2013. № 70

The first part of the paper considers the hypothesis about the influence of the non-uniform air flow on the aerodynamic drag of porous objects. The aerodynamic drag of porous objects such as radiators, mesh guards and screens, aerodynamic grids and grills, etc., which are installed in the air duct, is bigger than their resistance in ideal laboratory conditions. This difference appears when the air flow with non-uniform velocity distribution along the frontal surface flows through such an object. The research allowed to establish the factors, which influence the increase in the aerodynamic drag of the porous objects. The paper presents a formula for determining the aerodynamic drag of porous objects in a non-uniform air flow. The degree of change of radiator aerodynamic drag depends on two factors: the value of air flow non-uniformity and aerodynamic characteristics of the radiator itself (these characteristics are obtained during the tests with uniform flow).
The second part of the paper analyzes the reasons for this phenomenon. To carry out this analysis a radiator was considered as a porous object example. The suggested hypothesis was verified by using a Computational Fluid Dynamics (CFD) software module. It was found that the coefficient of variation of aerodynamic resistance of porous objects depends on the curvature of the graphical representation of their aerodynamic characteristic function. The larger the curvature of the function graphical representation is, the more the object is sensitive to the non-uniformity of the air flow.
Non-uniform air flow affects the aerodynamic characteristics of the radiator. In its turn the radiator also affects the airflow. As a result of interaction with the radiator the air flow is redistributed across its cross-section. This property of the radiator is important from the point of view of its heat emission. It was established that the larger the curvature of the aerodynamic characteristic is, the better the radiator equalizes the speed of the air flow across its cross-section.
The conducted research has shown that two interconnected processes can be observed during the passage of the air flow with non-uniform velocity distribution across the channel cross-section through the porous objects. On one hand, the porous objects affect the flow, on the other hand the non-uniform flow itself affects the aerodynamic characteristics of the object. The speeds within the air flow with a non-uniform velocity distribution across the cross-section of the channel are partially equalized while passing through the porous object.

Keywords: unique distribution of velocities of the air, porous object, radiator, aerodynamic resistance

In this paper the triblock and single supersonic turbulent underexpanded jets are numerically studied. The primary goal of the current research is to provide the numerical simulation for analysis of multiple turbulent supersonic underexpanded jets interacting with the barrier. Calculation grid consists of 2 million cells that makes possible to provide the quite good modeling of expansion waves and shock wave such as incident shock as well as the reflected shock and slip stream line in supersonic jets. Numerical simulation was based on the three-dimensional Reynolds equations with turbulent model SST. Three-dimensional compressible Reynolds equation was solved with the help of finite volume method of Godunov TVD scheme type with the second order of approximation in space discretization.
Depending on distance between nozzles and the barrier there are complex three-dimensional turbulent flow structures with areas of subsonic and supersonic flow, shock, rarefaction areas, viscous-inviscid interaction in turbulent mixing layer. This paper presents wall surface pressure distribution at various distances from nozzles to the barrier and its comparison with experimental data.

Keywords: numerical simulation, multiple jets, instability of flow

Ground launchers of unmanned aircraft vehicles with expansion machine for cold or hot working fluid and flexible transmission are main purpose of this article as well as formulation of optimization problem for ground catapult, which includes the boundaries determination of investigated system, the search for characteristic criterion, definition of intra-variable and model description.
Methodology consists of formulation of ground launcher useful function, drafting of factor matrix model, which consists of column matrixes of general and special settings and conduction of numerical experiments series for different values of parameters for common factor matrix as well as definition of dominant and secondary parameters of model by normalization of ground launchers configurations with the help of special modification criterion of artillery systems power.
It is selected the direction of dynamic performance improvement among the priorities directions of all types optimization of input systems in unmanned aircraft vehicles. Coupled physical-mathematical model of ground launcher was compiled. The numerical values of power modified criteria for various configurations of catapults were obtained. The conclusion about general parameters low impact on system useful function was made based on criteria. Inner optimization criteria were determined, which are the specific parameters of concrete input flight configuration (the transmission or drive).
Universal formulation of catapults optimization problem allows to use any method for solution to variational problem (direct or indirect). In the proposed formulation the optimization method allows to reduce the guided line length with maintaining of given flight velocity and start overload limit.
The problem of land dual catapults optimization based on gas-thermodynamically and mechanical model, which displays the non-stationary process in expansion machine, was posed for the very first time . The criterion of power caliber and super caliber artillery systems for terrestrial triggers for determination of their perfection was firstly used.

Keywords: ground catapult, characteristic criterion, intra variables, variation problem

In present work the basic results of acoustic tests low-sized pilotless vehicle (PV) are shown. Power plant (PP) of PV is arranged in a tail unit of a flight vehicle. The power plant consists of the two-stroke piston engine of air cooling P-032M and pusher ducted propeller. On the engine the reduction gearbox with the coefficient of reduction ≈ 2,102 is set.
Acoustic tests are conducted for three power conditions of PV power plant: an idle power condition, a cruiser power condition and full power condition. On each power conditions flow velocity in test section of wind tunnel is changed in a range of values from 15 m/sec to 50 m/sec with step 5m/sec.
As a result of an experimental research in a wind tunnel of acoustic radiation from low-sized aircraft power plant with pusher ducted propeller it is established that the total noise level of PP is determined, in the core, by first 10 harmonics of exhaust noise and first 5 harmonics of propeller rotation noise. The contribution of first five harmonics rotation noise to a total intensity of acoustic radiation of PV power plant makes on a mode of idle power condition of ~38 %, on a cruiser power condition of ~23 %, on a full power condition of ~3 %. The acoustic efficiency on full power conditions makes ~1,1 %.
On the basis of the carried out research it is possible to ascertain that in noise of PV power plant including the piston engine with air cooling and pusher ducted propeller, at absence in exhaust tube of the engine of exhaust silencer, a determining source of external noise is the system of an exhaust piston engine.
At the same time, the propeller arrangement in pusher configuration limits capabilities of decrease in exhaust noise of the piston engine as there are difficulties with accommodation of exhaust silencer.

Keywords: low-sized aircraft noise, piston power plant noise

The Magnetic Suspension and Balance System (MSBS) for wind tunnel models was created as the result of the collaboration of the Moscow Aviation Institute (MAI) and the Central Aero-Hydrodynamic Institute (CAGI). The MSBS was created for carrying out aerodynamic tests of the models with six degrees of freedom in a subsonic wind tunnel with the section size of the instrumentation chamber equal to 40 cm × 60 cm. Wind tunnel MSBS is being developed in order to solve the main problems of aerodynamics, which cannot be solved in the presence of the effects of mechanical support devices, and study the base pressure, etc.
The tested models are the bodies of rotation with the diameter of about 4 cm and 40 cm length (or less). The models could be equipped with a tail unit and low-aspect-ratio wings. A ferromagnetic core made of electrotechnical steel with the shape of a cylindrical tube was placed inside the model.
The MSBS has 6 degrees of freedom. The model angle of attack could vary within the limits from −45 to +45 degrees. The range of flow speed variation was 0 − 100 m/s. The MSBS includes 7 electromagnets. All electromagnets except for the one, which creates the longitudinal force, are installed symmetrically relative to the test section of the wind tunnel. The electromagnet, which creates the longitudinal force, is installed on one side of the test section of the wind tunnel. The shape and position of this electromagnet provides easy access to the model and a good visual observation. Each electromagnet is equipped with a separate system of current control. The upper coils of electromagnets have a common ferromagnetic core, which helps to increase the vertical force. The lower coils do not have a common core in order to obtain the required moments of forces and the necessary forces by extreme attitude testing.
The position and the attitude of the model are measured by the optical position sensor. The optical system includes the compensation of the influence of the change in the intensity of the light, which falls on the model.
The control system uses the control algorithms, which provide the maximal area of stability of the suspended object under the conditions of the control power limitation.
The roll stability around the longitudinal axis of the model is achieved by placing a permanent magnet inside the model. To provide the roll damping the magnet is fixed inside the model by a spring and a damp cell.
Investigation and improvement of the Magnetic Suspension and Balance Systems and techniques requires the creation of a mathematical model, which would correspond to the real system sufficiently. This paper adduces such a model, which is intended for simulation of the Magnetic Suspension and Balance Systems for a wind tunnel. The paper also presents some results of the research of the dynamic properties of the mathematical model.

Keywords: wind tunnel , magnetic suspension, electromagnets, mathematical model of the stabilizing system, wind tunnel aircraft model

When testing products of aircraft and space engineering various precise sensors that provide highly accurate measurements are extremely demanded. About 60% of all measuring devices are pressure sensors. Potential measurement accuracy depends particularly on the external effecting factors. Thus, strict technical conditions on stability against external mechanical action are applied to pressure sensors that are under modernization or development. More complex calculations require new methods for pressure sensors designing. In the article a new technique which integrates ECAD and MCAD systems for designing pressure sensor with differential capacitive measuring sensor is presented.
The proposed route of pressure sensors design procedure includes: simulation modeling in MATLAB environment, development of basic wiring diagram and printed circuit boards in ECAD MentorGraphics, design, development and execution of verification calculations in MCAD SolidWorks. Designing of 3D-model of the pressure sensor is made by data exchange between ECAD and MCAD systems using IDF files and CircuitWorks module. Analysis on vibration and shock action are executed in CosmosWorks module using the finite element method.
Simulation modeling allowed to determine resonant frequencies f0i along with pressure sensor waveforms, maximum relative movement zmax of the sensor elements, as well as shock acceleration ау. In order to confirm the correctness of selected circuit and design solutions, as well as the results of metrological characteristic studies there have been carried out preliminary tests of the pressure sensor model. As a result of testing it was determined that the proposed design of the pressure sensor provided proofness against mechanical shocks and vibration, and in total, the sensor accuracy meets the requirements of the specification. Estimated error of simulation does not exceed 5 %. Time required for pressure sensor development is reduced by 20 %. Results of the work can be used for designing of any class of pressure sensors.

Keywords: аircraft engineering products tests, pressure sensor design route, integration ECAD and MCAD systems, 3D-modeling, verification and design calculations in MCAD SolidWorks

In light aircraft design commonly used the spring gear, wherein the shock absorbers are used as springs working in bending and they absorb the main part of the impact energy during landing. At the initial stage of aircraft designing we must have a valid value of the relative weight of landing gear. For the following development of reliable methods for mass calculation of light aircraft landing gear the task of analysis of specific load on the wing and material mechanical characteristics depending on the relative mass of the spring main landing gear is assigned. An ideal equal stress spring of rectangular cross-section has been considered. The relationship between the strain energy and mechanical characteristics of springs material and main design parameters has been found out. Using Aviation Regulations AP -23 the work of external forces absorbed by a leaf spring has been determined. From the condition that the work of external forces and strain energy are equal an analytical relationship between relative weight of all the springs of the main landing gear and specific load on the wing and material mechanical characteristics has been found. The strong influence of the material on springs relative weight has been shown.

Keywords: light aircraft, landing gear, depreciation, spring

Operation reliability is one of the basic requirements to flying vehicles carried by launcher aircrafts. An intensive vibration causes high-level stresses in the areas of structural irregularity so that results the fatigue damages accumulation. These stresses must be considered to estimate accurately the strength and reliability of both carrier and carried vehicle, therefore the correct estimation of loading levels and of the durability of structures is one of the most practically important problems of aircrafts design.
The main goal of the presented investigation is the mathematical modeling of the dynamic stress and strain state of cinematically excited aircrafts’ structures and the estimation of their durability.
The random excitation caused by the flight loadings is transferred from the carrier to the carried vehicle through the suspension brackets.
The SolidWorks software is used to simulate the dynamic stress state. The stationary kinematic excitation is defined by several types of the acceleration spectral density. The properties of the random vibrations of the carried flying vehicle are computed using the constructed numerical models. The areas of the maximum amplitudes of vibration are found as well as the zones of the maximum stresses, and both the spectral properties and stresses are computed. It is shown that the maximum stress dispersion corresponds to two lowest eigenfrequencies.
The random stress processes corresponding to various spectral densities of stresses are computed using the statistical simulation algorithms as the harmonic series with random parameters for each time value. The obtained random processes are reduced to the sets of regular stress cycles that are equivalent by damaging capacity to the random ones on the basis of the standard algorithms. The repeatability diagrams for the regular cycles are calculated using the linear hypothesis of damage summation, and the fatigue damageability of the structure is estimated.
The developed models of random forced vibrations can be used to estimate the dynamic stress state, the vibration strength, and the durability of various carried flying vehicles.

Keywords: airplane, design, oscillations, finite element method, shell, dynamic response, vibration stress, stationary vibration, spectral density, statistical simulation, fatigue damageability

As a result of observations of the shape of the bird wings and the analysis of the positive impact on the aerodynamics of the wing vortices, it was decided to use a triangular protrusion for the wings of a modern flying boat. At that it is possible to observe the improvement of the wing aerodynamic efficiency. The circulation of the speed along the airfoil contour of the wing with a triangular protrusion was analyzed at different wing positions. The research of the wing with the different projection angles of protrusion deflection was also carried out.
The shape of the wing-to-fuselage joint within the flying boat wing layout influences the aerodynamic characteristics of the aircraft. We design two types of wing joints for the traditional high-wing flying boat and consider its aerodynamic characteristics and position of the mean aerodynamic center without the horizontal tail surfaces.
A number of numerical calculations was carried out with the help of the ANSYS Fluent 14,5 computational fluid dynamics software (license number 670351) to determine the overall aerodynamic coefficients and model the flow in the vicinity of the wing. These calculations confirmed the positive effect of the triangular protrusions on the aerodynamics of a flying boat wing.

Keywords: amphibious aircraft, fezyulyazh wing and flying boat, bird wing, wing with triangular protrusion, the Reynolds-averaged Navier-Stokes equations, the SIMPLEC method, turbulence models, aerodynamic coefficients

Future trends for space scientific-research activities make demands to the spacecraft and propulsion systems for them. Worldwide space programs and near-earth objects infrastructure evolution analysis enable the effective usage of electric propulsion for many of the delivered tasks. This paper focuses on analysis of one of the best possibility to increase engines performances and lifetime without bringing too many changes in the thruster design.
One of the main aims of this work was to find out if four electrode ion-optic systems could bring appreciable result to increasing lifetime, thrust and specific impulse of ion thrusters in comparison with its conventional three electrode systems. For that the analysis of state-of-the-art of ion thrusters was performed and all available results of integral performances were included into the summary table. As a result the current ranges for main thruster parameters were defined.
In the paper it is also presented some estimations of efficiency and expediency of usage of four electrodes ion-optic systems for one of most prospective technology for future applications – ion thrusters. Benefits arising from this issue delivered as a conclusion.

Keywords: electric propulsion, ion-optical system, ion beam extraction, beam focusing

A reusable electrically propelled cargo spacecraft with nuclear power plant, which should perform five round–trips between a low Earth orbit (800 km) and a low Lunar orbit (100 km) and transport components of a future permanently manned Lunar station, is considered. The spacecraft will weigh 61.8 tons, payload of 25 tons including, at the launch from the low Earth orbit. The nuclear–electric propulsion system will consist of a 1 MW nuclear power plant (of 20 tons in mass) and 4 clusters of 10 radio–frequency ion thrusters each. 28 of them will produce total thrust of 18.2 N and consume 840 kW of power. 10.8 tons of Xenon will be consumed per a round trip. The designed radio–frequency ion thruster “RFIT–45” with the ionizer diameter of 45 cm shall have the following performance: power consumption of 35 kW, specific impulse of 7000 s, and lifetime of 50 000 hrs.
According to the mission analysis, the reusable lunar transport spacecraft can deliver up to 128.5 tons of payload to the low Lunar orbit or 57.4 tons of payload to the lunar surface by 5 trips during about 7 years, delivery of payload of 205.2 tons to the low Earth orbit being necessary for securing mentioned operations, that being 2 – 3.5 times less than without the use of electric propulsions.

Keywords: radio frequency ion thruster, reusable lunar transfer vehicle, low earth orbit, low lunar orbit, nuclear electric propulsion unit

The improvement of methods (ways) and means of implementation and control of preload force remains one of the main directions of development of research on improvement of threaded connections reliability up till now.The research is focused on the process of assembly (tightening) of group threaded connections of aircraft units.
The paper is devoted to the problem of low accuracy of the control of the preload force according to the value of the torque moment during the assembly of group threaded connections of the repaired aircraft units.
The research aims at improving the tightening uniformity of the group threaded connections by increasing the accuracy of the control of the preload force according to the values of torque moments.
The accuracy of control of the preload force according to the value of the torque moment is largely determined by the magnitude of the friction forces in the threaded connection, which is the main flaw of such control method.
In actual practice the calculated values of friction coefficients can vary within a rather wide range from 0.05 up to 0.5, while the geometrical dimensions of a thread can deviate from the standards. These are the factors that result in low accuracy of such control method.
It was determined by an experimental approach that the condition of а threaded connection influences the relationship between the torque moment and the preload force. The experiments showed that after the threaded connection was exposed to corrosive action the previously calculated torque moment did not provide the required preload force even when a lubricant was used.
At present there are many ways and means of threaded connections tightening. However, the problem of providing the uniformity of distribution of the preload force in group threaded connections of aircraft units remains unsolved. This leads to numerous failures and decreases the reliability of the engineering.
It is possible to provide a more uniform tightening of group threaded connections by applying a more accurate method of preload force control by evaluating the relation between the unscrewing and screw-in torque moments.
The method consists in calculation of the exact torque moment needed for obtaining the required preload force for each threaded connection within the group. The relation of the unscrewing and screw-in torque moments, which is estimated for each particular threaded connection, makes it possible to take into account the actual condition of every threaded connection (i.e. indirectly estimate the values of the friction coefficients). At that it is necessary to take into account that the condition of the threaded connection may change during the term of unit operation.
The technological approach, which is described in this paper, allows to determine correctly the value of the relation between the unscrewing and screw-in torque moments for a particular threaded connection.
The paper contains useful information for scientists, who work on the development of threaded connection assembly technologies.

Keywords: threaded connection, torque, preload force, friction coefficient, torque wrench

This paper focuses on the problem of automatic landing of an unmanned aerial vehicle (UAV) on a runway in the presence of a strong crosswind.
This paper is suggesting the following landing procedure: an airplane deviates from the initial flight path adopting a course opposite to the wind direction from a given position, so as to minimize the lateral speed of the airplane relative to the ground from a certain ground proximity level, the rudder being used to align heading starting from the bank removal moment and through all the subsequent manoevers.
Ailerons and rudder linear and relay control laws are formulated, then a controller logic is devised, using both linear and relay controllers, for the task of monitoring the degree of coherence of the lateral and the longitudinal control channels.
In this work a new method of automatic landing was introduced. The landing maneuver is split in three parts. In the first part, the aircraft is following a precalculated path with an offset from the initial path. In the second part, the UAV follows a glidepath while performing bank control. During this stage, the lateral speed is nonzero and the course is directed towards the wind. In the third part, the bank deviation is eliminated by rudder control.
It should be noted that, for better accuracy, an adjustment for the wind value may be added to the control system, therefore complicating the controller structure, which is a suitable topic for a future work.
The paper suggests a new landing procedure, which may be of interest in the field of research in automatic control.

Keywords: unmanned aerial vehicle, landing, optimal control, crosswind, maneuver

This research is discussing a terminal aircraft control system of improved performance.
The aim was to develop the necessary logical structure for an integrated terminal aircraft control system, including linear and relay-based controllers.
Methodology: the tasks of guiding a vessel to a terminal point has always been a subject of primary interest in aircraft control. From the point of view of variation calculus, these tasks may be decomposed into a fixed and a mobile endpoints terminal control task. In the most common case, there is an endpoint with several undefined coordinates or with some additional constraints. This paper addresses one of these tasks — the problem of landing an aircraft on a given point of the runway, which is especially important for small allowable landing roll distances, and the problem of mooring an airship to a mast with a zero terminal velocity.
Achievements: this work proposes a structure for a terminal control system of higher performance, composed of a logical and an executive part. The logical part consists of a bloc estimating the remaining flight time, also in charge of the coordinated control in the lateral and longitudinal channels, and of two logic analyzer, tracking sign coincidences of deviations in position and velocity for lateral and longitudinal control channels. The executive part contains of two linear and two relay controllers, the first pair is used to assure cautious and soft docking, and the second pair — to improve the performance of significant deviation tracking.
Area of application: on aircrafts, expected to perform frequent landings.
Conclusions: The resulting control system structure is applicable on airships, where it’s important to shorten the time of approach to a desired endpoint, as well as on aircrafts, expected to perform landings in similar challenging conditions.

Keywords: jet, airship, logic, terminal point, optimal control, linear regulator, relay regulator, mooring, landing

In the article the analysis of the element of building systems was conducted according to the technical and tactical characteristics on the basis of the parameters range of functioning systems in the generated model. The purpose was a comparative analysis of the parameters and identifying the characteristics set out to solve the problems, viewed the functionality of onboard systems of the aircraft when they resist.
In this research was applied data mining method of construction an instant condition of the system. This method was used as a base on stage the collection and analysis information the system, and also combined with mathematical methods decrease the values of the parameters in the system.
Information about the problem obtained from the public media sources and Internet — list and description of systems used, outlining the problems and results of the application, relying on the chronicle of events 1999 — 2000, as well as on the basis of existing tactical decisions — as part of the initial information for the analysis and proposals of intelligent functioning systems with elements of decision support tools.
The method of analysis and the formation of the model have a significant meaning when necessary systematization of processes, large common scattered data of information flows and implies a basic knowledge of the problems of operation and development of the systems used.
The proposed element of the mathematical model and method of application methods, complements the initial approach in the analysis of system performance parameters of time, in order to decide on the criteria considered exposed original data generated as part of the block diagram.

Keywords: functions system, systems of fighting, generation new function of aircraft, control algorithms, functions model of systems, methods intellectual the analytics of data, experts systems

Performance analysis of an integrated strapdown INS based on MEMS sensors, and improving its accuracy by compensating instrumental errors are discussed.
Development of mathematical models for computation of MEMS sensors` errors; system errors algorithmic compensation by use of the mathematical models described below; development of calibration procedures for MEMS sensors, laboratory testing of calibration procedures and estimation of the results obtained; MEMS-based strapdown INS testing, consisting of the following steps: laboratory temperature tests and field testing. In the process of carrying out ground tests of MEMS-based strapdown INS a number of short term trips were performed. Each trip included a stage of initial alignment on the fixed base and a stage of motion between two waypoints. An initial angle of azimuth orientation was laid by means of magnetic compass. Azimuthal orientation was corrected according to GPS receiver track data during movement. The dead reckoning error of the system under test was estimated by comparing its data with output of a reference strapdown INS integrated with GPS receiver.

•  performance analysis of MEMS sensors and the system in whole was carried out;
•  an accurate mathematical model for computing errors of MEMS sensors for their estimation and compensation was developed;
•  calibration procedure for MEMS sensors allowing for minimizing the number of operations and automation using bench equipment was developed; testing of calibration procedures for a miniature micromechanical system with estimation of its performance was implemented;
•  enhanced precision performance of micromechanical INS in autonomous mode when compensating errors in the process of system calibration in accordance with mathematical model was achieved;
•  laboratory temperature tests of MEMS-based SINS have been performed;
•  field tests in the real vehicle motion mode were performed
  

•  temperature drift compensation of angular rate sensors and accelerometers (zero signal) is applied;
•  the state vector of the integrated data processing filter utilizes the mathematical model of angular rate sensors` drifts defined depending on temperature and linear accelerations;
•  shift parameters of accelerometers` zero signals are included in the state vector of the filter.

Keywords: navigation system, rate sensor, accelerometer, aggregation of information

Approaches to realization of the neural network (NN) regulator in the automatic control system (ACS) of an unmanned aerial vehicle (UAV) are studied in the paper. UAV flight in the approach and landing modes affected by wind disturbance in the vertical plane is considered. The structure of the NN and modeling of NN regulator training results are presented.
The UAV motion in the longitudinal plane is considered on the assumption of indicator speed stabilization by means of auto-throttle. At a stage of final approach maneuvering the preselected flight altitude is maintained, after crossing the glide path the UAV ACS eliminates the current height mismatch and provides its transition to a mode of holding of desired track coinciding with the direction of the glide path. The appropriate equations of UAV motion are presented in [1]. The UAV control in the vertical plane is realized as follows: under big deviations from the preselected altitude of flight when descending at the preset constant airspeed stabilization the control signal (1) on the elevator drive arrives:

(1)

After the deviation from the preselected altitude of flight becomes less than 10 m, the control law switches to the following one:

where  — coefficients of the control law, — increments of UAV coordinates,  increments of estimates of the UAV pitch and angle of attack caused by vertical wind derived as a result of estimation by means of the optimum Kalman filter [2]. The first control law (1) provides maximum system response speed when eliminating altitude mismatch, and the second control law (2) provides accuracy in maintenance of the desired track.
Transitions from one control law to another can cause loss of UAV stability. To prevent such occasion the presented regulators were realized on the neural network base which will provide the specified system response speed and continuity of control signal feeding the elevator drive, excluding thereby probability of UAV stability loss.
For creation of training data access there were used files of coordinate data included in control laws which are obtained by modeling of operation of the regulators mentioned above under various initial data: initial deviation from the desired altitude ΔH and speed of vertical wind Δ_αw. The following meanings of initial data were taken: for ΔH=0 m, 50 m, 100 m, 150 m, 200 m, and for Δ_αw=0 m/s, ±1 m/s, ±2 m/s, ±4 m/s. The NN output control signal feeds the UAV elevator drive.
The program of formation, training and testing of NN operation was developed. Results of modeling of NN training for initial deviation from the desired altitude of ΔH=50 m, and a vertical wind speed of Δ_αw =-1 m/s were analyzed in the paper. It was shown that on the 1000th epoch of NN training the value of a root mean square error amounted to 4.17 • 10^-4 .
Thus, NN solution of the control system developed reproduces a command signal with the adequate accuracy.

Keywords: unmanned aerial vehicle, automatic control system, neural network

Aluminum based chemical power plants use oxygen as an oxidizer, and that is why they are called as the aluminium-air power plants (AA PP). An AA PP is an energy efficient one among the currently known chemical current sources (CCS) in consequence of the high density of energy output (720-2340 kJ/kg or 250-700 Wh/kg), and it trails only the oxygen-hydrogen (О₂/Н₂) fuel cells (FC) with cryogenic storage system and some CCS with the lithium anode.
Based on the AA CCS study fulfilled by us it is possible to create power plants for different applications with a wide range of power output — from several watts to some tens of kilowatts. The construction arrangement of such power plants will obviously be substantially different — it depends on the power, the type of electrolyte and the design map of a power plant.
This paper presents the processes occurring in the AA CCS and AA power plants and the basic diagrams of AA PP designed for different purposes; besides, we offer main requirements to the auxiliary and service systems of such power plants.
The novel solutions for the main units of AA power plants are offered in this paper. We examine also the important process of the electrolyte clarification from the reaction products. This process is one of the limited processes in the CCS, which influences the working life and energy output of the AA CCS.
However, the operating time and the life cycle of the AA power plant are limited by the anode passivation and the destruction of the gas-diffusion cathode, which are the consequences of the reaction products (namely: aluminates and solid hydroxides). These characteristics can be improved by organizing the electrolyte cleaning system.
For increasing the AA PP operating time, we examined several ways of organizing the electrolyte cleaning system. Such ways are associated with a periodic or constant process of decomposition of the oversaturated aluminate solutions. They are associated also with a presence or lack of any crystallized component in the cells and the possibility to combine this component with the assembly of electrolyte cleaning system.
Parameters of the electrolytic circulating contour were assessed. That assessment showed that at the high decomposition rate (which conforms to the high concentration of the dissolved aluminum — the saturation rate is 1.5 and higher), the process of the crystallization proceeds in the electrolytic circulating contour.
This paper shows that sometimes for intensifying the process of electrolyte cleaning it will be required to use some special arrangement — crystallizer. The desired seed concentration in the crystallizer will be organized. The application conditions of the PP determine the optimal construction and the working diagram of the crystallizer. It is also shown that it is possible to use three types of units for cleaning the electrolyte from solid phase: sumps, filters and centrifugal separating devices.

Keywords: aluminium, anode, air, corrosion, organic inhibitor, polarization, electrochemical cell, electrolyte, electric power plant

This paper focuses on waveguide layout in the ground-based large-sized phased antenna arrays. Automation of the waveguide transmission lines designing is considered. Subject of the research includes routing methods and algorithms and its application to the waveguide layout meeting radio and technological requirements. The main purpose of the research is to optimize waveguide layout and to decrease design and engineering time and costs.
In the paper, it is justified that waveguide layout is a multicriteria optimization task. In this case, it is used decomposition principle. Waveguide layout is realized using topological-geometric method and consists of two steps: geometric layout and topological layout. The geometric layout stage includes layout under the condition of minimal lengths of the traces. At this step it is used linear-programming technique. The topological layout stage includes a correction procedure to obtain equal lengths of waveguide lines. A new method of equal length waveguide transmission lines routing is suggested. Proposed methodology includes calculation of the waveguide lines lengths and choosing the maximum length, definition of the clear connection field area to traces lengthening; unlocking of the traces fixation in the dedicated areas; correction of the each waveguide line length to achieve basic length.
Optimality criterion at this stage is minimum number of waveguide bends. There is proposed segment connecting field model which is based on adaptive radial grid. Using adaptive radial grid gives us following benefits: decreasing problem solving time, improving layout quality, random layout angle, adaptiveness of the grid.
The proposed methodology and algorithms are realized in the software module WDS (Waveguide Design Solution) based on SolidWorks system. The module is a Windows application which integrates both with MS Access (to import initial information) and SolidWorks system (to export output information). Field of Application of methodology, algorithm and WDS program described in the paper includes predesign and preliminary design of waveguide transmission lines located inside array structure.

Keywords: route tracing, topological-geometric routing method, waveguide transmission line, phased antenna array, computer-aided design system

The article is devoted to creation of fabric armor package mathematical model, its calculations and experimental study of protective characteristics. The steel balls were chosen as simulator of destroying impact. Each layer of armor package was separately considered as a sphere during the supersonic hit period or as a combination between sphere and tore during the following period. Interaction between layers was defined by pressure and friction. The mathematical model allows to determine the number of punched fabrics layers depending on mass and velocity of the ball. Corresponding dependencies are given in the article.
The slippage of the ball between threads and armor package overpunch was observed in preliminary tests. The placement of polymeric film in the package excluded the possibility of overpunch due to the slippage. The best results were obtained with the help of polyvinyl chloride film with thickness 0.5 mm placed between 25th and 26th layers of 30-layers package. The tests results concerning the number of punched layers mostly correspond to the calculations.

Keywords: airplane fabric armored jacket, steel ball, the mathematical model, results of calculations and tests

An important part of the budgeting system at the industrial enterprises is the budget for direct costs of basic materials. This budget reflects the need for basic materials and components for the production program, the size of the carry-over stocks of materials and costs of purchase of materials differentiated according to the months of the plan year. At the enterprises of aviation industry the formation of the budgets is associated with certain difficulties connected with a long production cycle and the unequal distribution of costs in stages of manufacturing. At the beginning of the production cycle, the need for basic materials is great. Then, at the stage of machining material costs are significantly reduced, as in the process of production auxiliary materials are mainly used. Another burst of need in the material, and hence the financial resources, is associated with the beginning of the build stage, especially Assembly facilities. This is because the production process involves expensive finished and semi-finished products, in particular various electronic equipment.
The article presents an algorithm of calculation of norms of material consumption per product, differentiated according to the stages of the production cycle. The algorithm takes into account the lead in time of the launch of structural elements of the product related to the date of completion of final Assembly. The proposed technology allows substantiate the need for material and financial resources more accurately, which creates conditions for the timely implementation of production tasks and increases the effectiveness of working capital management of the enterprise.

Keywords: budgeting, direct material expenses, direct material expenses, production cycle

This article describes the main features of fixed capital assets renewal, goals and algorithm steps strategy in natural and value forms. The results of the research can be applied in developing the methodology for strategic planning of space-rocket enterprises fixed capital assets’ renewal.
The implementation of the proposed concept will allow to develop effective strategic plan of renewal the enterprise fixed capital assets production of, which in turn will create the conditions for optimizing the cost of fixed assets reproduction, improving reliability and durability of the equipment for the competitiveness growth and for the enterprise as a whole.

Keywords: reproduction of fixed production assets, strategic planning, the space-rocket industry

Creation of the Russian national innovative system (NIS) while causes difficulties because of weakness of private business in innovative industries. However it is possible to begin with creation of the branch innovative systems (BIS) in aviation industry and electronic industry as the national security of the country depends on innovation of these industries in many respects. In the USA and other developed countries there are the state and state-private organizations carrying out coordination of research and development in aerospace industry and advising the military-political country leaders for decision-making. Besides, close cooperation between aerospace scientific centers, leading technical universities and aircraft manufacturing corporations is adjusted.
As well as for NIS, BIS it is formed within the general state macroeconomic policy and the regulatory legal base providing realization of this policy. Basic elements of BIS are the following subsystems: generation of knowledge (education and system of professional development of shots), science (at Universities, State Scientific Centers, R&D department of corporations), the transfer of knowledge in production and services, the effective organization of production, financing of all subsystems. Considering that functioning of BIS is under construction of market economy, the market of the hi-tech production and services can be considered as one of NIS subsystems.
The most real way of a transfer of innovations in the Russian NIS at the present stage is distribution of technologies in the form of the new equipment and forms of the organization of work. That fact is obvious that innovative activity of the Russian aviation firms depends on use of the technologies created out of these firms more and more. Knowledge of technologies can be received from information received in higher education institutions and at advanced training courses, publications in special editions, aircrafts shows. The help can render active use of outsourcing, in particular at introduction of information technologies.

Keywords: sectoral innovation system, NASA, innovation process, United Aircraft Corporation, innovation transfer

Factors that form cost of the innovative product at different life-cycle stages were considered, also cost formation approach "from the developer and the manufacturer" was used, that is the "prime cost plus rate of return" model, as well as scopes of life cycle’s cost index were considered.
The method of the system analysis and an assessment of products, and also method that allows calculating a cost index of life cycle in dynamic statement for the entire period of life cycle – the discounting method are used in the present article.
Main results of the present work are the following. The factors that form cost of the innovative product at different stages of life cycle are considered The accounting items’ structure for calculation of life cycle’s components cost is given, and also the general principles of calculation and resultant model of calculation of the life cycle cost of the innovative product at the aircraft industry enterprise in dynamic statement, taking into account time factor are stated, approach "from the developer and the manufacturer" is considered. The main practical approaches to formation of cost of the innovative product on the aircraft enterprise only are considered in the present article. The offered concepts and methodology of calculation may be used while carrying out the feasibility study on creation of innovative products on the aircraft industry enterprises ( assessment of requirement for investments; assessment of competitiveness of new development; assessment of efficiency of new development; assessment of expediency and feasibility of the project of creation of the innovative product), and, further, taking into account specifics of specific objectives, can form a methodological basis when carrying out the feasibility study on innovative products and processes in engine-building.

Keywords: aircraft industry enterprise, innovative product, cost, life cycle, discounting, competitiveness, conductor and producer point of view

Over the past decade enterprises of the aviation industry have successfully issued and placed corporate bonds.
The experiment of issuing bonds dates back to 2002. Analysis of bond issues by enterprises of the aviation industry demonstrated that the market for bonds of enterprises from the aviation industry is developing. Enterprises prefer to issue bonds for no more than three years, thereby reducing costs on the bond issue and maturity date. The maturity date of each issuer differs and depends on the goals of the issue and intended use of the proceeds. The main goal of a bond issue is to raise funds to finance a company’s investments and core production activities, and also to repay credits taken out by the company and also by its subsidiaries and associates. Notwithstanding the difficult financial position of the issuers, they have effectively managed their debt commitments in respect of the bonds. The coupon yields on bonds of the aviation industry over the past three years considered here are at the level of 8.25%, which complies with the refinancing rate of the Central Bank of the Russian Federation. The performance ratios of the bond issues of enterprises from the aviation industry are presented in a table in this article. All the bonds of the aviation enterprises were assigned a BB rating, which implies that investments in these bonds do not imply short-term risk. At the same time, there are significant uncertainties due to the issuer’s sensitivity to adverse business, financial and economic conditions

Keywords: bond, duration, modified duration, life cycle