems for automatic control of the aircraft lateral movement with applicaiton of fuzzy logic.

Kim N. V., Bodunkov N. E., Mikhailov N. A. Automated decision making by the onboard unmanned aerial vehicle system while road traffic monitoring. Aerospace MAI Journal, 2018, vol. 25, no 1, pp. 99-108.

The article presents the developed method for efficiency increase of the operator, performing traffic surveillance by an unmanned aerial vehicle (UAV) with the built-in computer vision system. Analyzing video information, received via the radio channel in real time mode by the human-controlled point, is associated with errors in decision-making. These errors are stipulated by the vast volume of information, which overburdens the operator, and, as a rule, by the so-called human factor. Productivity of such system can be increased significantly through addition of autonomous road situation estimation system. The UAVs equipped with surveillance systems, such as video cameras, receive images onboard (video sequences), and are able to extract from them the objects of interest: roads and transport means.

Estimation and analysis in this article are ensured by the road incidents consequences severity classification. The work employs the classification consisted of five classes. Each situation class is described by attributes' dictionary, which separates the attribute space into non-crossing areas, corresponding to the selected classes.

In addition, the article describes the developed hierarchical structure of “Description of the Scene Being Surveyed”. This structure relates to the so-called semantic descriptions, is rather universal, and ensures the possibility to describe various road traffic situations.

The article presents the technique for traffic situations classification over the images. It demonstrates the example of the situation classification based on the real image of the road accident.

Aglyamutdinova D. B., Sidyakin S. V. An object bounding box refinement algorithm while the tracking process initialization from the uav. Aerospace MAI Journal, 2018, vol. 25, no 1, pp. 109-121.

The presented article deals with the problem of semi-automatic initialization of the selected object tracking by unmanned averial vehicles (UAVs) or drones. Here, we proposed an algorithm of the position and sizes refining of the boundary rectangle of the tracked object at the start time (on the first frame) based on saliency detection algorithm, which simulates the map of human attention. The advantage of the proposed approach is that it applies the principles used by the human visual system: the color contrast, the main attention is centered on the central objects. The first stage of the proposed approach consists in preliminary image processing (noise removal) by the Gaussian filter and converting the image into the CIE LAB color space. The next stage is segmenting the image into homogeneous areas (superpixels) by simple linear iterative clustering (SLIC) algorithm. Undirected graph is employed as a container for information on segments storage. Based on information from the resulting graph, measures of identity, which assign superpixels to the background or an object, are computed. The resulting saliency measure is computed for each superpixel by optimizing the target cost function, which combines the measures of identity to the background, an the object and the smoothing component. The obtained saliency map of the image superpixels is binarized by the Otsu method. After that, the pixels belonging to the shadow can be determined. At the final stage, the operations of morphological filtering were applied to reduce fragmentation of objects and an algorithm for allocating coherent components, assigning the final dimensions and position of the object of interest for tracking initialization.

The algorithm was used to initialize a number of fast and effective methods of object tracking: DCF_CA, MOSSE_CA, SAMF, DCF, DSST, MOSSE, SRDCF.At the same time, the quality of the tracking was tested on the largest and most complex database of video clips, shot from an unmanned aerial vehicle – UAV 123.

The results of experimental testing allow conclude that the best tracking quality as a result of initialization by the proposed algorithm is achieved by tracking algorithms “SRDCF” and “MOSSE_CA”. In assessing the performance, you can notice that “MOSSE_CA” tracking algorithm is noticeably superior to the other algorithms. In this way, the most suitable algorithm for tracking objects by UAV, along with the proposed initialization algorithm, is “MOSSE_CA”, due to its least sensitivity to the of initial initialization accuracy and fastness among competitors.

The proposed algorithm does not require special hardware and can work in real-time. It is implemented in C ++. The average time required refining the object, occupying 40% of the image size of 256 × 256 pixels, equals 60 milliseconds on the Intel® CoreTM i5-3470 CPU @ 3.20GHz.

Abdulin R. R., Zudilin A. S., Obolensky Y. G., Rozhnin N. B., Samsonovich S. L., Stitsenko A. N. Developing of an electromechanical actuator of the higher reliability with redundancy. Aerospace MAI Journal, 2018, vol. 25, no 1, pp. 121-131.

Popular in recent decades concepts of an “all-electric aircraft” and “more-electric aircraft” assume full or partial replacement of centralized hydraulic systems by centralized electrical systems, and hence application electromechanical and electro-hydrostatic actuators alongside with electrohydraulic steering actuators. In the aircraft of the abovementioned class of the nearest perspective, the electric actuators with hydrostatic transmission are more applicable for the main steering surfaces controll. It is associated with higher indicators of wear resistance and non-failure operation of hydraulic cylinders. Electromechanical actuators application for these purposes is constrained by their insufficient reliability.

The article proposes solutions aimed at increasing the reliability of electromechanical actuators by both element-by-element and structural redundancy. An obligatory element of an electromechanical actuator optimized for weight and size indicators is a mechanical gearbox, which can wedge while operation. Multichannel electromechanical actuators can be constructed by one of the considered schemes, free of gearbox wedging. According to these schemes, each actuator channel must contain a motor shaft locking clutch, employed in case of a channel failure, and while channels serial operation as well. Alternative option are the schemes, requiring employing of clutches splintering the faulty channel off the common load – the steering surface. Such clutches should have reliability indices higher, than those required for an actuator all-in-all. The authors propose to construct them based on low power electromechanical actuators with redundancy.

Based on the comparative analysis results of the schemes options for constructing an electromechanical steering actuator with redundancy, three basic schemes were defined for which the preliminary failure rates were calculated.

The results of calculations allow us to consider the basic schemes of a electromechanical actuator with redundancy as an alternative to electro-hydrostatic steering actuators for a primary flight control system.

Lupanchuk V. Y. Navigation cartographic methods development for monitoring robotic complexes positioning in surrounding space. Aerospace MAI Journal, 2018, vol. 25, no 1, pp. 132-142.

The purpose of the study of this scientific article is accuracy improving of monitoring robotic complexes positioning in surrounding space while performing various types of motion.

The subject of the study are methods for joint navigational information processing obtained by on cartographic and instrumental data.

The article analyzes the approach and problems of high-precision positioning of unmanned aerial vehicles and ground-based robotic complexes in the surrounding space.

The initial data of the study are based on formation of local areas of the Earth surface employing cartographic data and instrumental measurements. The article presents the main stages of the methodology for the map high-precision local areas formation by mathematical processing of redundant navigation parameters at the base points.

The methodological approach differs from the known ones by the presence of correlations between the map errors and allows the accuracy increase of navigation parameters determination over the entire area of the local section by 1 m, and by 3 to 5 m at base points.

The studies can find application in various fields:

— when solving problems of high-precision positioning of air and ground robotic complexes in the surrounding space;

— when solving the problems of ensuring and developing the Earth Deformation Control Service within the framework of the Federal System of Seismological Observations, particularly in the highlands of the country;

— rapid creation of a multi-information cartographic basis of various scales with account for correlation dependence of navigation and geophysical information.

Ismagilov F. R., Vavilov V. E., Bekuzin V. I., Aiguzina V. V. Structure selection of synchronous motor with permanent magnets and asynchronous start-up. Aerospace MAI Journal, 2018, vol. 25, no 1, pp. 143-156.

Aerospace industry is in special want of high-efficiency electrical drives (motors), which allow reduce electric energy losses and rise productiveness of equipment. The number of electric drives onboard an aircraft varies from 50 to 220 pieces. With high tech development, the number of electric drives onboard an aircraft will only grow, and insignificant efficiency increase of all electric drives in the aggregate will lead to significant fuel savings. Three-phase induction motors with squirrel-cage rotor are in most common use in aerospace industry as fuel transfer drives. Asynchronous motors with maximum possible energy characteristics possess an efficiency below 80% and a power factor below 0.82. A possible alternative to asynchronous motors are BLCD motors, though their employing as pump drives becomes rather hindered due to cost intensive control system and large weight and size parameters. Another possible alternative to asynchronous motors may be a synchronous motor with permanent incorporated magnets and direct asynchronous start-up. The article is devoted to the analysis of structures of synchronous motors with incorporated permanent magnets and asynchronous start-up, for fuel-transfer pumps drives for the aerospace industry. The analysis was performed by computer simulation in the Ansoft Maxwell software package. The article proves the superiority of such motors over asynchronous motors. The structure of synchronous motor with incorporated permanent magnets and asynchronous start-up, which meet the requirements to fuel transfer drives for aerospace industry, was obtained based on computer simulation. The obtained results can be employed for the design of synchronous motors with permanent incorporated magnets and asynchronous start-up.

Reznikov S. B., Kiselev M. A., Moroshkin Y. V., Mukhin A. A., Kharchenko I. A. Electric power supply system with distributed differential high voltage dc-link and modular-scalable architecture for all-electric aircraft. Aerospace MAI Journal, 2018, vol. 25, no 1, pp. 157-166.

The promising concept of all-electric aircraft free from pneumo- and hydro actua-tors for flight control and stabilizing rotation frequency of main starter-generators supposes significant rise of power supply capacity up to 1.5 MW and more. To ensure high reliability indices and quality of supplied electric energy, the parallel con-nection of supply channels of onboard electric power system should be provided, as well as reversible (bidirectional) interconnection with stand-by low-voltage batteries.

To realize the concept of all-electric aircraft, the article suggests application of the so-called differential higher voltage DC-link with frame grounded averaged-potential (“zero”) wire.

Apart from the well-known benefits of the high DC voltage distribution subsystems, suggested high voltage DC-link has specific benefits, which allow substantiate particular requirements to the power supply systems for domestic all-electric aircraft.

As an example, the article presents the power circuit of the electric power supply combined channel with high voltage DC-link and standby battery based uninterruptible source for combined electric power systems with modular-scalable architecture. It also describes this channels operation. The reviewed structure of a single-phase power supply channel with high voltage DC-link may be recommended as an interrelated group of unified modules of switched mode converters applicable for synthesis of combined power supply systems with modular-scalable architecture and enhanced power supply capacity, and for all-electric aircraft power supply systems in particular.

The article suggest also the combined power supply system with distributed higher voltage DC-link and modular-scalable architecture for all-electric aircraft. Schematic and algorithmic solutions for three types of multifunctional switched mode converters, encompassing all specter of necessary conversions in onboard systems are considered. These solutions allow realize power supply systems with modular-scalable architecture for all-electric aircraft with account for import substitution of power electronics product range. The article presents the example of a simplified combined power supply system with differential higher voltage DC-link of hypothetical all-electric aircraft with four cruise engines and four main starter-generators.

Abramovich B. N., Sychev Y. A., Kuznetsov P. A. Electromecanical complex with high-frequency induction drive for gas-turbine engine. Aerospace MAI Journal, 2018, vol. 25, no 1, pp. 167-179.

The main issue of this article is modeling and reviewing of the possibilities of high-speed induction motors to modernize conventional and newly developed technical solutions.

The performed analysis reveals that by the year 2050 electrical energy production and consumption will practically double relative to 2015. Electric drive is one of its key consumers, and induction motors as the main motors. In this regard, presently the issue of meeting the rapidly developing industry requirements for developing highly effective reliable models capable of operating under conditions of drastically changing load arises.

The prospective arrangement to be modernized with such electric drive is a gearbox of a gas turbine engine. Difficultly controlled and tightly coupled with gearing-system, the pumps can be substituted by their lightweight, small-sized analogs in the form of electric drives. The authors offer two possible structures of modernization of various degree of complexity.

The induction motor modeling is complicated by computing parameters of its equivalent scheme. The article presents the review of the key values hard to calculate, and simplifications description, which were assumed while those parameters computing. The induction squirrel-cage motor with two rotor windings was selected as the basic model.

Two models for studying characteristics of high-speed motors were developed with MATLAB-Simulink. The first model simulates the motor with frequency regulator, and the second one is finished electrically driven gearbox of gas turbine engine aggregates. The problems of harmonic components generation by frequency converters are considered as well.

High frequency motors simulation results were compared to series-produced analog. They demonstrate the superiority of the new models compared to conventional, such as less jitter of the velocity curve, reduced inrush current, faster transients and increased torque. Comparison of variable-frequency control technique advantages with series-produced analogs was performed in the final part. The wider capabilities of the enhanced frequency range are demonstrated.

Shevtsov D. A., Poletaev A. S. Multiphase pulse-width modulators for devices with a multichannel principle of electric power conversion. Aerospace MAI Journal, 2018, vol. 25, no 1, pp. 180-189.

The energy path separation of switched mode electric energy converters into several channels with power switches control in different phases is a promising method for increasing the energy efficiency, reliability, and manufacturability of these devices. Selection of pulse-width modulator control mode for power converting cells while synthesizing the structure of a switched mode power converter with stabilized output voltage is of fundamental importance. Current mode has a number of significant advantages over Voltage mode.

They are as follows:

– better regulation dynamics;

– possibility of simpler overcurrent protection ensuring;

– automatic uniform distribution of currents between power converting cells. The main disadvantage of current mode is the possibility of subharmonic oscillations occurrence in continuous current mode. To ensure subharmonic stability, slope-compensation, or the duty cycle limited within the range of 0–0.5 are applied.

The article proposes three circuit solutions for multiphase basic frequency generator with a duty cycle equal to 0.5 and uniform time shift between phases for multiphase pulse-width modulators in Current mode.

A generator with a number of phases of N, defined as N = 2k, where k is a natural number, can be built employing T-flip-flops. The N-phase generator circuit, requires N–1 triggers. The disadvantage of the scheme

is the limited choice of the possible number of phases. Its advantages consist in realization simplicity and automatic restoration after a failure caused by external jamming.

A generator with any integer number of phases N – k can be realized employing a shift register. N-phase generator circuit requires an N bits register. The circuit is also insensitive to failures.

A generator with an even number of phases N – 2k can also be implemented employing a shift register. To obtain N phases according to this principle l – N/2 register bits are sufficient. The drawback of the last proposed scheme is inability of its automatic recover after a failure.

Egorova Y. B., Davydenko L. V., Chibisova E. V., Shmyrova A. V. The effect of chemical composition and heat treatment on mechanical properties of forgings from a pseudo-ß-titanium alloy. Aerospace MAI Journal, 2018, vol. 25, no 1, pp. 190-201.

The article presents the results of statistical studies of mechanical properties of the deformed semi-finished products from Ti-10V-2Fe-3Al titanium alloy based on analysis of literature, experimental and commercial data, by the “Stadia 7” software package. The effect of reheat temperature for quenching Th, as well as ageing temperature Tag on mechanical properties was evaluated by method of regressive analysis of the published tе data. Equations for computing polymorphic transformation temperature beta-transus temperature βtr and the quantity of primary α-phase, formed while quenching process on the temperatures interval from 700 °C to βtr were obtained:

βtr = 890 + 22,3Al - 13,9V - 8,0Fe,

nα = (0,3 ± 0,02)·( βtr-Тh), %


1608 ingots and die forgings, manufactured by the industrial technology in 2007-2016 were also the subjects of research. All forgings were subjected to thermal treatment, consisted of quenching (763-798°C for three hours followed by water cooling) and ageing (500-515°C for 8 hours followed by air cooling).

The following factors were selected for statistical analysis: alloying elements’ and impurities’ content, beta-transus temperature, alloy structural equivalents in aluminum    and molybdenum, hardening temperature Th and the aging temperature Tag, the mechanical properties (offset yield strength σ0,2, tensile strength σв, elongation δ , reduction of area ψ , fracture toughness K1C). Primary statistical processing and correlation-regression analysis were performed.

Correlations between mechanical properties with deviations of the brand composition and heat treatment modes were established. At the first stage, pairwise correlations between the investigated factors were analyzed. The results of the analysis revealed that each element separately either does not affect, or affects weakly the level of mechanical properties of forgings, which is most likely stipulated by small intervals of their change. The joint action of the elements, which was evaluated by and , appeared to be more significant, the coefficient of multiple correlation was R=0.3-0.5, the fraction of the of the properties variation was γ ≈10-25%. Coefficients of multiple property correlation with quenching temperature and aging temperature were equal to R = 0,3-0,6 depending on the year of production. The joint effect on theproperties of all four factors (, h, Тag) is evaluated by the coefficients R = 0,35-0,67, γ ≈ 12-45 %. The rest of the variation is stipulated by factors that could not be determined based on the data studied. The generalized regression dependence of the tensile strength of Ti-10-2-3 forgings on the chemical composition and heat treatment modes is:

Voronin S. V., Chaplygin K. K. A technique for determining aluminum alloy grains crystallographic orientation in polarized light. Aerospace MAI Journal, 2018, vol. 25, no 1, pp. 202-208.

The assumption on the possibility of employing interference pattern of aluminum alloys surface in polarized light for determining crystallographic orientation of separate grains was put forward. This assumption was tested on the example of aluminum alloy AD1. Optimum modes of electrolytic etching of AD1 alloy, under which the grains' boundaries were sharply defined, and necessary interference pattern of the grained structure was attained, were defined. Electrolytic etching was being performed in a 40% solution of hydrofluoric acid, boric acid and distilled water at 1.7-1.9 A, 100-110 V, and etching time duration of two minutes. It was established that the interference pattern of the sample surface changes with prolonged exposure in the open air. This was due to the oxide film's growth process. Employing literature data on elasticity modulus of aluminum mono crystals depending on crystallographic direction, the article defines the relationship between the grain elasticity modulus and its crystallographic orientation over three directions by the scanning probe microscopy method using NanoScan-3D device. Scanning of the studied section with a size of 128 х 128 µ m was carried out at a speed of 30 µ m/sec.During the scanning process, the signal from the indentation sensor was recorded and processed, resulting in a surface profile map (Zopt). Modulus of elasticity of separate grains was determined by the method of removing the curves of the indenter's supply to the surface of the sample for each grain in the section under study.

While comparing the interference pattern with the distribution of modulus values, it was found that the grains of blue color corresponded to minimum values of modulus of elasticity from 46 to 55 GPa. Maximum values of modulus of elasticity were in the range from 69 to 78 GPa, and corresponded to yellow grains. It was established also that pale orange grains correspond to modulus of elasticity from 55 to 64 GPa. As a result, the assumption was made that the blue grains have a crystallographic direction [100], since they have minimum modulus of elasticity in the array of obtained values. Yellow grains have a crystallographic orientation [111] and maximum modulus of elasticity from the values obtained. Pale orange grains occupied an intermediate position by value of modulus of elasticity, so it was assumed that their crystallographic direction corresponds to [110].

The developed technique is characterized by simplicity, low energy intensity, and less time consuming, in contrast to the methods traditionally used for this purpose. This technique can also be employed to determine the crystallographic orientation of individual grains of other aluminum alloys.

Grishin D. V. Development of effective forms of production process stuffing in aircraft building industry. Aerospace MAI Journal, 2018, vol. 25, no 1, pp. 209-219.

The aviation industry is one of the most high-tech industries not only in the product design and development, but also in the production process posing high requirements on personnel qualifications. The system of qualifications assessment and certification in the aviation industry helps to solve the issue of staffing of the production process due to:

  • Reconcilement of employers' requirements to graduates' qualification;

  • Independent and objective assessment of the qualifications' mastering level;

  • Accreditation of educational programs by employers.

The systems of professional certification in Europe and the United States has been functioning since the 1980s. In 2007 a large-scale activities on creation of National system of qualification started in Russia under the auspices of the RSPP. Since 2014, this work was being performed on the ground of the National Council on Professional Qualification under the President of the Russian Federation. The Council for professional qualifications in the field of aviation was established in 2016, it included representatives of major employers and their associations, public authorities and educational institutions.

At the first meeting of the Council, it was decided to consider the possibilities of employing and adaptation of the project sectoral qualifications frameworks in mechanical engineering for aircraft industry. For this purpose there are all preconditions, since the enterprises need the skilled workers in the first place, and it was for them that the Sectorial Council on Machine Building develops qualification requirements.

Based on Federal law No. 238-FZ “On independent qualification assessment”, 108 organizations, getting the status of centers of assessment of qualifications were selected, in which more than five thousand people have already confirmed their skills.

Currently CTCS in the field of aviation has not been not established, however the need for its creating is urgent. Though MAI cannot act as the CSC organizer, it is likely expedient to enlist the services of the MAI teaching staff to its activities In particular, the administrative tasks of CTCS can be transferred to MAI, which has extensive experience of performing similar procedures: admission company, examinations, etc.

The proposed system will help the industry economically, as well as strengthen the ties between the labor market and educational sphere.

Manvelidze A. B. Defining tнe demand for passenger airplanes in conditions of market saturation ву foreign-made aircraft. Aerospace MAI Journal, 2018, vol. 25, no 1, pp. 220-232.

The problem of Russian air companies' transition from employing foreign-made passenger aircraft to domestic ones is under consideration. The article analyzes the status of passenger aircraft being under operation or being ordered for the future. It also defines the aircraft ownership i.e. financial leasing, or operational leasing or airline's property. Мost attention is payed to the study of an aircraft operational leasing, since regulation of aircraft park being in temporary service will allow release market niches for domestic built aircraft.

The article presents a methodical approach, allowing assess variants of freight capacities volume formation based on statistics monitoring to implement new airplanes of domestic manufacturing.

Analysis of the rules of statistical accounting applied by US airlines led to the idea of the above saic proposed methodology. The Bureau of Transportation Statistics (BTS) of the USA publishes monthly the detailed airlines reports on distribution of aircraft types by airlines, reflecting distance factors, the number of planned and actually performed flights, passenger traffic, cargo and mail, available seat-miles, passenger-miles, ton-miles, flight hours and aviation fuel consumption. Based on the detailed data, brief reports on aircraft employing are being compiled and can be sent to aviation organizations, such as ICAO. The detailed presentation of information allows perform studies on modernization of aircraft fleet under operation adequately and without extra resources.

The air transport of Russian Federation publishes brief statistical forms on aircraft availability and usage (32 civil aviation and 33 civil aviation). To obtain detailed data on performed air service by airplanes of airlines the data on full schedules (SRS Analyser) and passenger transportation along the routes are being integrated.

The calculations simulating the workout resources of an aircraft in use are performed using the network modelled in such a way. In the longer run, the demand for airlifts rises, the aircraft in service drops out, and a niche of free seat-volumes for new aircraft implementation appears.

The source of information is the Transport Clearing House statistical database on aircraft fleet at disposal and employing this fleet by airlines and industry at large, transportation between pairs of cities, as well as international databases. The Flight Global database is used to analyze the state of passenger airplanes' park. It gives a comprehensive idea of air transporters' airplanes under operation and aircraft building industry perspectives. The airplane schedule and freight capacity were accepted according to SRS Analyser database. In calculation for perspective, the United Aircraft Corporation plans on aircraft building up to 2037 were accounted for. The market niche formation of new aircraft implementation is affected mainly by the demand on passenger aircraft seats as a whole and by segments; terms of passenger airplane disposal; signed contracts of airlines on the delivery of foreign-made aircraft; delivery plans of domestic manufactured aircraft.

Efimova N. S., Volenko A. K., Kanashova Y. G. High-tech production managing with account for requirements of economic security (on the example of aircraft building). Aerospace MAI Journal, 2018, vol. 25, no 1, pp. 233-242.

Presently it is necessary to develop highly effective assessment of high-tech enterprises economic security level as each aircraft building enterprise needs integrated self-concept of its production-commercial and financial-economic activities.

The main objective of assessment of the level of production activity economic security is assessment of risks on integrated system of indicators accounting for specific branch features at the enterprises of high-tech industries. The authors recommend employ the hi-tech enterprises' risks assessment based on a qualitative or interval method. Internal self-concept of state of production of aeronautical engineering development, and assessment of production development dynamics of the enterprise should be the main task of high-tech production monitoring.

The internal self-assessment of a production condition of creation of the aircraft equipment and assessment of dynamics of development of production of the enterprise have to be the main objective of monitoring of hi-tech enterprise. For this purpose, it is necessary to employ the technique, which would describe the main approaches and basic self-concept procedures for the risks of production departments of high-tech enterprises.

The article considered and developed managing system for high-tech production with account for economic security requirements, which will allow ensuring the raise of high-tech products competitiveness.It suggests assessment indicators for probability category of economic security factors coming-in, risks register, and risks map in high-tech branches of the industry.

Developing the system of economic security of production activity in aircraft building will allow forecast the aftermath of internal and external hazards on both production processes, and high-tech enterprises' activities at large. Implementation of the above said procedures at the aircraft building enterprises will allow ensure the necessary level of economic security, as well as optimize production processes at the enterprises of high-tech industry branches.

Kozlov A. E. Export potential of an aircraft building enterprise: development trends and predictive modeling (on the example of Progress Arsenyev Aviation Company). Aerospace MAI Journal, 2018, vol. 25, no 1, pp. 243-255.

Enterprises are constantly facing competition, both at domestic and foreign market. The competition promotes the development of export, but the risks associated with an unsatisfactory estimation of activity of an enterprise and improper planning may inflict damage to the company.

To consider the issue of products export one can employ the following method of export potential assessment to promote products and services at foreign markets:

  1. Assess the popularity of manufactured products or services at the domestic market. If they are successfully sold in the local market, they will be probably in demand abroad, at least at the markets of the countries with similar socio-economic conditions and needs;

  2. Evaluate the unique or most important features of the produced goods and services. If they are hard to be reproduced abroad, there is a possibility the company will enjoy the success, as unique goods do not face the severe competition and the demand for them is high.

Since the Holding Company “Helicopters of Russia” occupies the leading positions in military-industrial complex development, the article is devoted to the study of export potential of one of its enterprises, namely, Arseniev aircraft company “Progress”, which, plans to export military equipment in 2017.

To avoid the above said risks, the company was proposed to employ the model developed by the author for export potential predicting, based of complex evaluation of enterprises. The model is based on evaluation of the basic technical and economic, accounting and financial indicators, as well as indicators of the enterprise's management and its scientific and technical potential. It includes also statistics of qualitative and quantitative structure of personnel, gender and age structure, work experience of engineering personnel, managers, fellow laborers, workers, as well as information on the research work and capabilities of research projects performing.

Summarizing the results on the above listed factors, a model of the multiple regression describing the dependence of the export potential of human resources was built, which was subsequently automated with a software application.

As the result, the program was:

– effective, since the payback period was 45 days at a low annual economic effect;

– universal, since applying this tool is possible for other enterprises (not only the aircraft industry);

– easy-to-use, since the users need only necessary information on the enterprise dynamics, and, according to the theory of mathematical modeling and regression analysis, the more data will be used, the higher the model adequacy and forecast accuracy will be.

Abashev V. M., Demidov A. S., Eremkin I. V., Kiktev S. I., Khomovskii Y. N. Temperature stresses in a cylindrical shell of carbon fibers and the contact problem of heat transfer. Aerospace MAI Journal, 2017, vol. 24, no 4, pp. 7-13.

Cylindrical shells are the most common structural elements of rocket engines. When loading by the temperature gradient in radial direction radial temperature stresses occur in them. Such stresses in carbon-carbon shells can be rather dangerous notwithstanding that they are much smaller than the circumferential and axial ones. Moreover, they substantially depend on the thermal conductivity of the carbon fiber material and the shell structure.

The article suggests the equation for the structural thermal conductivity (contact thermal exchange) evaluation of a cylindrical shell in radial direction. When calculating with the equation the carbon fibers' roughness was not accounted for due to the presence of pre-preg matrix, and the shell was divided conditionally through-the-thickness into several layers. The contact forces acting on the fibers were determined based on a primary evaluation of the temperature stresses. The results of the shells' made of carbon fibers calculations with a diameter of 0.02, 0.05, 0.2, 0.5, and 1 mm are presented in the form of tables and graphical dependencies. It is shown, that the elasticity modulus of the first genus of carbon fibers' surface layers can be accounted for in the calculations. It was revealed, that in shells with internal warming-up the specific pressures at the areas of contact spots of the adjoining fibers could reach several tens of kilograms per square millimeter. There is a risk of the carbon fibers structure stratification in the shells with the external warming-up. Thus, we recommend conduct tensile or bending tests with small-sized samples, cut from the shell in radial direction. Tests of such samples should be carried out according to the methodological instructions.

Bezuevskii A. V., Ishmuratov F. Z. Quasi-static deformations effect on aeroelasticity characteristics of an aircraft with high aspect ratio wing. Aerospace MAI Journal, 2017, vol. 24, no 4, pp. .

One of the ways to increase the aerodynamic quality of modern and prospective aircraft consists in wing aspect ratio increasing. Such increasing leads to the occurrence of various new aspects of the structure loading, strength and aeroelasticity. One of these aspects is increasing of the wing flexibility, and as a consequence, possible in-flight structures deformations effect on aeroelasticity characteristics.

The paper presents a review of publications on the deformation effect on various aeroelasticity characteristics. It suggests and substantiates a computational method for studying the effect of quasi-static deformations of the wing on static and dynamic aeroelasticity of an aircraft. This method is based on automated generation of a set of aircraft computational models using the Ritz polynomial method. The paper presents the examples of a wing in-flight deformations effect on characteristics of static elasticity, frequency and shape of elastic vibrations, and flutter characteristics.

The results of the developed method application for aircraft of various configuration allowed establishing the main regularities of the effect of structure's deformation on aeroelasticity characteristics.

The effect of in-flight deformations on the characteristics of static aeroelasticity and load is determined by: 1) effective wing span decrease; 2) aerodynamic forces direction changing; 3) increase of the effective dihedral angle. Characteristics of longitudinal motion can reduce by 5-6%, while characteristics of lateral motion can increase or decrease by 5-15%.

The dynamic aeroelasticity characteristics change is determined mainly by the increase in the interaction of torsional oscillations of a wing with bending vibrations in the chord plane. For the unmanned aerial vehicles with a wing of extremely high aspect ratio, this effect can lead to a significant decrease in flutter speed (up to 30-50%). For modern airliners, the decrease in flutter speed due to the in-flight deformation does not exceed a few percent and lies within the accuracy of numerical methods.

An important feature of the method is its integration into a multidisciplinary design complex ARGON, validated while solving aeroelasticity problems in many practical applications.

Marakhtanov M. K., Pil'nikov A. V. On solar electric propulsion system application possibility for low-orbit small spacecraft. Aerospace MAI Journal, 2017, vol. 24, no 4, pp. 26-39.

A comprehensive material on operation of spacecraft with solar electric propulsion systems is accumulated by now. The latter are designed for spacecraft correction on both geostationary and circular low earth orbits.

At the same time, there is a tendency to developing of small spacecraft of various purposes, such as scientific, communication, Earth remote probing, navigation, hydro-meteorological etc. operating on low circular orbits with the height within the rage of 180–280 km. Such spacecraft are relatively cheap and possess the mass of 10 to 500 kg. However, such indicator as minimum orbit height, its relationship with the spacecraft weight and size, as well as parameters of its engine unit remain undetermined. System analysis and experimental data on spacecraft with solar electric propulsion systems, operating at the height of 140–280 km are practically inaccessible.

The paper considers the problem of small spacecraft transition fr om a higher circular orbit to a lower one. As far as the Earth atmosphere gradually transfers to vacuum, the aerodynamic drag force grows while a spacecraft descent. We suggest surmounting this force through the electric jet engine thrust power. It is obvious that while the spacecraft descent the aerodynamic drag grows, and such parameters as thrust power and electric jet engine power should be increased continuously. At large the problem becomes dynamical. Besides, the main cause of the orbit height limiting will be the drag force of the solar battery. Thus, the minimum orbit height hmin below which the spacecraft, equipped with the solar electric jet engine cannot exist, is limited by the spacecraft drag force due to the solar battery. At the lower altitude the battery's drag force will be greater than the electric rocket engine thrust force.

For the spacecraft motion analysis, we assumed that the solar battery takes the shape of an autonomous panel with rotation angle control to the sun radiation direction. The power flux density or the solar radiation at the Earth orbit is Q = 1400 W/m2 (solar constant). The efficiency of photoelectric transducers based on a three-stage gallium arsenide (GaAs) equals to ηSB= 0.22. The solar battery specific power is α = 308 W/m2. If the solar battery plane is oriented normally to the orbital movement velocity vector the drag factor equals to CSB = 2.15, and if it is oriented along this vector it equals CSB = 0.15.

If an ion thruster is used as an electric jet engine its specific impulse is assumed as ΙSP = 4500 s, and its efficiency equals to ηT= 0.7. In case of plasma engine of the SPT type ISP = 1700 s and ηT= 0.55 correspondingly.

The lower limits of the orbit altitude hmin = 200 for the solar electric jet system with the ion engine, and hmin = 180 km with plasma engine of the SPT type were established by the results of the performed analysis. The upper lim it of the altitudes descending from which requires continuous build-up of the electric jet system solar battery area to overcome the atmospheric aerodynamic drag is the altitude of hmax = 260 km.

The paper demonstrates that for a spacecraft continuous exploitation at the latitude of 180–260 km application of solar electric jet engine and atmospheric gas as a working agent is possible. Application of high- frequency ion engine of 4.4–5 kW is expedient for the propulsion installation such kind. With the specified power and solar battery weight, the weight of electric jet propulsion installation will be no less than 90–100 kg, and the total minimum weight of the spacecraft will be no less than 500–600 kg.

Kuz'michev V. S., Tkachenko A. Y., Filinov E. P. Effect of turbojet engine dimensionality on optimal working process parameters selection. Aerospace MAI Journal, 2017, vol. 24, no 4, pp. 40-45.

With turbojet engine thrust reduction, its small size begins affecting the effectiveness on its elements. Lower airflow rate results in blades size decrease and relative radial clearance increase. It affects the efficiency of axial turbo-machines. Due to this, radial and centrifugal turbo-machines become more effective at small thrust values. The main goal of this study consists in determining the most effective structural scheme of a turbojet engine for the thrust range from 0.1 kN to 100 kN. The problem was solved by performing the engine multi-criteria optimization employing ASTRA CAD, developed in Samara National Research University. The total weight of a power plant and fuel, as well as specific fuel consumption were selected as performance criteria. The optimized variables are the gas temperature prior to the turbine, and total pressure ratio. According to the optimization results the following inferences were drawn. With optimization of the engines with the thrust, lower than 25 kN, corrections on their small-size should be accounted for. With the engine thrust decrease, the optimal parameters of the working process are decreasing either, and the regions of compromises are contracting. The axial compressor is optimal for the thrust of 7 kN and higher, and with thrust decrease up to 1.3 kN, compressor of axial-centrifugal type becomes more appropriate. The axial turbine is effective up to 0.7 kN thrust value, and radial turbine is effective for small engines with lower thrust.

Ezrokhi Y. A., Khoreva E. A., Kizeev I. S. Determining the thrust of an aircraft gas turbine engine with flows mixing under condition of non-uniformity of total pressure at the engine inlet. Aerospace MAI Journal, 2017, vol. 24, no 4, pp. 46-51.

The article deals with the flight thrust determining method of a bypass engine with flows mixing in the presence of a non-uniform total pressure field at its inlet. The non-uniformity impact is taken into account for both air consumption due conventionally averaged total pressure at the inlet, and the specific thrust due to the overall pressure level reduction along the engine passage, and, respectively, the available differential pressure in the jet nozzle.

Earlier, the authors developed and patented the engine thrust determining method allowing evaluate its thrust while in flight under condition of the uniform flow at its inlet according to the measured operating conditions and external environment parameters. The presented work extends this simplified engineering method to the real case of a non-uniform total pressure field at the engine inlet. Moreover, it employs corrected values of the total pressure along the engine passage to compute the thrust.

Thus obtained, the value of the flight thrust can be used in both automated control system for its possible in-flight correction, such as partial or full flight thrust value restoration, and the complex engine diagnostics system to evaluate its deterioration rate and deterioration in performance of its separate parts and elements.

Calculated evaluations performed according the developed method with account for typical input total pressure non-uniformity revealed that the expected thrust fall will be of 8.9%, with about 8% herewith due to the air consumption reduction, and the rest is due to specific thrust decrease.

Belyaev I. V., Valiev A. V., Moshkov P. A., Ostrikov N. N. Studying the PTERO-G0 unmanned flying vehicles acoustic characteristics in AK-2 unechoic chamber. Aerospace MAI Journal, 2017, vol. 24, no 4, pp. 52-62.

Recently, more and more attention is paid to the problem of ensuring unmanned flying vehicles (UAV's) invisibility in various frequency ranges due to the wide application of the systems with small sized UAVs for solving special assignment tasks. To ensure the UAV's invisibility in the audible frequency range at the specified distance from the observer in conditions of known terrain of application, the qualitative data on the UAV acoustic characteristics is required.

The experimental study of the small sized UAV's “Ptero-G0” acoustic characteristics was performed within the framework of the presented work. The UAV's power plant consisted of a single-cylinder gasoline internal combustion engine (ICE) and a small sized two-blade propeller with the fixed pitch. The acoustic tests were performed in TsAGI unechoic chamber AK-2.

The following main results were obtained as a result of experimental researches.

  1. Energy, spectral and spatial characteristics of acoustic fields of a small sized propeller and single-cylinder four-cycle gasoline engine were obtained.

  2. The small sized propellers diameter effect on UAV's noise and signature characteristics was studied. Recommendations on acoustic signature reduction of the UAV “Ptero-G0” were elaborated. These recommendations were implemented and accounted for by the “AFM-Servers” company while developing new flying vehicles.

  3. It was demonstrated that a cowl mounting on the engine without both vibration and acoustic insulation could lead to significant noise increase of the power plant.

  4. The possibility of employing the empirical model while solving the problem of a single-cylinder four-stroke gasoline engine's noise evaluation was demonstrated.

Within the framework of the subject's of UAV acoustics development, the authors are planning to proceed this work in the following main trends.

  1. Studying the effect of the power plant's noise shielding by the airframe elements on UAVs noise and signature characteristics.

  2. Studying the noise caused by the UAV airframe flow-around.

  3. Development of semi-empirical model of the small-sized propeller's noise.

  4. Implementation of the existing computation methods for audibility and signature boundaries into practice of noise indices evaluation of various UAV's types.

  5. Software development for UAV flight trajectories' plotting under known weather and landscape conditions without the ability to detect it both by ear and with acoustic location finder.

Kruglov K. I. Numerical calculation of temperature distribution in power sypply unit of a radio frequency ion thruster. Aerospace MAI Journal, 2017, vol. 24, no 4, pp. 63-69.

Thermal flows emitted by power sypply unit (PSU) components lead to their heating, which, in its turn, may lead to changes of their operating characteristics up to their failure. Thus, the temperatures of these components should be maintained within the ranges ensuring maintenance of their operating characteristics. For this purpose a preliminary simulation of thermal processes in the PSU housing was performed.

The article presents a model for temperature distribution calculation in separate components of a radio frequency ion thruster's structure. These calculations were performed using ANSYS bundled software.

Due to the negligible effect of thermal flows from the thruster unit on the thermal state of PSU, thermal simulations of the thruster unit and the PSU were performed separately. The aluminum thermostatically controlled mounting flange, located above the gas-discharge chamber presents the boundary.

All PSU's structural elements in the computer model are simulated as simple geometric forms, such as cylinders or parallelepipeds with appropriate geometrical dimensions.

The total heat emission in the PSU unit from all its constituting elements is taken equal to 66.4 W. This value corresponds to the operating mode of a low-power radio frequency ion thruster.

To intensify maximally the heat removal by radiation, the emissivity factor of 0.9 was attributed to all external surfaces of the PSU unit components.

To maximize radiant heat removal, the outer surfaces of elements of PSU were modeled with the emissivity of 0.9. To increase the conductive heat exchange, a partial PSU components' potting (gersil) was performed.

The calculation used the real thermal contact between adjacent surfaces with corresponding values of thermal junction resistance. A series of calculations was conducted for various the compound's thermal resistance values from 1.2 to 2.7 W/(m·K).

The figure below shows the dependence of the temperature of the most heated component of the structure under various thermal conductivity coefficients of the compound.

The requirements for the thermal conductivity of the compound for filling the PSU's PCBs were determined.

When using materials with thermal conductivity exceeding 1.7 W/(m·K), it is possible to ensure the permissible temperature of electronic components at a temperature of the mounting flange reaching 50°C.

The developed physico-mathematical model can be employed at the stage of the ion thruster preliminary designing.

Chubov P. N., Saevets P. A., Rumyantsev А. V. Thermal calculation of the SPT-50 stationary plasma thruster. Aerospace MAI Journal, 2017, vol. 24, no 4, pp. 70-79.

Development of the SPT-50 thermal model, thermal calculations and study of the model sensitivity to changes and to various combinations of internal and external heat exchange parameters was carried out with account for the requirements of the Thermica software applications package (SAP) based on employing of isothermal elements method. The thermal model under development consists of 130 elements. The radiation couplings for the SPT-50 anode unit's thermal model were computed employing Thermica V4 SAP. To obtain the information on the thruster thermal state during thermal vacuum tests (TVT) it was equipped with temperature detectors, installed on the thruster in places with enough access to the surfaces for contact welding, glue and other ways of mounting. The SPT's thermal balance thermal vacuum and thermal cycling tests were performed. The thermal model correction with the testing results was realized by thermal calculations employing the developed thermal model. The calculations did not account for convective heat exchange (imitation of vacuum). The ambient temperature was set the same as the during testing, and SPT's optical and heat emission properties were set according to the operating mode during TVT.

The developed thruster thermal model, updated by testing results tests, allows analyze thermal processes inside the thruster in the places where installation of thermocouples is impossible. After the SPT-50 thermal model correction one can define the critical design elements, thermally affected by the thruster. Based on the thermal calculation results, the element of wire with critical temperature level has been defined, and this value approached maximum temperature value of 220°C. To decrease the wire temperature, we increase the wire core section area to enhance the heat sink from the wires critical element. The calculations revealed that the temperature of the SPT's critical elements does not exceed maximum admissible working temperature. It confirms correctness of the approaches to selection of thermal design and parameters of the thermal regulation system of the SPT-50 anode unit. The presented thermal model of the SPT-50 anode unit can be employed for developing other options of thermal and mounting interfaces for other discharge and magnetic thruster operating parameters.

Ermoshkin Y. M., Galaiko V. N., Kim V. P., Kochev Y. V., Merkur'ev D. V., Ostapushenko A. A., Popov G. A., Smirnov P. G., Shilov E. A., Yakimov E. N. Specifics of transients in the discharge circuit during the SPT-140D plasma engine starting. Aerospace MAI Journal, 2017, vol. 24, no 4, pp. 80-88.

The article presents the results of transients' in discharge circuit studies during SPT-140D plasma thruster starting while its operation together with power processing unit (PPU). SPT-140D is an electric thruster developed by the Design Bureau “Fakel”. This thruster was running on Xenon with the discharge voltage of 300 V and power of 4.5 kW, ensuring reactive thrust 280-290 mN and 1750 s specific impulse of thrust. At present, this thruster is ready for flight application for spacecraft motion control. The PPU unit was developed and manufactured by the Scientific and Production Center “Polus”. Since the main discharge is one of the powerful PPU loads, the main attention was payed to the study of transients in the power supply circuit of the main discharge. The obtained data was used for the development of imitation model of the named transients and electric imitator of the thruster for off-line PPU optimization and testing without the thruster. In addition, the information on the specifics of the thruster operation was obtained. The most interesting among them are the following:

  1. In the course of the thruster starting, after the main discharge ignition by the discharge voltage increasing with the rate of about 1 V/ms, the main discharge could ignite by various discharge voltages. Though after the discharge ignition its parameters during various start-ups vary according to one and the same averaged dynamic volt-ampere characteristic, close to the “static” characteristic obtained with slow voltage changing.

  2. Various oscillation modes of the discharge parameters were revealed, arousing at the various stages of discharge voltage variation, and changing drastically with the small variation of the discharge voltage. It allows evaluate the increment of their build-up.

  3. After reaching the nominal discharge parameters, the dominating discharge current oscillation mode frequency is 15-20 kHz. After an hour of continuous operation it reaches the value of 27-27 kHz, and its further variation is insignificant. It can be explained by the discharge chamber heating resulting in Xenon atoms velocity increase, decrease of their drift time through the ionization layer and acceleration leading to the frequency increase according to ionization-drifting oscillations excitation model.

Thus, the employed methodology of the study is useful also for conducting physical research of the processes in the thruster.

Shorr B. F., Melnikova G. V., Serebryakov N. N., Shadrin D. V., Bortnikov A. D. Calculation and experimental evaluation of damper efficiency for decreasing vibratory stresses in turbine rotor blades. Aerospace MAI Journal, 2017, vol. 24, no 4, pp. 89-99.

The subjects for study are dampers of various masses installed under the platforms of turbine blades.

The research issue is prevention of turbine blades failures caused by higher level of variable stresses.

The goal of the work consists in experimental and computational definition of effectiveness of shock-absorbing insertions' masses (3.3 g, 4.7 g and 5.8 g) for variable stresses reduction in the full-size turbine wheel.

The methodology of the work includes two trends: computation and experimental. The computation trend is based on modeling the damper using MSC.Nastran contact elements and estimating the reduction of vibratory stresses, by integrating the equation of motion in the time-domain employing the standard non-linear integration procedure by the Newmark method. The effect of the insertion on vibration frequencies of the blade was also studied. The experimental trend is based on a comparative analysis of the amplitudes of vibratory stresses in the blades both with installed damper and without them. Tests are performed on the CIAM bench test (manufacturer is Test Devices company). The turbine wheel is assembled for testing in a special way: one sector of the wheel is damper free, and the rest three sectors were equipped with dampers of various masses. The blades were prepared with strain gages, and in each sector the blades with maximum response to external excitation from the air supplied to the test chamber were selected. Tests were carried out for an unheated wheel.

The calculations revealed that the most effective reduction of vibratory stresses in the blade occurs when the holddown pressure of the damper to the bottom surface of the blade platform are 200–800 N. Such forces for damper mass of 3.3 g were caused by centrifugal forces at rotational speeds of the wheel in the range 35–70 % of the maximum rotational speed; this range is 29–58% for the damper of 4.7 g, and for the damper of 5.8 g, it is 26–51%. The affect of dampers weighing 4.7 g and 5.8 g is ineffective, starting, respectively, from 90% and 82% of the maximum rotational speed. According to calculations, the damper with mass of 3.3 g allows reduce the vibratory stresses by 22% at a resonant mode at the 87% of maximal rotational speed.

The tests revealed that, in comparison with damper and without damper, the blade frequency with shock absorber of 3.3 g increased by also 16%, and the oscillations' amplitude decreased by 25%. This correlates satisfactorily with the computation data.

Conclusions were drawn that the calculated and experimental results in these studies showed, in general, a satisfactory agreement with respect to both the reduction of vibratory stresses and the change in the resonant frequency when a damper was installed. Some discrepancy between the calculated and experimental data on the effect on the vibratory stresses of “heavy” inserts of 4.7 g and 5.8 g may occur due to the assumptions in calculations, as well as to the errors in the experiment and processing of the test results. To evaluate the effect of the amplitudes of vibratory stresses in the blade without damper, stiffness and mass of the damper, as well as friction coefficients on the effectiveness of the damper to reduce the vibratory stresses in the blade, additional experimental and calculated studies are required.

Bogdanov V. I. Research on realization of pulsating working processes in jet engines. Aerospace MAI Journal, 2017, vol. 24, no 4, pp. 100-109.

The article presents the results of the study in elaboration of scientific discovery No 314 “Phenomenon of the abnormal high growth of thrust in the ejector process with pulsating active stream” performed in MAI. The possibility of pulsating jet impulse increase through ejectorless addition of gas mass both from the external atmospheric environment and used up gas was shown experimentally. It increases the meaningfulness of the discovery. The physics of the process of the used up gas mass in pulsating stream is based the well-known phenomenon of wave interaction of cyclic masses with various velocities of front and tail parts.

Calculating and experimental studies substantiated the capability of creating a nozzle with the spherical resonator-thrust amplifier for air-breathing jet engine with stationary fuel combustion. The mechanism of gas masses adding in oscillating process is shown. The thrust amplification at certain gas-dynamic and geometrical relationships herewith can make 1.5 and more.

Carrying out of experimental studies on a vacuum bench (imitation of space conditions) has confirmed effective exhaust gas mass addition that opens new capabilities for increasing the thrust efficiency of space jet engines. According the test results, the constructive recommendations on the improvement of working process are given.

The results of computation and design working out of implementation of the obtained effects in the nozzle with the resonator of an optimum configuration for conventional air-breathing jet engine without its mass and dimensions characteristics derating present a great practical interest. Conditions and recommendations on calculation are given.

Possible perspective trends of further studies on implementation of the obtained effects of thrust increase at the expense of exhaust gas mass addition in liquid-propellant rocket engine and solid-propellant rocket engine with spin detonation fuel combustion, as well as in a gas turbine engine are determined.

Sha M. ., Agul'nik A. В., Yakovlev A. A. The effect of the computational mesh while mathematical modeling of the inflow of a subsonic flow onto the profile of a perspective blade with a deflectable trailing edge in a three-dimensional setup. Aerospace MAI Journal, 2017, vol. 24, no 4, pp. 110-121.

In the last decade, much attention has been paid to the studies conducted in the interests of mathematical modeling methods developing in 3D setup. It requires a detailed study of various computational meshes constructing methods and their effect on the obtained results.

The problem of the aerodynamic characteristics computation of the of a perspective blade with deflectable trailing edge profile is important for both the development of wind turbine blades, compressor design for advanced gas turbine engines, and aircraft structures.

The effect of the computational mesh is studied while mathematical modeling of the inflow of a subsonic flow onto the profiles of a perspective blade with a deviating trailing edge. Verification, the convergence and correctness checkup of the solutions obtained, as well as verification on tasks having reliable and detailed enough solutions are necessary.

The objectives of this article are as follows: determining the accuracy of the numerical solution of the aerodynamic profile of the perspective blade with the deflected trailing edge, and testing the computational mesh with the potential to achieve industrial applicability. The feature in common is the use of wall-adjacent blocks adapted to geometry, applying herewith various approaches for their coupling with the external mesh. Analysis of the solvers application employing the Cartesian mesh reveals also the necessity of constructing mesh layers adapted to the surface of the body.

Analysis of existing designs allows us to draw the following conclusions. A simple deflectable trailing edge increases the lifting force by increasing the curvature of the profile. This increases the pressure on the lower surface of the profile, as well as increases its load-bearing properties.

A mathematical model of the aerodynamic processes occurring on the profile surface of a perspective blade from the back deflected edge while its on flowing by a subsonic flow is suggested.

An acceptable correlation of the results of the calculations made using structured and hybrid meshes circuits was obtained. Analysis of the results of numerical simulation employing various meshes revealed that application the meshes under consideration considered allows obtain close results. The structured meshes applied herewith consume less computation time. Hence, we will use the structured meshes as the best way to solve the problem.

Thus, the proposed mathematical model and the first method of developing the mesh can be applied to determine the numerical solution accuracy of the problem of flow past the aerodynamic profile of a perspective blade or wing with a deflectable trailing edge, as well as the mesh testing

Abdulov R. N., Asadov H. G. Spectrozonal method for detection and optimal control of low-altitude rockets through the exhaust plume of a solid jet engine. Aerospace MAI Journal, 2017, vol. 24, no 4, pp. 122-128.

The problem of detection and control of rockets' launching and flying is topical from the viewpoint of functioning safety of various ground and aerial objects of both military and general assignment. At present, significant attention is paid to identification and tracking optimization of low-speed point objects' of various purposes. A method for spectrozonal detection and control of low-altitude rockets through the exhaust plume of a solid jet engine was developed and theoretically confirmed. The authors formulated new spectrozonal features for detection of launched low-altitude rockets, based on the well-known experimental results related to the study of spectral emission of rocket engines plume. To detect and control the low-altitude rockets a new spectrozonal feature, possessing experimental property useful for applying for on both axial and radial directions was formulated. The issues of identification optimization of low-speed low-altitude point objects under variable atmospheric conditions were also considered. The general mathematical problem of optimization of the entire cycle of optimization was formulated and solved. Its gist consists in achieving the maximum possible value of the averaged signal received from the object by the infrared identifier, through the accepted model of variation of atmosphere optical thickness. The article demonstrates that as applied to subsonic flying objects, such as cruise missiles, it is necessary to ensure direct proportion between atmosphere optical thickness and a certain time index. It should be considered herein that the minute scale of atmosphere optical density can be easily controlled, and it presents the result of rapid weather conditions changes due to natural or anthropogenic factors. A certain increase of the value of the functional, obtained by the performed optimization, can be interpreted as a possibility of a certain shift of the total time interval to the left. On this basis, the solution of the formulated optimization problem points to the possibility of realization of the much safe mode of an object detection and identification under revealed.

Some increase of accepted target functional caused by carried out optimization can be interpreted as possibility for some shift of whole time interval to left. Solution of formulated optimization task indicates the possibility of for safe regime for detection and identification of object upon revealed advantageous atmospheric conditions.

Kirsanov A. P. Kinematic properties of aircraft concealed motion trajectory in detection zone of the onboard doppler radar. Aerospace MAI Journal, 2017, vol. 24, no 4, pp. 129-136.

Onboard radars operating in pulse Doppler mode possess characteristic feature in the detection zone. This feature lies in the fact that at each point of the detection zone the aircraft has the sector of directions. Moving along these directions it, cannot be detected by the onboard Doppler radar. This sector is named as the sector of an aircraft concealed motion directions. Due to these features there are concealed trajectories, moving along which the aircraft becomes undetectable by the onboard Doppler radar, such as an aircraft airborne early warning radar (AEWR). Most of these concealed trajectories are curvilinear with variable curvature. The article is devoted to the study of the aircraft concealed movement trajectories curvature in the onboard Doppler radar detection zone. The study of the aircraft concealed movement trajectory curvature in the AEWR detection zone was carried out to evaluate the possibilities of flying over such trajectories with account for and aircraft maneuvering characteristics. The results of the study led to obtaining the equation for calculation the curvature of any concealed trajectory in any of its points. The equation allowing determine the shape and size of a region in which the movement over the concealed trajectory is impossible due to the fact that normal overload exceeds maximum aircraft operating overload. It was established that for the valid parameters of an aircraft movement and AEWR aircraft this region is located within the circle with radius not exceeding 10 km with its center coinciding the location of the AEWR aircraft. The region, where aircraft high manoeuvrability is required, presents utterly small portion of the detection zone of AEWR aircraft. Thus, the aircraft concealed movement is possible over concealed trajectory practically in the entire detection zone.

Kornilov V. A., Sinyavskaya Y. A. Parametrical synthesis of actuating mechanisms with dc motors. Aerospace MAI Journal, 2017, vol. 24, no 4, pp. 137-142.

One of the main problems of automatic control theory can be formulated as optimal functional links' forming between information and energy. The basic principle while control systems design consists in designing such systems, which are able to transmit or convert information with specified timing and phasing-in characteristics under condition of power consumption minimization for the given control law realization.

The main power consumption relates to actuating mechanisms while synchronous transmission fr om control system to the control object with concurrent increase of the energy level. The energy level limitations in actuating mechanism affect significantly such dynamic characteristics as stability, accuracy and noise reduction.

The problem of parametrical synthesis of the rudder servo drive actuating mechanism for the UAV's aerodynamic control system is interpreted as a system design optimization problem. The quality criterion of parametrical optimization problem is maximum effective power delivered to the control object from the power source, necessary to fulfill the required, most tough from the power consumption view point, motion laws of the control object (aerodynamic rudder) under specified parameters of aerodynamic load. Graph-analytic solution of the problem is based on plotting the dependencies Nmax(F), wh ere Nmax is the maximum effective power value; F = Mmax/Wmax is the robustness value of the actuating mechanism mechanical characteristic; Mmax and Wmax are the maximum torque and maximum speed of the actuating mechanism.

These dependencies allow define the optimal parameters of the actuating mechanism ensuring the fulfillment of all control object's required laws of motion, provided the minimized energy consumption for their realization.

Ismagilov F. R., Vavilov V. E. On eddy-currents losses determination in permanent magnets of high-speed electromechanical energy converters. Aerospace MAI Journal, 2017, vol. 24, no 4, pp. 143-150.

The industry demand for high-speed electric motors with rotation frequency of 48,000 rpm to 120,000 rpm and power of 5 to 250 kW increases from year to year. A number of technological problems exists herewith, which retards the high-speed electric motors market growth. These problems relate to the issues of their output voltage stabilization, which are solved by employing static converter, stator back magnetizing, or rectifier; ensuring bearing assembly reliability, which are solved by employing non-contact bearing assembly, as well as the problems of rotor heat evolution reduction. The latter are stipulated by the complexity of fast-rotating rotor's cooling.

To solve this problem, the article studies losses caused by eddy currents in permanent magnets of high-speed electromechanical energy converters.

The eddy currents losses in permanent magnets and rotor retaining shell are generated by spatial harmonics caused by electric motor structural specifics, stator serration, windings diagram and distribution ratio, as well as temporal harmonics, stipulated by the external circuit, such as inverter. Moreover,with the improper selection of the electric motor parameters eddy current losses may lead to the permanent magnets overheating and their demagnetizing under the effect of this overheating.

It is generally assumed, that the losses stipulated by temporal harmonics are higher than the losses caused by the spatial harmonics. This statement is valid only for a number of structural schemes of high-speed electric motors. For example, the electric motors with toothed windings feature significant spatial harmonics. And losses caused by these harmonics are higher than the losses caused by temporal harmonics.

It is found that with rotation speed increasing the losses in permanent magnets have maximum point, after which they start decreasing. This is explained by the fact that with rotor rotational speed increase, the magnetic field penetration depth into the permanent magnet body and bandage reduces. Thus, the losses reduce either.

The article shows also that the magnetic system does not exert a significant effect on the eddy current losses, created by spatial harmonics, in permanent magnets. The eddy currents losses in permanent magnets herewith may alter significantly due to load angle variation.

The article individually considers the losses caused by temporal and space harmonics. It also presents their numerical evaluation and describes of their minimization techniques.

Kovalev K. L., Tulinova E. E., Ivanov N. S. Comparative analysis of magnetomotive forces of the reverced structure synchronous motor with permanent magnets and excitation windings. Aerospace MAI Journal, 2017, vol. 24, no 4, pp. 151-158.

The article considers synchronous motors of a reversed structure with electromagnetic excitation of both conventional and based on high temperature superconductor tapes (HTST). The presence of excitation winding allows perform deep regulation, while current carry capabilities of modern HTS tapes of second generation allow create magnetomotive force (MMF) of the excitation winding, exceeding permanent magnets.

Synchronous electromechanical transducers of reversed structure with electromagnetic excitation open prospective application domains in wind-power engineering, low and middle power hydropower, special and military applications.

Based on analytical solution of the problems of magnetic fields distribution in active zone of electric motor of a reversed structure with both electromagnetic and permanent magnet (PM) excitation, the authors obtained equations for electromotive force (EMF) and inductive resistance. The obtained equations allow determine the dependence of motor's output parameters on the pairs of poles number, geometry of active zone, and excitation MMF. Likewise, in case of HTS tapes' implementation in the excitation winding (EW), it is possible to define the dependence of the motor parameters on the properties of the tape in use. Based on analytical equations the comparison of the motor of a reversed structure excitation MMF with PM excitation was performed. Besides, the analytical equation allowing compare these two types of excitation was obtained. It is shown, that the power of a motor with electromagnetic excitation can be greater with lower number of poles and HTS tape current close to 100 A. The obtained analytical equations can be employed for optimization calculations while defining the main sizes of a motor active zone. The combination of the presented fundamental solutions of theoretical problems and modern simulation methods will allow develop new calculation procedures for both traditional motors and motors based on HTS materials.

Reznikov S. B., Kiselev M. A., Moroshkin Y. V., Mukhin A. A., Kharchenko I. A. Combined electric power complex modular and scalable architecture for all-electric aircraft electric power systems. Aerospace MAI Journal, 2017, vol. 24, no 4, pp. 159-169.

The all-electric (more electric) aircraft (MEA, РОА, МОЕТ) concept is currently the main trend in the development of the perspective aircraft power system both in the Russian Federation and abroad. This concept assumes the replacement of aircraft pneumatic and hydraulic actuators by electric (or electro-hydraulic) ones, as well as transmission generators' constant speed drives elimination.

The total rated capacity of MEA aircraft electric power supply system can reach up to 1,5 MW. To ensure the specified quality of the electric energy at the consumers inputs and mutual backup, its distributing channels should e connected in paralles. Thus, each channel should contain a higher voltage (270 V or 540 V) DC link in addition to the low voltage (27 V) central distribution unit with battery. These higher voltages are no used for feeding the distribution unit buses due to the complexity of arcless commutation provision. Thus, each MEA electric power supply channel is an independent combined AC-DC complex with four types of the central and peripheral distributing units: 115/200 V, 360... 800 Hz; 115/200 V, 400 Hz; ± 27 V and ± 270 (540) V. Electronic secondary power supplies interconnect these units with each other.

The authors suggest the structure of combined electric power complex with secondary power supplies' modular and scalable architecture for all-electric aircraft power supply systems with increased power-to-weight ratio based on unified multipurpose switched mode converters. This structure ensures parallel operation of both supply channels to improve electric energy quality.It reckons multiple mutual redundancy of the circuits for essential consumers feeding, and allows unify multifunctional switched mode converters, constituting it.

The considered above nonconventional MPC circuit solutions provide the required electric power quality in static and dynamic modes, high specific power (per unit mass and volume) and increased functional reliability. These solutions are protected by Russian Federation priority and provide high extent of import substitution in listed products of power electronics.

Kirillov V. Y., Marchenko M. V., Tomilin M. M. Spacecraft elements and units benchmark test on electrostatic discharges impact. Aerospace MAI Journal, 2017, vol. 24, no 4, pp. 170-175.

The spacecraft onboard equipment electronic components and units, as well as cable networks benchmark testing on electrostatic discharges (ESD) resistant strength should be carried out conditions closer to the real spacecraft operation conditions, in which electrostatic discharges occur.

Benchmark tests are performed in the air medium, and electrostatic discharges are simulated an ESD-generator. Thus, drawing near real conditions of the outer space is possible only by insulating elements and units from grounding circuit and maximum offset from the conducting environment to reduce the capacitive coupling.

Establishing the standard requirements to the onboard equipment noise immunity to simulated electrostatic discharges impact allows ensuring the possibility of comparative analysis of the testing results of various space vehicles.

These standard requirements should specify the simulated ESDs types; the degree of the tests' robustness; characteristics of the working place for tests. The testing methods should account for the specifics of onboard elements and units, as well as cable network placing on the spacecraft structure.

The article presents the description and requirements for the spacecraft onboard elements and units, as well as cable network benchmark testing.

The authors suggest performing the benchmark tests in such a way that the elements and units under testing together with along with spacecraft shell element and measuring equipment would not have connections with grounding circuits and power network, and placed far from the conducting medium.

Didyk P. I., Zhukov A. A., Podgorodetskii S. G., Zabotin Y. M., Golikov E. A. Experimental evaluation of metallization quality of through holes in silicon wafers. Aerospace MAI Journal, 2017, vol. 24, no 4, pp. 176-183.

Metallization of through holes in silicon wafers has been investigated by electron microscopy methods. Dependences of the thicknesses of metallization at various depths of through holes in wafers for single-sidedand double-sided sputtering of chromium and copper, with thickness of 1 µm to 5 µm, as well as with successive galvanic deposition of layers of chrome and copper and chrome and SnBi alloy (tin 98-99%, bismuth 1-2%) on the films of chromium and copper were obtained by vacuum magnetron sputtering method. Optimal modes of through holes metallization in silicon wafers process with closest characteristics of film deposition along o all structure elements, consisting in performing the process in two stages were determined. Initially, employing vacuum magnetro n sputtering method prepare metallization with minimum thickness, ensuring formation of continuous metal film. With two-sided metallization by vacuum magnetron sputtering of chrome and copper, the derived films have minimum thickness in the middle of the through holes. The continuous film is formed at chrome and copper thickness more than 1.7 µm on the surface of the through holes. To ensure the thicknesses it is necessary to perform two-sided sputtering of chrome and copper by vacuum magnetron sputtering methods with thickness less than 4 µm. Then, by galvanic precipitation method refilling to the desired thickness should be performed by galvanic precipitation method. Thickness changing at one-sided metallization sputtering, obtained in through holes by vacuum magnetron sputtering method presents linear decreasing character with increase of the holes' depth.

The minimum thickness of metallization is determined, at which a continuous metal film is provided along the entire depth of the through holes in the wafers. With a thickness of less than 1 µm, the surface of the film in the through holes is not continuous, but an island one. When sprayed from the front side, a continuous film forms on the surface of the plate, but the metal structure it is not continuous on the chamfers and walls.

Solov'yanchik L. V., Shashkeev K. A., Soldatov M. A. Control method for electrically conducting properties of polymer compound. Aerospace MAI Journal, 2017, vol. 24, no 4, pp. 184-194.

This article subject of research are electrically conducting polymer compounds based on epoxy binder with non-covalently modified carbon nanotubes (CNTs). Such compounds can be applied as binders to create hybrid functional polymer composites. Fluoro-organo-silicon block copolymer was used as CNTs' modifier, which is organically compatible with epoxy olygomers. It allows regulate the interaction of the modifier with polymer matrix and study the nano-composite's functional properties under various distribution of the carbon tubes.

The goal of this research consists in developing a method to control electrical conductivity of polymer nano-composites by controlling the spatial distribution of CNTs in the bulk of the binder under development to create hybrid polymer composite materials with functional properties.

In the course of this work execution the experimental research on the development of a method of preparation of electrically conducting binder based on epoxy resin and non-covalently modified CNTs. Measurements of electric conductivity of hardened composition were performed. Since the non-uniformity of the CNTs' distribution over the nano-composite surface does not allow determine the value of the surface resistance with adequate accuracy by contact methods of conductivity measuring, the non-contact method was used based on measuring the electromagnetic wave reflection coefficient within the range of 20-35 GHz. The authors measured also the viscosity of the binder and determined the spatial distribution of nano-particles in the bulk of composition by scanning electronic microscopy and determination of element composition.

The effect of the modifier concentration on electrical conductivity and rheological properties of the binder was studied. It was established that the modifier concentration variation allows regulate electric conductivity of nano-composite and viscosity of modified binders under the constant concentration of CNTs. In the course of this work we obtained the values of electric conductivity of about 7.3 S/m with the viscosity of the developed binder comparable to the basic binder.

The results of the study allow solve technological problem of decreasing the viscosity of epoxy binder modified by carbon nano-particles, to produce electrically conducting hybrid polymer composite materials under conductivity preserving.

Nochovnaya N. A., Nikitin Y. Y., Gudkov S. V., Savushkin A. N. VT20 titanium alloy properties estimation after removing of operational carbonaceous impurities by chemical means. Aerospace MAI Journal, 2017, vol. 24, no 4, pp. 195-202.

The lack of information in domestic and foreign sources on the effect of carbonaceous impurities purification technology on titanium alloys' properties complicates for technologists selection of the most effective and safe methods of purification of a gas turbine engine compressor air-gas channel parts and units.

The purpose of this work consists in evaluating the property change of VT20 titanium alloy while removing carbonaceous impurities by chemical means.

The studies were performed with laboratory samples manufactured from a VT20 alloy sheet-billet. Caronaceous impurities, imitating operational ones, were applied on a number of samples according to the developed technology.

Eight foreign made and domestic chemical technologies (compositions) were studied as purification means.

The authors established that the most effective removal of the carbonaceous impurities from the surface of the heat-proof VT20 titanium alloy was ensured by domestic purifying solution No 1, a two-stage purification technology in alkaline and acid solutions (“loosening + etching”), and foreign made solution HDL 202. However, while purifying carbonaceous impurities with HDL 202 solution a general etching of the surface and its microstructure change might occur.

The surface roughness values of the VT20 titanium alloy do not change significantly after the removal of impurities. The relief and profiles of the purified surfaces have a shape similar to those of the original samples.

A slight increase in the microhardness of the purified samples (up to 5%) can occur due to gas saturation of thin surface layers, due to both formation of carbonaceous impurities and the processes of chemical surface purifying.

When purifying the surface from carbonaceous impurities, the activity of the surface decreases, regardless of the type of the solution used. The least decrease in activity is ensured by cleaning solution No 1.

There is no deterioration of moistening characteristics by the VPr16 solder of the surface purified from the carbonaceous impurities by purifying solution No 1 or two-stage “loosening + etching” technology and HDL 202 solution.

Purification of carbonaceous impurities by all studied solutions does not lead to VT20 alloy strength and plastic characteristics degradation, and to a change in the character of its destruction under conditions of static loading.

Grigor'eva Y. A., Omel'chenko I. N. Orders formation and realization process organization system in the context of their life cycle. Aerospace MAI Journal, 2017, vol. 24, no 4, pp. 203-212.

The article presents specifics of orders formation depending on their basic characteristics.

The presented work discloses the fact that at present the system of orders formation and realization process is of paramount importance. This is important so that the orders themselves should be dealt with in the context of their life cycle. The orders can be split between each other according to such parameters as volume, liability distribution while their fulfilment, delivery periods and formation method. The process of the order commissioning or the order lifecycle can be conditionally split into certain stages. The life cycle stages would differ from each other for various types of the order commissioning. Therefore, one should have an idea of an order life cycle specifics for various types to minimize the error occurrence probability in strategy selection, and, as a result, minimize the probability of financial losses risk occurrence.

The article consists of three main parts, namely, introduction, the gist of the work and conclusions. It presents the description of various stages specifics, and determines the relationship of the enterprises activities marketing component and the logistic one. An algorithm for the order maintenance was developed, and efficiency evaluation technique is presented for one of the order types.

Finally, the authors describe the orders formation technique through the Internet media and its efficiency evaluation tool. The article presents the basic technologies for the delivery of the order, best suited to the market demands.

Eventually, it is worth noting that companies should master orders fulfilment and maintenance methods through Internet medium, since at present this is the prospective trend of development. Besides, these methods are fast and economically sound as well.

Terent'ev V. B., Terent'eva A. V. Ideal point method modernization. Aerospace MAI Journal, 2017, vol. 24, no 4, pp. 213-220.

When solving the problem of objects' multicriterion selection and seriation, modern mathematical modeling technologies can employ various methods, including simple aggregate weighing (SAW) and “ideal point” (TOPSIS) methods.

When comparing alternative objects of research by their effectiveness, the necessity occurs to account for not only positive or negative indicators, but estimate the objects by the degree of proximity to a specified value, i. e. to the criterion. It should be noted, that the criterion value could lay in the middle of the range of the considered indicators. Besides, the object's effectiveness, as a rule, has non-linear dependency from the change of the indicator value. Inasmuch as the existing algorithms of SAW and TOPSIS methods do not allow perform such task, a certain modernization of the TOPSIS method is required. This method is top-of-the-line with respect to the ranking procedure.

In general, when the case in hand is the indicator's degree of proximity to the specified value, the attainability function is used. In multicriterion analysis, it is called the utility function. It allows realize transformation of the initial “decision matrix” system into normalized matrix, with account for the proximity to the specified values of the criteria. This operation is close in its meaning to linear (nonlinear) normalization. It is performed in the SAW method (determines the degree of the maximum or minimum attaining), and replaces the rationing using the TOPSIS method.

Earlier, the TOPSIS method could be applied, when a monotonic-increasing utility function existed for each criterion. In other cases, one had to apply the more simplified SAW method.

The presented TOPSIS method modernization gives, firstly, practically a comprehensive agreement of the computational results with the above said methods for positive indicators, and, secondly, a slight difference with the SAW method while using positive and negative indicators, when the unknown function of the relationship between efficiency and indicators is non-linear (linear and non-linear normalizing).

Thus, the proposed modernized version of TOPSIS method allows extend the scope of this method in the case of the specified criteria values (positive and negative), located within or outside the range of the indicators variation.

Korshunova E. D., Smirnov S. D. Methodical approach to industrial startup development management mechanism efficiency determination. Aerospace MAI Journal, 2017, vol. 24, no 4, pp. 221-225.

Startups become the benchmarks of innovation growth, and the government is interested in their successful functioning. It is proved by the formation of support infrastructure around startups. An integrated approach is required to startups' support realization.

The article describes the life cycle model of a startup. Irrespective of the startup's type and line of activity, each of them passes typical stages in the course of its development. The startup lifecycle model is associated with the I. Adizez's model of the organization lifecycle.

In the beginning of their functioning, most startups face with the necessity of solving the similar problems and performing the similar functions, particularly with the necessity of the substantiated selection of the way of their development.

The management mechanism of industrial startup development allows obtain a justified choice of the way for its further development. The participants of an expert group obtained a conclusion on the most effective method by carrying out the procedure of a startup development evaluation. However, it is necessary to ensure the effectiveness of the taken decision.

The expediency of applying the mechanism and the efficiency of the taken decision should be confirmed by calculations of economic efficiency. Thus, the methodological approach to the management mechanism of the industrial startup development becomes a crucial issue. All approaches consider the object of evaluation from certain sides, and are based on specific external and internal information.

The article describes the approach to efficiency determination of the industrial startup development mechanism management. A structure for calculating the total costs of a startup during the period of itscommercialization with the chosen development process was developed. A typical process for analyzing and calculating the total costs of a startup is presented.

Manvelidze A. B. Status analysis and forecast of operated aircraft writing-off. Aerospace MAI Journal, 2017, vol. 24, no 4, pp. 226-234.

Monitoring of the current state of the aircraft fleet is an essential component of the management process of the aircraft fleet renewal by introduction of new types of aircraft with improved technical and economic characteristics.

The branch (Federal Air Transport Agency -Rosaviatsya) does not supervise (keeps record) on such deals as purchase and leasing of foreign made aircraft operated in Russian Federation. We mean the monitoring with respect to concrete transaction number, the date of leasing commencement and its expiration.

In this connection, the problem of such aircraft retirement from the Russian air transportation market is difficultly formalizable. This analysis was based on publicly available databases Flightglobal (http://, contract data and published reports on big deals of the companies.

For the purposes of the analysis the operated aircraft was separated into several groups according to ownership types and aircraft age. These groups are as follows: the aircraft owned by air carrier; aircraft obtained by financial leasing (from which the author separated out the subgroups of “young” aircraft, the aircraft with life span lower than 12 years, and the aircraft with life span more than 12 years); the aircraft in back leasing and operational leasing.

The owned aircraft retirement was determined according to the expected life span, or maximum permissible flying hours and endurance cycles from the commencement of operation.

The retirement of an aircraft being in financial leasing or leaseback can be forecasted only by life the span and total operating time.

Meanwhile the economic mechanism of the transition of the second-hand aircraft from the big companies to regional Russian companies is not developed.

The article presents some results covering the general situation in passenger aircraft fleet in the branch at large, and more detailed on Aeroflot Russian Airlines.

Luk'yanova A. A., Kononova E. S., Belyakova E. V., Smorodinova N. I. Possibilities of sustainable social-economic development of the northern territories. Aerospace MAI Journal, 2017, vol. 24, no 4, pp. 235-240.

At present, the problems of sustainable development of socio-economic systems of various levels are in the sphere of close attention of Russian and foreign scientists.

The goal of the article consists in considering the possibilities of sustainable development of the northern territories, which have pronounced specific features. To clarify the concept of sustainable development of the northern territories, the article reveals limitations in the use of their resource potential, namely: high level of production costs, high vulnerability of the natural environment, preservation of the traditional way of life and the need to improve the quality of life of indigenous small peoples.

Based on the revealed limitations, the article formulates the principles on which the sustainable social and economic development of the northern territories should be based, the priority role of which is assigned to the implementation of the latest technologies to ensure high quality of life for the population and improve the ecological situation.

Remoteness of the Northern Territories from large settlements, its difficult climatic conditions, and the poor development of land infrastructure predetermine the high importance of air communication for the sustainable social and economic development of the North. Thus, the latest technologies used for the development of the Northern Territories can be concentrated in this area.

The article analyzes the experience of Canada and Alaska in the development of air communication. These areas were selected for analysis due to fact that natural and climatic conditions, labor endowment provision level, the distance from the economically developed regions allow establish certain similarity of these territories with the territories of the North of Russia.

Based on the performed analysis of international experience, the article suggests the following opportunities for using modern aviation and space technologies to ensure sustainable social and economic development of northern territories:

– Development of hub airports, as well as regional and local air transportation, supposing inclusion of regional and local airports into nodical structure of air transportation servicing;

– Implementation of unmanned aircraft for the delivery of goods;

– Implementation of high- capacity space communication vehicles to create a broadband satellite communication network;

– Remote sensing of the earth surface for natural resources development while meeting the requirement to preserve the unique ecosystem of the North.

The considered technologies are quite expensive and require significant investments in research and development. In this regard, the opportunities for sustainable socio-economic development of the northern territories are closely associated with the role of the state in socio-economic development.

The article concludes that the formation of an effective system of interaction of federal and regional authorities, business and indigenous people, the optimal combination of market mechanisms and public administration tools will ensure the implementation of opportunities for sustainable social and economic development of the northern territories and level the specifics of these territories that complicate this process.

Galkin V. I., Paltievich A. R., Shelest A. E. Modeling and evaluation of defects occurrence reasons while isothermal punching of ribbed panels from aluminum alloys. Aerospace MAI Journal, 2017, vol. 24, no 3, pp. 170-178.

Ribbed panels from aluminum alloys are widely used in aircraft industry as power structural elements, parts of the wing and fuel tanks, as well as in the form of the heat-exchange surfaces. Increased requirements on strength and reliability are rendered to such panels. The most rational technique for such kind of panel manufacturing, i. e. isothermal punching, may lead to clipping in the ribs and sink marks on the side, opposite to the ribbing.

Modeling and experimental results of the studies reveal that defects stems from the combination of manufacturing process control parameters, such as temperature and deformation velocity, as well as geometrics, i. e. blank thickness.

The main objective of the studies consists in developing design technique of the part blank design as a function of temperature and deformation velocity while isothermal punching.

The put forward problem is solved by control polynom development, linking manufacturing process parameters – the blank temperature, velocity and geometrics with the defect magnitude, i. e. sink marks in the ribbed aluminum panel while its manufacturing by isothermal punching technique.

The initial data for the required polynomial is the results of finite element mathematical modeling with varying initial parameters of the punching process and the magnitude of the forming sink mark or its absence.

The obtained modeling results were processed according to the three-way analysis of variance planning procedure. The regression equation was obtained to compute the sink mark magnitude in the ribbed panel in dependence of the process temperature and velocity, as well as the initial blank thickness.

The authors applied the analysis of variance, which allowed define the significant factors in the calculated polynomial, and, neglecting the rest, significantly simplify it.

The sink mark magnitude obtained with the calculated polynomial correlated well with the results of mathematical modeling and experimental studies.

The proposed method is universal and can be implemented for various cases of defect-free technological processes design, when evaluating the impact of the process's control parameters on and their contribution to the manufactured product's characteristic being studied is required.

Eremeev N. V., Eremeev V. V., Kondyukov S. L. Technological specifics of manufacturing of anodes based on aluminum-indium alloys system for chemical current sources. Aerospace MAI Journal, 2017, vol. 24, no 3, pp. 162-169.

Presently, one of the meaningful problems in modern machine building consists in creating new electric power sources. It is of special importance for such human field of activities as aircraft and spacecraft building.

Chemical current sources (CCS) based on aluminum anodes, where various solid, liquid or gas oxidizers used as cathodes, found an extensive application in spacecraft electric power systems. Very often, the modern methods of such sources design, however, turn out rather complicated and energy consuming.

One of the most successful electrical systems based on CCS, applied in modern spacecraft, is an oxygen-aluminum system with liquid electrolyte, consuming oxygen from the environment.

In most cases, anodes are made of aluminum alloying with such metals as Ga, Sn, In. Nevertheless, the high values of anode potential and current were obtained while the experiments with Al-In alloy anode. It was found, that aluminum doping with Indium ensures anode electrochemical activity with faraday efficiency no less than 90%.

Thus, this work was focused on developing the scientifically substantiated technology of anode manufacturing based on Al-In alloy to ensure highly dispersed, isotropic structure to provide a uniform anode dissolution, decreasing pitting formation, and, as a consequence, increasing energy and performance characteristics.

The main difficulty while Al-In alloys casting consists in organizing a uniform indium particles distribution (which is not soluble in a solid aluminum) over the solid base metal volume. The reason to it stems from too large difference between aluminum and indium melting points (659°C and 156°C respectively), as well as high density of the latter (6.5 g/cm3). Introduction of traditional modifiers into the alloy is unacceptable, since they (Ti, Zr, B) aggravate the electrochemical figures.

The studies conducted in MAI (laboratory UNPL “TOMD”) allowed develop technological scheme for obtaining anodes' blanks. The scheme includes obtaining ring blanks using additive technology of centrifugal casting and pressing by using shear deformation during pipe billet extrusion. It will allow work out sufficiently the alloy structure, grind up the phase inclusions and, as a result, ensure the necessary properties' level.

A distinctive advantage of the developed technology compared to the analogues is the possibility of regulating of a significant number of factors during the deformation process and, accordingly, to obtain the best possible material's characteristics in the final product. This technology is realized herewith using traditional equipment.

Kolesnikov A. V., Kolesnik A. V., Zabolotskii A. P. Pneumo-thermal molding of sandwich wedge-like panels from titanium alloy VT20. Aerospace MAI Journal, 2017, vol. 24, no 3, pp. 155-161.

The presented work deals with considering pneumo-thermal molding and diffusion welding (PTM/DW) technology for multilayer structures manufacturing from titanium alloys, including the ones of variable height. The paper represents the presentation of general theory of the above said technology, and analysis of the problems emerging while its realization.

The author separated out the stages of PTM/DW technology of multilayer titanium panels.

The main problem considered in the paper consists in the problem of non-removable defects formation, accompanying manufacturing of multilayer wedge-like panels. These defects are imaged on the appended plots and figures.

The reason of these defects occurrence while multilayer panels molding lies in the different displacement of the lower shell in various areas of the pack of sheets. In the area of diffusion welding this displacement is constrained by ribs of the filler, while in the zones which are not welded with the filler, the upper shell is forming freely under gas pressure. Its deflections are forming herewith between the areas of welding with the filler.

Solution of this problem consists in defining the managing program, necessary to form the ribs of the filler and the shell, whereby the shell deviation in the areas unreinforced by the ribs would not reach critical value.

The recommended range of the shell and filler thicknesses ratio in dependence on the shell deflection in the areas unreinforced by the ribs, as well as equations for determining critical deflection factor and molding pressure were obtained by mathematical modeling.

Application of the above said equations for the filler and shell thicknesses of multilayer wedge-like panels will allow avoiding defects occurrence, which was confirmed in practice. All fabricated panels comply with the calculated parameters. No defects were detected over the profile section dimensions.

In view of the foregoing, one may state that the problem consisting in determining the regularities for selection of design and geometry parameters of multilayer structures, allowing ensure qualitative molding process without defects formation was solved successfully, and the solution has practical applicability.

Ismagilov F. R., Vavilov V. E., Tarasov N. G., Aiguzina V. V. Integrated high-temperature starter-generators with intermittent concentrated winding. Aerospace MAI Journal, 2017, vol. 24, no 3, pp. 143-154.

The main objective of the research is the direct integration of the electric machine such as generator or starter-generator on the low-pressure and/or high-pressure shaft of an aircraft engine, and the gearbox elimination. This will allow reduce the aircraft engine's weight and size figures, as well as improve the aerodynamic efficiency of an aircraft as a whole. This article presents the design and experimental research of the scalable prototype of high-temperature starter-generator with the inner rotor for more electric aircraft. The fundamental difference of the developed generator from the conventional machine consists in no oil ingress into the rotor or stator cavity. The starter-generator is immersed in an aircraft engine oil chamber, containing the oil necessary for bearings lubrication at the temperature of 120-160 °С. The stator and rotor are not lubricated with oil, which does not circulate. Cooling is achieved by losses' heat sink into the surrounding oil. A scalable high-temperature starter-generator prototype model was developed in Ansys Maxwell software package. It revealed a high accuracy and close convergence with the experimental results. Moreover, the system efficiency assessment and computation of losses in starter-generator's elements were performed. Based on the experimental results and computer simulation the starter-generator full-sized model was developed, and tests at the temperature of 120 °С were conducted. This generator appeared to be less loaded from the viewpoint of electromagnetic and thermal loads. It proves the efficiency of the proposed conception and its effectiveness for implementation in more electric aircraft.

Komov A. A. IL-76MD-90А aircraft competitiveness recovery. Aerospace MAI Journal, 2017, vol. 24, no 3, pp. 7-12.

The paper compares competitiveness of Il-76MD-90А with the US C-17 military transport aircraft. The basic assessment criterion is the aircraft capability to perform landing on unprepared sites with restricted run length, which requires employing the engine thrust reverse.

The problems under discussion relate to employing thrust reverser of PS-90A-76 engine, installed on the Il-76MD-90А aircraft. These problems do not only increase the cost of an aircraft operating cycle and affect the flight safety, but reduce its competitiveness as well. The paper presents computation and experimental data, revealing that the main cause of the emerging problems consists in poor external aerodynamics of the power plant during an aircraft ground run employing the thrust reverse. By external aerodynamics the authors mean the gas jet discharge type form the engine reverse units, which may interact with the engine itself and control airframe surfaces while its ground run. Such interaction can lead to:

– gas dynamic instability in engine operation;

– damages to the rotor blades of the engine caused by foreign objects thrown from the surface of the aerodrome;

– Aircraft dynamic characteristics deterioration (wind drag, stability, controllability), and aircraft run-length increase. Unsatisfactory external aerodynamics of the Il-76MD-90A aircraft is the cause of its poor competitiveness compared to the US military transport aircraft S-17.

Ways to the aircraft external aerodynamics improvement are considered below:

– the engine reversal thrust value optimization;

– reverse jets discharge optimization in accordance with the aircraft layout.

Ways of the Il-76MD-90A aircraft external aerodynamics improvement were developed based on estimated and full-scale studies. The substantiation of the developed measures is based on the design features analysis of the S-17 engine reverser unit.

From the above said the author concludes:

  1. The level of Il-76MD-90A aircraft power plant external aerodynamics is not high enough.

  2. The Il-76MD-90A airplane competitiveness recovery requires carrying out studies on the power plant external aerodynamics improvement, which will allow competition with the S-17 aircraft in the foreign market.

Baklanov A. V. Stepwise gas turbine engine combustion chamber development in conditions of air velocity forcing at compressor outlet. Aerospace MAI Journal, 2017, vol. 24, no 3, pp. 13-22.

Development of new up-to-date actuating gas turbine engines based on fourth generation aircraft engines requires certain time consuming. Thus, one of the ways to series-production engines' parameters improving consists in their upgrading and forcing. A fifth series-production NK-16-18ST engine was developed hereupon at SC KMPO. Its more productive high-pressure compressor allowed ensure higher flow velocity (about 170 m/s) at the combustion chamber inlet. Combustion organization and provision of optimal level of toxic agents' emission in engines of such kind is hindered due to high-pressure parameters of the airflow.

Such situation led to the necessity for carrying out research and design effort consisting in altering structures of burner and flame tube with redistribution of air vents along its length. The approach, used in the above said structure lies in forming the «reach» mixture in combustion chamber primary zone with its subsequent sharp weakening to ensure “poor” content, which allows maintain low level of nitrogen oxides. Testing of this chamber together with the engine confirmed that the sel ected approach allows reduce nitrogen oxide content in combustion products. However, it requires a number of measures related to the structure changes to achieve the desired level of noxious substances emission. To increase penetration depth of a jet into combustion zone the chamber was upgraded by cylindrical hubs installation in the first row of vents. This measure allowed reduce concentration of oxide nitrogen emission in the engine's exhaust gases, but it was not enough to ensure the level, required by regulations. Having in mind, that residence time reduction of gas in high-temperature zone decreases oxide nitrogen formation, in the framework of the last version, measures were introduced to increase fuel mixture flow velocity fr om atomizers, mounted on the flame tube head. The atomizers have elongated nozzles and less diameter. Such configuration allowed ensure noxious substances emission in combustion products at the level complied with the State Standard requirements (GOST 28775-90).

Piunov V. Y., Nazarov V. P., Kolomentsev A. I. The upper stage oxygen-hydrogen rocket engine energy characteristics improvement by structural scheme optimization method. Aerospace MAI Journal, 2017, vol. 24, no 3, pp. 23-33.

The informationally and navigationally oriented spacecraft injection to the working orbit with high positioning accuracy, scientific and research spacecraft transition from support orbits to departure trajectories for deep space flight and other complex tasks of space exploration are carried out by rocket transportation systems. These systems include specialized withdrawal means, named “upper stages”. The following requirements, such as enhanced energy efficiency and reliability, long-term staying in starting readiness mode, protracted operating time and multiple starts are imposed on upper stages' cruise engines. The «liquid oxygen-liquid hydrogen» cryogenic pair burning engines possess maximum energy efficiency. The first home-produced oxygen-hydrogen LRE is 11D56 engine developed at Khimmash Design Bureau headed by A.M. Isaev. This engine can be considered as the basic one for ecologically clean upper stages for rocket carriers of “Angara”, “Soyus 2-16” and “Soyus 3” families presently under development. This engine's design allows modernization or modification (without significant time consumption) of its structurally stand-alone units, preserving characteristics, which define the engine workability and reliability at large. The KVD1 engine energy parameters and characteristics updating is realized by structural scheme optimization based on the structure technical analysis and effective options selection, related to the engine usage tasks.

Based on the experience in the KVD1 engine chamber design and development two options for chamber with retractable nozzle headers design were considered. For these options, corresponding to the two engine modernization variants, optimization of nozzle divergence geometric degree was carried out. Calculation of working process parameters and the main chamber characteristics optimization was performed.

The specific impulse's increase is analyzed by optimum relationship selection of fuel components consumption and selection of the maximum (optimal) pressure in the combustion chamber selection. The optimality criterion of fuel components consumption is payload weight maximum at geostationary orbit, at which, according to the specific impulse mass equivalent, the mass gain is equal to the fuel tanks of the engine unit the mass gain. the results of theoretical and calculating studies consists in defining principal design solutions of two variants of oxygen-hydrogen engines' chambers, under development based on KVD1 LRE.

Goza D. A. Development and investigation of laboratory model low-thrust thermal catalytic thruster on “green propellant”. Aerospace MAI Journal, 2017, vol. 24, no 3, pp. 34-42.

“Green fuel” is an aqueous solution of a high-energy oxidizer (hydroxylammonium nitrate and others), and a fuel, presented by various substances, such as alcohols, glycerin, etc. It offers a number of advantages, namely, a higher density, low freezing temperature and high specific characteristics. Such mixtures relate to low-toxic substances, whereas hydrazine is a high-toxic substance. Thus, the “green fuel” mixture implementation as a monopropellant for an aircraft correction and orientation thermocathalytic thrusters is up-to-date issue.

Hydroxylammonium nitrate was sel ected as a basis for the “green fuel”, to which a fuel and dissolvent (water) are added in calculated ratio. The energetic qualities of the fuel depend on its basis, though its output characteristics are strongly affected by the water content in the mixture.

The laboratory model consists of a heater for the structure's starting warm-up of the, combustion chamber fr om refractory metal with a special protective coating, catalytic bed, consisted of a combination of metallic and granulated catalysts, an injector unit ensuring operating pressure differential, and a system of thermal screens.

The laboratory model presents a disassembling model to monitor separate elements of the structure. Besides, such model allows quick replacement of the thruster elements, such as the catalyst bed.

The laboratory model was tested in air under normal climatic conditions. The thruster was tested on firing functioning both in impulse and continuous operating modes.

The tests of the thrusters were conducted in continuous modes at the inlet's dropping pressure. It is worth mentioning that with the inlet's pressure decrease, the pressure in the combustion chamber decreases proportionally, which demonstrates the stability of the thruster operation.

The K100E laboratory model maximum run amounted to 1.5 kg of consumed fuel over 1500 start-ups. The main reason for the thruster's failure relates to the tests conduction in atmospheric conditions, namely to the oxidizing and destruction of separate parts of the laboratory model (heater, screens) under higher operating temperatures.

Maximov N. А., Solodovnikova D. A., Sharonov A. V. Mobile system for fixing and accounting for aircraft external damages while preflight checkup. Aerospace MAI Journal, 2017, vol. 24, no 3, pp. 43-50.

The paper tackles the version of the mobile system for fixing and accounting for aircraft damages while preflight checkup. The benefits of the system are defined by the possibility to use mobile devices of a tablet type equipped with high-resolution cameras. These devices fix the detected damages, and convey the images of these damages into the server part of the system, which performs their processing and automatic logging to various exploitative documents, related to aircraft servicing. The basic tasks resolved by the developed and described in this paper bundled software of the mobile system are formulated as follows:

The developed mobile system's software solves all above listed tasks, which allows accelerate not only the preflight checkup itself and filling out the related documents, but also the subsequent technical servicing (such as repair, gathering of statistical reporting systematization on the condition of the certain aircraft

  1. Realization of the possibility to use mobile hardware (a tablet with high-resolution camera, or special hardware) for damage registration and the possibility to examine the tolerances on such damages by its location.

  2. Solving the problem of a picture of damage operative “attachment” to the “structure damages list”.

  3. Realization of automatic documentation generation to send the request for the element repair to the aircraft manufacturer, if such repair is beyond the scope of the Repair manual.

  4. Loading to the database the information on the performed repair and substituting of the damaged elements picture for the new one.

  5. Return the list of existing damages of the aircraft on the request of the program.

  6. Performing checkup and repair according to instruction manuals. Furthermore, the program should provide an opportunity for regulatory documents viewing to determine further activities for the damage elimination.

Chukhlebov R. V., Loshkarev A. N., Sidorenko A. S., Dmitriev V. G. Experimental research of an aircraft product's structure vibrations under flight loads action. Aerospace MAI Journal, 2017, vol. 24, no 3, pp. 51-59.

One of the main factors affecting reliability of aircraft articles is vibration effect during the joint flight with carrier. To obtain estimates of reliability characteristics flight test of products are carried out. Modern equipment for ground vibration testing, reproducing the flight conditions, allows substantially reduce the amount of flight tests by replacing them with laboratory tests. The actual problem here is formation of laboratory tests regimes to ensure the equivalence of loading in laboratory conditions and in flight. Characteristics of vibration loads are obtained usually based on measuring data obtained during flight tests of the product or its prototype. At vibration tests, a relation is established between laboratory test modes and flight dynamic loads by the levels of vibration accelerations or stresses.

The paper presents the technique and results of flight and laboratory vibration tests on definition of vibration stresses and accelerations characteristics of an aircraft product's structure in typical flight. Laboratory tests were conducted with random dynamic loading, corresponding to loading during flight structural tests. The purpose of laboratory tests is determination of characteristics of a structure's accelerations and stresses in the conditions of a spacecraft joint flight with the carrier. This requires reproduction of conditions exhibiting adequately enough the loading condition of a product according to the basic probabilistic characteristics during typical flight.

The authors developed the technique and modes of aircraft product's vibration tests, complying with vibration loading of a product at every stage of the host aircraft flight. Using the obtained modes the tests were conducted, whereby the random dynamic loading, corresponding to the operation conditions of the product on an internal suspension bracket of the carrier, was realized. Comparison of vibration acceleration probabilistic characteristics at laboratory and flight tests demonstrated conformity of these tests' results according to root mean square values of vibration acceleration.

The developed laboratory tests technique ensures correct reproduction of random vibration loading reproduction of and aircraft product structure during the flight on an internal suspension bracket of the carrier. The technique and results of the tests can be applied for estimation the structure vibration strength of an aircraft product of various applications during the joint flight of an aircraft product with the carrier.

Pashko A. D., Dontsov A. A. Guided missile trajectory and active protection element movement determination errors design procedure. Aerospace MAI Journal, 2017, vol. 24, no 3, pp. 60-71.

At present, the onboard aircraft defense structures for protection from “air-to-air” missiles are equipped with the systems of jamming cartridges of various calibers ejection. The existing algorithm of airborne defense systems application consists in practically continuous ejection of a series of jamming cartridges when the aircraft enters the area of the enemy's air defense. However, the existing techniques of jamming cartridges implementation do not ensure the aircraft protection from the missiles equipped with matrix photodetectors. There is a contradiction between the potential onboard defense systems implementation efficacy, and military characteristics of existing onboard defense systems. In this paper, the authors propose a technique for guided missile coordinates determination errors to ensure its neutralization on the flight trajectory.

A methodology for probability estimate of aircraft skipping the hit by a guided missile while realizing the active element ejection to the trajectory of the guide missile with its subsequent detonation was developed. The probability estimate of guided missile missing its target is based on probability calculation of the active element's detonation coordinates center will appear inside the dispersion ellipse with the main axes equal to mean square deviation of the active protection element detonation coordinates from the actual position of the missile. The equations for the active protection element detonation coordinates were obtained using measuring errors theory methods on the assumption of the aircraft and missile rectilinear and steady motion, after the initial miss elaboration, with an allowance for the aircraft and missile coordinates and flight speeds measuring error, as well as current missile's angle of attack determination errors. The paper shows that the missile's angle of attack determination errors depend on aircraft and missile current speeds determination errors, as well as missile bearing in vertical plane measured by optical radar station or specialized onboard radar station belonged to aircraft protection structure.

Popov A. S. Analysis of the capacity to use a repulsive two-mass space system with periodically formed coupling to perform interorbital flights. Aerospace MAI Journal, 2017, vol. 24, no 3, pp. 72-77.

At present, methods for orbit parameters changing with cable system by periodical changing of its length for the case, when the cable is located in the plane orthogonal to the orbit plane, or when it lays in the orbit plane, are known. However, the orbit parameters changing is possible only in case of non-central gravity field. The presented paper offers the structure of interorbital transfer of the space system, consisting of two masses repulsed and retracted in the orbit plane by periodically formed coupling. The flight is considered in the central gravitational field.

Originally, the system represents a single spacecraft, consisting of two parts of equal weight. Initially, the system s on a circular orbit. The mass repulsion occurs in the direction tangent to the trajectory. Hereafter, the masses being uncoupled move independently of one another over various trajectories. Performing a various number of turns around the attracting center, after a certain period of time they will turn up on the line coinciding with the radius vector. One of the masses herewith will pass the pericentre of its orbit, while the other – its apocentre. At this moment, the masses contraction occurs assisted by the formed coupling. Methods of coupling formation are not considered in this paper. The paper demonstrates that the eventually formed orbit differs from the original one.

The authors obtained analytically the dependence of the system final velocity in the point of masses contraction after their contraction versus the speed of their repulsion ΔV.

The dependence of the masses contraction point radius vector versus the initial repulsion speed ΔV for the final orbit.


Solution of this problem revealed a theoretical possibility of orbit parameters changing for the system of a proposed type.

The analytical dependence of the speed value at the time of contraction versus the initial masses repulsion velocity is obtained.

The equation determining the radius vector in the point of masses contraction of the formed final orbit created.

Vereshchagin Y. O. Deck-based aircraft aileron adaptive control technique. Aerospace MAI Journal, 2017, vol. 24, no 3, pp. 78-82.

Active development and application of digital technologies enabled realization of advanced algorithms in aircraft control systems, which could not be implemented earlier due to the limited capabilities of analog computers, and the more, so in mechanical control systems. The attempts to ensure aircraft control characteristics invariance to varying flight conditions, aerodynamic configurations, centering and mass-inertia characteristics led to the necessity of employing two classes of characteristics onboard the aircraft, namely, with reference model and with the identifier. Such algorithms are developed and successfully applied in the control systems' longitudinal channel of Su-30Sm, Jak-130, Su-35 and Т-50 aircraft. It is important to notice that the adaptive algorithms in lateral control channel have not found practical application, though the problems requiring solution exist there either.

Thus, the problem of lateral controllability deterioration caused by occurrence of the adverse moment in yaw during aileron deflection exists on all MiG-29 modifications. The aircraft heel moment caused by lateral static stability due to the sliding is directed opposite to the effective aileron roll control moment. Flight speed reduction and increasing angle of attack corresponding to it lead to reduction of available rate of roll, and in limit case to occurrence of roll back reaction to control stick deflection. The acuteness of the problem is partially reduced due to implementation of a unique structural solution, i. e. airspeed head wind eddies generators, which, however, does not eliminate the problem at large. The situation is aggravated in the case of external suspension brackets asymmetrical mounting, which becomes a standard situation in view of the increasing effectiveness of aircraft means of destruction. The aircraft herewith begins to react differently to the stick deflection to the suspension bracket side and to the side, opposite to it, by the rate of roll. It complicates substantially delivering air combat and ground targets attacking, accompanied by drastic banking maneuvers of positive sign to negative and vice versa. Aircraft landing approach control is rather complicated, especially in case of landing on a ship deck of limited size in conditions of oscillatory motion and atmospheric turbulence.

Gurevich O. S., Gol'berg F. D., Petukhov A. ., Zuev S. A. “Virtual engine” software usage for air bleed control in gte units' cooling systems. Aerospace MAI Journal, 2017, vol. 24, no 3, pp. 83-94.

One of the trends of gas turbine engines perfecting consists in “intelligent” engine developing. Within its control system, a so-called “virtual engine” functions in real time mode, i. e. a full range thermo-gas-dynamic GTD model. Its implementation allows, in particular, realize engine control by its critical parameters inaccessible for measuring. The gas temperature in the hottest part of the engine duct, i. e. the temperature at the turbine inlet, is one of such parameters. The paper presents the result of the study of new turbine cooling control methods, differing fundamentally from conventional indirect open-loop control of air bleed valves according to rotation speed, employed in modern automatic control system. A block diagram and algorithmic provision of adaptive closed loop control of turbine cooling units operating directly according to gas temperature prior to the turbine and rate change of turbine blade temperature are considered.

The result of such type of control estimation, carried out as applied to modern turbofan engine with high bypass ratio, revealed that its' implementation may allow:

– Engine efficiency increasing by decreasing the bleeding air consumption;

– Engine lifetime increasing by turbine inlet temperature decreasing by 100...200 К at steady-state modes, and the rate of turbine blade temperature decreasing by more than 20% at transient modes.

An adaptive control of air bleeding for turbines cooling associated with gas temperature limitation by effecting on the fuel flow in the combustion chamber was considered. The paper demonstrates that its implementation is possible:

– In flying conditions, when maximum engine thrust is required. It can be increased by 10% with the maximum allowable limitation of turbine blade temperature;

– Under operation conditions when engine lifetime is critical. It allows blades temperature reduction by approximately 50 K while maintaining the thrust value and specific fuel consumption.

Bilyaletdinova L. R., Steblinkin A. I. Mathematical modeling of electromechanical steering gear with ball-screw actuator with account for nonlinearities of “dry friction” and “backlash” types. Aerospace MAI Journal, 2017, vol. 24, no 3, pp. 95-108.

The paper addresses the multi-purpose mathematical model of the electro-mechanical actuator's (EMA) dynamics. It contains the general description of the EMA, which was the object for the modelling, the description of the mathematical model developed and mathematical modeling results. The actuator was developed in the frame of the Russian-European project called RESEARCH for the elevator deflection of a regional passenger airplane. The mathematical model was implemented within MATLAB/Simulink software.

The actuator model consists of four submodels of its physical constituent parts such as controller, power electronics block, electric motor and mechanical gearbox (ball screw transducer). Programmatically switchable models with various level of detail of physical processes were realized for each part. The electrics were realized by the submodels of a single-phase DC motor and a simplified controller corresponding to it. It also contains three-phase induction motor with permanent magnets, regulated by a controller, realizing vector control in {p, q}-coordinates. Power electronics is modeled either by simplified dynamic elements, or on a physical level in detail (electronic components level). Special attention was payed to mechanical part of the actuator modeling, i. e. various submodels of non-linear mechanical effects of a “dry friction” and “backlash” were realized. Thus, we managed to ensure a balance between modeling accuracy and speed within the framework of a single model.

Based on mathematical modeling results the paper demonstrates how the dry friction and backlash parameters, as well as software methods of their realization effect on the actuator's regulation quality and its characteristics. It shows that program splitting of the actuator states (idle, motion, initiation) based on velocity smallness without using the sign function approximation is optimal method of dry friction effect accounting. It ensures reproduction of the necessary actuator motion pattern with acceptable integration step (10-4 s). The paper demonstrates also that accounting for linear stiffness of the actuator's ball screw transducer has insignificant effect on the actuator's frequency response within the frequency range of control surface control. It is shown that the replacement of the three-phase motor with a single-phase one while reducing the EMA model leads to different regulation character even while using the similar regulator structure and comparable PID-regulator coefficients.

The developed model can be used while the electromechanical flight control systems design for various engineering tasks, characterized by significantly varyng requirements imposed on the model in use. These tasks include: 1) development of the actuator and its control system, including actuator digital regulator synthesis; 2) actuator static and dynamic characteristics express-analysis; 3) obtaining reference actuator characteristics including small control signals; 4) analysis of transient responses and stability margins of the closed “aircraft – flight control system – actuator” control loop, including in-line simulation; 5) study and optimization of actuator thermal conditions while operating in the closed bay of the outer wing.

Nadaraia T. G., Shestakov I. Y., Fadeev A. A. Aircraft landing gear wheels actuator. Aerospace MAI Journal, 2017, vol. 24, no 3, pp. 109-113.

According to the item of the State Program “Development of the aviation industry for 2013–2025” creation of scientific and technological capacity ensuring global leadership in aviation technology and product promotion of domestic aviation industry on the domestic and foreign markets should ensure high competitiveness of domestic aircraft by introduction of innovative developments. Operation and maintenance analysis of the existing civil aircraft park revealed that while aircraft aerodrome maneuvering the hundreds of kilograms of kerosene are wasted, and drive trucks waste tens of kilograms of fuel. When kerosene burns in an aircraft engine, and fuel burns in combustion engine the atmosphere is contaminated by noxious substances. While aircraft maneuvering on the runway the noise level is 90 dB. Using combined actuator in landing gearwheels will allow decrease negative effect on the environment and eliminate completely the majority of shortcomings.

The paper presents the schematic diagram of electromechanical landing gear wheel actuator in which brushless switched-reluctance motors are mounted inside cylindrical gearwheels. Due to low cost materials implementation, small size and weight, low energy consumption and high efficiency maintainability better design and operating characteristics of aircraft landing gear wheel actuator are ensured. While motor-reducer design, specifics of its operation in the landing gear wheel were accounted for. The results of motor-reducer computation, which demonstrated the wide specter of implementation of such kind of actuator for various types of aircraft components, such as landing gear wheels actuators, high-lift devices' elements are given. The presented motor-reducer possesses diversified structural concepts, which allows use it for various types of aircraft both civil and military oriented, as well as for unmanned aerial vehicles (UAV) and spacecraft of various kinds. The prototype of motor-reducer, used for UAV's high-lift devices, displayed its apparent advantage compared to the other actuators, such as design compactability, manufacturability and cost effectiveness. Implementation of the above-described structure will allow fuel consumption saving by both an aircraft, and airfield servicing facilities. The structural concept of the motor-reducer in aircraft landing gear wheel does not have counterparts either in Russia or abroad.

Dyul'dina N. E., Nekhoroshev M. V., Pronichev N. D. Developing additive technology of tool electrode manufacturing for aircraft engines parts machining. Aerospace MAI Journal, 2017, vol. 24, no 3, pp. 114-120.

The paper offers new technological solutions for gas turbine engines (GTE) manufacturing. These solutions are based on special processing methods using and additive technology implementation. To improve technological process of GTE parts manufacturing the authors suggest new technology of polymeric tool electrode (TE) fabrication with subsequent metal coating of its work surface using electrodeposition method. The most complex problem consists in ensuring accuracy of the profiled surface in the process of electrodeposition of a metal layer.

The objective of the work is developing computer model of the process of electrodeposition of metal on a dielectric TE for electrochemical machining (ECM).

This presented method consists in creating the information model, and studying the main process parameters of electrochemical deposition: electrolyte and electrode surface potentials, electrode reaction behavior, thickness and uniformity of the coating. While analytic model development parameters of electrode reaction, such as the exchange current density; electrochemical anodic and cathodic transfer coefficients; system electrode reaction equilibrium potential were determined. Besides, the above-mentioned method includes comparison of the formed profile with theoretical one.

The developed information model demonstrates that the metal coating possesses a variable thickness. On the boundary of cathode with electrode junction a thickening, stipulated by electrochemical processes, was formed. Here, in this zone the thickness deviation of the formed profile from a theoretical one is 355 µm. This implies that a minor mechanical processing is needed.

The developed technology allows carry out technological regimes and using the ECM obtain more detailed information on surface shaping.

Zhukov P. A., Marchenko M. V., Kirillov V. Y. Transition resistance effect on aircraft and spacecraft onboard cable network shielding efficiency. Aerospace MAI Journal, 2017, vol. 24, no 3, pp. 121-126.

To ensure the specified shielding efficiency electromagnetic screen should be homogenous to the maximum.

The uniformity of the shield depends on the resistance between the cable shield, the electrical connector and the onboard device case, i. e. transition resistances. High shielding efficiency can be ensured with small values of transition resistances.

The transition resistance is not a constant and can variable significantly during the life cycle of a product. These variations are caused by the effect of various factors: shields and cases bonding and connecting techniques; temperature and environmental conditions; operating conditions.

The results of the experiment that simulating a stay in a tropical climate revealed, that the magnitude of the transition resistance has increased up to 8 mOm, and in some cases it increased from 1 to 26 mOm, which significantly exceeds the standard value.

While temperature fluctuations effect testing, cable connectors subjected to thermal shock by immersion in liquid nitrogen with subsequent heating to 290°C by the stream of hot air. The results of this experiment demonstrate, that the transition resistance of the heated connector increases from 1 to 6 mOm.

In all these experiments, significant changes of transition resistances values in the direction to increase without returning to the initial values were observed. The reason for this consists in the thermal deformation of the parts' shape and contact failure due to the emergence of the oxidized layer.

The results of the shielding effectiveness study show that the magnitude of the transition resistance affects significantly the levels of induced interference voltage at the load, connected to the onboard instrument simulator cable.

Transient resistance value increasing reduces the onboard cables shielding efficiency. Thus, while electromagnetic shields designing, it is necessary to account for in shielding efficiency decrease on exposure to thermal and climatic factors during the life cycle of the product.

Vyshkov Y. D., Reznikov S. B. Supercapacitors applications for aircraft engine start systems. Aerospace MAI Journal, 2017, vol. 24, no 3, pp. 127-133.

The aircraft engine start fr om standstill to rated idle in ground conditions can be carried out by electric starting gear fed by either onboard or ground-based power sources. The onboard power sources herewith are accumulators, while voltage can be boosted in ground conditions. Accumulators limit the power of electric start systems. Thus, for starting high-power aircraft engines non-electric systems are used. To increase the power of the ground based engine start systems a high-voltage power supplies can be used. The goal of the article consists in demonstrating the possibility of supercapacitors application to start aircraft engine by electric system. As far as a supercapacitor specific power is greater than that of an accumulator, they can be effectively used for increasing the power rating of aircraft engine start electric systems, wh ere accumulators were previously used. Since the energy accumulated in supercapacitor increases with voltage rise, the supercapacitors can be effectively used in higher voltage systems to increase their power.

The goal of the presented work consists in studying and comparing characteristics and processes of the starting mode of electric motor, fed by power supply containing supercapacitor and without it, based on the results obtained by computer simulation with Electronic Workbench V5.12.

The simulation results confirm the possibility of increasing the aircraft engine start electric system's power by supercapacitor implementation. It means that in many cases it will allow replace aircraft engine start air-compressing and gas turbine systems by electric systems, which possess many advantages and are of great importance for all-electric aircraft.

Sukhachev K. I., Dorofeev A. S. Development and study of magnetic induction systems for micrometeorites' and cosmic particles' acceleration. Aerospace MAI Journal, 2017, vol. 24, no 3, pp. 134-142.

This work is dedicated to development of experimental test-bench based on magnetic induction rail system. The test-bench allows the ground testing of spacecraft materials and equipment on resistance to micro particles of natural and artificial origin impacts. It will solve the problems related to the costly and inefficient space experiments, and will significantly increase the repeatability, controllability and frequency of impact experiments. In the long-run this accelerator will be an essential part for developing effective protection of the spacecraft from the meteorite hazard, non-existent at the moment.

To solve the problem of low efficiency while converting electrical energy into kinetic energy, which is of great importance for acceleration of small bodies, weighing less than 0.1 g, the authors propose an experimental technique, allowing increasing the efficiency, and, thus, the impactor's maximum speed without increasing the stored energy in storage facilities. The unique feature of the proposed technique consists in increasing the accelerating force acting on the object from external magnetic systems. The authors propose to create a localized external magnetic field directly in the surrounding area of accelerated particles, and then move the magnetization area synchronously with the movement of the accelerated object over the path of the accelerator. This effect is achieved by using multiple-magnetic systems with independent switches and drives, and a single control system. To determine the switching time parameters and parameters of the railguns magnetic systems, the technique of the railgun computation, operating in combination with the multi-loop magnetizing system has been developed.

To test the proposed approach a prototype accelerator was designed and developed. The series of experiments confirming the effectiveness of the proposed method was carried out. Experiments were carried out with particles of various masses, a variety of energy storage levels, as well as for several options for magnetizing systems. The upgraded magnetizing system was 23% more efficient than the classic one, with the same energy storage. The developed accelerator allowed obtain the speed of more than 2100 m/s with a total energy of 11.6 kJ stored in the capacitor bank was reached.

The authors plan to apply the proposed methodology to the main circuit. According to the simulation results, the main circuit multi-step power supply will also contribute to the efficiency increase of rail accelerators.

Bespalov A. V., Petrov A. P., Sokolov A. V. Friction and surface phenomena when stamping hard-deformable alloys. Aerospace MAI Journal, 2017, vol. 24, no 3, pp. 179-194.

This work considers the issues of friction (dry, boundary, liquid) effect on the hot die forging process. It reveals the main sources of frictional forces formation.

High temperatures, pressure and permanent renewal of one of the friction deformable metal surfaces being in the plastic state characterize the external friction during the hot die forging. In the course of stamping, as the die fills, the surface area to body volume ratio is increasing. The destruction of oxide films thereupon on the surface of wrought workpieces and the outcome of the non-oxidized metal particles from them occurs. This event facilitates the development of the forces of intermolecular gripping of the wrought workpiece and the tool. The stainless steel, aluminum and titanium alloys are especially prone to sticking to the tool. Thus, their stamping is always carried out with lubrication.

In most cases, the friction at contact surfaces while stamping occurs together with intervening and isolation mediums (oxide scale, oxides, lubricant etc.). Thus, the interaction of lubricants with surface-active substances while stamping becomes of particular importance.

The types of lubricants, their composition and the additives effect on the difficult-to-form alloys of low-plasticity processing are considered.

The mechanism of action of surface-active substances in conditions of stamping and formation of plasticized surface layer with ultra-fine-grained and nano-sized structure was analyzed.

The article analyzed the results of leading Soviet and Russian scientists' studies in the field of nanostructured state forming in the surface layer of the material.

Based of the conducted analysis, we can state that the nano-structuring of the workpieces' surface, including pressure shaping, while applying surface-active substances, leads not only to the obtained semi-finished products' mechanical properties substantial improvement, but also to a significant improvement of their technological properties during the subsequent hot deformation, such as stamping. Thus, the compelling for the production possibility of difficult-to-form materials' super-plasticity deformation under lower temperatures and higher speeds of not only volume nano-structured workpieces, but also the workpieces with nano- structured surface is created.

Babin S. V., Fursov A. A., Egorov E. N. The study of intermediate plasma-sprayed layer effect on fiberglass-metal junction strength. Aerospace MAI Journal, 2017, vol. 24, no 3, pp. 195-201.

The study of laminate composite materials, compounds of dissimilar materials and hybrid composite materials for increasing their strength, fatigue strength and reliability is a topical problem for aircraft building.

This work studies the technique for increasing strength of fiberglass with AB-T1 aluminum alloy compound and fatigue strength of hybrid composite material by intermediate layer creation.

To reinforce composite compound intermediate rugged porous layer, obtained by plasma-sprayed method. The paper performed comparative analysis, sel ected materials and modes to such layer formation. Fatigue testing of hybrid composites samples was carried out. Temperature effect on shear strength of a composite compound was studied. The effect of fiberglass molding process (with glue or without it) on the components shear strength.

As a result of the conducted studies we found that:

  1. The presence of intermediate layer allows increase shear strength of a AB-T1 + (PN70U30 + EP741) +BK50 + VPS fiberglass composite compound up to 50%, and AB-T1 + (PN70U30 + EP741) + VPS composite up to 90%.

  2. Implementation of plasma-sprayed intermediate layer allows increase fatigue strength of fiberglass aluminum alloy compounds up to ≈ 120%.

  3. Implementation of plasma-sprayed intermediate layer ensures workability of hybrid composite materials under consideration at temperatures fr om – 60°C to +60°C. The temperature profile  repeats equidistantly the curve of basic technology, but at higher strength values.

The results of the study can be used for new composite materials development and hardening adhesive compounds of dissimilar materials. For example, to develop hybrid composites titanium fiberglass aluminum alloy, and new SIAL variants for fiberglass aircraft propeller blades design, compressor and turbine blades for gas-turbine engines.

Tischenko L. A., Kovalev A. A., Markin A. V. Photoresist thickness selection peculiarities to ensure and improve the lithography process stability during semiconductor devices structures manufacturing. Aerospace MAI Journal, 2017, vol. 24, no 3, pp. 202-211.

Photolithography is one of the main technological processes for obtaining on a special base a certain topology of various electronic components. The most important thing herewith is minimization of all errors in the course of image transfer fr om a photomask to the photoresist layer, and at the developing stage. In this case the most accurate mage transfer is achieved.

The paper is devoted to a topical problem, namely to the photoresist thickness selection specifics to ensure and improve lithography process stability during semiconductor components structures manufacturing.

The paper describes the experimental study of the dependence of radiation energy (E0) dose, necessary to full structures' development in photoresist, from the photoresist thickness (h) on the example of SPR700-1.2 photoresist. The energy dose for the structures' full development in photoresist, determining the quantity of energy affecting the photoresist, required for full photoresist elimination from certain areas, determined by components' structures topologies is one of the basic technological parameters of photoresist.

In the course of the study one area per each of 33 silicon wafers were detected, wh ere the photoresist was completely removed. Radiation energy, at which the exposure of these areas was carried out, is an energy dose, necessary for the full structures' development in photoresist. Thus, the plot of energy dose, necessary for full structure development, versus photoresist thickness was obtained In the course of mathematical calculation, approximation of experimental harmonic dependence was performed and equation of the given curve was obtained.

Rational operating points (thickness) were determined using the plot obtained while experimental curve approximation. These points represent extremums, since with minimum deviation from the rated value, inherent to the considered operating point, the energy dose for full structure development in photoresist would vary insignificantly.

Thus, nine operating points corresponding to a certain photoresist thickness were obtained as the result of the approximated curve analysis.

The result of experimental study of radiation energy dose dependence from photoresist thickness described in this paper consists in obtaining of a number of recommended photoresist thicknesses, which observing can lead to the most accurate image transfer from photomask to photoresist layer, which, in its turn, will improve lithography process stability during semiconductor components structures manufacturing.

Pokrovskii A. M., Chermoshentseva A. S. Experimental study of nano-additives effect on properties of composite materials with interlayer defects. Aerospace MAI Journal, 2017, vol. 24, no 3, pp. 212-221.

The subject of researh in the presented paper are interlayer deffects in composite materials (CM), prevention of the rate of their occurrence by strengthening the compoistes with nano-sized powder. It contributes to safety increase while aircraft operatioin and allows prevent emergency situations duting flights.

The goal of the studies consists in developing methods for manufacturing techmology of samples made of epoxy resin reinforcement of laminated CMs with interlayer defects and nano-dispersed powder, by adding nano-particles to the binder, and obtain maximum degrees of CM filling with nano-sized powder.

In the process of performing this work the experimental samples were produced, and the series of tests were conducted. It is noted that occurrence of interlayer defects contributes to lifespan reductiono of a product made of composite materials. The analysis was performed, thereby, on determining the character of a delamination type defects growth.

The properties such as mechanical characteristics anisotropy and the possibility of hidden defects presence in the form of material discontinuity over the separation surface are intrinsic specific properties of composite materials. The paper presents the experimental results of the study how degree of filling of the ED-22 resin by nano-sized silicon dioxide powder (“Taroksil” T-20) of various concentration affects the mechanical properties of a heterogene material. A brief description of production process technology of samples, made of epoxy resin and nano-disperced powder, is presented. The above said studies are used for solving the problem of interlayer defects hardening in laminated composiste materials, which occurrence is a consequence of the aircraft parts production technology imperfection and effect of operational loads of aircraft, by adding the nanoparticles to the binder. The optimal degrees of CMs filling by nano-sized dioxide silicon powder in dependence of mass concentrarion were found. The testing results of the samples made of CM with embedded interlayer defects with adding nano-dispersed additives with various volume concentration from 0.1% to 0.5% to the binder are presented.

These tests tresults' data would be offered for implementation by the enterprises of Holding JSC “Helicopters of Russia”. The work is prospective for further consideration and implementation in the future research activities.

Balyasov Y. A. Production management in conditions of multiproduct single-part and short-run production. Aerospace MAI Journal, 2017, vol. 24, no 3, pp. 222-227.

The scope of the article is production monitoring system developing at the machine-building enterprise in conditions of single-part and short-run production, allowing data provision of production progress, contribution to managerial decisions effectiveness increase and production lead-time reduction.

This goal implies organization of a production processes information system at the enterprise that accumulates initial order information, such as engineering and design documentation, route technological processes, operational labor standards. This data allows calculate a period of execution of works (with account for product structure) and draw an activity network as the basis for the day-to-day production planning. Actual production data is fixed in the strategic points relating to such production stages as resource supply, mechanical processing, finished items transfer to the picking store and to further assembly process. Analytical comparison of the initial and actual data serves as the basis for the management decision-making concerning de-bottlenecking and can be used in production scheduling.

It is supposed to use two groups of Key Performance Indicators (KPI) of production activity that characterize:

– the conformity of production progress with planned, estimated and directive periods of execution of entire work as well as intermediate production stages;

– production volume expressed in terms of production lots quantity or standard hours of work that makes it possible to estimate the current and coming labor content as well as production continuousness.

The key feature of the procedure is that all the data necessary for the monitoring is formed automatically with execution of standard working functions of a person responsible for their execution, disposing of the difficulty to obtain additional manpower resources.

This system serves as the basis for the production processes operative monitoring adapted to the single-part and short-run environment. It allows:

– systematize and analyze the real-time data on the basis of measurement of the strategic points that specify the execution of order;

– control the process on every level, from foreman to general manager;

– estimate the planned machine utilization.

This technology of production data organization is implemented at the DB “Armatura”-branch of FSUE “Khrunichev SRPSC”; the day-to-day production planning and monitoring algorithms are being tested. As a result of the partial implementation of the procedure there is a tendency to reducing of the throughput time. The suggested technology can be used at manufacturing enterprises with a high level of experimental development, such as engineering departments with a pilot plant.

Motyreva E. E., Tarasova E. V. Hedging of financial risks of developing enterprises. Aerospace MAI Journal, 2017, vol. 24, no 3, pp. 228-235.

Nowadays developing enterprises undergo hardship to find sources for innovative projects financing. Investors do not want to invest in risky programs, while enterprises are no able to bear the financial risks themselves. A distinctive feature of the high-tech defense developments production financing is preferential financing at the expense of budgetary funds. However, that does not absolve enterprises from financial risks due to insufficient financing or failure of terms. In this case, the enterprise is better to involve extra-budgetary financing. It would help: to exclude cases of forced attraction of own funds; to increase profitability and avoid loss of development; to reduce the probability of disruption in terms and penalties; to release a part of the developer's funds for own technical re-equipment and development; to reduce the impact of design, technological, financial and economic risks.

One of the distinguishing features of the development of high technology products is the presence of the objects of exclusive rights. These objects have a certain value, depending on the area of their application, utility, degree of elaboration and novelty. Additional funds for innovative projects implementation may be attracted by the sale of options for these objects to interested parties.

For the purposes of the developing company, it is necessary to sell such a number of options and at such a price that it would be able to ensure a risk-free return (i = 25%). For the company acquiring the option, the price is defined as the present value of future benefits taking into account the likelihood of a favorable outcome.

Including the price of sold options in the financial model of the project, you can achieve its payback.

Tikhonov G. V. Methodology for russia's small and medium-sized enterprises adaptation to crisis conditions. Aerospace MAI Journal, 2017, vol. 24, no 3, pp. 236-240.

In modern conditions of the Russian economy and restructuring of certain industries, the significance of enterprises’ management increases.

It is stipulated by the disruption of a great number of economic ties, active import substitution of products manufactured by certain industries, as well as the necessity for significant breakthrough in the field of military industrial complex (MIC).

Instability of external environment with high risks level is aggravated with account for the challenges facing the Russian industry in conditions of scale sanctions.

In this connection, the role of small and medium-scale enterprises, which significance in the modern world is steadily growing in both developed and developing countries, is increasing.

Methodology for assessing and monitoring industrial enterprises' adaptation level to the crisis conditions is analyzed in real conditions.

Monitoring system of adaptation level of separate enterprises to crisis conditions is an important element of anti-crisis policy, since it should contribute to the selection of the strategically important business-partners selection, subcontractors on production activities, etc. The enterprises with high-level adaptation to the crisis conditions should be primarily included in industry plans while preparing and implementation of the products critical for the industry.

The relevance of small and medium-size enterprises adaptation process is associated with the fact that their ordering parties are large enterprises. So operations of “business for business” (b2b) type are implemented. Thus, small and medium-sized industrial enterprises in most cases are not oriented on individual consumers, but on a big business, which establishes its own rules of conduct on the market.

The area for the study is steady development mechanism of economy of industries, complexes and enterprises.

The object of the work is the Russian small and medium-sized industrial enterprises overcoming the crisis phenomena in the economy.

The subject of this work is methodical and practical approaches to the Russia's industrial enterprises adaptation to negative conditions in the economy, including the assessment of the adaptation level of the enterprises.

The aim of this work consists in developing a methodology of Russia's small and medium-sized industrial enterprises adaptation to the conditions of crisis. The methodological basis of this work is a systematic approach to the small and medium-size enterprises adaptation assessment to crisis conditions.

Practical significance is determined by a comprehensive quantitative approach to the assessment of industrial enterprises' adaptation level to the crisis phenomena in the economy.

In theoretical terms, the overall conclusions on the adaptability of economic systems can be formulated:

  1. High level of adaptive properties of any system means that significant changes in the external environment cause insignificant reaction of the system.

  2. In the framework of market relations, the more so in crisis, the important characteristics of the external environment are mobility and uncertainty. Thus, the adaptability acts as a fundamental property of such dynamic systems as a subject of small and medium-sized industrial enterprises.

  3. Adaptability allows maintain an optimal level of internal processes flow in the system, while the system itself acquires stability and ability to survive in the existing environment.

Belov G. O., Stadnik D. M. Gear-type pump design procedure development providing its dynamic loading reduction. Aerospace MAI Journal, 2017, vol. 24, no 2, pp. 7-14.

Aerospace hydraulic systems generate pressure and flow-rate oscillations in the course of their operation, which in its turn leads to vibrations and noise level increase.

Thus, the problems of the study can be formulated as follows:

  1. Development of a model of hydrodynamic processes in gear-type pump, accounting for dynamic processes in a locked volume, two-phase nature and pressure oscillations of the working substance.

  2. Balancing grooves profile in gear-type pump front foot bearing design, allowing working substance overpressure in the locked volume.

  3. Determine experimentally the effectiveness of design procedures on the improvement of gear-type pump dynamic characteristics.

The authors realized numeric model with allowance for the two-phase nature of the flow and the pump's design features using programming language Delphi. Computations allowed obtain the fuel consumption patterns at the input and output of the pump, as well as cavitation phenomenon in the locked volume. Based on computation results, a technique for balancing grooves in front foot bearing was developed.

The effectiveness of such changes in construction was demonstrated experimentally at the Institute of Machine Acoustics. Using scada system LMS Mobile the authors fixated reduction of vibration, pressure oscillations and noise for NMSh-5-25-4 pump.

Thus, all the planned tasks of the research were fulfilled.

The results of this work were implemented at Wroclaw technical university (Wroclaw, Poland) and Institute of Machine Acoustics (Samara, Russia).

Komissarenko A. I., Kuznetsov V. M., Simakov S. Y., Muraschev A. A. Meteorological rocket “MERA”. Aerospace MAI Journal, 2017, vol. 24, no 2, pp. 15-23.

Rockets MMP-06 and MMP-06M are nowadays are the most popular rockets in Russia, while the “Dart” system is the most popular rocket in the USA.

The MMP-06M reentry vehicle maximum flight altitude is 60-80 kilometers.

Since 1988 the rocket is employed for the wind velocity and temperature measurement in upper atmosphere.

The rocket is equipped with the engine with a steel body as a thruster, which results in low thrusted efficiency for modern technology state-of-the-art (the empty engine with stabilizers to fuel weight ratio equals 0.56).

The altitude of probing is relatively low, and equals 80 km, and probing at the flight upward trajectory is impossible.

To avoid the above-mentioned drawbacks, the SC Instrumentation Design Bureau under contract with the Ministry of Environmental Monitoring and Research Developed “Mera” meteorological rocket with probing altitude greater than 100 km.

To serve as a thruster the surface-to-air bi-caliber missile engine, using solid propellant (with density impulse of 240 kgFs/kg) was developed and finished-off, with fiberglass body and the empty engine with stabilizers to fuel weight ratio of 0.3. It allows significantly reduce initial weight of the rocket and its size.

The meteorological rocket “Mera” was designed based on the above said surface-to-air bi-caliber missile engine, and MMP-06M “Dart” as reentry vehicle cruise component.

To provide requirements fulfillment (achieving altitudes over 100 km) meteorological rocket “Mera” has two-stage structure with passive cruise component and equipped with a booster.

Measuring and servicing equipment is allocated in the cruise component in the form of a container. The cruise component is equipped with parachute in a separate container.

On the assumption of stiffness conditions and required temperature the body of the cruise component is protected by combined coating.

To ensure radio signal of the equipment propagation the cruise component is equipped with radio-transparent insertion.

To ensure aerodynamic stability the cruise component is equipped with asymmetrical consoles.

The paper presents aerodynamic, weight, inertial and ballistic characteristics, impact zones and separated engine trajectories, as well as cruise component impact zones.

Dukhopel'nikov D. V., Vorob'ev E. V., Ivakhnenko S. G. Ion flux control in hall accelerators. Aerospace MAI Journal, 2017, vol. 24, no 2, pp. 24-30.

Hall thrusters are widely used for satellite orbit correction and marching operations for altitude change. At the same time the accelerators designed according to similar schemes acquired wide spread occurrence in vacuum ion plasma technologies as ion-cleaning and nano-scale surface treatment systems.

In a first approximation, in the design of such devices it is assumed that the magnetic field does not affect the ions movement in the accelerating channel. Actually, the ions deflected slightly in azimuthal direction under magnetic field impact, whereby the beam acquires the shape of one-sheet hyperboloid. With the thrusters, it might lead to the plume spread, derating and angular momentum occurrence. This leads to significant divergence of the ion beam in the technological accelerators operating on relatively lightweight argon. For surface cleaning before coating deposition such

divergence of circular beam is acceptable, since maximum processing area is required. However, for dimensional ion beam processing narrow ion beams with Gauss ion current density distribution are required. At the same time, effect of the ion azimuthal deviation does not allow focusing the ion beam of the Hall accelerator only by coning the walls of the acceleration channel.

In this paper, additional magnetic pole was installed for focusing ion beam into a spot with Gauss ion current density distribution along radius at the outlet of the cone acceleration channel of the ion source. This magnet pole produced the magnetic field which vector is opposed to magnetic field vector in the channel. Ion beam in the additional magnetic pole area turns in azimuthal direction, opposite to its turn in the acceleration chamber. As a result, the beam is coned and focused at a specified distance into a spot with maximum ion current density concentrated in the center.

The paper formulates the criterion of optimum ion beam focusing in accelerator with anode layer. The ion current density distribution along the radius of the focused ion beam was measured with the accelerator experimental sample. It was shown that the installation of additional magnetic pole allows focusing the ion beam completely.

The obtained results can be used in the design of ion sources for punctual ion-beam machining of the details for optical and electronic industry.

Ezrokhi Y. A., Kalenskii S. M., Morzeeva T. A., Kizeev I. S. Distributed power-plant concept with gas drive of external fan module analysis. Aerospace MAI Journal, 2017, vol. 24, no 2, pp. 31-41.

The paper analyzes the concept of distributed power plant (DPP) for prospective long haul passenger aircraft. This DPP is intended to provide deeper integration of a power plant and an aircraft, as well as increase its fuel efficiency.

Possible variants of drive realization for external fan modules, as a constituent part of the distributed DPP, are presented. The necessity of considering the gas drive, realized by introducing an additional transient duct between the turbine of dual-flow turbojet engine and the turbine of the external fan.

The paper presents the preliminary analysis of DPP with gas drive of an external fan module developing possibility in the simplest for realization version incorporating a single external module.

The authors developed the technique for numerical study, carried out the evaluation of the specified DPP parameters under various values of total pressure losses in the transient duct to the external fan module and performed preliminary evaluation of the distributed power plant weight.

Further development of the considered distributed power-plant concept the additional gas heating in the transient duct while the take-off mode is offered. Additional calculations of new type engines are carried out, and estimation of new distributed power-plant structure parameters improvement possibilities is made.

In conclusion, comparison of the considered distributed power plant structures basic parameters at various degree of total pressure losses in the transient gas duct to the turbine of the external fan module is presented. The conclusion is drawn on the necessity to assign transient ducts and intermediate heating systems technologies to critical category.

Biruykov V. I., Kochetkov Y. M., Zenin E. S. Determination of thrust specific impulse losses occurring due to chemical non-equilibrium in aircraft power plant. Aerospace MAI Journal, 2017, vol. 24, no 2, pp. 42-49.

As usual, thermodynamics and statistical mechanics deal with the problems, which suppose a system to be in equilibrium. Thus, the implemented mathematical tools could be rather conditional in cases of the systems with irreversible chemical reactions, as well as gas flows with thermodynamic non-equilibrium. Depending on how the system differs fr om equilibrium state, the great majority of practical solutions of combustion problems referenced in classical literature this condition is observed. In contrary cases, i. e. significant non-equilibrium in combustion in high-speed flows problems, and detonation in particular, the variation fr om stationary proliferation of chemical reaction fronts is unreasonably neglected. The traditional combustion in gas flows problem statement is unidimensional and based on consumption, momentum and energy conservation laws. Effects of viscosity force and thermal conduction are accounted for herewith. The basic difference of idealization consists in supposition of consistency and averaging of thermal capacitance value under constant pressure and volume. However, these values are dependent from chemical components composition and temperature in particular. Viscosity factor and other factors, characterizing transition are also functions of gas mixtures composition and temperature. As a consequence, gas constant and constitutive equation differ significantly from the idealized form. For complete analytical description of combustion gas dynamics, accounting for mutual diffusion of chemical components, regularities of components vanishing and occurring of new ones, as well as evaluation of total heat emission due to the completion of chemical reactions are required.

Systematic numerical studies of homogeneous and heterogeneous chemically non-equilibrium gas flows in aircraft power station nozzles are already conducted for many years. Various authors obtained results for combustion products of a number of fuels employed in aviation and rocketry. However, calculations of such flows do not satisfy modern practical requirements in all respects. Their main disadvantage consists in orientation on strictly defined set of substances and chemical reactions. To other shortcomings are neglecting the small concentrations of the reacting components, which compels to coarsen recombination mechanism. The variety of propulsion installations designs predetermines the presence of various units with non-equilibrium combustion in the area of lean and reach mixtures, such as gas generators; liquid rocket engines combustion chambers with complex mixture-formation systems; a number of pressurization systems and gas passages with gas flows; combustion chambers and afterburners of air-jet engines.

To a certain extent, determination of specific impulse losses in rocket solid engines due to chemical non-equilibrium with allowance for its effect on formation of Al2O3 and ALN condensed particles presents practical interest. The nowadays reality is the study of combustion detonation mode, wherein flows idealization is unjustified due to high conversion rates, and chemical reactions are principally non-equilibrium. The paper presents gas flows with non-equilibrium chemical reactions modeling in the form of conservation equations: uniformity of energy and impulses, wh ere impulses are presented as a product of gas mixtures density scalar and their velocity vector. As a result, in addition to the equation in Navier-Stokes form the authors obtained one more member, accounting for relaxation processes in thermodynamic system. Based on carried out analysis of the law of mass action the authors obtained interrelation between Gibbs thermodynamic potential with the equation member, accounting for non-equilibrium in gas flows with specified content in the form of normalized function. Based on it, the authors offer an engineering design procedure of a rocket engine specific thrust losses (aircraft power plant) caused by chemical non-equilibrium. The values of combustion products equilibrium and frozen compositions for the specified fuels are used for computation of adiabatic coefficients for lim it cases and normalized function. The paper presents graphs illustrating the computations for a wide spectrum of combustion products compositions. The examples of computation results of specific impulse for various cross sections of rocket engines nozzles.

The engineering method for calculation of the thrust specific impulse losses occurring due to chemical non-equilibrium allows estimate adequately their contribution to the common share of losses.

Orlov M. Y., Anisimov V. M. Computational study of compressor operation mode effect on gas turbine engine combustion chamber processes. Aerospace MAI Journal, 2017, vol. 24, no 2, pp. 50-56.

Improvement of modern GTE and power plants directly related to improvement of the combustion chamber. However, combustion chamber is one of the most problematic parts in terms of the design and finishing-out. To solve these problems the authors developed the technique for performing common computations of the compressor and combustion chamber together. In the framework of this work this method was used studying the effect of flow unevenness, occurring behind the compressor blades, and on combustion chamber workflow.

The method has been further developed in the way of implementation of common mesh model for the compressor, the combustor, and working out the boundary conditions setting principles. Geometrical model consists of four different geometrical volumes: guide vanes of the penultimate stage of high-pressure compressor, the impeller and guide vanes of the last stage and the flow path of combustion chamber. The sector of compressor and combustor was used instead of full-sized model to reduce calculation time. The sector angle kept constant for compressor and combustor.

Three-dimensional modeling software package Ansys Fluent was used for simulation of common operation of compressor and combustion chamber, since the combustion processes simulation was tested and verified for this package. Mathematical model and boundary conditions were set after mesh generation. Mathematical model included different calculative models, which were necessary for the combustion simulation. Boundary conditions were specified by temperature and pressure of the flow at the inlet and of the fuel. The flow blows the guide vanes at a certain angle. Hence, the direction vectors were set in cylindrical coordinates. The simulation was carried out in non-stationary arrangement. Thus, the certain time step and number of time steps, which are necessary for convergence, were set. The simulations were carried out for three engine operation modes (nominal, 0.7 of nominal and 0.5 of nominal regimes) with and without compressor. The least effect of the compressor detected at the the engine nominal mode, and the the largest was detected at 0.5 of the nominal. The obtained results were compared with the results from simulation without compressor.

Simulations revealed that that blade wakes extend up to the flame tube head. These wakes change the flame tongue, pressure field, temperature and velocity in the recirculation-mixing zone. It can affect combustion efficiency, ecological performance and temperature field at the combustor outlet. Thus, the simulations, which accounted for combustion chamber and compressor, more fully represent the characteristics of the working process of the combustion chamber and increase the efficiency of new products design.

Baklanov A. V. Low-emission combustion chamber of diffusion type employing micro flame burning process for converted aircraft gas turbine engine. Aerospace MAI Journal, 2017, vol. 24, no 2, pp. 57-68.

Combustion of fossil fuel is accompanied by a number of toxic agents' formation. Nitrogen oxide and carbon monoxide are the most ecologically destructive, for they hurtfully affect humans and the environment. For these reasons the paper solves the topical problem on creating a diffusion combustion chamber for a converted aircraft gas turbine.

For the purpose of efficient aircraft engine combustion chamber conversion from fluid to gaseous fuel, the author proposes the combustion chamber design and complex approach, including of engineering and design studies and experimental studies.

The experimental method includes three stages. At the first stage, the butners' outlet parameters are defined. For this purpose, a workbench for determining a burner throughput capacity and obtaining concentration pattern of the air-fuel mixture in swirling jet burner outlet. CO2 was used as a gas fed to the fuel ducts, instead of methane. Concentrations distribution over the sections after the burner presents the pattern, allowing trace the CO2 concentration level variation dynamics in whole area of measurements and in each point of the swirling jet. It allows evaluate the quality of air-fuel mixture preparation. The burner throughput capacity was evaluated at various pressure differences. Based on the performed work, selection of the burner geometry for implementation in the compustion chamber was performed.

While implementation of the flame tube head with a large number of atomizers, fuel distribution uniformity ensuring is of especial importance. It provides stable combustion process and mixture homogeneity at the combustion zone inlet. To determine the flame tube head flowrate characteristics, an installation with compressed air delivered to fuel ducts was implemented. Evaluation of air throughput deviation from its average value was carried out. It allowed working out the flame tube head from fuel feed ducts dimensions' optimization viewpoint.

The next stage consists in working with a full size combustion chamber. This stage includes two trends. The first one is the pressure loss determination in the combustion chamber, while the second one is determination of the non-uniformity of the outlet temperature field. Selection of combustor can degree of opening and air distribution along its length to provide optimal pressure losses and temperature field.

At the final stage the combustion chamber as a part of the engine functioning test was carried out. The engine throttle performance characterization and measuring the exhaust emissions of the engine was performed.

In accordance with the results of the studies, conclusions were made that the realized complex approach to toxic agents emission reduction allowed design the combustion chamber reducing nitrogen oxide emission by 40% and carbon oxides by 20% compared to a stock combustion chamber.

Zakharov I. V., Trubnikov A. A., Reshetnikov D. A. Functional control software/hardware complex master side model. Aerospace MAI Journal, 2017, vol. 24, no 2, pp. 69-78.

With implementation of existing methodological support of regular automated verification systems (AVS) the state and performance of a short-range missile of air-to-air class (AAM SRM) control system sensors are unobservable. Thus, while regular AVS typical control algorithms realization technical state of control system sensors, such as linear accelerometers (LA), or angular accelerometers (AA) can be estimated through indirect parameters, without their basic parameters determination (transfer factor, etc.). This could significantly reduce methodological fidelity of guidance system control.

To solve the above said problem the authors offer implementation of functional control (FC) method. This method can be realized based on software/hardware complex (SHC).

The paper suggests scientific basics of functional control. They are stipulated by implementation of harmonic balance of automated control theory. The FC structure, organized by duplication method, was used to realize AAM SRM guidance system FC control sensors.

To minimize the control structure dimensionality at the inputs a single primary impact  on the missile guidance system is applied using harmonic oscillation workbench (HOW). To close FC links one should be aware of HOW functioning as a master side of SHC.

HOW is the main preset part of SHC, generating a single primary stimulating effect  on the missile during FC of its guidance system sensors. To close FC system it is necessary to set correct stimulating action on an FC object. It is necessary herewith to eliminate FC resonant mode, and ensure FC main sensors functioning in linear range, i. e. exclude: guidance system signals overload limiting for LAs; missile body spin velocity limiting for AA detection unit, as well as angular target tracking rate and locating angle limiting for target-seeking head (TSH).

To ensure harmonic oscillation “comfort mode” for the missile guidance system, selection and adjustment of HOSs design values is carried out. For this purpose, the developed SHC FC master side model is used. In addition, the developed model is used for characterization of secondary stimulating effects on HOW, LA and AA detection units and determination of the signals of their reactions.

The process of HOW operation can be represented by a certain model in Laplace operator form. This model includes oscillating and measuring loops. The oscillating circuit dynamic model represents an oscillating link with time constant and damping factor, as well as nonlinearity of saturation type with known parameters, stipulated by HOW design specifics.

Measuring loop includes axial power transmission (PT) and inertia-free angular sensor. PT is free of reduction elements, and its gain KPT = 1.

The conducted experiments on a certain HOW embodiment confirmed the adequacy and performance capacity the developed models of SHC FC master side, as well as correctness of the HOW design values, allowing eliminate AAM SRM guidance system's signals limiting and their termination to stop.

Al'bokrinova A. S., Grumondz V. T. Gliding unmanned aerial vehicle flight dynamics at low speed and launch altitudes. Aerospace MAI Journal, 2017, vol. 24, no 2, pp. 79-85.

The authors conduct studies of gliding unmanned flying vehicle (UAV) flight dynamics at low speed and launch altitude. In the case under consideration the UAV flight dynamics significantly depend on initial flight speed and initial flight altitude which determine the total UAV energy and, as consequence, UAV's dynamic capacity while moving along the trajectory.

The paper considers the following two problems:

  1. Maximum flying range provision under initial UAV motional energy limitations.

  2. The UAV stability and maneuverability provision at all flight stages.

We assume the UAV is equipped with a certain booster engine with fixed total impulse, which can be realized by various thrust variation functions in the course of UAV movement.

Much attention was paid to the study of launching conditions and thrust behavior at the initial trajectory portion impact on the flight range under gross thrust impulse limitation, as well as studying of various possible technological deviations of thrust vector direction from UAV axis of roll impact on movement stability and UAV launching safety. The last problem was considered in the form of the following two problems:

  • ensuring such UAV angular stability at the initial passive trajectory segment, which would guarantee UAV angular orientation, eliminating the possibility of UAV collision with the carrier by the time of its engine firing;

  • ensuring the possibility of disturbances parrying, which occur during engine operation at the active trajectory segment and stipulated by technological errors of its mounting on the UAV. The results of the study revealed that the last factor could affect negatively as well on the UAV total flight range.

We assume that the UAV is launched in undisturbed air conditions so that at the starting moment it is not subjected to the additional aerodynamic impact, while the carrier is moving at constant altitude with constant speed. The authors developed a mathematical model of UAV spatial motion all over the flight. The control system accounts for pitch angle and angular velocity deviations. Solid fuel accelerator with fixed thrust impulse value, variable thrust value and operating time is considered as a boost engine. A time of engine ignition was computed. Movement parameters at the initial trajectory segment, booster thrust variation functions impact on the flight range and booster thrust misalignment impact on the UAV movement parameters and stabilization were evaluated. Extreme (guaranteed) values of solid fuel booster thrust misalignment caused by technological errors while booster manufacturing and mounting on the UAV ensuring the UAV flight safety at two stages – controlled flight without thrust and controlled flight with operating booster were obtained.

Zaichik L. E., Desyatnik P. A., Zhelonkin V. I., Zhelonkin M. V., Tkachenko O. I., Yashin Y. P. Mobility effect of flight simulator cabin on aircraft in-flight refueling problem modeling. Aerospace MAI Journal, 2017, vol. 24, no 2, pp. 86-94.

One of the serious problems of flight simulation with flight simulators consists in reproduction of accelerations experienced by a pilot while in flight, which play an important role in piloting. The paper considers this problem in the context of aircraft in-flight refueling. The goal of the study is quality estimation of cabin movability over various degrees of freedom effect on piloting, pilots nature of action and his judgement on simulated accelerations degree of adequacy to real flying conditions.

Experiments were conducted with TsAGI PSPK-102 flight simulator containing cabin with six degree- of-freedom mobility, collimated visualization system, instrument display, side-stick control with electrical loading system, and thrust control levers. The authors developed the in-flight refueling task simulation technique using flight simulator with movable cabin. The problem of cabin mobility system control algorithms optimization was fulfilled for the considered task.

The pilot's task consisted in performing closing-in with the refueling tanker and carry out the refueling cone in the course of the flight. Experiments were conducted with participation of an Honored military pilot, who had wide practical experience of refueling tasks in real flight conditions.

Experimental data on the accelerations effect on unbiased indicators of the cone tracking accuracy, pilots actions characteristics and aircraft movement parameters were obtained.

The study demonstrates that reproduction of accelerations affecting a pilot significantly increases the adequacy of in-flight refueling problem simulation to a real flight. According to the pilot's, opinion axial accelerations exert the strongest effect on refueling task.

Nevertheless, reproduction of vertical and lateral accelerations in the course of flight simulation plays an important role as well. The obtained objective data and the pilot's opinion accord well with overloads and angular accelerations over various degrees of freedom significance analysis performed based on earlier developed theoretical approach to the accelerations impact on piloting.

Tischenko L. A., Kovalev A. A., Chizhikov S. V. Basic process operations parameters impact of silicon electronic lens manufacturing and its storage conditions effect on electron beam shape studying. Aerospace MAI Journal, 2017, vol. 24, no 2, pp. 95-103.

Lithographic processing is one of the key operations of technological processes while semiconductor devices' and integrated circuits manufacturing. Its parameters effect strongly the precision of the devices structure creation, and, as a consequence, its output characteristics. Multi beam lithography is implemented in particular. Its technological equipment uses silicon electronic lenses for electron beam control, which electronic and optical parameters affect the accuracy of the manufactured product structure and, as consequence, their output characteristics.

The paper tackles the topical problem of ensuring the specified electro-optical parameters of electrostatic lens (including geometric sectional shape of electron beam) during its production, storage, and transportation, as well as repeatability of these parameters from batch to batch.

The research object of this project is electro-static lens representing silicon plate with a plenty of holes of circular shape. The lens under study is used in technological equipment for multi-beam e-lithography for a powerful beam splitting into a multitude of beams.

The electro-static lens parameters degradation causes in length of time identification, and their elimination technique development are the main tasks of this studies.

In the course of the study, a number of operations and factors that could affect the electro-optical lens parameters was revealed. According to the results of expert evaluation of electronic lens manufacturing technological process, these factors are oxidation and chemical cleaning operations.

The results of various technological operations and factors effect on electro-optical lens parameters variation were presented. While this research a series of experiments was conducted, which considered variation of electro-optical lens parameters in length of time.

The obtained results of the studies allowed revealing possible reasons of electro-static lens parameters degradation in length of time, and developing technological recommendations to prevent this degradation.

The plan of future studies is presented.

Tamarkin M. A., Verchenko A. V., Kishko A. A. Heavy-plate materials waterjet cutting effectiveness improvement. Aerospace MAI Journal, 2017, vol. 24, no 2, pp. 104-114.

A voluminous assortment of parts, characterized by higher requirements to accuracy and reliability, is used while aircraft manufacturing. They are fabricated fr om various materials, such as steel, aluminum, plastics and composites. Special attention is payed to developing new methods of the parts manufacturing and improvement of conventional technologies.

The majority of the parts is produced by pattern cutting of various materials of 0.5 to 200 mm thickness, followed by their machining or without it. It is interlinked with the development of CNC metalworking Machinery Park, where forged pieces or form workpieces are used in increasing frequency. The main question consists in productivity and quality of cutting blanks. There is a great variety of pattern cutting methods, distinguished by productivity and principles, with peculiar advantages and disadvantages. The authors consider the hydro-abrasive cutting, which is the newest and prospective metal cutting methods.

Hydro-abrasive cutting is the most up-to-date and efficient method for obtaining either blanks or parts from plate aviation materials. The cutting process is carried out by the thin water jet with abrasive grains mixture, emitted at high (supersonic) speed under high pressure up to 6000 bar. Garnet sand with 7.5-8 hardness is used as an abrasive material. The process represents erosion destruction under impact of working jet, wh ere the abrasive cuts the chips microlayers, while water takes them away from the cutting zone. The main advantages of hydro-abrasive cutting are high productivity ensured for high cutting speed (steel up to 300 mm), the absence of residual strains at the cut edge, the possibility of cutting practically any metal and non-metal as well as the ability of cutting figured profile and irregular shape parts.

Nowadays the process of hydro-abrasive cutting is poorly studied. Theoretical dependencies accounting for all technological parameters effects for the cut ruggedness and corrugation determination, and dependencies reflecting the value of cutting jet lagging.

The quality of hydro-abrasive cutting depends on the feed rate, the thickness and type of cutting material. It was found, that feed increase reduces the quality of cutting, increases ruggedness, and the area of smooth

cutting reduces, while the corrugation and obliquity of the cut increases. Deffects caused by jet lag cutting, such as formation of a burr on the sharp outer corners, forming holes in the inner corners, overcut and undercut at the beginning of the cut are also found.

The goal of this study was to explore the effect hydro-abrasive cutting modes, namely the feed effect on the cut roughness.

After a row of experiments the samples made of three different materials with 30 mm thickness, namely, steel 30HGSA, aluminum D16, multi-layer polymer composite such as titanium-fiberglass were obtained.

When cutting the feed was changed stepwise from 5 mm/min to 120 mm/min for a sample of steel, to 200 mm/min for samples of D16, and to 160 mm/min for a sample of the composite. The ruggedness of these samples was measured at the specific areas of the cutting section.

Analysis of ruggedness dynamics allowed suggest a mathematical model of cutting surface ruggedness profile forming. The ruggedness is formed by free abrasives, which remove repeatedly the micro-chip layers. The mathematical model is proved by experimental data, as indicated by a graph of the cutting ruggedness dependence from the cutting head feed.

The experimental data and theoretical curves allow predict the cut quality of the hydro-abrasive cutting. Based on this data, the possibility arises to select the most optimal hydro-abrasive cutting mode cutting, or a certain type of defects elimination. The cutting rate optimization is possible by slowing down the feed in areas of defects formation, or ruggedness unevenness.

Belov O. A., Berdnikova N. A., Babkin A. V., Kozlov M. V., Belov D. A. Composite shape-generating tool set for spacecraft antennae reflector manufacturing. Aerospace MAI Journal, 2017, vol. 24, no 2, pp. 115-122.

Irregular shape items manufacturing from polymeric composite materials (PCM) requires the tool set, which geometry duplicates geometry of the item. The material is spread on the shape-generating tool set, and then its polymerization is carried out at the predetermined pressure and temperature that can achieve up to 200°C. In this respect, the most complicated problem while forming precision items from PCM consists in temperature deformation occurrence while polymerization process.

For years, metal hybrid tool sets have prevailed in high-precision composite parts manufacturing. A hybrid tool set has invar (nickel alloy with CLTE close to zero) shaping plate and a support structure made of some other metal with sufficient thermal conductivity. The tool set of such kind involves shape-generating plate attachment to the support structure means, which ensure the possibility of their free thermal extension. The drawback of metal tool sets consists in their high cost, low material utilisation ratio and long manufacturing cycle.

The next step in tool sets for high-precision items made of PCM evolution was creation of composite shape-generating tool sets. Fiberglass and carbon reinforced plastics are implemented for such tool set manufacturing. Its surface can be coated with ceramic or gel coat layer of precise thickness, providing minimum roughness, maintainability, and increasing the items takeoffs. Composite tool sets does not have disadvantages of their metal counterparts, though several design problems are still stay unsolved.

This paper proposes a carbon composite tool set design for satellite antenna reflector producing. The main requirements to this tool set are precision and stability of the shaping surface. Design solutions are validated by thermal and static mechanical analyses based on finite elements method. In addition, the paper presents the results of autoclave operation simulation, which allows analysing the tool set optimal positioning inside the autoclave to provide uniform heating.

Shmidt I. A., Bormalev S. V., Mekhonoshin K. A. The concept of managing configuration and organization of technological preparation for assembly production of aircraft engines. Aerospace MAI Journal, 2017, vol. 24, no 2, pp. 123-131.

Aircraft engines manufacturers face the following challenge: on the one hand to ensuring the product configuration management, and, on the other hand, the necessity of effectively employing financial, material and human resources throughout the entire aircraft engine life cycle. This problem can be fixed due to information support of aircraft aviation engines' life cycle based on software solutions such as Product Lifecycle Management (PLM) and Enterprise Resource Planning (ERP).

The aircraft engine building branch intensively employs PLM solutions developed by Siemens PLM Software at the stage of design documentation development. aircraft engines design is carried out with allowance for the methodology of 3D electronic model design (EMD) of a product with NX system under control of PLM-system TEAMCENTER.

PLM solutions are not used at the stage of the technological preparation for assembly production. The technological preparation process is oriented on implementation of paper design and technological documentation. The existing process does not link the stages of design and technological documentation development. At the stage of technological preparation of production process, the electronic structure of the product is practically never used. The technological preparation and configuration management systems depend largely on the human factor. The production planning system is not working effectively due to the absence of connection with technological regulations and assembly production process.

The product assembly efficiency can be improved by creating a unified information environment for developing design and technological documentation.

The TEAMCENTER PLM system implementation for technological preparation of aircraft engine assembly production will allow develop a unified information environment for developing the design and technological documentation. It will enable also the product's configuration management problem fixing and reducing time and costs associated with aircraft engines.

A key feature of the new business process is the TEAMCENTER system implementation at all stages of production technological process preparation, and the products configuration and assembly are carried out according to electronic technological structure and technological process using 3D visualization and step-by-step account of assembly process. Technological structure of the product will allow fixing the problem of production configuration management throughout assembling process. The technological structure data should be transferred to EPR system of production planning.

The equipment and assembly must be carried out via the MBOM and technological process using the three-dimensional visualization and operational accounting of assembly production process. The operational account will allow monitoring the production progress and providing feedback to the production planning system. The step-by-step account will ensure documenting of the product configuration requirements carrying out and forming actual product configuration.

Introduction of the TEAMCENTER PLM system while preparing the assembly production will allow solving the configuration management problem.

Formation of actual configuration will provide a solution to the product configuration control problem, such as documentation, identification, and traceability of the requirements compliance status to products at all product assembly stages.

Implementation of the 3D product models facilitates understanding of new products design, allows exclude drawing working documentation to automate the technical documentation development process and, as a result, reduce the time of assembly and manufacturing costs.

Developing a unified information environment for design and technological documentation preparation via the PLM TEAMCENTER system will provide the market launch of new products with the specified characteristics in the shortest possible time and at the lowest cost.

Gabrelyan A. S., Ivanov N. S., Kondrashov D. A., Korenchuk K. Y. Superconducting electric motor with stator ring winding. Aerospace MAI Journal, 2017, vol. 24, no 2, pp. 132-140.

One of the promising trend of modern transportation systems development is transition to electric propulsion. This is topical for aircraft industry too. However, to solve this problem it is necessary to design electric motors with high power density over 20 kW/kg. To achieve such figures of the specific power is possible only using cryogenic cooling, and modern superconducting materials.

Design of the electric motors with superconducting inductor and armature windings, will allow obtain maximum benefits in terms of weight and size. This relates to the possibility of increasing the magnetic induction value in the motor air gap, as well as with the stator linear load increase.

Design a fully superconducting electric motors is complicated by the absence of any universal computation methods, as well as a number of design features and the critical parameters of high temperature superconducting tapes nonlinearity. All this requires the development of new computation methods for such kind electric motors.

The paper presents a fully superconducting electric motor with a ring armature winding and the method of determining the its specific power and the results of finite element modeling in three-dimensional formulation.

The obtained analytical expression for the main magnetic flux allows derive an equation for the power density of HTS machines with annular armature winding. It is shown, that this power may exceed the value of 20 kW/kg.

Kiselev M. A., Ismagilov F. R., Sayakhov I. F. Electric actuators for aircraft aerofoils control. Aerospace MAI Journal, 2017, vol. 24, no 2, pp. 141-148.

While increasing the aircraft degree of electrification hydraulic drives fed by centralized fluid power systems substitution by off-line electric drives is assumed.

Translational motion power actuators with ball-and-screw gear are widely used nowadays in aircraft flaps, slats and adjustable stabilizers control systems, and operate reliably for a few minutes per flight.

In the absence of strict requirements to the dynamic characteristics of electric actuators, such as high-lift drives, simple electromechanical actuators with controllable electric motors and mechanical gear are already in use.

During the flight of an aircraft, controlled airfoils are exposed to varying loads under the influence of airflows. These loads cause significant mechanical stresses in the electromechanical actuator, leading to accelerated wear of mechanical actuator components. Another problem with the existing electric actuators is their excessive weight and size as well as difficulty to ensure compliance with the stringent operational safety requirements.

Thus, the goal of this research consists in eliminating these deficiencies and improving the energy and operating characteristics of electric actuators. It is necessary herewith to consider the operation of an electric drive either in active mode, when the energy is spent to set the running gear in motion, or in passive mode, when the running gear is fixed in a certain position and exposed to significant mechanical loads caused by aerodynamic forces.

Based on the presented aerodynamic forces calculations, we analyze the designs that solve the stated problems. These designs allow implementing both the passive and the active electric actuator modes.

We propose a design that makes electric actuators more reliable and durable while operating in the passive mode. This is achieved by removing the output arm from the deadlock position to allow a limited range of deflection and by damping vibrations and oscillations caused by aerodynamic forces within that range.

However, oscillations damping by electromechanical dampers is not always efficient, since it may result in weight and size figures increase under high mechanical loads. This problem could be solved by implementing the electric actuator structure with flexible coupling between the ball-and-screw gear and remaining actuator components in the form of modified elastic compensating clutch. This proposed flexible coupling demonstrates small weight and size figures compared to with electromechanical dampers under heavy loads. Thus, such structure can be realized also in spacecraft.

Judging from the above said, the considered electric actuator construction arrangement allows reduce its weight and size figures. The resource increase in electric actuator passive operation mode is achieved by eliminating rigid fixation of the output arm in dead spots and limited oscillations of the output arm in the operating range of a position sensor.

Lisov A. A., Chernova T. A., Gorbunov M. S. Simulation approach to the study and modelling of electrical converters degradation processes. Aerospace MAI Journal, 2017, vol. 24, no 2, pp. 149-156.

Under real operation conditions of electrical industry products, the degradation variance of their features should be allowed for. The subject of research is various kinds of electrical converters which even slight degradation results in serious technogenic disasters. The paper considers and suggests basic principles for such type of problems solution, and establishes a number of degradation variances regularities.

Characteristic parameters variances analysis allows separate out four most characteristic types of functions for the named regularities description: entire irrational functions or polynomials, fractional rational functions, and functions for processes with description. Evaluation of degradation variances simulation results supposes tabulation of the measured values and selection of such an approximating function which would ensure it the least mean square deviation fr om the tabular dependence. The OLS method ensures the best results for solving the problems of such type.

Analysis the considered functions, describing the degradation process, allows state the following: all functions have the initial value, known for the unit in use fr om the its datasheet. Thus, in the course of degradation variance studies it is expedient to examine only the function degradation variations, instead of the whole function. Initial value of the deviations function equals to zero, and its plot passes through the origin of coordinates. While determining the number of parameters of approximation functions, their number would be one less for deviations function. Thus, the order of normal OLS system reduces.

The residual resource prediction was performed based on solving non-linear equation, wh ere degradation deviation function takes the normative allowable values. While solving the equation, the function arguments lim it value should be defined as an instant of failure. The evaluation of the residual resource was performed based on the instant of failure.

Kuznetsov P. A., Stepanov O. A. Reactive power compensation automated systems application to prevent blackouts. Aerospace MAI Journal, 2017, vol. 24, no 2, pp. 157-163.

The main task of the article consists in electric power grids basic emergency modes, leading to rolling blackouts, analysis, and high-speed reactive power smooth regulator (RPR) design for the existing domestic reactive power compensation systems (RPCS).

Failures analysis at industrial enterprises and substations revealed that one of the main reasons resulting in avalanche failures and blackouts is reactive power circulation, representing an integral part of complex electromechanical mechanisms functioning. However, the excessive amount of reactive power and circulation leads to complications and serves as the cause of failures.

Various reactive power-compensating systems, such as either static (capacitor installations), or dynamic (synchronous compensators), are widely used in the industry. However, preference is given to the static ones due to low price and durability. Their implementation for the most part pursues only economic benefits, namely energy cost reduction. Nevertheless, with certain updating these installations can be implemented successfully for failures, resulting in blackouts, prevention.

On the example of compensating installation, developed by the authors, they suggest to replace one of the critical elements for the purpose of regulating properties improvement. This element is Reactive Power Regulator of foreign manufacture. It has a number of disadvantages, which presence may result in cascade fault.

The proposed new regulator is a thyristor reactive power regulator consisting of a transformer, network mode sensor, regulator, control block, thyristor switches and a filter.

This paper presents the schematic diagram and computation algorithms for voltage level sensor parameters, network mode sensor parameters and filter parameters. The computation of the unit reliability is presented either.

Vavilov V. E., Bekuzin V. I., Aiguzina V. V. High-speed slotless generator, integrated into auxiliary power unit: design and experimental research of the scalable prototype. Aerospace MAI Journal, 2017, vol. 24, no 2, pp. 164-175.

The paper presents the design and experimental research of the high-speed slotless generator scalable prototype with strip-wound stator core, integrated into auxiliary power unit. The experimental research and computer simulation of the scaled-size prototype in no-load and on-load modes were conducted. They revealed that this generator demonstrates minimum rotor losses and voltage ripples, as well as high specific energy characteristics. The high-speed slotless generator scalable prototype computer model was developed with Ansoft Maxwell software. Experimental data deviation from computer simulation results does not exceed 5%. From the results of scalable prototype computer simulation a full scale computer model of high-speed slotless generator was developed. The main parameters of the high-speed slotless generator were defined and compared with the parameters of the slot-type high-speed generator. The comparison revealed that the slotless variant demonstrated lower losses (by 600 W) with minimal weight and size parameters (not more than 0.2 kg/kW), high efficiency, minimal negative high harmonics effect, absence of the slot ripples, and the simple production technology. Thus, the obtained data shows that the high-speed slotless generator with he strip-wound stator core made of amorphous alloy can be implemented as the generator integrated into the auxiliary gearless power unit. It proves also the possibility of its application in aircraft industry.

Le D. T., Averin S. V. Simplified spice vector pulse width modulation algorithm for asynchronous motor speed control. Aerospace MAI Journal, 2017, vol. 24, no 2, pp. 176-184.

The paper presents algorithms for voltage generation at induction motor (IM) windings in vector pulse width modulation (PWM) mode while IM rotation frequency regulating. These algorithms ensure smaller amount of computation, allowing eliminate through currents in the inverter power stage. Vector PWM (VPWM) employs 8 states of inverter switches for IM control. The paper considers the possibility of extra intermediate vector states of the switches, which would not cause through currents occurrence in the inverter and allow include them into IM speed control algorithm. For through current elimination in transition periods between zero vectors and basic and non-zero basic vectors the authors suggest implementation of intermediate switches conditions of the switches, which will be operated on as vectors. Let us consider treat these vector as variables. The authors analyzed the usage of a group of vectors V01, V02, V03, V04, V05, V06 or V10, V20, V30, V40, V50, V60. It allows obtain two most promising algorithm. To control IM output voltage and frequency parameters employing PWM mode, the assemblage of the transitions between acceptable ones, when at least one syllable of a control word would be inverted, will be referred to as vector subset of permitted dangerous bilateral transitions. The simultaneous switching of both switches of a totem pole corresponds to the dangerous transition. In the suggested algorithms, the transitions are in parallel with Karnaugh map sides, which means that they do not cause through currents. Vector PWM voltage and frequency parameters will be considered only in the time interval corresponding to the half of the sector.

Simulation was carried out varying time interval of vector V0 existence for regulating generated voltage value under invariable TV4/TV6 ratio (for s the first sector). The effect of n parameter on the quality of IM drive in VPWM mode. The simulation was performed with MATLAB Simulink. Simulation results are presented.

Kalugina M. S., Remshev E. Y., Danilin G. A., Vorob'eva G. A., Pekhov V. A. Combined thermoacoustic method for titanium alloy structure modifying. Aerospace MAI Journal, 2017, vol. 24, no 2, pp. 185-196.

The paper studies the possibilities of using acoustic emission and gas-dynamic processing (thermo-acoustic processing) methods for initial check of a material (titanium alloy) instead of a traditional method, i.e. optimal heat treatment mode selection.

Implementation of thermos-acoustic processing as an extra treatment of TC6, BT16 and BT23 alloys, demonstrating low mechanical properties in the initial state, ensures grains refining and improvement of property package up to the required level.

Physical features of titanium alloys and specifics of obtaining semi-products on their basis require that a manufacturer should know and allow for these semi-products initial state (mechanical properties, microstructure, etc.) while process design. Thus, the manufacturer should possess the technique allowing promptly estimate and correct mechanical-and-physical properties of the basic material, and in certain cases of a complete spring either.

For the experiment, the authors sel ected the alloys fr om various foundries (mechanical properties, microstructure, etc.).

The presented study area of application is titanium alloys implementation for springs, employed in airspace and other special equipment manufacturing, where the quality of basic material predetermines largely the quality of a final product.

The carried out studies in the area of the basic material quality in spring production allows draw inference on the possibility of a certain initial check modernization, as unattainable part of a component manufacturing process. It is established, that acoustic emission method allows qualitatively estimate the microstructure without labor consuming estimation methods and take a decision on treatment schedule of manufacturing process. ATAP implementation as an extra processing of TC6 alloy, demonstrating low mechanical properties in its basic state, ensures grains refining and improvement of property package up to the required level.

Kyaw A. L., Artemev A. V., Rabinsky L. N., Afanas'ev A. V., Semenov N. A., Solyaev Y. O. Monolayer properties identification in carbon composite with nano-modified matrix. Aerospace MAI Journal, 2017, vol. 24, no 2, pp. 197-208.

The results of monolayer elastic and thermos-elastic characteristics identification in carbon composite samples, produced by employing of epoxy matrix containing 0.2 wt.% of fullerene soot are presented. The composite samples with reinforcing schemes [02/904/02], [+452/-454/+452], [04], [904] were fabricated by vacuum shaping. The fullerene soot was preliminary added to a binder and disperse using mechanical and ultrasonic mixing.

The composite monolayer properties were obtained based on the analysis of the results of mechanical tests of the samples with various reinforcing schemes and inverse problem solution. The multilayer properties valuations were obtained, using micro-mechanical, analytical and numerical modeling and solving corresponding averaging problems. Mori-Tanaka averaging method was used for analytical computations for cylindrical embedding problem. Numerical calculations were performed using finite elements method at representative fragments, containing unidirectional fibers. The computations used initial matrix properties values obtained from the experiments, and matrix containing the fullerene soot.

The paper demonstrates that the results of numerical and analytical computations performed to evaluate the unidirectional layer properties are sufficiently close to each other. It follows from these computations that in case of impurities agglomeration, addition of nano-filler should lead in the first place to transverse elastic modulus increase and monolayer shear modulus due to matrix tightening. Pitch module should vary insignificantly since it is defined by filler properties. With the filler addition, the monolayer Poisson ratio practically should not change. These results do not correspond with the experiment, except shear modulus increase. Unlike the predicted monolayer transverse elastic modulus increase, the experiments revealed its decrease. It follows from the experiments that monolayer Poisson ration significantly decreases, which was not predicted by computations. The obtained results demonstrated the matrix embrittlement while implementing the selected nano-modification technique and the necessity of either filler volume fraction decreasing, or changing the technique of its dispersing in the binder.

The authors plan to use identified values of composites' monolayers elastic and thermos-elastic characteristics hereafter to describe the residue stressed-deformed state of carbon composite construction elements to reveal the possibilities of reducing residual stresses and shrinkages in the structures with asymmetric reinforcing schemes, using matrixes containing carbon nanoparticles.

Bychkov A. N., Fetisov G. P., Kydralieva K. A., Sokolov E. A., Dzhardimalieva G. I. Nanocomposite materials based on metallic nanoparticles and thermoplastic polymer matrices: production and properties. Aerospace MAI Journal, 2017, vol. 24, no 2, pp. 209-222.

A line of composite materials based on low-density linear polyethylene (LDPE) thermoplastic matrices, polypropylene (PP) and metallic nanoparticles was produced by mixing in polymer melt. The results of dynamic mechanical analysis of PP based composites with metallic nanoparticles, namely the product of Co (II) acrylamide nitrate complex and 2% FeCoAAm co-crystallizatant thermolysis, within the temperature range from −50 °C to +150 °C revealed, that low concentration of nano-filler (1 wt.%) does not lead to noticeable changes in dynamic elastic modulus, nano-composite mechanical losses and loss tangent. Thermooxidative degradation results indicated the increase of thermostability for above said PP-based composites compared to the initial PP at 4 and 8 wt.% of nanoparticles.

The authors obtained nanocomposite materials based on polyolefin matrix and pre-synthesized by chemical co-deposition magnetite nanoparticles such as LDPE-Fe3O4 and PP-Fe3O4. According to X-ray diffraction analysis, the major component in the system was magnetite nanoparticles with an average size of 15 nm. These results correspond to scanning electron microscopy data. The paper demonstrates that with the increase of nanoparticles content in polymer, and with magnetite high content in particular, the elastic modulus increases, and the tensile strength value decreases. Thermal behavior analysis in the PP-Fe3O4 (at 4 wt.%) system indicates that nanocomposite thermo-oxidative degradation reduced compared to the initial polypropylene, and the temperature of maximum degradation start-up increases from 300°C to 385°C.

Composite materials based on LDPE and Al65Cu22Fe13 with alloy (0.1 to 10 wt.%) were produced. The paper demonstrates that the presence of quasi-crystalline alloy as a filler leads to composites strength properties improvement. Unlike LDPE-Fe3O4 systems, a tensile strength of LDPE-Al65Cu22Fe13 increases with low filler concentrations.

Protective action of the nanocomposite systems under test in relation to beta-radiation was studied using dose metering method. It was demonstrated that with filler content increase in LDPE-Al65Cu22Fe13 and LDPE-Fe3O4 composites beta-radiation flux attenuation occurs. A high correlation between the portion of passing beta-radiation and relative dielectric constant of composite materials based on thermoplastic polymer matrix with metal-filler was observed.

Prosvirina N. V. Development and implementation of efficient production management principles based on lean production concept at the aircraft engine-building enterprises. Aerospace MAI Journal, 2017, vol. 24, no 2, pp. 223-232.

The paper tackles the topical issues of staffing training and forms of factory organization at the aircraft building enterprises based on lean production concept. While production development in Russia and its share increase in the global market, the issue of product company optimal management comes up. The lean production program leads to creation of learning organization with stable, continuously progressing processes, aimed at searching for non-productive losses and their minimization. The lean production becomes the topmost factor of efficiency increasing, competitive stability of an enterprise and reliable technique for all kinds of all kinds of expenditures. In an aggravated competitive struggle at domestic and world markets, the key factor of Russian engine-building companies' success is associated with their flexible response to rapidly changing market demands. This requires development and implementation of a number of measures aimed at improving the efficiency of production and enabling enterprises to enter the global market as providers of competitive aircraft engines.

The main problem at domestic engine-building enterprises consists in production systems modernization. Many companies take the mass production concept as a basis of their production system, which does not meet modern industrial requirements to goods and services production, and does not take the expected effect. Thus, it is necessary to carry out the production system modernization, taking more efficient and productive system as its basis, engaging all management and stuff of the company in this process.

Effective organization of production at aircraft engine-building enterprises is a significant and special component of the competitiveness analysis due to its magnitude to production encompassing and time scale parameters of their implementation. Thus, the organization of competitive aircraft equipment manufacture should allow for all kinds of losses and expenditures, and implement efficient production system, including the great majority of methods, techniques and tools.

Churilina I. V. Cost management at the space-rocket industry enterprises. Aerospace MAI Journal, 2017, vol. 24, no 2, pp. 233-240.

The paper tackles the issue of the cost management methodological approaches enhancement.

The main purpose of the paper consists in developing the cost management financial mechanism based on EVA concept for the space-rocket industry enterprises as an instrument of increasing the enterprise financial stability and competitiveness.

System approach to financial management theoretical basics summarizing and analysis composes a methodological base of the research. While conducting the research the author employs the methods of financial analysis and forecasting, economic-mathematical modeling and expert assessment.

The author identified space-rocket industry enterprises' specifics and explored cost management methods existent in economic science. The economic value added concept is a method on which basis the financial mechanism of cost management is developed.

As a result, the indicators of the EVA concept were modified. The adaptation of the basic indicators of the EVA method to the cost structure of the space-rocket industry enterprises and the technique of calculating EVA are specific to the organization of production and the budget process. They allow identify the basic elements of economic value added cost, i. e. purchase of materials (EVAм), staff salaries (EVAт), equipment handling (EVAа) and other expenses (EVAп), characterizing the efficiency of rocket space technique production process.

Moreover, the optimal cost financing structure was identified. The results of the research were proved on the example of the space-rocket industry enterprises. Finally, we conclude that the most expensive source of funds are borrowed funds, which effective use will consist in material and other costs financing at their expense, but costs of labor and depreciation deductions is preferable to be financed from the own funds.

Gyazova M. M. Cargo ramp aircraft implementation forecasting based on simulation modeling. Aerospace MAI Journal, 2017, vol. 24, no 2, pp. 241-248.

The paper is devoted to the issues of cargo air transportation market development in Russia, and exploration of ramp cargo airplane An124-100 operation in the market. At present, this plane carries out the major part of transportation of heavy oversize cargo. The plane demonstrates a unique combination of capabilities, as it is one and only air transport for oversize cargo alternative overseas transportation. The plain allows also increase safety and reduce damage probability of cargo, compared to overseas transportation. It is capable of delivering cargo to far-out regions, where there are no auto-road and railways, horizontal loading and unloading capacity through nose and tail ramps, as well as lowering the aircraft floor and unloading without specialized external cargo-handling equipment.

To forecast economic indices of the plane of a specified type simulation model run by Vensim program. The conclusion is drawn that with growth of heavy oversize cargo air transportation demand, the necessity for organizing in Russia the serial production of aircraft equal to An-124-100 increases. Analysis of cargo transportation world market modern tendencies revealed apparent significant potential of the sector of economy in question and its direct interrelation with such factors as the degree of technological development of the country, the State participation in the trading processes and general level of economic development. The branch of group air transportations should be considered as one of aircraft industry strategic orientations totally and enjoy the State support.

Bokhoeva L. A., Kurokhtin V. Y., Perevalov A. V., Rogov V. E., Pokrovskii A. M., Chermoshentseva A. S. Helicopter structural elements and components fatigue resistance tests. Aerospace MAI Journal, 2017, vol. 24, no 1, pp. 7-16.

The paper considered fatigue resistance testing of helicopter structural elements on the example of helicopter rotor blades samples testing. Endurance testing of aircraft equipment components and structural elements consists in laboratory reproduction of external disturbances corresponding to the standard operating conditions, cyclic loading and functioning. However, these tests do not include studies related to the gradual damages accumulation leading to cracks initiation and propagation and finally to structural damage. In this regard, studying the process of cracks growth while full-scale tests of the samples presents special interest. The paper presents the brief description of blades full-scale tests process with concurrent video shooting. The samples are subjected to static loading, with subsequent additional bending load moment of variable sign. Video records of cracks growth were processed, and data on the crack subcritical growth time was obtained. This information is presented by the diagram, illustrating the crack growth time dependence of the crack growth rate. The paper analyzes measuring and test equipment used while testing for recording values of tensions occurring in the studied samples, due to bending load of variable sign applied to them. Fatigue resistance characteristics were determined, and fatigue graph was plotted. Arithmetic mean and root-mean-square deviation of endurance limit stress are obtained also.

Marakhtanov M. K., Veldanov V. A., Dukhopel'nikov D. V., Karneichik A. S., Krutov I. S., Makarov A. A. Modeling a spacecraft fracture mechanism occurring as a result of its metal components inertial explosion at collision. Aerospace MAI Journal, 2017, vol. 24, no 1, pp. 17-25.

The accidents of two Earth satellites collision when impact velocity of the spacecraft abeam reached 10.5 km/s. This velocity is several times than that required for a crystal lattice inertial explosion of the metal, constituting the spacecraft body. Inertial explosion parameters of metal components, which can occur at the contact point of the collided spacecraft, are studied. The paper demonstrates experimental and computed data on the collision velocity, causing such an explosion, as well as motion speed and explosion vaporous products temperature, reaching 22 000 K. It shows that the time necessary for metal transition from the solid state to luminous atomic-vaporous mixture reaction excitation does not exceed 2 µs, if this transition was caused by mechanical shock. Mass ratio of the exploded metal was determined. All experiments were conducted using lead samples.

Metal preserves its solid state until the metallic binding energy  is enough to preserve its crystal lattice. This energy equals to the sum of a metal heat content from the temperature T = 0 K plus evaporation heat up to the sample sublimation. Acquiring the energy of  the metal ceases to be a condensed media and passes to high temperature vapor condition. Such transition occurs while siderite or nickel meteorite collision with Earth, or spacecraft.

The experiment procedure was as follows. The lead ram tester of a cylindrical form weighted 0.027 kg, had the diameter of 14.5 mm and length of 15.2 mm. Its velocity was v = 1128 + 14 m/s. The lead target was of a parallelepiped shape of 67 × 82 × 15.5 mm and weighted 0.91 kg. The target mass remained after the lead ram tester stroke was 0.68 kg. The rest lead target mass (as well as the ram tester) evaporated.

During the experiment, the velocity of moving elements was determined by images movement on video frames, recorded by Phantom V 16 model 10 video camera. The exposure time per one frame was 1 / 156000 s-1, and the shooting speed was 25 000 frames per second.

The shock waves pattern in inertial explosion vaporous products of the two lead structures was obtained. The Mach number measured in the open air equals 2.36.

Ezrokhi Y. A., Kalenskii S. M., Kizeev I. S. Double-flow turboprop with afterburner weight indices estimation at the initial stage of its design. Aerospace MAI Journal, 2017, vol. 24, no 1, pp. 26-37.

The paper considers perspective approaches to double-flow turboprop with afterburner weight estimation technique forming at the initial stage of its design, having potential of implementation with acceptable accuracy for new generation of engines.

The authors carried out analysis of the existing weight estimation techniques with different degree of their elementwise particularization, and under various methods of main regularities selection, linking engine gas-dynamic and weight parameters. “Modular” and integral engine weight estimation techniques were considered, and weights of 16 engines were computed using these techniques.

Based on carried out analysis, the sel ected variant of integral approach was updated with allowance for gathered statistical data on new generation of turboprop mass and gas-dynamic parameters. A correction factor, characterizing the generation to which a certain engine is related according to its weight efficiency, was determined.

Recommendations on weight estimation of an engine design based on the existing gas generator were developed. These recommendations imply implementation of correlation dependencies of the engine's separate modules weights fr om its operation parameters within the framework of the developed technique.

To determine the weight of turboprop with afterburner, developed on the basis of scaled or modified gas generator, a combined technique matching up either integral or “modular” approaches was formed.

Finally, the recommendations on implementation of the formed techniques with allowance for their future development by invoking additional data, including the data on newly developed engines, are provided.

Moshkov P. A., Samokhin V. F. Noise and acoustic signature reduction methods for unmanned aerial vehicles with engine-propeller power plant. Aerospace MAI Journal, 2017, vol. 24, no 1, pp. 38-48.

In recent years, the problem of acoustic signature has become particularly actual and a topical due to the extensive use of combat aircraft systems with unmanned structures, solving decisive reconnaissance and strike tasks, for which low figures of acoustic signature ensuring is of prime importance.

The paper considers basic techniques for engine-propeller power plant noise reduction of aircraft type UAVs, including single air propellers of various structures and configuration, as well as piston engines.

Based on semi-empirical model the authors proposed equations allowing evaluate the effect of the diameter and number of blades on tonal components of the propeller noise in the condition of constant thrust, aerodynamic and geometric similarity of blade profiles, as well as the Mach number of the tip speed. Acoustic testing of Yak-18T light aircraft with two- and three-blade propellers, F30 and MAI-223M, performed at the Moscow Aviation Institute airfield, generally confirmed these equations qualitatively.

The propeller diameter decrease of a small-sized UAV with piston engine was considered as one of the options for noise and signature reduction. It was found, that the diameter decrease by 3.3% resulted in approximately 300 meters reduction of the distance to the ground checkpoint, which a small-sized UAV can approach without the possibility of being detected.

The features of acoustic pusher propellers and proposed methods for noise reduction are described. Based on the flight test the aircraft noise reduction afield technique by axial clearance increasing between the pusher propeller and the wing located in front of it was proposed. The paper demonstrates that with the considered clearance increase by an amount greater than the wing chord, the negative effect of the propeller mounting in pushing arrangement is practically eliminated.

UAVs designers can implement the engine-propeller power plant noise reduction methods, presented in the paper. Finally, the authors outlined the ways of further studies aimed at solving the problem of developing low-noise power plants for small-sized unmanned aerial vehicles.

Vorob'ev A. G., Vorob'eva S. S. Liquid low-thrust rocket engine boundary layer numerical study. Aerospace MAI Journal, 2017, vol. 24, no 1, pp. 49-56.

The subject of the work consists in numerical study of the boundary layer on the wall of the combustion chamber and nozzle of a liquid rocket thruster. Using numerical integration method of the system of differential equations describing the boundary layer, the boundary layer parameters were computed as a function of the engine operating conditions and the pressure in the combustion chamber. To close the system of boundary layer equations, the values of turbulent moment and heat transfer coefficients are calculated by determining mixing length by the equation suggested by Prandtl with Van Driyst correction.

A numerical method for the boundary layer computation was realized as a software with the working interface in Excel. The program operates with relative dimensionless parameters.

Low-thrust LRE, burning such fuel components as nitrogen tetroxide and asymmetrical dimethyl hydrazine with the thrust of 200 N, parameters served as initial data for computation. The working flow parameters were taken according to the results of thermal and gas dynamics computation with average mixture ratio “on the wall” over the length of combustion chamber.

The paper presents computation results of the boundary layer parameters for the MAI-200-1 object engine: the displacement thickness, relative velocity profile, friction coefficient, nozzle flow rate.

The change of boundary layer thickness and flow rate coefficient for the object engine, and engines with working pressure of 2 and 3 MPa were calculated. The paper made clear that an increase in the combustion chamber pressure increases the relative thickness of the boundary layer, while nozzle flow rate falls.

Kolodyazhnyi D. Y., Nagornyi V. S. Electric field effect on kerosene-air mixture combustion products temperature distribution. Aerospace MAI Journal, 2017, vol. 24, no 1, pp. 57-62.

The paper presents the results the experimental study of appropriately organized electric field effect, using electric unit for aviation kerosene impact (EUAKI), applied to kerosene flow at the nozzle inlet, on the kerosene-air mixture burnout temperature. TC-1 kerosene was used as hydrocarbon fuel. The air was fed to combustion chamber at the temperature of 150°C. Fire tests were carried out on the Samara State Aerospace University workbench.

Experiments on gas temperature at the outlet of combustion chamber gas collector characterization were performed by direct gas temperature measurement with single-point chromel-alumel thermocouple (operating temperature range from 0 to 1,100°C) shifted in the plane of the flow cross-section at the distance of 20 mm from the gas collector cutoff of combustion chamber combustor can.

The electric field parameters, such as voltage type at the EUAKI electrodes, its amplitude and frequency, and EUAKI design parameters effect on gas temperature distribution at the combustion chamber outlet while kerosene-air mixture burning. Atomizer modules herein, consisting of SPA “Salut” fuel atomizer itself and various EUAKI design with electric fields organization from different electric power supplies were varied.

It was demonstrated that implementation of EUAKI directly connected to the fuel atomizer inlet as a part of atomizer module by rubber hoses with corresponding permittivity increases the average and maximum gas temperature at the gas collector outlet up to 4.09% and 4.88% correspondingly, reduces gas temperature field non-uniformity at the combustion chamber outlet by 10.34% relative to the base.

Finogenov S. L., Kolomentsev A. I. On solar thermal rocket engine structure and parameters selection. Aerospace MAI Journal, 2017, vol. 24, no 1, pp. 63-74.

The paper considers the solar thermal rocket engine (STRE) with isothermal (one-stage) and two-stage system concentrator-absorber system (CAS). It demonstrates their characteristics in the flight version, revealing rational parameters of the CASs under consideration, and inexpediency of attaining maximum possible hydrogen heating temperatures and maximal specific impulse with higher mirror booster accuracy in both structures.

For considered STRE schemes, implementation of heated hydrogen afterburning reveals the possibility of solar concentrator size reduction together with upper stage fuel compartment size reduction. Selection of expedient parameters of CASs under consideration may shift towards less accurate mirrors with less absorber heating temperatures followed by minor deterioration of upper stage ballistic characteristics.

To enhance STRE energy characteristics the authors suggest CAS with two-stage solar emission absorber, which heating level corresponds to the irradiance level in focal light spot. The highest hydrogen heating temperature occurs in the central part of the absorber. The specific impulse herein significantly exceeds the like when employing isothermal absorber.

Two-stage absorber efficiency computation regression model, based on energy balance of heating stages, allowing obtain rational temperatures relationship corresponding to maximum absorber efficiency, as well as optimal temperatures distribution along heating stages was developed. The obtained regression dependencies can be used for computation of real STRE, operating as a apart of space upper stage, flight characteristics. The paper demonstrates STRE flight characteristics with considered CASs, defines their specific flight It was demonstrated that in case of two-stage CAS mass efficiency exceeds the like for modern liquid means of interorbital transportation more than 2.3 times.

On oxidizer excess coefficient selection in case of hydrogen afterburning it is necessary take into account that for STRE with two-stage absorber each percent of concentrator diameter decrease corresponds to about one percent of payload weight reduction. This factor should be considered while practical design of various STRE structures.

Siluyanova M. V., Chelebyan O. G. Shadow particles anemometry method implementation for aerosol characteristics behind the flame tube heads of low-emission gas turbine engines. Aerospace MAI Journal, 2017, vol. 24, no 1, pp. 75-82.

The process of the liquid fuel atomization and vaporization is of fundamental importance for the GTE combustion chambers effective operation. Normally liquid fuels are insufficiently volatile, and therefore must be dispersed in large numbers of small droplets with an increased evaporation surface area required for the ignition process and combustion of the fuel-air mixture.

The paper presents the results of a new unique shadow particles anemometry method for studying parameters of the flame spraying nozzle unit of low emission combustion chamber (LECC) of the pneumatic type. A detailed description of the PSV measurement method and calculation algorithm when processing the data is presented. The special feature of this method consists in its relation to a method of direct measurement of various aerosols characteristics and provides highly accurate measurements of parameters compared to other methods. PSV method uniqueness consists in the fact that in addition to the spray basic parameters, it allows also define the shape of the particles, by freezing the shadows of droplets images in the measuring volume of camera matrix and high-speed pulsed backlighting. Tests were conducted on a CIAM laser diagnostics workbench in the open space behind the nozzle unit with fuel (kerosene TC-1) pneumo-spraying. During the tests distribution of fuel particles over size and shape at the distance of 30 mm from the nozzle section in the cross-section of spray pattern was obtained. Implementation of a new Shadow particles anemometry method (PSV) allowed verify experimental data, obtained earlier by the phase-Doppler anemometry, and the method itself has demonstrated its efficiency and effectiveness, as measured in terms of dense aerosols.

Desyatnik P. A. Optimization of highly automated aircraft handling characteristics in directonal control channel. Aerospace MAI Journal, 2017, vol. 24, no 1, pp. 83-95.

Topicality of stability and controllability characteristics selecting methods development, when employing rudder control, is caused by a number of incidents stipulated by the directional control channel drawbacks. An aircraft controllability in directional channel is completely defined by its dynamic characteristics, sensitivity and control efficiency characteristics, as well as by the characteristics defining yaw/roll motion interaction.

The paper considers earlier developed aircraft controllability criteria in directional control channel and analyzes them from the viewpoint of applicability to modern passenger plane with advanced automation.

One of the issues tackled in the paper consists in ensuring aircraft reasonable dynamic characteristics. All existing regulatory documents usually place demands on dynamic characteristics from the viewpoint of ensuring enough response speed in aircraft control channel. However, earlier studies revealed that unreasonably high response speed could become the reason of aircraft so-called «sharp response» on pilots effort. Thus, the requirements to should have upper bound. The paper presents the technique of criterion parameter determination, allowing determine an aircraft inclination to sharp response occurrence and the ways to its elimination by relevant selection of control system characteristics.

For modern aircraft with V-shaped wing and engines mounted on pylons, parameter, defining aircraft directional and lateral motions interaction, may attain rather high values. Automation introduction allows decrease this value, so that its equivalent value, i. e. the value with account for automation operation achieves an optimal value. The paper presents control system parameters selection technique ensuring optimal yaw/roll motion interaction.

The authors envisage two criteria to determine optimal control sensitivity. One criterion allows estimate sensitivity optimality in time domain, and the other in frequency domain. Both criteria give the same accuracy of the obtained results. The paper presents detailed technique for optimality evaluation of rudder control sensitivity in relation to aircraft dynamic characteristics and control stick loading characteristics.

The developed criteria give physical vindication of directional control channel characteristics optimality. They can be applied not only for preliminary selection of characteristics in directional control channel and ways of their realization on modern highly automated aircraft, but also for evaluation of mounted on the in-service aircraft.

Gelvig M. Y. Aircraft pilots actual external field of vision charting technique. Aerospace MAI Journal, 2017, vol. 24, no 1, pp. 96-102.

An adequate external field of vision fr om the pilot's station is one of the topmost conditions of safe and comfortable aircraft control including a helicopter.

Explicit numerical values of vertical and horizontal vision angles from the main sighting point “C” are specified by regulatory documents, in particular by Aviation regulations (FAR-29). For clarity, the normative field of vision (FoV) is usually represented in the form of a chart  in rectangular axes, wh ere  и  are vertical and horizontal FoV angles respectively. The opening outlines should comply with normative chart as much as possible.

Currently used methods of view assessment, including a measuring method (natural and virtual with 3D model), are rather labor consuming, as they require human processing of measured data. Besides, with the initial data change, such as main sighting point “C” position, all the measurements must be repeated.

The author has developed an interactive technique of structural FoV plotting by means of Siemens NX8.5 – the basic 3D CAD system of the company. However, structural FoV does not take account for pilot's head mobility and human vision binocularity. It results in overestimated, sometimes impracticable, requirements for geometry of cockpit openings.

As a continuation of the above said research, the development of plotting technique for so-called actual FoV, complied with Standard 1 00444-81 and with due account of the above mentioned factors, has been carried out.

This problem was also solved by graphical method with Siemens NX8.5 CAD in a similar way as structural FoV chart plotting. As a result, actual FoV chart in rectangular coordinates has been obtained. All plotting, like structural FoV, are fully associative. With input data change, the geometry is reshaped automatically.

The author also managed to solve the problem of normative actual field-of-vision boundaries on a crew compartment surface, based on reverse combination of projections and convolutions of normative FoV boundaries in rectangular coordinate system. This allows optimize the location and form of cockpit openings at early design stages.

Tatarenko D. S., Korsakov A. A. Aircraft aiming system ballistic support algorithm based on complete ballistic model. Aerospace MAI Journal, 2017, vol. 24, no 1, pp. 103-112.

The paper deals with accuracy increase of uncontrollable aviation ground target killers implementation. It was found that the existing aiming systems use approximating dependencies in onboard ballistic algorithm, which does not allow provide high accuracy in all combat conditions due to introduction multitudes of assumptions.

As is well known, the kernel of a ballistic movement complete mathematical model consists in the system of twelve differential equations, which solution requires the set of means, ensuring its numerical functioning. These include equations describing ambient environment parameters, the system of inertial, traction and aerodynamic characteristics of ballistic objects, as well as data on initial and terminal conditions of thrown bodies' movement. Until recently, the low speed of computing facilities hampered with obtaining solution of uncontrolled air-launched weapons movement differential equations in the course of aiming. However, todays level of onboard digital technology allows overcome this shortcoming.

Therefore, in these conditions we have the possibility to realize the onboard ballistic algorithm based on numerical solution of differential equations directly onboard an aircraft in the course of aiming.

The authors analyzed the ballistic problems solution accuracy during modern aiming systems terrestrial fire, implementing approximating dependencies in onboard ballistic algorithm, and revealed their main shortcoming, namely, impossibility of ensuring high accuracy of application in all conditions of combat operation, as well as with various operating lives. New technique and ballistic support algorithm for aviation uncontrolled destruction facilities were developed based on complete ballistic movement model solution. It allowed enhance the range of tactical employment due to firing initial conditions definition, atmospheric parameters, and aviation destruction facilities movement trajectories parameters definition; nutation angle prediction; flight trajectory parameters introduction into aiming system, and angular correction computation for aviation artillery-type weapon and uncontrolled aviation missiles, with allowance for the predicted nutation angle.

Zaichik L. E., Grinev K. N., Yashin Y. P., Sorokin S. A. Control stick force characteristics effect on pilot model parameters. Aerospace MAI Journal, 2017, vol. 24, no 1, pp. 113-122.

Notwithstanding the great number of publications concerning pilot models, none of them considers the issue of stick force optimal characteristics selection. It can be explained by the fact that pilot describing function in visual signal tracking is insensitive to loading characteristics variation and does not allow reveal any regularities and their effect on pilot model parameters.

This paper is aimed at studying the effects of stick force characteristics on pilot model and its components, such as limb-manipulator and neuromuscular systems, as well as finding objective proof of loading characteristics, selected by pilots, optimality.

The paper presents recently obtained experimental data on the effect of control stick force characteristics, such as gradient of stick-force damping on pilot model parameters. The effect is analyzed based on pilot model frequency response identified in the problem of compensatory pitch motion tracking. For limb-stick and neuromuscular systems characteristics identification, input strain signal is introduced in addition to visual input signal. Frequency response characteristic computation of various pilot model components was made according to specially developed program, based on fast Fourier transform.

Analysis revealed that the force gradient variation affects neuromuscular frequency response, demonstrating thereby a pilot's adaptation to the stick force variations. Due to this, the limb-stick cutoff frequency of the open-loop system remains constant for the force gradients assessed by the pilot as optimal. The force damping does not have any significant effect on limb-stick system frequency response.

The obtained results are of regular character and contribute to theoretical and practical aspects of pilot models implementation for aircraft sensitivity evaluation.

Pashko A. D., Dontsov A. A. Model of active protection element impact on guided missile in calculated space point. Aerospace MAI Journal, 2017, vol. 24, no 1, pp. 123-131.

The paper describes the process of spatial movement of the aircraft relative to the earth coordinate system by a system of differential equations, taking into account the dynamics and kinematics of translational and rotational motion.

The aerodynamic impact of the environment on the aircraft is determined by its configuration, position of the associated coordinate system relative to the velocity of the aircraft center of mass and vector of its angular velocity. To ensure an aircraft steady state flight mode the model solved the problem of balancing, consisting in the engine thrust values, angle of attack and the deviation of the aircraft control organs selection with subsequent solution of the system of differential equations. The output variables of the model are the parameters characterizing the actual position of the aircraft in space.

The calculated missile trajectory, represented in the form of differential equations and algebraic dependencies, describes the missile guidance to the aircraft. The result is the relative distance value of the aircraft defined by the elevation angles and azimuth. The rocket direction of motion measurement is made according to the method of proportional guidance. The control system sets the missile maneuver with an overload, directly proportional to the angular velocity of the rocket-target line of sight. Thus, any time it tends to ensure the direction of the missile movement to the set-forward point.

Based on the canonical equation of motion of the center of mass of the active protection element the terms of its ejection, to deliver it to the point of space where the guided missile is situated was calculated. By simulating the flow over the active element by turbulent incident flow, using finite volume method in Ansys CFX the authors defined the ballistic coefficient of the active protection element. It allowed us to calculate the resistance function value and produce the data on trajectory and projection parameters of active protection element to the control unit.

As a result, this model allows calculating, under different tactical actions of the aircraft crew, the target miss, the orientation angles and missile speed of convergence with the aircraft. When processing simulation results one can obtain the characteristics of the missile encounter with an aircraft, as well as active element ejection parameters for its encounter with the rocket in calculated space point.

Ryapukhin A. V. Innovative technological projects in the domain of aircraft and aerospace engineering quality management. Aerospace MAI Journal, 2017, vol. 24, no 1, pp. 132-137.

The paper deals with project quality management in aerospace industry. It analyses acting domestic and foreign Standards on project management, and offers problems grouping for promotion in the field of innovative technological projects aimed at aerospace products development quality management.

The paper suggest to eliminate technological gap between Russia and European and American industrial enterprises, developing samples of advanced technology in the sphere of aviation and astronautics through implementation of practices accumulated in international Standards on separate projects management, as well as programs and portfolios managing. Innovation projects management quality increasing for National design departments should base on quality provision and management integrated system development and putting it into practice based on both ISO Standards and proper Projects and Programs management Standards. The existing classification of design performance and other indices needs to be improved.

The author envisages concepts of technology, technological innovation project and technologies transfer. Depending on complexity, technology can be included into economic turn-round. Transfer key criterion means technology working efficiency in terms of technological complexity. The State puts forward the problem of research carrying out on improving the system of innovative projects in the area of technological processes provision and management. Its solution options are significant of applied research planning procedure renewal, and Hi-Tech innovation projects realization technological provision program substantiation at the life cycle early stages, and innovative constructive-technological solutions marketability preliminary estimation, as well as process design planning optimization based on Hi-Tech projects with allowance for economical production.

The results of the study can be implemented for new Hi-Tech innovation projects management quality methodology development.

The paper practical has practical importance for acting quality management system at the aerospace industry enterprises improvement. It can be implemented also in the process of specialists training in the innovation projects management sphere.

Reznikov S. B., Kharchenko I. A., Marchenko M. V., Zhegov N. A. Transformer multifunction switched mode converters for onboard airspace power sources. Aerospace MAI Journal, 2017, vol. 24, no 1, pp. 138-145.

The paper envisages circuit solutions for transformer multifunction switched mode converters meant for uninterruptible power sources as constituent parts of onboard aerospace electric power facilities and electric power supply systems. All solutions are protected by the Russian Federation priority. The paper is of interest for a wide range of specialists working in the field of aerospace onboard power electronic equipment design.

To power actuating brushless motors (aircraft onboard equipment in particular) the voltage higher than that provided by batteries, solar of supercapacitor (ionistor) elements is required due to the necessity of varying magnetic field space forming (either circular or linear) by currents flowing through flexible wires. Ensuring relatively higher voltage level only through series (stacked) low-voltage units, or series connection of the above said batteries with significant supply currents is hampered technologically, and leads to mass and size, reliability and cost parameters reduction. For example, in case of “stacked” units leads burning-out (or break) they should be shunted by diodes with low-voltage conducting junctions. In case of breakdown at the leads of a parallel link, it should be provided with disconnecting fuses. Thus, to increase the voltage level of a primary relatively low-voltage source switched mode converters (SMC or DC/DC converters) based on field-effect transistor switches (MOSFET) with low Rds(on) should be used. They should herewith be reversible to provide feeding batteries intensive charging. As a rule, such converters are included in so-called secondary power sources, or stand-by uninterruptible power sources (UPS) fed by batteries [1].

Aerospace uninterruptible power sources included in onboard electric power facilities and electric power supply systems, acquire primary energy from chemical or solar batteries, either form newly developed super capacitor (ionistor) batteries with relatively low voltage (28 V). As a rule, the UPS output voltages herewith are higher DC voltages (such as 135 V, 270 V, or 540 V), or higher AC three-phase (or single phase) voltage (stabilized or regulated) of constant or regulated frequency (e. g., within the limits of 115/200 V, 360-800 Hz, or 0-115 V, 0-400 Hz). Besides, UPS should provide fast feeding battery charging (accumulator or supercapacitor).

In this regard, at least specific requirements are placed on the above mentioned UPSs.

Kosolapov D. V., Kurbatkina E. I., Shavnev A. A. Mechanical alloying process specifics and factors affecting the processed material properties. Aerospace MAI Journal, 2017, vol. 24, no 1, pp. 146-159.

This article describes one of the powder metallurgy methods, namely mechanical alloying (MA), used for composite materials production. MA is a solid-phase process of deformation impact on the powder material. MA changes the structure and properties of the processed materials. The authors analyzed the effect of technological modes on the process of mechanical alloying. They described, in particular, the main types of ball mills, employed for МА process carrying out. The authors examined the effect of the impurities on fractional, chemical and phase composition of composite granules, which can both accelerate supersaturated solid solutions and amorphous phases building-up process, and prevent diffusion to form amorphous oxides and phases with work material. The authors demonstrated in the paper that the shape of the shape of the container and grinding bodies could also affect the MA process and its results, as well as MA effectiveness and fractional composition in particular. Shape, size and material of the grinding bodies selection depends on several factors. Generally speaking, the grinding bodies should correspond to two basic requirements, namely, they should possess developed superficial area to provide contact with the processed material, and have enough weight to possess enough energy for processed particles grinding. The grinding media can be not only in the form of a globe, but also cylindrical et. On the Al-50% Ta system example the authors envisaged the effect of globes weight to the weight of a material ratio on the MA process.

The authors demonstrated also that the MA rate is one of the most important parameters affecting the process of the processed material grains mixing and grinding, chemical reactions process and phase transformations occurring in solid phase. It is well known, that the greater the mill rotation speed, the greater the kinetic energy transferred to the bodies and particles, and, hence, the intensity of the process increased. However, excessively high rates might cause a number of complications, such as grinding bodies high degree abrading and overheating either of a drum mill, of processed material. The authors also studied the issue of temperature effect on phase and structural transformations during technological process. They noted, that high temperature contributes to phase transitions and chemical interaction, while lower temperature works towards nanocrystalline state and metastable phases forming, as well as allows process plastic materials effectively.

Thus, the materials presented in the paper help not only to select the initial charge materials processing mode, but also predict the obtained results.

Umarova O. Z., Pozhoga V. A., Buranshina R. R. Structure formation and mechanical properties of heat-resistant alloy based on titanium aluminide under heat treatment. Aerospace MAI Journal, 2017, vol. 24, no 1, pp. 160-169.

Titanium intermetallic Ti2AlNb (orthorhombic phase) based allows are promising materials for gas-turbine engine elements manufacturing operating at the temperatures of 650 — 700°С instead of heat-resistant steel due to their high specific properties, and also intermetallic super- - and  -alloys possessing low technological plasticity.

Orthorhombic alloys phase composition and structure strongly affect the final mechanical and technological properties of semi-finished products, which can be controlled by certain of thermo-mechanical and thermal treatment modes. Thus, the purpose of this study consisted in studying the effect of heat treatment on the structure and properties of heat-resistant alloy based on Ti2AlNb titanium intermetallic.

In this work, the effect of various heat-treatment modes on the structure, hardness and mechanical properties of the VTI-4 alloy based on Ti2AlNb titanium aluminide was studied. The samples were subjected to heat treatment, X-ray diffraction and metallographic analyses. Besides, the hardness of samples was measured by Rockwell method, and mechanical tensile tests were carried out at room temperature.

Based on conducted studies, data on the temperature ranges of phase regions in the alloy was obtained, and a scheme for the two-stage heat treatment was designed. It was stated, that the structure and hardness of the alloy are greatly affected by the cooling rate between the first (high-temperature) and the second (low-temperature) treatment stages. Increasing of the cooling rate from 0.01 K/sec to 10 K/sec resulted in fine-dispersed orthorhombic phase formation; the alloy hardness increased by 5 HRC units, and the strength grew by 100 MPa while maintaining a satisfactory level of 4 — 6% for the plastic properties. The cooling rate after the low-temperature stage had no effect on the alloy structure and hardness.

It was shown also, that temperature reduction of isothermal holding in the low-temperature stage by 50°C resulted in the tensile strength increase by 80 MPa, and plasticity decrease by 3%.

Designed VTI-4 alloy heat treatment modes on the example of rod semi-finished product allowed form in the alloy structure with different size of structural components. The obtained results allow also predict changes in the strength and plastic properties of other types of VTI-4 alloy semi-finished products according to the need for further forming operations.

Klimov V. G. Implementing laser pulse buildup for gte turbine rotor blades reconditioning process design development. Aerospace MAI Journal, 2017, vol. 24, no 1, pp. 170-179.

The gas turbine engine advancement goes hand in hand with the development of its basic component, namely, gas turbine as the key source of efficiency enhancement of the engine in aggregate. With each turn of gas turbine development, materials and technologies used for its manufacturing became more and more complicated and, as a consequence, expensive. Russia is one of global manufacturers of gas turbine engines.

The cost of engines for aviation and power industry applications is considerably high. Thus, on this background its reduction remains the main criterion of manufacturer's competiveness on the market. Besides, we should bear in mind that the gas turbine engines maintenance costs in the course of the engine life might exceed its original cost. Without effective maintenance technologies, manufacturing would incur permanent losses. One of the basic specifics of gas turbine engines consists in their significantly high production costs of a number of their parts and subassemblies with relatively short lifetime, requiring permanent replacement. Rotor blades present precisely these parts. They can be damaged by a great number of factors from changes in the structure to loss of geometry. The latter is the most frequent factor even in the case of insignificant geometry loss. From the maintenance technologies viewpoint turbine blades restoration is the most cost-effective, compared to the other parts of the engine. But the complexity of this task remains the major obstacle to its realization.

This article discusses the possibility of using high-temperature solder powders as wear-resistant layers applied by laser pulse buildup, as an alternative to classic wear-resistant composites with tungsten carbide admixture. These materials are undergoing testing for further pen height recovery on the example of the turbine blade of the turboprop starter for NK-12MP aircraft engine, and attaching wear-resistant to its end edge. Based on the conducted studies with Tescan VEGA3 LM electron microscope and Hardness DuraScan-10 micro-hardness meter, together with local abrasive wear tests and various powder materials, such as VPr11-40N, VPr24, VPr27 Rock-Dur 6740, analysis while pulse laser powder buildup, the authors confirmed the applicability of several solder powders as wear-resistant layers for turbine blades contact surfaces recovery. Further, comparative studies of the basic material, soldered and built-up structures of VPr11-40N (having the best figures) solder were conducted to detect hardening wear-resistant phases. The cooling rate dependencies of shaping and VPr11-40N solder strengthening phase size were revealed.

Zaharova L. F., Novikov S. V., Kudryavtsev M. S. Realization of system approach to the problem of large-scale scientific and technical competitive projects participants integration. Aerospace MAI Journal, 2017, vol. 24, no 1, pp. 180-191.

Ensuring competitiveness of the Russian Federation in the conditions of strengthening of the global competition defines the innovation development of national economy as the priority direction. This direction realization assumes dynamic and intensive development of the industry basis on development and implementation of radical, cardinal, breakthrough innovations, and primarily technology and product. This, in in its turn, imposes increasing requirements to research and manufacturing base of the Russian industry.

Development and upgrading efficiency, productivity of research and manufacturing base of industry represents difficult, complex, coordinated process of its participants' interaction. They involve all the basic, vital and concerned parties, and are aimed at innovative cycle reduction, primarily, at the stages of innovation development and innovation activity growth, and finally, holding leading global positions over key, priority trends of technological development.

Realization of this process in the framework of the Russian Federation of a scientific and technology complex development assumes continuous improvement of its organizational and economic mechanism. One of the main methods providing development of a scientific and technology structure of Russia consists in scientific and technical projects realization within the framework of the State order, requiring forming and carrying out competitive selections of such projects.

The solution of the above-stated problem proposed in this paper consists in forming specialized organizational and executive structure of a project realization within the system integration of participants-contractors working on the project.

This model of forming organizational and executive structure of the project is developed based on the criteria accounting for extra income on the project, possible from implementation of collateral and intermediate product output, while developing research and technology reserve and, thus, under otherwise equal conditions, supplements the project economy and reduces the risks in case of possible losses.

Zakharova I. V. Regional aviation claster evolution factors analysis. Aerospace MAI Journal, 2017, vol. 24, no 1, pp. 192-200.

The airlines employ flexible approach to strategic planning and conducting regular environmental monitoring of an unstable economy. The purpose of this paper was to adapt the SWOT-analysis to assess regional aviation clusters development.

This cluster incorporates enterprises, interconnected by the aircraft lifecycle: aviation enterprises, infrastructure enterprises, operating airlines, scientific and educational organizations. The method of the SWOT-analysis is applied for a specific enterprise and for the aviation cluster as a whole. The author analyzed the factors of the external and internal environment of the Ulyanovsk region aviation cluster.

In the studying process the priority of strategic decisions and the most significant capabilities of this socio-economic system, presented the basic economy indicators of this cluster was identified.

The study proved that using the SWOT-analysis requires quantification and ranking of external and internal environment factors of the regional aviation cluster.

Quantitative correlation of the factors reveals negative phenomena in the external environment of the aviation cluster. The paper offers the expert evaluation of factors according to four criteria such as, factors rating calculations, the relationship of the identified external and internal environment factors of the aviation cluster.

The greatest threats for the Ulyanovsk region aviation cluster are as follows: the gross regional product decline, the risk of the growth rates of loans to airlines, severer tax environment for business, dependence of the airlines development from State support. Risks that occur quickly and unexpectedly, devastating to the economic system. If adverse factors are inevitable, but not instantaneous (for example, the outflow of the region qualified personnel, reduction of the population), the production is adapting to them.

The research has practical value due to quantitative justification of the priority risks enabling the company to direct the limited resources more precisely.

Aminova G. A., Tikhonov G. V. Innovative-investment activities organization and management in small business. Aerospace MAI Journal, 2017, vol. 24, no 1, pp. 201-206.

The important role in development stability and efficiency enhancement of national economy belongs to small and medium business (SMB), as an echelon of economic dynamics. This is confirmed by the experience of economically developed countries, where the share of small businesses constitutes 56% of GDP. It should be noted that the of SMEs successful development in these countries became possible due to active government support (tax, legal, organizational, personnel, etc.). Unfortunately, in Russia small business is developing slowly, and one of the important reasons consists in the the lack of systematic State support. Today, in the conditions of economic crisis caused by the endless sanctions, special attention should be given to small businesses in the manufacturing industries, especially in machine-building industry. All the more so, in these industries, small business accounts for only 15-16% of all active small businesses. It is important to note, that development in these sectors should be based on close cooperation with large corporations. In this situation, small business can take the risk of the releasing new prototypes of high-tech industrial products. They can also take over the production of components for large enterprises, thereby reducing costs. Organization of small businesses in these sectors requires a fundamentally new approach. At the stage of economy modernization the SMBs need a more sophisticated system of Government support, which should include: development of programs for the development of SMBs cooperative relationships with large manufacturing structures, creating conditions for access to the scientific and technological achievements; assistance in professional staff training and retraining. Thus, for radical strengthening of small and medium business role in manufacturing industry it is necessary to develop a fundamentally new strategy of state support, that will contribute to the development of the organization and management of innovative-investment activities in small business.

Volgina K. M., Mineeva K. I., Nemchinov O. A. The ways to improve the transport and logistics activities of aerospace cluster enterprises. Aerospace MAI Journal, 2017, vol. 24, no 1, pp. 207-217.

Cluster policy has become the dominant trend in the development of many regions. The modern model of aviation industry enterprises consolidation assumes parceling of enterprises on several levels, such as suppliers of raw materials; suppliers of parts and components; suppliers of components and assemblies; sub-integrator; final integrator. Aerospace Cluster of Samara region is one of the high-tech sectors of regional economy. Bearing manufacturing industry seems to be interesting and prospective, since they present the represent components of every rotating mechanism, implemented in every branch of production (including aircraft and helicopter). In addition, products of the plants of the branch under consideration is required either in the region, in the country or other countries.

Currently, due to the marked production growth, produced products nomenclature increase and expansion of sales network enhanced the role of logistics significantly. Competent organization of logistics operations allows obtain quite considerable cost savings, which is an important tool for industrial enterprises production and commercial management activities.

Sales revenue from products sales factor analysis over three indicators, namely, product unit annual average cost; annual production output, which, in its turn, depends on the staff on the payroll and annual average yield by a single worker. The results of the analysis allowed make a conclusion on the necessity of transportation process optimization, since transportation costs constitute significant share of product cost and final product price.

In view of wide geography of sales, the decision was made on the necessity for establishing a distribution warehouse. In the course of calculations, the optimal warehouse location based on cities remoteness from a production point and their annual claims, was determined, and transport selection for production delivery was made. The structure of intracity production distribution on the example of Samara was offered, and the travelling salesman problem was solved, using the two-parameters accrual method, namely time and distance.

The study bears the applied nature, and the work has practical value when minimizing transportation costs and embodiment of transport and logistic activities, which will lead to effectiveness enhancement of the industrial enterprise.

The study is an applied nature, and the work is of practical value while minimizing the costs of the transport and logistics activities, which will totally increase the efficiency of the entire industrial enterprise as a whole.

Galkina E. E., Daynov M. I., Metechko L. B. Occupational safety and health care system economic efficiency at aircraft manufacturing enterprises. Aerospace MAI Journal, 2017, vol. 24, no 1, pp. 218-225.

A serious problem of modernity is a problem of flight safety promotion. This problem needs to be addressed not only during operation but also during the design and manufacturing of aircraft.

With this objection in mind, it is necessary to ensure implementation of Aviation Activities Safety Management System at the enterprises of aviation industry complying with the State Standard (GOST R 55848-2013), System of Safety management (GOST R 55585-2013), Quality Management System (GOST R ISO 9001:2015), Environmental Management System (GOST R ISO 14001:2007) and Occupational Safety and Health Care Management System (GOST R 54934-2012/OHSAS 18001-2007).

Currently, many aviation enterprises are putting into practice the system of Quality Management, but Environmental Management System and Occupational Safety and Health Care Management System are not so actively introduced in industrial enterprise management practice, notwithstanding that Russias annual underproduction due to industrial diseases and injuries goes as far as one trillion rubles.

Implementation of Occupational Safety and Health Care Management System will allow reduce these huge losses.

With implementation of Occupational Safety and Health Care Management System aviation enterprises acquire real economic effects by improving working conditions, reducing the lost work time as a result of injury and disease, reducing costs of benefits andcompensation for work in harmful working conditions, improve labor productivity and production growth.

The equations for economic effect and efficiency of Occupational Safety and Health Care Management System calculation demonstrate that the enterprise acquires not only social, but also real economic effect and social and economic efficiency.

The proposed equations are recommended for implementations not only for computing the economic impact and effectiveness derived from the introduction of Occupational Safety and Health Care Management System, but also for management decisions related to the implementation at the enterprises of aviation industry.

Tyutyunnikov N. P., Shklyarchuk F. N. Determination of aerodynamic characteristics of an elastic wing with end winglets turning in its plane. Aerospace MAI Journal, 2016, vol. 23, no 4, pp. 7-16.

Possibility of control by aerodynamic characteristics of a large aspect ratio elastic wing with the end winglets turning in the wing plane is investigated.

Controlled twisting of the elastic wing in flight subjected to aerodynamic load which depends on the wing twisting can be carried out by turning of small end winglets in the wing plane with the help of a small power drive.

The coupled aeroelasticity problem is solved using mathematical model based on the discrete vortex method for calculation of aerodynamic loads on deformable wing and the wing as a thin-walled weackly-conical beam subjected to bending, transverse shear and torsion.

The numerical solution of the aeroelasticity problem is obtained for the large aspect ratio wing with the winglets turning symmetrically forward or backwards in the wing plane. Due to turning of the winglets there appear the end aerodynamic moments which cause elastic twisting of the wing and change the distribution of the aerodynamic load along the wing.

For the example of a rectangular wing with the winglets it is shown that the turning of the winglets it is shown that the turning of the winglets in the wing plane creates the end torques and significant angles of twisting of the large aspect ratio wing and as a consequence significant change of the aerodynamic loads and the wing aerodynamic characteristics.

The results of calculation show that in a case of a wing which is sufficiently pliant in twisting in the wing plane at the angle δ can be effective for control of the wing aerodynamic characteristics . In case of a wing which is sufficiently rigid in twisting the winglets become ineffective.

Romanova T. N., Paschenko O. B., Gavrilova N. Y., Shchetinin G. A. Maneuverable aircraft horizontal empennage configurations multidisciplinary optimization. Aerospace MAI Journal, 2016, vol. 23, no 4, pp. 17-25.

The presented work is dedicated to horizontal empennage multidisciplinary optimization method development. Horizontal empennage is a complex technical system, described by the equations belonged to various scientific disciplines. That is why the developed method is called multidisciplinary. The horizontal empennage efficiency can be evaluated by the values of generated pitch moment and its gradient, guaranteeing the aircraft balancing and specified flight maneuver execution. The object region analysis was carried out and various parameters combinations for optimization within the framework of the given problem were determined. We determine optimization line and specify weighting factor for each parameter. Each of the parameters can be set either as a range-parameter, fixed-value, or a set of discrete values. Besides, the simultaneous several parameter setting by a set of tuples, containing discrete parameters values, is possible. The goal function is obtained (where the number of addends is determined by the number of optimized parameters). The goal function structure allows operate with all optimized parameters combinations, regardless of the way of their specifying.

Various approaches to the horizontal empennage optimization (methods employing the Pareto principle, and the Hurwitz criterion) were studied. The analysis of the obtained results revealed the insufficient efficiency of the implemented methods. To improve the obtained results, a new multidisciplinary optimization method was developed and suggested. This method employs several evaluation functions to obtain optimal solution. The efficiency of this method is demonstrated using various data sets and input data combinations. The effect of various weighting factors values on the obtained result was studied. The result of the suggested method implementation is horizontal empennage geometrics.

The suggested method was realized in the form of a Runtime library and integrated into CAD system Siemens NX 7.5 “Modeling” environment.

Kochetkov Y. M., Borovik I. N., Podymova O. A., Mavrov V. A., Ishaev R. O. Vortex effects in Ranque-Hilsh vortex pipes. Aerospace MAI Journal, 2016, vol. 23, no 4, pp. 26-35.

The paper presents the results of computational, experimental and analytical studies of gas-dynamic processes in Ranque-Hilsch vortex tubes. The presented review considers the relevance and need for employing vortex effect for aerospace engineering. It reveals the necessity for vortex tubes with varying geometrical dimensions design for the purpose of operation range enhancement. The authors developed vortex pipe 3D model in SolidWorks system. They realized a viscous gas in vortex tube computation engineering method, and demonstrated its implementation results in gas-dynamics computing FlowSimulation pack. To solve this problem Reynolds averaged Navier-Stokes system (RANS) of equations was used in this work. All computations were performed with orthogonal computing net using finite volume method. Two-parameter model of κ — ε type allowing sufficient flow core resolution was used as turbulence model. Several basic vortex effects, such as injection, heat stratification and vortex inversion, were obtained by computation. All calculation were performed for various structural versions. A series of experiments was conducted with custom-made experimental setup. Processing of the obtained results lead to obtaining hot and cold flows productivity optimums, injection ratio, temperature stratification, as well as adiabatic and temperature efficiency.

The experimental results fully confirmed the vortex effects of obtained by engineering computational method. The authors suggest new differential equations for parameters computation in these tubes. The obtained equations establish relation with flow rotation and whirling, as well as explain the enthalpy effect. Computational and experimental as well as analytical studies should continue with regard to optimal structural concept.

Moshkov P. A., Samokhin V. F. Propeller-driven light aircraft power plant noise Integral model. Aerospace MAI Journal, 2016, vol. 23, no 4, pp. 36-44.

The extensive development of small and unmanned aircraft together with existed requirements to permissible levels of noise generated by aircraft, make the noise prediction problem afield topical for prospective aircraft with engine-propeller power plant. The main source of noise afield created by aircraft of such kind is a power plant, consisting of single propellers of various design and configuration, and piston engines.

This work integrates and develops the authors’ previously developed methods of computing separate propeller noise and the piston engine noise for solving the problem of forecasting the characteristics of light aircraft and unmanned aerial vehicles power plants’ total acoustic field.

The authors suggest a semi empirical model for noise levels, generated by aircraft piston engines in the far field, evaluation with allowance for main noise sources. The acoustic field is considered as a superposition of fields, formed by propeller and piston engine noise radiation. For propeller audio frequency levels estimated evaluation implementation of semi empirical method developed earlier by the authors is recommended. To determine propeller’s vortex noise levels, presumably dominant in the broadband noise of tractor propellers, we propose to use one of analytical models of the trailing edge noise. To calculate the acoustic performance of the piston engine we suggest to use an empirical noise model.

The paper demonstrates close agreement between computed and experimental data on power plants with tractor propellers. Experimental data on power plants noise was obtained during light aircraft of An-2, Yak-18T, MAI-223M and F30 acoustic trials under static conditions at the Moscow Aviation Institute airbase. The acoustic field herewith was supposed axisymmetric relative to the propeller axis, while test microphones were located at the ground level. It allowed exclude the interference of sound impact on measured noise levels.

The future trends of the study concerning improvement of the above mentioned method and extension the area of its application were formulated.

Vorob'eva S. S., Vorob'ev A. G. Low-thrust rocket engine with internal boundary cooling combustion chamber thermal state analysis. Aerospace MAI Journal, 2016, vol. 23, no 4, pp. 45-54.

The paper considers the issue of low-thrust liquid engine powered by nitrogen tetroxide and dimethyl hydrazine components non-symmetrical dimethyl hydrazine thermal state theoretical study with account for boundary cooling. The goal of the paper consists in analyzing the results of combustion chamber wall thermal state computation at various operating modes, such as steady-state continuous mode with stationary and non-stationary thermal field, as well as steady-state pulse mode.

Liquid rocket engine MAI-200-1 developed in the laboratory of MAI “Liquid rocket thrusters” and undergone fire tes sel ected as a subject of research.

For thermal state computation, the authors used mathematical model based on the proposition of combustion chamber wall incoming and outgoing heat flows equality. To solve non-stationary heat problem the differential Fourier-Kirchhoff heat equation in cylindrical coordinates in the case of stationary environment and the absence of internal heat sources is used. Pulse mode of the engine operation is modeled by a quasi-steady approach when non-stationary modes during engine starting and voiding are replaced by the set of stationary modes with intermediate parameters.

Oxidizer and fuel were considered as boundary cooling components to protect the combustion chamber walls fr om hot combustion products impact.

Computation results prove selection of fuel as boundary cooling component with relative boundary mass-flow rate not less than 20%. Under these engine operating conditions it will allow sustaining the wall temperature within the limits of maximum permissible temperature for ХН60ВТ material.

The combustion chamber wall thermal state for pulse operating mode with various on-time and off-time values, such as on-time of 1 s, off-time of 1 s and on-time of 0.05 s, off-time of 0.05 s were analyzed.

Presented computation results may be interesting for specialists working in the field of liquid-propellant thrusters, as well as for specialists occupied with spacecraft propulsion systems design.

Kamenskii S. S. LPRE control algorithm based on computational-experimental mathematical model using check and proof test results. Aerospace MAI Journal, 2016, vol. 23, no 4, pp. 55-60.

The purpose of this work consisted in determining the type and functional content of the dependencies, constituting the two-component LPRE control algorithm and obtaining formal description of these dependencies for further use of this algorithm while implementing the engine as a part of a launcher during the flight.

It is shown that the task of maintaining the specified for flight conditions engine thrust level values R and mixture components ratio Km are clearly described by specifying functions of regulator assembly drives position in relation to the six parameters: R, Km and four conditions at the engine inlet (temperature and fuel components pressure).

This conclusion was drawn by analyzing the structure and functional dependencies of LPRE mathematical model. It was successfully proved by determining such dependencies using adequate fire tests results of a given single-chamber LPRE approximation.

To determine control algorithm for LPRE, undergone hot testing, the author suggested implementation of computational-experimental model (CEM), formed according to the results of this engine hot testing.

The properties of such model allow carrying out reliable forecast computations of the engine operating procedures parameters in a wide range of the six parameters under consideration, namely operating modes and ambient conditions.

The final form of control algorithm represents a polynomial, approximating computation results based on CEM, carried out over six-dimensional array of computed points, defined within the required engine operation range.

The adequacy of the proposed approach to the control algorithm formulation in the wide range of all six parameters is validated by comparing the values obtained by approximation with experimental data of a given single-chamber LPRE.

Kraev V. M. Present condition of unsteady turbulent flows study. Aerospace MAI Journal, 2016, vol. 23, no 4, pp. 61-67.

Heat and hydrodynamic processes are becoming determinant while creating new types of engines for space, aviation and nuclear power systems [1 – 9]. Unsteady hydrodynamic and heat transfer processes study is an extremely important problem of engine building.

Only the combination of fundamental and engineering studies provides most effective way to design precise unsteady process model for practical computation. Experimental studies carried out in Moscow Aviation Institute (MAI) hold a prominent place in this field [10 – 17].

The turbulent flow structure studies carried out in MAI reveal non-stationary conditions fundamental effect on turbulent flow structure.

Axial and radial velocity and temperature pulsations, average parameters and their correlations were measured as a part of the study. Generalized experimental data reveals significant impact of flow acceleration and deceleration on turbulent structure. Three specific zones in turbulent flow were identified: near wall area y/R = = 0...0.02 (y — distance from the wall, R — radius of the channel); maximal turbulent parameters modification area y/R = 0.02...0.4 and flow core. Significant difference of turbulent viscosity between steady and unsteady approaches up to three times was identified. Comparison of quasi-steady and unsteady approach to heat transfer and hydraulic resistance coefficients revealed the two-times difference. Undoubtedly, such huge difference is unacceptable for space, aviation and nuclear energetics. This result agrees well with experimental data obtained by other authors [18, 19].

Based on non-stationary conditions significant impact on turbulent structure a computation model was developed. With flow acceleration, hydraulic resistance coefficient exceeds relative quasi-steady value by 2 times and more. During flow deceleration, it is 35% less.

Experimental study results present reliable base for further theoretical studies to be carried out in MAI [17]. The existing high-Reynolds turbulent models are not able, in principal, to consider non-stationary effect. From turbulence models analyzed in [18], only Menters SST model, which is low-Reynolds model, gives the results close to the experimental. Generalized equations for non-stationary friction and heat transfer coefficients at flow acceleration and deceleration in a tube for engineering design were obtained. The advantage of such models consists in the possibility of their employing for any monotonous flow variation curve, as well as satisfactory convergence with experimental data on hydrodynamic non-stationary gas flow in through channels [20].

Among the works of theoretical character, the studies of Professor Igor Derevich should be noted in the first place. In reference [21] the author considers the gas flow with monotonous consumption decrease/increase, and reveals the causes of computation and experimental data mismatch.

For practice, we recommend to analyze the effect of non-stationary processes on a certain jet engine control system. In case, when the processes are principally non-stationary and the required accuracy must be high, a non-stationary model and/or other approaches, considering non-stationaries, should be used.

Kolodyazhnyi D. Y., Nagornyi V. S., Smirnovskii A. A. On effect of electrical charge on fuel drops surface tension at the atomizer outlet. Aerospace MAI Journal, 2016, vol. 23, no 4, pp. 68-78.

High-speed transport design, aircraft engines ecology and higher energy efficiency guarantee by improving fuel atomization and air-kerosene mixture quality in aircraft engines intensive research is carried out. To improve fuel atomization and air-kerosene mixture burning we suggest the use properly shaped electric fields in atomizer fuel supplying contours. For the first time the authors studied the effect of variable frequency AC electric field on combustion products chemical composition, when employing kerosene TS-1. Experimental results on the effect of variable frequency AC electric field on air-kerosene mixture combustion products burning rate were presented for the first time.

Post-combustion flow speed measuring at the simulative combustion chamber outlet were carried out at Samara State Aero-space University (SSAU).

Air-fuel mixture combustion products burning speed experimental determination technique was developed at SSAU. It forms the basis of the research on the effect of AC electric field on air-kerosene combustion mixture products speeds.

Employing the speed measuring data, computations of superficial velocity and mass flow ratio were carried out using well-known equations for gas-dynamic functions.

The result of experimentation consists in creating Microsoft Access database file with further possible export to Excel.

Experimental studies were carried out at SSAU on a single-burner bay of a simulative combustion chamber with operational OJSC “Klimov” duplex nozzle for liquid fuel. We employed a swirler with blades angle φ = 72°10′; gas collector with cone outlet diameter of 133 mm; square spacer plate with square cross-section shaped with square side of 180 mm and a baseline case of offset area holes, when mixer apertures were open. Kerosene TS-1 was used as fuel. Low-pressure compressed air was fed under pressure ≤ 0.75 MPa, and solid tracing particles were used for laser measurements of Ch-4 type.

When the AC electric field was applied to kerosene along each diameter, prior to feeding to atomizer, speed values move intermittently up and down. With this, air-kerosene mixture combustion products maximum relative speed reduction was 2.45%, while maximum relative speed of air-kerosene mixture combustion products with applied to kerosene flow AC electric field at the outlet of combustion chamber was 1.425.

Ivanov A. V. Study of genetic algorithm implementation efficiency while turboprop engine modeling. Aerospace MAI Journal, 2016, vol. 23, no 4, pp. 79-85.

Propellers design and development for modern coaxial propfans and their automated control systems are impossible without in-line simulation test benches, which allow reduce testing fee, imitate failure situations, work through control laws and algorithms and determine automated control systems stability margins.

Turboprop engine mathematical model plays key role while testing propellers and automatic control systems with in-line simulation test benches. The tests validity depends on accuracy of non-stationary processes reproduction by mathematical model. Due to turboprop dynamic characteristics errors when employing linear methods of modeling, at present, non-linear element-by-element models became widely used. In the course of SV-27 coaxial propfan and RSV-27 hydro- mechanical regulator testing bench, JSC SPE “Aerosila” employs D-27 turboprop non-linear element-by-element model. Implementation of gas turbine engines non-linear models results in significant processing power waste due to the multiple recalculation of the thermodynamic mathematical model while compressors and turbines joint operation point search. To optimize the computational process while using a non-linear turboprop engine mathematical model the authors suggest to use of a genetic algorithm. Genetic algorithm was developed with LabView software, employed with in-line simulation test bench and associated with the engine mathematical model. Genetic algorithm of various configurations and probability values of mutations and number of species in population with in-line simulation test implementation efficiency was studied. The results of the study allowed determine the optimal genetic algorithm configuration and parameters of its optimal operation. In its optimal configuration with a small number of species in population and increased calculating error, this genetic algorithm appeared to be effectiveness comparable to method of successive approximations by bisection. However, the genetic algorithm execution instability, leading to computational resources wasting for some calculated points, makes its implementation in turboprop engine mathematical model, used with in-line simulation test bench for air propellers tests and their automated control systems, impractical.

Siluyanova M. V., Chelebyan O. G. Pneumatic method for uniform air-fuel mixture preparation in GTE combustor. Aerospace MAI Journal, 2016, vol. 23, no 4, pp. 86-94.

The main objective of the research, aimed at developing combustors for civil aviation, consists in ensuring competitive level of engines emission characteristics. The presented work is dedicated to the development of technology for uniform air-fuel mixture preparation in the flame tube head with respect to aircraft combustion chamber.

Gas turbine engine aggregate characteristic guarantee, such as reliable start-up, wide range of stable operation, fuel combustion efficiency and low noxious emission depend in particular on combustor reliable operation. The researches in this field for the most part are agreed herein that achieving high-level of the above said characteristics in the combustor is stipulated, not after all the others, by liquid fuel crushing process quality and its preliminary mixing with air in the flame tube head. It is known that combustion of previously prepared homogeneous air-fuel mixture in model heat generators allows obtain low outlet noxious emission. However, real GTE combustor has no place or the time for such preparation. It stands to reason, that it is necessary the employ the available space and the residence time maximally to direct the air-fuel mixture characteristics drift towards a homogeneous composition.

This work presents the results of the designed flame tube head with liquid fuel pneumatic atomizer for low-emission combustor. The paper describes the air-fuel uniform mixture preparation technique in the flame tube head with fine-dispersed spray in swirl flow conditions.

Autonomous tests of the developed the flame tube head have been conducted. In the course of these tests the main characteristics of the air-fuel spray formed after burner by a non-contact laser diagnostics method in open space conditions were studied. According to the results of cold tests, the average Zauter diameter of the fuel droplets in the idle mode is about 23 microns. The wide and intense backflow zone is formed near the device axis. To test the developed device and method of air -fuel mixture preparation, fire tests in the model three-burner compartment under high-pressure environment were carried out. The ignition and blowout points under earth conditions have been obtained as the result of tests conducted. The efficiency of lean air-fuel mixture combustion technology has been confirmed.

Afanas'ev V. A., Tushavina O. V. Methods and means for thermal-protective materials development verification under climatic effects conditions. Aerospace MAI Journal, 2016, vol. 23, no 4, pp. 95-102.

Revealing climatic conditions effect in the course of pre-launch procedure of reusable space system is a necessary condition of thermal protection structural components ground development. The climatic tests experiment must simulate positive and negative temperature effects, as well as humidity and precipitation effects. The climatic tests algorithm is designed as a unified processing chain of test set up when a number of simultaneous or consecutive experiments are conducted at the experimental testing bench. The algorithm includes thermal-stability, low-temperature stability, moisture and weather resistance tests. The presented paper describes methods of reusable thermal-protective materials structure testing under the above- mentioned conditions as well as techniques for thermal protection structural elements testing for day-night and seasonal cycling.

Recommendations on carrying out the accelerated tests for climatic firmness are given. The approximate scheme of the main climatic factors affecting heat- protective material change in the experiment within the full-year cycle is presented.

It is noted, that experimental means for carrying out tests in the conditions of climatic influences must present a constituent part of the experimental means intended for the tests in the conditions of multiple-factor impact on of thermal protection materials.

The schematic diagram and photos of the test bench and its components used for heat-shielding reusable materials tests is provided.

The suggested methods and experimental facilities for conducting thermal-protective materials climatic tests on multivariable screen tests of tile-type thermal- protective structural elements can be used for consistent assessment of their working efficiency during ground tests. Ground tests of spacecraft units and plants can be conducted by simulating only the major external factors whereas secondary factors impact can be taken into account by introducing corresponding coefficients.

Zakharov I. V., Trubnikov A. A., Reshetnikov D. A. Airborne short-range air-to-air missile guidance system software/hardware complex technical layout and methodological support. Aerospace MAI Journal, 2016, vol. 23, no 4, pp. 103-110.

Based on the present-day maintenance conditions of modern missiles the paper reveals essential factors, which determine their reliability and readiness support at the required level. This, in its turn, allows establish that up to 70% of failures during instrumental control relate to guidance system, and more than half of them falls at the missile control system. The above mentioned problem manifested itself most acutely with short-range air-to-air airborne missiles. This implies the effective solution of missile control problem by employing guidance system functional control method and its realization based on hardware/software complex.

The paper suggests an original solution for short-range air-to-air airborne missile guidance-system loop technical condition, enhancing its functional control methodological efficiency, confirmed by methodological and hardware support synthesis.

Functional control scientific and methodological basics are determined by theory of similarity modeling and automatic control theory harmonic balance methods. The functional control effectiveness achieved with this method is determined by basic concepts inherent to the complete mathematical model structure, using the original inciting signal, generated by standard harmonic oscillations installation. These basic concepts include generation of such initial impact, which allow enhance missile guidance system controlled signals observability in system normal operation mode in space of parameters control.

The direct guidance system direct control time is one of the important parameters, related to its activation. This time is comparable to missile operation while intended application. It ensures the short-range air-to-air airborne missile specified life substantial saving.

Effectiveness of the methodological approach used by authors is supported by developing the guidance system software and hardware functional control complex that prevents introduction of changes to the guidance system hardware and sensors regular system. Thus, the possibility of practical implementation of the methodology, suggested by the authors, into field aerospace forces of the Russian Federation is guaranteed.

Sokolov N. L. Analytical calculations of a spacecraft motion path in atmosphere. Aerospace MAI Journal, 2016, vol. 23, no 4, pp. 111-121.

Employing analytical methods for computing spacecraft movement trajectories seems effective while solving a number of problems of practical importance. Analysis of the existing methods reveals that they are based mainly on mathematical models of spacecraft flight along a fore-and-aft plane, as well as some simplified spacecraft motion in space equations. It limits the possibility of their use while solving a number of space exploration problems of practical importance. The paper describes an analytical method for spacecraft atmosphere movement parameters computation. The scientific novelty of the developed method consists in transformation of a number functions in the form of recurrent piecewise-constant dependencies at the finite intervals of spacecraft flight trajectories.

After transformation of initial system of differential equations, we obtained the final computation dependencies for velocity and flight altitude, trajectory and course angles, longitudinal and cross range via the atmospheric density. Selection of such an argument, namely atmospheric density results from the fact that spacecraft flight situations can be identified based on calculations of this parameter with further recommendations for control decision-making. Based on the obtained equations we can compute not only the coordinates of spacecraft atmosphere movement, but evaluate the main characteristics, effecting design and technological decision making while a spacecraft design. Particularly, the fast evaluation of maximum overloads values, affecting a spacecraft in aerodynamic deceleration phases is provided. Analytic dependencies can be used while solving a number of variational problems in the conditions of preliminary definition of spacecraft control structure.

The tabular matter and graphical data are presented. Computation errors of spacecraft motion trajectory parameters are analyzed. It is shown that these computation errors do not exceed 2-3% with the total qualitative matching of obtained data and of differential equations numerical integration results. Employing of the developed analytical method allows obtain the highly precise computation results of spacecraft motion parameters in the atmosphere. The developed formulas provide high speed of calculations for a wide range of initial data, boundary conditions, and can form the base for spacecraft onboard control algorithms development.

Tatarenko D. S., Efanov V. V., Lobanov K. N. Uncontrolled object motion parameters algorithm based on radar data reprocessing. Aerospace MAI Journal, 2016, vol. 23, no 4, pp. 122-130.

This work relevance is stipulated by the necessity of airborne uncontrolled objet implementation accuracy to fulfill such tasks as forest fire extinguishing, large cargoes airlift delivery, etc. At present, conventional aiming systems do not provide uncontrolled object implementation effectiveness in full measure, since the onboard ballistic algorithm employs approximating equations and demonstrates low accuracy.

The authors suggest employ uncontrolled object motion complete ballistic model to improve onboard ballistic algorithm accuracy. The initial conditions can be obtained by determining uncontrolled object motion parameters based on radar signal reprocessing. These parameters determination can be realized with the algorithm, which description and structure are presented in this paper.

The paper presents computation results of the signal reflected from an uncontrolled object. These signals reveal that at the distances of up to 200 m secondary modulation harmonics of the first and second order are quite observable in the reflected signal spectrum, under condition of long-continued coherent integration of the signal.

The main advantage of this algorithm consists in the procedure of obtaining the unmanaged missile accurate initial conditions, based on the interpretation of the Dopplers effect together with complex application of known mathematical methods of signal processing. The reflected signal from uncontrolled object processing allows obtaining uncontrolled object launching (drop) angle, relative to the center of mass position, velocity and motion trajectory.

Moiseev K. A., Panov Y. N., Moiseev K. K. Study of overloads occurring while special long loads transportation, carried out by two-link tractors. Aerospace MAI Journal, 2016, vol. 23, no 4, pp. 132-136.

The paper presents a method for determination of overloads in the cross-section of long restricted articles, which can be employed at the initial stages of launching vehicles (LV) springing systems based on two-link tractor, while moving through rugged topography terrain, peculiar to Arctic zone of Russian Federation.

To evaluate overloads in in the cross-section of long restricted article the authors developed mathematical models of “two-link tractor — long restricted article” interdependent system, composed on the assumption of hitch mechanism infinite stiffness, when the LV presents infinite stiffness body, which practically eliminates the possibility of resonant speed modes occurrence while acceleration and moving with maximum speed.

The system of differential equations describing dynamic behavior of two-link tractor is divided into three less complicated systems of differential equations, which are solved by the original analytical method, namely combination method. This method is highly effective for dynamic systems study, if a differential equation does not exceed the sixth order. It presents an integral combination of symbolical and parameters variation methods. The symbolical method allows construct the resulting equation for the initial system of differential equations, and find dissipation and eigen frequency factors for the system under consideration. Parameters variation method, based on the solution obtained by symbolical method allows determine specific solution of the initial system of differential equations in the form convenient for the analysis.

The obtained results may be of interest to organizations involved in the design of viscoelastic suspensions not only for caterpillar tractors, but also for road and air transport, and exploring emerging overload of cargoes in extreme conditions.

According to the obtained results the conclusions on the expediency of operation of the hitch mechanism providing absolute rigidity of the coupling links of the tractor when moving on ground with periodic roughness in extreme operation conditions.

Kirillov V. Y., Tomilin M. M. Crosstalk calculation in electric circuits of aircraft steering gear. Aerospace MAI Journal, 2016, vol. 23, no 4, pp. 137-144.

Two types of electric drives — electro-mechanical or electro-hydrostatic — are supposed to be implemented for оnboard systems of “more electric aircraft” with a great number of various kinds of electrical equipment [1] for controlling various functional elements [2]. The increasing number of implemented electrical equipment, electro-mechanical steering gear in particular, which phase currents cause electromagnetic interference (EMI) in the form of electric and magnetic fields with high-level intensities. The main source of radiated EMI caused by electric drives systems are power circuits bundles. Electro-mechanic steering gear power circuits' bundles consist of a pair of twisted or axial conductors with currents' of tens of amps [5, 6]. Correspondingly, they generate radiated EMI, which may cause signal distortion in aircraft onboard system communication lines and, hence, deteriorate its functionality.

The presented study is dedicated to of radiated EMI levels in the form of magnetic field harmonic components computation. These EMI are generated by phase currents in aircraft electro-mechanical actuator motor powering circuits, and crosstalk in the form of voltages in open conductors of double-wire communication lines.

 The presented spacing charts allow deduce that voltages and currents, which amplitudes are commensurable or even greater than valid signals values, occur in aircraft onboard cable system communication lines in the form of harmonic electric and/or magnetic field. The charts allow determining the safe distances between power circuits and open communication lines, wherein the levels of induced conducted interferences are significantly lower than information and control signals peak values in aircraft onboard system communication lines. It allows provide electromagnetic compatibility of high-power and low-power circuits.

The presented paper is a part of the research work on computation and simulation of electromagnetic interferences, caused by transients in aircraft steering gear system.

Le D. T., Averin S. V. Generation of vector PWM ensuring through currents elimination in three-phase bridge inverter. Aerospace MAI Journal, 2016, vol. 23, no 4, pp. 155-163.

The рaрer suggests a control algorithm for voltage generation at induction motor windings by vector PWM. It reveals sрecifics of conventional vector PWM algorithm. Its is noted, that while through currents elimination with delay circuits a certain state occurs which allows identifying it as additional vector generation. The authors suggest a control algorithm with extra vectors generation as a through currents elimination technique. The рaрer comрares the suggested technique with conventional, and demonstrates that the develoрed algorithm using extra vectors allowed eliminate through currents of a first genus, decrease amрlitudes of high-order harmonics, and ensure рhase and рhase-to-рhase voltages рarameters similar to the conventional technique. Simulation of the suggested technique was carried out, and its results revealed that рhase current sine waveform could be ensured not only by increasing the number of generated vectors in one sector, but also by introducing extra vectors.

Conventional and suggested techniques reveal that рhase and рhase-to-рhase voltages characteristics as well as рhase current are similar, but the number of high- order harmonics is less than with conventional one. The breadboard tests revealed that the develoрed algorithm did not lead to shaft whiррing. Inverter inрut current herewith is less relative to the conventional vector PWM technique.

With on state of intermediate vectors significantly less than on state of base vectors the рossibility to attain рositive features рeculiar to the conventional technique, but eliminate a number of its drawbacks. The suggested technique, in рarticular, allows eliminate through currents, and gives more рossibilities of vector PWM imрlementation. Extra vectors on state duration control, rather than increasing the number of generated voltage vectors, allows ensure рhase current shaрe more close to sinusoidal.

Voronin S. V., Loboda P. S., Ledyaev M. E. Optimal porous structure determination to improve aluminum alloy mechanical properties. Aerospace MAI Journal, 2016, vol. 23, no 4, pp. 164-173.

Creating a competitive technology requires implementation of new materials with high specific mechanical properties. Conventionally, such materials are produced by introducing alloying elements, which form strengthening phases within the base metal structure. This approach usually results in the mass gain, because the hardening phase density is often higher than that of the base material. The mass of material can be reduced by introducing it into the volume of structural defects, such as pores. Due to high damping properties, low thermal conductivity, high sound-insulating ability and good moisture resistance, the porous materials are widely used in industry [1-7]. With existing porous aluminum, manufacturing technologies its strength properties decline takes place. However, with porous structure ordering the strength properties of finished products improve [8-10].

Thus, the goal of the presented work consists in improving specific mechanical properties, yield strength in particular, of the material by introducing orderly arranged pores.

This study employed deformation processes finite element modeling with engineering analysis pack MSC.Marc to determine an optimal porous structure [11-12].

The study of porosity and a type of porous structure effect on mechanical properties was carried out with the following types of porous structures: square, field-interleaved, square with a pore in its center, triangle and hexagonal.

With porosity of 0.4 to 0.5% porous samples FEM yield strength matching with compact material FEM samples yield strength is observed. With further porosity decrease growth of yield strength is observed for all types of porous structures. Maximum yield strength increase of 1 to 2% was achieved with porosity of 0.1%.

The blanks for all the samples were cut from the aluminum alloy A5 sheet using laser cutting complex. All the obtained blanks were decollateв into three parts. The first part was left intact as a compact material sample. In the second part of the blanks, the ordered porous structure was obtained by laser burning. In the remaining samples, the porous structure was obtained with CNC milling and engraving machines with the drill diameter of 300 µm.

The finite modelling and real uniaxial tensile tests results matching is observed.

Agafonov R. Y., Vilkov F. E., Kasitsyn A. N., Predko P. Y., Marchenkov A. Y. Aluminum based alloys with rare-earth metals additives application for rocket-and-space engineering. Aerospace MAI Journal, 2016, vol. 23, no 4, pp. 174-180.

Nowadays “AMg6”, “D16” and “AMn” aluminum alloys are traditionally used for space technology. Application of new advanced aluminum-based materials with of rare-earths additives instead of traditionally used alloys would enhance the electronic components protection from the space ionizing radiation due to alloying with high radiation absorbing elements. Whereas chemical composition manufacturing technique optimization will improve, alloys' mechanical properties compared to conventionally used, which will allow decrease weight and size parameters of the design.

Tests carried out by Russia's space industry leading organizations revealed significant preeminence of new alloys compared to conventionally used with regard to protection against outer space ionizing radiation properties, and corresponding to them ability to chemical electroplating. Aluminum based alloys specific mass with rare-earth additives is 2.9 g/cm3 on the average.

This paper is focused on the study of the three different alloying systems: 1 – Al-Dy-La-Cr-Zr, 2 – Al- Ce-Cr-Zr, 3 – Al-Mg-Sc-Zr-La-Ce; with rare-earths content not exceeding 11%, 7% and 9% by weight respectively. Each of the studied alloys, regarded as a material for spacecraft electronic equipment casing has a number of advantages and disadvantages. Increasing the rare-earth metals content in the alloy we can attain both better protective characteristics against space ionizing radiation, and aluminum based alloys with rare-earth additives welding properties improvement. Tough their density herewith will increase. Thus, it is necessary to pay special attention to improve mechanical properties of the basic metal and welding joints to prevent weight and size parameters of the design. Mechanical properties improvement with density reduction may, in some degree, be achieved by rare-earth aluminide phases' dispergating and increasing their density distribution in the alloy groundmass.

Betsofen S. Y., Osintsev O. E., Knyazev M. I., Dolgova M. I., Kabanova Y. A. Quantitative phase analysis of Al-Cu-Li-Mg system alloys. Aerospace MAI Journal, 2016, vol. 23, no 4, pp. 181-188.

The authors have developed a method for computing a number of intermetallic phases (T1 and δ′- phase) of Al-Cu-Li-Mg system alloys based on measuring α-solid solution lattice periods; Vegard’s law, linking up solid solution lattice period value with alloying ingredients content in it, as well as chemical and phase content equations. Lithium content in solid solution serves herewith as variable parameter. Quaternary Al-Cu-Li-Mg system alloys quantitative phase analysis method is based on the assumption that all magnesium resides in the solid solution. This fact is considered by introducing the relevant term into equation for calculating the solid solution lattice period. This is the only difference fr om the previously developed similar method for ternary alloys Al-Cu-Li. The paper shows that the developed method can be effectively used for quantitative interpretation of thermal and thermomechanical processing impact on alloys’ phase content study results, as well as while Al-Cu-Li-Mg system alloys content optimization. This method allowed us to compute the relation between periods of solid solution lattice and the amount of intermetallic phases for 29 Russian and foreign industrial alloys of various generations. The paper reveals the existence of linear dependence of relative quantity of intermetallic phases in alloys  from the atomic concentrations of lithium and copper (magnesium)  in these alloys. It shows also, that relation between δ′-phase and ternary phases is determined by the atomic concentration of lithium and copper. The authors suggested new Al-Cu-Li-Mg — alloys classification, wh ere all alloys should be divided into five groups, differing from each other by the double δ’-phase and ternary phase shares , or  ratio.

According to this classification, all the alloys are divided into five groups. The first group includes Al-Mg-Li alloys, for which the phases ratio . For the second group the ratio  varies from 2 to 3; for the 3rd group — from 5 to 7; for the fourth group — from 7 to 8, and the fifth group — from 11 to 17.

Soldatenko I. V. On titanium alloys semiproducts quality control. Aerospace MAI Journal, 2016, vol. 23, no 4, pp. 189-194.

The goal of the paper consists in titanium alloys semiproducts macrostructure quality evaluation technique improvement.

Active standard ten-point scale of macrostructures was developed based on – grains of strictly equiaxial shape specific to strain-free state of the alloy of sheet-like intragrain structure.

It is well know macrostructure we can see only those structure elements, which size exceeds 100 –150 micron (i. e. the ones exceeding the eye resolution capability).

Macro- and microstructure evaluation of a large number of serial semiproducts and laboratory samples revealed that not only – grains could be visible on a microstructure, but – colonies as well. It was established, that while checking a established, that the shape and size of the grains in the observed macrostructure depended on – grains and  – colonies in the microstructure.

Direct dependence of a macrostructure character from its microstructure was revealed. The paper shows that macro grain size and its tonality (degree of brilliance) depend directly on parameters of the microstructure, forming while deformation and heat treatment processes at temperatures of -or + – area. Correlation between the grain maximum longitudinal and diametrical sizes (the degree of non-equiaxiality – K) is clearly associated with physical degree of its deformation. This is another important parameter of macrostructure evaluation besides the grain size itself.

By deformation, the macro grain tonality or its degree of brilliance changes together with the macro grain shape. Interrelation between degree of brilliance of a macrostructure under study and with its microstructure was established.

The author suggests classify a macrostructure according to its tonality (degree of brilliance) by four types:

  • Absolute brilliant – a typical macrostructure peculiar to allows with recrystallized or slightly malformed – grains which size exceeds 100 microns.

  • Brilliant with fog elements – observed in alloys with medium degree of deformation (10-35%) in – area. Within on – grain one can observe micro areas withvarious degree of – phase spheroidizing development (from globular to practically non- spheroidized, plate-like shapes of the particles).

  • Fog with brilliance elements – peculiar to the alloys malformed in – area to the degree of 40-55%. For the most part this structure is globular or globular plate-like. In some locations it preserves oriented character of – phase excreta, which in case of their large size are responsible for appearance of these brilliant locations in the macrostructure.

  • Absolute fog-corresponds to globular or globular plate-like microstructure.

To improve titanium alloys evaluation objectivity and unambiguity the author suggest introduce quantitative estimation based on three parameters, namely grain size, and the degree of its non-equiaxiality and tonality.

The next step to titanium alloys production quality improvement consists in working out requirements to macrostructure based on quantitative estimation of its parameters.

Davydov A. D., Dianova E. V., Khmelevoi V. V. Fundamental and exploratory research priorities selection method. Aerospace MAI Journal, 2016, vol. 23, no 4, pp. 195-203.

The paper suggests methodological approach to fundamental and research priorities portfolio forming. Such an approach is based on expert selection procedure. The authors formulated thematically oriented verbal and numerical scales for qualitative selection criteria groups. This criteria grouping is organized according to thematically similar features, significant while new perspective aircraft systems design.

Due to the complexity of solved selection tasks it is reasonable to implement quality criteria system, described by 3-5 criteria groups with 3 to 7 criteria in each group. It allows convenient, transparent and comprehensive presentation of information to the expert in necessary and sufficient scale.

Group-1 represents usefulness and importance; Group-2 represents resource intensiveness and resourcing, and Group-3 represents stability and manageability.

With allowance for expertize complexity and supposed relative inconformity of experts opinions when evaluating significance of particular researches trends, we suggest selection procedure based on the Ansoff's theory of weak signals. Here, with allowable level of experts' nonconforming opinions, the individual opinions of competent experts with high estimate of particular FERs are taken into account. Here, core index (CI) and concordance index (DI) serve for the generalized selection measure. In this case we suggest FERs grouping in the following way.

FER-1 are the trends with experts' high estimation by CI with high DI value. FER-2 are the trends with relatively high estimation by CI with relatively low DI values. FER-3 are the trends which CI is better than this for FER-2, with DI lower relative to FER-1 and comparable to this for FER-1. FER-4 represents such FER trends, which received consensus on lower importance, either special opinion was expressed by experts with relatively low authority.

The authors suggested to form the portfolio not only by the trends with high CI and DI values, but consider FAR-2 trends (with priorities higher than this of FAR-3) as well. This approach allows us to identify and select among the priority research areas with potentially high efficiency, albeit with relatively high level of risk. The proposed approach also makes it possible to make informed decisions in a limited time based on authoritative (respectable) peer review. The method is oriented for use in decision support system.

Galkin V. I., Kuzina S. M. Building a model for optimal quantity determination of manufacturing facilities. Aerospace MAI Journal, 2016, vol. 23, no 4, pp. 204-210.

The paper presents a technique for a number of work places optimization at the enterprise with variable product release program. The developed technique is based on simulation and experimental design. The paper considers the operation of the enterprise manufacturing several kinds of products by assembling either purchased components or produced at this enterprise. The simulation model developed in the course of this study allows build and optimize manufacturing resources under various variants of enterprise's target figures.

The model was built with AnyLogic program, which allows specify time intervals at every stage of manufacturing either major product, or associated items. There is a possibility to model the situation with various number of assembling departments.

Based on the built model the authors carried out the optimization experiment, which allows compute an optimal number of equipment for the specified work-order quantity for all types of products. The paper suggests goal functions with productivity optimization. Using this instrument the results for each experiment were obtained by varying values of run-out production plan. It is found on what production volumes minimum quantity of equipment is optimal, and at what moment the number of working places should be increased. It is also determined that maximum possible quantity of equipment under specified production volume boundaries is not necessary.

The obtained results were processed according to the experimental design technique. The equation for corrected production effect computation as function of a number of assembling departments and products production volume. The proposed method is universal and can be applied for various types of production. The developed technique can be used as one of the instruments while developing the system of managerial decision making.

Efimova N. S., Zamkovoi A. A., Titkov A. M. Aircraft manufacturing enterprise innovative activities development with allowance for economic security requirements. Aerospace MAI Journal, 2016, vol. 23, no 4, pp. 211-217.

At present, creation and implementation of a system of indicators for monitoring R&D processes is necessary, since the degree of highly efficient estimation of economic security of an enterprise, and formation of economic mechanism implementing the complex of necessary measures on prediction and preventing a danger, corresponding to the scale and threat environment to aircraft industry in the aggregate.

The main objective of the innovative activities economic security level consists in timely analysis and monitoring of a complex indicators system, inclusive aircraft industry specifics.

Development and implementation of economic mechanism for innovative activities economic security provision in aircraft industry will allow reveal: insufficient certainty of a forecast at various R&D fulfillment at stages; excessively enlarged and averaged character of labor intensity rate and expenditures, new objects' operational service norms, reliability and durability; insufficient comparability of new objects with selected prototype objects, or the lack of scientifically substantiated techniques for this comparability economic evaluation while effectiveness indicators computation; the lack of exact information on all spheres of R&D results supposed implementation and their scope of their implementation; the difficulty of extracting the share of economic effect related to this particular technical solution, the specified object, used as a constituent part of a more complex technical system.

Chaika N. K., Gavrilova I. S. Corporate governance system development estimation method. Aerospace MAI Journal, 2016, vol. 23, no 4, pp. 218-224.

A corporate governance system developing method is offered within the framework of this paper.

This newly developed method aims at estimating the corporate development level of a space-rocket industry enterprise, as one of the crisis-proof enterprise governance tools.

American and German corporate governance models are distinguished as the most popular models. Russian corporate governance practices are equally distant ideologically from both.

To analyze the corporate governance system at the rocket-space industry enterprise the authors developed their own estimation method, consisting in accounting for a number of specifics of integrated structures and individual enterprises functioning.

Basic methodological approaches to optimal assessment method of corporate governance system development at the enterprises and integrated structures of rocket-space industry are considered.

The developed method account for specifics of the companies and allows carry out their comparative assessment in conditions of differences in structure, scope and lines of activity.

This method represents a multilevel assessment system by two blocks — the level of corporate governance implementation and the level of the corporate governance system formalization.

In accordance with the results of the s