2014. № 74

In the paper, we present a constructive algorithm for synthesis of the suboptimal information constrained control law for quasi-linear stochastic dynamical systems. This algorithm can be used directly for solving applied control problems. The information constraints are described as follows. We suppose that each control vector component depends on a prespecified set of precisely measured state vector components.
The aim of this article is to concretize our new results for the efficient practical usage. We propose to construct the stochastic quasi-linear system control laws as a linear in state function such that the linear parameter and the constant term of this function are polynomial in time. This control law is called suboptimal. If an optimal control can not be constructed for some reason, then the suboptimal one can be used. So, it is very important to have an algorithm for synthesis of the suboptimal control law. In the paper, we present the algorithm based on gradient descent numerical method.
The algorithm is used to construct the suboptimal control of a two-link robotic arm. The robotic arm control process is considered as a mechanical manipulator plane movement. The goal of control is to move the manipulator to a known state. The goal must be achieved in a specified time period. In addition, the control cost must be minimized.We show that the suboptimal control can be effectively used for this problem. Despite the fact that the result is not an optimal solution, it has a simple linear regulator structure, which can be directly used to control the robotic arm. The comparison of the suboptimal performance index with the optimal one shows that the difference between these solutions is very small.

Keywords: stochastic optimal control, quasi-linear dynamical system, information constraints, suboptimal control, two-link robotic arm

At the present time the working out and making of hypersonic flying machines (HFM) which can become a base of perspective systems for deducing of payload and vehicles is actual. The intensive aerodynamic heating of nose parts and forward edges of wings and other devices of construction HFV leads to the necessary presence of thermal protection systems.
The purpose of the given work is carrying out numerical estimation of functioning parameters of thermionic thermal protection (TTP) of HFV construction devices (CD) in condition of an aerodynamic heating. On the basis of obtained estimations it is possible to make a decision about expediency of integration TTP into a HFV composition of various types and determinate possible onboard consumers of generated electrical energy.
At the heart of TTP the phenomenon of emission of electrons by heated metal (a thermionic emission) is necessary. During to the motion in atmosphere with hypersonic speeds the exterior shell heats up to temperatures, at which because of the expense of thermal energy obtained by aerodynamic heating from its interior surface the hot electrons besieged then on an interior shell (anode) start to take off. It means that at the given stage electrons are carriers of heat or «coolants» of an exterior shell, which is the cathode. That’s how an electronic cooling of an exterior shell and principle TTP is realized.
The basic particularity of TTP is essential decrease in thermal action on construction HFV devices during aerodynamic heating by the expense of thermal energy transformation of heating CD into a significant amount of electrical energy onboard HFV. Devices realizing TTP can be parted on two types: with interior issue and with exterior issue.
The TTP mathematical model of wings forward edges and nose parts, and TTP design procedures were carried out for achievement of the specified purpose. Developed model and design procedures have the worldwide novelty confirmed by patents for the invention. On the basis of obtained effects the analysis of possible HFV perspective onboard systems (consumers of electrical energy) is carried out.
The TTP realization will allow to provide implementation of other systems guided on the maintenance of long-term hypersonic flight, for example, methods for magnetoplasma aerodynamics.
The given analyses is a basis for examination of thermionic thermal protection.
Thermionic thermal protection is a new object of space-rocket technics. Obtained effects provide modernity for new object of space-rocket technics.

Keywords: thermionic issue, electronic cooling, electric energy, the hypersonic flying vehicle, thermal protection

This paper provides a brief description of the software application designed for detailed strength analysis. Department of Development of algorithms and programs of Stress Division, which is a part of engineering department of IRKUT CORPORATION, has been developed this program for several years. This application is designed to reduce the time required to perform detailed strength calculations of the standard airframe construction.
To date, the regular part of the airframe MC-21-300 (about 60% of all construction) is automatically analyzed using software application by engineering methods. Reports are automatically generates for all calculations.
All calculations are based on a global finite-elements model of a plane, which is a maindata sources for strength analysis.
For the moment the application includes the following modules of the detailed analysis, which were developed in the department of Development of algorithms and programs of Strength Division.
FrameStrength is a detailed strength analysis of the frames. Analysis is carried out by engineering methods based on the guide for engineering TsAGI. This module has interactive mode for analysis of the joints with unspecified configuration.
StringerStrength is a detailed strength analysis of the join stringer panels. Analysis is engineering methods based on the guide for engineering TsAGI. This module has interactive mode for analysis of the stringer panels with unspecified configuration.
FastnersAnalysis is a detailed strength analysis of the joints of frames, stringers and skins. Analysis is executed by engineering methods based on the guide for engineering TsAGI. This module has interactive mode for analysis of the joints with unspecified configuration.
DiagonalField is a module intended for analysis of skin reduction in the global model (based on the methods of TsAGI). This module has not an interactive mode.
OrthotropicStability is a module intended for buckling analysis of orthotropic (design-orthotropic and composite) skins in a global airframe model (based on the energy method of the stability rating). This module has not an interactive mode.
CrossSectionAnalysis is a module intended for the strength analysis of the twin-walled bar.
The main part of this application is the coordinating unit (working title — AVALON). This unit synchronizes data generation from the global model to perform the detailed calculations and record of safety factor and results of detailed calculations in the database with the global model of the plane. Also, this unit controls the process of reports creation.
Due to the high level of automation the cycle of typical construction calculations is reduced to 36 hours.

Keywords: global finite-element model, strength analysis, airframe structure, detailed stress analysis

Purpose
Improvement of airborne system performance inevitably leads to more stringent specification for their testing equipment, in particular, for motion simulators. At present, there is an actual task to improve the static positioning accuracy of the end member of motion simulators. In this connection, the issues related to assessment, measurement and improvement of motion simulator positioning accuracy take on great importance.
Design/methodology/approach
The paper proposes a methodology which was developed for determination of the end member positioning error in the gimbal based motion simulators that have several axes with their initial orthogonal orientation. This approach allows considering influence of the disposition error of motion simulator axes, unit under test mounting error, as well as errors caused by actuator and the measuring system.
In the framework of this technique development, the mathematical model of the end member positioning error was composed with the help of matrix methods. The author proposed a technique for determination of the mathematical model coefficients for motion simulators that have from 1 to 3 axes.
Findings
As a result of the study, a mathematical model has been developed along with software which according to the given manufacture and assembling errors of motion simulators calculates the maximum error of the end member position. The program has a practical value because it can be used while forming the requirements to the motion simulators, and when choosing optimal design solutions for the motion simulators.
Research limitations/implications
The article considers the motion simulators built on the base of gimbal mounts that have orthogonal axes orientation as a nominal case. Mathematical model doesn’t take into consideration the linear error of dynamic test bench axes position.
Originality/value
Unlike the well-known methods of determination of the motion simulator end member positioning error, the method offered allows easily automate the process of error definition for any kinematic scheme of the motion simulator gimbal mount. This property was realized in the program for calculation of the maximum control signal execution error by the motion simulator end member. This program allows calculation of errors for motion simulators having from 1 to 3 degrees of freedom with arbitrary mutual position of axes, and considers the influence of the dynamic test bench axes position errors, as well as that of a mounting error of the unit under test, and that of the errors brought into by the actuator and the measuring system.

Keywords: motion simulator, gimbal mount, positional error, kinematic accuracy

The problem of landing dynamics investigation remains one of the most important tasks in airplane performance calculation. Its significance is especially emphasized by the fact that most aircraft accidents still occur during landings. It is necessary to take the following factors into account for the detailed investigation of the important physical aspects of the considered problem:

· the strongly nonlinear behavior of the landing gear shock absorbers,

· the Coulomb friction forces in tires and seals,

· the contact interactions between moving parts of shock absorbers,

· the contact between rotating deformable wheels and ground,

The developed mathematical model of the landing gear legs demonstrates all of the required functional parts of the real mechanism. Mass-stiffness beam based airplane model is linked to landing gear legs. The whole described system is solved by direct integration method with the usage of MSC Software products (MSC.Nastran or MSC.Marc). A specially developed program code for MSC.Patran preprocessor simplifies building and editing of both aircraft and landing gear legs models significantly. It also allows the designer to carry out the optimization.
To verify the accuracy of the developed aircraft model its calculated natural frequencies were compared with the results of the modal tests of the real structure. Drop tests were simulated to verify the operation of the shock absorbers of the main landing gear legs. Both of these verification checks showed good agreement between the data obtained from the model and during the experiments, which allows the researchers to use the developed methods for further practical research.
Both symmetric and asymmetric landing cases are investigated on the example of Tu-204SM passenger airplane. Calculation results are compared with the flight test data and information from several published works [2, 3] (the comparison was made, for example, according to the shape of the curves that describe the dependencies between the reaction forces, accelerations and time during the landing). The developed model was also used for calculating landing loads and investigating the operational efficiency of landing gear shock absorption.
Application of the Patran Common Language for the automatization of the model building process and usage of the widely employed MSC.Marc solver make the created methodology more clear and friendly to new users. The author expects that this feature would help to widen the scope of practical application of the developed methods and models. Besides the author considers solving the optimization problem aimed at attaining maximum operational efficiency of landing gear shock absorption as a very important practical task, the appropriate investigation of which is planned in the future.

Keywords: airplane, landing dynamics, mass-stiffness model, shock-absorbers, numerical methods

The main purposes of the operation of the complex for endurance bench testing of the full-scale attack airplane structure, which provides an extension of the design service life of the aircraft fleet, are the following:
— experimental verification of the remaining service life of the aircraft’s airframe and landing gear structure;
— location of the elements and units, which are important from the point of view of fatigue conditions, within the structure of the airframe and landing gear during the recreation of the operational spectrum of the variable loads;
— determination of the duration of fatigue cracks evolution and residual strength of the damaged airframe structure;
— determination of durability after overhaul.
The endurance tests of the two-seat combat trainer attack airplane were carried out according to these goals wherever possible. Airframe fatigue tests were carried out for the following types of cyclic loads: A) in-flight loads; B) takeoff and landing loads; C) ground loads. Additionally the following fatigue tests were performed: rudder control system linkage tests; elevator control system linkage tests; aileron control system linkage tests; tests of undercarriage doors in closed position; tests of landing gear up locks; tests of the main and nose landing gear legs by applying the loads that emerge during their retraction and extension.
After the fatigue tests were finished, the residual strength of the airframe structure, which was damaged and contained fatigue cracks, was tested.
Test methodology
The tests were performed by using a full-scale serially manufactured two-seat combat trainer attack airplane. The following airframe components were subjected to simultaneous fatigue tests: wings with leading-edge slats and trailing-edge flaps; fuselage; horizontal tail; elevators; engine mount fittings; mount fitting of brake parachute lock. Horizontal loads were balanced by the loads, which were applied to the brake parachute lock and main undercarriage legs. Variable loading was carried out by a flight cycle program block, which corresponded to 100 flight hours or 90 flights. At the same time the fuel tanks of the wing and fuselage were stressed with excessive pressure. The program block consists of 5 flight cycle types, which differed from one another in terms of maximum and minimum overloads in maneuver configuration. Each flight cycle included two different loading modes: in-flight configuration (FC) mode and maneuvering configuration (MC) mode. MC mode loadings were carried out with the extended high-lift devices of the wing (δ_slats=9º, δ_flaps=20º). Retraction and extension of the high-lift devices were performed after the appropriate overloads of n_y=2,5 were attained. In FC mode the high-lift devices were retracted and loaded as an integral part of the wing.
A specialized test site was created to properly perform all of the required endurance bench tests.
The testing facility included the following main systems and components:
— airframe loading system;
— fuel tanks pressurization system;
— oil pump station and hydraulic system;
— automated system of multi-channel electro-hydraulic loading;
— information and measurement instrumentation system.
Research results
Strain-gauge sensor bridges were mounted at the right half of the wing to control the variable loadings during the fatigue tests. The sensors were placed according to the same layout as the one, which was used during the measurements of the real in-flight loads. The sensor bridges were calibrated during the application of the step-stress loads. After that their readings were recorded during the application of cyclic loadings. At the same time the dynamometer readings were measured and used to calculate the bending moments within the wing sections where the sensor bridges were mounted. The comparison of the sensor bridges readings and calculated bending moments allowed the research team to estimate the amount of the error.
The tests of the airframe structure residual strength were initiated due to multiple structural damages. During these tests the strain-gauge sensors were mounted on the undamaged ribs in the zones with fatigue crackings. These sensors were used to trace the step-by-step sequence of fractures of the fitting attachment during the residual strength tests.
Practical implications
The endurance tests have allowed the research team to increase the durability of the transverse joint of the central wing box lower panels and the detachable wing half considerably. In particular the following measures should be taken:
— the tests should be carried out with certain loosening (to ¼ of the turn) of the tightness of the joint bolts, which surround the fractured rib of the fitting attachment of the central wing box and detachable wing half;
— it is necessary to break corners around the flange holes with the improved surface smoothness, which are intended for fitting bolt mounting, to increase the durability;
— it is necessary to use the tension-control bolts in rib holes, which are intended for installation of covering hinge bracket.
Summary
The presented complex for fatigue and survivability tests of attack airplane structure has confirmed the possibility of the substantial increase of airframe and landing gear structure life on the condition that the airframes of the operated airplanes are subjected to the appropriate modifications.

Keywords: дефектоскопия, дифрактометр, нагружающее устройство, напряжение, объект испытаний, переменное нагружение, ресурсный стенд, тензометрия, усталость, фрактография

Preventing injuries of passengers and crew members of the modern helicopters is a key factor for providing competitive success on the transport vehicles market of today. However, the existing equipment that ensures injury prevention for helicopter passengers under emergency conditions is represented by complex engineering systems, which are based on the obsolete energy absorbing chair design. Such chairs are built according to a concept with a uniaxial shock absorption system and safety belts, which cannot ensure complete passenger safety during an emergency landing.
Neck and vertebral column injuries are the most dangerous and probable passenger traumas, which can occur during an emergency landing due to limited physiological capabilities of the human body under the overload exposure. Such injuries are caused by the «head nod» (rapid downward movement of the head) effect. «Head nod» is caused by a moment due to the misaligned positions of centers of mass of human head and body. A new concept of injury prevention system has been developed based on leading edge technologies to eliminate this moment and prevent nodding. This concept focuses on a passenger chair design, which incorporates a neck protection system, programmable deformable elements and shock absorption system. Neck protection is based on the use of the D3O innovative material. D3O is a newest-generation energy-absorbent material, which is capable of protecting passenger’s neck against injury without causing any discomfort. D3O is also used in chair upholstering, which contributes to further improving passenger comfort and additionally relieving his vertebral column regions. The shock absorption system employs programmable deformable elements. These include shear pins and chair frame slide with controlled destruction, which occurs when the loads exceed critical level. The guide rails of the shock absorption system are formed into arch shapes. Such shapes ensure partial change of the direction of the overload force vector from vertical plane towards horizontal and thus provide better compensation of the moment. The method of safety belt attachment has also been modified to improve passenger fixation.
Introduction of these elements combined with the use of shock absorbers with adjustable rigidity will allow the designers to both prevent neck injuries and reduce loads on the upper, middle and lower regions of the vertebral column.

Keywords: injury prevention, programmable deformation, neck protection system, minimization of the acting force, energy-absorbent material, D3O

Object: Parameters of reliability and target efficiency of the spacecraft;
Subject: Spacecraft performance evaluation expressed in taken images of the Earth surface area and monitoring periodicity;
Purpose: Development of simulation models and algorithms for influence of onboard supplying systems partial failures and target equipment on parameters of spacecraft target efficiency;
Methodology: In the article are used: logical-and-probabilistic methods for the analysis of onboard systems functionability, a method of simulation modeling and statistical tests. The analysis of statistical material involves a classification of failures with respect to the target operation of spacecraft stand-by time. Construction of failures mathematical models is based on the construction of onboard systems failures trees. Failures trees construction of observation spacecraft is based on the construction of onboard systems failures trees.
Findings: On the basis of algorithm for an estimation of shooting productivity expressed by quantity of finished shooting objects of supervision, the estimation algorithm of shooting productivity is developed. An estimation method for monitoring periodicity of Earth remote sensing spacecraft taking into account the reliability of on-board systems is suggested. This method is based on the spacecraft target functioning simulation with consideration of onboard system failure.
Practical implications: Results of the provided researches are used at early design stages of the spacecraft. Research results can be used at early stages of onboard systems development and their parameters estimations.
Originality: Mathematical models of onboard systems failures, algorithm for simulation modeling of partial failures influence on parameters of shooting productivity and Earth monitoring periodicity of the spacecraft and the corresponding software are developed. By means of the software it is possible to construct relation between shooting productivity and monitoring periodicity with respect to the reliability parameters of spacecraft. Such relations allow to estimate losses of output effect from reliability level of spacecraft. From the other hand it is possible to set spacecraft reliability parameters norms taking into account onboard systems failures.

Keywords: spacecraft, Earth remote sensing, parameters of target efficiency, shooting productivity

This article deals with carrying out missions of employing a large number of small spacecraft (SSCs) including those of "CubeSat«format. This task is associated with an increased number of «CubeSat» small spacecraft being developed by educational, scientific and industrial organizations which will result in the need of expanding the market for launching «CubeSat» spacecraft. Currently, the growth of the number of «CubeSat» projects allows to predict the launch of 30 to 50 satellites annually within the next 6 years. Given the growing number of SSCs projects having practical value and developed in the interests of government agencies it can be said that the task of launching these spacecraft becomes highly specialized.
Currently, «CubeSat» SSCs launches are mainly run within larger expensive missions that have several disadvantages, such as the difficulty of choosing an appropriate mission; being bound to the main spacecraft orbit; long preparation cycle of main payload launch.
Samara Space Center is developing and producing a number of products that can be used for employing «CubeSat» SSCs cluster missions. Above all, this is «Soyuz- 2» Phase 1B rocket ( two-stage light class booster) designed to launch satellites from «Soyuz- 2» launch complexes. To ensure employing the payloads into orbits with a wider range of required accuracy a «Volga» deployment unit was designed. Application of «Soyuz- 2» Phase 1B rocket together with «Volga» deployment unit increases the maximum payload mass. The first launch of «Soyuz- 2» Phase 1B rocket together with «Volga» deployment unit was held on 28 December 2013.
To organize and cluster launch «CubeSat» SSCs we have developed «The Matrix» launch pad designed for simultaneous orbit employment to 512 SSCs from 1U to 6U, format totaling up to 1536U. This provides the ability to set the launch pad and from 1 to 4 SSCs with the specified parameters (different from «CubeSat» format) to separate. Meanwhile, the launch pad is a place to accommodate transport and launch from 1U to 6U «CubeSat» SSCs containers.
«The Matrix» launch pad is set onto the unit being launched in order to rotate on one axis. Thereby it provides a desired angle of separating the SSCs. In turn, the «CubeSat» SSC are installed onto the launch pad in a unified transport launch canister which separates the SSC with the specified linear and angular velocities.
Launching such a large number of SSCs requires a special approach to the organizing the launch: it require establishing a special section on the website of the Samara Space Center www.samspace.ru ensuring the possibility of submitting online applications and online tracking the process of formation of a cluster launch online.
Thus, Samara Space Center has a number of technologies to provide both single and group launch for SSCs including the «CubeSat» ones. Development of special tools, such as «The Matrix» launch pad, provides the simultaneous employment of a large number of SSCs on a wide range of operating orbits and thus ensuring high technical and economic efficiency of SSCs missions.

Keywords: small spacecraft, CubeSat, general-purpose platform, launch development, web- technologies, transport and launch container

Today, small-sized Spacecrafts with mass up to 500 kg are applied to fulfill a wide range of special space tasks. Electric Propulsion Systems based on Stationary Plasma Thrusters (SPT) are mostly used aboard such small-sized Spacecrafts.
This paper is dedicated to the development of a perspective scheme and creation of a Hall Effect Thruster ensuring high thrust performances, and also to research of influence of constructive parameters and operating mode on thruster efficiency.
The most challenging design is SPT with a hollow anode, which experimental model SPT-1 has been developed by EDB Fakel and has been tested in Russia and USA. Development of the perspective high-efficiency Plasma Thruster of low power is based on the constructive scheme tested on SPT-1 by its mass-energy optimization.
Thruster operating efficiency is determined by operating efficiency of its magnet system, and also by organization efficiency of ionization and acceleration processes in its discharge chamber. This paper presents the results of development of a perspective constructive scheme and the results of optimization of the main elements of a magnet system and discharge chamber of the thruster with a hollow magnet anode of low power PlaS-40. As a result of optimization, the reliability of thruster operation and thrust performances level are increased by loss decrease in magnetic circuit and increase of gas distribution and ionization efficiency in discharge chamber.
As a result of PlaS-40 researches, it is determined that thruster operates stable in the range of power of 100 to 650 W at a discharge voltage of 100 to 500 V and at a discharge current of 1.00 to 2.25 A, and also during life test.
Thruster PlaS-40 allows to reach high level of performances, which are ensured with the thruster of a bigger dimension type as, for example, SPT-70. An application of the new PlaS type thrusters for the similar tasks solved by SPT allows to reduce mass and volume occupied by them aboard S/C by ~40%. Implementing several PlaS type thrusters aboard the S/C Electric Propulsion System, the overall benefit will significantly increase.
Based on the obtained results of the researches and the mass-energy optimization of the PlaS-40 design, it is supposed to further improve the PlaS type constructive scheme and to develop a bigger dimension type thruster model, such as PlaS-55 and PlaS-85 and to research theirs performances.

Keywords: electric propulsion (EP), small-sized spacecraft, stationary plasma thruster with a hollow magnet anode

The problem of wide range F ion sources (IS) constructing is very relevant nowadays. These sources are used in technological processes such as sputtering, cleaning, ion assistance, surface modification and as an electric propulsion system. The gridded ion sources are used in the applications where the ions with energies 200-1000eV are necessary. However such sources have limited service life. Possible alternative is an RF gridless ion sources developed by R.Boswell and C.Charles during the last decade. Its operational principle is based on the dense plasma production under conditions of the helicon waves excitation and ion acceleration in the double layer at the outlet of the IS. The maximal achieved ion energy is about 120eV.
The goal of the present work is to develop the low-energy IS with long service life and wide energy range (10-300eV). To do so three modifications of the RF gridless IS have been constructed and studied experimentally. The RF inductive discharge located in the external magnetic field was used as the operational process of the IS. The magnetic field value corresponded to the resonance conditions of helicons and oblique Langmuir waves excitation. The acceleration of ions took place due to the potential drop formed near the outlet of the IS. Besides the double layer formation due to the magnetic plasma contraction similar to R.W.Boswell experiments, potential drop was enhanced using an additional geometrical plasma contraction. For this purpose two IS configurations where provided with constrictions of different diameters near the outlet. The third IS configuration was constructed for checking the possibility of ions acceleration due to the potential drop appearing near electrodes in the capacitive RF discharge. To do so a coaxial electrode was mounted near constriction and connected to the inductor lowest coil. Between electrode and inductor a variable separating capacity was placed. This provided the existence of RF voltage between the lower part of the inductor and coaxial electrode. The RF voltage in the capacitive channel of the discharge could be controlled by the separating capacity value. In order to organize the acceleration of ions parallel to the IS axis the diameter of constriction was made smaller than the characteristic dimension of the near electrode sheath.
Experiments showed that the existence of the constriction at the outlet of the IS leads to the mean ion beam energy increase. The smaller the constriction and longer, the energy of the accelerated ions is the bigger. The idea to increase the potential drop near the outlet of IS due to electronegative self-bias effect using IS with capacitive channel was experimentally confirmed with average improvement of 20-60eV. As the result of the work an effective ion source with independent flexible control options has been developed. Such device can vary ion beam energy from 10 to 300eV and ion current value — from 0 to 250µA/cm2

Keywords: electric propulsion systems, ion thrusters, radio frequency, inductive discharge, helicons, oblique Langmuir waves

World ocean problems solution nowadays is becoming the most actual directions of earth remote sensing (ERS) from the space. The role of the world ocean in natural processes as well as in humankind life is commonly known. However, the exploration of the ocean is the most difficult problem of the science. The conservative methods for scientific study with the help of research vessels can not solve the majority of oceanographic problems immediately and on the rims. That is why the ocean remote sensing methods have begun to develop.
The purpose of this article is to present justification of the main technical, design, and process solutions to be implemented at building Meteor-M #3 SC developed by ‘VNIIEM Corporation" JSC.
The developed space platform for perspective oceanographic satellite Meteor-M #3 has been designed taking into account modern tendencies, great experience, and process stock of ‘VNIIEM Corporation" JSC accumulated during the long-term period of creation and operation of Meteor series SC.
The main objective of this satellite according to performance specifications is all-weather illumination-independent radar monitoring performed by on-board active array radar system (AESA). The perspective SC is planned to be equipped with some other special purpose hardware: scatterometer, ocean and nearshore zone color grade scanner, atmosphere radio translucence equipment.
The mentioned above types of equipment are differing exclusively by variety of physical principles, mass and dimension features, demand for power supply, requirements for maintaining thermal regimes and highlighting fields to be observed, as well as their electromagnetic compatibility shall be provided. All these significantly complicate their mutual embedding and operation in one SC.
Additional difficulties arise at selection of design-layout scheme and provision of spacecraft with electrical power if any sufficiently big and energy-intensive devices are in payload.
One of the most serious problems at determining configuration of space platform is achievement of high toughness and strength of the body and inadmissibility of small resonating frequencies of SC on adapter separation subsystem from «Fregat» upper-stage of «Souyz-1» booster of 1b stage.
The second problem of the configuration challenging for the given type of SC is a requirement to align the satellite mass with a minimal deviation from the center of its coordinates system.
An extremely complicated engineering problem is to provide a thermal rate of SC as a whole and of AESA in particular, as well. Besides, turns around of longitudinal axis to transfer AESA from the right board of SC to the left one are foreseen in the program. In this mode the SC thermal control system operates in extreme regime, since the sun rays fall on the exchanger and the solar array is found turned out from the Sun.
The present article deals with a justification of selection of optimal configuration of SC based on construction stringer mating ring main circuit, as well as on reticular prism and primary structural elements stiffness calculations.
Considering the mentioned above requirements and calculations, a horizontal configuration scheme based on an octagonal carbon fiber prism was accepted for this Meteor-M #3 SC.
The requirements and solved design solutions put into this configuration will enable to build a unified space platform of new-generation. Thus, this space platform shall become a baseline for development of the next generation of Earth remote sensing satellites at ‘VNIIEM Corporation’ JSC.

Keywords: space platform, satellite, stringer and transverse-based supporting structure, carbon-fiber cylinder, World ocean

— Purpose 
One of the most important components in the structure safety is the quality of aviation fuel. The presence of impurities in the fuel, such asmoisture and mechanical fine impurities of different physical nature, can lead to failure of the fuel system and aircraft incidents. 
Analysis of physical-chemical parameters of fuel directly before refueling the aircraft with the help of devices express control isincreasingly important. 
Thus, the task of developing high-precision analytical devices rapid quality control of aviation fluids is relevant. 
— Design/methodology/approach 
For rapid analysis of the quality of aviation fuel, an approach based on the application for information purposes microwave waveguide-antenna techniques. The essence of techniques is that measure of the measured value is the evaluation result of the topological deformation space-time structure microwave radiation interacted with the control object 
— Practical implications 
As a result of theoretical and experimental studies the layout of the instrument for rapid controlling quality aviation fuel was designed 
Advantages of the proposed device is to provide noncontact measurement, high efficiency and accuracy of measurements, a large number of simultaneously controllable parameters, the low cost of the primary transmitter and the minimum elements of the device. 
— Originality/value 
The originality and novelty of the proposed method are to improve the accuracy and efficiency of the control parameters aviation fuel through the use of for informational purposes change the mode ofthe waveguide transmission line, as well as the effects of the interaction of electromagnetic waves microwave with liquid dielectrics and ferrites.

Keywords: aviation fuel, microwave method, waveguide, electrophysical parameters

The assignment of some flying vehicle (F/V) sets specific conditions for propellants gassing — Nitrogen Tetroxide (NTO) and Unsymmetrical Dime-thylhydrazine (UDMH). For example, for space crafts it is strongly recommended the fuel has no dissolved gas. F/V assigned for permanent standby in a wide range of environment temperature drop, should be filled with propellant of certain gas rate concentration in it.
A range of technological methods are accepted all over the world to solve this problem. The most widespread methods are: for degassing — vacuum gas volume, heat release, bubbling under the vacuum; for gas saturation — mixing, bubbling, fuel spraying.
Basing on analysis results it is easy to see that each degassing and gas saturation method has a range of disadvantages. They are process duration, huge vapor loss.
While designing a modern filling system it is necessary to develop new highly effective technologies along with structural variations which permit to speed up the propellant degassing and gas saturation process with low medium loses (fuel, compressed gas).
To solve this problem authors performed an analysis of principles for increasing the heat and mass exchange efficiency which are basing on the following concepts: increasing of phase contact surface; increasing of mixing performance; improvement of phase contact process; using of unsteady interphase exchange regimes which allow to reach the immediate values of mass exchange coefficients; performing a mass exchange processes within the hydrodynamic instability of interphase surface.
Finally a new easy technology was suggested — fuel bubbling through the gas-jet rod radiators. Structural and metering characteristics of device also as a time span for a mass exchange processes were estimated for applying this technology.
Specific recommendations were given for efficiently applying this technology
Installation of gas-jet rod ultrasound radiators in a tank provides an efficient propellant mixing by means of high medium turbulization which splits the gas bubbles and intensify the mass exchange.
Gas-jet rod ultrasound radiators allow to reduce the propellant processing period and minimize the medium loses (fuel, compressed gas).

Keywords: mass exchange, nitrogen tetroxide, unsymmetrical dimethylhydrazine, degassing, satiation, gas-jet rod radiator, concentration, nitrogen, helium

Purpose
Coupled gas dynamic modeling allows taking into account the mutual influence of neighboring components in the design phase, improving the quality and reducing development costs to overcome the identified problems.
Methodology approach
Two approaches to CFD simulation of gas turbine engine were stated by the authors:
— simulation approach using a number of special programs each of which is best suited to describe the workflow of a particular engine component;
— simulation approach in one universal program that allows to modeling all the core’s components simultaneously at once.
The first approach allows calculating each component’s workflow in the most appropriate program with the optimal model and solver settings and involving the most appropriate physical models. This provides a better simulation of the processes and requires less computational resources because the GTE elements are calculated separately. The disadvantage of this approach is the necessity of data exchange between engine components that are modeling in different programs.
The second approach is free from such disadvantages. The computational model is created in a single universal CFD software package, consisting of several separate components, and data exchange is organized easily with standard tools of the program. However, the settings of the model are «universal» and certainly not optimal for each component in this case.
Practical implications
To illustrate the feasibility of end-to-end simulation of GTE workflow in universal software package, the authors calculated in CFD program the gas flow in single-shaft gas turbine engine. The computational model consisted of grid models of intake, the centrifugal compressor rotor wheel*, vaned diffuser*, reverse-flow combustion chamber**, axial turbine nozzle guide vane and rotor wheel* and also nozzle (*designed by Turbomachinery research group, headed Baturin O.V.; **designed by Combustion processes research group, headed Matveev S.G., SSAU
The same GTE workflow was also investigated at takeoff mode with a one-dimensional thermodynamic model developed in the program ASTRA. This program was developed by A. Yu. Tkachenko. Comparing the results of thermodynamic calculations with the CFD data it can be concluded that they are in good agreement with each other. The difference between the results is no more than 7%. Thus it can be concluded that the CFD calculations results do not contradict existing physical concepts of GTE workflow and can be used for its parameters calculation and simulations under various conditions.
Conclusion
The numerical modeling of engine core workflow is very promising. It allows predicting the mutual influence of engine components on each other, to investigate the impact of any working conditions and changes of passage elements on the GTE characteristics and all the components. For this reason, the research in this field will be continued in the future.

Keywords: gas turbine engine, CFD, mesh, coupled simulation, the balance of power

Objective
The analysis of the nanostructured high strength materials for sliding bearings of the rotor support in aero — gas turbine engines with the help of the vibro-diagnostic methods.
Methodology
To estimate the opportunity of implementing rotor support ceramic bearings in gas turbine engines, different versions of smooth hydrodynamic ceramic sliding bearings with similar to the roll bearings geometry in the series production were developed and tested.
The impact of the couple materials friction on the bearing operability was analyzed.
During testing the complex approach to the diagnostic estimation of the tested bearing conditions was optimized. The work was fulfilled in real time scale including spectral analysis of the bearing oil, evaluation of the thermal component conditions of bearing parts and its oil, vibration analysis of the experimental part on the test rig.
The Results
The efficient design of the smooth hydrodynamic ceramic sliding bearing for the rotor support in gas turbine engine was developed.
The complex method of evaluating technical conditions of the ceramic sliding bearing was designed what allows the timely estimation of the experimental part conditions during testing and instantly change the designed outer loading.
The sliding ceramic bearing design of a new generation confirmed its operability with lower oil consumption to 0,1 l/min.
The best operability factors were achieved with new generation ceramic couple used (composite on the basis of silicon carbide with the composite on the basis of titanium carbo-nitrid).
The area of implementation
The achieved results allow starting testing ceramic sliding bearings in the gas turbine engines. The defined features of the working processes would be used for further improvement of the bearing support design.
The achieved results may be implemented in the diagnostic system and used by the gas turbine operators for making decisions on the conformity of the tested bearing with the specification.
Conclusions
The advantages and shortcomings of the new generation ceramic bearings and means of diagnostics versus conventional roll bearings were analyzed.
The implementation of the nanostructured high strength ceramic composite materials for design and manufacture of the gas turbine engine bearing supports is acknowledged superior in comparison with the existing ones.
The implementation of the diagnostic accompaniment of the bearing analysis including real time scale decreased the number of stops and test rig partitioning for visual and instrumentation estimate of the tested part, thus reduced the total cost of work and duration of the experiment.

Keywords: gas turbine engines, bearings, vibrodiagnostics

This paper is focused on development of the test rig dedicated to aircraft engine turbine vanes study with various design measures to optimize turbine vane passage flow aerodynamics.
First test cycle was conducted for nozzle guide vanes (NGV) sector using advanced optimization technology for secondary turbulent flows near the walls with application of non-axisymmetric end walls. NGV sector was designed based on computational study conducted as part of the work [1].
The rig was designed to meet the below required performance and boundary conditions for tested NGV:
a) required flow quality at tested NGV inlet: uniform parameters profile, required boundary layer thickness and NGV gas stagnation angles;
b) required flow quality at tested NGV exit: no deviation of parameters across turbine passages, required radial distribution of parameters.
The rig was designed using, among other methods, up-to-date numeric flow modelling techniques. First test cycle steady-state modelling was conducted using ANSYS CFX 14.0 to estimate rig flow parameters.
This test rig is a wind tunnel consisting of the following key modules:
a) inlet duct: diffuser, settling chamber, straightening grid and three screens, contraction, NGV sector of 7 vanes (7-NGV sector);
b) test item: 7-NGV sector;
c) exhaust duct: test item exhaust duct, rig case, radial traverse, window.
Rig basic components are manufactured using rapid prototyping (RP) technology.
Required pressure ratio adjustments for the test item is reached using compressed air source (compressor) as well as control and bleed valves in working fluid supply and discharge lines.
Required boundary conditions for the test item are provided using dedicated design features as follows: flow straightening grid and screens, NGV, boundary layer bleed plates, optimized test item exhaust duct, exhaust duct ring step etc.
Various flow measuring and visualizing devices were integrated into the rig to study aerodynamics of vanes, secondary flows in vane passages and total pressure and momentum losses through the rows:
a) measurement of flow velocity and turbulence profiles using LDA (Laser Doppler Anemometry);
b) measurement of static pressure (up to 300 points) throughout the vane passages using AIR-20 pressure transmitters;
c) visualization of smoke particles movement paths in vane passages using high-speed video camera;
d) visualization of vane and end walls streamlines using oil film;
e) measurement of total and static pressure profiles and test item gas exit angles using automated system combining industrial robot, laser tracker and 5-hole probe.
The rig provides wide-range test capabilities focused on assessment of various methods of vane and vane passage 3D geometry optimization, such as:
a) non-axisymmetric vane passage end walls;
b) trailing adge local divergence (flow exit angle increase);
c) flow path mean line optimization;
d) vane lean angle change;
e) using bowed vane configuration;
f) arc-shape trailing edges etc.
Besides the above mentioned, there are also capabilities to study present-day aerodynamics issues, particularly: effects of roughness of the thermal barrier coating, Reynolds numbers, effects of vane nozzles cooling flow and geometry on NGV losses. Design and implementation of test rig dedicated to turbine vanes study featuring various design measures to optimize turbine vane passages flow aerodynamics is completed.

Keywords: test rig, design, test, aerodynamics, turbine vane, non-axisymmetric end walls, secondary flows

Purpose
The purpose of the paper is to determine possible expansion of applicability bicalibre missiles owing to usage of an integral ramjet as a motor.
Methodology
The research was conducted by means of multi-parametric optimization using numerical simulation methods. The research tool is based on bundled software describing conjugate processes of an aerial vehicle movement in the air and ramjet functioning. The structure, composition and level of the mathematical models used to construct the research tool are determined by its orientation towards the conceptual design stage, which requires solving optimization problems of structural-parametric synthesis.
Findings
The article shows topicality of the issues related to flight range increase while keeping the dimensional and weight characteristics unchanged. It is extremely difficult to solve this problem within the framework of using solid-propellant motors, which are classical for this type of aerial vehicles, due to the fact that their today`s thrust characteristics are close to their limit values. In this connection it was suggested by the authors that a solid-propellant integral ramjet motor should be installed onboard the aerial vehicle. Three layout solutions are considered in the paper, and predictive estimates of the maximum range of flight are given for each of the solutions with respect to the selected prototype upon the results of multi-parametric optimization. Results of a series of computing experiments show that 60 — 80 % increment of the flight range is provided for the most preferable layout in case of a ballistic trajectory of flight, and 50 — 70 % increment if there is a gliding phase. The following values are optimal to obtain the maximum range of flight: 25/75 — propellant weight ratio between the ramjet and the solid-propellant motor, 0.321 — relative area of the air intake cross-section and 0.425 — relative throat of the secondary nozzle. As a result, an operation algorithm of a bicalibre missile with an integral ramjet motor has been obtained, which enables to expand the applicability of such a missile minimum by 80 % in comparison with a version equipped only with a solid-propellant motor.
Practical implications
Practical application field of the research results is associated with structural-parametric synthesis of advanced bicalibre missiles at the conceptual design stage. Both the obtained results the computing experiments and the research tool, which can be used for further works in this area, are of obvious practical importance.
Originality
The originality of the obtained results is that a way for expanding the applicability of bicalibre missiles thanks to incorporation of ramjet motors has been shown and that functional relationships between the ranges of flight of the aerial vehicles under consideration and the main design parameters of the motor assembly have been obtained

Keywords: operation algorithm, bicalibre missile, rational parameters, ramjet motor

One of the main issues while designing classical control and monitoring systems for gas turbo engines is the presence of wiring and multiple wiring harnesses used to couple the sensors to the central processing unit reducing overall system reliability (around 30% of system failures are caused by various connection defects) and scalability and increasing weight and cost. The objective of this work is to present an application of energy efficient wireless sensor networks (WSN) in engine control systems replacing classical wired connections between sensors/actuators and engine controller. Implementing an engine WSN for coupling sensors/actuators to the engine controller enables to reduce the system weight, increase system reliability due to lower connections and, if needed, flexibly vary the number of system components without any need of redesign. It also provides capability to standardize modules of control and monitoring to use them with different types of engines. Methods and tools for the implementation of wireless data transmission and power supply organization are researched, justified and selected. A scheme of energy efficient wireless system node containing thermoelectric generator and impulse ultra-wideband transceiver is proposed. The direct amplification scheme making able to reduce transceiver design complexity resulting in power consumption and dimensions reduction is used. Calculations made allow to determine the maximum data transmission speed in the network requirements for the given transmission range, signal duration and probability of errors. The possibility of further optimization of the energy efficient transceiver scheme allowing to increase the transmission rate is shown. The node scheme allows to create energy-efficient wireless networks for data transfer in the gas turbo engine control system.

Keywords: wireless communication, ultra-wideband signals, sensor networks, sensors, energy harvesting, thermoelectric generators, gas turbo engines

The article is considered issues of mathematic simulation of flow performance of working medium in combustion chamber of gas-generator and duct of fuel flow controller under computational solution of equations of internal ballistics in nonstationary scenario and calculation in CFD. Examples of such calculation results are given.
Major units of CSRM (Combined Solid-propellant Rocket-ramjet Motor) are air intake, gas-generator of solid propellant, fuel flow controller, afterburner chamber with charge of booster and nozzle. Signal of predetermined consumption is transformed into required pivot angle of fuel controller drive. Drive pivoting changes area of nozzle controller openings, and as a result of it, causes variation of pressure and consumption of sustain engine. After gas-products of incomplete combustion of solid propellant of gas-generator charge got through fuel flow controller, they fell into afterburner chamber where burnt down in the air flow from intake.
During the design phase without reliable experimental data of combustion processes in engine passage under thermodynamic calculation we apply for equilibrium condition model of combustion products. Calculation of the equilibrium condition is numerically carried out by computer. The task optimization-minimization of Gibbs energy of combustion products is solved. Further equation system of internal ballistics is solved for combustion chamber of gas-generator. The solution is got by Runge-Kutta method of fourth order.
The results of preliminary ballistic calculation of gas-generator are used as initial data for flow simulation in fuel flow controller. Study object is unit of consumption regulator of gas-generator products Flow control of gas-generator products is carried out by variation of area of internal cross section in central opening of nozzle insert by means of moving center body along axis of controller.
Controller study was conducted at the limit mode, with almost completely closed central hole. This mode is characterized by a large pressure drop (gas-generator chamber 12 MPa and afterburner chamber 1,4 MPa) and maximum flow of gas-generator products.Computation is done in program complex of СFD in stationary scenario.
Calculated case is characterized by overexpansion of flow in expanding part of nozzles, in consequence of which, shock wave is formed and flow rate transforms from supersonic to subsonic. Movement in passage is uniform, some small eddies are present, major consumption falls on two constantly opened nozzle orifices.

Keywords: CSRM (Combined Solid-propellant Rocket-ramjet Motor), SFDR (Solid Fuel Ducted Rocket Ramjet), gas generator, fuel flow controller, computer simulation, propulsion system, rocket ramjet, ducted rocket engine, solid propellant, mathematic simulation, experimental method

The presented research lies in the field of automation of production planning during the product manufacturability control. The paper describes the general functional structure of the developed system, which is required to carry out the manufacturability control. This control is based on the manufacturability recommendations, which are structured within a knowledge base. The description of the production environment objects and their interrelations is implemented within the database module.
The automated system of product structure manufacturability (PSM) analysis acts as a guideline, which provides feedback from the manufacturability modules to the structure design modules. This system allows the designer to detect the non-manufacturable combinations of structural shapes in the detail and change its design according to the manufacturability recommendations in an interactive fashion. It also allows the production designer to present the recommendations, which are aimed at ensuring the PSM, in a formalized way.
The product manufacturability testing is a complex task. During its solution the designer must ensure both the high technical level and required performance of the created product. He should also fully take into account the manufacturing requirements (i.e. ensure the product manufacturability).
In practice the processes of ensuring the product structure manufacturability (PSM) can be conducted by using geometric modeling systems. The usage of these systems is closely linked to modern information technologies, which are used for the integration of the processes that take place during the entire life cycle of the product and its components. Therefore it is obvious that the provision of PSM, which is one of the tasks of production planning, should also be considered in the context of the CALS technologies application.
Thus, the application of «PSM analysis system» at the conceptual design stage can help to achieve the creation of a competitive product with high target values of the product structure manufacturability and maintainability in comparison to the alternatives. The project is based on the idea of formalizing the knowledge of the production designer and building a decision support system on its basis. Such system should help to reduce the number of errors in product manufacturability control during production startup. It should also allow the enterprise to reduce the manufacturing costs. This reduction is attained through the analysis of several design alternatives and selection of the optimal one. The selected alternative should be optimal in terms of its structure elements composition with taking the given level of manufacturability and price into the account. The software configuration is very flexible and it can be used at various machinery production enterprises.

Keywords: product model, manufacturability control, manufacturing process, knowledge base, database

The paper is devoted to building of the control system for high-adaptive and autonomous mobile robots (CSAMR) based on the use of the new software-controlled computer device, so-called voxel computer (VC).
The functioning process of the autonomous mobile robot (MR) assumes continuous monitoring of the MR internal state and space conditions around the robot. The more often such monitoring is performed, the better machine models synthesized by CSAMR reflect real state of MR and surrounding conditions.
The high productivity of VC allows to monitor very large number of the MR sensors within shortest time, to build the current models of the surrounding conditions and the robot itself, to build the model of the «excited» areas of the robot`s «skin» very operatively, and to alert timely about the possible clashes of the robot with obstacles, etc.
The high performance of VC is achieved by using deep parallelization of the operations and procedures most frequently used in the processing of the surrounding scenes and images, namely:
— set-theoretic operations with the objects of surrounding scenes;
— geometric transformations;
— analysis of the objects as possible obstacles when robot moving;
— calculation of the square and volume of the objects;
— determination of the objects’ position in space.
The article provides the block diagram of CSAMR, describes its configuration, purpose and principle of the functioning of the CSAMR separate components. Also, it presents the results of the modeling of VC and CSAMR in order to develop basic technical decisions and to give evidence of their hardware and software feasibility on the example of the MR’s operational model.

Keywords: high-speed processing, voxel computer, parallelization, pixel, voxel

Spacecraft onboard control system (SOCS) is a complex multicomponent set of devices. It includes both hardware and software components.
Strict requirements applied to the reliability of such systems cause the necessity of creation and further development of quick methods for technical condition assessment during onboard control systems operation. SOCS operates under rough conditions of the outer space, so it is exposed to such disturbances as space radiation and wide range temperature changes.
The new approach to solving problem of SOCS reliability analysis based on combined application of fuzzy and neuro-fuzzy models is proposed in the paper.
The use of classical statistical approaches to the analysis of SOCS reliability at the operational phase appears to be inefficient. But, good results were obtained by using fuzzy logic to describe weakly formalizing relationships. Also, artificial neural networks could be applied to use data accumulated during operation for further reliability analysis.
The proposed hybrid model of reliability takes into account dispersion of characteristics of onboard control systems’ elements and influence of external factors. On the one hand it helps to compensate the lack of data for analysis during initial period of spacecraft operation and to perform more accurate analysis in the future using collected during operation data on the other hand.
The proposed approach is an alternative to the generally accepted at the present time methods of analysis of system reliability based mainly on statistical approach.
The technique was implemented as a PC software application and proved its effectiveness in SOCS reliability analysis during maintenance of existing spacecrafts.

Keywords: reliability, fuzzy logic, neural networks, spacecraft, control system

The justification of functional tasks and structure of the perspective problem-oriented control system for complicated dynamic objects was conducted. This line of research is aimed at improving efficiency of intended purpose of the objects under consideration (flying crafts of different classes and types). The main approach to achieve the desired objective was intellectualization of the problem-oriented control system of complex dynamic objects, as well as the method of system integration of dissimilar problem-oriented systems.
The conceptual model of problem-oriented control system was developed. In accordance to the ultimate purpose of development of functional capability of problem-oriented systems when developing the conceptual model, the issue of information-system security of the problem-oriented systems throughout their lifecycle was examined. This problem was solved by using intellectualization of problem-oriented control systems based on the use of the medium of language scheme radicals. In this medium it is possible to develop and expand the range of solvable regular tasks of ensuring information/system security of problem-oriented systems by training them to solve some emergency tasks. Using medium of scheme radicals all objects and relations in the problem area of a complicated system, along with all processes taking place in it can be definitely and visually presented. Systems and subsystems, components of all types of provision (software, technical and other), tasks under consideration can be also presented conveniently.
The results of this work are intended for solving applied tasks in cybernetics. Practical use of the results obtained showed their effectiveness at all stages of the life cycle of a complicated system.

Keywords: problem-oriented control system, structure, system integration, intelligent control, information-system security, medium scheme of radicals, subsystem, functional tasks

This research work was aimed at creation of a system automatically monitoring the event occurrence in the ground-based control system software for the «Souz-2» launch vehicles.
The errors detected in the programs during the checkout have been analyzed: the event time intervals are mainly specified incorrect; more rarely, the wrong commands are sent or the wrong command/signal addresses are specified. During the program execution, the main events are recorded into a log file: the event occurrence time and the comment to the event are reflected there. This work requires great attention focusing on the event time diagram which should be worked out, then the recorded time intervals should be compared to the calculated ones; it is also important to check whether all the required events have occurred and none of the unnecessary actions have been performed. The developed automatic control mechanism makes it possible to simplify this process.
The solution of the original problem was based on the double programming principle. As the result of applying the developed testing procedure, errors were detected in the tested programs at the first checkout stage. The log became more understandable due to the highlighting of the positive and negative control results, whereupon the protocol analysis time reduced considerably.

Keywords: testing, software unit, automatic checkout, event, background task

In [3] a plan is proposed to optimize nested-query processing time for a single-processor database.
In this paper, optimization plan to optimize nested-query processing time is developed for a multiprocessor database and the minimal processing time is determined by Theorem.
Theorem. In a multiprocessor database the minimum time of the nested-queries for ordered or unordered data tables is achieved by processing all elementary queries, forming a nested — query.
Minimum query processing time for the unordered data table of all the elementary queries in the order of elementary queries is specified by condition

is determined by the expression

Minimum query processing time for the unordered data table of all the elementary queries in the order of elementary queries is specified by condition

is determined by the expression

Where
— processing time of the i-th (i = 1 , ..., r) processor of j- elementary query,
probability of success (data correspond to a specified condition) of j- elementary query,
j = (i, 2r +1 — i, 2r + i, 4r +1 — i, 4r + i, 6r +1 — i, ... , k- i+1).
Comparative execution times nested-queries for different numbers of processors are obtained for the joint and uncooperative order processing ordered and disordered data tables with parameters corresponding arithmetic and geometric progressions.
Moreover in this paper we determine the minimum number of processors, where execution time of the nested -queries is minimal.

Keywords: database, multiprocessor computer, nested -queries, distribution of elementary queries, optimization

Test radar impulse developing questions have very high priority, as they are needed when hardware is calibrated, tried out and approval and preliminary tests of the digital radar impulse processing device of perspective radar. The fact that blocks of test data were generated out of real radar signals allows essentially expand range of use. A number of important parameters of developing radar can be estimated within ground-based measurements verifications without flight testing.
Interpolation of the original signal sample for getting test data is determined by difference in signal data format in ongoing and developing radar. Noise interpolation special features take place because of finite accuracy of the interpolation signal sample. It ensues from the results of the conducted research, that interpolation errors are multiplicative and correlated. Energy of interpolation errors is concentrated in narrow band with correlate passive noise spectrum that is regarded non-working and rejected at spectrum processing. Thus, influence of interpolation errors for target detecting can be neglected
Output data was verified spectrographically and revealed that using interpolated input signal arrays comes up to increasing distance and frequency resolution. Generated data can be successfully used for testing target detection algorithms in complex background conditions.

Keywords: surveillance radar, interpolation, digital signal processing, spectrography, resolution power

A method of finding the mechanical characteristics of a composite wing is considered. A preliminary computation of the flutter safety is carried out for several variants of layered composite structures. A Ritz method combined with the Ritz polynomial approximation is used for the computation of the flutter stability in subsonic flows. Two variants of the layered structure are considered and the flutter-safe variant is selected.
The mechanical properties are found for each of layers of laminated composite, and then the properties of the layered structure are computed taking into account the winding angles for the fiber reinforcement on the groundwork of the common coordinate frame. Using the found properties the flutter stability of the considered composite wing is estimated by Ritz method.
The flutter-safest variant of the wing structure is found out. The efficiency of the proposed method is based on the simplicity of the mathematical model, the quickness of data preprocessing as well as the computation speed. The preliminary computing realized for various composite structures allows one to benefit financially by identifying the possible weaknesses of structures without expensive real testing.
The proposed calculation method can be used for linearly elastic and perfectly bounded layered structures represented in thin plates, particularly wings, rudders and other bearing surfaces of the aircraft. Ritz polynomial method can be applied to estimate flutter safety of separate parts of the aircrafts as well as of all aircrafts in general. For the well-posed problems with accurate input data the calculation error is up to 1-5%.
The proposed method has been implemented as the «Flutter Analysis Program» software developed by author using the «Wolfram Mathematica 8» package. This software combines the high-precision computations and the simplicity of the input data. Another important feature of the software is the possibility of the calculations not only for composite materials, but also for orthotropic and isotropic ones. The main restrictions are the linear elasticity of the materials, the requirement of the perfect bounding of the layered structure of thin plates and applicability of the Kirchhoff hypothesis.

Keywords: flutter, Ritz polynomial calculation, composite, modeling of the aircraft

Article type: Research and practical
Objective: Experimental investigation on the study of protective properties of Microarc oxidation (MAO) coatings
Subject: Development of anti-erosion coatings for spacecraft power elements
Object of research: MAO coatings on aluminum alloys
For the operation of the spacecraft (SC) during the entire period of lifetime need that degradation of a number of parameters of the equipment and systems in the operation will not lead to a breach of its intended use. Stricter requirements on weight, available power, and increased resource demands and the reliability of spacecraft lead to compact its layout scheme. As a consequence, the erosive effect of stationary plasma thrusters (SPT), which are used as correction thrusters, on the elements of the spacecraft structure is increasing.
Erosive effect of plasma jets SPT consisting in ablation material of construction as a result of long-term gas ion bombardment leading to a reduction in the thickness and contamination of the external surfaces of spacecraft. The main characteristic of this type of exposure is the depth of erosion, in another words the thickness of the sputtered layer.
Another negative effect of plasma jets SPT is erosion of conductive layer from thermal control coating of spacecraft, which leads to the accumulation of static electricity on the surface and occurrence electrical breakdown of surface SC.
Thus, in order to improve the quality of manufacturing products of space technology the development and application of critical spacecraft surfaces coatings with improved resistance to erosive action of plasma SPT is required.
A preliminary analysis showed that substance which has a high resistance to the plasma jet of inert gas (argon, xenon) is aluminum oxide (Al₂O₃).
Assessment calculations show that the required thickness of the protective coating of Al₂O₃ for active existence SC 15 years should not be less than 30 microns.
In this paper we propose to protect the spacecraft design elements fall under the plasma SPD the thin (100 micron) aluminum foil coated with aluminum oxide thickness up to 30 microns, applied microarc oxidation (MAO), which is acceptable for weight characteristics. On the spacecraft design, made from aluminum alloys, the coating may be applied directly to the surfaces to be protected.
The methodology of the study. Coatings were deposited on the IAT -T installation, power supply which allows independent adjustment of the anodic and cathodic current components and simultaneously stabilize the average values of these currents, which greatly simplifies the process of MAO and leads to an improvement in the quality of the coatings.
Source has the following specifications: range of adjustable voltage — (0-800 V), the range of regulated currents — ( 0-120 ) A/dm² ; error stabilization of current to 5%.
As substrate samples used foil AD 160×130 mm size and a thickness of 100 microns.
Microarc oxidation of samples was carried out in slightly alkaline aqueous electrolytes of various compositions.
The coatings were formed at a ratio of I_k / I_a from 0.6 to 1.4, the current densities in the range of 10 to 40 A/dm², the treatment time was 10-60 min.
The findings of research. For testing of protective coatings on the impact factors of storage and operation the specimens with protective coating on aluminum foil AD GOST 4784 74 microarc surface oxidation were prepared. The sample size is 100×100 mm.
The following tests were carried out:
— Cyclic bending at a diameter of 20 mm;
— Accelerated environmental testing;
— Radiative exposure;
— Thermal cycling.
Studies samples obtained showed that after exposure to plasma samples from the aluminum foil with surface microarc oxidation appearance remained unchanged. Maximum weight and thickness changed in samples polyamide film. MAO samples lost weight greater than the samples of aluminum foil; it is possible to explain some hygroscopic MAO coatings. Coating thickness remained virtually unchanged. Studies have shown that coatings obtained by MAO may be used as protective coatings for spacecraft.

Keywords: spacecraft, antenna reflectors, radio reflection coatings, antierosion protection, microarc oxidation

The honeycomb structure forming’s technologies often result the displacements of the honeycomb elements. The deviation of the structure’s form from the structural norm reduce the quality of composite structures and result the rejection of the defective structural elements. The deviation of the form may be detected wore often after the sandwich structure’s disassembling and cannot be eliminated.
The solution of the mentioned problem requires the analysis of the main causes of the displacement of honeycomb cores, the analysis of existing models allowing the theoretical investigation as well as the practical methods of elimination of honeycomb cores’ deviations, and the optimal design solutions increasing the technological efficiency of honeycomb structures of high structural depth and bias angle 45°.
The deviation of honeycomb’s forms during the forming are caused by the following reasons:
— the insufficient stiffness of the block subjected to the horizontal static loads being result of the molding pressure;
— the insufficient adhesion between honeycombs and covering layers.
The preliminary analysis shows that the main cause of the displacement of honeycombs is the horizontal loading during the forming. Therefore, taking into account the various methods of the honeycomb displacement elimination, the specimens for the experimental study were constructed.
One of the main methods of elimination of honeycombs’ deviation is the arising of the adhesion conditional factor on the groundwork of the technology of the prepreg fixation at the matching attachment by use of supplementary materials such as abrasive grids or perforated steel foil.
To estimate both the longitudinal and lateral stiffness of honeycomb cores the tests are implemented using the special test attachment. It must be noted that the longitudinal and lateral forces of bearing are three times different values. The computation based on the results of the mentioned tests has shown that the re-orientation of the honeycomb core in the skew area allows one the efficient resistance to the forces resulting the structure deviation.
Thus, the design solutions and the technology of elimination of honeycomb core structures during the forming of polymer sandwich composites are proposed and their efficiency is shown.

Keywords: Polymer composite materials, honeycomb structures, stiffness, contact retention force, horizontal component of pressure molding vector

The improvement of the manufacturing process of complex products using PDM- systems for aviation enterprises determines the relevance of the study. However, the results of studies in this field require individual adaptation to the particular enterprise. This study developed the technique of introduction of electronic data interchange, including the algorithm for paperless issuing of engineering and design documents (EDD) of the aviation enterprises, the model of а placement information structure and the typical scenarios in the Teamcenter medium. The study showed that the competent organization of a unified information space can significantly reduce the time and cost of introduction the PDM- system that makes it possible to reduce the labor intensiveness and number of errors in the preparation of EDD.
The purpose of this study is to increase the issuing of engineering and design documents efficiency through the introduction of PDM-system.
To construct an algorithm for the paperless issuing of engineering and design documents (EDD) at JSC «PO «Strela» were applied provisions GOST 19.701 — 90 «Schemes of algorithms, programs, data and systems». To reflect information interrelations of paperless issuing of engineering and design documents developed UML- sequence diagram.
Scientific novelty of the study is theoretical justification and methodological basis of the information descriptions aircraft engineering products.
In the technique clearly defined stages of inspection, verification, approval and coordination of engineering and design documents.
In developing technology paperless issuing of engineering and design documents in the system Teamcenter Engineering applied methodology for functional modelling IDEF0.
Placing information in the PDM-system is implemented as a hierarchical model structure in the medium Teamcenter Engineering.
The main conclusions and findings.
1. Organization of a unified information space will allow JSC «PO „Strela“ significantly reduce the number of errors in the preparation of engineering and design documents.
2. The proposed algorithm of design documentation of the aircraft enterprise makes it possible to carry out both the control and coordination of the technical electronic documents.
3. The developed technology of paperless issuing of engineering and design documents in Teamcenter Engineering is clearly presents and explains the choice of templates and procedures for creating and start the process of issuing of the technical electronic documents.
4. The hierarchical model of the structure of placing information in Teamcenter Engineering environment allows providing uniformity and convenient access users PDM-system.
5. On the basis of the algorithm issue of engineering and design documents, UML-diagrams, mathematical description of posting information was developed practical technique automation of design processes of „PO“ Strela», including the stages of development of policy documents, the algorithm of the management details, the algorithm of the management assemblies and the algorithm of the data management geometric materials.
6. The developed method allows «PO» Strela" to reduce terms setting of products in the manufacturing, to ensure that the design process in parallel by multiple divisions or implementing, to structure product information, to avoid duplication of documents, to facilitate interaction with partner companies.

Keywords: PDM-system, a unified information space, electronic data interchange, engineering and design documents, life cycle of products

In order to maintain competitiveness on the world market aircraft manufacturers need to plan and implement development projects. Many of the Russian aerospace industry companies are currently experiencing shortage of resources, so they (resources) should be spent as efficiently as possible. The vital need for performance management techniques in aerospace industry is based on the fact that planning and investment are complicated by the long production cycle, high resource consumption and relatively low profitability of production.
One way to solve the problem of efficient resource management is development of the enterprise strategy and the creation of a system for strategic portfolio management. An important part of this system is the Strategic Project Office (SPO) which is special software that supports the Executive Review Board.
To implement the SPO it is necessary to construct methodology for formation of strategic portfolio and evaluation of strategic performance. We have developed such methodology. The steps of persons involved in the management of strategic portfolio are specified; the information flows that define sequence of the steps are determined. The proposed approach is illustrated by the activity flow diagram to clearly describe the process of strategic portfolio management.
Algorithms for building a strategic portfolio were proposed by scientists before (there are special standards for project portfolio management, e.g. PMI — The Standard of Portfolio Management). However, most of these algorithms only verbally describe the activities which should be performed. Our methodology differs from them, because it relies on usage of specially developed mathematical models and methods, so it is well suited for incorporation into the SPO.
Currently, we have desktop applications that support evaluation of strategic performance. Also we have provided the templates for web-interface of SPO and analytical reports.
The proposed methodology has been used to analyze the strategic performance of an enterprise which produces light aircrafts and has shown its practical usefulness.

Keywords: strategic project office, portfolio of strategic initiatives, optimizing the allocation of resources