2013. № 71

The paper proposes approaches to solving the problems of one of the priorities of the strategic development of the Ulyanovsk State University (ULSU) «Aviation technology and aviation mobility» in order to implement the model of advanced training of specialists that is aimed at the satisfaction of needs of the CJSC «Aviastar-SP» and other companies realizing their activities in the aviation field.
First of all, it is an adaptation of training plans for the following branches of science: «Aircraft engineering», «System analysis and control», and «Automation of technological processes and manufactures» to the modern requirements of dynamically changing labor market in the aviation industry.
To do this, the ULSU department of mathematical modeling of technical systems, as an experiment, in the formation of professional competence in the above areas of training has offered a step-by-step methodology, which is based on the object competence of students. Within such methodology, the competence matrix was developed for the basic subjects, general-technical ones as a rule. Professional requirements were formed by expert assessments on the appropriate subject among graduates working at aviation enterprises and plant specialists being the department heads.
Competencies identified by the subjects were structured and transformed subsequently to be within the framework of the existing educational standard. As a result of that work additional educational paths have been offered for the implementation of purpose-oriented training, as well as the options for adjusting curricula.
Based on the analysis of competencies demanded by the CJSC «Aviastar-SP», the subjects were defined, training in which should be carried out with the involvement of resources of the leading universities.
Second, it is an improvement of education quality through the introduction into the training process of the modern methodological and laboratory support with the aviation focus and of the point rating system of monitoring and appraisal of students.
Important are the probations of lecturers and staff at the leading Russian aviation universities to study modern methods of teaching, advanced training in the field of methodological support, insight into to the modern researches in the field of aviation-related mathematical modeling, study of new technologies of production of structural aircraft materials, insight into the laboratory facilities and methods of their application, etc.
The paper discusses the participation of students in research and action-oriented work that is a necessary condition for the high-grade digestion of educational program. To regulate and systematize such work, a Student Scientific-and-Technological Bureau was established at the ULSU, in which the interested students study modern digital technologies, gain additional competences in the use of software and the solution of specific problems.
We also consider the experience on enlisting the services of the leading lecturers of aviation universities and specialist of aviation companies to assess the quality of graduate training.
One of the results of such work was the establishment of the Competence Center «Aviation technologies and aviation mobility» that allowed bringing together efforts of experts from aviation enterprises and leading scientists from the ULSU departments and institutes. We plan to create a joint center of competence under the CJSC «Aviastar-SP» and ULSU to solve problems within the framework of the federal purpose-oriented programs.
Thus, the proposed and implemented methods for solving the problem considered in this paper include the concept of a full cycle of training, ranging from the occupational guidance for school children — future students at the stage of their enrollment in a higher education establishment, and to the graduates (bachelors, masters) with the knowledge and skills satisfying future demands of the aviation industry.

Keywords: academic mobility, professional competence, advanced training, training, aviation orientation

The training of students in professional work for the industry is an important task for the regions with developed aircraft industry. Contracts of domestic enterprises for the aircraft production are the long-term ones. The process of human resources formation for the modern airlines begins with the work with students.
The theoretical basis of vocational guidance and training was laid by the modern representatives of psychological-pedagogical school. Typology of professional occupations, vocational guidance, self-determination in the profession, professional choices, educational technology, and professional diagnostics are meaningful for the elaboration of methodology in psychology and pedagogy.
Implementation of professionally designed programs for students is based on a large number of practical approaches. Domestic universities practice meetings with students (study, information, case, etc.), activities to engage young people in the world of professions, their individual training. The experience of foreign universities is interesting due to the career-oriented programs for successful adaptation to the labor market.
The Ulyanovsk State University has set a goal to develop a modern model of primary training school for the needs of the regional aviation industry. The main approaches were identified as follows: the formation of network interaction between the university and general-education institutions, industrial enterprises, government bodies of different levels, and aircraft producers on the basis of contractual relations; organization of continuing centers for occupational guidance of schoolchildren in aircraft construction; systematic carrying-out of competitions for the selection and training of the most capable school graduates in the aircraft manufacturing; formation of a multi-channel mechanism of information support of programs for school children.
Testing of the model allowed us to determine and detail the technology of successful students training. First, individual components of the career guidance programs are to be coordinated on a distributed basis. Thus, the solution of the individual problems of the career guidance centers highlights the work of the organizers of appropriate activities. Second, vocational guidance centers should be prepared to work in the offline mode. This means the possibility of duplication of certain functions assigned to the company or university, if the representatives of the latter are not able to attend classes with students. Third, the competitive exams for students should end up with different forms of stimulations for the most successful participants.
These technologies have passed the checks. Their joint use has significant socio-economic results — formation of a «core» of promising students — future students in aircraft engineering disciplines, proliferation of potential applicants interested in obtaining promising specialty in aircraft engineering, increase in the total number of students of secondary and high schools involved in the research, creative and practical activities.
Obtaining of such results promises intensive development of the university, municipalities, the region as a whole.

Keywords: professional orientation system, primary training, organization of the process technology

The problem of heat conductivity for an arbitrary thin-walled shell exposed to a plane-parallel radiant heat flux from the infinitely remote source of radiation is considered. A convective heat transfer between the shell and the environment is described by the Newton law. Thermophysical parameters of the shell depend on the temperature.
Nonlinear heat conduction problem is described by the following differential equation of heat conductivity:

with the initial condition

and the boundary one:

Here the following notations are used:

is the function of temperature distribution along the shell; is the ambient temperature; and are the curvilinear coordinates on the shell midsurface; n is the normal unit to the boundary of the shell; is the time; is the angle of incidence of the radiant heat flow on the surface of the shell; is the shell thickness; are the function radiant heat flux and its maximum value; is the coefficient of heat transfer from the shell into the environment; — are the heat capacity, the mass density, the thermal conductivity, and the thermal diffusivity coefficients for the shell material; are the coefficients at the initial temperature; A, B are the coefficients of the first quadratic form of the shell midsurface; R is the least radius of curvature of the shell midsurface.
At the initial moment the temperature of the shell is equal to zero. Excluding the process of thermal conductivity in the shell midsurface various asymptotic solutions are constructed.
The formula for the temperature field in the shell in the case of a linear dependence where are constants is obtained in the following form:

The obtained asymptotic solutions provide the accuracy sufficient for practical use for various aircrafts structures.

Keywords: nonlintar heat conduction problem, aircrafts, radiant heat flux, convective heat transfer, asymptotic solutions

A temperature field in a reinforced by frames and stringers arbitrary-shaped shell exposed to a plane-parallel radiant heat flux from an infinitely remote heat source is investigated. A convective heat transfer exists between the shell and the environment. Geometric parameters of cross sections of reinforcing elements are small as compared with both the distances between them and the radius of curvature of the shell midsurface. With these assumptions, the normal component of radiant heat flux moves smoothly over the surface of the shell. Therefore, the gradients of temperature field will reach the extreme values in the vicinity of the reinforcing elements in the directions perpendicular to their axes.
The asymptotic dependence of the distribution of the temperature is found from the solution of the heat conduction problem: 

where is the distribution of the temperature with height ribs,  are the thickness and the height of section of ribs, are the coefficients of thermal diffusivity and thermal conductivity of material ribs, — coefficient of heat transfer from the reinforcing element in the environment, is the temperature distribution along the shell, is the shell thickness, are the coefficients of the thermal diffusivity and thermal conductivity of shell‘s material, is the coefficient of heat transfer from the shell into the environment, is the time dependence of the heat flux, is its maximum value, is the cosine of the angle of incidence of the heat flux on the shell surface, is a curvilinear coordinate extends from the lower base of the cross section of the ribs to the top on the axis of symmetry and next to the shell middle surface perpendicular to the edge.
Using Laplace transform the various asymptotic solutions are constructed. At small times the convective heat transfer between the shell and the environment has no significant effect to the temperature field, and for the practical calculation the following expressions can be used:

Keywords: shell, temperature field, supporting elements, convective heat transfer, radiant heat flux, the Laplace transform, asymptotic solutions

The plane transient problem of the linear elasticity theory for the thin layer of the aircraft paneling represented as an homogeneous isotropic layer of uniform thickness is studied. An arbitrary distribution of the surficial load is considered.
A rectangular Cartesian frame is used; axe is directed deep into an elastic layer, and one is directed along its free surface z=0.It is supposed that the load depends not on the coordinate y therefore a plane problem can e formulated. Using the superposition principle for the fundamental solutions we have the displacement vector represented as follows:
where is a matrix of fundamental solutions (or a transient function matrix).
For the transient functions we have the initial-boundary value problems for the layer under the concentrated loads distributed according the law where is the delta function. To compute the transient functions analytically the Fourier integral transforms with respect to the coordinate x and the Laplace one with respect to the time variable are used. The images of the transient functions are represented as the exponential power series; each member is the sum of products of homogeneous rational functions of the degree (-1) in terms of the transformation parameters and square roots  and exponential functions in terms of the linear combination of these roots. Here  are parameters of Fourier and Laplace transformations respectively, and is the dimensionless parameter defined by the properties of the material. This approach allows one to use the joint method of Laplace and Fourier transforms’ inversion that’s based on the construction of images’ analytic representations. If the exponential function’s degree contains one root only the inverse transforms can be represented explicitly. If the degree of exponent has more than one root the modified algorithm for joint inversion of Fourier-Laplace images has to be used to obtain the inverse transforms.
To obtain the solutions the numero-analytical algorithm based on the method of quadrature and Simpson formulae is developed. Some examples of solutions are shown.

Keywords: thin skin of the aircraft, transient effects, elastic waves, integral transformation, the influence function of the elastic layer, the numerical-analytical algorithm

The paper considers the application of the thermodynamic functions method for the research of the Magnetic Suspension and Balance Systems (MSBS) for wind tunnel models with six degrees of freedom.
Electromagnetic suspension system includes subsystems of different physical nature — electrical and mechanical.
Interdependencies and interaction between mechanical and electrical variables in the electromechanical systems are defined by the equations of mechanics and electromagnetic field. The difficulty of establishing relationships between the variables on the basis of these equations is conditioned by the fact that the electromagnetic characteristics of the materials (magnetic permeability and dielectric permittivity, conductivity), which are included in the equations, depend on the energy density in the matter. It is also necessary to note that these dependencies are not amenable to theoretical calculation while the experimental characteristics can only be obtained in the simplest cases (which are quite far from the considered Magnetic Suspension and Balance System for wind tunnel). Complexity of the configuration of the electromagnetic system of a real Magnetic Suspension and Balance System for wind tunnel aircraft models also leads to serious mathematical complexities in obtaining the equation solutions.
In this paper, the relationships between mechanical and electrical variables are established on the basis of the general energy principles. This allowed to establish a number of new common relationships between the state variables and functions of the energy state of the electromechanical systems, and, in particular, for the suspension system for wind tunnel aircraft models.
The method of analysis of electromechanical systems, which is also called the «method of the energy state functions», is based on the energy conservation law. This law is expressed in the form of differential input functions, which describe the energy state of the electromechanical systems. The general relations between the changes of the state variables and the energy functions values are established based on the condition of existence of the total differentials of these energy functions.
The developed method was used to create a model of the electromagnetic suspension system for wind tunnel aircraft models. There are n electromagnetic forces in this electromagnetic suspension system, which are applied to the aircraft model with six variable spatial coordinates.
The relations, which were derived from the developed method, allow us to express some system characteristics in terms of the others. This result can be used during the design and study of the electromagnetic suspension system. The adduced research helps to deal with the fact that while the electromagnetic force depends on the current values of the spatial coordinates, which are available for measurement, the measurement of the magnetic flux is very difficult. This is especially true for the multicomponent electromagnetic suspension with several electromagnets, which are situated in the magnetic field with a complex configuration such as the one that exists in the suspension system for wind tunnel aircraft models.

Keywords: wind tunnel , magnetic suspension, wind tunnel aircraft model, method of thermodynamic functions

The paper proposes a method for simulating the two-dimensional motion of an arbitrary holonomic rigid-body system. This method allows to automate the process of constructing models of aircraft landing gear legs with arbitrary kinematic schemes. It is based on the solution of the Lagrange equations of the first kind. The method allows to model equality (bilateral) and inequality (unilateral) constraints, which restrict the motion of the rigid bodies system. The ideal holonomic constraints are imposed on the positions of the bodies. These constraints are described by the following equations

where x is the position vector of the bodies of the system ; for equality (bilateral) constraints ; for inequality (unilateral) constraints; d is the number of constraints , which are imposed on the system.
The equations of system motion are written in matrix form as follows:
  where F is the vector of active forces; a is the acceleration vector; is the vector of Lagrange undetermined multipliers; J is the Jacobian matrix of the constraint vector; M is the diagonal matrix of lumped masses and moments of inertia of the rigid bodies.
The software, which was developed on the basis of the proposed method, was used to simulate the compression dynamics of the landing gear during landing impact. The main landing gear leg of the Ka-62 helicopter was chosen for modeling (Fig. 1). The model consists of the following five rigid bodies: wheel (1), lever (2), shock strut piston (3), shock strut cylinder (4), and load, which is applied to one leg (5). 

Fig. 1 The model of the main landing gear leg of the Ka-62 helicopter
Swivels a, b, c, and d and sliding joints e and g restrict two degrees of freedom. These joints are defined by two equality (bilateral) constraints. The cylinder rod stop f, which restricts one degree of freedom, is modeled by one inequality (unilateral) constraint. Thus the model includes a total of 13 constraints. The model also includes three forces, which act on the system: the tire-compression force P_T, the axial force in the strut shock absorber P_sh.a. and the lifting force P_L.
Figure 2 shows the calculated dependence of the force, which acts on the piston of the shock strut absorber, on the compression of the shock strut absorber as well as the diagram of the polytropic compression of gas in the first chamber of the shock strut absorber.

Fig. 2 Diagram of the axial force, which acts on the shock strut absorber, and polytropic compression of gas depending on piston stroke

Keywords: helicopter landing gear, oleo-pneumatic shock strut absorber, shock-absorption model, method of Lagrange undetermined multipliersModeling the compression dynamics of main a landing gear strut of a helicopter

Modern fan design for aviation big size gas turbine engines is supposed to use hollow wide-cord rotor blades. Hollow fan blade developing requires complex problem solution in the fields of aerodynamics, design, technology, vibration damping, fatigue resistance and operation damage durability. The article is devoted to the problem of turbo engine hollow fan blade resistance to foreign object damage.
The investigated hollow fan blade is a sheet welded structure with gofered filler. The blade airfoil is shaped and assembled with the method of super-plastic molding and pressure welding combination.
It is proposed the technique of numerical simulation of ballistic damage by stone of hollow fan blade. The technique is based on joint using of finite elements and smoothed particle method. Johnson & Cook model was used to describe high speed plastic behavior of blade titanium alloy under impact by stone.
There were numerically simulated several cases of ballistic damage of blade leading edge and airfoil shell by different size stone moving with speed 266-515 m/s. Damage geometric parameters, residual stress and stain fields were found as result of simulation.
Experimental modeling was executed to verify the calculation technique and results. Specimens of blade edge and airfoil were subjected to impact by granitic ball accelerated in gas gun up to velocity 400 m/s. To estimate residual strains at damage zone special net was wrote on the specimen. It was found good agreement of experimental and calculated data (damage size and residual strain).
The developed numerical simulation technique may be used for comparative analysis measures of blade foreign objects damage resistance improving.

Keywords: fan blade, foreign object damage, finite elements method, smooched particles method

In recent years, required Time Between Overhaul (TBO) of centrifugal pumps for TPS, NPP, marine and other facilities is up to 40,000 hours. One of the main factors limiting the growth of the TBO is cavitation erosion of impeller blades. NPSH is usually chosen with respect to the cavitation flow break operation mode. However, this NPSH choice not guaranteed to work without cavitation erosion due to unsteady flow behavior at the impeller inlet. In this regard, it suggests ways to calculate NPSH of the cavitation inception. On the basis of the published experimental data the inception cavitation coefficient empirical formula is derived. Furthermore three-dimensional unsteady turbulent flow in the impeller validated with experimental data published shows cavitation inception operation modes on different flow rates including back flow phenomena range. The described method of computer modeling can be recommended to optimize the pumps against cavitation erosion. Cavitation flow break NPSH value calculated from the equations of flow, energy and momentum for the plane cascade of plates of finite thickness. The data obtained are generalized to the circular cascade of centrifugal impeller. For optimization by flow break NPSH the Rudnev analytical formula and Shemel & Shapiros empirical formula are recommended. All this data give one the opportunity to reasonably choose the centrifugal pump NPSH required.

Keywords: сentrifugal pump, cavitation erosion, NPSH, cavitation inception, cavitational flow break

The research subject of the present paper is a single-stage axial turbine of the turbocharger TK-32. Turbocharger is manufactured by LLC «Penzadieselmash» (Penza, Russian) and used as unit supercharge for locomotive diesel. The aim of this work is to ensure the turbine work capacity when rotor speed is increased by 10% without significant reduction of safety margin and coefficient of efficiency.
The workflow gas-dynamic analysis of turbocharger TK-32 turbine stage and the stress strain state calculation of its rotor blade in base variant at nominal mode (n=25500 rpm) and at forced mode (n=28000 rpm) were carried out in the given work. The conducted strain-stress state analysis at forced mode indicated the region of high stresses on turbine rotor blade at the level of 2/3 from root. The stresses are exceeding permissible. In addition, it was revealed the presence of plastic deformation in the attachment of the disk and blades.
The variant of the turbine modernization was proposed based on the performed strength and gas-dynamic computational studies. The peripheral rotor blade section tangential displacements were implemented to create a modernized turbine design. It is allowed to reduce the level of stresses almost by 20%. The derived value of safety margin of modernized turbine at the forced mode is not below the value of safety margin for base turbine version at nominal conditions. The flow in the turbine modernized in accordance with the received recommendations was investigated using Ansys CFX. The calculation showed that the coefficient efficiency of an improved turbine is increased by more than 1% compared with the original value.
Experimental results showed an increase in turbine efficiency which fully confirms the conclusions drawn by the authors.

Keywords: turbocharger, stress-strain behavior, assurance factor, the gas-dynamic analysis

Liquid propellant rocket thrusters (LPRT) are the executive bodies of spacecraft reactive control system. Modern liquid propellant rocket thrusters have a relatively small value of required thrusts 0.01 — 1600 N, which is explained by small level perturbations acting on the spacecraft in free flight and operating at pressures in the chamber Pa, which corresponds to the minimum total weight of the propulsion system, and advantageously in a pulsed mode with a huge number of short inclusions (10⁵ — 10⁶). Power supply LPRT is carried out mainly from the pressure feed system (PFS) the fuel.
To control spacecraft propulsion by pair of force on three axes is necessary 12 engines, however, actually it can be from 8 to 40 intravehicular.
To assess the design perfection liquid rocket propulsion system energy-mass characteristics are used. They are important in choosing optimum parameters of spacecraft, propulsion systems and their separate aggregates.
One of the possible ways to improve the mass-energy characteristics of LPRT and total propulsion system is to increase the pressure in the combustion chamber, whereby there are the decreases in weight and size, the increase in specific impulse and an opportunity increase the degree of expansion of liquid propellant rocket thrusters nozzle.
Article is devoted justification necessity of the account the dimensions and mass characteristics liquid propellant rocket thrusters for reactive control system in the design phase, developing a mathematical model dimensions and mass characteristics of liquid propellant rocket thrusters operating on propellants nitrogen tetroxide and unsymmetrical dimethyl hydrazine with thrust 200-1500 H and operating pressure 1-4 MPa.
The article provides a comparison of the computational model with data from a real engine, developed at the Department of MAI 202 «Rocket engines» laboratory «Liquid propellant rocket thrusters ». In the article satisfactory convergence of the comparison results is given. It is planned adjustments model and its further application for a different kind of fuel with a wider range thrusts and pressures.

Keywords: liquid propellant rocket thrusters, the mathematical model, dimensions and mass characteristics, increased pressure in the combustion chamber

Ion thruster relates to one of the electric propulsion types, in which the plasma-forming gas is accelerated in the form of a beam of positively charged ions, subsequently neutralized by electrons at the thruster exit. One of the basic units of the ion thruster is the grid system designed to extract ions from the gas discharge plasma and to accelerate them up to the required speed. The slow secondary charge-exchange ions originate during the primary beam motion in the interelectrode gap and in the neutralization zone; they bombard the accelerating electrode, causing its erosion, which limits the thruster lifetime.
The article deals with the modeling of the formation of the primary ion beam and charge-exchange ion flows in a three-electrode grid system. The purpose was to obtain a detailed pattern of the primary ion trajectories for the selection of optimal configuration of the grid system providing necessary parameters of the beam efflux, as well as of the trajectories of the secondary ions for further modeling of the electrode erosion.
Modeling of trajectories of the primary beam ions was made by the software package IGUN. In addition, a program block was developed to simulate the trajectories of the secondary ions, using which an area was defined within the volume of the primary ion beam, from which the originating charge-exchange ions hited the accelerating electrode of the grid system. The results obtained can be used to calculate erosion of the accelerating electrode of the thruster grid system.

Keywords: radio frequency ion thruster, ion-optical system, erosion, ion flux recharge, neutralization zone, space charge

The objects of the study are structural systems and structural elements of gas turbine engines recuperators.
Recuperator is gas-air heat exchanger where compressed air in the compressor is heated by exhaust gases, which increases the thermodynamic parameters of the gas turbine engine cycle as a heat engine.
The reliability and service life of the regenerative gas turbine engine is largely dependent on the reliability of the recuperator, which in turn is defined by its structure.
Currently improving the design recuperators of GTE is in several ways:

• For stationary gas turbines, where the weight and dimensions ofthe heat exchanger are not critical, used remote tubular heat exchangers. Design ofsuch heat exchangers isestablished and reliable.
• Toreduce the weight and dimensions ofheat exchangers ofaviation and transport gas turbine units inthe design ofrecuperators the heat exchange intensifiers ofvarious designs and high-performance heat transfer surfaces ofdifferent types are used
• The improvement ofmethods ofcomplex thermal-hydraulic calculation ofheat exchangers.

Insofar as there are many classifications of structures recuperators GTE, there is a need to organize them somehow, i.e. separate them using the most important and fundamental design features. The main aim is to familiarize readers with the most effective design techniques to improve the efficiency and reliability of gas turbine recuperators.

Keywords: recuperator, air-gas heat exchanger, efficiency and reliability of GTE

Energy balancing of LRE is originally produced during developing of the new rocket engine, determining its characteristics, choice of pneumohydraulic scheme. At the initial stage of development, during the balancing of energy parameters previously established experience is taken into account. Based on this experience, the new developed engine has attainable values of pumps and turbines efficiency, the needed pressure for reliable cooling of the combustion chamber, the fuel spray nozzle components, regulatory units, etc.
The procedure of energy balancing of parameters of LRE is performed from the beginning of the new engine to the ideas of his moral death and removal from the production. It is necessary because many of the measured or calculated parameters must be constantly refined. Determination of the thermodynamic characteristics of the products of components combustion, especially in gas generator is main task for energy balancing. It is because the gas flow burn in the gas generator at mixture ratio of fuel components significantly different from the stoichiometric, and taking into account the real properties of turbines gas. It can significantly change the energy balance of the engine and its influence on the choice of pneumohydraulic scheme.
The article presents a comparison of operating temperatures in oxidizing gas generators of engines 11D58M SG, RD-120 and RD-191 and calculated temperatures by programs CEA, RPA and Astra.4/pc
The calculations were performed taking into account the temperature of oxygen and hydrocarbon fuel entering in the gas generator. Comparison shows that, in actual temperature oxidizing gas of generator differs significantly from the calculated values. Comparison shows that, in actual conditions the temperature oxidizing gas of generator differs significantly from the calculated values. This is due to at least two factors, firstly the fact that the kinetics of the combustion of fuels at large or small level of mix ration is not well understood. This is confirmed by the results of calculations. Secondly, the organizations of mixing inside gas generator providing a substantial influence on the final temperature of the gas at the outlet of the gas generator at the large levels of mix ration.
Conclusions: during energy balancing of LRE non- ideal of chemical reactions (for calculation of the thermodynamic parameters of the combustion products) and difference between the properties of a real gas from ideal (for determining the power of turbines) must be considered.

Keywords: liquid rocket engines, energy balance, thermodynamic performance, non-ideal chemical reaction

Here we examine the problems related to update the technological process of die tooling manufacturing for section extruding in the process of production of gas turbine engine compressor blades.
As far as specialists [1, 2] are concerned the factory labour hours of producing the gas turbine engine compressor blades are reaching nearly 30% of factory labour hours of producing the whole engine.
At the present moment different manufacturing technologies are applied in producing the gas turbine engine compressor blades, which are based on three methods of blade airfoil shaping. They are rolling, milling and electro- erosion machining.
All these technologies have their advantages and disadvantages [3].
In this project the authors examine the airfoil shaping blade manufacturing using cold rolling (for blades, which were made of titanium alloy and heat resistant steal). The main attention is drawn to raw material manufacturing by die tooling.
First of all some milling processes regarding semimatrix were changed from universal equipment to CNC.
To increase accuracy of blade locking piece manufacturing the semimatrixes there were held the experimental works on tolerance value of main semimatrix material and hard alloy insert.
For reaching definite manufacturing of fillet places and hard alloy insert transition to the main semimatrix material the electro-erosion machining was carried out in several operating steps.
The number of processing datum surfaces was reduced by using of the special equipment. New processing datum surfaces have become common for electrode CNC milling and for burning pieces on the spark-erosion equipment.
Some parts of inspection operation were replaced by digit instrumentation.
The authors have carried out in JSC «Moscow Machine Building Enterprise named after V.V.Chernyshev» procedures to meet the improvement of wear resistance of die tooling through implementation constructional and technological procedures without using the special processing work surfaces, thereby reducing the complexity of its manufacturing.
The suggested procedures may be used both for the new tooling and modernization of the existing one.

Keywords: section extruding, electro-erosion machining, profile extrusion, CNC milling

The structural scheme of the spacecraft (SC) with the multi-purpose space large scale laser system (MPSBLSLS) onboard placed in an unpressurized compartment of a non-refundable SC (target module) with 180 days lifetime on earth orbit was developed. The estimation of weight and size characteristics of SC own systems was carried out. Weight and size characteristics are defined and implemented in the layout of the compartments of the target module spacecraft components MPSBLSLS such as systems of periodical-impulse generation of radiation, storage and supply of fuel components of the laser forming optical systems and laser ranging.
The general assembling scheme of the SC is shown. It is shown that SC with the MPSBLSLS (based on the cw chemical HF(DF) laser) onboard with weight of ~19700 kg can be locate under a “standard” aerodynamic panel of the carrier rocket “Proton-M” that allows to put it into the orbit of artificial satellite of the earth. Thus the total duration of MPSBLSLS will 30 minutes at necessary time of impact for space debris ~ 1 s pulse-periodic mode of radiation on HF molecules with pulse energy of 0,8 J and 180 minutes in a radiation mode on DF molecules with a pulse energy on the right vibrational- rotational transitions ~ 5 µJ. If it is necessary MPSBLSLS can be refueled by supply spacecraft with reserve component of the laser fuel. In this case the total duration of the MKLU mode radiation will increase about 2.5 times.

Keywords: space-based laser system, cw chemical HF laser, space debris, monitoring of the ground atmosph, launch vehicle, layout

One of the most important indicators of development level of rocket-and-space technology along with the energy and accuracy characteristics is its accident-free operationing. The solution to this problem affects the total set of program measures to ensure reliability, algorithms development and operational documentation. However, many issues closely connect the work and characteristics of product units with work of flight control system.
The task of ensuring trouble-free flights is solved at different stages of development and use of launch vehicles, boosters and spacecraft.
This starts with a design stage of project works in the form of items conservation and creation of redundancy in manageability, i.e. stocks manageability and reserves of design elements strength, and installation of required onboard computer informational-measuring system
The article presents the ways to decrease the probability of accidental outcome flight products of rocket and space technology and considers an assessment of fall points dispersion of exhaust stages.
A significant decrease in the size of ellipse dispersion for the 2nd steps associated with the presence of compartment intermediate structure, which has a shape of truncated cone, which in some ways fulfils a role of brake pads during the descent. Thus, there is so-called effect of side center adjustment influence on the decrease in size of falling fields. This phenomenon is about fact, that upon entering into the dense layers of atmosphere due to aerodynamic forces, which are directly proportional to velocity head, arising aerodynamic moments spinning away a stage with angular speed with inversely proportional to their axial moments of inertia. Due to the spin the aerodynamic forces are averaged and, as consequence, the area of ellipse dispersion is reduced. To reduce the dispersion of exhaust stages points of fall we have considered the impact of stabilizing constructive devices in the form of the plume, as well as the additional controlled descent algorithm for the onboard control system of launch vehicles. According to the results of statistical modeling the table for maximum values of exhaust stages characteristic parameters in descent trajectories was obtained.

Keywords: environmental damage of rocket, space system, off-nominal situation, exploitation of space rocket technology, probability of emergency outcome, accident-free operation

Aerothermodynamics investigations of perspective atmosphere descent vehicles (DV) are the main point of interest of Russian and foreign scientists, who have deal with design of modern space technologies. In these descent vehicles the aerodynamic and thermal protective shields made to the shape of aeroelastic constructions in part or in whole, and in particular made to the shape of hermetic gas pressurized shells are used for effectiveness deceleration.
The one from the main projects realizations problems of such DV with ballute is creation of thermal protective system for ballute shells forming frontal aerodynamic shield (FAS). In comparison with thermal shields of hard structure the peculiarity and main advantages of FAS are the possibility to pack ballute in well-knit capacity size during the transportation of DV with ballute under the launch vehicle fairing and on the board of spacecraft.
It is reasonable to use complex mathematical description of all processes going with vehicle atmospherically movement at all regimes of gas flow for operating aerothermodynamics calculation of DV with ballute during the reentry and for solution to main target — determination of structural characteristics of flexible thermal protective cover of frontal area.
This paper presents the short description of such complex mathematical model, directions to applied numerical methods of solution to model integrated parts, and some results of parametrical calculations of reentry main aerothermodynamical characteristics of chosen DV with ballute design having different ballistic parameters of atmospheric entry — vehicle and angle.

Keywords: lander, inflatable braking device, a thermal barrier coating, trajectory, ablative materials

One of the main tests of RST (rocket and space technology items) is the test of their resistance to pyrotechnic shock loads (actuation of pyrolocks, pyrovalves, etc). Shock loads from pyrotechnic devices are characterized by a broad range of frequencies and amplitudes. This shock loads can’t lead to the entire structural failure, but can dactivate electronic components which are widely used in modern spacecraft and launch vehicles. As a result shock experiment testing is very important for the designers. Testing of RST items in accordance with reliability requirements and state standards determine the necessary maximum similarity between real and prototype shock loads. One of the most important tasks is the construction of shock test machines that are capable to reproduce the normalized load in case of shock response spectrum. Shock test machines can’t be used in the testing of large-scale models. It is explained by the fact that source of shock loads are mounted in different points of structures. In order to meet industrial specifications or industrial space requirements it is necessary to develop mobile devices that can produce the required loads. Therefore mobile shock pulse generator which is capable to simulate the pyrotechnic shock loads on large-scale models was designed in the TSNIImash. This paper presents the techniques of modeling the effect of pyrobolt actuation on the launch vehicle fuel tank in plane of separation with the help of developed generator. Shock response spectrum was calculated and compared with the previous experimental data. Using the finite element package Abaqus (Explicit) numerical simulation of shock load impact on launch vehicle fuel tank was conducted. Numerical results show high coincidence in comparison with experimental data. The proposed method of calculation allows numerically choose the parameters of a mobile shock pulse generator to generate the required loading conditions. Developed generator and established methodology of numerical simulation make it possible to carry out qualification testing of structures without the use of standard pyrotechnic devices.

Keywords: pyroshock measurement, pyrovalve, experimental pyroshock simulation, experimental development, numerical simulation

Methods for complex technical-economic analysis of perspective rocket-space system (RSS) are continuously improved. First of all it is connected with the new directions of its technical and technological development and with change of technical-economic tasks of research. It is also should be underlined that development of such RSS objects as space system (SS) of Earth’s remote sensing (ERS) is carried out in condition of significant indeterminations and connected with great cost. It makes the acceptance of project decisions much harder. Project mistake can lead to the great financial loss and others negative moments. According to this there is a problem of methods creation for project decisions estimation, in which indeterminations would be taken into account and possibility of risk would be minimized in order to provide stability solutions and increase efficiency of space systems exploitation with restricted cost.
The task about SS ERS technical-economic characteristics estimation with its modernization in planned period with indeterminate conditions are considered in presented paper. It is assumed, that indetermination of task solution is connected only with lack of knowledge about project model coefficients. The peculiarity of this problem is counting design-and-engineering decisions (DES) of modified spacecraft (SC) subsystems and adaptive models realization. The methodology for pursuance of the researches is proposed and it allows to carry out the effective specification of SS ERS project decision in condition of modernization in limited planned period with the aim to decrease space system realization risk. Two-level model of development control and method of two-level agreed optimization including guided adaptation of upper level project decisions were applied for solution to given task.
The problem statement of upper and lower development control, two-level agreed SC parameters estimation and algorithms for its solutions were given during the realization of model prototype research devoted to technical-economic characteristics investigation of modified SC being a part of SS ERS. Algorithm for research realization includes solution to upper and lower lever project tasks with execution of project decisions agreement. From one side this approach gives possibility for counting peculiarity of design-and-engineering decisions of spacecraft subsystems without expansion of project model structure, from another side this approach gives possibility to carry out estimation of apparatus subsystems parameters at lower control level (with detailed project model) taking into account dynamic of functional restrictions (mass, size, informativeness and energy). Approach specification gives possibilities to organize multivariate researches (with limited work schedule) and provides optimal project decision-making with the help of expansion of possible solution area. It is assumed that parameters of instrument subsystems module are known and detailed project analysis was carried out at lower level of development control and only for module of target equipments (MTE) taking into account its DES peculiarities. In project models applied for investigation of perspective SC there are underlined key parameters (coefficients of cost dependences). Coefficients are random values and have dispersion. During the solution to model task the statistical method of agreed two-level optimization of SS ERS with presence of random uncontrollable factors are used. Research on solution precision estimation and realization risk of space system project are presented and analyzed.
Application of proposed methodology with appropriate customization of project models allows to investigate influence of DES of SC subsystems on technical-economic characteristics of vehicle as well as on space system including reliability, quantity of system SC and system recovery period, besides, it allows to carry out technical-economic estimation of SC modification variants and evaluate space system realization risk in planned period.

Keywords: space system, earth observation systems, indetermination, realization program, multilevel process, two-level model

Analysis of the flight scheme (Earth — Earth — Venus — Venus — Venus — Venus — Venus) for the spacecraft injection into the system of working heliocentric orbits with low perihelion and large inclination.
The optimization problem for the heliocentric trajectory Earth — Earth — Venus using deep space maneuver and gravity assists from Earth and Venus with a predetermined value of the hyperbolic excess velocity after Venus approach is formulated as an unconstrained minimization problem. Active set method and sequential quadratic programming method are used to minimize the objective function.
The analysis of the flight scheme (Earth — Earth — Venus — Venus — Venus — Venus-Venus) for SC injection into the system of working heliocentric orbits with low perihelion and large inclination with deep space maneuver on the Earth — Earth heliocentric trajectory.
The characteristics of spacecraft trajectory into the system of working heliocentric orbits using «Soyuz-2 launcher», the chemical upper stage «Fregat», and liquid rocket propulsion system with a specific impulse of 310 seconds are obtained. The mass estimations of the spacecraft are presented.
The development of space research programs with using the heliocentric trajectories. For example, it can be used for study on inner heliosphere of the Sun from close distances and positions of outer ecliptic plane.
Conclusions
The one flight scheme of space mission to study the Sun using a system of working five heliocentric orbits is presented. It is not intended to use main propulsion on these orbits. The transfer from one orbit to another orbit is carried out by gravity assist at Venus.
Heliocentric trajectory of Earth — Earth is realized with one deep space maneuver. Gravity-assist maneuver near the Earth is implemented in such way that the spacecraft transfers without the running engine from Earth to Venus and the first gravity assist from Venus is carried out.
Analysis shows that the considered transport system which is based on the «Soyuz-2», the chemical upper stage «Fregat» and liquid rocket propulsion allows to insert the spacecraft into the system of working heliocentric orbits with relatively large mass of spacecraft to study the Sun.

Keywords: spacecraft, Sun exploration, flight scheme, gravity assist trajectory, heliocentric orbits, pulse rate

In this paper we consider the direction of improving ballistic calculation algorithms (BA) and point trajectory of uncontrollable aviation missile (UAM). It is shown that factors leading to a decrease in accuracy is not quite accurate count in the BA value of the initial angle of nutation UAM occurring in the presence of attack angles and aviation system slipping at start-up.
Algorithms for amending the calculation of the initial conditions of the UAM movement give a significant reduction in BA systematic errors. Take into account the particular application of problem under solution the specific set of coordinate systems were chosen, for which there is a deduction of the necessary equations in the case when equations are not available in the approximation of amendments by ground-based stationary electronic computers and subsequent transfer to the ballistic takes place. Calculations for UAM model that show the effectiveness of the proposed method are presented as an example.

Keywords: ballistic algorithm, uncontrollable aviation missile, the initial angle of nutation

This paper analyzes types and sources of powerful electromagnetic interference (PEMI), affecting elements and systems of an aircraft’s electrotechnical complex (ETC), and describes a method for calculating the current flow through the braided shield of a cable, and the common mode voltage on inner conductors of a shielded cable, inside a wing made of composite material, at the impact of a lightning’s electrical field and the magnetic field of current flowing along the aircraft’s fuselage.
The calculations are performed in two steps. At the first step, the model calculates the impact on a shielded cable, located in wing of composite material, of a lightning’s electrical field and the magnetic field of a current of 200 kA, flowing along the aircraft’s fuselage, together with the current flowing through the braided shield of a cable, all using the MathCAD software package. At the second step, the model calculates the impact of current flowing through the braided shield on inner conductors of a shielded cable, along with the common mode voltage on inner conductors of a shielded cable, using the MATLAB software package. The calculation of the current flowing through a braided shield and of the common mode voltage on inner conductors of the shielded cable, performed for different braid types, show that the lightning’s electrical field and the magnetic field of the lightning’s current, flowing along the aircraft’s fuselage, induce a significant amount of electromagnetic interferences (EMI) on the inner conductors of shielded cable, located inside a wing of composite material, that can substantially deteriorate the ability of aircraft systems, cause the loss of operability of some elements or systems of the aircraft’s ETC, and trigger a situation of emergency.
Calculation results give the designer a possibility to guide the choice of braid types with regard to the amplitude of induced common mode voltage on the inner conductors of a shielded cable and to the mass requirements for braids, and shows the need for additional measures to protect element and systems of an aircraft’s ETC from the impact on shielded cables, located in a wing of composite material, of the lightning’s electrical field and of the magnetic field of the current flowing along the aircraft’s fuselage.

Keywords: electrical wiring interconnection system, lightning, braided shield, jet, electromagnetic interference, electrotechnical complex

The authors propose a radically new implementation of a circuit synthesis technology of controlled sine wave inverters on the basis of standardized modules made from reversible impulse converters with pulsating output voltages.
In the task of regulating the power of an induction motor through a three-phase inverter, the quality of the inverter output voltage is of primary importance, namely — the sine shape and symmetry of the phase voltages and the speed of regulation of their main parameters: amplitude, frequency and temporal phase. Consequently, the prospects for a successful energy-efficient aircraft drive are largely determined by the output parameters of the controlled sine wave inverter. Currently applied technologies of circuit synthesis for regulated sine wave inverters cannot simultaneously provide satisfactory parameters of: mass and dimensions (unit weight and efficiency), reliability (heat resistance, durability and repairability), electromagnetic compatibility (levels of conductive and field electromagnetic emissions), regulatory and dynamic parameters (switching speed for modes of power conversion and it’s direction, value of transients in the regulation of amplitude, frequency and phase shift voltage), costs and operational (maintenance costs and repair work, taking into account downtime transport).
Therefore it seems most advantageous for research to study new principles of conversion of DC voltage into a sine wave form, excluding some of the described drawbacks and, if possible — to create a fundamentally new technology of circuit design.
The technology developed here can significantly improve parameters of mass and dimensions, reliability, dynamic regulation, cost and operation, and provide high quality of output power, improving the efficiency of onboard electric motors and the interference emission level. The main focus of the technology proposed here is an overall improvement the energy efficiency of the aircraft electricity system.
The article may be useful for developers of aircraft power supply systems, as well as a wide range of experts in the field of power electronics and electric drives.

Keywords: regulated inverter, standard modules, reversible impulse converter, aircraft electricity system, power supply

The authors propose a new circuit solution for the design of multi-phase reversible frequency inverters for aviation applications with double-acting direct connection of cyclic converters with active correction of power factor for electromagnetic compatibility with power supply system of an aircraft owing to the high quality of current being consumed. It is assigned for use in the secondary power supplies of prospective aircraft with power-intensive, in particular, fully-electrified equipment.
Electronic frequency converters with direct connection (cyclic converters) carry out single conversion of electric power in both directions. They are applied in aviation, and other onboard transport autonomous power supply systems, but still have several drawbacks.
The authors propose a principally new circuit solution for creating an aviation cyclic converter free of the drawbacks. The main difference from similar (multiphase) circuits is that a function of the pulse modulator for power factor correction was added. The proposed circuit solution for the design of onboard aviation cyclic converters with active power factor correction and high quality input current (i.e. with good electromagnetic compatibility) allows:

• reduction ofthe smoothing output filter size;
• mitigation of requirements both for the value of transient reactance, and calculated power of aircraft onboard generator;
•  essential increase of the power consumption factor compared to that in conventional thyristor cyclic converters;
•  reduction of the voltage loss in the input wire connections and lowering their weight;
•  relatively simple realization of switching synchronization with angular position of the rotor of electrically-driven motors without special position sensors.

Keywords: cycloconverter; inverter; converter; PFC; power supply

Optimization problems are the most often solvable in aerospace design engineering. That’s why tools for solving such problems are really requested and used both in autonomous engineering procedures, mainly aided at search for Pareto solutions, and in their results agreement. Various optimization methods are developed and widely used in practice including in interactive operation. But a designer should know and understand specific conditions, which the constituent elements of formal optimization problem should follow. He also should be able to manually convert his initial applied problem to the form of the relevant optimization algorithm class. If one takes into consideration that the majority of practical problems enable only computational solution, a designer should have knowledge in programming to get the integrated program. The objective of this paper is to develop an approach to programming components construction for solving optimization problems in CAD system so that to assure automation of the whole solution process starting from a verbal description to receiving final results.
The approach is based on the fact that current automation facilities are one of CAD system components. Another obligatory CAD components are mathematical models of design products, e.g. models implemented as application package . It is proposed to use structural properties of mathematical models variables and relations to define the elements of optimization problem formal definition: criteria, controlled variables, target function, functional constraints and to check its validation.
For the used mathematical model for one verbal description there might a lot of correct formal definitions. Thus it is necessary to evaluate their computational complexity. Such evaluation can also be made by analyzing structural properties of the model based on the possibility of formal problem decomposition into the tasks of smaller dimensionality and on exclusion of required calculations out of an optimization cycles. There are no doubts that the final choice of the variant to be used should be made only by a designer himself.
In this paper the requirements are formulated to the computer-aided tools of optimization problems solving in CAD system. The main operators are identified, they implement the following:

• conversion from aninitial verbal description oftask, which was performed byadesigner, tothe formal definition predefined inoptimization methods;
• analysis of initial task validation and its correlation with the used model and the formal conditions of optimization;
• identification of incorrectness reasons;
• generation of many correction variants with evaluation of the necessary computable labor content;
• algorithmization ofthe general procedure ofthe solving.

The general operator scheme of automation instruments operation is formed.
The integration of optimization methods and algorithms of calculation planning by using mathematical models of aerospace engineering makes it possible to sufficiently extend the possibility of solving different engineering tasks in CAD systems.
Taking into consideration the proposed approach in realizing CAD program components one can form computer-aided procedures for searching optimal project decisions and provide their high efficiency as well as convenience for designers.

Keywords: design process, aerospace engineering, CAD system, optimization problem, method of optimization, mathematical model, structural properties, operator scheme

Presented work is focused on conditions determination for educing of informational «bottom-up» links in the process of aerospace engineering designing in integrated automated systems, and it is based on the following statements:
· the design process is considered as consistent stage-by-stage specification of the product under design;
· synthesis and analysis of decision-making at every level of detail are based on the use of mathematical models;
· models in use represent a common database of the automated design system made-up of a set of elementary information-related (in the limit — peer/scalar) models.
The specifics of «bottom-up» relations, for which identification this work is aimed, are what they can be determined only in the synthesis of decisions at lower levels. Moreover, they have «shimmering» character, i.e. they can appear and disappear depending on the solutions variant synthesized at these levels.
The main result is determinated properties and structure introduced in consideration of formal construction — «tandem model». It is defined as a set of various levels models, the main feature of which is the possibility to correct model of higher levels based on the calculations results for underlying models. Formal condition for tandem-like groups of elementary models used for synthesis and analysis of design solutions on i and j levels (i, is ,  — united vectors of relevant models groups variables. This condition allows to considerate the results of calculations by models-level as a computational experiment for models identification of level. The link between models of different levels is implemented as a solution to the problem of identification.
It is shown that on the basis of the number of elementary models, forming the base of models of integrated automated system, the different tandem models can be defined. In such case the tandem model levels are formed according to the standard procedure of two-level models formation. The fact of formation of each of these models should be followed from the introduction into consideration variables in the process project details specification, while doing one or group of design operations.
The proposed representation of tandem models allows to quickly identify the presence of «bottom-up» links between levels of project details as a fact that at these levels the models are forming the tandem model. The existence of such relationships makes it necessary to have vertical coordination of the design decisions at different levels. The formal conditions of tandem-like models allow algorithmization of this procedure and control the vertical consistency of decisions made in the design of complex products of the aerospace industry.

Keywords: tandem model, project specification, identification task, coordination of the design decisions, aerospace engineering

This article describes a Preprocessor of an educational training software program intended for preparation of developers of engineering analysis systems used in CAD systems in aerospace industry. The purpose of the Preprocessor is preparation of the input data for calculations where proposed subject of calculations is a flat bar. The Preprocessor generates an optimal finite element mesh using various triangulation methods and evaluates its quality before any calculations take place.
The geometric model of the designed object is created with the help of standard primitives such as point, line, arc, circle, and is split into 8 nodal zones. The finite element mesh is generated using isoparametric coordinate method, frontal method, and several Delaunay methods. In the frontal method the density function is specified in numeric form interactively using graphical elements. In Delaunay methods first the set of support nodes of the finite element mesh is generated (in 5 different ways), then triangulation of the area is carried out using either S-Hull algorithm or the Paul Burka’s method. Results of all triangulation algorithms are attached. Additionally, the Preprocessor has a subsystem to optimize the mesh using the increasing minimum angle method and a subsystem to regulate the mesh by positioning each node in the center of gravity of the polygon. There is also a subsystem to analyze the quality of the mesh, allowing to review the full specification of the mesh with graphical highlighting of areas with high and low mesh quality. As a result, the Preprocessor is a complex tool designed for generating meshes of various types and for analyzing their characteristics. The UI of the system was designed with the purpose of training the user. It was developed using standard WinForms library. The application is using MDI (Multi Document Interface) technology allowing users to work on multiple projects. The development was done in C# programming language using .NET platform (IDE Microsoft Visual Studio 2010 Ultimate) and Tao Framework that allows .Net and Mono developers access to OpenGL and SDL.
During the use of the Preprocessor significant scientific results and lots of hands on experience were obtained.

Keywords: CAE, finite element procedures, triangulation procedures, optimization procedures of finite element model

The article describes the methods of building a single image from images bands obtained by separate CCD sensors satellite cameras with overlapping areas. The aim was to develop methods that provide the more accurate matching and focused on working with large images (of the order of several gigabytes) which have a small area of overlap. It is also necessary to obtain a refine model of satellite acquisition for new joined image. This refine model can be used for photomap building when necessary.
Proposed photogrammetric method taking into consideration the multivariate model of satellite acquisition is such a refine parametric model that all its images are matched in projection of Earth. The iterative Gauss-Newton method with some additions for matrix regularization was used to refine the model. This optimization problem works for minimize residuals in binding points in Earth projection. Support points on the Earth for stability of flying model refine was also added. For regularization task, two weight matrices for model parameters was built which are supporting and binding points using prior information about satellite parameters and points precision. Using backward photogrammetric formula single image in focal plane is obtained.
Offered photogrammetric method for image joining have more easily implemented algorithm. This is achieved by reducing the number of parameters to refine. The method also has good image joining precision and gives a good potential precision for image transformation in projection of the Earth. This is obtaining by using together binding and supporting points.
The methods can be applied to obtain a single image for automatic geometric correction from a set of neighboring images at the stage of the initial processing of satellite images. These methods belong to the level of processing 1B according to the international classification which means that it includes radiometric correction and geometric correction of systematic errors of CCD sensors in scanning system.
The advantage of this method is that it provides not only joint image but also new refine parameters for satellite acquisition model. These parameters provide good joining of images and coordinate positioning of joined image in Earth’s projection. So one can obtain more accurate photomap from source images.

Keywords: satellite images geometric correction, image joining, photogrammetric method, satellite model parameters refine

The research purpose is to examine the approaches to implementation of analytic adjustment and calibration procedures of accelerometer and gyroscope blocks which are included in a strapdown inertial navigation system’s inertial measurement units, and also to prove the efficiency of the presented analytic adjustment and calibration procedures and algorithms.
Procedures for analytical adjustment and calibration of groups of sensors are performed during predefined sequences of standard actuations (rotations to certain angles), applied to groups of sensors in the IMU basic coordinate system, and with the recording of the subsequent sensor’s output signals. The algorithm identifies the orientation of measurement axis and finds the estimates of some of the instrumental errors, that remain constant during the procedure, such as sensors bias (zero signal shift), scaling errors and output nonlinearity.
This article describes the analytic adjustment and calibration algorithm, which provides bias, scale coefficients, axis direction cosines and output nonlinearity coefficients of the sensor measurements. It is shown that the adjustment and calibration accuracy depends on the number of measurements and on instrument spindle center line mounting angle above the horizon, α. With up to 24 measurements, the optimal range of instrument spindle center line mounting angle is from 40⁰ to 55⁰, in accordance with developed adjustment. It is also noted that the accuracy of adjustment and calibration significantly depends on sensor measuring axis orientation in the basic instrument coordinate system. The calculations have shown that by using an optical dividing head with accelerometer zero signal variation not exceeding 2*10^-5g and α = 40⁰, within 24 measurements, the limiting errors (with probability 95%) from all perturbing factors are the following: the bias error is 0.89*10^-5g, the scaling factor error is 0.0020%, the adjustment error is 5.4 seconds of arc. For the gyroscope block with gyroscope bias variation not exceeding 0.04 degree/hour and α = 39⁰, within 24 measurements, the limiting errors (with probability 95%) from all the perturbing factors are the following: the bias error is 0.0038 degree/hour, the scaling factor error is 0.025%, the adjustment error is 26 arc seconds.
These analytical adjustment and calibration algorithms may be used during production or operation of accelerometer and gyroscope blocks or in data acquisition systems or complexes built upon them, e.g. strapdown inertial navigation systems. These algorithms provide a possibility to improve resultant accuracy of navigation equipment and to lower the technological requirements for block production, as in this case, the complexity level of some of the requirements on the accuracy of the sensor’s construction may be considerably reduced. Beside this, such devices may even be produced with standard available equipment.
This data demonstrates the high efficiency of analytical adjustment and calibration procedures on simplified tests, using available standard equipment, for periodic inspection of functionally redundant inertial measurement units during their operation.

Keywords: strapdown inertial navigation system, inertial measurement unit, adjustment, calibration,procedure, modeling

The present research paper focuses on the task of integration of inertial and magnetometric sensors for navigation of an autonomous flying vehicle. The actuality of the problem is based on specifics of widely used strapdown inertial navigation systems (INS). In the absence of additional navigation aids to supplement INS, zero drifts of gyroscopes lead to errors in the angular rate measurements, which result into so-called «angle drifts», accumulated proportionally to the vehicle operation time.
Authors of the article consider the opportunity to reduce these errors by augmenting INS by magnetometric subsystem. For this purpose, a three-axis magnetometer should be installed on board. The magnetometer measures total magnetic-field vector at the current point of the operation. The developed algorithm of coprocessing of inertial and magnetometric measurements is described in the body of the paper. Providing that pitch and roll drifts are compensated, the algorithm gives improved estimate of the course angle. It is assumed that pitch and roll drifts could be estimated with satisfactory accuracy by processing measurements of the accelerometers during the periods of the vehicle straight-line motion.
To verify the algorithm performance the simulation model of the integrated system operation was used. The model of magnetometer measurements implements components of magnetic field conditioned by the Earth and the vehicle frame. Mean value, variance, and variation range of the course angle estimation error were used as performance indexes. These statistical characteristic were estimated with respect to simulation time.
Simulation results given in the concluding part of the paper prove the operating capability of the algorithm developed. Width of the variation range of the integrated navigation system resultant error is reduced in 40 times in comparison with the variance range for the sole strapdown system error. It was shown that efficiency of the algorithm is degraded in presence of inexact estimates of pitch and roll angles. Based on the results, authors suggest modification of the scheme of system integration and improving the designed model. The conclusion also contains recommendations on using the integrated system on board of flying vehicles.

Keywords: integrated navigation system, magnetometric system, simulation model, strapdown inertial navigation system, aircraft

A structure of an electronic circuit element is considered. It is shown that the physical failure model for a circuit element is based on two types of binary events: a conductor break and a dielectric breakdown. The physical failure model is formalized in the form of the binary failure distribution law and the binomial failure distribution law. It is proved, with the Lyapunov limit theorem, that the failure distribution law for electronic components and systems is normal. Since no failures are available on the negative part of the event time line, the failure distribution law for a circuit element is truncated normal. Some estimates of expectations and probable deviations of failures are obtained for electronic circuit elements. These estimates allow us to state that the failure probability distribution density and failure rate are equal and constant over the period up to one million hours (hundred years). Therefore, the circuit elements of electronic systems operate practically on the stationary part of failure rate. However, while the theory of redundant fail-safe systems with a great number of circuit elements is built, the application of the Poisson failure law produces relationships became less adequate to the proposed model. The results presented in the paper can be interesting for specialists in reliability theory.

Keywords: electronics, element, probability, expectation, dispersion, random quantity, distribution law, normal law

There are many scientific works about the problem of the moving air object spectral signs reception and processing. The range portrait is one of air object spectral signs. The special range portrait processing can assure the identification and moving target detection in the perspective frequency tuning radar. One known moving target detection method is based on special processing of air object complex frequency response characteristics and air object range portraits. The scientific purposes of this work are the development of the offers on realization of this moving target detection method for perspective frequency tuning radars and the development of the air object spectral signs reception and processing algorithm for identification and moving target detection.
As a result of the research, it was developed the air object spectral signs reception and processing algorithm for identification and moving target detection in the perspective frequency tuning radar. For its realization in circular review radar stations, the detection zone breaks on azimuthal and range channels. Probing signals represent sequence of packs pairs of frequency hopping signals with the repetition period between steams, equal to three packs duration. The law of frequency change for each pair of packs is distinguished but in pair it is identical, the period of impulses repetition changes from a three of pair packs to a three. Division of signals into frequency channels is offered to be spent by means of a set of strip filters. The decision on the presence of air object is accepted on the basis of processing results of each three steams of frequency hopping signals packs with in case the difference between three received estimations of speed does not exceed the established threshold. The air object azimuth is defined under number of the azimuthal channel in which there was an aerial normal at the moment of radiation of the second pair processed frequency hopping signals packs from three structures.

Keywords: identification, range portrait, moving target detection, range rate, frequency hopping signal

It is shown theoretically that the failure flow for electronic elements and systems is not a Poisson one. In order to the failure flow being a strictly Poisson flow, the probability of a binary event must tend to zero, with the number of elements approaching infinity, that dis-agrees with the physical model of a failure. At the same time, multiple statistical data for electronic element and system failures obtained with rigorous calculation methods, con-firms the Poisson character of their failure flow. The paper discloses and gives a strict ex-planation for this contradiction between the practice and the theory. The contradiction oc-curs because the operational time is less than the mean failure time for a circuit element by a factor with value from ten thousand to hundred thousand. Over such a long interval related to the maximum value both the failure distribution density function and the failure rate function have in practice the same values and they are stationary during the entire operating life. The proven theorem declares that the failure rate for a system, which consists of the infinite number of absolutely reliable elements (element reliability characteristics and their quantity tend to infinity), tends to a constant value. The theorem shows that it is not reasonable to improve system reliability for a system with a great number of elements by improving the reliability of system elements. The results obtained in the paper will be in-teresting for researchers in reliability theory.

Keywords: failure flow, Poisson flow, limit, theory, practice,, circuit element, probability, distribution law, normal law, exponential law

The aim of this research is to formulate guidelines on the use of different criteria of preference when forming a set of equivalent alternatives (Pareto set) depending on the capacity (size) of the initial set, the number of specific criteria, etc. for the creation of the Pareto set of minimal size.
We research the dependency of the Pareto set size on the size of the initial set of alternatives, the criteria of preference, the number of specific criteria and their distribution of random values. Because the values of these parameters in different tasks vary, the research was carried out on the basis of the statistical approach, covering a wide range of problem situations. To realize this approach, we proposed and implemented a statistical algorithm of forming a Pareto set from random values of the source data. The function under study was the dependency of the size and other parameters of the formed Pareto set on the size and other parameters of the initial set of alternatives.
A generic procedure of the preference criterion description was proposed, based on the predicate representation of the concept «better», which allows setting different criteria of preference algorithmically, with minimal modifications of the program code. The concept of the severity of the criterion of preference was introduced, a number of new criteria were formulated, and computer research of their effectiveness was carried out. The guidelines for using different criteria of preference in forming the Pareto set depending on the initial data were also formulated.
The results of this study may be used in electronic CAD Results in structural-parametric synthesis method based on the morphological box
The study revealed that different methods of Pareto set forming are not equally efficient in alternatives sifting, on the number of alternatives, number of specific criteria and their values distribution. Because of that for tasks with different initial data, it is recommended to choose the criteria of preference following the guidelines formulated in this study.

Keywords: Pareto set; specific criteria; equivalent alternatives criteria of preference

The aim of this work is to study the methods of calculating the parameters of the digital phase locked loop (PLL) and the creation of its program implementation.
At present are widely used digital information transmission systems. In such systems, the receiver contains a digital part where algorithms of filtration, demodulation and synchronization at various levels are implemented. PLL is used as a basis in a many synchronization systems used in digital telecommunication systems. For example, a PLL is often used for carrier frequency synchronization and establishing of symbol timing.
In this article presents the results of a calculating and software implementation of the digital phase locked loop, which has in its structure proportional-integrating filter — PIF (Lag-Lead passive filter); shows the structural schemes of traditional (analog) and digital PLL systems. Based on the transfer function of analogue proportional-integrating filter structure of its digital realization has been synthesized. The technique of calculation of the PLL of this type is considered. Based on presented structures and the calculated values ​​of the parameters, software implementation of the device has been performed.
As an example, calculation of phase locked loop system for carrier frequency synchronization system of demodulator of digital telecommunication system has been presented. In the calculation were taken into account the following initial data: system data transmission rate is Rb = 600 kbit/s, signal/noise ratio at the demodulator input is 8 dB.
Testing of the developed system with different signal/noise ratio and different frequency offset of input signal from the nominal value has been performed. Analysis of the results showed that the performance of software realization of PLL system meet the requirements of the steady-state phase error, synchronization time and acquisition bandwidth of signal.
The results obtained indicate the possibility of using the considered techniques of designing of the PLL system and the ability to use the developed software model as a research sample, or prototype for software-hardware realization of PLL system as a high frequency synchronization subsystem of quasicoherent demodulators of digital telecommunication systems.

Keywords: PLL system, quasicoherent reception, software implementation, digital part of receiver

A stationary problem of the thermoelasticity based on the strain-gradient theory of laminated composite structures is considered. A formulation of the plane strain problem of strain-gradient thermoelasticity is formulated and an solution’s algorithm for the uniform in-plane heating are proposed. It is shown that use of the strain-gradient thermoelasticity models providing deformations continuity in contact areas allows one to consider some phenomena that are not described by the traditional thermoelasticity and can be due to the effect of stress localization in contact areas. As a result the strain-gradient elasticity theory predicts the local additional tensile stresses in the layer with a lower coefficient of thermal expansion. The obtained solution is an generalization of the traditional thermoelasticity solutions and allows one to take into account the effect of scale parameters on stress-strain state of materials.
Some test problems are considered, a heating of a single and a double layer structures. The temperature field in the layers of the considered structure is computed on the basis of the classical thermal conductivity theory. It is shown that the strain-gradient thermoelasticity allows one to take into account the dependence of stresses and strains in layers of varying thickness with the same temperature gradient so that is impossible using the traditional theory. It is shown that the resulting «reinforcement» effect is obtained, the decreasing of the thickness provides the decreasing of strain and stress intensity in layer, so that the material can be interpreted as a material with higher «effective» elastic modulus and yield strength. If the layer thickness is greater than the scale parameter of the material the classical thermoelastic model is obtained, but for the thinner layers the local stresses increase near to the boundaries so that can provide a significant variation of the stress-strain state, particularly the tension in plane of the layers can increase.
It is shown that for two-layer structures the strain-gradient elasticity predicts the occurrence of additional local tensile stresses in the layer with a lower coefficient of thermal expansion. As compared with the solution of the similar problem obtained numerically on the basis of Ansys and with the classical analytical solutions it is shown that the resulting solution generalises the classical thermoelasticity problem and allows one to take into account the effect of the layer thickness and scale parameters of the materials on the stress- strain state of the layered structure.

Keywords: strain-gradient elasticity, thermal stresses, layered structures, scale parameters