References

The article describes the “Experimental complex of supersonic wind tunnels for aero-physical tests”, which is operated in the MAI. It is intended for educational process, as well as research and development works. Both external and internal aerodynamic blowdowns are being performed with it. The complex includes two supersonic wind tunnels with interconnected systems: pneumatic system, exhaust vacuum system, as well asmeasuring and control systems. Aero-physical tests are being conducted for the models of a 30-300 mm diameter and a 0.35-1.5 m length. The duration of the experiment is 0.2-3.0 s. The airflow velocity is supersonic, ensuring the tests for modern operation conditions of atmospheric aircraft. The air consumption is up to 5.0-150.0 kg/s at the temperature up to 720-750 K.

The wind tunnels are of the same structural scheme and differ only in sizes. The principle of “sequential experiment” is being realized. Two series of aero-physical tests are being performed after preliminary numerical thermo- gas-dynamic study. At first, the required number of low-cost approximate tests on a small-size autonomous wind tunnel is conducted. The adjustment of equipment, rigging, various systems and measurements necessary for the main tube functioning is performed. Preliminary test results of the small-size model are being obtained. Further, a small number of full-size model tests are conducted in the main wind tunnel.

The tests specificity consists in their high economy and low cost due to the short time of the experiment and availability of the autonomous pressure systems.

The article describes the sensor, measuring static pressures of the supersonic flow in the inner duct of the experimental model. Small orifices serve as sensing elements, operating as stress concentrators. The stresses are being measured with polarizing-optical method of photoelasticity. Polarizing-optical installations intended for visualization and fixation of the principal stress difference bands pattern in the experimental model are presented. The stresses determining accuracy is 1-3%.

Static pressures are being determined by measured principal stress difference near the orifices.

The task of ensuring an acceptable level of stress in all structural elements of the main rotor blade both in flight and in the parking lot of the helicopter is one of the paramount ones while its design. It is well-known that the stresses arising in the main rotor blade spar from the forces of the blade’s own weight and wind loading may reach significant values, and lead to the residual deformations appearance.

The blade tail sections are less strong than the spar elements. With the achieved spar strength level, ensuring an equal level of strength of the tail sections under the action of wind in the parking lot, especially for a blade with a large chord and width of the tail sections is necessary. Creating a light and durable tail sections design is a constituent part of the task on the main rotor blades designing.

In this regard, the strength computing method developing for the tail sections and, in the first place, its skin as the most loaded and significant by weight presents interest.

The problem on determining the stress-strain state of main rotor blade tail sections skins is being solved in the open press mainly for the cases of the in-flight loading.

The presented article proposes a method for stresses computing in the tail sections skin of non-rotating main rotor blades under the impact of wind in the helicopter parking lot, based on the numerical solution of the plane problem of elasticity theory, as well as computing stresses in the blade spar under the static impact of the wind. The obtained system of differential equations describing the skin stress-strain state by the grid method is reduced to a system of linear algebraic equations with respect to the sought displacements. The SVD-algorithm for the pseudo-solution construction was employed for this system numerical solution.

The article presents the results of computations performed for the main rotor blades skins of the Mi-38 type helicopter. The wind speed limit is determined by the condition of the tail sections skins strength of the blade being considered at the given blowing direction. Comparative calculations of longitudinal stresses in the tail section skin under the action of the blade’s own weight forces demonstrated close convergence with experimental data.

For more than 60 years of space activity, more than 6 thousand launches have resulted in the appearance of about 56,000 objects in orbit, out of which about 26,000 can be tracked from Earth. According to the Main Information and Analytical Center of the automated warning System about dangerous situations, about 3,000 objects listed in the catalog are active satellites, and the remaining objects represent space debris. In recent decades, the problem of near-Earth space pollution by technogenic objects is being worsen in connection with the space activities expansion, i.e., the tendency to the spacecraft miniaturization and launching of numerous small spacecraft groupings instead of a single large spacecraft. As of now, methods for space debris cleaning-off are being actively developed, as well as measures preventing in prospect the possibility of contamination itself.

As of passive techniques for nano-satellites withdrawal from low near-Earth orbits, the most realizable are the method of aerodynamic braking by the inflatable devices and braking devices from foam polymer materials of a foamed plastic type. The inflatable braking device possesses the following disadvantages:

– a high probability of both internal (during disclosing) and external damage (micrometeriorites, space debris particles, solar UV radiation), which will lead to rapid loss of gas composition and the shell shape deformation;

– the loss of shape will be occurring while interaction with atmosphere and, as a consequence, braking probability reduction.

Polymer material coating by foam for a braking device creating has the following disadvantages:

– the foam coating formation is of a high polymerization rate, thus, the coating spherical shape obtaining in vacuum is rather difficult to control;

– there is no proof that the foam coating will retain its structure in a vacuum;

– technical device for the foam creating is more complex than the inflatable mechanism.

Our proposal supposes foam feeding into an elastic thin-film tank of a rubber ball type. The walls of this ball will perform two functions: expand under the impact of the foam to the large sizes and, on the other hand, will limit the foam material escape into space. To realize the said method, , the activities on numerical modeling and model experiments on disclosing and filling the braking shell with the foam materials under conditions close to the operation on low near-Earth orbits are required besides developing a special polymer foam for operation under vacuum conditions.

The intensive exploitation of the near-Earth space is accompanied by the accumulation of high-speed space debris (SD) in this area, which disposal is becoming an important problem for our civilization. The methods developed by us for the of orbital space debris disposal propose employing large-sized artificial plasma formations (APF), with which help of the process of the spacecraft aerodynamic deceleration is intensified for subsequent thermal disposal. With this end in view, a large-diameter APF is being formed around space debris, which creation is being accomplished with the gas-dust environment generator. The impact of the outer space radiation forms ionization processes in the generator gas and dust environment, which, in its turn, ensures formation of a carrier system consisting of solid space debris objects and the gas and dust plasma surrounding them. Due to the fact that the gas-dust plasma medium consists of charged particles, differing greatly from each other in mass, they have proportionally large differences in the speed of movement and the associated intensity of condensation on the surfaces of solid objects of space debris. The said fact leads to dominating condensation of the light – electronic plasma component on the space debris surfaces, creating a negative charge on them, which, in its turn, leads to the positive charge forming in the APF volume. This selective charge distribution stipulates the electrostatic (Coulomb) interactions forming that attract the ingredients of the IPO structure (CM and plasma atmosphere) to each other. The sources of extra ionization are being employed at the expense of radionuclide additives application in the generator plasma, spontaneously radiating ionizing radiation, to intensify electrostatic interaction in the APF. Besides, the degree of the APF medium ionization is being increased due to utilizing easily ionizing alkaline and alkaline earth substances, possessing low ionization potentials, in its composition. Thus, the external dispersing impacts of the aerodynamic forces of the Earth atmosphere traces are being surpassed by the Coulomb electrostatic attractions inside the APF. The process of intense deceleration of such a large formation leads to a multiple decrease in the period of its ballistic existence, terminating when it reaches the dense layers of the Earth atmosphere, where its thermal disposal happens.

When a soft wing profile selecting, it should be borne in mind that the data adduced in the atlases of airfoils appears to be insufficient. This is associated with the fact that these documents reflect the data of blowing rigid models, which preserve their shape even when the area with the reverse, directed downward lifting force, is being formed at the nose of the profile. The profile of the soft wing is losing stability under these conditions. Thus, the additional parameters such as the form of the graph of the total pressure coefficient along the upper and lower surfaces distribution at its nose, the angle of attack of this value transition to the negative region and the length along the profile chord captured by this transition affect the soft wing profile selection. This length indicates how expansive the profile extra turn is when it goes beyond the critical angle, and hence the degree of danger. Besides the above said parameters, the range of the accessible angles of attack for the soft wing depends upon location and size of its air intakes and slots (if any), and obtained as the result coefficient of pressure on the surface and in the wing cavity as the profile housing stability criterion to the local crushing. This criterion should be less than one everywhere for the stable profile shape. When the soft wing yeilding of the negative angles of attack, for example due to entering the down-flow, its air intakes lose their ability to keep up the excessive pressure. The upper leading edge herewith crushes, and airfoil deforms in such a way that its centerline in the nose attains reduced or reversed curvature, and, consequently, its aerodynamic force, acting on the wing leading edge sharply changes direction turning the front segment of the carrying plane. The extra effect while the profile deformation introduces the center of pressure shift, which forces the wing forward and additionally, reduces its angle of attack (this movement is being compensated to a certain extent by the deformed profile resistance).

Relevance of the research related to the unmanned aerial vehicles is being confirmed by an increased number of their application and planned financial expenditure for the development of modern and prospective complexes with unmanned aerial vehicles. Control methods are being permanently improved, and activities on the performance enhancing of the complexes with unmanned aerial vehicles are in full strength. However, due attention is not paid to the issues of the complexes technical availability, though many patterns of unmanned aerial vehicles are as good as manned aircraft concerning their mass and volume characteristics.

The existing contradictions in theory and practice indicate the need for modeling and analysis of various types of maintenance performed on complexes with unmanned aerial vehicles. One of the ways to this problem solving consists in developing a simulation model of the maintenance system for complexes with unmanned aerial vehicles. Simulation modeling is by far one of the most effective tools for studying complex systems. Simulation modeling application in many areas of activity has a number of undeniable advantages. Modeling helps to find optimal solutions to problems and ensures a clear understanding of complex systems.

When forming a maintenance system, it is necessary to account for a large number of factors that may affect the timing and quality of work. At the same time, the adjustment of the maintenance program during operation is being practiced as well. Simulation model was developed with the AnyLogic System with a view to increase the efficiency of employing complexes with unmanned aerial vehicles. The said model allows substantiating technological process, rational periodicity of maintenance, adopting rational decision on the maintenance specialists selection, assessing their workload, as well as determining the requirements to the rational set of necessary maintenance equipment.

The developed model accounts for the effect of an extended number of the input indicators and possible states of the technical operation of complexes with unmanned aerial vehicles. The proposed model may be further employed for solving the problems of rational distribution of available resources, increasing the coefficient of technical readiness and forming a rational maintenance system for complexes with unmanned aerial vehicles.

Non-uniform airflow enters due to various reasons the engine inlet under flight conditions while flowing around the engine nacelle of the power plant with the bypass gas turbine engine. The said non-uniformity presence affects negatively its key parameters such as engine thrust, specific fuel consumption and gas-dynamic stability margin of compression elements (fans and compressor stages), and, as a consequence, the engine stability at large. Two parameters, which estimate the total pressure field non-uniformity at the engine inlet, were considered. The first one is the generally accepted parameter W (stationary component, which estimates the difference between the minimum pressure at the inlet plane and its average value). The second one is the criterion parameter ER, which estimates not only maximum and minimum pressure values, but relative sizes of zones with different total pressure value as well.

A bypass two-shaft turbojet engine with the design parameters level corresponding to the fourth generation was selected as the object of study. The calculated esteem of the inlet flow non-uniformity effect on the engine thrust- cost performance was performed with 1D mathematical model employing the well-known method of parallel compressors at the three characteristic flight modes, such as takeoff, climbing and cruise supersonic mode with various engine control laws. Rotation frequency sustenance of both engine rotors n₁ and n₂, as well as sustaining gas temperature T_t* at the turbine outlet were considered as such laws.

The study of the total pressure non-uniformity at the engine inlet effect on its basic parameters at various control laws revealed that the less effect on the thrust-cost characteristics the non-uniform airflow exerts at the gas temperature sustaining behind the low pressure turbine. The maximal effect of the non-uniform total pressure on the thrust and specific fuel consumption was revealed while realizing the program of high-pressure shaft rotation frequency n2 control. The share of the extra losses in the compressing elements due to the thrust reduction increases with the flight speed increasing and climbing and may reach up to 20%.

At present, temperatures at the inlet to the turbines of gas turbine engines reach 1500-1900°C, which exceeds the melting point of the materials from which the turbine blades are made. Despite the fact that for the most heat- stressed blades of gas turbine engines, the main cooling is achieved through the film cooling systems, convective part is present there as well, which removes a significant amount of heat. With this regard the issues of developing a convective part of cooled turbine blades, as well as the heat transfer intensification inside the blades are up-to-date. Intensifiers in the form of several rows of pins are traditionally widely used in the cooling channels of the blades located in the middle part and the rear of the airfoil. Generally, a staggered arrangement of pins relative to the direction of the cooling air flow is employed. However, a change in the direction of the airflow along the height of the feather may lead to the pins flow-around at different angles, including a flow corresponding to their in-line arrangement, which may significantly reduce heat transfer.

For the purpose of further heat transfer intensification in the blade cooling channels, this authors propose application of the pins installed in holes, as well as pins installed in transverse grooves. These modified pin intensifiers allow substantial heat removal intensifying at trifling hydraulic resistance increase, as well as reducing the shadow stagnant zone behind the pins, where heat transfer decreases, due to extra vortex formation in the cavity zone.

The article presents the results of a study of several design solutions for heat transfer intensification: pin intensifiers, pin-hole intensifiers and pin intensifiers located in the transverse grooves. The method of calorimetry in a liquid metal thermostat, consisting in the thickness measuring of zinc crusts formed while thermohydraulic cooling of the studied channels models and the heat transfer coefficients and Nusselt numbers determining by them, was employed to study heat transfer characteristics in the channels.

A basic channel with pins without recesses was selected as a channel for comparison with the results described in the literature. The experimental data obtained while the basic channel studying revealed a high degree of agreement with the Metzger data, the average deviation was less than 10%.

The experimental studies results of modified cooling channels with pins revealed that cooling channels with pins in the transversal grooves display maximum throughout among the channels being considered due to the minimum flow passage area increase. The average by length Nusselt numbers for the given channel herewith are 36% more compared to the basic channel with pins, and 22% more compared to the channel with pins placed in round dimples.

Metal rolling bearing are employed traditionally in rotor supports. These bearings disadvantages are high friction coefficient, limited rotation frequency and their susceptibility to the severe wear.

Implementation of new technologies and materials enabled application of ceramic bearings. These bearings advantage over the metal ones consists in:

– low adhesion of mating parts, low friction coefficient;

– non-magnetic properties, high operating temperature;

– chemical resistance in aggressive environments, high strength.

Traditionally, the performance check of such bearings are the tests, which require heavy economic and time spending that may be reduced by numerical simulation with modern software packages.

Numerical computation of ceramics strength characteristics of represents a problem, since it is associated with the need to build an adequate micro-cracks propagation model in inhomogeneous structures.

This article presents a model of Johnson-Holmquist (J-H) ceramics deformation and fracture, which allows estimating a micro-fracture, as well as the time and place of cracks initiation.

The fracture mathematical modeling in the J-H model is based on introduction of the fracture parameter (D), defining the degree of material continuity loss, as well as equations describing the D parameter changes in the loaded material. The fracture parameter growth is associated with deformations accumulation.

Simulation of several options of the internal ring corresponding to the real structures (conventional ring and a ring with a slit) by the finite element method was performed for the technique for the bearing strength estimation try-out. The model was being loaded by the centrifugal force in time, applied linearly, from zero to full destruction. The ceramic ring material was Carboprom-K. The properties of the analog, namely silicon carbide, were employed for the damages analysis.

The authors considered a notional sum of the friction moment and a moment from the asymmetry in the articulated steering unit under condition of the upper limit of the said aggregate moment. For the first time in the practice of studying steering units torque characteristics of liquid rocket engines (LRE), an aggregate torque parameter was introduced, characterizing the above mentioned conditional sum. The obtained parameter components may be herewith independent variables. The previously declared method of this conditional value adjusting by eccentric hubs may be supplemented by an additional parameter optimizing, i.e. the hubs eccentricity and compensating moment directly associated with it. The currently available analytical dependencies define concretely only the boundary of commencing application of hubs as adjusting elements. The numerical boundary value herewith is unambiguously defined as a half of the compensating torque, created by the eccentric hub, value. In the furtherance of the subject, the value itself of the compensating torque was considered in detail. The dependence of the said adjustable value in the form of the simplest function, which argument is the entire permissible range of the torque, on the thrust asymmetry of the steering unit was established based on the graphical solution. Joint determination of regulation commencing and the adjustable value allows elaborating a universal technique for the aggregate torque correction applicable for the articulated steering units under any limitations for the torques composing the sum. This technique may be applied herewith both at the design stage and in the process of the existing structures modification while operational conditions changing. The newly obtained analytical dependencies allow determining the margin of the friction torque increase without increasing its upper set limit.

Implementation of the new technique for torque characteristics adjustment allows reducing the process of serial structures fine-tuning to the required friction torque values by simple increasing of the admissible value. The said possibility contributes to the number of costly repeated tests number reduction.

The world leading aerospace industry organizations show interest in developing and upgrading the ultra-low- power engines, characterized by the power less than 100 W, for the small spacecraft (SC) including the CubeSat format spacecraft. This interest can be explained by the possibility of obtaining new knowledge and deriving of commercial profit while the small-size SC, equipped with propulsion units with high values of the total burn, for orbital maneuvers performing. The projects of commercial companies, aspiring covering the low-orbit space around the Earth by the information transmission systems, which represent orbital groups of the small SCs, constituting formation and jointly performing the flight task, so-called satellite constellations, may be adduced as an example of the considered interest.

Application of the small spacecraft of the CubeSat format may lead in the future to the change of the basic approach to the Solar system exploration due to the high ratio of the obtained scientific knowledge to the financial costs. Thus, the growing interest of the world market in the movement control systems for the small SC is being observed, which is proved out by the presence of scientific works and publications. Nonetheless, according to the «World’s Largest Database of Nanosatellites» European database information, more than 1300 nano-satellites were manufactured by the middle of 2020 (including the SC of the CubeSat format), and only 5% if the small SC from this number had a propulsion unit as their part.

Propulsion units for nanosatellites of the CubeSat format can be formed both on an electric rocket engine (ERE) and on a gas-powered engine (GPEU), which has a minimum volume and mass, which, in its turn, complicates the extra thermostating system placing on it.

The article describes the technique and stages of the GPEU thermal design, and adduces its thermal mathematical model, consisted of detailed thermal models of all the constituent elements of the installation, placed on the spacecraft frame, around, around which the screens with photocells of the solar battery are placed.

The article presents the results of developing and employing the thermal model of a nanosatellite with gas propulsion system of orbital operation. The said model was used for the temperature field computing, internal and external conductive and radiative heat fluxes determining. It allows as well determine gradients and rates of temperature change in stationary and dynamic operation modes with subsequent recommendations on improve the nanosatellite thermal design and reliability.

The results of thermal computations on determining temperature ranges and thermal fluxes among the GPEU elements for the considered options of its placing on the SC frame at the extreme combination of thermal loads and thermal conditions of the GPEU application set for the thermal computations are presented. The authors gave recommendations on the thermostating system improvement.

A mathematical model and results of calculation for leakage currents and floating potential of a high voltage solar battery (SB) of a spacecraft (SC) in plasma generated by electric propulsion thruster (EPT) is presented. The floating potential of the solar battery is determined as a potential of SB minus bus with respect to plasma potential generated by EPT. Floating potential value is determined according to Kirchhoff ‘s law by using electron and ion currents coming through open electrodes of solar elements and SB frame. Electron currents are calculated according to relationships known in probe theory for positively and negatively charged electrodes depending on their potential, surface area and electron temperature. The ion current is determined according to jet parameters at electrode surface without considering electric field effect to the trajectories of accelerated ions.

With the help of the presented model we calculated the floating potential and leakage currents for an abstract SB with working voltage of 150 V and with current of 16 A. SB panel has pipe frame with size of 2.5 × 3.2 m. It contains 40 strings with 60 solar arrays (SA) with size of 40 × 80 mm, working voltage of 2.5 V and current of 0.4 А. The area of SA open electrodes is set equal to 0.05, 0.1 and 0.2 cm2. The array is rotated round its own axis and subjected to the impact of SPT-100 jet. Ion currents are calculated for the worst case without considering ion incident angle to open electrodes and SB frame. As a result of calculations we reveal that SB floating potential is defined mainly by leakage current trough SB frame and its value runs up to 100 mА in the point closest to EPT jet axis. SB potential ranges from −140 up to −40 V depending on the angle of SA rotation. Maximal value of leakage current is 1400 mcА and it takes place at positively charged electrode in the area where plasma concentration is maximal. SB power loss due to leakage currents through plasma is not higher than 1%.

Leakage currents heat impact to electrodes is estimated for heat removal by radiation. We reveal that leakage current through positively charged electrodes can heat electrodes up to high temperature, cause secondary arc discharges, which can destroy electrodes and failed some SB strings. Microarcs can appear at negatively charged electrodes and they can transform into powerful arc discharge, which also can destroy SA.

The obtained results show that EPT plasma impact onto high voltage SB of the SC can be great and it should be consider under designing and testing of power plant of the SC based on high voltage SB.

The article presents the results of the correcting propulsion unit development of two nominal sizes for nano- satellites of CubeSat 3U and 12U format based on a low-thrust Freon-running thruster. Working medium selection analysis for the engine type being developed is presented. Freon R-236fa, R-227ea and RC-318 are being considered as a working medium. This Freon special feature consist in the possibility of its storage in the saturated state under the pressure less than 1 MPa (10 atm). The average specific impulse and thrust of the engine being developed are of no less than 392 m/s (40 s) and 0,015 N respectively at the temperature of the Freon being considered of T =293 K.

In the course of these propulsion units development, the following elements were newly developed, namely tanks, small-sized low-pressure control valve, small-size feeding device, receiver and a low-thrust engine, representing gas-dynamic nozzle. Application of Freon as a working medium allowed abnegating the high-pressure fittings. The pressure in the saturated state Freon, considered in the article, is no more than 1 MPa within the temperature range from 273 to 313 K.

The overall dimension of the developed propulsion unit are of no more than 1U. Its weight is about 1.4 kg for the CubeSat 3U format nano-satellite with the propulsion unit peak energy consumption of no more than 17 W. Based on the estimation, the total thrust impulse of the unit will be about 138 N × s. Characteristic velocity margin will be of 24 m/s with the tank volume of 0.25 liters for the satellite of the 5.6 kg total mass.

The overall dimension of the developed propulsion unit are of no more than 4U. Its weight is about 5.0 kg for the CubeSat 12U format nano-satellite with the propulsion unit peak energy consumption of no more than 10 W. Based on the estimation, the total thrust impulse of the unit will be about 1250 N × s. Characteristic velocity margin will be of 24 m/s with the tank volume of 2.2 liters for the satellite of the 20 kg total mass.

The result of the presented consists in the development of the propulsion units of two different standard sizes based on Freon propellant, which allow performing such maneuvers as satellite position on its orbit correction, and correction of the parameters of the orbit itself, as well as the satellite de-orbiting.

Annular nozzles are competitive with traditional central nozzles in a number of characteristics. This is stipulated by the presence of a central body, which determines the required flow structure during supersonic expansion. In the narrowing section of the nozzle, the central body contributes to the friction losses increase, and its geometric characteristics will determine the uneven flow parameters distribution and total pressure losses up to the critical section. The authors conducted numerical studies of the central body shape and of inlet section geometric parameters impact on the gas-dynamic component of the flow coefficient of an annular nozzle with a direct critical section.

The outflow processes simulation was performed with the ANSYS Fluent software within the framework of the axisymmetric approximation in the adiabatic formulation of the quasi-stationary problem, assuming that the structural supports that secure the central body do not significantly change the flow coefficient. The approach based on solving the Reynolds-averaged Navier–Stokes equations closed by the k–w SST turbulence model widely used in engineering calculations with a typical set of model constants was employed. A homogeneous gas was considered as a working fluid.

The simulation results revealed that the flow coefficient gas-dynamic component of the annular nozzles with a straight critical section can be comparable to the value of traditional central nozzles, and exceed it for certain geometric parameters of the central body, which is stipulated by more uniform distribution of parameters in the critical section. A linear dependence of the washed area increase of the central body with its ellipsoidity increase, and a nonlinear nature of the change in the total values of the friction stresses with an extremum for the spherical shape of the central body are shown. The most optimal shape of the central body is a spherical one.

The flow coefficient of annular nozzles with a straight critical section depends significantly on the conjoint distribution of the central body geometric parameters and the outer contour of the narrowing section. With the optimal shape of the central body, and the ratio of central body diameter to the outer contour diameter in the minimum nozzle cross-section of the order of 0.7, the flow coefficient gas dynamic component acquires maximum value, exceeding this value of the conventional central nozzle by 0.3%.

In contrast to the flow coefficient of conventional central nozzles, the flow coefficient of the annular nozzles increases with pressure increasing in the combustion chamber.

The presented article describes a computational study of a modernized hybrid cryogenic power plant. Liquid nitrogen was selected as the working fluid of the cryogenic installation. The schematic diagram of a hybrid cryogenic power plant consists of an internal combustion engine (ICE) and a cryogenic plant (CP); the heat source is liquid antifreeze, which collects the heat from the internal combustion engine and delivers it to the liquid nitrogen.

The installation principle of operation consists in the following. The cryogenic working fluid (liquid nitrogen) from the tank enters the heater through the cryogenic pump, where nitrogen obtains thermal energy from antifreeze. The antifreeze, in its turn, is the coolant for the ICE. From the heat exchanger gaseous nitrogen enters the piston expander, where the polytropic process emanates. The resulting work is transferred to the screw actuator.

The cryogenic power plant operates according to the open Rankine cycle. The open circuit of the power plant, which employs the low-potential heat of liquefied nitrogen, is quite simple and economical. Both nitrogen and air, liquefied natural gas, etc. can be employed as a working fluid.

The Rankine cycle was constructed in T-S coordinates (temperature-entropy coordinates) of nitrogen with the Coolpack application software package [15]. Thermodynamic parameters of the basic points were computed employing an algorithm for conducting a computational study of the hybrid cryogenic power plant parameters [14, 18].

The working body is being heated in the heat exchanger-evaporator to the temperature of the upper heat source [19]. Technical specification indicates that the flight altitude of the unmanned aerial vehicle (UAV) is 2000 m, and the temperature of the hot coolant is 363 K [14]. Computational study of the UAV aerodynamic characteristics revealed that required power would be 15 kW at the cruising flight.

The results of the computational study demonstrated the necessity of both temperature and pressure increasing at the piston expander inlet for the hybrid cryogenic power plant efficiency enhancing. Temperature increasing up to 363 K may be achieved through employing the heat removed from the ICE, employing liquid cooling system. It will allow reducing the cryogenic working body consumption to 0.053 kg/s while ensuring the power output of the UAV power plant at the level of 15 kW.

The article presents the study of optimal control programs for spatial relative motion at near-circular orbit.

Two spacecraft, namely maneuvering, equipped with engine of finite thrust, and passive, located in a circular orbit are being considered.

The problem of bringing the maneuvering spacecraft to the specified position relative to the passive one is being set. Equations in cylindrical reference frame, which origin is placed in center of mass of the passive aircraft, and equations of motion are linearized near the passive spacecraft orbit, are used for construction of the dimensionless and invariant to the datum orbit mathematical model of relative motion.

New variables, describing the relative motion in the orbit plane in terms of the secular and periodical motion, and in the form of the maneuvering spacecraft oscillations amplitude and phase relative to the passive one, are introduced.

The authors demonstrate that longitudinal motion in linear approximation is associated with the lateral one only through the controlling accelerations, in which connection two control options are considered. The first one is joint, when both longitudinal and lateral motion components change simultaneously, and no limitations are imposed herewith on the thrust vector orientation of the maneuvering spacecraft. The second one, which is no less common, supposes sequential longitudinal component elimination of the relative motion, and then the lateral one.

Time optimal controls are obtained with the Pontryagin maximum principle application. The optimization problem is reduced to a two-point boundary problem for a system of differential equations, which is solved for three qualitatively different boundary conditions, namely the longitudinal periodic motion correction dominance, the requirements of longitudinal secular motion correction and the requirements of the lateral motion correction dominance.

An additional calculation of the required turning speed of the active spacecraft was performed to the optimal control program accomplishment, which indicated the necessity of introducing the passive sections on the trajectory at time instants corresponding to an almost instantaneous turn of the spacecraft by one hundred and eighty degrees around its axis.

The article is reviewing the history of emergence and development of descent vehicles with inflatable structural elements. Descent vehicles equipped with inflatable braking devices possess the following advantages:

1. The payload volume fraction increase under the launch vehicle fairing.

2. The diameter of the inflatable braking systems is not limited by the size of the launch vehicle fairing.

3. The folded inflatable braking system does not block access to the payload

The article presents also specifics of the descent vehicles with inflatable braking devices. These specifics are entailed by the inflatable braking devices deformation occurring while their motion in the atmosphere. They are:

1. The descent vehicle aerodynamic characteristics changing.

2. The descent vehicle the dynamic stability changing.

The authors educed the ongoing research relevance, which was confirmed by works of Russian and foreign scientists.

The object of the research is a descent vehicle with a conical inflatable braking device, which control is being perpetrated by the center of mass shifting. The hypothesis in the ongoing work is the control method, namely, the center of mass displacement on account of the payload rotation.

A study of the angular motion that occurs during the descent vehicle control was conducted to confirm the said hypothesis.

A mathematical model, accounting for the considered control method specifics, was developed to study the angular motion of the descent vehicle. Solution of the equations of the mathematical model was performed for several cases of initial conditions of motion. Simulation results are presented in the form of graphical dependencies, reflecting the points’ movement trajectories on the descent vehicle surface, as well as angular velocities changing in the process of movement. Inference was drawn for each of the considered cases of the initial conditions of motion.

Solution of the mathematical model equations was performed by the 4th-order Runge-Kutta method.

Analysis of the results allowed drawing the inference on the descent vehicle angular motion stability, as well as revealing further trends of studying the control method being considered.

Light vertical takeoff and landing (VTOL) are being regarded in many countries as basic means of rectifying the tasks of urban air mobility. Rotary-winged aircraft may be employed as both aero-taxis for passenger transportation and by various city services including police, ambulance and fire-fighting service. Conventional helicopters, quadcopters, multicopters, including those with aerodynamic surfaces for the flight range and endurance increasing, as well as transformable (convertible) in flight aerial vehicles were being proposed as aerodynamic configurations. Scientific studies in the field of design, flight dynamics and control systems of convertible aircraft or tilt rotors with 90° swiveling rotors, are in full strength all over the world (predominantly in the USA) since 1950s. As of now, the tiltrotors are widely applied in the military-oriented aviation (Bell/Boeing V-22 Osprey, V-280 Valor) and being prepared for application in civil aviation (AgustaWestland AW-609). In the article being presented the tiltrotor scheme with two swiveling rotors was selected as an aircraft scheme for urban air taxi, as the one combining the advantages of both helicopter and airplane. Its main advantages are:

– the ability performing hovering mode, vertical takeoff and landing;

– high speed of horizontal flight;

– higher flight endurance and range.

The presented article considers nonlinear mathematical model of light convertible rotary-winged aircraft flight dynamics with a view to this aerial vehicle application for solving the task of urban air mobility. This flight dynamics mathematical model development was being accomplished with the MATLAB/Simulink software package. The alike VTOL is being supposed to be equipped with a traditional power plant, such as internal combustion engine and gas turbine engine, or electrical (hybrid) one. A system of differential equations of solid body motion was used for the flight dynamics model description. Mathematical modeling of the tiltrotor main rotors was being performed employing the blade element theory. For the modeling accuracy enhancing of energetic maneuvers with drastic changes of the flight parameters, such as overloads, as well as translational and angular VTOL velocities, the mathematical model accounted for both angular and translational displacement of the main rotors. The algorithm for aerodynamic calculation of the airframe elements, such as wing, fuselage and empennage, of the convertible aircraft using analytical models was proposed. Synthesis of automatic control system (autopilot) for all flight modes (“helicopter”, “airplane” and transitional) was accomplished. Tiltrotor trajectories computing for the main flight stages (hovering and a flight with low translational velocity, transitional modes from “helicopter” to “airplane” and back, the flight along the rectangular route, steady turns, as well as upward and downward spirals) was performed.

The article considers the up-to-date problem of the power plant parts resource enhancing. One of this problem solution consists in synthesis of duplex coatings, representing multi-layer coatings being created by successive surface modification, and coating precipitation in the unitary operational space. The surface layer ion-plasma nitration in the plasma of non-self-maintained high-current discharge, generated by the “PINK” plasma generator is employed as a surface modification method. A method of condensation with ion bombardment is applied as a plasma-ion method for protective coatings obtaining. The authors proposed obtaining duplex coatings of the TiZrAIN system on the HNV 6.6-I1 modernized chamber-type installation equipped with the electric arc evaporators with Ti, Zr and Al cathodes.

With a view to vacuum ion-plasma coatings with the complex of enhanced operational properties creating a technology for duplex coatings of the TiZrAIN system, including successive employment of ion-plasma nitration and ion-plasma deposition in the unitary operational space was created. These two technological processes combining in the unitary operational space is performed without vacuum chamber depressurizing and overloading substrates being processed. The TiZrAIN system coatings synthesis was being performed under conditions of plasma assisting by the “PINK” plasma generator. A system of the magnetic-arc filtration for the electric arc evaporator with Al-cathode is being employed while coatings deposition, which allows accomplishing separation of the drop phase of low-melting aluminum, and contributes to substantial drop phase reduction in the coating and initial substrate roughness retention.

Samples of the 20 mm diameter and 3 mm height were obtained for studying micro-hardness, adhesive strength and corrosion resistance. The studies of the duplex coating of the TiZrAlN system synthesized by the developed technology were being performed in comparison with the multi-layer coating obtained by the vacuum ion-plasma method, as well as with the duplex coating obtained by successive pursuance of the ion-plasma nitration, and synthesis of the vacuum-ion coating (without plasma current separation).

The synthesized coatings surface micro-hardness measuring revealed that duplex coatings micro-hardness was higher compared to the multilayer coating, which is being associated with the coating surface layer application on the surface already strengthened by the ion-plasma nitration. Surface micro-hardness of the duplex coating being synthesized, obtained without plasma separation was 45.2 GPa, 46.6 GPa fabricated by the developed technology, while the multi-layer coating micro-hardness was 35.3 GPa.

The study of micro-photos of scratches and profile protocols of the fracture zones obtained while scratch-testing revealed that duplex coatings synthesized by the developed technologies was being characterized by the enhanced adhesive strength. Loading of the first cracks origination in the duplex coatings is 20 mN, whereas the one of multilayer coatings is 14.83 mN.

Corrosion rate studies revealed that with the duplex coating with plasma flux separation it was 11.7% less than the samples with duplex coating without plasma flux separation and 30.1% less than for the samples with multilayer coating. Consequently, the surface of the coating synthesized by the developed technology is more passive, which indicates its higher corrosion resistance.

The conducted studies results confirmed the prospective of the developed technology of duplex coatings synthesis application for the power plants parts protection from abrasive, corrosive and erosion impacts.

Both conventional technologies for workpieces obtaining and additive technological process of direct energy and material feeding (DED) are being employed for manufacturing bulky workpieces for gas turbine engines parts from heat-resistant nickel-based alloys.

The DED technology allows managing a highly coordinated energy impact on the micro-volume of the alloy, which ensures the material structure obtaining with higher working characteristics compared to castings. As of now, nickel materials application in the area of additive technologies is limited by the ultrafast crystallization processes specifics that cause accumulation of significant internal stresses, which leads to micro- and macro-defects forming. Heat treatment is recommended for residual stresses reduction in the products after the DED process, but optimal modes of such kind of the workpieces processing are not clearly specified. On the other hand, heat treatment implies obtaining high mechanical properties. For the products fabricated by additive methods of surfacing powders with non-equilibrium structure, the similar recommendations are of rather small volume.

The place of heat treatment in the general cycle of parts manufacturing is being set depending on the requirements for the product properties. In most cases, heat treatment is being performed after mechanical post-treatment. This is associated with the requirements to high strength, hardness and wear resistance of the product material.

The article studies the effect of various heat treatment modes on the hardness, microstructure and residual stresses of the samples made of the HN50VMTUB heat-resistant nickel-based alloy obtained by the DED technology.

The DED technology of workpieces manufacturing from the HN50VMTUB alloy leads to a fairly high hardness of about 190 NB. It is well-known that the products growth from the highly-alloyed powder of non-equilibrium structure proceeds by rapid cooling, which causes structural changes similar to the aging while heating by the laser beam. Heat treatment of the grown products may be aimed at increasing the machinability by cutting and reducing the of products warping herewith, as the result of the residual stresses redistribution. In this case, the decrease in hardness may be the goal achieving criterion.

The results of the presented study demonstrate that the most economical mode of heat treatment for the residual stresses removing is the mode consisting in products heating up to 1180°C, holding for four hours with subsequent air cooling, which allows reducing hardness from 191 ±1 HВ to 135 ±1 HВ. The lowest hardness values of HB 128 ±1 were obtained after heating to 1140°C, holding for 4 hours and cooling with a furnace. Air cooling allows obtaining hardness of HB 130 ±18. On the one hand, this indicates slightly higher hardness values, but deviations are of a higher level, the level of residual stresses in the annular samples herewith are of the lowest values, which follows from the results of samples geometry changing after cutting.

The highest hardness of 311 ±8 HB was obtained at the end of heat treatment, which includes heating up to 1100°C; holding for 4 h; air cooling, and then heating up to 950°C, 3.5 h holding, air cooling; then heating up to 800°C, exposure 7.5 h, air cooling, then heating up to 700°C, holding time of 14 h, air cooling.

The microstructure analysis of the grown samples reveals that after all types of heat treatment, an inequigranular structure is being formed in the samples, and the layered structure characteristic for the deposited particles is lost.

The article considers the problem of an unmanned aerial vehicle (UAV) with a solid propellant rocket engine designing. One of the ways for the UAVs efficiency enhancing consists in qualitative improving of the design choices being made. The issues of the design links restoring between the project parameters and the UAV functioning conditions are of special meaningfulness while the UAV modeling. These design links are being restored from the samplings obtained by the UAV mathematical model probing. The project links are being constructed in the class of harmonic polynomials by the static regularity criterion, which optimization is being performed by the original method of the global extremum seeking. Mass, flight performance, economic and operational indicators as well as other criterion characteristics may be accepted as a goal function. The article being presented assumes the UAV flight range as an optimality criterion. The UAV efficiency increasing is associated with highly accurate small-sized and moving targets hitting, which leads to the necessity of the UAVs power plants further improving. The UAV efficiency, like any other aerial vehicle type, is a complex indicator, determining the UAV flight range. The highest augment in the UAV flight range may be reached through the solid-fuel rocket-ramjet engines application (SFRRE). Such engines improvement is being accomplished by way of working process and power plant structure, as well as specific-mass and energy properties of solid fuels selection. The supersonic air intake device makes significant contribution to the working process quality. In this regard, the air compressing process efficiency in the air intake is of significant importance.

Parameters selection of the power plant with the SFRRE as well as the UAV parameters, ensuring the maximum flight range of the rocket with a fixed launch weight and specified fuel margin was performed. The fuel-flow rate at the cruising section as well as inlet ant outlet cross-sections areas of the air intake are assumed as variable parameters. Optimal selection of the inlet and outlet cross-sections area of the air intake and fuel consumption allowed increasing the UAV flight range by 5.842% for the 1000 m flight altitude, and by 12.283% for the flight altitude of 10000 m.

The authors consider the issues of obtaining composite materials, widely applied in aircraft building, by the vacuum molding method. Special attention is paid to the stage of semi-preg stack impregnation by the polymer binding material melt. The processes of filtration impregnation and capillary impregnation are being distinguished while its realization process. On the assumption that the semi-preg stack is being horizontally set, mathematical modeling of the reinforcing fabric filler filtration impregnation under the impact of pressure difference in the vertical direction and gravity with account for the filler carcass damping is being performed. Provided that the reinforcement material does not swell or shrink, and discontinuity deficiency, the super capillary pores filling by the binding melt is quite rapid, and the upper, intermediate and lower semi-pregs are being distinguished. It is assumed in its turn that while the binding melt flowing in the vapor space of the filler it represents an uncompressing viscous uniform liquid, which viscosity and density do not change while the filtration process, and the melt flow is laminar and isothermal. As the result of the generalized Darcy filtration law integrating, an expression for the filtration speed of the melt in the filler layer and its full impregnation time were obtained. The article demonstrates that the time of the filtration impregnation can be reduced, and the productivity at this stage of production can be correspondingly increased. It can be achieved in the first place by the pressure drop increasing at the filler layer thickness, additional loading action on the semi-pregs stack surface and the melt viscosity reduction due to both temperature and density increase, as well as super-capillary porosity enhancing of the filler. The set regularities represent the possibility of rational technological modes selection for woven composite materials obtaining by the vacuum molding method.

Layered composite materials (CM) are of a wide application range in the design of aircraft. These materials advantage consists in the ability of changing the package physical characteristics by the reinforcement angle varying. Physical properties degradation under various types of loading [1-4], which, in its turn, affects the aircraft strength, should be accounted for while the aircraft structures design.

The presented article studies characteristics degradation of a composite material of different structure. The hypothesis that transversal cracks leading to the physical characteristic degrading and residual deformation appearance, occur in the composite material monolayer while loading is accepted. The issue of the transversal cracks occurrence in the matrix structure of a composite material is being considered on a wide scale in [1-18].

The article considers the samples from a woven organoplastic and unidirectional prepreg of carbon fiber- reinforced plastic, with various stowing of 0 – 90 and 0 – 90 – ±45 degrees, as well as with various geometric characteristics.

The article presents the results of the experiment on composite panels cantilever bending under normal climatic conditions. The samples were loaded by the forced displacements of the stop along the mounting axis with a step of 2 mm, in the direction of the profile. Unloading and measurement of residual deformations of the uttermost edge were performed after each loading step.

Stiffness characteristics degradation of the material is being determined in this article by residual deformations measuring after the sample loading. A more accurate method of cracking detection in the CM matrix structure is non-destructive testing with roentgenography methods application. The said method will allow detecting cracks in the CM structure with normalized accuracy. The issue of non-destructive defects testing in composite materials is being considered in [19-20].

The full-scale tests allowed establishing the presence of residual deformations in structurally similar flexible elements of all types of cross-section. It was revealed that the stiffness properties degradation in the composite material occurs at the cantilever bending of the sample.

Structurally, such flexible elements with reinforcement angles of 0 – 90 – ±45 display the smallest increase in residual deformations, compared to the samples, which reinforcement angle corresponds to 0.90 degrees. It is associated with the fact that organoplastics are of a braided structure, and at reinforcement angles of 0 – 90 degrees half of the fibers are not beingincluded in the overall bending of the structure. The reinforcement angle of 0 – 90 – ±45 degrees herewith allows including all the fibers of woven organoplastics in the general bend and reduce the package stiffness characteristics, which, in the aggregate, leads to the stresses drop in the monolayer of the composite material package and, as the result, the least progression of stiffness characteristics degradation.

Geometric layout affects significantly the aerodynamic characteristics of the helicopter main rotor in various operating modes. When designing a rotor with a given diameter and solidity, various geometry layout solutions are possible, such as, the number of blades, blade twist and relative position of the blades (single or coaxial rotor).

It is common knowledge that the geometrical layout has a significant impact on the efficiency of the rotor in hover [1]. For the other flight modes, the geometry impact on the rotor aerodynamics is of practical interest as well.

The study considers the effect of various options of geometrical layout on rotors aerodynamic characteristics at a vertical descent in the vortex ring state modes range in the range of descent speeds Vy = 0 — 28 m/s. The sharp rotor thrust reduction compared to the hover mode and its non-stationary pulsations are characteristic to the «vortex ring» modes, which makes these modes unsafe for the helicopter.

Wide-scale experimental studies of the vortex ring modes are extremely difficult, thus application of the state-of-the-art computational methods is rational. The studies being presented were performed based on the nonlinear vortex rotor model developed at the Moscow Aviation Institute [19].

Single two- four- and six-bladed rotors, as well as coaxial six-bladed rotor with the same solidity, airfoils and blade twist were considered. The blade twists of 0°, 8° and 16° were considered for the 4-bladed rotor as well.

Computations have been performed at the fixed blade pitch angles, ensuring the equal thrust in hover. The total and distributed aerodynamic characteristics have been obtained and analyzed, including the shapes of the vortex wake and flow-around patterns.

The smallest obtained thrust drop in the vortex ring modes was demonstrated by the two-bladed rotor. The thrusts of the equivalent six-bladed single main rotor and six-bladed coaxial main rotor have similar dependences on the rate of descent, but the coaxial rotor herewith has had lower values of the thrust pulsations amplitude. With the blade twist values growth, the thrust drop and thrust pulsations in vortex ring state increased for the four-bladed rotor. The blade twist effect on the rotor aerodynamic characteristics at the vortex ring modes is in good agreement with the available experimental data [6].

Thus, the considered technical solutions on the rotor geometrical layout (that improve its aerodynamics in hover [1]) do not have a positive effect in the vortex ring modes.

The obtained results may be handy in the rotor aerodynamics analysis in vortex ring state modes.

Airframe elements laminarization is considered to be one of the further aviation development paths. Airframe elements laminar flow can be provided passively (by Natural Flow Laminarization, NFL), as well as by means of boundary layer suction through the perforated aircraft skin. The NFL utilization proves to be rational as small regional aviation regards. A regional aircraft with its wing to be laminar and its engines to be arranged over the upper surface of the wing trailing edge is one of the most promising layouts for the NFL positive effects to be realized and for the engine noise to be shielded by airframe elements. This paper presents experimental studies which were conducted on a laminar wing supplied with Krueger flap and for the wing lift performance to be improved.

The large-scale semi-span model of a regional aircraft with a small-swept wing in landing configuration was tested in TsAGI T‑128 wind tunnel in the wide range of the Reynolds number. The Krueger flap had eight different root inserts which covered a gap between the flap root and the fuselage; and the fuselage had one vortex generator.

The test results revealed that the root inserts & vortex generator application leads to flow separation diminishing in the wing root region, flow pattern improving and the layout lift performance increasing. The root inserts proved to be more efficient than the vortex generator, and the most effective of the former ones significantly augmented the magnitude of stalling incidence, reduced C_D and increased the maximum lift of the layout by ΔС_Lmax = 0.21. With this Krueger flap root insert being applied to the layout configuration the resulting lift magnitude (ΔС_Lmax = 2.78) proved to be not worse than the ones achieved on layouts with a common slat, though a Krueger flap used as high-lift device for wing leading edge is characterized by its lower efficiency.

Technologies of layer-by-layer laser sintering by the SLS-printing method are being increasingly employed in modern mechanical engineering and instrumentation. The gist of the technology consists in layer-by-layer sintering of powder materials (polyamides, plastics) using a laser beam.

Relatively low labor intensity and cost, as well as the achievable speed of products manufacturing allows applying this technology to aerodynamic models creation used for experimental testing of aerospace engineering products. However, the development of these technologies is hindered by the poor studies of the internal structure of the parts’ material.

There is an assumption based on the study of the outer layer of printed parts that a high porosity presents in the parts, caused by incomplete melting of all powder particles. This effect of incomplete sintering is visible on the outer surface. The problem lies in the fact that when sintering powder particles with a laser, neighboring, i.e. nearby particles that do not completely melt, forming a kind of a"relief" of the surface, are baked to the outer molten layer. It is obvious that such surface is not set in advance at the design stage, and the formed surface layer of stuck particles can be called undesirable. The external roughness control is especially up-to-date when creating aerodynamic models, since the external structure of the product surface may greatly affect the structure of the gas flow and the change in aerodynamic characteristics. The study of this layer and the roughness parameters will help designers to set and evaluate the necessary design requirements.

The research conduction is based on the results of a series of experiments performed with the EOS FORMIGA P110 SLS printing unit, in which laser is the main heat source with a power of 200 W-1 kW. The PA 2200 polymer was used for the samples production.

One of the problems while the research conducting is the impossibility of cutting samples or obtaining sections by mechanical or other methods without damaging the material structure. To solve it, an approach was adopted, according to which the operation of the installation was «emergently» terminated until the next layer of powder was applied. In other words, the newly obtained sample layer was not being filled with powder to form a new subsequent layer. It is possible to fulfill this by the printing emergency stoppage. Thus, it provided an opportunity to study the surface of the sample by the microscopy and measuring the roughness parameters of the formed surface. After processing the obtained images, the inference is being drawn that the internal structure is rather homogeneous and differs significantly from the outer layers of the samples. The outer layer of the products is of high level of roughness, which limits the possibility of their application in the field of aerodynamics. The article presents possible options for improving the surface layer of products.

The conclusion is made that the technology of selective laser sintering is utterly promising for creating aerodynamic models, provided that recommendations on improving characteristics of the outer surface roughness will be issued.

Vortex wakes studies after various aircraft is of both scientific and practical interest since the other aircraft entering the vortex wake is fraught with catastrophic consequences. The vortex wake is being characterized by the perturbed velocities field, as well as the shape and position in space. Tangential velocities W_τ are of the most interest, as long as their impact on the aircraft, got into the vortex wake, is quintessential.

The article considers approaches to the perturbed tangential speed determining at an arbitrary point in the vortex wake area. That task emerges at the aerodynamic characteristics determining of an aircraft, got into the vortex wake area, by the discrete vortex method, when the perturbed velocities determining is required at the point of the aircraft surface, where the «no-flow» condition is fulfilled and at the node points of the vortex sheet.

The problem of the perturbed tangential velocity computing at the point is being considered in the dimensionless form, i.e. coordinates, velocities and time are dimensionless. The way of the said dimensionless values obtaining is similar to the way employed while setting the problem of the aircraft aerodynamic characteristics determining by the discrete vortex method. The problem solution is being considered in the «frozen» field of the perturbed velocities approximation.

Three types of interpolation are under consideration. They are linear interpolation with mean value calculation of tangential speed at a point close to the given one; linear interpolation for determining the speed differential; non-linear second order interpolation. The authors disclose the advantages and disadvantages, and propose criteria determining selection of this or that interpolation. Finite-difference solution schemes of their differential representation were obtained for each type, and procedures, represented as algorithms and realized in the algorithm for aerodynamic characteristics computing of the aircraft entering the vortex wake by the discrete vortex method were proposed.

A comparative assessment of the computation results with the analytical solution was performed to assess the adequacy of the interpolations. The problem was being solved in a two-dimensional setting. Expressions for a pair of Renkin potential vortices modeling end vortices from the wing was selected as analytical expressions.

The presented work recognized that with the slight gradients of the perturbed tangential velocity changing, the linear interpolation should be used, while with substantial alterations of the velocity gradients and large velocity gradients values the second order nonlinear interpolation procedure should be used.

In modern supersonic aircraft, air intake units (AIU) exert a key effect on the entire power plant operation. The AIU main purpose is the gas flow supplying to the engine with minimal total pressure loss. The AIU development is a complex scientific and engineering task, which solution is being put into effect with computational and experimental methods.

The presented article considers methodological issues related to validation of the ANSYS Fluent software package (TsAGI license No. 501024), and provides a detailed description of the physical processes occurring in the AIN channel while throttling.

The authors performed numerical simulation of the flow in a flat supersonic AIU employing various turbulence models. The AIU geometry was borrowed from [1]. The oncoming flow parameters were as follows: Mach number М_∞= 2.41, angle of attack α = 10°, Reynolds number Re_х∞=5.07 × 10⁷ [1/m], total pressure P₀ = 540 кPa, total temperature T₀ = 305 К. Data obtained by computing the static pressure distribution on the AIU channel walls were being compared with the experimental results from [1]. The authors revealed that the best match of computed and experimental data on static pressure distribution of the AIU upper and lower walls are ensured by the two turbulence models, namely k-ωSST-CC (CC stands for compressibility correction) and Reynolds Stress Model.

The turbulence model k-ω SST—СС, considered in more detail in this article, allows reproducing a qualitative flow pattern with stationary separation zones, shock waves (including those from separation zones), rarefaction waves, and vorticity regions.

The two-dimensional calculation comparison with the three-dimensional one revealed that the Mach number fields were practically the same for both 3D- and 2D-flow in the AIU symmetry plane. An angular vortex is being formed near the AIU side wall, which drastically changes in the sections close to the wall the flow field and static pressure distribution on the AIU channel lower wall in the transverse direction compared to the flow in the AIU plane of symmetry.

To study the effect of backpressure being set at the channel outlet boundary on the flow field properties of the supersonic air intake, throttling of the model channel was being executed. The backpressure coefficient d was equal to d = P_back/Р_∞, where P_back is the static pressure set at the outlet boundary of the channel, and Р_∞ is the static pressure of the incident flow.

The studies revealed that with the opened throttle (d = 0) the flow in the AIU channel was supersonic. The local zone of the boundary layer separation originates herewith behind the break in its contour and a fan of rarefaction waves in the area of interaction of falling compression shock from the cowl with the AIU lower wall boundary layer.

With the backpressure coefficient of d = 5.5, an extensive separation region appears in the expanding (diffuser) part of the channel and a transition from supersonic to subsonic flow occurs.

At backpressure coefficient d = 8.5, a flow similar to a Mach disk is being formed at the AIU inlet: a direct shock wave is located in the central part of the inlet, and on top (near the shell) and below (near the compression wedge) two λ-shaped shocks are formed.

With the backpressure further increase (d ≥ 9.25), the direct shock wave is shiftiing forward and locating prior to the shell, while the upper λ-shaped of shock waves disappears, and the lower one moves forward, increasing in size.

Modern trends in aviation development lead to the emergence of new aircraft types and layouts, such as unmanned aircraft of the «flying wing» scheme and supersonic administrative aircraft. The layout limitations imposed by the adopted design decisions in terms of the modern aircraft appearance, lead in some situations to the non-standard ratios of the undercarriage base and the track. Changing proportions of the landing gear leads, in its turn, to the ground motion characteristics degradation. The existing techniques for the aircraft landing gear design do not imply the aircraft stability and controllability assessment in the process of design solutions selection for landing gear systems directly responsible for the ground motion control: the spectrum of these characteristics are evaluated already in the process of flight testing. Thus, the purpose of this work consists in proposing a technique for rational design solutions selecting in terms of ground motion control systems for the three-leg aircraft landing gear employing predictive modeling of runway going, which would allow identifying the aircraft negative specifics of controllability and stability, as well as eliminating them even prior to the aircraft creation.

The proposed approach is based on a predictive evaluation of stability characteristics, controllability and the range of operational limitations in ground motion, performed by mathematical modeling of the aircraft ground motion. To verify the results, experimental methods with the flight experiment data processing by means of mathematical statistics are used.

As the result of the suggested technique application for modernizing potential determining in terms of the landing gear of the aircraft being developed, it becomes possible not only to form the project decisions rational from the viewpoint of the weight efficiency (weigh reduction by the braking and steering-and-damping systems by several dozens of kilos), but obtain reliable estimation of the ground run characteristics with misalignment from the flight experiment by by 7–10% average as well.

The proposed methodology uses initial data in the scope of conceptual design and may be applied without significant modifications to the aircraft with the tricycle landing gear with nose support with takeoff weight not exceeding 40,000 kg and no more than two wheels on each landing gear leg. Predictive evaluation allows not only, if necessary, correcting the adopted design decisions at the stage of product development, which requires an order of magnitude less time and financial expenses than elimination of remarks after the flight tests, but estimating permissible operating conditions and restrictions on basing on various runways as well.

At present, upper stages are the main means for implementing a wide range of transport tasks of delivering payloads to various near-Earth orbits, as well as to the planets of the solar system. The «D» upper stage is the basic one in our country. In a number of cases, the two-staged upper stage, incorporating the «D» upper stage (the first stage) and «Frigate» as a second one, is proposed to be employed. The upper stage «Breese» is being employed of late as a part of the «Proton» launch vehicle to solve a number of transport tasks. A new oxygen-hydrogen upper stage is being planned to be developed as well. The fact that upper stages are equipped with liquid propellant rocket engines is associated with their higher thrust impulse compared to the solid propellant rocket motors. However, a very simple design and relatively high reliability make solid propellant rocket engines practically indispensable in solving a number of especially important transport tasks. A solid propellant engine, with which final acceleration up to the speed corresponding to the speed of movement on the final circular orbit, engine may be employed for bringing a spacecraft from a transitional orbit to the final circular one. It should be noted that such launching scheme application allows increasing the launched vehicle mass when employing the same space rocket (compared to the direct placement of a spacecraft into a circular orbit). It determines the said scheme relevance, since the obtained information allows to improving the spacecraft design quality as a whole and increasing the range of target tasks it solves. Computation of the charge geometric parameters is of special importance while the upper stage parameters selection. It is well known that flight tests allow confirming compliance of the design and other characteristics of subsystems with parameters and requirements for the spacecraft developing. Mathematical models are being corrected, the system settings are being refined by the flight tests results, and changes are being introduced in the design if necessary. However, flight test are rather costly experiment, and the number of such experiments is strictly limited. Thus, the more accurate mathematical model and its subsystems are, the less experimental launches will be required in the future.

The small upper stage (SUS) is a new type of technology being developed in the world since early 2010s to solve the problem of launch vehicles and payloads disproportionality, as well as provide peripheral launch services. Such spacecraft are launched as a part of a cluster launch, separate from the launch vehicle and, maneuvering independently, form the orbits micro and nano-satellites located on them. The article considers the layout specifics of upper stage running on the gaseous fuel components. The purpose of the article consists in searching for a new layout scheme of reduced size over the longitudinal axis, in which the tanks with the rectilinear axis are employed instead of toroid tanks. The gaseous Oxygen—Methane propellant propulsion system was selected for the SUS and the original «Sphere—Toroid» layout scheme was applied. The spherical tank of 87 liters capacity is being used for methane storage, and toroid tank of 158 liters capacity is for the oxygen storage. The main engine is inside the central orifice of the toroid tank. The possible schemes of the main configuration items, namely high-pressure tanks, placing were analyzed using geometrical model. Mathematical dependences expressed in a system of linear algebraic equations are obtained. The equations show the design parameters range that may be applied to design the new SUS layout scheme. Based on the analysis, both rational design option and the small upper stage layout scheme on its basis are proposed, which employs two pairs of spherical bottom cylindrical tanks. Compared to the original design, the new scheme is reduced by 40% along the longitudinal axis. The 20% reduction of cylinders volume new layout-may be compensated by a fuel pressure increase. The results of the study were applied while the new design-layout scheme of the «BOT» (Bauman Orbital Tractor) SUS development. The activities on the «BOT» SUS development are in full strength according to the ANO «Aeronet Center» technical requirements in the framework of the National Technological Initiative contest since 2020.

The current reality of the relationship between the manufacturer and operator is developing in the direction of ultra-fast response to the needs of the customer or the conditions of use of products, both in the civilian sector and in the military. The mature production structure does not allow such interaction being accomplished effectively. Application of the new model of such system of interaction between the designer, producer and operator «Mobile Technological Platform» is being proposed as a way of this structure changing.

The mobile technological platform (MTP) is a technological process brought out from the large industrial sites (enterprises) and employed under conditions of critically important rapid design and production of small-batch products.

The technologies used in MTP are essentially objects of the Industry 4.0 space artificial intelligence, Internet of Things, additive manufacturing, robotics, cloud storage, augmented reality, etc. Accordingly, this space, representing a digital production environment, forms a toolkit with ultra-fast computational technologies, self-developing and interconnected intellectual interactions in which decisions are made on the basis of self-learning data exchange systems in an automated mode.

The MTP existence is possible only in the space of Industry 4.0 using the appropriate tools and technologies. A space, in which quick product manufacturing according to ever varying requests of customer satisfaction, is possible due to the technology and production methods being used.

A participant in the MTP reality does not necessarily have to possess gigantic industrial resources for development, but must be integrated into a system of indicators determining his belonging to modern production processes through the use of appropriate technologies. If considering the MTP exclusively in the aircraft building industry, the wide geography of various purposes aircraft operation and aftersales servicing system (ASS), including maintenance and repair (M&R) should be accounted for.

More efficient ASS and M&R may be provided by integrating into this MTP system to produce parts locally, timely and in small quantities, needed all of a sudden planned according to changing operating conditions of the product. Thus, deploying rapid production (MTP) next to the operators, it is possible to form a more efficient and at the same time no less reliable supply system compared to the one that exists in various implementations today.

The mobile technology platform is a model of a new system of relations between the designer manufacturer and the operator in industry 4.0. Due to the rapid deployment of production of small—batch products, it will be possible to reduce the volume of warehousing, simplify the inventory management system of parts and components, eliminate long-distance transportation and total time spent on the supply chain, as well as save financial costs for paying for a large the number of specialists accompanying these processes.

The need for the MTP forming is obvious in conditions of competitive requirements of the rapidly changing environment. The place of the MTP existence is the space of Industry 4.0 with its technologies and tools.

There are requirements for the resource strength of aircraft in the aviation industry. It is necessary at present to perform costly and lengthy full-scale tests to meet the said requirements, virtual tests are being implemented hereupon. Just now, techniques for the resource computation require extra studying and incapable of replacing completely the full-scale tests.

The presented article is studying the accumulated damages in composite materials at the low-cycle cantilever bending. The full-scale experiment on cantilever bending of a structurally congruent flexible element is adduced. Residual deformations were being observed on the samples after the load was removed. The hypothesis on the residual deformations origination due to the cracks formation in the transversal layers of a composite material is being put forward. The majority of the known works on the said subject [1-9] consider behavior of the simplest plates. The presented article studies a flexible element from a composite material with complex physical and geometrical characteristics.

The object of the study is a structurally congruent flexible element from the composite material, which serves as an overlay between the tail part of the wing and lift devices. This element is being installed in preload and deformed, tracking the deviations of lift devices, being in constant contact with it. This allows hiding the gap between the tail section of the wing and the lift devices, creating thereby a continuous aerodynamic contour. In the works [10-11], various designs, in which a closed loop is implemented between the wing and the lift devices in various deflection modes, are presented.

The author solved the following tasks:

1. The cantilever bending calculation of a structurally congruent flexible element in a geometrically nonlinear formulation by the finite element method. The finite element model employs three-dimensional (volumetric) elements. The monolayer of the composite package was modeled into one element by the height. The model is fixed on the end face over all degrees of freedom. The stop was modelled in the form of а cylinder, to which a hard loading was applied in the form of the vertical displacement. All remaining degrees of freedom of the stop member were prohibited. A contact was applied between the structurally congruent flexible element and the stop.
2. Identification of theoretical calculations based by the full-scale experiment results.
3. Analytical calculation of the composite material package stiffness characteristics degradation. The process of micro-defects accumulation is assumed to be mostly a corollary, which is being regulated by physical thermodynamic laws, based on the entropy approach [12-15]. Micromechanical approach [16-19] is being used to associate the material damage with its appropriate properties and exact description of the degradation effect. A technique for the composite material properties degradation computing is described in [20].

The obtained results of the package stiffness characteristics degradation demonstrated a behavior similar to the full-scale experiment. Based on the obtained results, the inference may be drawn that the hypothesis on the cracks origination in the transversal layers of a composite material describes the material behavior

Based on the data obtained, it can be concluded that the hypothesis of the occurrence of cracks in the transversal layers of the composite material describes the behavior of the material quite acceptably.

The engine cut-off may occur in flight in a number of cases such as compressor surging, a bird ingress, or the crew error. In spite of the fact that the engine cut-off does not occur frequently, the possibility of its restarting in flight is one of the certification requirements, and comprehension of axial turbo-machines operation and characteristics at the extremely off-design modes gains more and more significant importance.

Following the engine cut-off in flight, rotation frequency of the engine rotors decreases to the steady-state value called the windmill rotation speed n_windmill. The compressor rotates herewith due solely to the impact pressure of the air incoming to the engine (the combustion chamber is off, the engine does not produce power). There is free windmilling, as well as locked wingmilling (the auto wingmilling at the rotor cranking by the starter). In the first case, which is being considered in the article, the engine shafts rotate with the speed depended on the flight Mach number, friction losses, angle of attack, the flow separation etc. In the second case, the initial shaft rotation is hampered since the ram air creates a torque not enough for the rotor cranking. At the modes where the speed is lower than at the rated idle compressor may run as a compressor (the energy is being transferred from the rotor to the liquid, which leads to the total pressure and temperature increase), a stirrer (total temperature rises in the compressor, but the total pressure falls), of a turbine (the temperature and pressure at the outlet are lower than at the inlet, and the power is being taken off the flow).

Numerical modeling in the 3D setting to obtain the subsonic ventilator characteristics at the windmill modes of was performed with the software complexes FlowVision 3.12.01 and NUMECA Fine Turbo 8.9.1 using the Spalart-Allmaras (SA) turbulence model. The simulation was performed for the following flight mode: the flight altitude of 11 000 m, and Mach number of 0.2–0.6.

The availability of the engine subassemblies characteristics is necessary to elaborate a technique for the parameters estimation of the turbojet engine at the windmill modes. As for now, there are no exact mathematical models allowing reliable description of these modes.

The purpose of this work consists in developing a technique for creating characteristics of a low-pressure compressor in windmill modes. Further research will be aimed at obtaining performances of turbines and other engine subassemblies as well as the development of the above-mentioned technique.

The technology of Ni-based granulated alloys has become widely spread for gas turbine engine disks manufacturing. This technology concedes the presence of internal defects of metallurgical nature. These defects may cause the crack growth under conditions of vacuum at cyclic loading. It is necessary to know the fatigue crack growth (FCG) rate in vacuum to assess the lifetime of the disks.

Special samples of cylindrical shape have been developed to solve this problem. A non-metallic defect is placed in the center of the working path of the samples, serving as a crack initiation source. The defect placement in the sample occurs while the powder filling into the sample capsule. Cyclic fatigue tests are being conducted until the sample destruction.

The two types of samples are used, namely vented and unvented. The vented sample differs by the presence of a through axial hole, which serves for the air supplying to the crack tip and the growth rate testing in the air. The unvented sample is necessary for testing the crack fatigue growth rate in vacuum.

The fractures of the samples are being examined by the fractography. The search for and measurement of fatigue striation spacing and the crack growth fronts reconstruction are performed at this stage. The presence of fatigue striations indicates a stable period of crack growth. The width of the fatigue striation spacing corresponds to the crack growth in one loading cycle, i.e. the fatigue growth rate. Thee fatigue growth rate is necessary for plotting a crack kinetic diagram. The crack growth rate is necessary to build a relationship between the rate and stress intensity factors (SIF), which is being computed after the crack shape reconstruction by the finite element method.

The sample and the defect diameters are being selected so that elastic stresses prevail in the section with the crack at the given maximum load. The ANSYS software was employed for determining optimal sizes of the sample with the finite element method.

Cyclical test of the special samples from the EP741NP alloy at the maximum loading of the cycle and the temperature of 400°C were conducted. Maximum load in the cycle ensures nominal stresses in the section with the crack, which is 0.58 of the proportionality limit at the beginning of the tests.

The average number of cycles to failure for unvented samples is 12.7 times greater than for vented ones. This indicates a significantly slower crack growth rate in vacuum.

Preliminary fractographic analysis of the specimens surface fractures were performed. The areas of fatigue striations location were identified. Fatigue striations are being observed almost throughout the entire crack growth area for the vented samples. This indicates that the crack growth occurred by the stable growth mechanism. For the unvented samples, fatigue striations are located in a narrow zone near the boundary of stable-tearing crack growth region. In this case, the crack growth in vacuum occurred mainly at a low rate, corresponding to the unstable growth mechanism.

Simulation of the fronts of the cracks started based on the obtained data to determine the values of the stresses intensity factors ranges and plotting kinetic diagram of the crack growth rate in vacuum and in the air.

EDB «Fakel» performs modernization of a SPT-70 type thrusters family, on which basis the TM-70 thrust modules (propulsion units), which were being employed at the «Yamal-100» and «Yamal-200» type spacecraft, and are in use at present at both «KazSat-2» and «EgyptSat» spacecraft.

The results of the research presented in the article were obtained during the EM1 engineering model of the SPT-70M thruster (hereinafter referred to as EM1) testing, which purpose consisted in studying the thrust and specific parameters, the thruster model lifetime characteristics and parameters of the plasma plume. These parameters studies were carried through in the course of the thruster model on Xenon and Krypton testing in the power range from 300 to 1500 W with discharging currents from 1.0 to 4.5 A and a voltage range from 150 to 500 V for various configurations of the discharge chamber channel exit part, which are simulating various lifetimes. These parameters studies were carried through in the course of the thruster model on Xenon and Krypton testing in the power range from 300 to 1500 W with discharging currents from 1.0 to 4.5 A and a voltage range from 150 to 500 V for various configurations of the discharge chamber channel exit part, which are simulating various lifetime durations.

Operation parameters fields of the thruster model, which may be employed while operation points parameters selection when operating both on Xenon and Krypton were determined by the results of the tests. Besides, the results of the EM1 direct and reduced endurance testing in the mode of the discharge power of 900 W (discharge current of 3.0 A) revealed the predictable total thrust impulse with Krypton would be no less than 1.0 MN, and 1.3 MN when operating on Xenon. The results of the tests on plasma plume parameters determining revealed that when operating in the mode with discharge power of 900 W (3.0 A/300 V) the divergence angle of the plasma plume while operation on Xenon was in the ranage from 35° to 37°, while with Krypton it was in the range from 48° to 50°.

The issue of liquid dispersion is of necessity in the design of air-jet engine combustion chamber and power plant. The article solved the solution of the two-phase gas-drop flow structure outflow from the cylindrical orifice to determine both velocity and gas consumption coefficients, as well as dispersed jet behavior. In the issues of the two-phase gas-drop flow forming with subsequent liquid phase dispersing (disintegration) in the combustion chamber of the jet-air engine, determining values of the velocity coefficients and phases consumption coefficients simplifies such devices designing for the intended result obtaining.

Preliminary design of spraying devices, such as mixers, injectors and devices involved in mixture formation is necessary when the air-jet engines combustion chambers designing. These devices operate on a two-phase working fluid, where the volume fraction of the gas phase concentration is equal or greater than the liquid concentration. Knowing the values of the velocity coefficients and phase flow rates allows solving the inverse problem. Thus, the purpose of this task consists in developing a technique for determining the velocity coefficients and phase flow rates.

The solution was performed by numerical methods employing the monodisperse heterogeneous model of two-phase flow. The flow of a two-phase flow through a cylindrical orifice of a 2 mm diameter in a jet nozzle with a given geometry was being simulated, where the nozzle length to diameter ratio equaled approximately to one. While simulation, the grid-independent solution was obtained with an error not exceeding 5%, which demonstrates the high degree of the computation accuracy. As the result of simulation, velocity coefficients and phase flow rates were determined. The obtained information on the liquid phase velocity coefficient and flow coefficient allows solving the inverse problem of the two-phase gas-drop flow dispersing as is shown in the additional one dimension computation of parameters. It is worth noting that the velocity coefficients exceed one, which is shown for the first time. Such values of quantities are being explained by physics of the complex interphase interaction. As far as the gas phase velocity at definite values of the initial parameters appears to be much higher than that of a liquid, which leads to extra acceleration, so that the velocity coefficient adopts a value greater than one.

The results obtained in this work may be applied not only in the combustion chambers of the air-jet engines, but in the design of other atomizing devices operating on a two-phase working body of a gas-drop structure as well.

To ensure the turbine blades operability and the required life cycle the blades cooling is implemented. Cooling channels should ensure intensive and uniform heat takeoff to ensure the necessary temperatures level at minimum hydraulic resistance. Traditionally, the blade leading edge zone and the middle of the blade airfoil zone are being marked out. The article being presented analyses the existing design solutions for both leading edge and the middle of the blade thermal exchange intensifying, defines the prototypes of design solutions for the reverse engineering with account for the existing patents. Cyclone cooling models, ensuring heat takeoff by forming stable vortex structures, were selected for the leading edge. Finned radial channels and vortex matrices were selected for the middle of the blade. Thermal-hydraulic models employing these design solutions were computed by the numerical simulation in a wide range of mode parameters and aspect ratios. Experimental studies were conducted using thermal imaging methods and calorimetry to confirm the obtained results. By the results of numerical studies, the temperature at the leading edge did not exceed 1270K, which confirms the necessary efficiency achieving. At the same time, the validation of the hydraulic model showed a discrepancy between physical and numerical simulations no more than 7%.

The solid propellant propulsion units depends designing entirely on the solid propellant selection. Solid propellant, especially a composite one, is of a complex composition, which includes oxidizer, binding propellant and an extra additives pack. Each component interaction determining and its burning kinetics studying are the first stage of the combustion process description. Thus, for example, the ammonium perchlorate is the most common oxidizer in solid fuel.

A number of N.E. Ermolin and Puduppakkam kinetic mechanisms were defined based on the experimental works of N.E. Ermolin and O.P. Korobeinichev on the composition determining of the stable individual substances of combustion products and the ammonium perchlorate decomposition by the distance from the combustion surface. N.E. Ermolin’s mechanism included 79 reactions, while Puduppakkam’s mechanism included 611 reactions and 105 substances.

The presented article considers the ammonium perchlorate combustion kinetic mechanisms, studies the temperature change and concentration of individual substances by the distance from the combustion surface. N.E. Ermolin’s mechanism (modified + 1 reaction 2NO = O₂ + N₂); Puduppakkam’s mechanism.

N.E. Ermolin’s boundary conditions (component composition of the decomposition products of the ammonium perchlorate condensed phase) were applied. Modeling was performed with the ANSYS CHEMKIN software in the one-dimensional PFR (Plug Flow Reactor) formulation.

The reduced mechanism (128 chemical reactions) was obtained based on the Pudupakkam mechanism reduction by the insignificant reactions determining. The results obtained by the less mechanism correspond satisfactory with the most detailed one, which includes reactions of octogen and hexogen combustion besides the ammonium perchlorate combustion. Thus, the mechanism. which may be applied in gas dynamics modeling was obtained.

The article presents the flame temperature profiles and individual stable chemical compounds concentrations at the pressure levels from 0.6 atm to 150 atm. The results obtained for the three mechanisms under study were compared by the end combustion products with the equilibrium calculation. N.E. Ermolin’s mechanism determines the list deviation from the equilibrium composition by the combustion products. Pudupakkam mechanism predicts the combustion products temperature much closer to thermodynamics.

This article considers techniques for the thermal state and radial displacements computing of the GTE turbine hull in the context of their application as a part of the mathematical model of the active clearance control system (ACCS) integrated into the electronic engine controller. The first technique is the of displacements computing based on the direct measurement of the hull temperature with the running engine. This scheme is realized on the CFM56-7B engine. It was revealed by the results of the analysis that it was quite enough to determine deformations of the external turbine hull only, and deformations caused by the pressure difference could be neglected, since were of no more than 3% of the temperature ones. These simplifications are being applied for the analysis of the rest techniques. The result of the hull displacements modeling results at the known temperature at the single point comparison with the results of axisymmetric modeling by the field verified by the temperature determined that the said technique ensures enough accuracy and computing speed. The second technique, namely displacements computing based on the temperature state, determined with the finite element method. Modeling results and technique verification are presented in the opened sources. They demonstrate that the error of the stationary thermal state modeling relative to the experimental data reaches 25% for the ground based gas turbine engine hull. Hence, this error will be much higher at the transient modes of the non-stationary computations. On the assumption of the performed analysis, the second technique does not satisfy the accuracy requirements to be integrated into the engine electronic controller. The third technique was developed by the authors, and based on the turbine hull displacements determining by the heat dissipation, calculated by the time constants dynamic computing. The hull temperature computing is performed by the two parameters, such as predicted stationary temperature and time constant, which are being computed at each time instant of the engine parameters registration in the automatic control system (ACS). Parameters computing is divided into the modes at turned-on and turned-off ACCS , since the time constants computing is based on the intrinsic heat transfer coefficients determining, and substantial heat exchange intensification occurs at the ACCS turning on, since the blow-off type changes from the smooth channel to the jet with the barrier. Stationary temperatures are being computed at the turned-off ACCS by the engine operation modes, while with the turned-on ACCS the air consumption and temperature in the blow-off collector are being accounted for additionally since these parameters are being used directly for the hull thermal state control and, hence, radial clearances. The calculated temperatures are compared with the data on the turbine housing thermometry on a full-size engine from the two test cycles. It is confirmed that this technique reliably reproduces the values and dynamics of temperature changes. Thus, it can be integrated into the ACC mathematical model. On the assumption of the accuracy and quick response, the first and the third of the considered methods can be applied to simulate the hull displacements on a real-time scale, and account for the control parameters of the ACCs, such as temperature and airflow in the blow manifold. Thus, they may be integrated into dynamic ACC, optimizing radial clearances in the turbine at all engine operating modes, and as the result, enhance its efficiency.

The up-to-date aircraft employ generally conventional engines, either piston or gas turbine, which operation efficiency onboard an aircraft has been studied quite well. Further efficiency enhancement of air transportation and aviation application for the new types of works requires implementation of new solutions and technologies, one of which may be a hybrid power plant (HPP).

The number of flights increasing predicted in the world in conjunction with the requirements of the Paris Agreement (2015) stipulates development of such solutions, which will allow significant reduction of hazardous emissions compared to the 2005 level. The aspect of no less significance is the fact that electric power units together with batteries are tenfold heavier, than turbofan engines commensurable by the power. As of today, the best way out of the current situation consists in the HPP application in aviation.

The purpose of the research is studying the HPP impact on the aircraft performance characteristics. Computations for the light class aircraft parameters optimization by the specially designed HPP integration into the aircraft structure were performed. Conditional HPP includes thermal engine, generator, electric motor, battery and for control, telemetry and information display systems. The layouts options of the two light aircraft in basic cases without the HPP and with the integrated on board HPP were studied, and analysis of basic performance characteristics was performed.

The projects of aircraft, such as EAG HERA, Zunum Aero ZA10, Heart Aerospace ES-30 and Faradair BEHA, originally designed with the HPP were studied. Four standard sizes of the aircraft most popular among the companies-operators were studied. The most popular aircraft models of similar passenger capacity were used for the comparison. As long as propeller should be a part for the power plant herewith, only the well-known aircraft with the HPP were employed for the configuration effectiveness comparison of the aircraft with the turboprop engine.

Inferences on the practicability of various standard size aircraft design for searching their weight-and-size parameters and performance characteristics were drawn by the results of the study. The necessity of the new aircraft projects development «from the scratch» for the most complete realization of the HPP potential as a part of the aircraft was substantiated as well.

The HPP components base, namely batteries, electric motors, generators etc., being employed presently, does not possess the parameters, which would ensure substantial supremacy of the aircraft with the HPP compared to the performance characteristics of the aircraft with conventional layout. However, other design aspects, such as hazardous emissions value the aircraft noise level, as well as the flight hour cost of the aircraft with the HPP, which should be less than this of the akin by size conventional regional turboprop aircraft of similar passenger capacity are essential for the aircraft with the HPP development.

The task of putting space rockets (space intended objects)into the target orbit is an extremely responsible add complicated task since exact delivery to the specified orbit with respect to the orbit parameters characteristic for each particular launching is required for the a satellite constellation deploying, or human scientific activities implementing in space onboard a spacecraft.

The perturbing factors impact at the stage of leading out affect adversely the object navigation by the satellite systems since it contributes to the distortion and loss of the navigation satellite signal. Autonomous object leading out has to be performed thereby. As long as this leading out is being performed autonomously by the inertial navigation systems (INS) readings, the total error of leading out would be read-out stipulated by the initial setting accuracy. The coordinates of the object being launched are known herewith with geodetic accuracy, and initial velocities are negligibly small. Thus, the initial error will be formed by the initial orientation error of the inertial measuring unit, including the triad of accelerometers and gyroscopes relative to a certain geographic basis.

The object of research in this work is the algorithm for initial setting of the platform class inertial navigation system for the objects of various classes and applications.

The purpose of the study consists in elaboration of the algorithm for the goniometrical initial setting of the platform inertial navigation system based on application of mathematical programing methods, and noises effect estimation of inertial sensors (gyroscopes and accelerometers) on response time and accuracy of the setting.

The authors proposed a fundamentally new approach to the algorithm elaboration for the platform INS initial setting to reduce its response time and enhance its accuracy. Simulation modeling of the proposed algorithm, as well estimation and analysis of the noises effect on its efficiency were performed.

A large number of aviation accidents, referred to as Loss-of-Control in flight (LOC-I), led to the keen interest to motion cueing fidelity while on-ground simulation of civil aircraft stall cases.

Analysis of flight records reveals that any of stall cases may be divided into two stages, namely before and after the stall, which differ by motion cues amplitude-frequency content. This difference is determined by the difference in types of piloting task since before the stall the piloting task relates to the stabilization type task, when a pilot operates in the closed loop, while after the stall it relates to the maneuver-type task, when a pilot operates in the open loop.

According to the approaches developed at TsAGI, the types of motion cues distortions differ depending on piloting task as well. In stall simulating, it is the stage of stall recovering which causes the main concern, since the false cues arising in the course of this maneuver simulation can considerably distort the real aircraft behavior.

Thus, the following assumption on the required volume of the motion cues for the stall simulation was put forward:

1) the stage of aircraft approaching the stall needs motion cues reproduction in full;

2) the stage of aircraft stall recovering needs motion cues limited by those typical of buffeting in heave and sway.

To prove the assumption, special experiments were conducted on the TsAGI flight simulator, in which four experienced flight-test pilots participated.

Three different combinations of motion cues were considered in the course of experiments:

— no motion cues were reproduced (immovable bench);

— only buffeting was reproduced;

— motion cues were reproduced in full volume (full flight simulation).

The model of a hypothetical line-haul aircraft (SUPRA) was employed.

The results of experiments had proved the assumption concerning the required volume of motion cues for aircraft stall simulation and led to the following conclusions:

1. With account for the amplitude-frequency content of motion cues, as well as motion cues role in piloting and types of motion cues distortions, the simulation process of aircraft stall can be divided into two stages: before aircraft stall (stall approaching) and after aircraft stall (stall recovering).
2. Based on the experiments, the article demonstrates that the «stall approaching» stage requires full flight simulation, buffeting included; «stall recovering» stage requires reproduction of buffeting only.
3. For the «stall recovering» stage, the full reproduction of motion cues was assessed by the pilots as the least acceptable one among the considered options.

The obtained results may be employed in aviation design bureaus and research institutes, as well as in aviation crew training centers.

The article deals with the up-to-date problem of the aircraft engine parts service life increasing. One of the problem solutions is protective coatings application by methods of plasma flows condensation from low-temperature plasma. The presented work performed the analysis of surface protective layers creation for target cathodes for gas-discharging systems employed in practice and the ways of their preparation. The authors proposed employing high-entropic target cathodes obtained by spark plasma sintering to generate plasma in electric arc and magnetron sources. Technological process of spark plasma sintering high-entropic target cathodes for protective coatings synthesis incorporating five stages was developed. These stages are powder composition preparation, pilot experiment, sintering of high-entropic target cathodes, post-sintering process.

Technological process of high-entropy cathode-target synthesis of the Al₂₀-Ti₂₀-Zr₁₅-V₁₅-Cr₁₅-Nb₁₅ system composition with reference to the KCE-FCT-HP-D25-SD facility was realized, with account for specific aspects related to the features of the of a multi-component mixed plasma flow generation ensuring, being generated by vacuum-arc and magnetron discharge in the vapor high-entropy target cathode for uniform coating deposition of the given composition. The samples with the diameter of 20 mm and height of 3 mm were obtained to perform preliminary studies, assess the powder composition elements compatibility. The samples of 80 mm diameter and 8 mm height were obtained for studying physical and mechanical properties and assessing the target cathodes performance characteristics.

The results of energy dispersive microanalysis of target cathode samples obtained by spark plasma sintering of powder composition Al−Ti−Zr−V−Cr−Nb revealed the presence of all components of the initial powder composition, which confirms the possibility of obtaining high-entropy target cathodes by the said method.

Regularities of sintering technological modes effect (temperature, extrusion pressure, holding time at maximum temperature reaching and heating rate) on the target cathodes properties and structure were determined. Dependences of the physical and mechanical properties of high-entropic cathodes on the technological modes of the spark plasma sintering process were revealed. With sintering temperature increasing from 600 to 1000°C, an increase in hardness and electrical conductivity is being observed, and further sintering temperature increase does not lead to a significant change in the controlled parameters, and the values of hardness herewith correlate with the values of electrical conductivity. The sintering temperature effect on the structure of high-entropy sintered target cathodes samples was determined in the course of the performed experimental study. The article demonstrates that the structure of the samples sintered at the higher temperatures is characterized by higher homogeneity.

Modes of spark plasma sintering of the Al20-Ti20-Zr15-V15-Cr15-Nb15 system composition of the high-entropic cathode-target with reference to KCE-FCT-H-HP-D25-SD installation were determined based of the conducted studies.

The results of the conducted experiments confirmed the perspective of spark plasma sintering application for producing high-entropy target cathodes for the protective coatings synthesis on the aircraft engine parts, but further studies on the the geometry and configuration of the powder particles effect on the composition and properties of sintered high-entropy target cathodes are required.

The article presents the results of tribology properties experimental studies, particularly the dependence of wear over the back surface of the cutting edge on the cutting path length and the miller durability period for various coatings. Full-scale tests of innovative composite nanostructured wear-resistant coatings were being performed at high-speed milling of the VT3-1 and VT6 titanium alloys widely employed in the parts and units of the state-oof-the-art gas-turbine engines (GTE) and rocket-space technology. The coated millers wear was being measured with the «Carl Zeiss Stereo Discovery V12» stationary motorized stereo microscope with telecommunication capability and 3 mp «Zeiss Axiocam 503 Color» video camera based visualization system. A series of experiments on studying contact processes while milling, such as temperature and cutting force components were conducted with modern equipment and facilities. While the electro-physical parameters control and registration, accompanying blade machining process, the natural thermocouples method with mercury current collector application and PC recording was employed for the cutting temperature determining. The temperature tests results while milling by wear resistant coatings revealed the average 20% reduction in cutting temperature with the «nACRo3+TiB2» coating compared to the others, and the VT3-1 machining was less heat intensive compared to the VT6 one. The cutting forces components were being determined with the «Kisler» dynamometer complex consisted of the 9253B23 model three-component dynamometer, amplifier with ADC and a PC. The cutting edge force loading of the miller with the «nACRo3+TiB2» coating has lower value compared to the others and is 25-30%. The results of the conducted studies revealed effectiveness increase of the titanium alloys high-speed milling by application of the state-of-the-art nano-structured wear resistant more than twofold. Express-evaluation of the machined surface quality indices, such as roughness and hardening, was being extra conducted while these tests. The results of the performed express-evaluations revealed the required machined surface quality indices improvement, which is important while the GTE structural elements such as flanges, disks, compressor rotor shafts etc. blade cutting machining.

During these tests, express-evaluation of the quality indicators of the machined surface (roughness and naklep of the machined surface) was additionally carried out. The results of the express-evaluations showed the improvement of the required quality indicators of the machined surface, which is important for blade cutting machining of structural elements of GTE parts (flanges, disks, compressor rotor shafts, etc.).

Selective Laser Melting (SLM) represents an additive manufacturing technology meant for metal powders alloyage by a high-power laser. Powder materials application ensures more uniform chemical composition over the product section and the absence of zonal segregation. Titanium alloy powders application for selective laser alloyage is a prospective trend in additive manufacturing.

The possibility of the parts production with configuration of any complexity, simultaneous growth of several samples, high material utilization coefficient and products prototyping simplification are among the SLM technology benefits. The presence of residual porosity in the part being manufactured, limitation of materials being used and laser radiation sources, as well as the size of the products being manufactured are related to the said technology drawbacks.

The purpose of the article consists in studying the Ti6Al4V alloy microstructure forming while manufacturing the gas turbine engine compressor impeller by the selective laser alloyage method.

The samples for studying were fabricated with the installation for the SLM 280 HL metal powder selective laser alloyage installation. They were synthesized both horizontally and vertically relative to the building-up platform. The microstructure studying after etching was performed with the METAM LV-31 metallographic microscope. Electron-microscopic analysis of the samples and original powder substance was conducted with the TESCAN Vega SB scanning electron microscope. Chemical composition of the original powder material was determined with the INCAx-Act Energy Dispesive Spectrology (EDS) device. EDS analysis revealed that the original titanium alloy powder chemical composition corresponds to the standard with an excess of aluminum and silicon content. The electron-microscopic analysis results revealed the spherical shape of the powder particles peculiar to the method of obtaining the dispersed molten. Metallographic analysis of demonstrated that after the SLM the samples had a microstructure of the α-phase plates, and the β-phase was not noticed. The electron microscopic analysis of samples fractures after the tensile testing revealed the mixed character of the fracture mechanism. The non-uniform fracture contains the sections corresponding to various stages of destruction.

The ultimate strength of the samples after the SEA is 1117 MPa. It is more than for the material obtained by stamping. Relative elongation of the vertical sample is 3.08 percent. Relative elongation of the horizontal sample is 6.11 percent, which is less than for the stamped one.

An aero-physical experiment conducted in wind tunnels (WT) is not aimed only at applied research for determining aerodynamic characteristics of the aircraft models of the widest purpose and their power plants elements. It is also aimed at fundamental research in the field of fluid, gas and plasma mechanics, physics of condensed state, applied electrodynamics and detailed testing of new physico-mathematical models being developed based on the molecular-kinetic theory, as well as modern computer codes for computational fluid dynamics. Among all types of the WT, a special place is occupied by continuously operating variable density wind tunnels, which simultaneously create subsonic and supersonic dry airflows in a wide range of Mach and Reynolds numbers, as the most universal experi-mental setups in this area. There are only very few of these worldwide. Typical examples include the NASA John H. Glenn 8×6 supersonic wind tunnel, ONERA S2MA, DLR TWG, TsAGI T-128, and European ETW cryogenic tunnel. A similar wind tunnel is available within the walls of Moscow Aviation Institute (National Research University). This is the T2 multi-mode variable-density wind tunnel, which first draft design was completed at MAI back in 1947. This wind tunnel advent is largely associat-ed with prominent Soviet scientists and engineers such as B.N. Yuriev, G.V. Kamenkov, B.I. Mindrov, K.M. Drobyazgo. This large and unique MAIN experimental facility, put into practice 1959, allowed Soviet designers create advanced aeronautical, rocket prototypes and spacecraft technology, and produce them at the highest world level. Management of MAI aerodynamic laboratory, named after N.E. Zhukovsky, consisted of industrial units T-1 and T-2, was performed through the USSR Ministry of Aviation Industry, which allowed staffing it with highly qualified engineering and technical personnel. At its core, scientific research performed in the laboratory was mainly of experimental nature, dealing with various aircraft models at their design stage, production or flight tests. As the result of the long-term activities based on self-financing, strong ties were established between the laboratory and aviation industry companies, the Min-istry of General Mechanical Engineering, the Ministry of Mechanical Engineering, the Ministry of the Shipbuilding Industry, the Ministry of Electronics and others from its conception and well into 1991. Since the moment of its establishing, the average annual production of scientific and technical re-ports from the laboratory was about 25–30 reports annually. The scientific and technical staff of the laboratory was awarded the State Prize and the 25-th MAI anniversary Prize for their deep scientific contribution to the development of aircraft aerodynamics.

In today’s economic conditions, the volume of scientific re-search conducted in MAI experimental laboratory of Aircraft Aerodynamics Department has significantly decreased. Along with this, the opinion that wind tunnels will be substituted in the nearest future by the mathematical models and Computational Fluid Dynamics software packages is being increasingly introduced today into the mass consciousness. This article comprehensively proves that wind tunnels cannot be replaced by their digital counterparts in principle. A key matter is the fact that along the decade from 2010 to 2020, TsAGI has undergone a deep modernization of the entire complex of its wind tunnels. The similar wind tunnels upgrades were performed in the West. According to authors opinion, in the distant future the physical experiment is most likely will be harmoniously combined with the computational one, including promising educational, scientific and applied platforms for aerodynamic design. To solve the problem of sustaining a ceaseless aerodynamic laboratory operation, the authors referred to the experience of leading domestic and foreign industry and academic research institutes. The volume of scientific research can be increased by expanding the range of tasks to be solved, which is possible only after a MAI wind tunnels modernization, measuring systems and methods improvement, both conducting experiments and managing the laboratory with a deeply scientific approach in the field of management.

Evaluation of the aerodynamic loads distribution along the wingspan for the aircraft with an extra-high aspect ratio wing is an up-to-date task, and in the future, it will allow developing measures for the negative impact reduction of wing deformations while the flights in a turbulent atmosphere. Another problem for the said aerodynamic layouts with the extra-high aspect ratio wing is the flight at a crosswind, which may lead, among other things, to the «Dutch step» phenomenon occurrence. The presented article considered the crosswind impact on the load distribution along the wing, including running pulling propellers. The aerodynamic loading distribution along the takeoff and landing mechanization elements and control organs on the wing was obtained.

Numerical studies of the side slip angle (crosswind) effect on the aero-dynamic characteristics of the aircraft model with an extra-high aspect ratio wing with propellers running at the wing ends were performed with a program based on the Reynolds-averaged Navier—Stokes equations solving. The computations were conducted with incoming flow velocity of V = 50 m/s and Reynolds number of Re = 0.35 × 10⁶ at non-deflected wing mechanization of d = 0 to compare the computational and experiment results within the range of side slip angles b from 0 to 20°, as well as in takeoff position with d = 15°. The article shows that with the side wind the flow bevels increase and local angle of attack on the wing changes. Computational studies revealed the interference of the running engines at the side wind has a significant impact on the aircraft model flow-around, its aerodynamic characteristics and hinge moments of the wing mechanization. The lift coefficient dis-tribution along the wingspan shows that the lift force reduction at the slide angle increase is being strongly affected only by the left wing console, while at the windward right console of the wing the lifting force drop at the slide angle increase is just local in the area of the propeller slipstreams blow-around. The slide angle change increases, in general, the hinge moment of the external aileron only at the right windward wing console especially with the propeller blowing. This is being stipulated by the fact that the slide angle affects the flow bevels in the propeller blow-around area, and only the windward console gets into it. The article shows that at the blow-around of the undeflected mechanization and the slide angle increase the flow bevels of the external aileron are large enough, while with deflected wing mechanization the they decrease, and the pressure on the lower part of the wing increases.

Computational studies revealed that the interference of running propellers in a crosswind significantly affects the aircraft model flow-around and its aerodynamic characteristics. With a crosswind, the flow bevels are increasing, and the local angle of attack on the wing is changing. With the side slip angle increasing this effect strengthens on the windward side, and the interference zone with the propeller spreads along the span of the windward side of the wing, where the hinge moments of the wing mechanization increase hereupon even in the retracted position.

The article presents the study of the aircraft mutual effect while subsonic for-mation flight at minimum distances with account for the height and interval variation of the wingman relative to the lead aircraft. The issue on aerodynamic characteristics changing in the formation flight is up-to-date for many years. The trail aircraft movement in the wake vortex is known to lead to the wingman aerodynamic characteristics changing. The wake vortex impact on the aircraft and its subsequent disturbed motion depend on the whole number of factors such as aircraft performance characteristics and a flight mode of both wingman and lead aircraft, spatial position of the aircraft relative to each other and the state of the surrounding atmosphere. For this reason, the problems of flight safety while moving in a wake vortex after the lead aircraft emerge. However, the flight at minimum distances between the aircraft formation flight may incur the aerodynamic quality growth as well, which will allow increasing the flight range and duration. Determining optimal position between the lead aircraft and a wingman will allow meeting the controversy of the requirements in the formation flight. This is especially up-to-date for the highly automated flight control systems, which are being installed including the unmanned aerial vehicles. The study was being conducted using Solid Works, Numeca Hexpress and Ansys Fluent software packages. The article presents the dependencies of the parameters being studied, namely the lift and drag coefficients, the pitch, roll and yaw moments on the interval and height of the wingman relative to the «flying wing» type lead aircraft. The authors show that computational methods application for the aerodynamic characteristics determining allows supplementing the results of experimental modeling in wind tunnels. Thus, with the interval between the aircraft axes of symmetry decreasing, the lift force increment increases and reaches its maximum at Δz/l = 0.9. The increment of the moment coefficients changing changes slightly herewith. Further, while further transversal spacing decrease, the sharp changing of coefficients of aerodynamic forces and moments starts due to approaching the lead aircraft vortex wake. While the trailing aircraft movement in the vortex wake (which symmetry axis coincides with the spatial position of the vortex core Δz/l = 0.5) the moment coefficients and drag force are maximum, and the lifting force increment is negative.

Airframe units mating is a vital stage of the aircraft final assem-bly, which should ensure high accuracy of the aircraft external aerodynamic surfaces. As of today, two foreign-made systems of civil aircraft jig-free assembly are being operated in Russia: technological process of detachable wing part and fuselage of the SSJ-100NEW is being realized in Komsomolsk-on-Amur with the BROTJE automated bench, and the German «ThyssenKrupp» production line is being employed for the MC-21 aircraft in Irkutsk. As for domestic equipment, production line for the Il76MD-90A aircraft automated assembly is functioning in Ulianovsk at the «Aviasatar-SP» aircraft building plant.

In the presented article, the authors consider technological process of detachable wing part to the aircraft fuselage mating employing an automated bench. This con-tributes to reduction of the number of personnel in charge of the routine technological operations of material production.

Process automation is being implied as the industrial robotics applica-tion, much as the numerical control machine tools were employed as the production automation tools. With account for the fact that robotics operate on the assumption of the electronic information, managing programs are being written, products electronic models are being developed and processes are being modeled for it. The article gives an account of the method for the product compliance with design documentation validating, and describes the employed rigging necessary for the bench operation and ensuring high accuracy of measurements. The basic structure of the technological process of the wing detachable part mating with fuselage is presented in the form of the table with the basic operations description. The process of the automated bench with measuring bases is described. The authors propose to employ the considered operation principle of the automated bench while creating a mobile version of the mating bench. The article gives an account of the requirements to the mobile bench structure and its basic technical characteristics.

Application of the automated bench mobile version will allow increasing the volume of released production with the possibility of producing various aircraft configurations at the single production site.

The article deals with the problem of design an aircraft guided missile (AGM) with a solid propellant rocket engine (SPRE), and performs comparative analysis of the of the AGM motion along the trajectory in the multiple activation mode. The authors demonstrate that the engine may be regulated to a certain degree by the thrust cutoff at certain time in-stants. This is being implemented with the specially designed dampers. To realize the passive flight seg-ment, the passive flight segment parameters duration, selected from the flight range maximizing condition, is being introduced to the design parameters vector. Particularly, alongside with the AGM flight range increase, the passive segments inclusion into the flight trajectory may lead to the AGM flight altitude, its opera-tion time and other optimality criteria losses.

In essence, the AGM trajectory consisting of both active (with running engine) and passive (with dead engine) segments is being determined by the AGM motion mode. This mode, alongside with the other design parameters and the SPRE parameters, constitutes the design solution vector, which is being selected by the vector criterion.

The final design solution selection is being performed employing convolution with variable weight coefficients. Substantiation of this application of convolution is being derived from the principle of the complex technical system rational organizing. The gain from the passive segments application herewith is 10% in range on average. The additive principle with optimal weighting coefficients allows selecting design solutions without involving any information hypotheses. In case of the preferences presence of the project designer, correction of the obtained solution is being performed in accordance with this system of preferences.

As of now, space exploration is accompanied by the de-crease in mass and an increase in the number of artificial satellites To create and employ a huge number of satellites, it is necessary to change the principles of development, production and operation of space technology. The article proposes using the modular principle of universal space platforms (USP) building. As applied to the design, it consists in assembling the product from standard multi-functional modules of complete readiness. The modular principle application in the USP design consists in creating a certain standard communication interface, somewhat erector kit, which allows assembling an apparatus for the concrete mission.

As evidence of the feasibility and effectiveness of such an approach application, the USP demonstrator layout has been developed, which includes three modules (de-pending on its purpose): a service systems module, a payload module and a reusable upper stage module. The reusable upper stage module is presented in two options for delivering the payload of various masses of 250 kg and of 500 kg. The authors propose four versions of the upper stage demonstrator for different missions: a spacecraft to deliver equipment to a body with a low gravity field, a spacecraft for landing and takeoff from the surface of small planets of the solar system or satellites of small bodies, satellites for LEO and GEO. The article considered an option of electrolytic propulsion unit application as a part of USP demonstrator. The presented design omnitude lies primarily in the possibility for creating options of satellites, and spacecraft at the request of the customer at short notice by employing unified design solutions based on the USP. The module principle application will allow significant reduction in time of the spacecraft development and manufacturing.

As of now, the interest to the functioning processes studying of the aircraft with ramjet power plants (RPP) has increased, which is being explained by the en-hancing possibilities of such systems operation analysis based the new qualitative tool of numerical mod-eling of complex gas-dynamic processes.

The fulfilled study performed the search of the said configurations based on the approach in the form of comparative analysis of merits and demerits of various structural schemes of the airflow duct of the air intake devices, applied on the certain class of aircraft and in various speed ranges.

The basic methods of the study are the methods for the gas-dynamic processes numerical modelling in the air duct of the air intake unit. The article presents the results of various structural schemes design. Variations of throttle characteristics as well as coefficients of extra aerodynamic drag, introduced by the air intake unit installing, were obtained for the developed schemes with the CFD modeling methods. Based on the energy efficiency analysis of the developed schemes, the most effective air intake units were selected.

The air intake unit operation effectiveness determines at large the RPP energy parameters. Besides the boundary layer, the disturbances caused by the aerodynamic surfaces and angles of attack relate as well to the number of factors capable of reducing the air intake unit gas-dynamic perfection.

It is found that the presence of angles of attack leads to a significant reduction of the air intake unit characteristics. Various degrees of sensitivity to external disturbances and angles of attack were obtained for the considered configurations as well.

The authors analyzed the airfoil and rudders effect on the air intake unit char-acteristics. It was found that vortex formation in the wing trace and shock waves, as well as unsteady perturbations led to the vortex trace forming, turbulized the flow and reduced its energy, affecting the air intake unit operation. As the result, rational position of the wings relative to the air intake unit has been selected.

To eliminate the said drawbacks, a modification of the internal compression air intake design has been proposed. The technical task of the proposed layout scheme consists in ensuring maximum possible throttle performance in the range of angles of attack of the aircraft from 0 to 5 degrees with minimum extra aerodynamic drag.

As the result, a method for the air intake unit functional specifics evalu-ating, which allows priority solutions selecting by their configurations, with account for the aircraft flight specifics and limitations imposed on it, has been created. A theoretical foundation for the SPP implementation on aircraft with specific flight conditions (the dominant energy-passive section of the trajectory) and stringent mass-and-size limitations in the design was created thereby.

Despite the rapid development of mathematical modeling methods, the stage of experimental work-out in the design and creation of aerospace vehicles still plays a particularly important role. On the one hand, it allows exploring the most difficult modes for mathematical modeling, and on the other hand, it allows validating numerical methods. The products functioning under conditions of a possible turbulent mode creates certain difficulties for the correct con-duct of experiments on aerodynamic models t with a view to further correct transfer of the obtained data to a full-scale object.

In earlier studies of one of the authors, the issues related to the possibility of accounting for additional entropy production due to the stochastic perturbations excitation while a turbulent flow mode implementation were considered in details. This allowed modifying the Navier-Stokes (NSE) equations by posing them in an expanded phase space. In this case, the left part of the NSE, i.e. the full derivative in time, is being supplemented by a term characterizing the change in velocity when an additional «stochastic» variable changes. Inclusion of an additional term characterized by entropy production (which is always non-negative) in the equations allows, in particular, to account for the irreversibility of physical processes in time in cases where this production is non-zero. Based on this approach, both «laminar» and «turbulent» solutions for the Hagen-Poiseuille problem [8], the plane Couette problem [7] and the plane Poiseuille problem [6] were analytically obtained for large values of the Reynolds number.

This article shows that the «modified» Navier-Stokes equations allow obtaining extra similarity criteria, which, in fact, are analogs of the well-known similarity criteria ob-tained for «classical» NSE, but wielding a multiscale character: starting from the scales of a viscous boundary layer and ending with a macroscale flow.

Multiscale similarity criteria can be useful for more complete and accurate ex-perimental and numerical modeling of liquid and gas flow, in particular, when creating new products of aerospace equipment operated under conditions of possible turbulent mode. This approach will allow selecting the «right» size of the surface roughness and appropriate technological approaches when creating aerodynamic models for experimental research.

The article considered the issues of creating aerodynamic models with a controlled surface roughness size for conducting multiscale hydro- and aerodynamic experiments. It is noted that the most promising methods for such models creating can be technologies based on the SLS printing [13-17].

With the advent of microelectronics and implementation of microcontrollers in aircraft fire protecting systems, application of automation units based on electronic-digital components instead of relay components is becoming increasingly up-to-date. This allows interacting with aircraft equipment via the code line and perform more effectively the functions of detecting and employing fire alarms in aircraft compartments. Besides, application of digital systems provides new opportunities for creating fire extinguishing system control algorithms and generating signal information logic, which cannot be implemented when arranging blocks with relay components, resulting in a large mass and the need to arrange complexly branched computational circuits for the imple-mentation of algorithmic and computed sequences.

The authors determined that most aircraft employed balloon-type fire extinguishing systems. At the same time, fire extinguishers are being initially divided into sequential fire extinguishing queues consisting of several cylinders, regardless of the compartment in which the fire eventuated.

The purpose of this study consisted in developing a new ap-proach to the fire extinguishing queues selection in terms of creating the most effective conditions for extinguishing a fire in each fire hazardous compartment of an aircraft.

The numerical method was applied to compute hydraulic losses and find the average pressures created at the outlets from the orifices of the spray collectors in the fire hazard-ous compartments of the aircraft.

While further scientific research, a fundamentally new, combinatorial ap-proach to the fire extinguishing queues selection was developed, which allows increasing the fire protection system efficiency in the event of a fire in the aircraft compartments, and meets the latest trends in the development of digital fire-fighting automation systems. An algorithm for the fire extinguishing queues forming has been developed within the framework of the combinatorial approach, which is of adaptive character, where the combination of fire extinguishers can be changed with account for possible leaks in the cylinders.

Designing the state-of-the-art aircraft requires new structural solutions and applica-tion of fundamentally new materials and technological processes for their manufactiring.

The aircraft hardware compartment was selected as the object of research. Temperature indicators on the aircraft hull are directly related with its speed. Thus, among all design tasks the authors chose the task of temperature level reduction inside the onboard hardware compartment to ensure its uninterruptible operation. Mathematical modeling of intensive aerodynamic heating impact on the hardware part of the aircraft hull performed by the authors allowed obtaining steady state vapues of its hardware temperature at the level exceeding the marginal allowable value. The article regards a method for the aircraft hardware compartment temperature reduction employing aircraft onboard hardware passive thermal protection means based on the new class materials application.

A discrete fiber material based on aluminum oxide and quartz fiber (aerogel) is under study as an internal thermal protective coating (TPC). The article considers the hardware compartment structure with account for internal TPC (aerogel) and external TPC (composite erosion-resistant material), and presents the temperature values obtained for various TPC types, which ensure the necessary temperature level inside the compartment.

Analysis of the results of mathematical modeling, performed by the authors, of the intensive aerodynamic heating impact on the aircraft reveals the effectiveness of the aerogel application. This material allowed the aircraft hardware temperature reduction to 86°C. The stress-strain state modeling confirmed the strength of the load-bearing aircraft compartment structure involving external composite material (CM). The article demonstrates that fundamentally new material of the internal TPC, namely aerogel, leads to the onboard hardware temperature reduction by 4 °C without the external TPC application, and by 12 °C with the CM application as the external TPC. Despite the heat-protective layer reduction of the internal TPC, introduction of the external TPC from the erosion-resistant CM leads not only to the temperature level reductioin inside the aircraft compartment, realizing the temperature operative range, but it reduces the temperature on the titanium hull as well, which allows varying the material, both hull and external CM thickness for the hull mass reduction.

According to paragraph «33.65 Surge and Stall Characteris-tics» of the Aviation Regulations, part 33 (Aircraft Engine Airworthiness Standards) the fol-lowing is stated. «It is required that while engine operation according to the Op-eration Manual the engine startup, power or thrust changing, power or thrust forcing, limit non-uniformity of the air flow at the engine inlet should not cause surging or flow separation, which might lead to the flame breaking, destruction of the structure, temperature rise or breaking the possibility of recovering power or thrust in any point of the operation modes range». In this regard, the issues of simulating the required type and level of the flow non-uniformity at the inlet prior to the engine while bench testing to confirm sufficient margin of the gas-dynamic stability of the compressor and relatively trifle impact on the basic parameters of the engine vibration-strength characteristics, become up-to-date and of practical meaningfulness.

The field of velocities and pressures at the inlet of the engine as a part of the power plant is being determined by the aircraft flight conditions (altitude and flight Mach number, angles of attack and sideslip, etc.), the engine operation mode and the air intake design. In general, this field is non-uniform, and the flow prior to the engine is non-stationary. Thus, its imitation while bench testing of a gas turbine engine is a difficult technical task.

The following basic requirements are being imposed on the simulators of a non-uniform flow prior to the engine:

the values of the total pressure coefficient s_in, averaged over the channel section prior to the engine and the values of the parameters (criteria) of the non-uniform flow (the circumferential non-uniformity criteria of the total pressure and the total pressure pulsations intensity) behind the power plant air intake and simulator, should be the same;

the simulator should generate a total pressure field prior to the en-gine, similar to the real field of total pressures behind the air intake at equal values of the reduced air mass flow through the engine.

The main reasons causing the uneven flow prior to the engine are associ-ated with the local separation zones occurrence in the air intake duct, accompanied by the total pres-sure loss and an increase in the flow turbulence.

Various methods and technical means are employed in practice to reproduce characteristics of the uneven flow at the engine inlet. However, the main ones in practice are as follows:

hydraulic grids with different density installing in the bench inlet device for the engine testing, which are employed in case of simulating a low level intensity of the pulsations full pressure prior to the engine (less than 2%);

installing interceptors of various configurations in the bench inlet device prior to the engine inlet, which allow simulating a high level of flow non-uniformity, including the of pulsations intensity.

This article presents the main results of simulating the flow non-uniformity basic parameters at the engine inlet by two interceptor-segments with different values of the relative flow shading area depending on the value of the reduced mass flow density q(l).

The article presents also the experimentally obtained correction factors flow non-uniformity impact on the tested engine basic parameters.

On the assumption of permanently growing product complexity, stipu-lated by the requirement toughening to the functional characteristics, additive technologies play the key role in industrialization of the new production methods of aviation engineering. However, the ex-isting barriers caused by limitations of the additive production technologies and powders properties, ca-pabilities of hardware and software are hindering active reengineering of the products to additive production technologies.

Technological process developing for parts manufacturing by the selective la-ser fusion (SLF) method is a multifactorial and multivariate task. Decision-making on the SLF technology implementing for manufacturing hot part of the industrial gas turbine engine installations is based on the following criteria definition. They are productivity; level of detail, i.e. the possibil-ity of minute fragments manufacturing; plotting accuracy; work-out labor intensity; geometric parameters stability and reproducibility; reliability and endurance of the additive production; main assemblages lifespan prior to replacement or refurbishment.

However, achieving the above said criteria is being ensured by technological processes optimizing based on parametric and structural methods.

The purpose of this work consists in developing algorithms for creating a «smart» structure (configuration) of the «Burner device» assembly unit when the SLF technology design for the VG159 heat-resistant alloy metal powder. This algorithm for the aircraft engi-neering products design will allow obtaining «on the first try» a workpiece according to the required quality parameters of the SLF technological process. The article presents the recommended sequence of work in the design of gas turbine engines assembly unit, being manufactured according to the SLF technological process.

The SLF process being developed for the of the «Burner device» assembly unit manufacturing is based on a complex «product-material-process-properties» digital twin, which allows ensuring a «free» geometry with the assembly unit accuracy provision, in contrast to the conventional methods of designing based on the me-chanical properties optimization. The algorithm specificity consists in accounting for the values of residual stresses, the magnitude and direction of deformations in the designed workpiece ob-tained by the SLF technology by modeling the SLF process and predicting the level of residual stresses. The simulation result is corrected 3D model of the workpiece with the geometry, which, after manufacturing according to the SLF, heat treatment, separation from the structuring platform and removal of supports, will ensure minimum deviation of the shape, size and location of surfaces from the specified values, i.e. the «smart» design of the assembly unit.

The article demonstrates that the existing mathematical models for gas turbine en-gines (GTE) thermodynamic computing do not allow accurate simulation of the start-up and windmill operation modes. The reason lies in the impropriety of compressors and turbines characteristics, set in traditional form of their representation for these elements parameters determining under conditions close to the quiescent state when the pressure ratio is close to 1.0.

A method for calculating the start-up and windmill modes of aircraft gas turbine engines using thermodynamic mathematical models is demonstrated. The said method is based on employing the performance maps of compressors and turbines in a transformed form. The authors proposed to use the compressor torque normalized to the total inlet pressure instead of the traditional compressor efficiency. At the compressor operating modes, at which the air pressure is being increased, this parameter is unambiguously derived from the values of pressure ratio, normalized flow rate, adiabatic efficiency and normalized rotation frequency of the compressor. For this reason, the reduced torque is a criterion parameter that ensures the similarity of compressor operating modes. The similar conditions are being ensured for the characteristics of turbines, where, the torque at the turbine shaft normalized to the total pressure in the nozzle assembly throat is proposed to be used as well instead of the turbine efficiency. For the «near-zero» modes, turbines and compressor characteristics recomputed for employing normalized torque instead of efficiency, may be obtained by either extrapolation or computing using state-of-the-art 3D CFD methods.

This method operability for the steady-state modes is demonstrated on the examples of computing the windmill and motoring modes for a two-shaft turbojet engine. The article shows the possibility of a nonlinear mathematical model employing to determine max-imum amount of power that can be taken from the windmilling engine shaft. It was demonstrated as well that it was preferable to swing the high-pressure rotor by the starter, since it allows obtaining noticeably greater air consumption and pressure at the combustion chamber inlet with the same delivered power.

The example of the two-shaft turbojet start-up on the ground with the starter, swinging the high-pressure rotor, as well as in the flight conditions by the wind milling without starter employing, was given for the non-steady-state operating modes.

It is noted in conclusion that the developed method application allows significantly expanding the application scope of element-by-element thermodynamic nonlinear mathe-matical models of gas turbine engines for solving real-life problems in the field of engine build-ing.

Traditional method based on definition of the most rational engine and its units project parameters proceeding from intended purpose and features of operation is usually used when the new aviation gas turbine engine (GTE) creation in practice. Besides, another method, supposing max-imum possible use of some engine units and elements from its predecessor already manufactured and checked up in operation, is widely used.

The rest of engine units of the new engine are designed anew, most of-ten at higher technical and/or technological level. In this case, it is possible to expect occurrence of the new engine (usually of the same generation) in a shorter time and at a lower cost.

In practice, preserved engine units are usually considered the high-pressure compressor (HPC), as the most labor-intensive in designing and operational development GTE unit, or engine core, consisting of the HPC, the combustion chamber and the high-pressure turbine (HPT).

For the successful realization of this method when creating a new en-gine (or families of engines) of the required thrust or power rate, it is necessary, that initial «engine-donor» has a core with the necessary parameters, first of all, core size parameter and compressor pressure ratio.

Because such a condition is not always executable, the problem of creation new engine core, capable of meeting the thrust and power requirements of a number of engines for various purpose constructed on the basis of this unified core, is set.

The results of parametrical research of three the most widespread schemes engines variants, having same base engine core, are presented in the article.

As an example, options for replacing some foreign engines, applied on domestic aircrafts, with new alternative engines, constructed on the basis of this unified core, are shown.

The present paper is devoted to the modification of the structural model of pilot behavior, which allows to evaluate the influence of the characteristics and type of inceptor on the properties of the pilot-aircraft system. To this purpose, a series of experiments was performed employing MAI’s ground-based simulator to determine the regularities of the pilot-aircraft system when using a center and side stick, different types of control signals sent to the flight control system (proportional to the displacement and proportional to the forces applied), as well as different characteristics of the inceptor (stiffness and damping). Two configurations from the HAVE PIO database were selected as the controlled element dynamics, one corresponding to Level I flying qualities, the other corresponding to Level III. In the experiments, the operators performed a compensatory pitch angle tracking task.

Studies have shown that in terms of piloting accuracy, the optimum value of center stick stiffness for both types of control signals and controlled element dynamics corre-sponds to 10 N/cm. With a side stick, the optimum stiffness takes the value 20 N/cm. In all cases considered, the best piloting accuracy is achieved with minimal damping.

Studies have also shown that the piloting accuracy for a Level I flying qualities configuration is 1.5 and 1.6 times better when using a center and side stick, respectively, com-pared to displacement sensing control in a Level I configuration. For a Level III configuration, this transition is accompanied by an improvement in accuracy by 1.25 and 1.3 times. In addition to piloting accuracy, force sensing control reduces the equivalent phase delay introduced by the pilot and improves other parameters of the pilot-aircraft sys-tem.

Overall, the transition from a traditional DSC-type center stick to a FSC-type side stick results in a 2.3-fold improvement in piloting accuracy when controlling configurations which belong to the first level of flying qualities and a 1.9-fold im-provement when controlling configurations which belong to the third level.

Based on the results obtained, a modification of the structural model of pilot behavior was proposed. This model takes into account the models of visual cue perception and the neuromuscular system, inceptor dynamics, correction of information received from proprioceptive feedback which closes the «neuromuscular system + inceptor» system. When a command signal is proportional to the displacement, the inceptor model is in the direct loop of this system, and when the pilot’s force input is used as such a signal, the inceptor model moves into the feedback loop. Different models and parameters of neuromuscular dynamics are used in the study of the effects of the center and side stick. In addition, the model takes into account noise in the perception of visual and kinesthetic information. The spectral density of the latter is proportional to the variance of inceptor displacement. Due to the results of experimental studies having shown that the noise components of signals are inversely proportional to the stiffness and directly proportional to the damping, the parameters of inceptor stiffness and damping are introduced into the model of this spectral density.

The parameters determining the correction of visual and proprioceptive cues are chosen by minimizing a functional consisting of the sum of the error signal variance and a summand proportional to the variance of the forces applied to the inceptor and its stiffness. This summand was added to match the optimal values of stiffness obtained in experiments and in mathematical modeling.

The use of the proposed model makes it possible to obtain results close to the results of experimental studies, as well as to assess the influence of inceptor characteristics, inceptor type, and the type of control signal on the characteristics of the pilot-aircraft system.

The cruise flight is the main phase of the flight of long-haul air-craft, which mainly determines the effectiveness of entire flight. The cruise flight effectiveness depends on the selected flight mode. Typical cruise modes include maximum range mode, maximum cruising mode as well as compromise modes. Compromise modes selection are being performed by the flight costs indicator of the flight at the given range. This indicator employing is possible only at known values of the fuel cost and cost indicator, which are the uncertainty source in the tasks of the long-haul aircraft effectiveness studying.

The article proposes considering the problem of compromise modes selection under uncertainty conditions for a certain range, employing flight costs indicator presented in analytical form. The search for the compromise modes is being performed on a set of modes, limited by the maximum range mode and maximum cruising mode, which we will call the set of optimal modes. Partial criteria of the effectiveness indicator such as fuel consumption and flight speed are deter-mined on such set. Analytical effectiveness indicator is the sum of normalized partial criteria with weight coefficients that are the parameters of the task. The flight mode selection under uncertainty conditions is being performed in the minimax problem setting using the analytical weight coefficients. The weight coefficient in this indicator can be interpreted two-fold, which allows considering the problem of compromise mode selection in two formulations, such as operational and trajectory. In the operational formulation of the problem, the weight coefficient is the normalized value of the cost index and does not change along the flight path. In the trajectory formulation of the problem, the weight coefficient is a measure of relative importance between fuel consumption and flight time and can vary along the flight path.

The studies of the compromise conditions achieving in the trajectory formulation of the problem for various values of the cruise range allowed identifying the optimal range, different from the maximum range, for which the compromise mode can be considered optimal. The optimal range obtained by the trajectory method is an objective criterion for change flight level at the compromise flight modes. The said criterion allows objectively selecting the point of transition to another flight level and improve thereby the operational performance of the entire flight (such as the required flight fuel margin and the flight endurance). The optimal range in the operational formula-tion of the problem is the maximum range.

The article presents an example of cruise flight optimization under the flight conditions at different flight levels, which results demonstrate the ability to reduce the required fuel and flight endurance compared to this flight implementation in the maximum flight range, maximal cruis-ing and operational compromise flight mode. The effect of the flight altitude and the payload (the aircraft weight at the cruise flight termination) on the optimal range value in comparison with the maximum range was established as well. The results of the cruise flight effectiveness studying, obtained by the trajectory method, may be useful for the development of a flight manual and flight paths optimization problem of long-haul aircrafts. The object of research is the Il-96-300 long-haul aircraft.

Spacecraft orientation determining and angular motion control are among the crucial tasks being solved in the space-rocket engineering area. Measuring modules, including gyroscopes based on microelectromechanical systems (MEMS), are employed to solve this problem in the nano-class spacecraft. However, MEMS gyroscopes belong to the type of sensors of relatively medium and low measurement accuracy. Besides, space factors, such as cosmic radiation, solar activity, aerodynamic forces, or temperature gradients, lead to the sensor reading drift over time, depending on its stability. The sensors of inertial navigation systems are calibrated thereby automatically in flight. Despite this fact, the pre-flight ground calibration, which is necessary to be performed to confirm all sensors integrated into the system correspond to the minimum requirements placed on the space mission, occupies an important place. There are special turntables on the market for gyroscopes calibration, which set predefined turns at certain velocities and orientations, though they are rather costly. As of now, robot manipulators are widespread all over the world, and they are most often employed to perform certain motions with high precision. In this sense, robot manipulator represents a possible option for solving this issue. Thus, the article proposes reliable technique for bench calibration employing robot manipulator to eliminate systematic errors of commercial MEMS gyroscopes. The main idea of this technique is based on using the wrist of robot manipulator as a high-precision rotary device. The author proposes a modified six-position method in the form of the sequence of rotations to perform laboratory calibration. This technique allows determining systematic errors of the sensor output signals, particularly bias, scale factor and the axes non-orthogonality. Bench tests form a set of experimental data for subsequent processing by the calibration algorithms, and allow identifying all systematic errors and assess the degree of applicability of this bench. For this technique testing, a Strapdown Inertial Navigation System was manufactured, and bench tests were performed, which revealed the possibility of employing a robot manipulator as a calibration instrument. The features of the results of processing experimental measurement data during tests of commercial gyroscopes using this technique are described. The application of the developed approach leads to a five-fold reduction of the error by five times compared to to raw measurements.

The purpose of the presented work consists in developing automated technique for the product structure manufacturability (PSM) assessment based on its 3D model. The following hypothesis was put forward for its solving: formalized PSM of the product may be realized employing initial data from the product design documentation (DD) in the form of the product electronic model (PEM). This will allow obtaining the output data in the form of technological recommendations on preproduction recommendations to the production engineer such as tools selection, typical technological process (TTP), as well as providing quantitative and qualitative indicators of the PSM analysis.

Based on the said hypothesis and purpose the following tasks were put for-ward:

1. Developing concept of the technique for the PSM analysis in the form of the flowchart.
2. Selecting the part and its definitely significant structural elements (SE).
3. Performing information formalization necessary for the part manufacturability assessment.
4. Developing aggregative concept of the PSM analysis technique in the form of the flowchart.

The study consisted in analysis of the conventional methods for the manufacturability assessment, i.e. how this assessment is being realized at the modern industry. In other words, to analyze the technique for analysis performing and reveal its problems. Based on the problem and industry and data digitalization (the product electronic model is a design doc-ument) the concept of the technique for the manufacturability assessment of the machine-building product structure base on its 3D model was put forward.

This result may be implemented at any state-of-the-art machine-building enterprise while preproduction of a new product.

The following inference can be drawn. Traditional method for the PSM as-sessment has become obsolete and does not match the digital industry criteria. Besides, qualified production engineer is required to perform the assessment of the part structure manufacturability. The need for such specialist would be eliminated with the application of the new method for the assessing the structure manufacturability of the part of the machine building production. The said method will significantly re-duce preproduction period of a new product, minimizing human factor, as well as drastically simplify the work of the production engineer at decision-making on either manufacturability or non-manufacturability of the product at the given type of production.

The article regards the problem of external multi-factor impact on the basic parts of machines, particularly the impact of corrosion on the operational and technical characteristics of plunger pumps, and the stages of its solution.

In the aviation area, hydraulic systems, to which exclusive requirements on the structural reliability are being imposed, are susceptible to the greatest corrosion impact. Thus, rational method selection for the basic parts protection of aviation products is the up-to-date task, and requires corresponding technique development. This technique is considered on the example of the protection method selection of the part of the aviation hydraulic plunger pump.

The authors performed the analysis of technical requirements to plunger pumps and revealed dominating factors affecting the basic parts wear, as well as considered the ways of deposition of protecting metal and non-metal coatings. The article presents the developed technique for protective corrosion resistant coatings deposition on the parts of the «Case» type. The said technique implementation was performed on the example of structural and technological criteria assessment on each of proposed coating deposition method for the basic part. By the results of this technique, the method of polymer powder coatings deposition is preferable.

The proposed method is being widely employed in production, in particular, while the aircraft and machine-building products manufacturing. This proves the proposed tech-nique fidelity. For the reason that the group of structural and technological criteria is being considered, the said method can be employed not only for the basic parts of the aviation plunger pumps, but for the variety of other products of aviation industry, including gas turbine engine blades, rotary engine stators and other products of aircraft engineering.

The complex goal of the research consisted in reducing the superplastic moulding temperature of the VT14 alloy by alloying with β-stabilizers with a high diffusion coefficient, which effectively lower the temperature of the β→α phase transformation to achieve optimal phase ratio under low temperature conditions. Fe and Ni wielding the high dif-fusion coefficient in the β-phase were selected as alloying elements. As the result, the effect of alloying by the β-stabilizers of various concentration on the microstructure evolu-tion, indicators of superplasticity, as well as on mechanical properties at the indoor temperature are being studied.

The alloys under study were additionally alloyed by the minor additive of boron (up to 0.1 wt.%) to enhance mechanical and technological characteristics by the fine crushing of the grain structure in the presented of the dispersed ToB particles while the molten crystallization, as well as while thermo-mechanical treatment. The results of the microstructure evolution analysis while annealing in the temperature range of 625–850 °C revealed that the Fe content growing from 0.5 to 2% and Ni from 0.5 to 1.8% led to the β-phase volume fraction growth, and, hence, shifting of the optimal temperature range to the lower temperatures.

Analysis of the uniaxial tension tests results with 1´10^—3 s^—1 velocity in the temperature range of 625–775 °C revealed that due to the β-transus temperature reduction and dif-fusion coefficient increase, the increase in the Fe and Ni content significantly improved the superplasticity indicators. Superplastic deformation of the modified alloys was characterized by the high values of the strain rate sensitivity coefficient m = 0.45–0.5 for the alloys with Fe and m = 0.5–0.6 for the alloys with Ni, as well as with high relative elongation of 500–1000% at the twice as low flow stress compared to the alloy without addi-tives. It was demonstrated as well that alloying by 0.9% Ni and 0.5% Fe was quite enough for ensuring high relative elongations and indicator m = 0.5 at the deformation temperatures of 700–775 °C, and temperature reduction to 625 °C, required concentration increasing of Ni up to 1.8% and Fe up to 2%.

Alloying allowed increasing the level of mechanical properties at the indoor temperature after the superplastic deformation at the temperature of 775 °C with the rate of 1´10^—3 s^—1. The increase of the β-stabilizers content contributed to the strength margin and yield stress margin by 100–250 MP, as well as minor plasticity reduction relative to the industrial VT14 alloy.

As the result, optimal alloys contents, wielding increased strength properties at minor plasticity reduction, and characterized by the high superplastic indicators under conditions of the lower temperatures (625—700°C) were proposed. These alloys are Ti-4Al-3Mo-1V-0,9Ni-0,1B and Ti-4Al-3Mo-1V-0,1B-0,5Fe-0,1B.

The question area of the work tackles with one of the aircraft building state-of-the-art problems, namely shock and bullet damages diagnostics of the structures from polymer compo-site materials for subsequent selection of the technique for their refurbishment. Visible damages on the surface do not give a comprehensive idea of the destruction inside the structure. Instrumental control methods application allows studying both character and sizes of the damage to define the type and scope of the repair job in case of the damage confirmation.

The possibilities of the X-ray computer tomography for the composite struc-tures studying were considered in the course of the work. Both shock and bullet damages were inflicted to the samples for the operative refurbishment technology work-out. Non-destructive control was performed with the X-ray computer tomography (CT) to determine the character and sizes of the damages.

The studies of the internal structure of the samples was being per-formed with various X-ray computer tomographs. The presented work studied the character of bullet damages of the two helicopter composite structures, namely the fragment of the steering rotor blade and a part of the experimental spring of the skid landing gear. A fragment of the helicop-ter rotor blade was subjected to the shock damages.

Computer tomography allowed considering the layer of interest in details, scaling the pattern, and determining the defects sizes and their location in the structure. The sizes of the visually registered dent on the surface were much smaller than the fracture zone inside the sample. The fibers destruction, fibers damage with stratification and stratification without fibers damage are being observed. All these damages alter the structure of the material and increase the porosity in the damage zone, which reduces the mechanical characteristics. The size of the shock damage depends on the characteristics of the material and the impact energy. The inference can be drawn from the shock damages analysis that even low impact energies on a honeycomb structure lead to the dent forming in the skin and a honeycomb filler crumpling with partial destruction. At the higher impact energies the skin bursting and honeycomb filler destruction occurs. The issue of performing non-destructive control of the damaged zones after refurbishment and its quality assessment is an up-to-date one.

Among various environmental impacts on the aircraft, icing is the most dangerous one. Despite the almost century-old history of this problem research, accounting for and elimination of icing is still an actual task.

The purpose of the presented numerical study consists in researching the impact of the airscrew interference and a straight wing of a high aspect ratio of a solar battery powered aircraft on the aerodynamic characteristics and hinge moments values of the wing-flap system deflections under icing conditions.

Numerical study of the airscrew, installed at the wing tip of a high aspect ratio wing, impact on aerodynamic characteristics and hinge moments of the wing-flap system, deflected to the takeoff position (= 15°), was performed by the program based on the Reynolds-averaged Navier-Stokes equations solving, at the aircraft under the icing conditions. Calculated study was performed with the aircraft, which aerodynamic layout was realized by the classical scheme with cantilever high-set wing with the aspect ratio of = 23.4. Engine nacelles were placed on the wingtip. The airscrews rotation frequency was of N = 15000 rpm. The airscrews rotating direction corresponds to the vortex sheet folding from the wing tip.

Numerical studies were conducted without airscrews and with operating two-bladed airscrews, both without aircraft icing and with it. Initially the ice shapes without blow-off and with the blow-off by the airscrew were calculated. The calculation revealed that the presence of a rotating airscrew had a great impact on the ice growth formation on the wing. The ice thickness on the wing without airscrew is almost the same over the entire surface, while a high barrier of horn-shaped ice is being added to the existing one on the wing beside the tip of the airscrew blade.

Further, aerodynamic characteristics were calculated, and a hinge moment was obtained for each part of deflected wing-flap system. These calculations were performed at the angles of attack of −5°15° with the Mach number of М = 0.15 and Reynolds number of Re = 0.35·106.

Calculation results revealed that aircraft bearing surfaces icing reduced maximum lift force and increase pitching moment on pitch-up, as well as contributes to the aircraft drag increase, especially with the airscrews blow-off beyond stall angles.

The airscrew running under conditions of icing leads to the detachable zone size increase, which grows with the angle of attack increase.

The article demonstrates that icing may decrease the hinge moment of the wing-flap system. This occurs as a consequence of the overgrown ice forming such a shape below the surface of the deflected wing-flap system, which decreases pressure on its windward side. The value of the total force, acting on the deflected wing-flap system, decreases herewith, and the center of pressure of the deflected control surface is being shifted closer to the rotation axis.

The wing of modern aircraft is one of the objects of control. Depending on the purpose, type, class and aerodynamic layout of the aircraft, it is equipped with various means of mechanization — devices and systems designed to control aerodynamic characteristics without changing the angular position of the aircraft in the stream. Mechanization is used at all stages of flight: during takeoff, climb, cruising, level change, descent, landing approach, movement along the glide path, landing and landing run. In order to increase the lift force for a supersonic passenger aircraft in takeoff mode, a blown flaps control device has been developed, that is, near the trailing edge of the wing, it is carried out by imparting additional kinetic energy to the retarded flow by blowing off the boundary layer with a gas jet. This article presents the results of an experimental study of the influence of the jet momentum coefficient and the flap deflection angle on the lift coefficient СL and the drag coefficient СD. Using the PIV (Particle Image Velocimetry) observation system, the flap blowing control mechanism was studied. The lift measurement results show that is too large to effectively increase СL when air circulation control is not applied, while the effective can be increased after air circulation control is applied. The maximum lift force of the model wing can be obtained with a small and = 30°, and with an increase in , the maximum point of the lift force of the model gradually shifts back at = 40°. The results of the PIV experiment show that in the absence of airflow control on the surface of the flaps, a clear flow separation is observed, and after turning on the flow control at = 30°, the flow reattachment can be completed with a small . With an increase in , the flow velocity on the upper surface of the wing further increases; when is less than 0.04 and = 40°, the flow joins, at which СL and СD increase; when is greater than 0.04, the flow joins, at which СL increases, and СD decreases, the lift-to-drag ratio K increases, and the aerodynamic characteristics improve significantly.

A spacecraft touchdown on the surface of the planets and their satellites is one of the crucial stages of the flight, since the surfaces of the planets are insufficiently studied, and the spacecraft motion kinematic parameters may vary over a wide range.

Landing gear, which should ensure touchdown with acceptable overloads and stable spacecraft position on the surface, is employed while touchdown for the spacecraft dampening.

The landing gear consists of three or four supports, depending on the power scheme of the landing mechanism.

The spacecraft dampening is being realized by the energy absorbers, placed in the landing mechanism shock absorbers. A rod, honeycomb, pipe and tape (flat rod), which absorb the energy of a spacecraft while touchdown due to the plastic deformation, are being applied as one-time operation energy absorbers.

Accounting for the landing gear elasticity will allow concrete determining of the dynamic loads and the spacecraft stability area while touchdown, which is specially important while touchdown on the comets or small-gravity satellites.

When solving the touchdown dynamics problem, the equations of motion of the landing gear supports are being used, with account for the elastic deformation of the structure.

Accounting for the elastic deformation energy accumulated in the landing devices elements and the places of their attachment to the body will allow determining the dynamic loads on the device and structural elements, as well as correctly determining the area of the spacecraft stability to overturning.

The presence of the developed shock absorber structure with the energy absorber, such as tape; kinematic scheme of the landing gear support, as well as an algorithm for the problem of a spacecraft touchdown solution allows selecting basic design parameters of the landing gear with account for limitations. These parameters will ensure safe and stable touchdown without overturning of a spacecraft with low inertial characteristics.

When determining the spacecraft touchdown stability area, preliminary design parameters of the supports selected according to statistics are used, and a series of calculations are being performed on the touchdown dynamics, varying by the factors listed above.

An energy absorber is a tape employed in the design of the landing gear supports, which ensures cushioning of a spacecraft with low inertial characteristics when touchdown on either planets or their satellites, as well as asteroids, with account for restrictions.

Computing cases of overloads and a spacecraft stability were determined by the landing gear design parameters varying obtained from the spacecraft touchdown dynamics problem solution. If the landing mechanism layout of the spacecraft with low inertial characteristic does not allow placing the landing gear support with the recommended requirement to the base to the center of masses ratio, then it is advisable to employ clamping emgines.

For instance, when the Rosetta spacecraft landed on the Churyumov-Gerasimenko comet, a landing device containing a harpoon and clamping engines was applied.

The work deals with the study and modification of the Automated Design Dialogue System (ARDIS) for the subsonic passenger aircraft design to create an extra possibility of calculating the cargo aircraft characteristics.

The ARDIS is meant for performing calculations at the initial stage of the design (concept selecting, requirements formulating and draft proposal developing), when the state of the project is marked by many uncertainties, and it is necessary to consider a large number of options and perform their parametric studies.

The methodological basis for the ARDIS modification is application of computational algorithms for cargo aircraft characteristics, including ramp ones, and their software implementation.

Proceeding to the ARDIS software package modifying, a part of modules, subjected to the changes, was separated out, while the other part of the modules, which are not planned for modification at this stage, but their application may affect the result of characteristics computing of the transport category aircraft, remained unchanged. Such modules as Geometry, Aerodynamics, Power Plant, Flight Performance and Mass relate to these kind of modules. They were studied with description and detailed block-diagrams compiling.

The ARDIS modification assumes not only direct editing of the program source code, but also introduction of new variations of the features that will allow the ARDIS to switch algorithmic branches for calculating characteristics of both cargo and passenger aircraft.

The new types of transport aircraft introduction to ARDIS allowed modifying the program code responsible for computing characteristics corresponding to these types. Specifics of new types of aircraft affect the change in the mass of the aircraft and, first of all, the change in the mass of the fuselage. Algorithms for computing the weight of the longitudinal framing, windows, doors, hatches, sealing and the weight of the floor of the passenger cabin or cargo compartment have undergone partial modification. Algorithms for computing all other characteristics unique to the cargo-type aircraft have been redeveloped.

Computations in the modified system of computer-aided design (ARDIS) were performed on the example of the prospective transport aircraft in passenger version, and a cargo aircraft. As the result, the aircraft specifications were obtained, which were verified with the prospect characteristics.

The article considers application of the wire electric arc technology for additive forming in manufacturing a structural element of a manned spacecraft. The structural element represents a typical bracket for attaching equipment to the spacecraft body. For manufacturing by the additive growing technique, the source element was optimized for the printing technology capabilities and limitations. After optimization, a manufacturing process was formed, in which course the electric arc parameters were changed at various stages of printing. Manufacturing was being performed employing the AMg6 aluminum wire of a 1.2 mm diameter. As the result, the obtained structural element was tested for harmonic and static impacts. The purpose of the tests consisted in determining damping, strength and stiffness properties of the structure. The obtained test results were being compared with the computed finite element model. According to the analysis, under harmonic action, the frequency and form of oscillations of the first tone coincide with the computed ones (66 and 69 Hz, respectively). The damping coefficient in determining the amplitude-frequency characteristic and vibration impact was 2% and 2.5%, respectively, which allows accepting the obtained value as the damping coefficient of the material itself (in the first approximation). During static tests, the structural element collapsed under a load of 15300 N, the displacement herewith reached the value of 19.6 mm. The destruction occurred at the place of attachment to the power floor along the thinnest part of the bracket base. The fracture emergence is of a similar character with the maximum stresses occurrence in the finite element model, except of the primary fracture in the fusion zone of the material layers. Analysis of the material microstructure revealed the presence of gas pores from 20 to 500 microns. The chemical composition corresponds to the AMg6 alloy, though without manganese (Mn) on its surface. The results of the study revealed that the manufactured structural element withstood operational loads, and the printing technology was possible and efficient to be employed with certain assumptions for the studies of additive technologies under conditions of low gravity at the orbital space station.

The article deals with the issue of thermal insulation properties improving of the unmanned aerial vehicle (UAV) operating under conditions of extreme temperatures. Basic conditions for the structural layout elaboration ensuring high indices of the UAV thermal insulation without application of thermal insulation protective means are being considered.

For this issue solving, theoretical studies of the unsteady heat exchange of the UAV unit were being performed with account for various diameters and sections under the impact of extremely low and high temperatures. The said studies solutions were being performed by a numerical method, namely a finite difference method.

The results of the theoretical study point out that the UAV high thermal insulation indicators require that its structural layout ensuring the gas interlayer between the hull and the unit constituent parts. For the small diameters being considered in the article (less than 500 mm), the average thickness of the gas interlayer under the impact of the extremely low temperatures should be 8 mm, and for large diameters (500 mm and more) it should be 12 mm and higher. Insuring high indicators of the UAV thermal insulation under the impact of extremely high temperatures require the average thickness of the gas interlayer between the hull and the unit constituent parts by the following dependence: it should be 20, 30, 60 and 120 mm for the diameters of respectively 100, 200, 300 and 500 mm.

The said changes in the UAV unit structure allow its thermal isolation indicators sixfold improving without application of thermal insulation protective means.

The article presents a model for assessing the presence and hazard degree of defects based on acoustic emission invariants. The analysis of acoustic emission criteria of destruction is presented from the viewpoint of their application possibility while diagnosing power elements of aircraft structures in real time to determine the degree of deformation and the hazard of structure defects. As of now, a number of the destruction criteria of the controlled object (CO) material has been developed, based on the various approaches to the acoustic-emission (AE) information processingn and analyzing. However, there is no criterion allowing diagnosing the cracks being developed with high probability. This is being associated with the CO material inhomogeneity and the presence of the residual stresses. Under these conditions, to increase fidelity of the acoustic-emission method of non-destructive control and defining the degree of the defects hazard, it is rational to develop and employ destruction criteria based on statistic invariant dependencies that characterize pulse flows of the acoustic emission. The article presents the results of studying the acoustic emission parameters relationship with the early stages of destruction specifics of a layered composite, as well as iron and aluminum alloys employed in the design of power elements of the aircraft airframe. The studies on destruction of the standard cylindrical samples from the 40 steel and flat samples from D16 duralumin were conducted for experimental test of the drawn inferences validity. These types of samples selection is stipulated by the wide-spread occurrence of steels, having the yield point, and aluminum based allows in the power elements of the structures. High acoustic activity at the yield point, i.e. avalanche-like density increase of the mobile dislocations, is intrinsic to these types of materials. The developed approach provides the possibility of assessment in the process of control of dynamics and degree of change of the emission informative parameters, characterizing the degree of the pre-destructive state of the structure. Assessment with the relations being presented allows evaluating both initial and «rarefied» acoustic-emission flows of any order and does not depend on the loadings pre-history, shapes and sizes of the structures, which allows perpetrating constant and periodic acoustic-emission control.

In this study, we explored and analyzed how the design issues stemming from the Mars specific conditions have been addressed by the previous authors. The identified design trends, as well as the presented historical data on the previous Mars aircraft projects can be used as a basis for determining a future Mars aircraft mission scenario.

For several decades, scientists have been exploring Mars using orbiting spacecraft and rovers. Orbiters cover large areas and provide images of the planet surface with a resolution limited to a few meters, while rovers can analyze the composition of soil and rocks. In contrast, an aircraft flying at a low altitude above the surface of Mars will carry out a whole range of specific scientific research, mapping an area several orders of magnitude larger than a rover, with a resolution much higher than the resolution offered by modern satellites, as well as gathering valuable atmospheric data at different altitudes.

In contrast to the previous publications, the focus of the current investigation is to identify the relation between the Martian specific conditions and the design options adopted for exiting Martian aircraft projects. This will enable us to justify the design of a new fixed-wing Mars aircraft and to compose a set of relevant requirements to start the design process.

The recent improvements related to aerodynamic design, concepts of engines, energy storage and materials, have expanded the range of options for Martian unmanned aerial vehicles.

Possible missions of a future Mars science aircraft include performing a climatic, mineralogical, thermophysical and magnetic study of Mars.

The design process will be guided by the specific Mars environmental conditions (density, speed of sound, temperature, Reynolds number, dust storms, electrical phenomena, carbon dioxide carving). For a lander, Martian rugged terrain will exclude the conventional take-off and landing option. The need to deliver the aircraft to Mars and expose it to the space radiation will affect the aircraft aerodynamic layout, structural design, weight specification. The expected operating area, altitude, and season may significantly affect the design decisions in terms of aircraft configuration, geometry and total mass.

Finally, the flow field on a Mars airplane is expected to be highly complicated with a strong interaction of viscous and compressibility effects. This makes the numerical simulation of the aircraft operating in Martian atmosphere extremely challenging.

Nevertheless, the concept of a long-endurance aircraft, either solar or radioisotope powered, featuring foldable or inflatable wings and capable of flying in the Martian atmosphere seems feasible and can be considered as an option for future Mars exploration missions.

With all the variety of the existing programs for the of aircraft structures behavior mathematical modeling in real operating conditions, bench tests of fuselages and individual structural elements still remain the basic method, substantiating the confirmation (increase) of the resources. One of the tests elements is the reproduction equivalence assessment of operational loads on the bench. As of today, the basic method for equivalence assessment is the tested sample strain-gauging with the bench prior to the testing commence, the assessing criterion herewith is a measure of the discrepancy between the stresses at the structure critical points on the bench and those measured in previous flight tests. This criterion is sufficient for testing aircraft fuselages and, all the more, for individual elements of the aircraft structure. However, for the helicopter fuselage, particularly for the single rotor scheme, subjected to asymmetrical loads in a wide range of frequencies, availability of the technique allowing assessing the structure behavior in total even indirectly would be extremely useful.

The authors propose a new method for the reproduction equivalence assessing of the operational loads during bench tests of the Mi-26(T) helicopter fuselage.

The Mi-26(T) heavy transport helicopter was released with a declared assigned resource of 12,000 hours. However, the assigned resource up to date confirmed by tests is 4,200 hours with the possibility of a phased increase to 4,800–6,000 hours for helicopter instances, depending on the year of manufacture and technical condition. To achieve the designated resource declared by the Developer, it is necessary to continue bench tests of the fuselage. Up to the present day, two sets of such tests have been conducted.

A great number of cracks (up to the 1000 items per a single item, and about 10000 over the whole fleet) of a stringer from the 01420 aluminum-lithium alloy is being detected from the very beginning of the Mi-26 (T) helicopter operation. This alloy has not been applied since 1992, but more than 90% of the Mi-26T helicopter fleet in the civil aviation of the Russian Federation consists of helicopters produced in 1987-1992. Thus, their airworthiness maintenance is an urgent need.

Since 2002, the Federal State Unitary Enterprise GosNIIGA has been keeping records of these cracks, namely the location, the helicopter operation time at the moment of detection, etc. are being recorded, and the generalized map of stringers cracks for the entire fleet and crack maps for the separate samples of helicopters have been created and constantly updated. With all the negative impact of this defect on the operation, its mass character (if the crack occurrence is considered as an event from the viewpoint of the probability theory) allows full application of the mathematical statistics methods for its description. It should be noted particularly that distribution of the number of stringers cracks along the fuselage and in individual compartments qualitatively reflects the stresses distribution in specific zones.

The presented technique is based on a periodic comparison of distribution of the number of stringers cracks on the tested sample with the distribution of the number of cracks on the fleet of helicopters operated or previously operated in the civil aviation of the Russian Federation. The said technique suggests employing the Kolmogorov hypothesis likelihood estimation method for this comparison. This technique application allows assessing the structure behavior in total while the testing process, bringing it as close as possible to real operating conditions. Timely correction of the loading program allows increasing the sample durability on the bench. The said technique herewith does not require costly equipment and great time consumption.

The article demonstrates the technique approbation on the example of technical condition assessing of a specific helicopter fuselage (RA-06015) and by the example of a sample on a test bench.

The presented article deals with the matters related to operation of the roller bearings, functioning as part of rotors of the single-shaft and multi-shaft aircraft gas turbine engines, as well as methods of their hydraulic resistance reduction by the mating surfaces profiling. It presents examples of the developed roller bearing structures and results of their examining.

The goal, consisted in developing measures for the energy losses reduction while the bearing operation, has been achieved. For this purpose, the authors are solving the problem of the hydraulic resistance and internal friction reducing in the oil layers. To develop a physical model of the oil wedge hydrodynamic process they used the results of thermal imaging and temperature measuring on the operating bearing instrumented with the fiber-optic sensor, which is a new approach to this matter.

The developed roller bearings structure with the oil-removing grooves (which realizes oil bypass from the oil wedge zone with high pressure to the zone with reduced pressure) enables losses reduction on the internal friction in the oil layers, and avoid cavitation in the zone of oil wedge rarefaction. Analysis of the experimental determination results of the bearing temperature variation, that demonstrated its notable reduction, serves as a confirmation of this conclusion.

The obtained results attest to the possibility of employing the roller bearings with grooves, made on the mating surfaces of rolling elements and bearing tracks, to increase their operation efficiency by reducing the energy losses, as well as decreasing the heat liberation and functional noise.

Such bearings are expected to be employed in the structure of rotor supports in the aircraft and ground-based gas turbine engines. The proposed design may be expanded as well to the high-load bearings of different engineering products, especially operated under conditions of higher requirements to the absence of vibration and noise.

The presented article deals with two approaches to the durability calculation of the inter-shaft roller bearing (IRRB), which was passing life tests with the bearing test bench at the Central Institute of Aviation Motors (CIAM).

The durability computing is being performed by the contact crushing stresses obtained by analytical and numerical methods.

Various trends in the works on creating techniques for bearings durability determining exist nowadays both computational-analytical and experimental. Life tests of bearings by the equivalent programs relate to the experimental ones.

Computational-analytical techniques, in their turn, are also being separated into the two trends, namely analytical ones with the equivalent loading computing with further durability evaluation, and techniques, employing numerical finite element models of the bearings supporting subassemblies for their stress-strain state computing.

As is known, the reliability of machines and mechanisms operation largely depends on the performance of their bearing subassemblies. This is of special importance for the aircraft products, as the bearing subassemblies of aircraft engines, gearboxes, aircraft units and assemblies are one of the most critical subassemblies, limiting, as a rule, their resources. The inter-shaft bearing is one of the most problematic engine components. When detecting defect symptoms of the inter-shaft bearing, the engine is being withdrawn from service, as this can lead to the rotors jamming and the entire engine failure. The main reason for the rolling bearings failure under normal operating conditions is the contact stresses originations and, as the result, wear-out of the rolling surfaces.

Most of the well-known analytical methods for bearing collapse stresses computing are based on the Hertz’s theory of static contact between two bodies. However, there is a number of simplifications for this theory:

• no friction;
• the contact area is small compared to the radii of curvature;
• the materials of the contacting bodies are homogeneous, isotropic and absolutely elastic.

Numerical calculation allows solving contact problems without simplifying the Hertz theory:

• simulation of friction;
• accounting for the nonlinear properties of the material;
• accounting for the roughness of the contacting surfaces by selecting the size of the finite element mesh.

A comparative assessment of the stresses in the contact of the rollers with the raceways of the bearing with opposite and unidirectional rotation of the rings is performed, with account for the above said factors.

Dynamic calculation of the IRRB as a part of experimental bench was performed in two options to determine the contact crushing stresses and, as a consequence, durability estimation. The presented article compares the result of the study obtained by the engineering technique are being compared with the results of numerical analysis. The elastoplastic computations were performed using the LS-DYNA code.

It is noted that the dynamic formulation of the problem, realized in a numerical approach, allows obtaining more accurate results on stresses and, hence, bearing life.

The article presents the results of studying parameters and geometric ratios for combustion chambers of narrow-body aircraft engines with different combustion technologies.

Due to volume reduction of passenger transportation by world aviation, the wide-fuselage aircraft employing is being decreased. Narrow-body aircraft are once again becoming the most common in civil aviation. With a view to the political situation, the development of national narrow-body aircraft and their engines is becoming an up-to-date task for Russia. It is going to be solved in the form of the concept of import substitution. In terms of time consumption, the engine design is a more durable process than the aircraft development. Thus, it is important that even at the stage of preliminary engine design its optimal structure is selected. Combustion chamber is one of the problematic ones at the preliminary design of the engine components. This fact is associated with the presence of combustion process in it. It is impossible to compute the combustion chamber workflow and characteristics without its detailed geometry. Thus, the authors propose wide employing of statistical data on the existing products at the preliminary design stage. Within the framework of this work, the data on more than fifty narrow-body aircraft engines was accumulated and analyzed. Technical data, diagrams and drawings of their combustion chambers were analyzed. The authors considered chronology of the combustion chambers development of both domestic and foreign engines of narrow-body aircraft. The ranges of the thrust changing, total pressure ratio and gas temperature prior to the turbine were determined.

Thus, it was found that the pressure ratio increased 3.5 times while transition from the third to the fifth generation engines, and the gas temperature prior to the turbine by 800K and more. This was achieved, among other things, by combustion technology improving. Analysis of the change in the ratio of the combustion chamber length to the maximum height of its profile revealed that it decreased by about 1.7 times from the mid-1960s to 2015.This is the result of the low-toxic combustion chambers creation. Evaluation of the ratio change in the combustion chamber length to the engine length has been performed for the same period. The distance from the fan blade inlet (inlet device) to the turbine outlet was used as characteristic length of the engine. This distance is of interest in terms of the engine work process implementation. The lowest achieved values correspond to the TAPS and RQL combustion technologies. The value of this ratio is 1.5 times higher for the conventional combustion scheme. The data presented in the article allows performing weight-and-size-characteristics evaluation of the engines being developed with the specified parameters (the pressure rise degree, temperature at the engine inlet) at the preliminary design stage. Evaluation of various technologies capabilities for the engines operation efficiency enhancing can be performed as well.

There is a demand nowadays for small aircraft engines of a power up to 500 hp. Piston engines possess competitive edge in this category due to their light weight, low fuel consumption and decent weight-to-power ratio.

The most feasible way of ensuring this demand consists in converting automobile engines to aviation application and standards. Aviation engines are running for the most part at greater crankshaft rotation frequency and higher loads. It leads to the necessity for conventional systems alteration, including inlet manifold.

Earlier, the adapted piston engine was developed. In the framework of the engine-demonstrator, the input manifold, ensuring dynamic supercharging, was installed. Its size and shape were non-optimal from the gas exchange viewpoint. That is why structural refining of the manifold was required.

The greatest problem with the conventional manifold consisted in the uneven power distribution among the cylinders, due to the difference in filling up to 20% from the average value. The manifold was being designed for the lab tests as well, and fitted poorly the aircraft layout.

The purpose of the presented research consisted in equalizing mass flow through each cylinder with achievement of more even filling, which would ensure more even operation. It was desirable as well to ensure more aerodynamic shape and minimize pressure losses.

The core method of flow analyzing in manifold was the 3D CFD modeling. The non-stationary RANS model with realizable k-epsilon turbulence model and enhanced EWT was employed.

The main problems, such as dead zones in the back part of the manifold, the swirl in the front one and mutual effect of the branch pipes were determined by the geometry analyzing.

The following solutions were applied: the dead zone filling; the front part expansion for the swirl dissipation, and separators introduction. Each solution was applied iteratively with the search of preferable dimension and geometry up to the potential solutions exhaustion.

The resulting manifold design allowed achieving 50% and 30% reduction of maximum and average air consumption correspondingly. More aerodynamic shape was achieved. Pressure losses changes were within the error margins.

The article deals with the issues of the appearance forming of aviation hybrid power plant based on an internal combustion engine and an electric power plant for light aircraft. This is an up-to-date subject in both Russia and the world with a view to the environmental requirements tightening and the possibility of new aircraft schemes realizing on the assumption of the latest technological achievements in the engine building, electrical engineering, and electronics. The article regards the results obtained earlier with an experimental all-electric light aircraft. The effort on the electric power plant brought the authors to the thought of creating a hybrid engine on its basis.

The hybrid system was being considered in terms of the possibility of increasing flight time while keeping the advantages of the electric propulsion system. The effort on the hybrid propulsion system was preceded by extensive experimental activities on the bench of the electric propulsion system with a propeller and further flight tests on a light aircraft. All pros and contras of the electric power plant were revealed in real flight operation. The simplicity and operational reliability appeared to be positive features, while negative features were low battery capacity and short flight duration, as well as relatively large battery charging time. When considering the hybrid power plant appearance, the article analyzed various internal combustion engines and characteristics of the electric motors suitable for application in aviation. The scheme of direct drive for both types of engines to the propeller was selected as the most advantageous in terms of the system efficiency. Operating modes of the hybrid power plant at various flight stages were selected when analyzing the light aircraft flight cycle. Special attention is paid in the article to the practical operation and assembly of a bench sample of a hybrid power plant prototype. An important task consisted in revealing all the problems and the possibility of synchronizing the operation of two engines of different types, combined in a hybrid power plant. A suitable easy in operation and of reasonable price home produced engine was selected. A type of transmission, and reducing gear were selected. A test bench with the possibility of its mobile transposition was produced. In the process of idea try-out, a testing plan had been formed, which was being adjusted as and when necessary while the bench experimental works. A possibility of synchronous operation of the two engines of various types was proved and several characteristics on the propeller rotations and thrust were obtained while the experiment. The obtained results may be employed in the future for the larger class aircraft. Experimental works are being continued.

The presented work studied the impact of impurities in aviation fuel on the parameters of the working process and efficiency indicators of gas turbine engines (GTE) and power plants. The authors performed the analysis of various environmentally sound impurities and revealed the expediency of converting gas turbine unit to more eco-friendly gas fuels.

Computations were being performed with the «ASTRA» computer-aided system for thermal-gas-dynamic computation and analysis of the power plant. The fuel being used and components ratio in the obtained mixture fed to the GTE combustion chamber was being changed with the integrated fuel block. The study was being conducted for the mixtures based on the TC-1 fuel and hydrogen. Possible combinations of fuel mixtures were modeled with no regard for the requirements to their storage and application.

The results of the study are presented in the form of dependences of the engine working cycle basic parameters on the changes in the impurities concentration in the fuel. It is found that that methane is the optimal choice as an impurity, since with effective power increase the specific consumption is significantly reduced. It was obtained that hydrogen significantly affects the parameters, but its application in its pure form is not profitable and practical.

As the final stage, the similar research was conducted for the hydrogen based mixtures. This computation allowed defining the fuel that reduces the hydrogen concentration in the mixture for its cost reduction, though it does not affect significantly the inflammable mixture cost and effectiveness degradation of the hydrogen fuel.

The presented study demonstrated that concentration increase of the gas fuel affects beneficially the efficiency and economy indicators of the aircraft power plant. The inflammable mixture composition in its turn does not practically affect the engine effective efficiency coefficient. The methane and TC-1 fuel mixture is the best composite aviation fuel by its indicators. The liquefied natural gas application may result in significant aircraft characteristics improving, though the rational solution would be process commencing of phase-in natural gases implementing, starting from the gaseous impurities addition to the standard fuel types.

The object of the present study is a combustion chamber for oxy-fuel power generation technology, in which carbon dioxide at supercritical pressure is employed as the working body and oxygen is the oxidizer. The article presents the results of the combustion chamber designing of a gas turbine power plant, obtained with the techniques and scientific-and-technical solutions applied while combustion chambers for the aircraft engines creation. The techniques selection is stipulated by the following criteria: oxygen application as an oxidizer, high values of the flame tube thermal factor, reaching temperatures above 2500°C in the combustion zone, which brings the oxy-fuel combustion chamber of the power plant closer to the combustion chambers of aircraft gas turbine engines. A cannular combustion chamber with slot-type cooling of the flame tube was selected as a prototype.

Recommendations on the combustion chambers designing for carbon dioxide power plants, accounting for difference of the employed oxidizer, cooler and components of ballast, were proposed based on the studies being conducted. The dependencies of the flame normal propagation velocity value and adiabatic combustion temperature were determined with the Chemkin-Pro software complex for the combustion chamber being developed. Computational results allowed determining the carbon dioxide fluxes distribution along the flame tube length. According to the criterion of normal flame propagation velocity and adiabatic combustion temperature the value of CO₂ mass content in the combustion zone was selected as 0.6, which corresponds to supplying 12% of the total CO₂ consumption in the combustion chamber into combustion zone.

After substantiated carbon dioxide flows distribution in the combustion chamber, a constructive profile of the combustion chamber system, including a slot-type cooling, was obtained employing one-dimensional computations.

Numerical modeling of combustion and hydrodynamics processes was performed with the Ansys Fluent software package, which proved itself well for the design. The velocity vectors fields and temperature distribution plots along the wall of the flame tube of the combustion chamber were obtained. The authors revealed that the vortex did not form while the flame tube diffuser flow-around by the flow, and, as a consequence, the cooling section was locking did not occur. The cooling film steadily grows from section to section along the axis of the flame tube at the obtained design characteristics of the cooling system. The authors determined that the mass flow rates of carbon dioxide flows for cooling and mixing should differ by no more than 10% to maintain a stable film along the entire flame tube.

Aviation turboprop engines’ gearboxes are the most stress-intensive assemblages. Their main defect is the teeth side surfaces wear. The main hazard of the said defect consists in the vibrations generation that cause fatigue failures of the engine structural elements. Application of the widely exercised methods of vibroacoustic diagnostics for aviation turboprop engines has certain limitations. Mostly, intensities of vibration spectrum components and their combinations are employed as diagnostic signs of the defects. When developing diagnostic techniques, the required statistical data obtaining is being executed for the most part under conditions of a test bench of the engine manufacturer, whereas the diagnostics is being performed under operating conditions at the facility. However, a number of studies have shown that the engine re-installing from the bench to the facility led, as a rule, to the intensity increasing of the vibration process components. Respective conversion factors evaluation leads to the substantial material and time costs increase. Application of various types of propellers on both test bench and facility is possible for the turboprop engines. Evaluation of the engine re-installing from test bench to the facility and changing the propeller from one type to the other with a slightly higher thrust was performed on the example of the turboprop engine differential gearbox.

The following parameters were in use:

• Intensity of the two spectral components;
• The depth of the amplitude and frequency modulation indices of the narrow band process near the tooth harmonic of the «solar gear — satellite» pair at the solar gear rotation frequency;
• The width of the tooth spectral component at the level of the half of its maximum value;
• Deviation dispersions of the rotation frequencies values of both input and output shafts of the gearbox.

The authors revealed that the engine re-installing from the test bench to the facility led to the components intensities growth from 24 to 137%. Parameters changing, plotted on the frequency deviation characteristics stays within the measurement errors limits. The propeller type impact on the intensity based parameters was not revealed. Installation of the propeller of the higher thrust has not led to drastic changing of the parameters, based on the shaft rotation frequency deviation, up to the engine operating mode up to 0.85 of the rated value. Their significant difference was marked at higher operation modes. The obtained results demonstrate that application of the parameters based on rotation frequencies deviation characteristics of the engine shafts are insensitive to the engine re-installing from the test bench to the facility. While the propeller type changing, it is necessary to define the area of the engine operating modes, insensitive to the said change. The obtained results allow the gearboxes technical state evaluating under operation conditions.

One of the critical tasks in aviation gas turbine engine (GTE) creating consists in its design process efficiency increasing, which lies in the design period reduction while the project high quality and competitiveness ensuring.

The article considers conceptual design stage, which includes external design of the engine in the aircraft system, layout forming of the gas turbine engine workflow and its structural and geometry layout. This stage is being characterized by the substantial uncertainty level, which source might lie in the initial data incompleteness or generalization.

The initial data uncertainty impact on the engine parameters basic figures in the aggregate with tightening these figures permissible deviations from the project requires maximum possible transition from the initial design data values, predicted by based on the statistics, to the computed ones while successive solution of the project tasks. Mathematical models application of various complexity levels and dimensionality allows reducing the level of the initial data uncertainty as the project development forward and thereby cutting the terms of searching for the effective design solutions.

The need for employing system analysis, multidimensional optimization, the object modeling hierarchy principle and CALS-technology led to the idea of multilevel modeling. The GTE multilevel model represents the set of all engine elements and systems, employed at the various stages of the life cycle.

Accounting for the requirements for both multilevel model and design process allowed determining the most rational structures of the model being applied for the standard set of the design tasks. Conceptual design approach to the gas turbine engines designing with the multilevel model was elaborated on this basis.

The said approach application allows cutting the terms of computations due to the initial data uncertainty level reducing and the iterations number cutting between computations since the assembly units are being optimized in the engine system.

Development of modern machine-building production urgently requires solving the problems of predictability and reliability ensuring of the technological process of hard-to-cut steels and alloys blade cutting by cutting tools with innovative coatings based on studying the effect of the cutting process main modes on the temperature and force conditions and wear resistance; bringing to light the effective and informative parameters for control and diagnostics with subsequent development and introduction of adaptive control systems. Primary attention is payed in the article to the issues of cutting processing diagnosing by emplloying basic physical and chemical phenomena manifested in the process, i.e. the cutting temperature by thermo-EMF measuring; the cutting force components by determining the electrical conductivity of the «tool-part» contact, etc. To study the wear patterns of cutting tools with multilayer composite coatings during turning, characteristic representatives of three various groups of structural materials widely applied in aircraft engine structure, with significantly differing physical and mechanical properties, chemical composition and, as a consequence, different machinability by cutting were selected. They are 15X18N12X4TYU heat-proof, heat-resistant and acid-resistant austenitic steel of the IV group of machinability by cutting; HN73MBTU heat-resistant, deformable nickel-based alloy of the V group of machinability by cutting; VT18U titanium alloy of the VII group of machinability by cutting. Experimental tests were performed on the I6K20F3NC lathes with normal hardness, and 16K20 universal lathe, equipped with thyristor converter for stepless spindle speed regulation. Turning was carry through with carbide inserts of VK10 OM and T15K6 grades with different composition, thickness and architecture of composite wear-resistant single-component coatings, multi-component composite coatings based on double compound nitride systems, as well as triple compound nitrides (TiAlCr)N, (AlTiCr)N, (AlCrTi)N, (Ti,Al,V)N, (Ti,Zr,C)N). The coatings were obtained with both domestic and foreign Platit 311 and Platit 411 installations. The results of the contact processes experimental research, such as temperature and cutting forces, cutting tool wear-out etc., revealed that cutting process can be effectively diagnosed and reliability of the cutting tool with wear-resistant coating can be effectively predicted by the values of thermo-EMF and electric conductivity of the «tool-part» contact.

Additive manufacturing of products from metal powder materials is being put into effect in two ways, namely the powder bed fusion (PBF) and direct energy deposition. The first method is being realized in both laser beam powder bed fusion and electron beam powder bed fusion technologies in a powder layer.

With this method, the powder is being evenly distributed over the structuring platform, with selective scanning whereafter. Such approach leads to the increased powder consumption due to the need for filling the technological volume of the structuring chamber with it. This disadvantage may be eliminated by the method of direct energy and material supply, particularly, laser beam direct energy deposition (DED) or direct metal deposition (DMD) technology.

The purpose of the presented work consists in studying the effect of the direct laser growing modes, such as laser radiation power, transporting gas consumption and the speed of growth, on the shape and geometry of single rollers and walls obtained as the result of surfacing.

The additive installation for direct laser growing, including the Fanuc M-20iA_20M industrial robot, surfacing head and the Fanuc 2-axis Arc Positioner two axes positioner, on which table the samples were being grown, was employed for the study conducting. This installation is equipped with the three kilowatts YLS-3000 IPG Photonics ytterbium fiber laser and a FILED 30 IPG Photonics laser head with a removable four-jet coaxial nozzle for surfacing. The powder feeding to the fusion zone was realized by the Sulzer Metco Twin 10C powder feeder.

The powder from the HN50VMTUB brand heat-resistant nickel alloy, produced by the JSC «Composite» and JSC «Experimental Plant «Micron», was employed as the studied material.

While the study conducting, the single rollers were being surfaced on a substrate, which represented a sheet of ordinary grade St3 carbon steel of a 3 mm thickness. The surfacing was being performed with the DMD installation. The samples represented single tracks with the 30 mm length and a width of about 2.6 mm. Two series of experiments were performed herewith. Single rollers were being grown during the first series, while the wall consisted of five layers was being surfaced during the second series.

It can be seen from the measurements results analysis that the deposited material is being melted into the substrate to the average depth of 0.1–0.4 mm. The quenched layer of the 0.3–0.5 mm thickness is being formed in the substrate material owing to the fast heating under the impact of laser radiation and intensive cooling. The best convergence of the set and actual geometric parameters for single rollers, depending on the powder used, is being observed in mode 5, and 6 for fivefold tracks in mode.

The study of micro-hardness on the fivefold tracks revealed that the thermal impact zone had the same micro-hardness as the deposited material. The lowest microhardness occurs for both powders in mode No. 4. The maximum value of micro-hardness for the «Micron» powder is being ensured in mode No. 7, and for the «Composite» powder in mode No. 3.

Resource-intensive CFD methods, requiring both significant time and computing costs, are usually being employed to compute the aircraft aerodynamic characteristics. Thus, it is reasonable to apply fast semi- empirical methods for the aircraft conceptual design.

The article considers the existing semi-empirical methods for calculating the aircraft aerodynamic characteristics, and compares these methods with each other and verifies them with experimental data. Special focus is given to techniques that allow estimating the flaps and slats effect on the aircraft aerodynamic characteristics. Thus, the Arep’yev and Raymer methods are the two basic methods for the cruising aerodynamic characteristics estimation being considered in this article. To verify the mathematical models, computations of the cruising aerodynamic characteristics of the three aircraft with a maximum takeoff weight from 6600 to 21000 kg were performed by the Arep’ev and Raymer methods, and their results were compared with the experimental data. The high efficiency of the modified Arep’ev method for calculating the aircraft coefficients of lift and drag up to the angles of attack of 12° is demonstrated.

Among the techniques for the takeoff and landing aerodynamic characteristics estimation, the two methods that yield the most correct result were selected as well. Additionally, the article suggests a simple dependence of the additional drag coefficient caused by flaps deflection depending on the angle of their deflection. Comparison of the takeoff and landing aerodynamic characteristics computing results of the three aircraft with maximum takeoff weight from 6600 to 21000 kg with the experimental data was performed as well. This comparison demonstrated the high efficiency of the methods under consideration.

Airplanes powered by the Sun energy for the flight supporting and ensuring conventionally has a special structure and a wing of a wide span, while their aerodynamic surface are covered with photovoltaic cells. The wing of such flying vehicle consists of several sections to control its motion. Numerical study of the effect of interference of the airscrew and the solar battery powered airplane’s straight wing with extra-large aspect ratio on the hinge moments values of its deflected mechanization was performed. Computations were run at flow rates of V= 25 and 50 m/s and Reynolds numbers of Re = 0.17 and 0.35 10⁶ by the program based on the Reynolds-averaged Navier-Stokes equations. The presented work considered two options of mechanization equally deflected over all the wingspan, namely δ_mech = 15° and δ_mech = 30°, without airscrews and with two-bladed airscrews placed on the wing tips and rotated symmetrically in the fuselage direction with the rotation frequency of N = 15000 rpm.

The flow-around patterns and pressure distribution are presented in dependence on the propelling screw blow-off. The authors gave a comparison of the computational results in the 2D and 3D problem setting, as well as airplane aerodynamic characteristics comparison of the without blow-off by the airscrews with the experimental data.

Numerical studies reveal that the presence of airscrew effect on the hinge moment value of the mechanization deflected depends on many factors, such as airscrew diameter and its design features, rotation frequency, its location, as well as blow-off and deflection angles. With the blow-off by the propelling airscrew, placed prior to the wing, local angles of attack on the wing and mechanization change, and the pressure at the windward side in the area of the blow-off by the airscrew

The blow of pulling airscrew, which mounted in front of the wing, influence on change local angle of attack wing and mechanization, decrees height of separate zone and increase pressure on windward side in the area of blowing airscrew.

Analysis of computation of profile and wing revealed that hinge moments computating in the 2D problem setting without blow-off may be employed for fast predicting the straight wing mechanization hinge moments values.

Spacecraft, as a rule, have outboard structures of low rigidity, such as solar panels. When simulating a dynamic model of elastic spacecraft and selecting control system settings, it is necessary to set up a system of equations of motion, as well as determine its coefficients characterizing both elastic (eigenforms, oscillation frequencies and inertial coupling coefficients) and dissipative (logarithmic decrements and damping coefficients) properties of the structure. As the result of frequency tests, the dynamic characteristics of the solar panel were obtained, namely the spectrum of natural frequencies, shapes, decrements, as well as the dependence of frequencies and decrements on the amplitude of the panel oscillations. It should be noted that with the amplitude changes, the spread of the decrements values of the oscillations might be rather significant. It is stipulated by the fact that at small oscillation amplitudes, energy dissipation is mainly determined by internal friction in the material and structural damping, which is characterized by friction in kinematic pairs, as well as in splined, threaded, etc. joints. While loading, small slippages in such joints occur over the contact surfaces, which may lead to drastic energy dissipation increase, and does not meet the flight conditions in outer space. Besides, the weight-killing tether system, which introduced extra stiffness, weight and damping, was employed while these tests for the docking nodes with gaps offloading. The system with vibrators makes as well its contribution through the attached masses of coils and their attachment points. The elastic and dissipative characteristics refinement of solar battery wing of the “Spectrum” type spacecraft based on frequency tests results of the panels and analytical studies with account for the weight- killing system impact was performed. The solar battery wing herewith, consisted of four panels, was the object of the studies.

The results of frequency tests of the wing of the solar battery of the «Spectrum» type spacecraft were analyzed. The obtained results demonstrate that the oscillations are of nonlinear character. The presence of backlashes in the drive are stipulated by the dependence of natural frequencies and decrements of oscillations on the oscillations amplitude. This is also the reason for the oscillations forms deviation from the obtained calculations.

The modal parameters of the solar battery wing of the spacecraft were identified based on the results of the frequency tests with account for the weight-killing system impact. A good agreement herewith between the calculated and experimental characteristics with the offloading system was obtained, which allows feasible selection of the dynamic model parameters values of “Spectrum” type spacecraft.

The article recounts the technology of the fuselage internal layout by the automated system of three- dimensional layout for passenger aircraft (AVTOKOM) developed in TsAGI.

AVTOKOM allows forming passenger cabins and cargo bays of the fuselage with account for the specified comfort standards and safety requirements to the deployment of passenger seats, common and service premises, operational and emergency exits, luggage compartments etc. in both interactive and automatic modes.

The stages of fuselage layout by AVTOKOM are as follows:

1. Formation of typical elements that meet the specified standards and requirements.

2. Optimization of the cross section of the fuselage regular part.

3. Passenger and cargo decks layout.

4. Creating a parametric model of the fuselage outlines.

5. Three-dimensional surface model of the entire aircraft outlines.

6. Calculation of the center of mass of the elements comprising the layout.

7. Visualization of the studies results.

8. Output data formation for the subsequent calculations.

An iterative technology of passenger aircraft geometric model formation has been developed, on which basis further research in the areas of aerodynamic layout, structural strength and aircraft control systems are being conducted. As the result, the aircraft mathematical model that meets the layout requirements and numerous physical criteria is being formed.

The article presents the examples of the AVTOKOM application while performing the layout studies

of:

  – A long-haul aircraft with 200, 400 and 600, 800, 1000 and 1400 passenger capacity for medium and

long-haul airlines;

  – A long-haul aircraft concept with an integrated power plant;

  – A long-haul aircraft on liquid hydrogen fuel;

 – An aerospace plane with a capacity of 5-7 passengers.

As the result of these studies, the external geometric contours, layouts of passenger cabin and cargo bays of fuselages with elements of equipment and interior and specified nomenclature of service and cargo equipment, as well as layouts of the landing gear and fuel tanks have been formed. The article demonstrates that the standards of passenger comfort and safety requirements are met in all of the considered aircraft projects.

The article tackles with the problem associated with the possibilities of rocket engineering development based of innovative fuel charges for rocket engines of the guided aircraft (GA) in accordance with the conventional set of requirements to the fuel, which selection is being defined as one of the most crucial stages of creating the state-of-the-art samples of the rocket engineering.

On the assumption of the set problem, the authors conducted analysis of the existing developments and types of the gel-like rocket fuel to define the effectiveness of such fuels application as a prospective type of the fuel charge for the engines of GA.

Special attention was paid to the rocket fuel selection technique for application in the engine of the GA, based on the relative indicators of the fuel itself. The article presents the basic dependencies, which associate the rocket fuel parameters with tactical and technical characteristics of the aircraft itself. Being guided by the method of the rocket maximum ideal flight velocity, the authors define the basic parameters of the prospective fuel charge of a rocket engine, which enhancing will allow developing the sample of rocket engineering capable of surpassing the existing analogs by the set of important characteristics. Thus, the article confirms the effectiveness of the gel-like fuel application, which possesses high specific energy indicators and capable of ensuring increase of the tactical and technical characteristics of the missile itself, for the engines of the GA.

In this connection, the article describes the basic features of the gel-like rocket fuels uncovering the possibilities of the rocket fuel of this type application to solve the problems facing modern rocket building. The generalized technique for forming the gel-like rocket fuel composition employing equivalent formula for possible realization of maximum energy characteristics with the required operational parameters preserving of this type of rocket fuel was considered. The authors present herewith characteristic of the well-known gel-like rocket fuels contents, and define the possibilities of their improving by application of high-energy additives or well-defined relationship of the basic components.

The article regards the basic problems while creating the pilot sample of the rocket engine with the gel- like fuel charge stipulated by the specifics of this kind of the rocket fuel, which allow ensuring characteristics surpassing conventional analogs. In this connection, the article presents the description of the important design-engineering solutions, which are developed for the innovative sample of the rocket engine with the gel-like fuel charge realization. These solutions ensure also this engine effectiveness as a part of the missile due to the stressed state of the gel-like fuel charge, as well as the possibility of changing geometrical configuration of the combustion surface to achieve the appropriate values of the required thrust.

In conclusion, the authors give a brief characteristic of the possibilities of the gel-like rocket fuels application and adduce recommendations on their further upgrading for employing in the prospective rocket engines developments.

When an aircraft fly through a rain zone at high speed, the windshield and the advancing parts of other components, as well as the coating of the aircraft skin are being easily destroyed due to raindrop-shock erosion. In the studies of the aircraft damages from the raindrop-shock erosion, which is the most common at the subsonic speed, due to the low speed, the value of pressure generated by one impact is assumed negligible. Thus, hundreds or thousands of successive impacts are often required over a time period to cause damage to the surface of materials or structures. In this case, all researchers are paying attention to the mechanism of damage from the fatigue loading. Although the probability of raindrop shock of a supersonic speed occurring is low, its peak water hammer pressure impulse (up to the GPa level) far exceeds the strength of many materials, and one or more impacts are enough to damage the material or structure. At this time, much greater attention is being paid to the mechanism of the damage from shock loading.

Due to the advantages of the small size, ease of operation, and controlled test conditions, the single-jet generator is most widely used in the studies on the mechanism of damage to materials and the interaction of raindrop-shock erosion. The presented work considers a single-jet impact platform, based on a gas gun, which is capable of stable water jets generating with the speeds of 90-700 m/s and arc-like front section diameters of 4-7 mm. Then the test on the jet shock upon the oriented and non-oriented aviation organic glasses (Polymethyl methacrylate – PMMA) for are being conducted at various speeds. According to the experience, the optimal position of the organic glass sample setting while the raindrop-shock erosion testing is 10 mm from the nozzle.

The results indicate that at the high-speed jet shock impact damages in the form of surface stratification manifest themselves with the oriented organic glass, while with the non-oriented organic glass these damages are the surface ones. With constant impact velocity increasing, the surface stratification appeared on both organic glass samples, and stratification of the oriented organic glass at that was more serious. Observing the stress wave propagation and damage expanding inside the sample revealed that the shear waves prevailed in the subsurface stratification of the oriented organic glass.

The presented article deals with analysis of integration interaction of a wing and wingtip-mounted propellers.

The main purpose of the study consists in defining the useful effects originating when engine mounting in pulling, pushing or tandem scheme in the specified position relative to the wing due to the interference interaction.

The author performed variation of several parameters, defining mutual arrangement of the wing and propussors, as well as size and parameters of the propellers.

The article shows that relative increment of maximum aerodynamic quality K_max through wing-tip propellers installation increases with the wing aspect ratio λ decrease. The absolute value of K_max, in its turn, is higher at the propeller diameter and B parameter increase. Thus, with λ = 10, D_prop/b_wing = 1.0, the aerodynamic quality increment ΔK_max reaches 19.5% at B = 0.4. Maximum increment of aerodynamic quality with λ = 6, B = 0.4 and D_prop/b_wing = 1.0 reaches 33% of the K_max value of an aircraft without propellers.

Under conditions close to the real cruising flight (M = 0.4, B = 0.2), in case of the wing aspect ratio of λ = 10 and D_prop/b_wing = 1.0 obtaining the increase of ΔK_max ~6.4 is possible. Witht the wing aspect ratio decrease up to λ = 6, the increment ΔK_max increases up to 11%, though, the level of ΔK_max absolute values decreases from 17.1 to 14.1 compared to the case of λ = 10. It was established that propeller installation behind the trailing edge affects slightly the aerodynamic characteristics changing.

The article considers as well the possibility of installing tandem propellers, i.e. one prior to the leading edge and the other behind the trailing edge of the wing. Thus, installation of only the front propeller at λ = 10, B = 0.2 and D_prop/b_wing = 1.0 leads to the K_max value increase by 6.4%; while the additional installation of the rear propeller leads to a certain K_max decrease up to 5%. Rear propeller diameter varying at the tandem location of the propellers does not affect practically the value of the aircraft K_max.

The main advantage of the tandem propellers compared to a single one consists in the increasing aircraft safety, wince in the event of the front or rear propeller failure, the system thrust only approximately halves, rather than falls to zero.

The development and creation of unmanned aerial vehicles is the most dynamically developing trend of the aviation industry worldwide. This is being facilitated by the continuous practice of their application in solving a wide range of diverse tasks. In this diversity, the military purposes unmanned aerial vehicles occupy a special place, since demonstration of their capabilities by law enforcement agencies while solving combat tasks in modern local conflicts in the most obvious way reveals the advantages of their application.

Against this background, a steady trend of the unmanned aerial vehicles development is being observed in our country with a forecast for decades to come. To reduce terms and costs for the unmanned aerial vehicles development the authors propose to realize the targeted development of prospective unmanned flying vehicles by the principle “Task — solution option — facilities — terms — cost”. The issue of the power plants developing still remains herewith the most complex one, which is being associated with the lack of the stat-of-the-art substantiated methods and techniques combined with the criteria, on which basis the assessment of the power plant efficiency with various types of aviation engines characteristic for application on the unmanned aerial vehicles.

The article presents a unified methodological approach to the development of the military purposes unmanned aerial vehicles with hybrid power plants and power plants based on the engines of conventional types and schemes, such as gas turbine, piston and electric. Special attention herewith is paid to the disclosure of problematic issues of scientific and research nature, and production straightforwardly when creating aircraft engines for the power plants of unmanned aerial vehicles. These issues relate to the stage of external design of military purposes unmanned aerial vehicles and their power plants, and affect the fundamental and applied foundations of design and production, which should be accounted for while preliminary design.

The article describes the following issues developed by the authors:

— The methodology for the precursory technical appearance forming of power plants for the military purposes unmanned aerial vehicles;

— The technique for substantiating optimal parameters of both engine and airframe;

— Classification of military purposes unmanned aerial vehicles;

— A complex mathematical model of an unmanned aerial vehicle for computational and theoretical studies of the “Unmanned aerial vehicle — power plant” system using computer software.

For further development of the complex mathematical model, the authors plan to finalize the mathematical model of the power plant based on both turbo-screw and piston engines, as well as hybrid options of power plants, including an electric generator in addition to the “thermal” engines, an electric motor and a separate propulsor.

The practical value of this work, which consists in the fact that its results may be employed in both scientific and design organizations, preoccupied with developments of prospective unmanned aerial vehicles and their power plants, as well as ordering organizations and industry while substantiating requirements to the new samples of aviation engineering, is worth mentioning.

The article proposes a method for reconstructing the average integral heat transfer coefficient as a function of temperature for axial heat pipes. This method is based on the studied parameter representation in the form of its parameterized value, multiplied by the corresponding basis function that describes its dependence on the temperature. Linear-continuous function was selected as the basis one. Further, with the selected initial approximation of the heat transfer coefficient parameter, the “direct” problem of the theoretical temperature field determining is being solved under known initial-boundary conditions and thermo-physical properties of the material. Based on the flight thermal elaboration of the axial tube, the root-mean-square deviation between the theoretical and experimental temperature field at the sites of temperature sensors installation is being composed. The obtained functional is being minimized by the conjugate directions method, with preliminary selection of the descent step. The descent step is being selected from the condition of the residual functional minimum at all iterations, starting from the second one. Likewise, one of the most important tasks prior to minimization is finding the gradient component of this functional. For this purpose, the statement of the “direct” problem of heating the pipe is being solved again with a preliminary differentiation of this statement of the problem by the parameterized value of the heat transfer coefficient. The sum of errors, namely systematic, statement of the research problem, rounding and the set problem solving method, was selected as the iteration process termination criterion. Reaching the termination criterion assumes that the searched for parameterized value is found, otherwise the above described routine should be repeated again. To check the adequacy of the developed method, the obtained result was compared to the method for the heat transfer coefficient determining from the thermal resistances analysis based on the experimental temperature field. Analysis of relative errors shows good convergence in the case of this coefficient averaging over time with its experimental counterpart, otherwise, a greater number of considered time blocks and a more accurate thermal model of an axial heat pipe are required.

The article emphasizes the relevance of the Moon problem exploration, namely construction of a long-term lunar orbital station and a habitable base on its surface. Attention is also focused on the fact that these programs implementation will require more than 1000 tons of payload. One of the ways of a payload delivery to the lunar orbit is a reusable inter-orbital transport vehicle (RIOTV), which propulsion system should be capable of multiple turn-ons. With a view to the unit cost minimizing of the payload leading out, the RIOTV propulsion unit should be optimized. In this regard, the article defines the technical appearance of the liquid-propellant rocket propulsion unit (LPRPU), optimized by the following criteria: minimum mass and minimum unit cost of the payload leading out.

Complex mathematical model, conjugating mathematical models of “rocket” and “engine” basic design parameters (BDP) impact on the effectiveness criteria of RIOTV and STS, was developed to define technical appearance of the LPRPU of the RIOTV in total. Computation of the two options of optimal LPRPU ROTV for the concrete of leading out task, namely the 16500 kg payload insertion into orbit, was performed with the developed model.

A technical appearance with high values of both turbine efficiency and pressure in the combustion chamber was obtained by the computation results. Next, the turbine efficiency in the obtained layout was reduced to much realistic values. The authors established that application of optimized option of the LPRPU with less effective fuel turbine, reduced pressure in the combustion chamber and reduced rotation frequency of the fuel turbo pump unit for this transportation operation will allow, with otherwise equal parameters, increasing lifecycle of the LPRPU as a whole and reducing the unit cost of the payload leading out.

The article demonstrates that the basic LPRE, which technical appearance and basic project parameters are not optimized to the problem of leading out being considered in the presented study are ineffective by the majority of criteria.

The presented article recounts the results of the studies on the flame tube walls temperature determining of the gas turbine engine (GTE) running on the gaseous fuel.

The flame tube walls cooling is one of the essential components of the processes organizing in the GTE combustion chamber. The combustion chamber operation reliability and the engine lifetime in the aggregate are fully dependent on the effective cooling of the flame tube walls. One of the most widespread cooling systems is convective-film one, consisting in the air film forming, which does not allow the heated gas interact with metal and removes the heat from the opposite side of the wall due to the convection.

The article presents the description of the test bench equipment. It considers thee options of burners that differ by the nozzle attachment design, the geometry of the swirler and atomizer herewith remains unchanged. The results of fire tests studies of three burners with various nozzle attachments are presented. Comparison of the flame structure of the two burners was made.

The article presents the combustion chamber design of converted aircraft gas turbine engine, meant for the supercharger drive of the gas-pumping unit. Dissection of the combustion chamber walls in its various cross-sections was performed, and combustion chamber testing as a part of gas turbine engine was conducted.

Temperature of the walls at the modes being considered does not exceed 800°C, which is indicative of the ample flame tube cooling.

Based on the results of the work being conducted, the inferences were drawn on the most acceptable option of the burner for implementation with the engine.

The most common scheme of bench tests of an aviation gas turbine engine in the test stand box or on an open test stand is a layout, which includes a test-bench input device with a lemniscate headpiece.

When the turbofan engine testing in the thermal pressure chamber, an inlet connected pipeline or inlet device with a lemniscate headpiece (testing scheme “with a baffle”) is being employed. Such inlet devices include a flow rate metering manifold for measuring air mass flow rate, which measuring section is set at a relative distance of not less than 1.5 gauge (L/D) from the inlet to the tested turbofan engine according to the requirements of the Industry Standard OST 1 02555-85.

When conducting turbofan engine tests in the thermal pressure chamber of the high-altitude test bench under both climatic and altitude-velocity conditions by the scheme with the connected pipeline at the inlet, as well as at the autonomous low-pressure fan testing on the compressor test bench with the air heating or cooling at the inlet, the heat inleak forming to the subsonic flow at the turbofan engine inlet (or it removal) is possible.

The process of the flow energy additional increase (decrease) while heat input (removal) leads the thermal boundary layer forming in the in the cross section of the mass air flow meter and at the engine inlet. It leads as well to the change of regime parameters values (total pressure p*_lN and total temperature T*_IN ) at the inlet to the turbofan and to their certain difference from the corresponding values of p*_M and T*_M, measured according to the OST 102555-85 from the inlet section into the air flow manifold at a relative distance of at least L_M-M÷IN-IN/D_M≥ 1,5.

However, for the turbofan engine with high bypass ratio, mode parameters measuring in front of the engine (p*_IN (Pa), T*_IN (K)) is difficult to ensure under bench conditions for a number of reasons:

– due to the absence of fan inlet guide vanes for a low-pressure fan of a turbofan engine, which moght be employed for measuring values of (p*_IN, T*_IN) ;

– owing to the possibility of resonance stresses occurrence in the fan working blades while installing radial chasers for p*_IN, T*_IN measuring nearby in front of them in the inlet bench channel.

Neglecting accounting for the heat inleak (or heat removal) as applied to the turbofan engine with the large degree of bypass and reduced fan pressure rise degree may, in some cases, lead to noticeable errors in the turbofan engine basic data estimation. It relates,in particular, to the fan efficiency value, as well as the values of the turbofan basic parameters reduced to the international standard atmosphere.

The article recounts the technique for determining the values of regime parameters of breaking temperature T*_IN and total pressure p*_IN directly at the inlet of the turbofan engine of high degree of bypass. This technique accounts for the heat inleak (removal) to the airflow in the pipeline, attached to the engine in the section between the measuring section in the flow manifold and the section prior the engine inlet, by reference to the condition of mass air flow rate value preserving G_AIR.M=G_AIR.IN and accounting for heat line of ΔE in the total flow energy value of E_IN =E_M + ΔE at the turbofan inlet.

The article presents the examples of the heat supply effect (or heat removal) on the nature of the thermal boundary layer changes in the flow measuring section of the inlet pipeline by the results of tests of turbofan engine under thermal vacuum chamber in the altitude-speed conditions. The example of estimating the value of the heat flux to the airflow and the corresponding change in the main regime parameters at the inlet to the turbofan engine of high bypass ratio is recounted.

High frequency ion thrusters are one of the electric rocket thrusters schemes employed in spacecraft as low thrust engines. Initially, electrojet thrusters were applied for geostationary satellites orbit stabilizing and correcting. Recently, the range of problems being solved in space engineering by dint of the electrojet thrusters has expanded significantly. It is worth noting that such thrusters’ application for bringing satellites into calculated orbits, as well as their successful employing as cruising propulsion systems for implementing missions into deep space, for flights to the Moon and minor planets of the Solar System.

High frequency ion thrusters (HFIT) are the variety of electrojet thrusters. Plasma in the discharge chamber is being sustained by the high frequency electromagnetic field, in contrast to the more world-common Kaufman DC-based scheme, in which plasma is being generated by high-energy electrons injection into the discharge chamber.

Initially, relatively simple configurations were employed for the HFIT structures basic elements, which were the discharge chamber and ion-optical system (IOS) electrodes. In the current practice, the HFITs were of cylindrical, semispherical and conical form, or their combination. The flat IOS electrodes were being selected for the thrusters with the ion beam diameter less than 10 cm. For the thrusters with greater ion beam diameter electrodes with relatively small outward buckling were employed to avoid significant thermoplastic deformation of electrodes of the ion-optical system, being heated by the plasma while the thruster operation. With that, the task of determining the most optimal from the viewpoint of the engine thrust, the plasma volume shape, limited by the surfaces of the discharge chamber and the IOS electrodes was not directly set.

The article proposes employing the discharge chamber with reduced surface curvature and noticeably convex IOS electrodes in the HFIT structure. Numerical model for computing plasma parameters in the HFIT discharge chamber allows setting an optimization problem on determining the best geometry of the discharge chamber and the IOS electrodes. It is being planned to employ the engine thrust, being computed from the calculated basic plasma parameters distributions over the volume, namely electron density and electron temperature, as the optimmization criterion.

An interest of the leading aerospace enterprises [1, 2, 3] in development and improvement of very low-power electric propulsion thrusters, which are characterized by a discharge power less than 100 W, for small spacecraft, including CubeSat standard small spacecraft, can be explained by a predicted possibility of getting new scientific knowledge and earning a commercial profit by using small spacecraft equipped with propulsion systems with high values of a generated total thrust impulse. Due to the interest of the world market in the availability of propulsion control systems for small spacecraft, the works on creation of very low-power plasma thrusters were initiated at EDB “Fakel”.

This paper gives the results of research work with experimental laboratory models of very low-power plasma accelerators U-M1 and U-M2 created with the purpose of searching and subsequent optimization of new technical solutions for very low-power plasma thrusters which are developed at EDB “Fakel”. The accelerators U-M1 and U-M2 are built on the basis of two principal schemes which differ by the configuration of their magnetic and discharge systems, what allows to expand the available range of magnetic field parameters and electric discharge parameters defining the studied operational processes’ organization in a discharge chamber. The accelerators’ models were created based on the principle of achievement of maximum simplified systems configurations at a minimum possible geometry enabling stability and sufficiency of the operational process.The results of the U-M1 and U-M2 accelerators performance research works are presented. A long-time functioning of two models of plasma accelerators has been demonstrated, which functioning is characterized by a stable operational process for a long (for this dimension type) time of a total firing and by the sufficiency of accelerators’ thrust parameters:

• U-M1 accelerator: thrust is 0,77 mN, anode specific impulse is 523 s at the discharge power of 27 W;

• U-M2 accelerator: thrust is 0,5 mN, anode specific impulse is 313 s at the discharge power of 20 W.

Specific features of the U-M1 and U-M2 accelerators’ operational process related to a very low geometry of systems and very low discharge power have been studied, and as a result, an assumption of a position of the ionization core and acceleration layer outside the spatial limits of the discharge chamber has been formulated. In case of an experimental confirmation of this assumption, the possibility of using the known assessment criteria of the ionization core and acceleration layer position for the conditions of very low geometry and very low discharge power is put in doubt.

The article deals with studying the effect of the inter-rotor bearing numerical model characteristics such as:

– characteristic size of elements, determining the computations step;

– numerical formulation of finite elements;

– integration method (explicit or implicit) on the computations time and accuracy.

Reliability of both machines and mechanisms is known to be largely dependent on the bearing assembly operability. This is of special importance for the aircraft engineering products, since bearing assemblies of aviation engines, reducing gear, units and products of aircraft are one of the utmost crucial assemblages, which determine as a rule their resources. The inter-rotor bearing is one of the most problematic assemblages of the engine. The engine is being taken off the operation while the inter-rotor bearing defect symptoms diagnostics since in may lead the rotors jamming and failure of the whole engine. The main cause of bearings failure under normal conditions is an emergence of contact stresses, and consequently rolling surface degradation.

Most known analytical methods for computing the contact crumpling stresses in bearings are based on the Hertz theory on the static contact of the two bodies. However, there is a number of simplifications for this theory:

– nonexistence of friction;

– the contact area is small as against to the curvature radius;

– materials of contacting bodies are homogenous, isotropic and perfectly elastic.

Numerical computation allows solving contact problems without simplification of the Hertz theory:

– friction simulation;

– accounting for the material nonlinear properties;

– accounting for the contacting surfaces roughness by the finite-element mesh size selection.

The authors performed comparative assessment of the stresses in the rollers contact with bearing roller ways with the opposite and unidirectional rotation of rings with account for the above-listed factors.

The effect of the inter-rotor bearing rings misalignment on the contact stresses of crumbling was studied in this work as well.

The factors assessment was performed in the LS-DYNA software.

The presented work was accomplished for the dynamic model preparation, where the bearing rings rotation is accounted for.

With each year, the space debris poses increasing threat to the functioning spacecraft, as well as people and property on Earth. Dozens of large-size spacecraft enter annually the atmosphere and reach the Earth surface, and there is always a risk herewith of inflicting damage to the people or property. Several collision have occurred by now in the near-Earth space, which can be avoided in the future, if appropriate characteristics of the systems, which ensure warning about such events, will be guaranteed.

The basic method of the threats parrying associated with the space debris is a warning about dangerous situations, namely time and place of large objects re-entry, a possible collision of a spacecraft with space debris or some other spacecraft. For realizing this method and solving corresponding problems, the refined data on the spacecraft orbits parameters by measurements are being required. Accuracy improving of the orbits parameters evaluation and their further prediction is necessary for safety ensuring of space activities under conditions of a large number of spacecraft.

The article presents basic mathematical relationships of optimal measurement filtering method (OFI), and shows that the OFI method application may significantly improve the results of the re-entry time evaluation and the space objects collision probability compared to the conventionally employed least square method. The results of the OFI application while predicting the time and place of the Tiangong-1 orbital station re-entry are demonstrated using the available accessible data. A posteriori evaluation of the prediction results accuracy showed that the OFI application allows sevenfold accuracy increasing of the estimates, without increasing herewith the computational complexity.

One of the ways of new space debris forming mitigation consists in its active removal. Presently, the works on the space debris active removal have been transferred from research to the ones being realized in daily practice of space activities. In the years to come, a number of projects will be implemented to remove spent upper stages, rocket bodies and spacecraft from orbits. The article presents the results of comparing the areas of the space debris active removal obtained by the technique, which accounts for the OFI with a concrete list of objects, obtained by a group of international experts. As is seen from the comparison, 48 out of 50 objects get into the calculated areas, which indicates a good correspondence of results obtained earlier with estimates of international specialists group. In this regard, it can be considered that both the ranges of orbits in altitudes and inclinations, and specific objects have been determined to prevent collisions that could lead to a large formation of new objects in the near future.

The OFI method application in monitoring and warning systems for hazardous events related to the space debris will increase efficiency of their functioning with the existing measuring instruments.

The strength of compact sample at normal separation (fracture mode I) was studied within the framework of the Neuber–Novozhilov approach. A model of ideal elastoplastic material with ultimate relative elongation was selected as a model of a deformable solid. This class of materials includes, for example, low-alloyed steels applied in the structures operating at temperatures below the cold brittleness threshold.

The crack propagation criterion is formulated with the modified Leonov–Panasyuk–Dugdale model, which employs an additional parameter, namely the plasticity zone diameter (the pre-fracture zone width). The two-parameter (twinned) criterion for the crack quasi-brittle fracture in the elastoplastic material was formulated under conditions of small-scale yielding with the presence of the stresses field singularity in the vicinity of the crack tip. This twinned fracture criterion includes the deformation criterion, formulated in the crack tip, as well as force criterion, formulated in the model crack tip. The lengths of the original and model cracks differ by the pre-fracture zone length.

Diagrams of quasi-brittle fracture of a sample under conditions of plane strain and plane stress are plotted. These diagrams consist of two curves, which divide the “crack length–stress” plane into three regions. The first region corresponds to the absence of fracture. In the second region, damages are being accumulated in the pre-fracture zone under the repeated loading. In the third region, the sample is being divided into parts under monotonic loading.

The constitutive equations of the analytical model are analyzed in detail depending on the characteristic linear size of the material structure. The authors obtained simple formulas suitable for verification calculations of the critical fracture loading and the length of the pre-fracture zone. The analysis of the parameters included in the proposed model of quasi-brittle fracture was performed. The authors propose model parameters selecting by approximation of the uniaxial tension diagram and stress intensity coefficient.

The subject of the article being presented is a helicopter-type unmanned aerial vehicle (UAV) with a coaxial rotors design. The research issue is landing procedures automation on a site unprepared with respect to engineering. The purpose of the work consists in developing a set of basic requirements for an air-defined landing site based on current aviation standards, as well as implementing neural network classifier of the underlying surface. The authors considered the existing methods of landing performing for the UAV. As the result of the analysis, the method of autonomy enhancing by implementation of information systems and sensors of various operation principles was defined as the most promising. With account for acting Federal Aviation Regulations (FAR), as well as norms adopted by the International Civil Aviation Organization (ICAO) and European Aviation Safety Agency (EASA), the list of requirements for the prospective landing zone characteristics, accounting for the specifics of the UAV studied in the work, was developed. The main complexity here consists in the lack of the standardized regulations of performing landing procedures for the UAVs of this weight class of 325 kilos. The review of the conventional methods for the underlying surface quality determining was conducted. By reason of small overall sizes of the aerial vehicle being studied, meso- and micro-relief of the terrain are of special interest. The authors decided to split the algorithm for appropriate landing site determining into the two logical stages. Optical survey of the terrain and determination of several optimal prospective landing zones based on color semantics, characteristic structure patterns, presence of obstacles and proximity of the terrain regions transition are being executed at the first stage. Next, the descent to the most optimal site to the altitude exceeding the critical decision point is being performed, and relief scanning by the compensated laser-radar system is being executed to obtain the relief model and determine the soil characteristics. Both technique and software development was being performed in the course of this work for the first stage of the underlying surface primary inspection. The main problem of the video fixation cameras application onboard of aerial vehicles consists in strong dependence of the obtained data processing results on the environment state. Variability of both weather conditions and Earth surface lighting conditions may exert drastic parasitic effect the result of the algorithm execution. Various methods of preliminary image processing, such as contrast ratio improving, segmentation and noise filtering, allow partially solving this problem. However, the greatest invariance to the shooting conditions can be achieved using neural network methods for image analysis. The authors proposed an optical recognition method of the prospective landing zones employing self-organizing Kohonen maps. The neural networks of this kind advantage is the simplicity of the training sample preparing, as well as simplicity of the synoptic weights distribution process in the course of the casual observer training. The selected approach allows evaluating not only the color specters distribution on the image, bug tracking characteristic patterns of the texture as well. The training sample contained 2700 fragments of the terrain topographic snapshots, and the neural network training time was 10,000 epochs. Computer tests revealed 21% of the alpha errors and 0% of the beta errors, which is specific for the neural networks of this class as well. The results obtained in the course of this work are simultaneously indicative of this approach exploitability to the underlying surface clustering and the need for further research on the considered issue.

The article deals with theoretical and experimental studies of cutting tool wear intensity while machining chrome-nickel alloys under temperature-force conditions employing modern wear-resistant complexes. Application of modifying multilayer-composite multicomponent nanostructured wear-resistant complexes is one of the most promising ways to improve the cutting properties of edge tools. The authors defined basic trends for edge cutting tools wear-off intensity reduction, associated with the friction coefficient value decreasing by application of the lubricant-cooling technological agents (LCTA) and wear-resistant complexes, as well as cutting temperature impact on the wear intensity in time. The cutting mode, temperature-force factor value in the working zone and contact phenomena at the cutting wedge affect the tool complexes origination (the tool material and wear-proof coating) with the effect of adaptation in the process of friction.

The article presents data on a series of experimental studies on the cutting process thermo-physics and mechanics, regularities of the cutting tool wear process while chromium-nickel parts lathe work for the qualitative estimation of the wear-resistant coatings effect on the machinability. Quadrihedral carbide plates (10 × 10 mm) and solid tools from the materials (BK8, BKIOOM) with various wear-resistant coatings were employed as cutting tools. The life testing and temperature-force tests were conducted with the I6K20 universal lathe machine of normal stiffness with stepless spindle rotation frequency control.

Temperature measuring in the process of metal cutting processing with a view to identify the average contact temperature with a sufficiently high accuracy and reliability was being performed by the natural thermocouple method. The thermo-EMF values registration and evaluation were accomplished by the mercury current collector and «Elemer» digital voltmeter. Estimation of friction coefficient and stress state of contact zone at various temperatures was conducted with the adhesion installation.

It has been established that the most favorable temperature-force state is being ensured at deposition the TiAlN of multilayer coatings after magnetic-arc filtration (MAF). Relative linear wear and its intensity decrease are being observed herewith, which can be explained by forming protective amorphous-like (aluminum oxide) and lubricating (titanium oxide) structures on the cutting wedge surface.

It has been revealed that the increase of cutting temperature and tangential component of cutting force with subsequent decrease of cutting tool wear resistance when using chromium-containing coating is associated with the phenomenon of chemical affinity of contacting materials at increased temperatures in the cutting zone.

It has been established that application of chromium-nickel alloys in the contact zone under conditions of the increased thermal power load at blade machining with tool wear-resistant complexes allows the twofold increasing of the durability period.

Keywords: wear-resistant complexes, friction, nano-structured coatings, cutting temperature and force, cutting tools wearing-out, thermo-emf, adhesive bonds strength.

Efficiency improving of the state-of-the-art techniques of welding aircraft thin-walled pipelines is an urgent task of the modern aviation industry. The main trends are the following: the welding procedure robotics, implementation of welding techniques and technologies with thermal cycle stabilization, and, accordingly, the structure and properties along the entire seam length, costs reduction of materials and electric energy, increase of productivity and quality of final products. The article presents the results of the studies conducted on the effect of the blow medium and pulsating arc while robotized argon arc welding of thin-walled elements of stainless steel piping systems for aircraft by non-melting tungsten electrode without application of the filler wire on the structure and properties of welded butt joints.

Welding was performed on an automatic welding installation for rotating bodies developed at Komsomolsk-on-Amur State University and programmable controlled by Mach3 via G-codes. The installation includes a welding rotator, a Kemppi MinarcTig Evo 200 MPL power supply with a TTC 220 burner, a positioner for the burner transverse movement, a welding wire feeder, a laptop, and a control unit. The G-code was employed for welding, the value of the standby current herewith was 15 A, the maximum current was 35 A, and the pulse duration was being reduced from 1.3 to 1.0 s within 0.1 s decrement. The extent of the first sector is the smallest with the maximum pulse duration, and is meant for stabilizing welding modes and seam geometry. The second and the third sectors are of equal extent, but with different values of pulse duration. The fourth sector is of the greatest extent with the minimum pulse duration.

A pipe from AISI 321 steel of a 50 mm diameter with a wall thickness of 1 mm was employed as blanks. The edges of the welded blanks were trimmed on a lathe prior to the assembly. The butt assembly for welding was performed manually with a gap of 0-0.1 mm on the prism without filler material application.

The developed and manufactured protective device, tightly installed in the internal cavity of the pipes being assembled through the packing rings, which seal the limited space of the butt edges, were employed for the blowing.

Geometric parameters of the obtained welding seams (the height of the reinforcement of the roller front side) were being determined by the MCAx laser scanning and 3D model processing in the Focus 10 Inspection software. Welded samples of thin-walled pipe blanks were tested for static tension and are subjected to microstructural studies and microhardness measurement.

The obtained welded joints meet by the geometric parameters the requirements of regulatory documentation governing the welding procedure of the aircraft pipeline systems. However, the joints obtained with the air atmosphere inside the pipe are characterized by a reduced tensile strength of up to 20% and elongation. Argon and nitrogen application as a blowing is being characterized by the lack of the oxidized layer, and mechanical properties closeness to the basic metal ones. Besides, a possibility for controlling the value of the root and front roller strengthening by the blowing gases pressure appears. The results of the work can be applied in the aircraft industry for both automatic and robotic welding of thin-walled stainless steel pipelines.

The properties of thin hard films with a thickness of about 30 μm deposited on a polymer coating take a significant effect on the operation properties of such composite compounds. At the same time, there are no reliable and generally accepted methods for revealing the mechanical properties of such composite compounds and their claddings, especially for the case of multilayer coatings. The mechanical tests method, which is rather sensitive to the properties of these materials, is required for the quality control of such coatings. A special method for micro-fracturing viscosity at the local loading with the Vickers pyramid was tested earlier for the single-layer composite compound.

The presented study describes a new method for the micro-fracture viscosity coefficient computing of the multilayer composite compounds. The composite compound consists of the thin hard nano-crystalline metallic films and polymeric material. The micro-fracture viscosity of a multilayer composite is being determined by analyzing the features of the system of cracks formed under local loading by the Vickers pyramid. The authors show that the recommended formulas and algorithms for the micro-fracture viscosity determining may be employed for multilayer composites mechanical tests. It is demonstrated that the micro-fracture viscosity determining of the two-layer amorphous-nano-crystal film compounds may be applied to the multi-layer composite compounds with account for correction of the fracture micro-patterns analysis method and computational formulas.

Based on the experimental data, specificity of determining the coating micro-fracture viscosity of the multy-layer composite compounds is considered for the cases when local loading with the Vickers pyramid does not allow creating the standard pattern of cracks, united into symmetrical nested figures.

The article proposes the technique and formulas for micro-fracture viscosity calculation for the cases of linear and exponential dependence of the bulge height on loading on the indentor. Specifics of the micro-fracture viscosity coefficient calculating of multi-layer composite compounds when the bulge height depends non-monotonically on the loading on the indentor, which is the feature of many multilayer composite compounds is being considered separately.

Based on the previously developed technique for determining the crystallographic orientation in polarized light, the authors propose evaluating the change in the interference pattern after the sample loading. For this purpose, a sample with decreasing cross-sectional area along the tension axis was fabricated, for loading its parts on various degree of deformation. The sample was being stretched until the yield stress was reached in the smallest section of the sample. After stretching, the local degrees of deformation of the sample sections were calculated. Three main sections with deformations of 1.5%, 5.5% and 17.5% were identified.

Metallographic section, subjected to electrolytic etching for the surface observing by the polarizing microscopy, was fabricated from each section. As was established earlier, three basic colors, namely blue, brown and yellow, which volume fractions changed depending of the deformation degree, were being observed on the sample.

The dependence of microstructure interference pattern on the degree of deformation was determined in the course of the studies for the AD0 alloy microstructure. It has been established that with an increase in the degree of deformation, the volume fraction of blue and yellow grains increases. The volume fraction of brown grains decreases, which can be explained by the fact that these grains correspond to the [110] crystallographic direction, which is more amenable to plastic deformation in the FCC lattice.

It should be noted that the volume fraction of blue and yellow grains increases by 25% at a deformation of 5.5%, while that of brown grains decreases by 44%. At the degree of deformation of 17.5%, the volume fraction of brown grains becomes smaller by another 17% compared to the 5.5%, while the volume fraction of blue and yellow grains slightly increases by 4 and 6%, respectively.

The authors propose employing the obtained dependencies to control the anisotropy and degree of deformation in the production of aluminum parts and products, as well as the express method for controlling the crystallographic orientation.

The presented article proposes a technology for various area wing design to create a family of prospective heavy transport aircraft with two and four PD-35 type engines with a thrust of 35 tons. Payload of the first aircraft could be of 70–80 tons, while the large aircraft can carry up to 150 tons. To simplify and reduce the cost of a large aircraft creating, the outer wing consoles with their engine were borrowed from the wing of the «junior» member of the family, and the area was increased due to the new center wing equipped with two extra engines. The aerodynamic layout of the wings of both aircraft was designed applying various CFD approaches, including the fast direct and robust inverse methods as well as multi- mode optimization technique.

The article presents the description of the aerodynamic design procedure and some specifics of each of the aerodynamic layouts. It is shown that the designed wings with a sweep of χ_1/4=24° do provide cruising flight at a speed of M = 0.77 ÷ 0.8 (820 ÷ 850 km/h). Two aerodynamic models of the considered airplanes have been manufactured (1:32 scale was selected for the two engine aircraft and 1:50 scale for the four engine one) and tested in the large TsAGI T-106 transonic wind tunnel. The experiment confirmed the achievement of the design goals for both cruise and takeoff-landing speed modes.

An expert assessment of the L/D ratio losses due to proposed approach to the design of a composite wing was performed. For this purpose, a free optimization of the wing of an enlarged area with the same planform and relative thicknesses distribution along the span was conducted. The article shows that the high-speed characteristics do not degrade. At the same time, the maximum L/D-ratio of the composite wing layout is ~1.5% less.

Development of aviation and weapons envisages the speed characteristics enhancement of newly developed aircraft. The requirements for test bench equipment, including braking devices employed on the rocket-rail track, are being increased. Braking expands the high-speed track tests functionality, increases their efficiency and informativity, reduces the preparation time and cost due to the reuse of the retained material part. Solution to the problem of braking rocket sleds moving along a rocket track at a speed of more than 1.200 m/s envisages the development of braking devices ensuring effective and safe braking in the entire speed range. Selection of the braking type for the promising braking device on the assumption of its technical capabilities is being required.

The article describes various types of braking employed on the rocket track facilities when testing objects of aviation and rocket technology. Technical capabilities of the conventional types and means of braking are determined including their advantages and disadvantages, as well as their application scope. Analytical study on the types of braking acceptable during high-speed track tests is adduced.

In the course of the conducted research, it was determined that braking of high-speed rocket sleds is advisable to be performed not by a single type of braking, but by several ones, applying a set of braking devices. A single type of braking is effective and safe only in a limited speed range.

Achieving hypersonic speeds on the rocket-rail track requires modernization of the technological equipment, including braking devices, as well as developing new techniques for the tests conduction.

Solutions should be elaborated to ensure braking of the objects moving under conditions of a rocket-track facility at new high-speed boundaries, as well as methods of mathematical computation of operation of the braking devices being employed should be determined.

The article provides a brief analysis of the additive technologies global market development, and bespeaks the need to activate the Russian market segment, which currently occupies no more than 2%. It regards the problems of introducing developments of the new elements of the structures of spacecraft antenna-feeder systems (AFS) and technological processes of their manufacturing employing selective laser alloyage (SLA). This said topic is insufficiently studied, since the conventional techniques are being limited only by the development of the technological process for parts manufacturing with the SLA application. The article presents the technique algorithm from technological analysis of the technical assignment for the SC AFS development to the end product manufacturing and testing. The authors note that the important feature of this technique consists in interrelation of the development process and capabilities of the parts manufacturing technology (SLA). This allows AFS manufacturing with the geometry corresponding to the rated one, which is being determined by the electro-dynamic modeling, without adjusting the part structure to the conventional manufacturing technologies capabilities. Thus, the principle of «from function to the design» is being put into practice. The technique was developed based on the authors’ experience on the SLA technology implementation and analysis of scientific publications on the issue. The authors tested the technique on the example of development and manufacturing, applying the SLA technology, of new structures of the helix antenna and waveguide corners for the spacecraft. The technique includes certification of the newly implemented material, performed according to the industry standard and consisting in conducting tests of necessary operational properties of the new material by the corresponding program.

The following documents were drawn up by the certification results:

— a certificate containing data on the properties of the material, the results of its performance evaluating under conditions as close as possible to operational conditions and recommendations for testing in production and operational conditions;

— technical specifications containing technical requirements for the material of part blanks manufactured by the SLM method.

The technique provides also the development, based on the organization Standard, of the Program for experimental try-out of technological process for parts manufacturing employing the SLA technique,

The results, obtained while developing the feasibility study, such as reduction of mass, material utilization factor, labor intensity, and cost, as well as the SC AFSs elements operational characteristics improvement, including active life increase, and new structures try-out period reduction afford ground to consider the presented article as up-to-date not only for the space industry, but for the radio-electronic industry as well.

Aerodynamic models, assumed as a rule to be very rigid, are actually subjected to noticeable elastic deformations under the wind tunnel (WT) testing conditions, which distort the measurement results. The article studies the dependences of elastic deformations of «rigid» WT models on their geometric and structural parameters to develop requirements for the model stiffness characteristics and determine rational modifications of the primary structure, which allow minimizing the model elastic twist angle for various wing layouts and flow-around modes

The procedure specifics for developing a of a steel wing computational model of the aerodynamic model in the NASTRAN software package for solving static aeroelasticity problems are considered. Parametric dependences of elastic deflections, twist angles and the lift coefficient on the wing sweep angle and position of the stiffness axis are studied.

Analysis of the results obtained for the wing model of a typical mainline aircraft reveal the following:

– the elastic streamwise twist angle is mainly determined by the bending angle;

– the angle of torsion around the stiffness axis for the model under consideration increases the streamwise twist angle.

Thus, the streamwise twist would be possible at the twist angle sign changing due to the shift of the axis of stiffness. It may also be seen from the comparative analysis of the center of pressure position of the sections along chord for the three different problems, which correspond to the pressure distribution depending on the curvature and twist at the zero angle of attack, unit angle of attack and these problems combination at different angles of attack.

For the model under consideration, in the middle and end parts, where significant deformations occur, the stiffness axis is located at a distance of (0.4-0.45)c from the leading edge (here с is the wing local chord). The sections’ center of pressure position is much further, and reaches a value of 0.6с at the wing end. This rear position of the pressure center is stipulated by the specificity of the employed supercritical airfoils with a strong undercutting of the lower surface near the trailing edge.

Thus, the possibility of reducing the elastic deformation impact on aerodynamic characteristics for a certain range of wing sweep angles and test modes due to the model layout modification was revealed as the result of parametric studies.

To minimize the elasticity effect on aerodynamic characteristics while WT test, modifications of the model layout may be considered in two aspects: 1) the relative position changing of the pressure centers line and the stiffness axis; 2) torsional stiffness reducing.

The said areas of research are supposed to be developed in the further activities on this issue.

The article presents the results of complex studies on parametric dependencies of the strength, stiffness and weight characteristics of the wing structure on the values of the set of design parameters for the regional aircraft with both strut-braced and non-strut-braced layout. A new version of the four-level designing algorithm, which employs the decomposition principle of loading cases within the framework of parametric strength and aerodynamic models while searching for the computed loading cases were used while computational studies conducting.

The article presents the description of the algorithm, which realizes the principle of inflight loading cases, employing the bond of finite element model of the airframe structure and aerodynamic parametric model based of the single vortexes method. The ability of both models for automatic dimensionality changing of loading cases allows ensuring dividing the acceptable loading cases into the groups by the degree of criticality. This ability allows also the possibility of realizing a multi-stage search procedure, when strength and aerodynamic models with low dimensionality are being used for all alternative loading cases at the first stage of the analysis, while at the subsequent stages, the models with higher dimensionality are being used to analyze the critical cases selected at the first stage.

The modified version of the algorithm demonstrated high performance and reliability for the strength analysis and design of the wing structures with high level of elastic displacements.

The efficiency of the loading cases decomposition principle in conjunction with other decomposition principles, such as structure decomposition and decomposition of the strength problems, used within the framework of the basic four-level algorithm, is demonstrated within the framework of this article on the example of the hypothetic regional aircraft of 15 tons take-off weight and passenger capacity up to 50 persons.

The values of the wing structure weight, as well as the values of the strut attachment point position on a wing (which are 50-65% of the semi wingspan depending on the aspect ratio) were obtained. The better weight efficiency of the wing structure based on the strut-braced layout compared to the non-strut-braced one was confirmed for the hypothetic regional aircraft under consideration.

Weight savings for the wing structure option with the aspect ratio of λ₀ = 11.7 is 12.3%, whereas for the alternative options with λ₁ = 15 and λ₂ = 20 the weight savings are 31% and 37 % respectively.

The labor intensity analysis of the parametric strength studies, associated with significant parameters variations of the airframe external geometry and high levels of elastic displacements of lifting surfaces, revealed that application of the loading cases principle of decomposition allows no less than tenfold labor intensity reduction of the strength analysis procedures.

The results of the performed studies have proved the efficiency of the modified four-level algorithm application for solving the design tasks for:

• An aircraft with non-conventional aerodynamic layouts;
• Regional aircraft, for which the elastic displacements impact on the external aerodynamic loads is significant.

Mathematical modeling of the aviation gas turbine engine (GTE) is one of the most important instruments, which is being employed at all stages of its life cycle. Foremost, it is being applied at the stages of engine design and its engineering follow-up

The efficiency of the engine mathematical model (EMM) application depends on the accuracy and adequacy of the working process description in an air-gas channel of the engine and its components. The accuracy of the basic engine components defining is an essential factor that determines the accuracy of the gas-turbine engine mathematical model. The engine gas turbine is one of such basic GTE components.

The firsts-level mathematical model of the engine the gas turbine represents a single-stage (one nozzle assembly and one impeller). The turbine performances are being represented as the dependence of the normalized gas consumption in the first nozzle assembly throat and efficiency on the turbine pressure ratio and reduced circular velocity value on the impeller average radius.

As is known, the efficiency reflects the difference between the real and ideal processes (without thermal losses, i.e. adiabatic expansion) in the engine turbine. In other words, it is the ratio of the power generated by the turbine to the turbine adiabatic power.

The article presents various options of the turbine efficiency determining, which differ each other by the accounting for the cooling air energy.

Analysis of the engine parameters impact on the difference between the efficiency value determined by the parameters in the nozzle throat and the efficiency value determined by the parameters in the gap between the nozzle and the impeller blades was performed. The article demonstrates that incorrect accounting for the efficiency while the aircraft GTE model computing may lead to significant errors in determining its parameters and performances.

The presented article considers the design of two gas turbine engine combustion chambers running on natural gas. There are 32 burners in the first combustion chamber, while the second one contains 136 nozzles placed in two rows in the flame tube head.

In accordance with the fact that carbon dioxide is being formed as an intermediate substance in the process of carbon-bearing fuels oxidizing, the CO emissions control is being reduced not to this substance forming prevention, but to the problem of completing reaction of its oxidation by ensuring maximum combustion efficiency.

Technical substantiation for the multi-flame fuel combustion application was set forth. If assume that the torch length is proportional to the nozzle diameter, including the number of nozzles, which equals 136, into the calculation, the torch length will be half the length of the torch length with the number of 32 pieces.

The article adduces the results of studying two combustion chambers differing by the design of the flame tube head, presents the test-bench equipment, and describes the experimental research specifics. The results of the studies on concentration measuring of the final gas mixture components at the outlet of both combustion chambers are presented. The fuel combustion completeness was determined, and inference was drawn on most acceptable flame tube head design, which ensures maximum combustion completeness and minimum concentration of carbon oxides. This design represents the multi-nozzle combustion chamber.

The inference was drawn that the combustion efficiency growth with the combustion sources increase was associated with bothr chemical reacting acceleration and substantial improving of the air-and-fuel mixture preparation prior to its feeding to the combustion zone.

The article being presented proposes the structure of the impeller peripheral part of the high-pressure single-stage centrifugal compressor with high degree pressure ratio of π_κ* = 0.9 and η ≈ 0.78, which allows reducing gas overflowing from the concave side to the convex side of the blade in its opened radial gap, as well as efficiency increasing of this compressor stage. With this end in view, the impeller end surface is bent relative to the radial direction rather than having radial direction.

As is known, the opened gap in the centrifugal compressor is much more meaningful due to its large outstretch compared to the outstretch of radial gap above the impeller of axis compressor. This efficiency reduction is being aggravated also by the fact that pressure difference in the radial gap above the impeller of the high-pressure compressor under consideration is essentially higher, and, hence, there is larger overflowing of the air being compressed from the concave side to the convex side of the blade. Installing covering disk, fixed on the high-pressure compressor impellers end butts does not solve the problem.

Firstly, in the presence of easily worn-out coating applied on the stator housing above the blades end butts, the high-pressure impeller runs with small values of the radial gap, which, in itself, reduces the air overflowing in the radial gap. Secondly, the so-called secondary airflow the concave side to the convex side of the blade passage appears on the inner side of the covering disk. This unordered secondary airflow transfers to the reverse convex side of the channel and moves along the height into the depth of the channel, which distorts significantly the computed trajectory of its flowing as well as computed exit angles from the impeller and compressed air inlet to the vaned or slot diffusor. The area of variously directed airflows shifting and their intermixing appears, which leads to the centrifugal compressor efficiency reduction.

Computational studies of seventeen options of the working blades design of a high-pressure centrifugal compressor with various angles of inclination of the peripheral part of the working blades were conducted. The inclination angle value varied herewith in the range from α_rk= –40 to α_rk= +40°. The step value of the slope changing was 5°. Geometric models of the centrifugal wheel were developed in the Ansys system. The two-dimensional model was created using the Vista CCD program, and a three-dimensional geometric model was created based on the results of the two-dimensional calculation and optimized in BladeGen.

The isentropic and polytropic efficiency of this centrifugal compressor demonstrate significant increase of about 0.2% for every 5° up to the point corresponding to the model with α_rk= +35 . Further, the efficiency growth in the computational domain decreases. Thus, the article demonstrates that there is a range of values of the inclination angles of the working blades in their end part, where gas flowing in the radial gap is reduced, and is a significant gain in compressor efficiency is obtained.

Promising engines parameters improving can be achieved primarily by significant parameters upgrading of the units and their components, such as labyrinth seals. The gas turbine engine efficiency depends on air leaks in both compressor and turbine, for which various types of seals are being used in the cooling and bleed air system. Labyrinth seals are the most common in aircraft engines. The state-of-the-art labyrinth seals are of high quality and their further improvement requires application of computer aided modeling and optimization.

The author conducted gas dynamic and strength computing of the labyrinth seal operation, and performed the labyrinth seal geometry optimization with account for the strength properties. The article demonstrates the optimization technique, which may be applied while labyrinth seal design to ensure minimum air consumption and meeting the strength criteria.

The gas flow in the labyrinth seal computing was being performed with the 1.1 pressure ratio at the rated rotation frequency of 16,000 rpm. Analysis of the circumferential speed impact on the labyrinth seal operation was performed. The circumferential speed impact on the air consumption was up to 3%.

With the circumferential velocity increase, the absolute value of the velocity in the seal gap increases, and the axial component decreases, which results in the air flow decrease through the seal. Prior to optimization, the total mass air consumption through the labyrinth seal was 8.46 g/s.

The strength calculation used boundary conditions with the pressure field on the labyrinth seal surface, obtained as the result of the gas-dynamic computation of the flow in the channel and rotation frequency. The following parameters were being calculated: total deformation, von-Mises equivalent stress, and safety margin.

As the result of optimization, the space between the ridges increases. Vortex structures emerge in the space between the ridges, caused by the action of viscous forces between the flow core and the gas between the ridges, sufficient space between the ridges ensures the vortex structures unhampered formation. More intensive vortex structures ensure, in their turn, more intensive energy dissipation, which leads to the air consumption reduction in the labyrinth seal gap. Besides this, emerging of the radial component of the velocity prior to the top of each ridge leads to the air consumption reduction as well.

After optimization, the air consumption reduction through the labyrinth seal by 16,8% was achieved at the rated speed of 16,000 rpm. Deformation and strength margin criteria were met as well. Deformation decreased by 6 %, Mises stresses decreased by 13,66 %, and the safety margin of the labyrinth seal increased by 16,13 %.

The presented calculation technique may be applied in solving problems of labyrinth seal optimization for searching for the labyrinth seal configuration ensuring minimum air consumption and meeting the strength criteria.

This article presents the results of preliminary study of influence of the Stationary Plasma Thruster (SPT) ground-based test conditions in the typical vacuum chambers on the SPT output parameters and discharge current oscillation characteristics. The SPT’s are operating already many years in space as a parts of the spacecraft (S/C) motion control systems and ensuring the S/C operation during 5-15 year service time. To reach their reliable SPT operation they undergo complex of the ground tests including that ones in the vacuum chambers imitating thruster operation conditions in space. And that it is impossible to reproduce these conditions fully. Then it s important to understand how the difference in the test conditions and those in space can influence on the thruster operation and performance. Particularly this difference could be responsible for the increase of the discharge current oscillation amplitudes after thruster switching on and further operation of the two SPT-100B in pair during their tests in the vacuum chamber in comparison with the case of switching on and operation of one such thruster in the same vacuum chamber. This event was obtained at the Russian JSC "Information Satellite Systems"(ISS) and initiated this study. Preliminary analysis had shown that the possible reason of the mentioned event is an increased release of the gases and sputtered material of the vacuum chamber wall due to their bombardment by the increased accelerated ion flow from the two thrusters. These products are able to penetrate into working volumes of the thruster parts and change the properties of surfaces of the mentioned parts such as a cathode emitter or discharge chamber walls. As a result they can change the thruster operation and its characteristics. The rate of the mentioned gas release of the adsorbed or absorbed gases and sputtered products depend on time and state of the internal vacuum chamber wall surfaces being in contact with atmosphere. Then, it is to be dependent on the history of the earlier electric thruster test before the given one because the accelerated ion flow are cleaning the mentioned wall surfaces. Taking all the mentioned into account the given investigation consisted of the study of the SPT-100 type thruster output parameters variation and discharge current oscillation characteristics in time during at least 100 hours of operation in the two different vacuum chambers of 2 m in diameter and 3m (chamber 1) and 5 m (chmber2) in length, respectively. The internal walls of these chambers had different state of their internal surfaces because the walls the chamber 1 was staying in contact with atmosphere around 6 months after test of the SPT with powers not exceeding 1kW. And chamber 2 stayed in contact with atmosphere around 3 months after test of the ion thruster model operated with power 10-13 kW and ion energies 5 keV around 50 hours. Thus, the vacuum chamber wall internal surfaces of these chambers were cleaned to different state due to different intensity of their bombardment by the ion flows during previous tests.

To estimate the rate of the sputtered products condensation on the discharge chamber wall internal surfaces there were installed the removable ring-shape internal reference samples (IRS) into the external and internal discharge chamber wall parts in between anode and eroding their parts in such a manner that they were not changing geometries of the mentioned walls. The IRS were made of the same ceramics as that of the discharge chamber. Then, to estimate the condensation rate of the sputtered from the vacuum chamber wall products on the exit side surfaces of the external magnetic poles there were installed the external reference samples (ERS) made of the same ceramics as that of the discharge chamber or made from the stainless steel. There was mounted also one Langmuir probe near one of ERS to estimate the plasma parameters near the surface of the external magnetic pole. The experimental study was made during 150 hours in the chamber 1 (cycle 1) and during 100 hours in the vacuum chamber 2 (cycle 2). During each cycle thruster model was switched on and thruster operated during 3-5 hours with the discharge voltage 300V and mass flow rate ensuring the discharge current 4,5A optimized by magnetization currents. And there was realized registration of the pressure in the vacuum chamber, thrust, discharge parameters and discharge current oscillations. There were made also periodic measurements of the plasma parameters by probe. After every 25-30 hours of thruster operation the vacuum chamber were opened and IRS and ERS were weighed.

Obtained results had shown the following:

— during 1st ~50 hours of thruster operation in the vacuum 1 there were obtained jumps of the vacuum chamber pressure after thruster switching on and there was obtained increased level of the discharge current amplitude which was regularly reduced in time. Such jumps was not observed during the cycle 2 tests;

— at the internal part of the acceleration channel walls the flows of the sputtered from the exit parts of the discharge chamber wall are drastically dominating in comparison with the vacuum chamber wall sputtering product flows;

— performance level was a little bit higher during cycle 2;

— the ceramic and metal ERS samples are slowly sputtered.

Finally, it was concluded that the most probable reason of the oscillation amplitudes increase during 1st period of the thruster pair operation in the vacuum chamber after its internal wall contact with atmosphere is the increased release of the active gases from the vacuum chamber walls being long time in contact with atmosphere under their bombardment by the increased and widened ion flow.

Development of low-emission combustion chambers for modern and advanced gas turbine engines at this date is impossible without experimental determining of their pulsation state. At the same time, ripples measuring with existing sensors at typical temperature conditions common to modern combustion chambers represents a rather huge problem. An alternative approach to this problem consists in the waveguide-type acoustic probe application, which allows removing the said sensor from the high-temperature area. The presence of a pneumatic information transmission channel places high demands on the probe frequency characteristics determining accuracy. The main feature of the probe operation as part of the combustion chamber is the temperature inhomogeneity along its length. However, the effect of the temperature distribution along the probe length on its frequency characteristics has not been fully studied by now. Thus, the main goal of this research consists in developing a mathematical model for frequency characteristics computing of the acoustic probe at the arbitrary temperature distribution along its length. The impedance method was applied when developing its mathematical model. It is assumed that the chamber represents an ideal source of pressure fluctuations, i.e. pressure ripples in the combustion chamber do not depend on the probe acoustic characteristics. The acoustic probe computational domain consists of four elements, such as waveguide, matching pipeline, sensor cavity, and adapter channel. Frequency characteristics of the sensor cavity and adapter channel, which form the Helmholtz resonator, are being computed with lumped-parameter models. This article herewith does not consider the effect of the cavity shape and the sensor impedance on the Helmholtz resonator dynamic characteristics. The waveguide and the matching pipeline are being computed with distributed-parameter models and presented as sections of the same length, within either of which the temperature is assumed constant. The temperature values for each section are being determined by interpolating the temperature distribution law along the length of the probe, which, in its turn, may be obtained by computing or experiment. Each individual section is being presented in the form of a passive quadripole. The wave process propagation constants and wave impedances for each section are being computed depending on the frequency either by applying a low-frequency model or a high-frequency one. The results obtained with the developed mathematical model were compared with the experimental data obtained at the elevated pressure. Comparison of computational and experimental data demonstrated their good convergence.

The processes of designing, fine-tuning and modernization of aircraft gas turbine engines require credibility of the mathematical models (MM) reflecting physical picture of the engine functioning processes. The latter can be achieved by the model parameters calibrating based on the engine test-bench and flight experiments results.

The MM calibration process of modern aircraft gas turbine engines is rather time-consuming task due to the need for identifying the main parameters obtained while experimental studies, which depend on a large number of parameters uncontrolled during the experiment, which values may vary while the identification process.

The presented work studies the combinatorial calibration method of the engine mathematical model. Four virtual experiments are pre-conducted, presented in the form of a model computation with introduced correction coefficients on the nodes characteristics. Global array of correction coefficients is being formed in the ThermoGTE software for the existing engine structure by the results of virtual tests. Further, the problem on the calculated parameters and experimental results minimization is being solved for each combination of correction coefficients by the ThermoGTE software built-in simplex method. As the result, an array of resulting functions is being formed for each combination of corrections, and the most accurate groups of corrections are being determined. The selected solutions operability is being checked thereafter by correction coefficients substituting into the engine mathematical model. As the result, the research engineer obtains several scenarios for the mathematical model calibration. It is assumed while solving that the parameters being measured have no deviation from the real ones (zero measurement error). The correction multipliers constancy is being assumed as well that at all engine operation modes.

The presented MM calibration method may be employed to refine mathematical model of any engine with any number of measured parameters. However, it should be noted that the presence of a large number of correction coefficients of the model under study leads to an exponential increase in the computation time, which in its turn leads to the need for the problem parallelization.

This work is up-to-date since cutting time of defects detection and elimination of civil passenger aircraft allows substantial reduction of departure delays and airlines losses associated with them. Statistical data analysis reveals that defects detecting and eliminating are the dominant causes of delays of civil aviation aircraft. The defects detection herewith takes 90% of the time. Modern aircraft is equipped with the onboard diagnostic systems. Their main purpose consists in controlling the aircraft technical state. They report on the presence of malfunction. However, they do not allow for the most part automatically localize the malfunction within the accuracy of the defect, which was its cause. The necessity for manual checking methods application employing specially developed software and hardware means arises. The time of defect detection depends on how well the algorithm for performing checks is selected. This time can be reduced if pre-elaborated searching algorithms are being placed at the technical staff disposal.

A significant effect will be achieved if and only if these algorithms are optimal by the criterion that reflects the real dependence of losses on the delay time. As statistics show, the losses grow exponentially with the increase in time spent on manual detection and elimination of a defect being the cause of a malfunction recorded by the onboard monitoring systems. In as much as the objective function is not additive, classical methods are not applicable for finding the desired algorithm. Heuristic methods do not guarantee the an optimal algorithm elaboration. Its finding by the brute force search is unrealistic, due to the huge number of possible options. The purpose of the article consists in proposing a computationally efficient method for optimal algorithms elaboration for defects detecting and eliminating, considering the exponential dependence of losses on the time of the defect detection and elimination. The algorithm is considered to be optimal if the average losses caused by the flight delay are minimal. The method for elaborating the desired algorithms based on the Bellman optimality principle proposed in this article for the first time. Previously, this approach was used only with a linear dependence of losses on the time for defects searching. Note that each combination of indications of the onboard diagnostic system has its own set of defects, with an accuracy up to which the defect that is the cause of the malfunction is being localized. The number of possible combinations of indications of the onboard diagnostic system is large. Each of them should correspond to its own manual search algorithm. Naturally, the time of its elaboration should not be too long. The proposed method satisfies this requirement. The algorithm elaboration and its presentation to a specialist may well be performed by a modern mobile device, which is not even necessarily to be a full-fledged PC. The materials of this article are of practical value for managers and employees of civil passenger aircraft operation servicing.

The spacecraft orientation stability these days is of utter importance for both public and private space agencies and companies. The growing interest to the Red Planet increases the number of space missions, which include orbital apparatuses, landers or Mars rovers. Since 1960s up to now, more than forty nine missions were sent to Mars from different countries. The majority of them end in failure, either fly far away from the Mars orbit (did not enter an orbit), crash upon its surface, do not reach the target, or connection is being lost prior to the target reaching. This indirectly indicates errors at the stages of navigation, control, stabilization or design.

The following missions are the example of failed missions to Mars, which are either lost or crashed due to failures in the navigation system, or incorrect orientation. They are 1M, 2M, 2MV, 3MV and 3MS (1960-1971), Mars-1 (1962), Mars-2 lander (1971), Mars-6 and Mars-7 landers (1973), Phobos-1 (1988), Mars Observer (1992), Mars-96 (1996), Mars Polar Lander (1999), Deep Space-2 (1999), Beagle-2 (2003), Yinghuo-1 (2011), Schiaparelli EDM lander (2016).

The presented article considers a dynamic model describing the spacecraft perturbed motion as a rigid body with significant aerodynamic and mass asymmetries relative to the spacecraft center of mass in the rarefied atmosphere of Mars.

The purpose of this work consists in obtaining an approximate discrete optimized control law of a spacecraft attitude employing dynamic programming and averaging methods. The system of quasi-linear equation was considered and averaged to obtain a simpler system of equations, which can be modeled applying the dynamic programming method.

Optimal control laws were determined based on the quadratic optimization criterion by Bellman principle, and, besides, the system of discrete equations, employing analytical Z-transform, reverse Z-transform and numerical discrete Euler method, was developed and solved. Reliability of the obtained analytical control laws is being confirmed by the results of numerical integration by the numerical Euler Method.

Euler method integration was being performed employing fixed and variable integration steps. The results obtained with a variable step appeared to be more exact than those obtained with the fixed step with the Z-transform method. The conversion behavior of both the angle of attack and the angular velocity at comparing them with the found solutions while similar studies for a significant aerodynamic and inertial asymmetry relative to the center of mass come closer to the results of this study.

The numerical results of this work confirm that the obtained approximate discrete expressions for control optimization ensure the in angular velocity and spatial angle of attack reduction to the required small values in a time commensurable with the time from the free movement start of the spacecraft uncontrolled descent to the braking parachute system initializing.

By applying these laws to a lander with asymmetries in both vehicle aerodynamics and mass, the values of angular velocity and the angle of attack will converge to zeros enforcing the stabilization.

The practical significance of the obtained discrete laws of the two-channel control is being confirmed by application of the small jet engines running in discrete mode.

The article deals with the failures of spacecraft attitude thrusters applied for angular maneuvers and rotational motion stabilization. A failure of the thruster as a part of the spacecraft attitude control loop may lead to failure of the attitude mode, high fuel consumption, exceeding the allowable loads on the structure, and even to the loss of the spacecraft. The system functioning reliability requires continuous monitoring of the thrusters running correctness in real time and timely failures parrying them in case of their occurrence.

The article considers the two possible types of failures, such as the thruster start failing, i.e. it does not start at the starting command, and the thruster turn-off failing, i.e. it keeps on running at the turnoff command. The proposed algorithm is based on the analysis of the difference between the actual behavior of the angular motion dynamics of a real control object and its onboard model. The mismatch between the vector of the measured angular rate and the vector of the angular motion estimation is being analyzed. This type of mismatch is well suited for the fact of the attitude thruster failure detecting, since it will be close to zero at any stroke of the onboard computer, while it will differ greatly from zero at the certain strokes. The thruster turnoff failure is possible to detect by analyzing the mismatch only at the strokes, where the turn-on commands for the failed thruster do not present, while the thruster turn-on failure is possible to detect by the mismatch analyzing only at the strokes where the failed thruster turn-on commands do present.

The article issues recommendations for the algorithm parameters selection. The proposed algorithm operability is being demonstrated by the results of mathematical modeling.

The article substantiates the need for the development of motion control systems for industrial robots, CNC machines and other mechatronic systems, defines the requirements for ensuring trust in such systems from the viewpoint of functional reliability and information security. One of the most up-to-date trends in the development of motion control systems for digital production is a significant expansion of their functionality for managing complex multi-coordinate nonlinear objects in real time. Practical meeting of the requirements for improving and ensuring the trust of motion control systems of industrial robots, CNC machines and other mechatronic systems can be achieved by improving the architecture of motion control systems, in particular through the application of memory-centric architecture of motion control systems. On the assumption of the specifics of control systems with memory-centric architecture, basic requirements for programming such control systems can be set. According to these requirements, the programming language should be:

— subject-oriented and specialized for motion control;

— declarative with elements of functional and logical language, optimal for setting algorithms of operation, i.e. for distributing tasks between autonomous functional modules of the control system;

— interpreted (or assembly language), ensuring the speed and compactness of the program code, as well as optimal use of shared memory resources of the control system when running in real time.

In addition, the program in the language being defined should implement the model of actors and ensure confidence increasing in the motion control system. To meet the specified requirements, the authors created a domain-oriented declarative interpreted language of a modular digital system. The key elements of the language are a set of syntactic elements, as well as application programming interfaces built from syntactic elements of the language and serving for integration into the language of external libraries (in the same or other languages). The program in language includes the following basic elements: operators, structures and expressions formed from syntactic elements of the language; actors formed as instances of additional programs emulated by the (main) program at startup or during the process of running.

The motion control system, programmed in the language, consists of four main structural components:

— A human-machine interface, through which the program code generated by the human operator, describing the algorithm of operation of the equipment, as well as a configuration file that provides program configuration for the tasks being formulated;

— A central processor responsible for the overall management of the system and distribution of tasks;

— Functional modules, performing data processing of sensitization, computations and control of regulators of actuating devices;

— Communication networks, ensuring communication between the structural elements of a computer, as well as with external devices.

As the result of the research being conducted, the mechanism of implementing the actor model through meta-programming, as well as tools for increasing confidence in the management system through management decentralization and data localization, were determined.

Machining by milling and laser cutting are widely applied in blanking-and-stamping production. However, despite the availability and simplicity of the milling method, laser cutting is being increasingly employed in production.

The laser cutting method ensures high productivity of the process in combination with high precision and quality of cut surfaces, as well as a small cutting width. However, one of the significant disadvantages of laser cutting consists in the presence of a temperature-affected zone in the area of laser beam impact, which leads to a change in the material properties at the edge of the billet and, as a result, to a decrease in the fatigue resistance of parts.

Fatigue test samples were cut from a 2.5 mm thick cold-rolled sheet of the D16AT alloy across the rolling direction. One part of the samples was fabricated by milling, while the other part was produced by laser cutting. The samples were tested for high-cycle fatigue in bending at a symmetrical cycle, and the test base was of three million loading cycles. The loading threshold of the three samples without their destruction was being estimated. Besides, after laser cutting the samples, were subjected to etching in Keller’s reagent to eliminate the defective layer formed as the result of laser processing.

The result of the samples fatigue testing revealed that the conditional endurance limit of the samples obtained by the laser cutting method was 55MPam which was 60% lower than the one for the samples manufactured by milling, which was equal to 90 MPa.

The metallographic results allowed revealing that the end-butts of the samples manufactured by the laser cutting method contained the defective layer associated with the metal overburning, which was the cause of the conditional endurance limit reduction. To remove the metal layer with overburning etching was employed, which allowed partial restoring of the conditional endurance limit of the material equal to 80 MPa. In this case, the conditional endurance limit is 18% less than that for the milled samples.

Thus, the conducted study reveals that during the products operation obtained by laser cutting, premature fatigue failure may occur under cyclic loading conditions. To eliminate this possible defect, formed as the result of manufacturing by the laser method, the defective metal layer with overburining should be removed. The defective layer removal will lead to the increase fatigue resistance of the products.

The article presents the results of the study of inertial friction welding of the EP741NP nickel heat-resistant alloy. Since not all metal materials can be welded by melting methods, there are alternatives, for example, mechanical welding, namely friction welding. A literature review oriented to studying the preceding similar experience on optimal welding modes selection, which was the prime task was accomplished. Due to the friction welding optimal parameters selection more strong and qualitative joints are being produced at the lower temperatures and less heat-affected zones (HAZ) than while fusion-welding methods employing. The major part of works was being conducted on the foundry alloys, since their grain structure is more pliable for these kind of impacts, while our work studied the alloy, being obtained by the metallurgy of granules method. The task is being aggravated by the fact that the alloy itself is liable to the great risk of crack formation while moulding methods employing in view of utterly complex chemical composition. Based on the above said we come to the conclusion that the purpose of the presented work consists in achieving stable, high level of strength, no less than that adduced in the said review. It was established in the course of the study that this welding process allows obtaining the joints, which are not being obtained by the melting welding methods. Welding was being performed on the PSTI-120SW installation. Afterwards, rods were cut out for the samples production and templates for the obtained joint structure studying. Further, experimental study of the samples’ mechanical properties was conducted. The experiments were being performed with the Instron universal breaking machine. The samples were being subjected to the tests on the short-time strength at the room temperature, and the long-time strength (for 100 hours) under the load of 900 Mpa at 650C. As the result of the test, the samples demonstrated rather high qualities as it was predicted. This fact allows our alloys and equipment competing with their foreign counterparts. In the course of this research, the authors studied the microstructure of the weld seams and weld-affected zones. Transverse metallographic samples containing welded joint were prepared for the research. The microstructure analysis was being performed with metallographic microscope. The structure of the samples from the EP741NP alloy is granular. Three types of the strengthening -phase are being observed, and it is noted that no defects on the macrostructure are detected in both joint and weld-affected zones. The study of the weld workblanks from the EP741NP alloy revealed the absence of porosity and cracks in the basic material and thermally affected zone. The welded joint up to 200 microns, irrespective of the final material shortening. Transition zone (a zone of the thermal affect) of 500 microns to 1000 microns is being observed. The heat treatment conducting after the welding contributes to the strengthening phase exudation in both seam and weld-affected zone. As the result, the welded joints become equal in strength to the basic material, which will be the next stage of materials treatment in the further research.

The article deals with lattice thin-walled load-bearing shell elements equipped with an external sealed shell and applied in civil aviation as aircraft fuselages. Analysis of the existing experience in lattice composite structures design and application as appllied to both spacecraft and atmospheric aircraft is being performed. Composite skin together with composite bearing ribs, ensuring the structure aerodynamic quality and the aircraft internal volume tightness are being employed as a rule in the said structures.

The flight speeds increase, as well as possible shock impacts from objects of various nature, do not only hinder, but also make composite skin of aircraft elements application potentially impossible, whereby the authors propose to apply metal alloy skin in a lattice thin-walled shell structure.

The article proposes a technique for the design stage calculation of stiffness characteristics of lattice anisogrid structures with metal sheathing, which allows solving the problem of optimal design of this kind of structures by increasing their weight perfection. Comparison of the results obtained by analytical solving with those of the numerical experiment is being adduced.

As it follows from the results obtained, the presence of a metal edging does not only serve as a solution for creating a reliable mechanical linkage between the metal sheathing and the composite load-bearing element, but gives some increase in both flexural and membrane stiffness as well. The proposed method for stiffness characteristics determining and its verifying employing the finite element method (FEM) demonstrates the fundamental possibility of designing and calculating composite elements, such as beams, anisogrid plates and shells containing a metal edging or metal sheathing. It can be applied not only in aerospace designs, but also in the field of ground structures developing, as well as shipbuilding.

Synthesis and study of nano-powders properties of various oxide materials is an important problem in modern materials science. The quartz glass possesses a unique combination of characteristics, such as high melting point, high heat resistance, chemical inertness, and optical transparency. The said stipulates the interest in the preparation and study of nano-powders from this material.

One of the successful methods for obtaining nanopowders is the solid inorganic substances evaporation under the action of electron beams with subsequent condensation. The purpose of this work consists in to analyzing the process of obtaining silicon dioxide nano-powders based on high-intensity evaporation of quartz glass under the impact of CO2-laser radiation (10.6 μm), and studying the products of this reaction.

The subject of the study in this work is the process of evaporation from the surface of quartz glass under the impact of focused laser radiation from a CO2-laser of a 100 W power. The evaporation rate was being controlled by changing the laser beam power, the rotation speed, and the linear feed of the quartz rod. With the CO2-laser power increasing up to 5 kW, Si02 nano-powder obtaining with a productivity of up to 5 kg/h is possible.

Radiation, absorbed in a thin layer of quartz glass, heats it up to evaporation temperatures without the liquid phase formation. Quartz glass evaporates in the plasma state.

A brightly luminous plasma, which leads to the NO₂ gas forming, is being formed in the process of quartz nano-poweders obtaining in the zone of material evaporation under the impact of the focused laser radiation.

The sizes and phase composition of nanoparticles, as well as the specific surface area and optical properties of nanopowders, were studied. The spherical structure of the quartz powder particles is visible, which indicates a liquid-drop mechanism of evaporation. The size distribution has its maximum at 80 nm. The chemical composition of the silicon dioxide powders corresponds to the chemical composition of the feedstock, and, unlike industrial grades of silicon dioxide powders, they do not contain chlorine and fluorine.

Analysis of the obtained silicon dioxide nanopowders application revealed the possibility of their employing in high-quality polishing, cleaning, grinding friction pairs of high-precision mechanisms technologies, as well as an additive in composite polymer materials and lunar soil simulators.

The article considers the thermo-gas-cyclic nitration method, consisting in alternating the process stages with high and low nitrogen potential, being conducted at temperatures, respectively, below and above the temperature of the eutectoid transformation in the Fe-N system. At the half-cycle with high saturation capacity of the atmosphere at the low dissociation degree of ammonia, a high-nitrogen nitride zone is being formed on the surface. It transforms into an extended γ’-zone and an internal nitrating zone due to internal diffusion in a cycle with a low saturating capability of the atmosphere, or at a high degree of ammonia dissociation. The processes with alternate changing of the nitrogen potential contribute at certain stages to accelerated growth of the nitrided layer. Besides, this allows controlling the process and obtaining the required combination of phases, determining these or that product properties, necessary for various operation conditions, namely:

— The presence of a high-nitrogen nitride zone on the surface contributes to the running-in of friction units and, some cases, increases the corrosion resistance;

— Under wear-out condition at the increased specific pressures, the multi-layer structure from the surface nitride zone bearing on the internal nitriding zone, appears to be the most steadfast one;

— The extended zone of internal nitriding with minimal surface nitride layer should be formed for the parts operating in the dynamic wear-our mode and shock loading.

In some cases, such as corrosion-resistant steels nitriding, a diffusion layer based on an internal nitriding zone (solid solution) without a nitride zone is advantageous.

The control principle is based on maintaining the nitrogen potential at the level of values corresponding to the solubility of nitrogen in a given phase of the Fe-N system. Chemical-thermal treatment with alternate supply of ammonia and air (gas-cyclic process) allows fourfold duration reduction of the diffusion layer forming process of a specified thickness in alloyed steels. The phase composition of the surface layer after various nitriding process modes and kinetics of its individual sections growth were studied.

Possibilities of intensifying nitriding and controlling the phase composition of the layer by a rational choice of process parameters, namely the number of half-cycles and their duration are shown.

Electron sources have found their application in many fields of science and technology. In ion-plasma technologies and electro-propulsion engines (EPE), the electron source is applied as a cathode-neutralizer. Besides, it is employed as a plasma contactor that ensures the electric charge discharging from the body of a spacecraft, such as the (International Space Station) ISS.

Most electron sources, being applied, are based on the thermionic emission phenomenon. The disadvantage of such emitters is many factors limiting their resource. The resource of such electron sources decreases even more when the latter are employed in the processes with reactive gases.

However, there are gas-discharging electron sources or plasma cold-cathode electron sources. A glow discharge or a Penning discharge are being most often used in such sources. The effect of a hollow cathode is being used as well. Thus, such an emitter is referred to as a cold hollow cathode (CHC) in many applications. The disadvantage of the CHC based on self-sustained gas discharges is high operating voltages.

The CHC presents interest when working with reactive gases. The studies of alternative working substances for electric thruster (air, iodine) require the design further development of the thrusters including cathodes.

The presented work conducts the studies of the cold hollow magnetron cathode performance (CHMC) for the electric thruster, and performs energy efficiency comparison of various cathode material – working gas combinations.

The following factors affecting the CHMC energy efficiency were studied in the presented work:

1. The working gas flow rate. The article shows that maximum energy efficiency is being achieved by maximum possible flow rate of the working gas.

2. The magnetic field magnitude in the hollow cathode. The study revealed that maximum energy efficiency is achieved at maximum value of the magnetic field.

3. Combination of the cathode material and working gas. The article demonstrates that the CHMC performance characteristics depend significantly on the cathode material and the working gas type. To demonstrate capabilities of the cathode applied consumption as a cathode-c neutralizer for the electric thrusters, the unit operating characteristics were obtained while running on gases, such as xenon and air.

Thus, the experiments on the presented design of a hollow magnetron cathode have revealed the fundamental possibility of obtaining an electron current to compensate for the charge of the ion beam of the electric thruster. However, the device efficiency compared with the thermionic cathodes employed now is low. It has been demonstrated experimentally that all the ways, being described, of the energy efficiency increasing are limited by the operating voltage of 300 V. This limitation corresponds to the theoretical models of magnetron discharge.

To reduce the operating voltage threshold, the authors are planning the electrode system modification, such as, extra ionization stages application with non-self-maintained discharges.

The problem of aviation gas turbine engines protection from foreign objects damage (FOD) casted into them when the aircraft taxiing on the airfield surface is well known. The article regards one of the reasons of foreign objects casting into the engines, namely foreign objects casting by the aircraft landing gear wheels on takeoff and landing modes. To avoid engines damage by foreign objects during operation, it is relevant to assess the engines protection already at the stage of preliminary aircraft design. The conducted airfield testing studies revealed a relationship between the of engines protection from the damage by foreign objects casted by the landing gear wheels from the surface of the airfield and the power plant layout. Thus, the of the power plant layout on the aircraft allows assessing the engines protection at the design stage. If the assessment reveals that the engines protection is not ensured, then it is necessary to develop structural measures aimed at achieving the necessary protection level. Protective devices installed on the front landing gear wheels to protect the engines from the FOD casted by landing gear wheels have become widespread. However, it is necessary to assess the possibility of ensuring the protection of engines by changing the power plant layout, before employing such protective devices. There is a throw-out zone of foreign objects behind the landing gear wheels when the aircraft is taxiing around the airfield. If the inlet edges of the engine air intake unit are in the throw-out zone, the foreign objects may be casted into the engine.

The distance between the front landing gear wheels and the inlet edges of air intake unit has a great effect on the probability of foreign objects thrown-out by the landing wheels, into the engine. The probability of casting the foreign objects decreases while the inlet edges of the air intake unit approaching the front landing gear wheels. At a certain distance between the front landing gear wheels and the inlet edges of the air intake unit, the probability of foreign objects being thrown-out becomes zero. Such power plant layout should be considered as the most appropriate for the engines protection ensuring. However, the problem of engines protection ensuring by the front landing gear wheels approach to the inlet edges of the air intakes is closely connected with the landing gear scheme, namely with the location limits of the landing gear struts relatively to the aircraft center of mass. The power plant layout changing by shifting the front landing gear at the required distance to the inlet edges of the air intake unit may lead to an unacceptable change in the aircraft landing gear scheme and going outside the accepted restrictions. If the aircraft power plant layout changing is impossible, the only way out remained is employing protection devices installed on the front landing gear struts.

The article considers a bypass burner device design for a low-emission combustion chamber of a gas turbine engine running on natural gas. The results of the two burners differing in the swirler flow area studying are presented.

The burner device modification consisted in changing its design by installing a cowling on a swirler, which allowed reducing its flow passage area. As the result of the cowling installation, the swirler channels overlap by 38% occurred compared to the original option. The basic idea of such modernization consisted in forming an expanding channel from the swirler inlet to the nozzle outlet.

The article presents the bench equipment and specifics of the experimental study. The results of the studies on the final gas mixture concentration measuring along the length of the flame of the two burners are presented as well. The said studies revealed that the modernized burner device allowed twofold CH level reduction, i.e. the fuel underburning reduction. Thus, the discussed burning device has been selected for installation into the combustion chamber.

The combustion chamber fire tube refining was performed by organizing an extra air feeding on the walls through elaborating an extra number of orifices. Pressure losses in the combustion chamber, as well as temperature field at the outlet of both stock and modernized combustion chamber were determined. As the result of computation, the excess air ratio behind the flame tube head in nominal rating mode for the NK-38ST gas turbine engine was 2.1 for the for the stock combustion chamber, while it was 1.8 for the modernized one.

The results of the tests revealed that efficiency increase in the whole range of the ambient temperature was being traced for the engine with modernized combustion chamber.

The presented article deals with studying the possibility of applying the free fall launching method of a small-size UAV for application from the UAV-carrier. This task is up-to-date since the possibility of the UAV application in air operations depends on its solution.

The research is being conducted by a binding of two programs, namely Euler and SimInTech. Euler is being used for cargo flight dynamics analyzing and displaying output values of angles and speeds. SimInTech receives the output data from Euler and applies it to computer aerodynamic and interferential forces and moments that are being transferred back to Euler.

The results of the conducted studies under various conditions revealed that, the UAV starts rotating rapidly while free falling. At the initial stage of the flight, the UAV rudders are ineffective and unable to compensate the increasing angular velocity of the cargo. This leads to the fact that on achieving the speed enough for the rudders become effective, the UAV angular speed will become so large that the stabilization system would be unable to stabilize it. The application area of the obtained results is military one.

Based on the obtained data, a proposal to employ gas-dynamic devices for the cargo stabilization at the initial segment of the flight was put forward. This method seems more feasible since of ailerons or wings installation on a small-size UAV is problematic due of its small size. Besides, in contrast to the other methods of stabilization, gas-dynamic devices do not increase the UAV weight that much, which is an important factor for aviation engineering.

Stability loss of the thin skins under loads close to the operating level is allowed for the upper panels of the low-capacity aircraft wing-box. The article proposes an applied technique for determining optimal parameters of thin metal skins with account for the two levels of loading. At the first level, the problem of stability ensuring of a rectangular panel with a minimum margin is being considered. The relations of geometrically nonlinear optimal design problem of the panel under postbuckling behavior are being written for the second level of loading. The article presents also analytical relations explaining the place of the design methodology for the supercritical state in the general theory of optimal design of thin-walled aircraft structures. It considers the design technique, which accounts for the interrelation of the two above-said problems. The panel thickness and width were selected as the variables of the general optimization problem. It is noted, that the optimal design problem proposed in the article differs from the traditional options by the said features. The article presents the panel design techniques based on analytical solutions of geometrically nonlinear problems when considering various options of loading a thin rectangular panel with hinge support. For the cases of compression and shear, compact analytical relations for the optimum parameters determining, which can be recommended for use in the early stages of design when selecting design solutions, are obtained. The longitudinal compressive and shear flows impact at combined loading was considered. In this case, a general option of the optimal design methodology is presented. For the second level of loading, the article regards also various static strength criteria and presents corresponding analytical expressions for computing optimal width of the panel at compression and shear. To illustrate the technique, the article presents numerical examples of determining optimal thickness and width of metal panels in compression. Conclusions and possible variants of the practical use of the technique are presented. As an example, an option of determining optimal parameters of a multi-web flap is given.

The ground-based facilities are being subjected to flight check-ups at putting into operation, in the process of operation and certain special cases for checking parameters and characteristics of ground-based flight support facilities correspondence to the specified operational requirements. The existing techniques application is, in some cases, cumbersome, for example at operational airfields, where operational deployment of radio-technical flight support facilities and their putting into operation is required. The situation may be drastically aggravated under condition of various intended and unintended destabilizing factors impact, including terroristic groups. Not only the failure of technical facilities herewith, but losses among the crew of the aircraft-laboratory are possible.

In this regard, the purpose of the study consists in developing a technique for flight check-ups to ensure their running under conditions of possible destructive impacts on the aircraft-laboratory, its crew, as well as flight check-ups operative organizing.

The set goal pursuing is being achieved by an unmanned aircraft application instead of a manned aircraft-laboratory, as well as by excluding ground means of trajectory measurements from the flight check-up procedure.

The basis of the proposed method of flight checks of ground-based radio-navigation means is to determine the module of difference between the measured value of the ground-based means parameter and its set value for each set point of the unmanned aircraft flight; to correct the flight trajectory taking into account the value obtained at the previous step; to re-flight the unmanned aircraft on the corrected trajectory.

The following items underlie the proposed technique for the flight check-ups of the ground-based radio-technical aircraft flight support utilities:

– Determining the absolute value of the difference between the measured parameter (of characteristic) value of a ground-based facility and its set value for each set UAV flight point;

– The flight trajectory correction with account for the value obtained at the previous step;

– The UAV reflight along the corrected trajectory.

The number of repeated flights is being determined by the required measurements accuracy.

The article presents a technique for flight check-ups conducting of ground-based radio-technical aircraft flight support facilities employing the UAV, which does not require the ground-based trajectory measuring facilities. A flight control device and a simulation model for the glissade radio beacon testing have been developed. Analysis of its application possibility was performed based on the simulation. The article demonstrates that the landing glissade coordinates determining accuracy is being determined by the coordinates determining accuracy by the UAV.

The proposed method allows

– Excluding the ground means of trajectory measurements application during flight checks;

– Control equipment deployment onboard an unmanned aircraft;

– Performing the UAV flight control of an unmanned aircraft during flight checks-ups without signals from the ground-based radio-technical aircraft flight support facilities.

This will allow reducing operational costs, the number of personnel involved and ensuring high operational readiness of the facilities involved.

Maintaining the specified safety, reliability and availability characteristics of the aerial vehicles (AV) with long operation life and after-sales service, can significantly exceed their purchase cost. Conceptually new approaches are required nowadays in the industry to ensure the quality improvement level, increase in the safety and economic efficiency of the AV for the aviation industry enterprises. Highly efficient AV with low life cycle cost (LLC) and high utilization factor are economically viable for the aircraft operators (consumers). One of the ways of the LCC reduction consists in optimizing the aircraft maintenance system during operation, refurbishment and overhaul.

Manufacturing companies that are among the first in the aviation industry to integrate predictive maintenance (PM) into the after-sales service (AS) and maintenance and repair systems (MRO), all other things being equal, will be able to provide the most competitive product in the aviation industry. This concept implementation is complicated since the PTO concept involves continuous monitoring of a large number of parameters, which does not allow fully implementing it in the aviation industry due to the lack of global broadband data transmission from the aircraft throughout the entire flight.

Mathematical method of artificial neural networks (ANN) application is the least costly for the incoming big data streaming analysis.

The gist of the ANN utilization consists in processing the information array obtained from the product state monitoring system to predict the available solutions on the product maintenance.

The way to the MRO optimization is integration with the Aircraft Health Monitoring (AHM), in which, the ANN employing as a tool is one of the concepts.

The authors propose application of the developed model of the aircraft maintenance and refurbishment for the ANN utilization, with the ANN employing as a predictive maintenance tool.

As of today, small gas turbine engines are of significant commercial potential in minor power engineering and aviation sectors. However, little attention is being paid in Russia to the issues of the small engines creating despite of the significant experience in the gas turbine engines design and wide infrastructure for their production. A small-sized engine creation, meeting requirements of both power engineering and aviation, will allow necessary energy generation in close vicinity of the place of its consumption. This will significantly reduce transportation losses, and allow, in prospect, making both heat and electric power supply system’s more dynamical and adaptable to the needs of a certain consumer, as well as loading idle production capacities of many aviation plants.

The proposed method for radial clearances determining allows identifying the compressor and turbine rotor and stator behavior more accurately under conditions of high temperature and pressure differences, as well as at various operating modes. With account for the obtained deformations, the radial clearance optimal value may be obtained, as well as both compressor and turbine thrust and efficiency can be computed. This method may be applied as well to the full-sized gas turbine engines and gas turbine plants. However, transient operating modes are characteristic for the gas turbine engines, which necessitates non-stationary gas-dynamics computations performing.

The rotor and stator 3D models obtained in NX CAD and being imported to the ANSYS, where finite element models were created, are being employed for the computational time reduction. Next, computation of gas dynamics is being performed in Fluid Flow (CFX), in which the heat exchange between the working fluid and rotor and stator parts is accounted for, is being performed. The obtained results are being transferred to the Steady-State Thermal for temperature fields distribution computing over rotor and stator, and further to the Static Structural for determining rotor and stator deformations from various factors impact, such as thermal expansion, pressure differential at the back and trough of the vanes, as well as centrifugal forces.

It was determined while computations that the compressor and turbine parts thermal expansion exerts the greatest impact (up to 99%) on the radial clearance. This is associated with the materials employed, as well as high temperatures and large drops in the engine operation.

It is necessary to ensure a radial clearance of at least 0.15 mm to prevent the rotor from touching the stator during transient operating modes at the maximum operating mode. With account for the obtained deformations in the compressor, this condition is being fulfilled at the maximum operating mode with the radial clearance is of 262.04 µm from the side of the leading edge and 274.95 µm from the side of the trailing edge. The authors suggested increasing the mounting radial clearance to 0.4 mm in the turbine. In this case, radial clearance in the turbine at the maximum operation mode will be 250.46 microns from the inlet side, and 183.2 microns from the outlet side.

Seen from front, the wing shape is being characterized by the wing deflection angle, which usually has negative values in the aircraft parking position for the swept wing aircraft, which is realized according to the high-wing of mid-wing scheme. The wing root herewith is located higher than its cantilever (end) part. With the said shape, changes in the deflection angle sign from negative to positive are possible in process of the flight.

One of the negative consequences of this change is the residual fuel flow-over from the cantilever part of the wing to its root.

The following tasks are being solved in the course of this study:

– Analysis of the wingtip displacements on the ground and in flight from the loads affecting the aircraft wing;

– Detecting causes of fuel mass readings changes in the non-fueled wing tanks;

– Clarification of fuel automation mathematical models based on the results of the analysis.

It was analytically proved by the analysis results of the loads affecting the wing in the aircraft parking and flight position, as well as in the takeoff and climbing modes, that:

– A possible fuel mass increase in the wing tanks in the aircraft flight position was not associated with the fuel automation operation errors, but it was stipulated by the residual fuel flow-over in the wing tanks from their cantilever part to the root one due to the positive wing deflection in flight as affected by the lifting force;

– A possible fuel mass decrease in the wing tanks in both takeoff and flight modes is being stipulated by the residual fuel flow-over in the wing tanks from the root part back to the cantilever one due to the negative or zero wing deflection, formed by the force of inertia under the aircraft vertical acceleration impact.

The obtained results may be employed for clarifying the mathematical models, by which the fuel automation computes the fuel mass in the tanks, with account for the fuel flow-over in the wing tanks during the aircraft flight.

Currently, the role of unmanned aerial vehicles (UAV) has risen sharply in the field of aircraft building, and the scope of their application herewith is regularly expanding. This type of aerial vehicles is not at a stop, and has been actively developing in recent years. One of the ways of the UAV development consists in enhancing its resistance to the multifactor uncertainty. Multifactor uncertainty is being understood as uncertainty, stipulated by the uncontrolled factors action. It is worth noting that uncontrollable factors incur a significant impact on the design procedures results and design as a whole. In the most general case, the set of possible states of the uncontrollable factors vector will generate an equal to itself by the size set of optimal solutions.

In retrospect, this problem was being solved for the UAVs and aircraft in general by introducing a number of assumptions and special project regulations being formed based on the experience and designer’s subjective perception. The “standard atmosphere” model, rated values of the materials strength etc. may serve as an example of such approach, though, objectively, there are always certain differences from these conditions. For such difference compensation and possible degradation of the aircraft operation, an excess (safety margin) is being admittedly provided in the aircraft capabilities with respect to the design conditions, which frequently leads to the aircraft weight and cost increase. These safety margins are not scientifically substantiated and being elaborated purely empirically. In general, this approach is distinguished by subjectivity. This subjectivism may be eliminated to a certain extent, if the UAV possesses the properties of uncontrollable factors resistance.

There is a whole number of stability studying methods, however, the most convenient and widespread method is Lyapunov function method, though it is imperfect and has a number of disadvantages. The most grave disadvantage of Lyapunov theory consists in the fact that in the general case the Lyapunov function should be guessed. The direct Lyapunov’s method in the stability theory is basic for the stability studying of dynamic systems. However, the Lyapunov function definition does not directly relate to structural properties of the system under study, and, thus, there are still no exhaustive regular ways to its construction according to the given equations of the aircraft motion.

This work novelty lies in the fact that the UAV stability is being studied by a new constructive method of the Lyapunov function statistical synthesis. The statistical synthesis method is being applied to restore functional dependencies from the statistical data. Actually, the original problem of the UAV stability studying is being reduced to a nonlinear programming problem with a statistical stability criterion, by which the optimal design solution is being selected. Statistical synthesis is based on the three basic elements such as statistical sampling, basis functions and statistical criteria. As the result of the conducted study, the following results were obtained:

1. A method of stability studying for a wide class of the UAV-type aircraft has been developed.

2. The stability of the UAV movement was studied according to the developed statistical criterion.

The thermal control system (TCS) is intended for maintaining the required thermal conditions of all spacecraft elements and onboard equipment.

The spacecraft TCS designing is a significant part of the spacecraft general engineering. This is due to the fact that the TCS is a deeply integrated spacecraft system interrelated with main onboard systems, environment, structural elements and flight tasks.

It is necessary to account for the thermal loads from the onboard equipment, radiation and re-radiation from the Sun and planets, and many other factors while designing a spacecraft. With relatively small thermal capacities, the spacecraft has a leaky design and the TSR is being designed on passive means of thermo stating. Application of thermal models with lumped parameters is widespread in the design of spacecraft onboard equipment. This approach appropriateness is confirmed by the practice of various units of a spacecraft TSR electronic equipment designing, analyzing and testing. The presence of telemetry parameters creates the possibility and directions for techniques optimization for the spacecraft TSR with improved qualitative mass-energy characteristics design.

The most common liquid TSRs display the essential fault in terms of specific mass-energy characteristics due to the greater mass of a coolant fueling, employing only heat-capacitive heat accumulation, as a consequence of the vapor phase inadmissibility at the contour centrifugal pump, though both models and heat balances of such systems are elaborated enough.

The presented article deals with an approach to the design of structural schemes for the spacecraft thermal control system with passive coolant pumping with of at least 3 kW of thermal power productivity. Three options were considered herewith.

The first option studies application of the thermal control system based on heat pipes, installed on the radiating panels. The heat-emitting devices herewith is installed on the backside of the radiating surface, and heat pipes distribute the heat along the panels’ surface transferring heat from one panel to the other.

The second option suggests the device in the form of the central heat bus, in which the heat-emitting devices are located on the common cooling panel, and uncontrolled heat pipes are embedded into the board being cooled and carry the heat from the electronic equipment to the passive heat transfer device in the form of the capillary pump.

The heat transfer unit of the third option does not contain flexible pipelines, and connects the electronic equipment board with the emitting radiator by the rigid pipelines. To provide the possibility for temperature control of the board being cooled, the heat pipes’ condensing zones of the cooled board and emitting radiators are connected by the gas-regulated heat pipes.

As far as the system with passive coolant pumping is under consideration, such criteria as energy consumption, operability range, control accuracy and reliability for all options are practically the same, and dominant evaluation criterion is the mass, which computing for all three options is presented. The computational results revealed the first option advantage, for witch specific mass-energy characteristic was ~33 kg/kW (without considering the ration of a certain part of the mass to the load-bearing structure mass).

The results of the performed comparative analysis allow drawing a conclusion that at the spacecraft equipment thermal load up to 3 kW, the most optimal is the thermal control system, which design scheme is based on application of the exclusively axial heat pipes.

Since its advent, the multimode military aviation evolution, both in Russia and in other countries, tends to expand the boundaries of aircraft flight characteristics. The impressive range of modern engines operating conditions for super-maneuverable modern aircraft fighters incessantly increases all types of loads on the load-bearing elements of turbojet bypass engines with an afterburner. The task of military aviation consists in the capability to operate under conditions of frequent and sharp operation modes changes, as well as ensure long term fault-free operation under the impact of maximum loads on the engine. Thus, the progress of aircraft engine building is impossible without enhancing the structure stability to the increasing loads, or, if possible, reducing the impact on the load bearing elements of the engine. The purpose of this work consists in studying methods for constructive reduction of axial forces acting on the high-pressure rotor bearings, and defining the most effective one. For this purpose, comparative analysis of various types of turbojet engines air systems was performed from the viewpoint of the axial forces balance. As the result of studying the load-bearing schemes and various structural solutions, the gas generator of the engine-prototype with the most effective air system was selected. The hydraulic design procedure of the air system was performed according to the presented technique. Computing of axial forces, acting in the engine-prototype at four different modes was performed on its basis. The computational results reveal that the axial force values acting on the high-pressure rotor bearing comes closer to their limits, acceptable for the required service life ensuring. Further, a comparative analysis of the axial forces distribution in the engine optimization techniques was conducted. This allowed selecting the most effective one, according to which measures on the axial pressures changing in the inter-disk cavity were proposed. This, in its turn, allowed obtaining tangible increase in the force, acting on the rear part of the high-pressure turbine disk necessary for the reduction of the resultant loading of the high-pressure bearing, without principal, laborious and costly structure changes, as well as significant increase in the cooling air consumption. This solution is optimal for the set problem of the bearing unloading from the axial forces, and will allow prolong the engine fault-free operation under conditions of maximum loading or sharp changes in the operating modes.

In the paper the concept of the distributed power plant (DPP) is considered at its integration with the long range aircraft (LRA).

The given propulsion system consists of a turbine bypass engine (TBE) which turbine is connect with two taken out fan modules with the help of the mechanical transmission. The mechanical way of power transfer is the level of airplane 2030 and based on results of the researches CIAM of P.I. Baranov of new circuit designs.

As the advance design of the long range airplane with DPP is observed the aircraft type “hybrid flying wing”. Two distributed propulsion systems take place on the top of an aft tail of the plane.

The DPP parameters definition is the result of computer model of the given power plant system. According the calculation, the average cruise value of inlet total pressure recovery coefficient is about ~0,958.

In the paper is presented the adaptation of the computer model for distributed propulsion system to adapt for the process of multidisciplinary optimization.

For heightening efficiency of remote fan’s modules on different conditions of flight are examined controllable blades of these fans.

In view of the big magnitudes of total compression ratio of perspective DPP (≥50) core engine was considered the two-shaft scheme. TBE has the two-position nozzle of bypass duct for displacement of an operating point on performance of the fan to have near optimum of efficiency.

The component efficiency level of the DPP is defined on the base of the forecast of development of aircraft engines for perspective long range aircrafts of commercial aviation 2030 years.

The computer model of the DPP is developed using the block-structure and separate blocks created earlier in CIAM first level mathematical model of turbine engines.

Thus the block-structure of a bypass unmixed engine has been changed by accessing blocks of remote fans. The DPP compressor and turbine groups’ calculation is added by the corresponding equation of balance of fans and turbines powers.

In the paper the system of defining equations for DPP computer model of the design and off-design modes as aero thermodynamic characteristics is presented.

The description of computer model of estimated DPP turbo machinery weight and weights of gearboxes and transmission shafts is given.

The given adaptation of model provided possibility in an automatic regime to vary the basic data on settlement (cruiser) regime DPP. Also it provided the calculation of aero thermodynamic and ecological characteristics for further researches of LRA and DPP and receiving results in the necessary aspect.

With given computer model optimizing DPP for aircraft type “hybrid flying wing” researches has been conducted. Carried out researches have allowed to determine two alternative versions of the DPP providing smaller runway length (on 4 %) and the best parameters on issue СО₂ not conceding base version on range of flight and expenses of fuel.

The State economy effective functioning largely depends on the transport capacities of civil aviation, which ensure the required volume of passenger and commercial cargo transportation. It is especially important for Russia, with its large and remote regions of the Far North and the Far East. Establishing dozens of new routes on domestic and local routes will predictably lead to the significant growth of transportation by regional and short-range passenger airplanes.

In the current situation of the domestic air transportation development in Russia, the problem of the aircraft line expansion of all needs of this market segment coverage has not been completely solved. Thus, the development and creation of new regional and short-haul aircraft and aircraft engines for their power plants keeps on being an urgent task.

The article solved a complex task of searching for the optimum set of design parameters and characteristics of the technical system “Aircraft-Power plant”, in which capacity a twin-engine short-haul (regional) aircraft with the flight range of 2000 km and a power plant based on the two-bypass turbojet engine in the takeoff thrust class of 25 kN was taken.

The universal technique for technical layout forming and efficiency evaluation of the aircraft power plants of various purpose, developed and many times officially accepted at the Department of Aircraft Engines of the “Air Force Academy named after professor N.E. Zhukovsky and Y.A. Gagarin” was employed as the technique for the studies conducting. The instrumental “Airplane-Engine” software package, which realizes the complex approach while forming the engine technical layout, i.e. the engine, power plant, airframe and flight trajectory parameters and characteristics are being regarded in the aggregate, underlie the said technique.

Development of the power plant with two-bypass turbojet engine was performed based on the TV7-117C gas generator turboprop engine, and the Yak-40 aircraft as the airframe prototype, to which structural changes were introduced to meet the specifications on the flight speed and height.

The technical parameter of an aircraft level, namely average fuel consumption per kilometer, which directly depends on the specific fuel consumption and determines the flight range, was selected in the presented work as an optimization criterion according to the problem conditions.

The performed optimization studies conducted employing the indirect statistical optimization method based on the self-organization resulted in the selected target function increase by 7%.

The practical value of this work lies in the fact that its results may be employed by:

– scientific and design organizations involved in the development of advanced passenger aircraft and engines for their power plants;

– ordering organizations and industry while justifying the requirements for new aircraft models, as well as in aviation engineering universities to improve educational process.

Flight tests of new parachute systems often lead to an increased landing speed of weight models with an unacceptably high value of landing overload and loss, along with the layout, of both test materials and expensive flight test equipment. This makes employing a rescue parachute system as a part of a weight model along with the parachute system being tested. The said rescue system should be in constant readiness to its application, and the experiment should be planned so that urgently identify a critical failure and run the rescue parachute system in case of emergency. The presented work is devoted to the cargo rescuing parachute systems development.

The issues of flight test equipment certification for large-area parachute systems were considered in detail in [1], particularly, the requirements for weight models that act as weight equivalents of the landing cargo. Weight models are also being equipped with costly sensors, measuring and recording equipment employed for qualitative and quantitative assessment of the tested parachute system functioning.

Flight tests of new parachute equipment, as a rule, are of a high risk of the parachute system failure during its operation with all subsequent negative consequences following this, i.e. accidents of weight models and irretrievable loss of valuable information and expensive equipment.

To preserve the integrity of the weight models, besides the parachute system being tested, which characteristics have to be studied, they should be equipped with the block of parachutes of the rescue parachute system, which is being run in case of the tested parachute system failure.

The task consists in assessing the possible causes, as well as scenarios of the emergencies occurrence and development, possible outcomes in cases of failures in the functioning processes of the tested parachute systems, options for the emergency parachute systems bringing into action and the rescue system selection for the weight model.

The studies of weight models rescuing were being conducted for the first time in [2-4].

The presented article regards in detail the following issues on the task being considered:

– The requirements laid for the rescue parachute system and its functioning specifics;

– Ballistic calculations performing and phase trajectories developing for the weight model free motion;

– Cascading of the system, and determining the canopies areas of the parachute cascades;

– Examples of computations and phase trajectories plotting;

– Minimum permissible height determining of the introduction of the main and braking parachutes of the parachute rescue system;

– Specifics of phase trajectories plotting with account for possible emergencies;

– Development of the flight operations implementation programs logic for the automatics of the rescue parachute system operation control system.

The goal of this work consists in continuing and developing the studies started in [2-4].

The subject of the article relevance is stipulated by the presence of fundamental possibility of solving the axis of sight stabilization problem of the optical means positioned on the movable base of the unmanned aerial vehicle under conditions of low stabilization accuracy of the gyroscopic platform at rapid u-turns, vibration and aerial vehicles maneuvers.

The purpose of the research of the article consists in accuracy increasing of the axis of sight of optical devices installed on a gyro-stabilized platform of an unmanned aerial vehicle.

The object of the study is the optical surveillance system of an unmanned aerial vehicle.

The subject of the study is the process of objects determining by the optoelectronic system of an unmanned aerial vehicle.

The novelty of the research is stipulated by the development and scientific justification of an optical surveillance system of an unmanned aerial vehicle, as a part of television and thermal imaging information channels, a laser rangefinder-designator, as well as mathematically described method for optical surveillance system stabilizing.

Practical significance lies in application of an unmanned aerial vehicle optical surveillance system for objects capturing and tracking by the operator, as well as for objects automatic capture and tracking.

The article presents a block diagram of the gyroscopic stabilization system, as well as mathematical formulation of the problem of the optical surveillance system stabilization of an unmanned aerial vehicle.

The stabilizing method of the optical surveillance system of an unmanned aerial vehicle for determining objects, which allows independently estimate the speed and angles of departure of the biaxial gyrostabilizer platform based on the information on the nature of the platform stabilization system gyroscopes movement is substantiated. The stabilization problem solution is based on building an asymptotic optimal observer (identifier) of the biaxial gyrostabilizer state variables with incomplete stabilization coupling. It was assumed herewith that the system was under the effect of statistically indeterminate disturbances.

In general, the simulation revealed the possibility of employing the said algorithms to evaluate the initial position of the platform and calibrate systematic components of the platform departures of the biaxial gyrostabilizer under conditions of a movable base. 

Further trends of the research are the methods for images informativity increasing for identification and auto-tracking of the target detection objects by the unmanned aerial vehicle optical surveillance system in abnormal conditions associated with periodical images distortions.

The article presents a brief overview of the accidents occurred in the past due to the aircraft pilot induced oscillation (PIO). It proposes the alternative algorithm for the nonlinear pre-filter (oscillation suppressor). Compared to the other pre-filter versions, the proposed filter is being installed inside the flight control system contour, and its output serves as an input signal for the actuator. According to its algorithm, this signal does not decline when its output signal (δ) is equal or less than rate limiting(δ_max). However, when δ exceeds δ_max, δ decreases according to the developed algorithm.

Effectiveness of the proposed pre-filter is being compared with the other two pre-filters versions. One of them is the traditional nonlinear pre-filter, which algorithm corresponds to the simplified actuator model. Its input signal is proportional to the control stick deflection. Another nonlinear pre-filter is so-called “rate limiter with feedback and bypass” developed by the SAAB Company for the JAS-39 aircraft.

The following two types of experiments were conducted:

– PIO suppression effectiveness comparison by various nonlinear pre-filters and of error reduction in the tracking task in case of precise knowledge of the actuator model parameters;

– Robustness evaluation of the proposed pre-filters.

All experiments were being conducted at one of the MAI flight-simulators. The piloting task consisted in pitch tracking task with the tracking error-minimizing goal. The dynamic configuration corresponded to the statically neutral aircraft with feedbacks ensuring the HP2.1 dynamic configuration from the Have PIO database with no nonlinear effects impact. The actuator simplified model parameters corresponded to ±15 deg/s and gain coefficient K = 10.

The experiments revealed that in case of piloting without pre-filters, the unstable PIO process occurs. Installation of whatever pre-filter allows suppressing the diverging oscillation. However the proposed nonlinear pre-filter ensures the of the of error variance decrease by2.35 and 1.95 times and higher bandwidth of closed-loop system compared to the conventional pre-filter and so-called “rate limiter with feedback and bypass”.

The experiments on robustness studying demonstrated that the inaccurate knowledge of the actuator model employed in all pre-filters algorithms does not affect practically on the results of experiments in the case of the proposed pre-filter. As for the other pre-filters, the inaccurate knowledge of actuator model parameters considerably affects the error variances and other pilot-aircraft system characteristics.

The author proposes an innovative option of emergency fly-by-wire servo-control to preserve controllability at both hydraulic systems failure for a prospective regional aircraft with fly-by-wire control system and two hydraulic systems. Two electro-hydraulic servo-actuators (EHSA), fed from the two independent hydraulic systems, and servotab with electromechanical actuator (EMA) are being installed on each main control surface. With both hydraulic systems failure, all EHSAs enter the passive mode (damping mode), and switching to servotabs emergency control occurs. The servotab deflection produces a hinge moment, which in its turn deflects the control surface. The aircraft handling qualities in the servo-control mode should ensure the capability of the safe flight termination.

Mathematical model of the control surface rotation under the impact of the external hinge moment, originating while the servotab control, was developed for computational and test-bench studies with account for the specifics caused by friction and damping effects from the electro-hydraulic servo-actuators operating in passive mode. The damping force value significantly affects the aircraft handling qualities in servotab control mode.

The results of numerical studies revealed that in order to meet the AMC CS-25 25.671(c) requirements for manoeuver capabilities after failures and the MIL-STD-1797 recommendations for maximum allowable phase lag between control stick pilot input and control surface response, the servotab control laws should contain speed-up pre-filters on pilot control signals, pitch rate feedback (elevator servotab control law), roll and yaw rates feedbacks (rudder servotab control law). The emergency servotab control algorithms parameters selecting, ensuring the set requirements meeting at various values of the EHSA damping coefficient, was performed.

To confirm the possibility of the safe flight termination with the selected servotab emergency control law parameters, the test-bench tests on the flight simulator with participation of test pilots were conducted.

The approach and landing tasks with glideslope offset correction and with crosswind W_z = 5 m/s were under study. According to the pilots’ opinions, the aircraft handling qualities in servotab control mode correspond to the Cooper-Harper rating PR=4.5...5. Slight PIO tendency noted mostly in roll channel corresponds to the PIOR=3...3.5. The obtained pilot ratings confirm the correctness of the emergency servotab control algorithm parameters selection and the possibility of the safe flight termination in this mode.

Materials forming or forging is being complicated with their development. This complexity concerns the movements that need to be performed by the output link of the machine (press or hammer). Besides the purely translational movement, which was characteristic to the first hammers, as well as the purely rotary movement, which dates back to the time of the first rolling mills (XIX century), forming machines of the early XX century were able to combine translational and rotary movements. This is how the processes of spherical or orbital forming, based on incremental or sector approach, allowing producing the parts of hub and flanges type without the need to employ the equipment of high deforming force, appear. On the other hand, the development of heavy machinery and control systems allows creating presses with mechanical and hydraulic systems that form one or more output links, to apply servo control as well as schemes from robotics and create flexible forming systems. The material flow can be improved by increasing the total deforming volume per time step or the intensity of deformation, for example, by torsion with forging.

As the article shows by the finite element (FE) simulation in the QForm of the “bevel pinion” forging without teeth working out, rotating tools allow:

– Reducing peak deformation force,

– Creating in material media the required thermal characteristic for the material propitious flow;

– Obtaining the shape with specified contour offset from the required geometry;

– Reducing the stress-strain state and tools’ wear.

The 3D geometry of both the tool and the workpiece, boundary conditions setting, corresponding to the technological conditions of process and non-linear characteristic describing of the material hardening in the process of its deforming are being required for numerical simulation. The computations duration depends upon the basic computing duration and duration of the problems being additionally solved, such as simulation of the stress-strain state of the forming tools. In other words, numerical simulation by the finite element method depends on the number of equations of the system being solved in the mesh points, which number is being determined depending on the degrees of freedom, characterizing the actuator movement, as well as rheological description of materials.

The article deals with the development of the structural theory of strength and design on its basis of various technological schemes for surface hardening of steels and alloys. The basic principles of dislocation theory are also presented here, according to which the resistance of real metals to plastic deformation being expressed by the strength characteristics (yield strength σ t and tensile strength σv), is higher, the lower the dislocation mobility is, i.e. the more barriers are in its path. On the other hand, the ductility and toughness of metals are being reduced herewith, leading to the brittle fracture as the result of the possible initiation and progressive development of a crack. Hardening of real metallic materials is being considered as the result of the dislocations interaction with a certain combination of several types of obstacles, or as a combined effect of several structural mechanisms, namely hardening by interstitial or substitutional atoms (solid solution hardening), hardening by grain and subgrain boundaries, hardening by dislocations, and hardening by dispersed particles. Contribution of these mechanisms to the overall hardening may vary greatly depending on the class, brand of metallic material, as well as on the technology employed. The approximation of linear additivity of various mechanisms is generally accepted and confirmed by the concurrence of calculated and experimental results for certain classes of steels.

This article adduces a calculation of the of the alloying elements impact in austenitic steels and alloys on the level of solid solution hardening, which is the predominant mechanism of structural strengthening in this class of austenitic steels while nitriding. It is worth noting that nitriding is one of the most widespread chemical-thermal treatment processes in mechanical engineering. The structural strengthening while formation solid solutions forming occurs due to the deceleration and blocking of dislocations by atoms of the dissolved element owing to the Cottrell atmospheres formation, which increase the stress required for dislocation glide, i.e., cause hardening. Hardening level prediction based on computational models allows associating the material structure with the yield strength and fracture toughness as the main indicators of the structural strength of a product, as well as maximally implementing the main of hardening mechanisms order to develop new effective technologies for creating materials with desired properties.

The objective of the presented article consists in studying impacts of various options of the boundary layer control (BLC) system on characteristics of spatial uncontrolled air intakes. The spatial supersonic uncontrolled air intake of external compression with an oval inlet was developed in the course of this work. Three different options of the boundary layer control system were developed for this air intake. They are:

1. The transversal slit on the compression wedge in the throat area.
2. The transversal slit in conjunction with perforation on the side surfaces in the inlet area.
3. Perforation accomplished in the form of the open-ended elliptic ring on the compression wedge and side surfaces in the area of the air take inlet.

Numerical study of the flow-around physical specifics and characteristics of the isolated oval-shaped air intake without the BLC system, as well as with all developed options of the BLC system was performed. The air intake flow-around was modeled based on numerical integration of the Reynolds-averaged Navier— Stokes equations (RANS) employing non-structured computational meshes, generated in the areas of the flow outside and inside of the air intake. The air intake duct throttling was modeled by the active disk method.

The results of the computational modeling are presented in the form of graphs of the air intake characteristics dependencies and flow patterns in various sections of the air intake channel. These graphs present dependencies of the total pressure recovery coefficient v on the air mass flow rate through the engine f, as well as circumferential distortion parameter dependence on the specific reduced air mass flow rate through the engine q(engine). The Mach number fields in both longitudinal vertical and longitudinal horizontal sections of the air intake channel, as well as fields of the v coefficient in the channel cross section, corresponding to the inlet of the engine compressor, are presented in the flow patterns.

Analysis of the obtained results of the computational study revealed that all developed options of the BLC system ensured the air intake characteristics improvement. The v coefficient herewith increases, and the parameter decreases compared to the basic option of the air intake. It was determined that the third option of the BLC system ensured the greatest characteristics augmentation. Besides, this option of the BLC system ensures maximum length of the horizontal section of the air intake throttle characteristic.

Based on the results of the performed computational study, the high level of characteristics of the air intake, equipped with the third option of the boundary layer control system was established. This is associated with the positive effect of the total pressure losses reduction, when the part of the flow passing through the diagonal shocks of the -structure of the terminal shock wave, leaning against the BLC system element, namely the perforated section of the air intake internal surface.

The article presents also the results of the computational and experimental studies of the isolated spatial trapezoidal air intake of the external compression, equipped with the BLC system in the form of perforation on the surfaces of the compression wedges in the area of the channel inlet. It is demonstrated that the detected positive effect of the -structure is being realized while the trapezoidal air intake flow-around as well.

Elimination of the influence of the wind tunnel test section walls on the flow over the model is one of the important problems in experimental aerodynamics. The flow near the model placed in the test section of the wind tunnel is different from the flow existing over the model in the unbounded flow. The shape of the streamlines is distorted at the location of the model due to the presence of the test section walls. The problem of interference between the model and the walls becomes most urgent due to the phenomenon of the test section blockage in transonic wind tunnels with solid walls. The using of permeable (perforated or slotted) walls of the test section is the most common method to reduce wall interference. However, permeable walls allow only to reduce their influence on the flow over the model, but not to completely exclude it. In addition, perforation is a source of low-frequency noise, large-scale eddies are generated due to slot boundaries.

Jet boundaries have been shown to be effective compared to existing methods to solve the wall interference problem in transonic wind tunnel. However, this approach has not become widespread due to the technical complexity of the jet installations implementation.

The approach based on the using of a controlled boundary layer is quite effective and technically easy to implement that is shown both experimentally and numerically. However, in some cases, the tested models are oversized, and the thickness of the boundary layer turns out to be insufficient to eliminate the solid wall interference.

A new approach to solve the wall interference problem is presented in the paper — combined jet-perforated boundaries. The proposed method combines the advantages of perforated boundaries and the controlled boundary layer. In addition, it is technically easy to implement, economically profitable and does not exclude the possibility of using it in existing wind tunnels.

Experimental study was carried out with a drained symmetric NACA-0012 airfoil with a chord 150 mm in TsAGI T-112 wind tunnel.

The experiment was carried out in solid walls with spoilers, in perforated boundaries with an open-area ratio of 0%, 2%, 10% and 23% and in jet-perforated boundaries with similar permeability coefficients and the spoiler height of 30 mm. The Mach number was 0.6; 0.65; 0.7 and 0.74. The angle of attack varied from −4° to 6°. As a result, the pressure distribution was obtained. The main aerodynamic characteristics of the model were calculated based on the obtained data on the pressure distribution.

This article presents the results of the airfoil model characteristics under the unbounded flow that was conducted in ANSYS CFX software by numerically solving the Reynolds averaged Navier — Stokes (RANS) equations. The SST turbulence model was used for the approximation. Numerical calculations of the flow over the NACA 0012 airfoil were carried out under conditions corresponding to the experimental one (Mach number: 0.6; 0.65; 0.7; 0.74; angle of attack: 0°, 1°, 2°, 3°, 4°).

The analysis of the results made it possible to draw a number of conclusions about the possibility to reduce the wall interference in transonic wind tunnel by using jet-perforated boundaries. It is shown that with relatively moderate level of disturbances introduced into the flow by the model (at Mach numbers up to 0.74 and angles of attack from −4° to +4°), the optimal combination of the perforated wall with the open-area ratio of no more than 2% with the controlled boundary layer generated wedge-shaped spoilers with a height of 30 mm (10% of the test section half-height of the T-112 wind tunnel). The selected combination of parameters made it possible to practically eliminate wall interference when the models’ chord does not exceed 25% of the test section height. The perforation ratio or boundary layer thickness should also increase with the increase in the model size or lift force.

At the present stage of aviation development, the main way to the transport aircraft wing load-bearing characteristics improving is application of high-lift devices of the leading and trailing edges of the wing. By now, the high-lift devices of the trailing edge with the Fowler type single-slotted flap became widespread. The endeavor to simplify the high-lift device structure at preserving its effectiveness led to the advent of high-lift device of the wing trailing edge, in which the tilt flap and descending spoiler are being applied. Equipping modern long-distance aircraft with bypass turbojets of high and ultra-high bypass ratio complicates the high-lift device layout in the «low-wing monoplane» scheme. Ensuring the required minimum clearance between the nacelle and runway surface leads to the distance reduction between the wing and the engine, while the wing interaction and the high-lift device interaction with the jet exhaust leads to the drag increase at the cruising flight and noise increase on the takeoff-landing mode.

The article presents the results of experimental study on the application effectiveness of adaptive high-lift device employing the model of aircraft with high-wing monoplane, equipped with two solid propellant engine nacelles of ultra-high bypass ratio.

Aircraft model tests were performed in a subsonic wind tunnel at a flow velocity of V = 40 m/s, corresponding to the Reynolds number value of Re = 0.89·10⁶, on mechanical six-component balance in the range of angles of attack of α = –6 ÷ 24° at zero slip angle. The model tests were conducted for the following options of the flap: δ_F = 30°, δ_F = 40° and δ_F = 30°/20°. The spoiler droop (adaptive element) in the tests deflected by the angles δ_SD = 0, 8, 12°, the relative height herewith of the gaps between the wing and the flap was 2.5%, 1.2%, 0.6%, respectively.

The above said experimental studies revealed that the adaptive element application together with a single-slot retractable flap allows obtaining high load-bearing characteristics close to more complex double-slotted flaps at lower drag. The adaptive element deflection leads to a significant increase in load-bearing characteristics by 25–45% in the area of takeoff and landing angles of attack α = 8·10°, and maximum wing lift increase coefficient compared to configurations without deflected adaptive element. Disadvantage of adaptive element application consists in critical angle of attack value decrease by  Δα = 2÷4°. However, the lifting force coefficient changing herewith of large angles of attack goes smoothly. Geometric parameters optimization of the adaptive element may reduce the above said negative impact.

Optimization of the geometric parameters of the adaptive element can reduce this negative impact.

The presented article is devoted to the studies being performed on rotating strain-gauge balance calibration measuring six components of the total aerodynamic force and the moment of forces acting on the aircraft propeller during an experiment in wind tunnels.

The article describes basic principles of multicomponent aerodynamic scales calibration, working formulas computing, errors determining and other criteria for calibration quality evaluating.

The calibration machine prototype, by which calibration of the strain-gauge balance was performed, was considered. The article presents the technique for the strain-gauge balance working formulas obtaining by the least-squares method in the matrix form for three types of mathematical models, namely 6×27, 6×33 and 6×96. Analysis of the mathematical models quality was being performed by such criteria as absolute, reduced and relative and errors, authenticity and standard error of the regression coefficients.

The authors indicate and analyze the trends of methods and tools development for processing the results and strain-gauge balance loading to improve calibration accuracy. Methods of optimal experiment planning and artificial neuron networks application both for calibration results processing and calibration work benches control relate to these trends.

The largest reduced error was 0.50% for the mathematical model with the 6×27 dimensionality. The error for the 6×33 model was 0.32%, and 0.2% for the 6×96 model. Calibration error of 0.2% conforms the best world samples of rotating strain-gauge balances.

The obtained results allow developing a technique and recommendations for static calibration of rotating strain-gauge balance for characteristics measuring of aircraft propellers and can be accounted for while developing new design schemes of strain gauge balance. Besides, the obtained data are the scientific and technical groundwork for creating a dynamic calibration machine for strain-gauge balance calibration in rotation. Such work bench is necessary, for example, to account for the centrifugal force impact on the strain-gauge balance readings.

The article performs an integrated assessment of the Earth remote sensing (ERS) spacecraft (SC) rational parameters and development program in the period under consideration with account for technical-and-economic limitations. The problem of rational parameters assessment of the ERS space system (SS) modernization program is being solved. The problem specialty consists in the fact that the initial state was determined, namely the base object (ERS SC).

The authors proposed a technique for integrated assessment of a spacecraft rational parameters and development program, based on the multilevel design management multilevel project study models and statistical method of multilevel consistent optimization. This technique includes a stagewise solution of rational parameters integrated assessment of a spacecraft as a part of the ERS SC in the considered period. The first stage solves the problem of parameters assessment of the ERS SC modernization program. The second stage solves the problem of the spacecraft rational parameters assessment with account for design work solutions for its subsystems.

The article presents the developed algorithm for integrated assessment of the spacecraft rational parameters and development program, as well as basic relations of the project models. The design work analysis specialty of the spacecraft development program in the considered period is a complex nature of the research. A system rational structure is being determined herewith simultaneously with the subsystems (spacecraft modifications) project parameters, as well as the system modernization program, namely the date and terms of modernizations performing in the considered period. The dependencies reflecting the basic ERS SC characteristics (weight and cost) changing on the system technical characteristics were formed by both correlation and regression methods based on the posteriori (statistical) information of the ERS SC samples-prototypes characteristics. The article adduces the results of the various options of the modernization programs studying. The considered (being forecasted) time period is of twenty years. In contrast to the third one when only one modernization is being performed with four spacecraft modifications, the first and the second options comprise performing two modernizations. The difference between the first and the second options consists in the number of the spacecraft modifications. The first option contains four modifications while there are three of them in the second one. The performed quantitative esteems of the total reduced expenditures on the modernization program realization in the course of twenty years reveal that the second option, at which the expenditures are minimum and of 1.154 billion of conventional unit is rational. The cost saving is 12.5–30% compared to the first and third options of the modernization program.

The article demonstrates that the system modernization in the considered period and the search for rational project work solutions is being performed in a complex and consistent manner with the spacecraft parameters assessment as well as parameters of the spacecraft subsystems being replaced. This complex studies allow accounting for the functional relationships (both internal and external) dynamics, and determining rational solution on the term extension of the ERS SC effective application at the restricted costs.

The developed technique allows performing technical-and-economic analysis of the ERS SC modernization program alternative options and obtaining necessary quantitative assessments while project solutions of the spacecraft modifications assessment and selection, as well as assessing the unified space platforms application effectiveness and enhancing the operational life of subsystems and a spacecraft as a whole. The developed technique may be applied for the ERS SC development programs correction and determining requirements to the prospective spacecraft and its modifications.

The presented study is focused on the experimental study of impact resistance of integral polymer composite panels with lengthwise framing. In the course of the work, the character of impact damages in the area of the skin attachment and stringer under the impact of various kinds of the impact energy was studied, and these damages effect on the panels residual carrying capacity was evaluated. The effect of adding the extra layers of polyethylene plastic with higher energy absorbing properties on the panels’ impact resistance was estimated as well. Samples of panels were fabricated from the two types of materials, namely carbon fiber-reinforced polymer (type C) and a combination of carbon fiber reinforced polymer and polyethylene (type D).

A testing methodology selection substantiation was performed in the course of this work. An ins ert with cuttings for integral panel for longitudinal framework was fabricated for the testing with standard rigging. From the incomplete destruction conditions of the integral panels, the impact energy was of 2 and 10 J. The impact is being inflicted in the zone of the skin reinforcement to the stringer, since the damage in this area should lead to a greater strength reduction of the panel at the post-impact loading. Tests of integral carbon reinforced plastic panels revealed no visual damages on the panels at the impact of 2 J. The impact of 10 J leads to the partial internal and interlayer damages from the opposite side in the place of the skin transition to the stringer.

Static tests on longitudinal compression were conducted after the impact resistance test to determine residual strength of the panels. As far as the samples are of various shape and cross-section area, comparison was being made by the absolute maximum loading val ue, sustained by the sample at the longitudinal compression. The impact of 2 J did not affect practically the strength properties of the samples. Maximum force reduction while all type of samples destruction is no more than 10%. The impact of 10 J leads to drastic damages of all types of panels. The residual strength of integral carbon panels is 63%, while it is only 60% for the combined panels.

The results of the experiment demonstrated that combination of materials with different properties, such as carbon fiber-reinforced polymer and polyethylene, may increase impact resistance of the part as it prevents crack growth and fracture of the material from the damage initiation area on the skin to the frame.

Waveguide ducts are the integral units of microwave devices in space technology, and, besides the specified radio-technical parameters, they require ensuring their dynamic state with account for heating. One of the most important parameters determining the dynamic behavior of the extended waveguide structure under the combined impact of forced vibrations and heating is the values of the first natural vibration frequency and the critical temperature of stability loss. The presented work considers the issues of controlling the first natural vibration frequency and critical temperature as applied to the spacecraft straight waveguide ducts by the developed technique of the supports arrangement substantiated choice. The author suggests the techniques for solving direct and inverse problems, allowing both determining the first natural vibration frequency and critical temperature at the specified fixations, and selecting the structure of the supports arrangement, which will ensure these parameters of the waveguide dynamic state. The example of the straight waveguide duct computation and comparative numerical calculations, which demonstrated good convergence of the results, were performed with Ansys software. The developed techniques are of a general character, and they may be employed at both checking calculation and developing any kind of straight beam structures for controlling their dynamic state by the supports arrangement.

For loading reduction while landing the aircraft landing gear are equipped with the damping system, consisting, as a rule, of shock absorbers and tire pneumatics. Various landing gear structural schemes are employed on modern aircraft. Dynamic calculation of the landing gear is one of the most important tasks of the aircraft design. It is advisable to employ numerical simulation method of an arbitrary holonomic system motion of rigid bodies using the Lagrange equations of the first kind to simulate the damping system of the landing gear of various kinematic schemes.

This approach differs from the previously used techniques, such as application of the Lagrange equations of the second kind, written in generalized coordinates by:

• The versatility of the approach when modeling landing gear struts of various kinematic schemes;
• Representation of the landing gear strut model in object form, e. as a set of objects: rigid bodies, force factors and mechanical constraints, which allows formalizing and automating the process of a landing gear model developing, and ensures modularity and extensibility of models.

The article considers the landing impact simulation of the mainline plane main landing gear. The landing gear model consists of the three rigid bodies: the wheel, the shock absorber rod, and the shock absorber cylinder, together with the loading on one strut. The model includes seven mechanical constraints. Three force factors are set in the model as well. They are the force of pneumatics compression Pw, the axial force in the shock absorber Psh and the lift force Pl.

The landing impact calculation of the landing gear was performed for the case of absorption at normal operational work. Computational results were being compared with the experimental data of impact tests being performed in the Department of dynamic strength of Siberian Aeronautical Research Institute.

The landing impact parameters of the landing gear calculated by the proposed technique are consistent with the results of drop tests within the experimental error, which confirms the good agreement of the mathematical model with the real object.

The theoretical studies considered in this work reflect the development of thermal insulation protective means applied on the aircraft. The purpose of the work consists in studying the possibilities of enhancing thermal insulation characteristics of the aircraft being operated under extreme temperatures. Namely, the article tackles the option of a multilayer structure suggested as a thermal insulator for its application on the aircraft. This structure consists of the composite material layers, porous material and aluminum-magnesium alloy layers. Theoretical study of heat exchange of this structure and existing thermal insulating structures employed on the aircraft is being conducted for comparison and evaluation of the considered multilevel structure application effectiveness.

The extreme temperatures are being determined in this work from the aircraft flight mode conditions, at which these excessively high temperatures occur.

The thermal conductivity studies of the proposed multilayer structure and conventional heat-insulating structures considered in this work were being performed numerically by the finite-difference method.

The numerical study results of the unsteady thermal conductivity revealed that a multilayer structure was twelve times superior in thermal insulation to all other existing thermal insulation structures considered in the work. Besides, the results of studying thermal conductivity of the structures under consideration demonstrate that:

• The layers of materials in the element do not operate separately from each other, but they all operate in the common heat exchange system;
• The monotony of the temperature distribution in the elements depends on the of the materials’ thermal conductivity coefficients ratio.

The results of this work may be recommended for application in real designs of the state-of-the-art aircraft.

The ingress of foreign objects or birds into the engine, interacting with structural elements of gas turbine engines, leads to the compressor blades damaging and, depending on the degree of the damage, contributes to the incidents or accidents occurrence in the process of gas turbine engines exploitation. Due to the leading edge damaging of the compressor working blade, the profile chord reduction and radius changing of the entry edge occurs, which finally leads to the damaged blade flow-around by air character changing.

The article presents computations for determining the compressor characteristics changing, its gas-dynamic stability margin and the mass flow while operating in the engine system under the impact of damages in the form of dints. The NUMECA Fine/Turbo CFD code, which realizes the numerical solution of the Navier-Stokes equations averaged by Reynolds for computing the three-dimensional air flow in the compressor, is employed for this problem solving.

The commercial NUMECA Fine/Turbo software product allows quantifying the impact of damage on the compressor operation quality.

Damage in the form of a dint leads to the reduction of local values of pressure increase, efficiency and gas-dynamic stability margin of all compressor operation modes. The gas-dynamic stability margin lowering increases with the blades chord length decreasing. The modes, at which the gas-dynamic stability decrease takes maximum values occur at n_pr = 80%, 85%.

The dint curvature affects the quality of the compressor, that is, it leads to the gas-dynamic stability margin decrease due to a change in the character of the damaged blade flow-around by the air.

An increase in the number of damaged blades leads to a decrease in the compressor gas-dynamic stability. In the modes when n_pr = 80%, and n_pr = 85%, the gas-dynamic stability decreases significantly.

With a sequential arrangement of damaged blades, the gas-dynamic stability of the compressor decreases, compared to the case of inconsistent arrangement due to the turbolization of the boundary layer intensity increase.

The introduction to the article is focused on the problem of early diagnostics of the aircraft gas turbine engine bearings. Particularly, the gas turbine engine bearing functioning period disrupts namely at its early developmental stage, which does not always succumbs to estimation by the conventional methods. The authors suggest employing the apparatus widely known in medicine practice to analyze the occurring quasi-periodicity, namely rithmogram and scatterogram.

A rithmogram plotting is being realized based on the developed technique. The technique in its turn bases on the correlation processing principles, wavelet transform theory and Hermite transform. Briefly, the gist of the technique consists of the following: mutual correlation function of the studied signal of the bearing and reference function is being computed. The reference function is being plotted based on Hermite transform, and represents mirror reflection of the impulse characteristic of the complex quasi-matched filter. Wavelet processing principles application (scaling parameter variation) allows refining positions of the correlation function peaks. After the cross-correlation function threshold processing we obtain rhythmogram and scatterogram of the signal under study.

Further, the article considers processing of real signals of gas turbine bearing. Spectral and statistical analysis of the obtained rhythmograms and scatterograms is being performed. Inferences are being drawn on the state of the bearings under study.

Conclusion considers further prospects of the rhythmograms and scatterograms application as diagnostics tools for aircraft gas turbine engines.

Parameters measurement accuracy while gas turbine engine (GTE) tests is incurring direct impact on the tests quality and engine parameters setting-up during its pilot and serial production. Considerable attention while testing is being paid to the accuracy of the engine main output operating parameters determining such as thrust and specific fuel consumption, since these parameters directly affect the aircraft flight characteristics. However, accuracy of these parameters actual values determining while the GTE bench testing is being affected by many factors, the main of which are the aerodynamic characteristics of the test-bench box. Determining the test-bench aerodynamic characteristics impact on the engine thrust is being performed in accordance with the Industry Standard OST 101021-93 «Test-benches for aircraft gas turbine engines. General requirements» and according to the «Aerodynamic force at gas turbine engines tests on the ground-based closed test-benches» measuring technique adduced in the OST 1 02781-2004 Standard. However, this technique is applicable only to the turbojet and turbofan engines with common nozzle on the supercritical operation mode at π*_nozzle ≥ π*_{nozzle crit}.

The purpose of this work consists in developing a technique for the aerodynamic force value determining as a correction to the force from the engine thrust. This value is being measured with the force measuring system in the (closed) box of the test-bench based on comparing the bench-testing results of the GTE with a large degree of double-flow with separated circuits under condition of H = 0 and M = 0 at two layouts of the inlet lemniscate device. This technique proposes determining the reduced value of the aerodynamic force determining for the selected GTE type on the steady-state modes of the engine operation at the constant value of the reduced rotor rotation frequency n_r cor = const in the (closed) box of the test-bench in two options. The first option supposes the layout with mechanically connected lemniscate (the reduced thrust of the test-bench R_eng.cor is being determined with no account for the values of the input impulse ΔR_inlet and aerodynamic drag ΔR_windage), employed while acceptance bench-test. The second option employs the layout with the lemniscate mechanically disconnected by the labyrinth seal. The reduced thrust of the test-bench R_0eng.corr is being determined herewith with account for both the input impulse in the section of the labyrinth seal of the inlet test-bench device and external aerodynamic drag ΔR_windage with connected pipeline at the inlet, applied while the test-bench box calibration, as the difference between the thrust values ΔR_{air_force cor} = R_0eng.corr – R_eng.corr. The article presents the technique for test-bench thrust reduction to normal conditions H = 0 and M = 0 of GTE with large double-flow degree with split circuits at subcritical modes of the jet nozzles. This is being done at the total pressure loss σ_in in the inlet device difference from 1.0, as well as total pressure at the inlet P_in^*, damped temperature T_in^* and the moisture content d difference from the standard values.

The aerodynamic force value (ΔR_AF) determining error estimation according to the technique being suggested was performed in the article.

The article estimates the error in determining the value of the aerodynamic force according to the proposed method.

The article demonstrates the possibility of employing, if necessary, a certified high-altitude test-bench for the aerodynamically non-certified box of the test-bench to determine the aerodynamic force reduced value (ΔR_{air.force.cor}) for the selected turbofan type. The demonstration is based on the example of satisfactory comparison of the experimental values of the reduced test-bench thrust of the turbofan of large double-flow degree with separated circuits in the mode n_fan.cor = const in the certified (closed) box of the test-bench. The experiment was conducted in both layout with mechanical coupling by the input lemniscate, and in thermal pressure chamber of the certified high-altitude test-bench with mechanically detached lamniscate under conditions of H = 0 and M = 0.

The technique for the aerodynamic force determining as a correction to the force from the engine thrust, recounted in the article, may be applied for aerodynamic calibration of the non-certified closed box of the text-bench to account for the value of aerodynamic force. This can be done while both development tests of the pilot item and acceptance tests of a stock-produced turbofan of a large double-flow degree with separate circuits.

The article presents the results of a study aimed at enhancing accuracy and computational efficiency of algorithms for working fluid thermodynamic properties and functions determining used for the gas turbine engine workflow computing.

The working fluid of an atmospheric gas turbine engine is a mixture of seven general individual components such as nitrogen, oxygen, water vapor, carbon dioxide, sulfur dioxide, argon and helium. Setting values of relative mass fractions of components allows calculate the working fluid parameters depending on the properties of the above-said components.

Expressions and corresponding coefficients for a mixture thermodynamic properties and functions computing were obtained based on the existing dependencies of the isobaric heat capacity on temperature for the above-listed components. A new thermodynamic function j was introduced, which allowed establishing a relationship between the total and critical temperatures of the working fluid, with account for its composition and variable heat capacity.

The expressions being presented allow replacing conventional isentropic functions based on the assumption of a constant heat capacity. Application of these new expressions for isentropic relationships between total, static and critical state parameters ensures higher adequacy and better reliability of a gas turbine engine thermodynamic model. This became possible since the isentropic functions are accounting for the dependence of properties on working fluid composition and temperature as well.

The developed approach for the working fluid properties numerical modeling allows creating the time-efficient algorithms for thermodynamic and gas-dynamic process simulation. It has a wide range of applications and scaling capability to create more complex working fluid models.

«ExoMars» is an international project intended for studying the Mars surface, obtaining geological samples and detecting traces of possible life existence by delivering a Russian-made descent platform to the surface with a Mars rover onboard.

The structural elements and systems of the «ExoMars» spacecraft should function reliably under the impact of Martian atmosphere factors, which characteristic feature, is constant presence of dust in particular. The presence of the above said operating conditions leads to the necessity of increasing the volume of ground-based experimental tests and functioning check-up of the spacecraft structure unfurling elements

after exposure to dust. Such «ExoMars» spacecraft structural elements include: — The Mars rover ladders;

— Low-directional antenna boom (LDA); — Solar panels (SP).

Dust settling on the structure of mechanisms may lead to clogging the gaps in rotation nodes, abrasive impact on rubbing pairs and, as the result, to the decrease in functional characteristics of mechanisms.

Since the dusty conditions lead to the increase in the energy capacity losses of the springs in the rotation nodes, and the presence of dust on the mechanism structure leads to the increase in its moments of inertia, the angular velocity of the mechanism under dusty conditions should be less, and the unfurling time should increase.

Tests of sand dust impact on the unfurling elements of the «ExoMars» spacecraft structure were performed in a sand-and-dust chamber, representing a device equipped with a closed wind channel and including an internal working volume and a unit for the dust feeding.

To achieve the required dust concentration, a calculated amount of dust was introduced into the chamber, and air was supplied.

The components and elements of the unfurling structures of the «ExoMars» spacecraft intended for laboratory and development tests were subjected to dust exposure tests. They were two ladders for the Mars rover exit, two SAT panels, and an MNA boom. The task of the tests consisted in operability checking of these structures after exposure to dust, as well as to assessing the unfurling time changes prior and after the dust exposure.

The dust exposure tests were conducted in the following order:

— Accelerometer sensors connected to the measuring station were fixed on the structural elements of the unfurling mechanisms, and mechanisms were transferred into the furled position and locked by pyro nodes simulators. Testing ladders opening, the MNA boom and the SB panels was performed manually prior to the dust exposure. The unfurling time was being determined according to the graphs from the sensors;

— The unfurling structures were returned to the folded and locked position. The inner volume of the sand and dust chamber was hermetically sealed. The test objects were being exposed to the dust particles of no more than 50 microns in size for 15 minutes;

— The ladders, the MNA rod, and the SB panels were unfurling after the dust exposure in various spatial positions provided for by the test programs and techniques. The unfurling time for each product was determined according to the obtained graphs from the sensors.

The test results reveal that the dust impact (similar to the Martian dust impact) does not significantly affect the performance of the unfurling structures. The unfurling occurs in the normal mode, the opening time increases herewith by no more than 3% compared to similar tests prior to the dust exposure. Consequently, the energy consumption of the springs of the mechanisms is sufficient for full-scale operation of the spacecraft in Mars conditions.

The topic of the article being presented is trajectory estimating algorithms and subsequent initial state vector determining of a small-sized aerial vehicle based on measurements obtained on the BT CM3 type ballistic tracks. At the beginning, the article considers general issues of the small-sized aircraft studying by full-scale tests on ballistic tracks, presents the features of their instrument equipment, and touches upon the issues of the trajectory restoring based on the measurement results.

The technique proposed by the author is based on the least squares method application for a trajectory forming according to the measurements of the aircraft flight coordinates through the certain sections of the test facility. The efficiency of these algorithms is illustrated by the solution of a numerical example simulating experimental data. It was proved by additional computations and comparative analysis that the most effective way to restore the trajectory is the least squares method using the second-order approximating polynomial. Theoretical justification of this phenomenon is presented.

Besides the algorithm for the initial state vector detecting, inclusive coordinates of the flight initiation in the selected coordinates system, the initial trajectory inclination angle, initial track angle and initial velocity value, the article suggests the trivial technique for the single anomalous measurements rejection. It presents also theoretical justification of the full-scale experiment results, and defines the requirements for conducting research on the ballistic track with target frames application. A typical algorithm for the initial angular velocity determining and estimating the derivation value is described as well. An empirical algorithm for finding the drag coefficient value based on the results of experimental shooting is presented. Among other things, the article presents the main characteristics of the ballistic track of the «Dynamics and Flight Control of Rockets and Spacecraft» Department at the Bauman Moscow State Technical University.

The final part of the article formulates a number of practical remarks and recommendations to the experimental studies organization on ballistic tracks for the initial state vector reliable determining and flight trajectory restoring.

The article performs the analysis and definition of the in-flight refueling as a problem of the high-precision piloting, and considers refueling system of a «hose—cone—link rod». Statistical characteristics evaluation of the piloting process was performed based on experimental data obtained while full-scale and semi-natural experiments on the flight simulator employing various dynamic configurations. A widely known Neil-Smith database as well as the data obtained by the identification results in flight experiments with the Russian planes underlie the basis of the dynamic configurations structure.

The experiments were being performed with the TM-21 flight simulator at the Moscow Aviation Institute. The semi-natural model for the refueling imitation was structured so that the electro-hydraulic loading of the central control stick corresponds to the range of the steering levers loading of modern maneuverable aircraft as well as speed control characteristics. A totality of 263 experiments was performed with participation of six professional test pilots. The gross amount of runs was 897. Conditions of the experiments corresponded to the average values of flight speeds and altitudes.

The simulation system verification revealed rather high correlation coefficient value (k = 0.834) between the «simulation» and «real» ratings, which confirms the obtained results authenticity. Besides the pilots participating in the experiment, three more test pilots, highly experienced in the refueling flights, were being engaged additionally as experts to estimate the flight simulation adequacy. The pilots-experts stated the high level of the simulation congruency.

The following indicators were adopted as the basic quality indicators of the refueling performing and aircraft controllability characteristics:

• by a particular experiment — the target accuracy characterized by the radius of deviation fr om the cone center at the instant its shear plane crossing, and subjective pilot estimation;
• by a number of experiments — the relative frequency of hitting as the hitting probability estimation. The results of the experiments revealed that according to the expert-pilots esteems the piloting characteristics qualities are being correlated rather closely with the relative number of hits. The boundary of the first level of flying qualities (PR = 3.5) corresponds to the relative number of hits of about 60%, and the lower lim it of the second level of flying qualities (PR = 6.5) corresponds to the relative number of hitting of about 30%.

The obtained results are recommended to be employed for the requirements forming to the aircraft piloting characteristics at the in-flight refueling modes.

Cyclonic systems for the leading edge cooling are an effective way of heat transfer intensification, which ensures low pressure losses in the cooling channels and the lowest possible coolant consumption. One of the basic tasks the designer faces when developing a cooling system for a gas turbine blade with the leading edge cyclonic cooling consists in determining rational diameters of the intake and outtake orifices and the step of their placement, which allow ensuring maximum heat removal from the surface with a minimum temperature field asymmetry. An important feature of cyclone cooling is the high sensitivity of the heat transfer intensity and the nature of the heat transfer coefficients distribution over the surface of the cyclone chamber to the geometric parameters of the cooling system. These parameters are the orifices diameters ratio, their step, the cyclonic chamber size and shape, and the orifices shape. In this regard, numerical studies conduction is required for each particular blade structure to determine geometry parameters of the cyclonic chamber to obtain the required cooling efficiency. The presented work deals with numerical study of the heat transfer in the closed cyclonic channel, which is assumed to be applied for convective cooling of the turbine blade leading edge.

The thermal and hydraulic characteristics studies of a closed cyclone have been conducted to ensure the nozzle blade development for the high-temperature turbine with convective cooling of the leading edge. The intake orifices diameter was being varied from 1 mm to 2 mm, the outtake orifices diameter was being varied from 2 mm to 3 mm, and the cyclonic chamber was of 6.2 mm diameter. The article shows that area increasing of the intake and outtake orifices in the cyclonic chamber changes the heat transfer coefficients distribution profile. The local heat transfer coefficients were computed, and criterion equations for the dependence of the Nusselt number in the cyclone chambers on their geometric and operating parameters were elaborated.

It was found practical to reduce the outtake orifices diameter with conjoined step reduction for the heat transfer coefficients values increasing, which would ensure the non-uniformity reduction in the heat transfer coefficients distribution over the cyclonic channel height.

With the fixed pressure drop in the outtake and intake channels, the throughput of the cyclone channel is determined mainly by the area of the intake orifices, which allows the leading edge cooling efficiency enhancing, by increasing the outtake orifices area.

The article recounts the significance of real-time traffic control systems for global competitiveness and technological security ensuring amidst the fourth industrial revolution realization. As far as the growth potential of computers elements base running speed is close to exhaustion, and further development in this trend is being associated with significant technical complexions and economic efficiency reduction, computers architecture improvement should be considered as the main trend of the computer productivity increasing. The article considered pressing tasks of the computations productivity increasing, which may be solved at the cost of computers architecture improvement. These tasks include the processed data flow volume reduction; increasing data transmission speed between computer elements; eliminating queues while several computing devices simultaneously accessing the same memory. The authors propose conceptual model of the industrial robot movement control based on the analysis of the possible ways of the set problems solving. The problem of the processed data flow reduction is being solved in the system built according to the conceptual model being proposed by application of extra computing modules, such as coprocessors and accelerators, performing parallel computing. The main portion of computations herewith is being performed without control from the system core. The problem of data transmission speed increasing between the system functional elements and blocks is being solved by the memory-centric architecture employing, with which all devices requiring high speed of data exchange with memory for their operation, are being integrated into the memory. The queues elimination problem is being solved by dynamic random access memory (DRAM) splitting into local areas accessible only by a single device. Interaction between devices is being implemented in the high-speed static random access memory (SRAM) employing minimum data volumes, as well as through the communication network ensuring direct communication between the devices without delays occurrence. The actor instrumental model, ensuring emulation of parallel computing and functional modules interaction, is being selected to describe the industrial robot movement control system operation built according to the presented conceptual model.

This article considers the problem of thermal spray coatings adhesion strength assessing to the part surface. Performing numerical modeling of heating and acceleration processes of the sprayed material particles, as well as their collision with the base surface of the set micro-relief employing the ANSYS CFD Premium software is being suggested as the problem solution. The plasma spraying process is being considered as an example.

At the beginning, the article performs the analysis of the literature related to the problem of adhesion strength determining of gas-thermal coatings, obtained by the plasma spraying, with the base surface. The rationale for the need to model the sprayed material particles transfer and collision processes with the base surface is rendered.

The work separates out the stages and general approaches to the plasma spraying process modeling. The main process parameters are being defined, and description of the plasma jet outflow from the nozzle with the flow of particles being sprayed onto the base, is being presented. The curves of the spraying temperature and particles velocity dependency on time were plotted. Comparison of the obtained values with the experimental data is being performed.

Simulation of a single sprayed particle collision with the base at various combinations of temperature and the particle velocity at the moment of the particle approach to the base surface is performed in the work. The micro-relief geometry and size are being determined herewith. As the result, various particle shapes after collision and the value of the specific contacting area for each case under consideration were obtained. Finally, a qualitative assessment of the interaction between a particle of the sprayed material and the sprayed surface is presented. The most optimal combination of the temperature and particle velocity is identified.

Quality assessment of the military-oriented aircraft engineering (MOAE) samples is being performed by one of the following techniques: complex, differentiated, mixed, integrated as well as by the economic practicality. Each of these methods has its pros and contras.

The complex technique allows assessing the quality level in aggregate, but it does not allow accounting for all meaningful indicators.

The differentiated technique computes simple quality indicators with account for the meaningful ones affecting the quality of the MOAE samples. This method application causes difficulties in the quality level assessment by the large quantity of simple indicators.

The mixed method allows quality assessment of the MOAE sample at large aggregate of the simple, meaningful and generalized indicators. Accounting for the large quantity of indicators requires complex mathematical calculations.

The integral method is applicable for assessing the MOAE operation efficiency. This method application is practical only when total costs of the sample creation, operation and useful effect of the sample operation are determined.

The sample quality assessment technique by the economic effectiveness is applied only when economic assessment is necessary. With this technique application, a large quantity of data on the sample should be necessarily accounted for.

All these techniques are applicable for the assessment of a single-type MOAE samples, namely of the same type and purpose. For assessing diversified MOAE samples quality indices are being employed

A brief analysis of the above listed techniques allows inferring that their application for the MOAE sample is not always practical. It is stipulated by the following reasons:

• The difficulty of reducing a wide nomenclature of indices to the resulting value expressed in a numerical form;

• The absence of the possibility for accounting for the external factors; 

• The absence of the full pattern of the MOAE sample quality.

All these reasons instigate the search for new approaches and techniques of quality assessment accounting for the MOAE sample specifics.

According to the article «Application of analytical methods of open complicated systems for assessing the quality of designs of weapons, military and special equipment», MOAE is an open complicated system. Hence, the most suitable quality assessment technique for the open complicated systems is the technique for express-assessment of the open complicated systems functioning.

With account for the suggested technique and the approach, applied at present, the algorithm for the quality level assessment of the production was developed. The algorithm for the MOAE quality level assessment consists of two basic blocks. The first block is universal, and it is applied for quality level assessment of practically all kinds of products. As applied to the MOAE the first block consists of the following stages:

1. Setting the goals and tasks for the MOAE quality level assessment at all life-cycle stages. The main life-cycle stages are development, production and operation.
2. Defining the quality indicators nomenclature of the MOAE sample under study is a very important stage for its quality assessment. It is necessary to regard for the composition, structure, operation conditions, design specifications specifications and a number of other parameters while defining the quality indicators nomenclature of the MOAE sample.
3. There are six main techniques for defining the values of product quality indicators. They are measuring, registration, calculation, organoleptic, expert and sociological. All these techniques may be employed as applied to the MOAE samples.
4. Quality indicators values determining of the MOAE samples depends on the selected technique, and the tools used by this method.

The second block of the MOAE samples quality level assessment consists of the following stages:

1. The MOAE sample quality formalization represents its expansion into fundamental composite indicators in the form of hierarchical structure. The algorithm distinguishes internal and external formalization. External formalization means the studied object extraction from the external environment. In this particular case, the object of study is the MOAE sample quality indicator. Internal formalization means the MOAE sample quality indicator representation in the form of the hierarchical structure of the indicators, affecting its quality. Let call these indicators factors, since each lower-level indicator in the hierarchical structure affects the upper-level one.
2. Assessment of all factors of the hierarchical structure, as well as those of different physical nature is being performed according to the unified criterion scale, which envisions the factor state assessment on the assumption of the direct assessment principle on the interval from 0 to 1.
3. A neural-like network is being set based on the hierarchical formalization. The neural-like elements of this network and connections formed between them simulate individual factors. Each layer of the neural-like elements simulates factors of one hierarchy level. A neural-like network can work in two basic ways:
• Deterministic, when all neural-like elements operate according to a deterministic option;
• Statistical, when at least one neural-like element operates using simulation by to one of its characteristics. 

4. The initial data for the MOAE sample can be determined on account of the purpose and structure, qualitative and quantitative characteristics of the operation processes, characteristics of external impacts of various physical nature factors, tactical situations options, characteristics and composition of means interacting with the sample, and characteristics of active counteraction means.
5. According to the pre-determined operating option of a neural-like element in the neural-like network, the compliance level of the MOAE sample with the intended objectives is being calculated.
6. If necessary, factor analysis is performed to check correctness and reliability of the resulting operating model of the neural-like network.
7. Decision making on the compliance level of the MOAE sample with the intended objectives (the requirements of tactical and technical tasks or technical conditions) serve as a basis for:
• Preparation and formation of suggestions and conclusions on the possibility of adopting the developed (tested) MOAE samples with putting them into production;
• Assessing the degree of the MOAE sample employing in real combat conditions; 
• the possibility of the MOAE sample employing in various weather conditions./li>

8. Conclusions on the MOAE sample quality level (in conjunction with its purpose) compare the obtained quality indicator either with the basic one or with quality indicators of the foreign samples computed earlier. If the quality indicator appears less to be than the basic one or the foreign sample, suggestions are being elaborated on the indicators (factors) improvement of the first, second, third etc. hierarchical levels.

The suggested approach to assessing the quality level of MOAE sample possesses the following advantages:

• Apprehensible and accessible formalization (structuring) of the object under study;
• A comprehensive assessment of the MOAE samples quality is being performed with account for the external factors of various physical nature;
• The quality level assessment authenticity is being determined by the possibility of employing all available information (deterministic, calculated, expert);

The ability of quick initial data setting and producing the results in real time.

Оn May 13, 2021, the Department of the “Rocket Engines” (depanment 202) of the Moscow Aviation Institute (MAI) in collaboration with colleagues from other universities and industry bodies held the All- Russian Scientific and Technical Workshop “Bladed pumps and turbopump units”. The Workshop was dedicated to the 100th anniversary of Boris Viktorovich Ovsyannikov, ап outstanding scientist, tutor, founder of the scientific school of high-speed turbopump units of liquid-propellant rocket engines. Doctor of technical sciences, Professor of MAI B.V. Ovsyannikov, has been working as the head of the Department 202 for а long time; he educated а whole galaxy of scholars. Не is the author of the famous textbook оп liquid-propellant rocket engines turbopumps, which gained the world recognition.

The Workshop was attended by the colleagues from NPO Energomash, SSC “Center Keldysh”, UDD “Kristall”, St. Petersburg Peter the Great Polytechnic University, Siberian State University named after M.F. Reshetnev and others. The content of the Workshop were memories of B.V. Ovsyannikov’s colleagues and relatives about him, modern scientific and technical information оп topical problems of bladed pumps, as well as liquid propellant rocket engine turbopumps units. А selection of artricles in the Aerospace MAI Journal was prepared based оп а number of reports.

The scientific heritage of В.V. Ovsyannikov, his artricles, textbooks, author’s certificates total more than а hundred titles. They are being used heretofore by students, postgraduate students, and engineers.

To improve centrifugal pumps cavitation qualities of turbopump units (TPU) of liquid-propellant rocket engines, a centrifugal impeller with increased throat area at the inlet is being developed, a booster pump with rotational speed lower than that of the main pump is being employed, an upstream axial wheel, i.e. a screw inducer, is being applied. This allows reducing the required cavitation margin. However, along with high cavitation qualities, the upstream inducer displays significant disadvantages. When the screw is operating at the inlet at feeding modes less than 0.5 of the optimal value, backflows are being formed, increasing with the feeding decrease. These backflows lead to the increased vibration, unstable operation, and low-frequency pressure pulsations of the self-oscillations nature. Cavitational self-oscillations attain a large amplitude and may lead to the pump and even the entire feeding system failure. One of the promising ways of the pump cavitation qualities improving, and reducing noise, vibration and low-frequency pressure and flow pulsations consists in the axial-vortex stage installing at the pump inlet. The axial-vortex stage (AVS) represents a pump consisting of an axial screw wheel and a fixed helical cascade on its periphery. The AVS advantages are being manifested most substantially at the flow rates less than the optimal one compared to the screw inducer. The axial-vortex stage (AVS) wields a higher pressure coefficient, better cavitation qualities, and ensures stable operation in the entire flow range and on the stalling branch of the cavitation characteristic. Further studies on the possibility of pressure pulsations, vibration and cavitation damage reduction while the AVS application are required.

As fire tests (FT) practice revealed, defects leading to destruction of engine structure elements, such as radial-thrust bearings, parking seal and blade wheel hub of the centrifugal pump occurred and developed with time in the automatic relief valve (ARV) circuit and parking seal (PS). Very often, such defects were developing in the course of several and even tens of seconds. These defects may be detected at the early stages of their development by the functional diagnostics methods employing slowly changing parameters being measured while the FT and mathematical model of the engine workflow processes.

Until recently, the computational-experimental analysis of accidents occurring in the ARV circuit and PS was performed locally, using only a mathematical model of this circuit, where the boundary conditions were assigned by empirical or approximation dependences. It is clear that integration of the ARV circuit and PS mathematical model into the math model of the engine workflow processes gives an opportunity of obtaining more complete diagnostic information about the circuit being considered. It is worth noting the inexpediency of neural network involving for this purpose due to the necessity of its training on a large number of FTs.

To increase the depth of engine diagnosing and confident control of the ARV and SS circuit state, the system of ARV and SS equations is closed by the parameters, by which this circuit is being conjugated with the engine parameters. By the model obtained in this way, a step-by-step process of the ARV and SS circuit state diagnosing is presented, starting from the moment of identifying the time of a fault occurrence and up to its localization. At each stage, special algorithms are being used to confirm the decisions made at the previous stage. The control begins with determining the moment of malfunction occurrence by measured parameters of the malfunction occurrence time instant. After this, deviations of measured parameters from the ones computed with the model are being controlled. Then it is necessary to proceed to the control of the engine characteristics deviation from those obtained while autonomous tests of units. Finally, if necessary, the control of functional relations violation by the structural exclusion method is being performed. On the example of liquid rocket engine state control during test bench fire test, the sequence of diagnostic procedures resulted in the malfunction, which caused forces unbalance on the radial-thrust bearing of the oxidizer pump and pressure increase in the cavity of the oxidizer pump control system, was detected and localized, was presented.

The stated diagnostic procedures may be employed in the analysis of a wide class of complex technical systems functioning.

Pump seals of liquid rocket engines turbo-pump units are the key element defining the pump volumetric efficiency. The seal type selection herewith affects not only characteristics, but the pump operability as well. Both contactless and wearing-in seals are being employed in the liquid rocket engines turbo-pumps. The article considered the contactless seals, such as seals with floating and semi-movable rings, groove seal with fixed smooth wall and labyrinth seals, as the seals most frequently employed in the pumps structure.

Very often, the gap in the seal is being considered as a constant value while the pump operation analysis on the engine regulation modes. This was substantiated for the pumps of the engines operating without the generator gas afterburning behind the turbine, when delivery pressure and peripheral velocities were relatively small and, consequently, the level of seal elements deformation, both rotor and stator, was not high. It allowed not accounting for their impact on the gap value and leakages (consumption) through the seal. Transition to the engines with generator gas afterburning was accompanied by the pressure and peripheral velocities growth. It led to the necessity of accounting for the deformation of seal structure elements impact on its characteristics. The necessity for the engine operation regulation, including both forcing and throttling modes by thrust from 25 to 120% of the rated value required knowing the pumps parameters on all operation modes.

Another task during design is selection of the clearance size, ensuring the contactless operation of seal in all engine’s operating modes, from chill-down to its shutdown.

Thus, while the seals design of the pumps’ air-gas channel, the two types of gaps should be determined on all operation modes: the working gap determining consumption characteristics of the seal, i.e. the pump volumetric efficiency, and minimal guaranteed gap between rotor and stator seal elements, defining contactless operation conditions of the seal.

The article provides the dependencies for estimating the seal gap at the initial design stage.

The performed analysis demonstrates that already at the early design stages it is necessary to account for the seal gap impact on the pump efficiency with dependence on the operation mode.

The seal type selection exerts a substantial impact on the value of the seal guaranteed minimum gap. Thus, the analysis of its changing and permissible value should be performed beginning from the early design stages. The errors in the seal gap size selection may lead to modifying and necessity to the crucial changes of the structure.

The object of the study is an oxygen-kerosene liquid rocket engine (LRE), realized according to the scheme of the generator gas afterburning in the combustion chamber.

The method proposed in this article is a method for current state monitoring of modern high-power LRE in real-time scale of the test-bench fire tests. It allows estimating its actual state in both stationary and transient modes.

The method does not require pre-estimation of the fail-safe operation criteria boundaries of the LRE being monitored, and adapted to the operation modes and external conditions changing.

The current state of the engine is being monitored at the rate of measurement results receiving of the slowly changing engine parameters, determined with certain rather small time step.

Each specific situation is being considered as a continuation of the previous engine operation in the mode under consideration, for which purpose, both conformity and inconsistency of the current engine state to the «prehistory» of this state, which was recognized corresponding to the successful operation of the engine, are statistically confirmed.

Formally, this “prehistory”, as well as information about the current state of the engine, is a set of measurements of its parameters obtained from the initial control point to the one under consideration.

To make a decision on a malfunction occurrence, a statistical analysis method is used, developed to identify and exclude the results with abnormal inaccuracies. In case of current statistical characteristics threshold values are exceeded by their current values, the fact of malfunction occurrence is being registered, and the test is being terminated to development of the revealed malfunction.

For stationary LRE operation modes, the instant of a malfunction occurrence can be defined as the moment of a distinct change in the stability of measured parameters. In this case, for making a decision on the malfunction occurrence and test termination, the time series of measured parameters are subjected to statistical evaluation based on the Student’s criterion.

In transient modes, the time series values of changes gradients in the measured parameters, possessing the property of stationarity, are subjected to a similar analysis. This property is stipulated by the fact that during bench tests conducted according to a given cyclogram, the engine control in transient modes is being ensured by changing the drive angle of the control unit by the linear law.

The developed method for assessing the current state of the LRE during bench tests allows preventing the LRE malfunction development, and generate an appropriate signal to the engine control system in real time of the test-bench fire test.

Powerful energy-propulsion units differ from the others by the elements, subassemblies and structures associated with powerful turbo-machines and gas generators as a part of them. The high rpm of the turbines shafts rigidly affects the structure, which may lead to its destruction. High-frequency vibrations, which occurrence is possible in the turbopump units of liquid propellant rocket engines, are of especial danger.

The purpose of the study consists in the following:

– the problem setting of high-frequency instability prediction in powerful energy propulsion units on the example of the turbopump unit of a liquid propellant rocket engine, determining instability parameters in this subassembly and required ratio of the turbulent gas field parameters;

– formalization of vibrations automatic monitoring condition by the digital methods of multi-step discreet Fourier transform without performing hardware-consuming multiplication operators.

The presence of constant free volume is necessary for setting constant stable turbulence mode for the high frequency stability ensuring. This fact actualizes the study of the additional possibility of setting constant stable turbulence mode with the gas or liquid flow velocity increase. Namely turbulence is in charge of high-frequency instability, and, hence, vibrations occurrence. Turbulent flow originates practically always in turbopump units.

The occurring high-frequency instability of the process, accompanied by the oscillation of the working fluid particles inside the turbopump unit, impacts the walls of the apparatus that restrains the working volume. The walls of this apparatus begin reacting to the force impacts of the gas and naturally impede it, generating vibrations of the structure. The effect on the system occurs as the impact of a compelled force in the form of a harmonic component coming from the gas. The equation of the oscillating link for the structure will look like a second-order differential equation with respect to the walls displacements.

The study employed the principles of vibrations diagnostics of liquid propellant rocket engines on the example of a turbopump unit by digital methods of a multi-stage discrete Fourier transform.

An increase in the vibration level of liquid propellant rocket engines may lead to the increased thermal loads with subsequent possible burnouts of the walls of the turbopump assembly units. This requires quality improving of the vibrations diagnostics of liquid propellant rocket engines and increasing the information content of methods employed for the level control of these vibrations.

Vibration diagnostics may and should be ensured with the software and hardware for digital signal processing from signaling sensors using digital filtering and discrete Fourier transform of such signals. The term «unerroric» (from the Latin «errare») in relation to such digital signals deductive processing defines an active process of the errors level reducing in digital signal processing when setting various values of integer difference coefficients of digital difference filters applied for multi-stage discrete Fourier transform. Such unerroric reduces the error of automatic vibration control.

Gradual tightening of the requirements for the liquid rocket propellant engines reliability contributes to the problem actualization of such engines vibrations diagnosing under conditions of their mass production.

The turbopump unit (TPU) solves the problem with the flow rate and, thus, the problem of overcoming the power threshold necessary for long-distance flights into space. All modern space rockets employing a turbopump as an alternative device for supplying high fuel consumption to the combustion chamber, ensuring the necessary power and thrust of a liquid-propellant rocket engine (LPRE).

The V-2 rocket was created in Germany during the Second World War. It was being deeveloped on an initiative basis by a group of specialists within the framework of the German Ministry of Defense. It took a lot of time and trouble to convince the leaders of Nazi Germany of the need to create powerful space rockets that could cross continents and go into outer space. As the result, on July 7, 1943, the decision was made to assign the Peenemunde project the status of the highest priority in the German armament program. After that, the original name of the rocket “A-4” project was changed to “V-2”, and under this name, it became a history.

The basic invention of the V-2 (A-4) rocket was the centrifugal pumps application. Werner von Braun solved the problem of pumps by using fire pumps in the LPRE. Thus, he anticipated the beginning of a new era of LPRE – the era of turbopump.

It seemed almost impossible to design such a pump. After all, it had to perform a number of complex functions, such as supplying liquefied gas, which was one of the fuel components, at a pressure of about 21 atm, and pump herewith more than 190 liters of fuel per second. In addition, it should be quite simple in terms of design and quite light. Besides, the pump had to be started and switched to full power within a very short period of time (~6 s). Explaining to the pumping factory staff his requirements for rocket pumps for the V-2, von Braun involuntarily expected objections from people, but they did not follow. The entire staff of the pumps producing factory was ready for such requirements. Instead of objections, everyone listened, silently and approvingly. Specialists immediately offered a specific solution – the necessary pump was in many ways similar to one of the fire centrifugal pump types. A gas turbine and a steam generator were proposed to be employed as a drive.

The V-2 turbopump represented a single structure in which a two-stage turbine powered by steam gas and two centrifugal pumps for fuel components supplying were mounted on one shaft.

German scientists have created a truly unique unit, and together with it a unique rocket. In fact, a new branch of the industry was created, namlely, rocket engineering under the general leadership of V. R. Dornberger. Subsequently, many V-2 solutions were used by Soviet and foreign rocket engine developers in their latest products, in particular, when creating the R-1 medium-range ballistic missile under the leadership of S.P. Korolev and V.P. Glushko. The historical significance of the A-4 and R-1 missiles cannot be underestimated. This was the first breakthrough into a completely new field of technology. It is impossible to derogate the merit of domestic scientists, their dedicated work, but German scientists V.R. Dornberger, V. Thiel, V. von Braun and others were the first at that time.

Nevertheless, the main finding of German scientists, the turbopump, along with a revolutionary leap, brought a lot of worries into the life of rocket scientists. The impartial analysis of the failures associated with this unit revealed that in most cases the main cause of engine failures was due to the turbopump. It is well-known, that one of the most insidious causes of rotary machines accidents is the so-called fatigue, i.e. the gradually accumulating effect of cyclic dynamic loads, leading to the breakage of shafts, turbine blades, machine rods and other parts.

Thus, it seems rather relevant to apply new methods of analysis, including a combination of various methods of rotary machines diagnostics (primarily, methods of vibration diagnostics) to determine the source and nature of increased dynamic loads to eliminate them or reduce their impact on the structure.

As practice has revealed, hard-to-detect furtive defects, which were not detected by the other methods and control means, specified by the regulatory documentation, were detected, identified and eliminated by the TPU vibration diagnostics. Malfunctions of the turbopump subassemblies caused increased vibration-pulsation loads, leading in some cases to the LPRE failures and emergencies.

The effects and phenomena that were not previously encountered with in the practice of domestic and foreign LPRE-building were identified and studied in detail.

About 70% of the stratum fluid is being extracted by submersible installations of electric centrifugal pumps (ESP). To increase the oil recovery coefficient (ORC), the depression on the stratum increases, the bottom-hole pressure decreases, and technological operations for stratum hydraulic fracturing are being employed.

In this regard, the content of free gas and mechanical impurities increases at the ECP installation inlet. It is necessary to improve the free gas separation efficiency and of gas separators reliability. New, more accurate techniques of bench tests are necessary for the new design solutions testing and developing.

Conventional techniques for gas separators testing on the gas separation efficiency may be conditionally attributed to the two basic techniques. According to the first technique, the gas-liquid mixture (GLM) is being fed into a pipe that simulates the annular space, while according to the second one it is being fed directly to the gas separator inlet.

The first pneumo-hydraulic scheme simulates integrally the gas separator (GS) operation in a well. Some part of the gas misses the gas separator inlet. The efficiency of this pre-separation depends on the design of the base, protective grid and the size of the gas bubbles' average diameter. The larger the diameter, the more likely the bubbles will not get into the gas separator. In this regard, the devices for the gas phase enlargement are relevant.

If the separator is installed inside the pipe, it is difficult to measure the flow parameters inside the flow part, although, namely, this information on what percentage of gas entered the GS, and what percentage missed it due to the pre-separation is necessary to improve the flow part. Difficulties in obtaining the information necessary to improve the flow part inside the GS may be assigned to the disadvantages of the first technique.

The advantage of the second technique consists in the fact that the gas-liquid mixture is being fed directly to the tested gas separator inlet. The quantity herewith of the free gas entering the GS is precisely known. Information on the efficiency of the free gas separation inside the GS, and the capability of measuring the flow parameters inside the GS, allow evaluating the operation of the flow part elements. The disadvantage of this technique consists in the problem of accurate differential pressure maintaining between the areas of the GLM at the gas separator inlet and the separated gas at the outlet, which should correspond to the difference in annular space.

Based on the analysis, the third promising technique and the pneumo-hydraulic scheme of the new test bench were developed and presented. By the authors opinion, the technique combines pros and aligns cons of the conventional techniques. It allows fully simulate tests in the well, and perform measurements in the flow part of the separator.

When optimizing and searching for new design solutions for the flow part elements to increase the separating properties efficiency, the new technique allows installing pressure gauges and special taps for sampling on the gas separator housing, determining the pressure gradients along the length of the separation chamber and the degree of mixture dispersion. The separation efficiency is higher for structures with the higher pressure gradient and larger average diameter of gas bubbles.

Based on the engineering mathematical models the article considers the issues of filling and defining the dome (wing) filling criteria of a double-shell gliding parachute, which is directly interrelated with such important parameters and characteristics as aerodynamic load on the parachute, parachute strength, the filling path, altitude loss while filling and the wing geometry stability.

The double-shell wings fillability of the gliding parachutes means their capability of taking its aerodynamic fully filled shape (from the state of the wing stowed in a package) under the impact of velocity head of the incoming flow in a definite time called the filling time.

The article regards certain basic moments and structural specifics, significantly affecting the filling process of the double-shell gliding parachute.

Great attention is paid in the work to the air intake operation efficiency, depending upon the whole number of factors, such as:

– Divergence angle of the system velocity vector line of action with the normal to the air intake plane, depending on its location on the wing. It defines the wing filling efficiency and maintaining sufficient excessive pressure in it to keep the wing filling geometry;

– Air intake area;

– The Strouhal number, which determines the pulsation nature of the mass of air emissions from the wing through the air intake into the external flow, which causes the pulsation nature of the entire pattern of the external flow, significantly increasing the resistance of the wing and reducing the speed of the system.

The article presents engineering calculations for estimating the filling time of the sections and the wing as a whole, with account the for structural air permeability in the wing ribs. The differential equation of the masses balance of the air entering the section and flowing out of it was formed. Integration was performed, and the dependences for determining the gliding parachute wing section filling time were obtained. The time dependence for the volume of the section being filled was obtained as well. Graphs for the obtained dependencies are presented and their analysis is performed.

The article considers in detail the gliding parachute filling criteria, such as filling time and the Strouhal number, characterizing the wing filling efficiency. These criteria may be employed while comparing filling processes and optimal option of the gliding parachute structure selection.

The article deals with the task medium-term forecasting of rational characteristics (imaging hardware spatial resolution, weight and cost) of a prospective spacecraft for remote Earth probing with optoelectronic imaging hardware. It proposes a method for the task solving employing extrapolation methods based on the statistical data on the products prototypes. Forecasting is being performed by extrapolating into the future the regularities revealed in the process of studying characteristics up to the present moment.

For the proposed method realization, the searching algorithm, including such blocks as initial data, extrapolating prediction and a spacecraft characteristics evaluation, was developed, and the results of its technical-and-economic characteristics at the medium-term forecasting are presented. The source data block includes information on the characteristics of the Earth remote probing spacecraft with optoelectronic imaging hardware of various types. Statistical data processing on the characteristic (parameter) under study is being performed in the extrapolating prediction blockIt is assumed herewith that parameter realization is a random function of time (a forecast function).

Characteristics predicting of the Earth remote probing spacecraft is being performed for the following types of optoelectronic imaging hardware: panchromatic range; multispectral visible and near-infrared ranges; combined (panchromatic and multispectral) visible and near-infrared ranges. The article presents the computational results of Earth remote probing spacecraft characteristics being predicted, such as spatial resolution of imaging hardware of various types, weight and cost of the spacecraft creation up to 2030.

Computational results show that the following improvements are forecasted for the spacecraft with panchromatic and combined imaging hardware:

– The spatial resolution improvement up to 0.19–0.22 m with maximum diameter of the Korsch type optical system up to 1.3–1.4 m;

– Weight improvement up to 3000–4000 kg;

– Insufficient cost of creation increase up to 235 million of conventional units.

For the spacecraft with multispectral imaging hardware:

– The spatial resolution improvement up to 3.0–4.0 m;

– Optical system diameter up to 0.25–0.32 m;

– Weight improvement up to 500 kg, and cost of creation increase up to 60 million of conventional units.

Thus, the method proposed in the article and developed design models allow predicting technical-and-economic characteristics of prospective modifications of the Earth remote probing spacecraft for 7–10 years, and ensuring necessary research accuracy.

Thermal mathematical models with distributed and concentrated parameters of the AIST series small spacecraft were developed. Verification of these models was performed based on telemetry data obtained while he spacecraft experimental operation. Verification possibility of theoretical calculations of the supposed small spacecraft temperatures and obtained telemetry parameters allows improving the technique for finding parameters of the thermal control system with improved qualitative indicators. The authors developed the technique for the small spacecraft thermal control system design. Computation of mathematical model of a small spacecraft with distributed parameters was performed with the Simcenter 3D Space Systems Thermal module of the Siemens NX specialized software. Computation of the spacecraft thermal state mathematical model based on differential equations with lumped parameters was performed with MATLAB software package in Simulink environment for the complex technical systems dynamic interdisciplinary modeling.

The developed technique of the thermal mathematical model was applied for developing a computational mathematical model of the thermal state of a prospective small spacecraft for environmental monitoring tasks. Thus, the main objectives of the study are as follows:

– obtaining and analyzing a real picture of the thermal regime of the «AIST» series small spacecraft based on the telemetry data;

– developing thermal mathematical model of a small spacecraft in distributed parameters;

– developing thermal mathematical model of a small spacecraft in lumped parameters;

– verifying computational models by the telemetry data;

– developing design technique for the small spacecraft thermal control system, with appropriate mathematical models application;

– solving partial design problems employing the developed technique.

This publication briefly discusses the possibility of high-quality improvement of the power plant performance, built on the turbofan basis, by injecting water into the inlet device. The probability of this power plant introducing into the space transport system, instead of the first stage at the flight speeds up to six Mach, was considered as well. The expert analysis of the existing research solutions was performed. This technology realization solves the problems of cargo transportation to the International Space Station (ISS). There is a possibility of creating a passenger spacecraft with an immense flight speed in the future.

It is necessary to find a solution, with which the speed characteristics of a turbojet bypass engine with an afterburner are an order of magnitude higher with water injection than without it, and find out the required amount of water necessary for air-cooling to 120°C and 300°C at the engine inlet.

The basic requirements placed for the engine are the low weight and cost at a comparatively high power. Accordingly, the power plant should be operational at all speeds up to six Mach, as well as its operation must meet all the necessary conditions at altitudes within 25-40 km to implement a full flight cycle. The engine herewith should be of the lowest possible specific fuel consumption. Maintenance should not be impeded, since it is necessary to expand the number of airports at which this aircraft can be based, expanding thereby its flight routes.

Water injection of into the flow part increases the engine speed characteristics and its application at the speeds up to six Mach. However, this technology has its minuses as well. Takeoff weight increase and complication of the design negatively affect the flight range and the ease of operation. Due to the cooler injection application, the the power plant device becomes more complicated, which leads to the complication of all technological operations, from manufacturing to setting up the unit.

Nevertheless, the idea is rather promising in practical application, but it requires an utmost high-quality detailed refinement of both the power unit itself and the aircraft.

Today, industry, especially knowledge-intensive branches, is experiencing an active growth of well-deserved attention to digital technologies. In support for the Aircraft Building Development Program of the Russian Federation realization, and the strategy for the civil products in the sales and service segment the United Engine Building Corporation goes along the path of comprehensive innovations implementation while conducting research, research and development work, manufacturing and after-sale services.

Among the priorities of the innovative development of the Corporation the following areas may be highlighted:

– A concerted strategy of scientific and technical development of the industry, which defines the list of critical technologies and the trends of the corporation industrial model transformation;

– The key product programs of engine building in the trends of aviation, ground and seaborne aggregates;

– Transformational projects, which task consists in achieving the strategic goals of the Corporation, including the terms reduction for launching new products to the market.

Digital technologies allow not only the current processes automation, but also formation of the new ones with new qualities and contributing to the products of the United Engine Corporation being competitive and in demand on the world market.

For this goal achieving, accumulation of the best technologies, best resources, operating in the high-tech field such as engineering centers, startups, research teams at the Universities, and the institutes of the Russian Academy of Sciences is of utter importance. This is an ambitious task, practically proclaiming that it is important to become twice as effective to meet the customers’ needs. A digital twin is a prospective trend for this problem solution.

The concept of a digital twin was proposed by Michael Grieves, a professor at the University of Michigan, back in 2002. As he notes in his work, it was primarily called the «Mirrored Spaces Model».

The definition of a digital twin from Greaves can be found in the same place: «The digital twin is a set of virtual informational structures that fully describes potential or actual manufactured goods: from its atomic functions to geometry. Under ideal conditions, all the information that can be obtained from the product can be obtained from its digital twin».

Employing digital modelling of high-level correspondence to real test within the framework of the «digital twins» technology, as well as standardized techniques developing for mathematical models validation and analysis of the computational results will allow significant increase the completeness of comprehension. Besides, It will increase the quality of field tests, and reduce their volume, and, in some cases, substitute them by computational substantiation based on the mathematical models validated by the results of multiple experiments. As the result, the possibility originates to reduce the time and costs of the engine certification.

Despite the fact that almost all gas turbine engine units and systems can be modeled, the accuracy of some mathematical models does not yet allow replacing the tests, but not even ensuring acceptable accuracy for making a technical decision on the design change.

Various structures of swirlers, differing by the blades installation angle within the range of 15–60 degrees, were developed for experimental study of mixing processes fr om the vane swirler by the layer-by-layer deposit welding technology.

The manufactured swirlers were blown-through on the experimental test bench with heated air.

The experimental study results indicate a general regularity characteristic for mixing in a swirled jet with surrounding air, consisting in the fact that:

1. With the swirl intensity increase (the vane installation angle increase), within the limits of the studied vane rotation angles, the ejection ability of the flow increases;

2. With moving away from the swirler mouth, the share attached (ejected) air mass increases in the axial direction of the swirled flow.

Based on the works of Akhmedov R.B., Lewis B. and Lefebvre A., mixing in a swirling flow, depending mainly on the turbulent mass transfer process, can be represented as a dependence on turbulent diffusion. It allows forming analytical dependences for mixing process calculation using the following assumptions:

1. The average radius of the swirler R_AV is the radius of the annular source R_CS;

2. A mixture of air and fuel is a gas flowing out of an annular source;

3. The flow swirling effect is being determined by its impact on the coefficient of turbulent diffusion.

Comparisons of the swirlers experimental data with various vane installation angles with analytical calculations reveal satisfactory qualitative and quantitative convergence. Analytical dependence is described by a power function close to linear.

In practice, the impact of the swirler vanes shape on the mixing process is of interest. An experimental study of the vane shape impact on the mixing ratio was conducted. The profiled vanes demonstrated a more uniform temperature field and the highest mixing ratios. Obviously, this is due to the fact that the profiled vanes application allows obtaining a more uniform flow behind the vanes due to the absence of separated flows in the inter-vane channel of the swirler. As the result, a pressure losses decrease occurs during the flow passage through the profiled vanes and, accordingly, an increase in the ejection ability of the jet occurs. It is worth noting that the same result was obtained in the work of Lefebvre A., wh ere the vanes profiling significantly reduces the pressure loss in the swirler.

The conducted experiment and analytical calculation aimed at studying the change in flow parameters depending on the installation angle and the vane profile allowed obtaining the following generalizing results. With an increase in the vane installation angle in the range of angles under study, the ejection ability of the swirling flow increases. The blade profiling strongly affects the temperature field. Unlike the flat ones, the profiled vanes create more uniform flow at the outlet without significant separation zones, reducing thereby hydraulic losses in the flame tube head and ensuring a high mixing ratio with secondary air. A change in the number of profiled and flat vanes has an insignificant impact on the hydraulic resistance change, in contrast to a change in the vane installation angle. Thus, the obtained results of the work may be handy while designing the effective flame tube head of the gas turbine engine combustion chamber.

Five schemes of double bypass engines with changeable working process were considered in the work:

1. A double bypass turbojet engine with an afterburner chamber (DBTEAC), in which the air flow from the third circuit is being supplied directly into the common afterburner chamber;

2. The double bypass engine consisting of the two gas turbine engines. One of the engines is a turboshaft one with a free turbine, which represents the additional turbine of the second engine, which is a turbo-eject one;

3. The double bypass engine with independently controlled third circuit;

4. The Rolls-Royce company double bypass engine with changeable work process, consisting of a central bypass engine and additional modules placed around it, such as bypass turbojet engine or turbojet engine with afterburner.

5. The FLADE VCE double bypass engine of changeable work cycle with extra modules.

Computer simulation of three models of double bypass engines was performed with the ASTRA CAE system, which covers the entire cycle of thermo-gas-dynamic design of a gas turbine engine. The prototype engine was the RD-33 turbojet engine with an afterburner. Besides the thermodynamic calculations, computations of the full flight cycle, mass characteristics of the power plant and aircraft as well as efficiency criteria were performed.

Variation of the degree of both bypass and double bypass values allowed obtaining the values of the total mass of the power plant, and fuel required for a flight at a given range — M_su+t, as well as the fuel consumption in kilogram per one ton-kilometer of transported cargo — C_t.km.

In the course of this computation the conclusion was made that the most rational and favorable ratio of efficiency parameters was obtained from the double bypass gas turbine engine of the FLADE VCE variable duty cycle.

The resulting parameters exceed the values of efficiency parameters of the prototype engine by 13%. These parameters may be employed to perform structural-parametric optimization of parameters to reduce the fuel costs and increase the engines efficiency with a complex cycle, designed for military aviation, on the cruising section of the flight.

The computational models used today require unambiguous (deterministic) values of the initial data in order to obtain a solution. In reality, however, the researcher often does not know the exact value of a given quantity. He knows the results of their direct or indirect measurement, which has a margin of error. Awareness of the fact of the initial data uncertainty may lead to a complete rethinking of the computational study process and the interpretation of its results.

In the conducted study, the authors created a stochastic thermodynamic model of the AI-25 gas turbine engine that accounts for the initial data uncertainty.

As the result of the available set of experimental results generalization the most probable values of the measured engine parameters have been found. Based on these, a deterministic thermodynamic model of the AI-25 engine operating process for the selected operating mode was created. Further, an algorithm was developed and implemented, which transformed a deterministic mathematical model of the AI-25 engine operating process at the operating mode of interest into a stochastic one. It allows determining the scatter of outlet parameter values, knowing the scatter of several inlet parameters. The stochastic model has been built on the assumption that the scatter of uncertain inlet data complied with a normal distribution law. Notwithstanding that the thermodynamic model is relatively simple and fast, it requires a huge number of calls to the initial deterministic computational model, which does not allow obtaining stochastic results for all variables of interest in a reasonable time frame.

For this reason, a stochastic solution was being searched for in two stages. At the first stage, a sensitivity analysis was being performed. As the result, the initial data was ranked according to the degree of the end result affecting. A study, in which computation of specific fuel consumption scattering for 2, 3, 4, 5 and 6 first variables of the series was being performed sequentially, was conducted for the sequence obtaining. The scatter of specific fuel consumption values and other important parameters at the selected engine operation mode was changing insignificantly after accounting for more than five affecting variables. The obtained data was transformed into the bell-shaped bivariant distribution on the graph of the parameter of interest dependence on the air consumption. The obtained data herewith was compared with the similar bell-shaped graph, obtained by the experimental data.

With the conventional deterministic approach, computational and experimental results obtained for the same mode are the points of the graph. Their mismatch is being computed in the form of the two differences (deviations) along the two coordinate axes. Given that the errors of the two points being compared determining are not accounted for herewith, the obtained mismatch has an error, which value is unknown. The stochastic approach allows giving a quantitative description of the mismatch. It represents a bell-shaped bivariant distribution, described by the two parameters: the expectation of the difference and the mean-square deviation for the two coordinate axes.

The unmanned aerial vehicles integration into modern infrastructure systems operation is one of the most urgent tasks in the modern transport industry. Such integration requires the solution of a whole range of problems, including technological, managerial, legal, etc. Among others, the problem of traffic safety can be highlighted, since namely this unresolved problem of the unmanned aerial vehicles traffic is the cause of a number of restrictions on their application. The authors of the presented work proposed a system of directional stability, allowing preventing the unmanned aerial vehicle with movable wing (multicopter) escape from the air passage boundaries available for its movement, which reduces the risk of emergency occurrence with its participation. The system solves the safety ensuring problem for multicopter movement, operating along the preset routs, such as in technological process monitoring systems, goods delivery systems, object video surveillance systems etc. Technological elements of the system being proposed are of small size and do not need electric power supply, which maximally simplifies their implementation to the existing infrastructure.

The proposed system may be of interest to large chain retailers with the goal of employing it in such applications as the goods delivery operating according to the scheme “central logistics center → points of goods delivery in the city”. The system may be employed in applications for industrial facilities monitoring, providing for the movement of unmanned aerial vehicles along certain routes over the territory of the enterprise with additional equipment installed on them, such as scanners, thermal imagers, video cameras, emission detectors, etc. to control technological processes of the enterprise. An additional application trend of the proposed system is safety ensuring of interaction between multicopters and aircraft in the airport area, which is being currently closed for their flights. The system allows ensuring the movement of the multicopter strictly in a given air corridor, which solves the problem of splitting the involved multicopters and other air traffic participants in the airspace.

The world civil aviation development, the traffic volumes increase, and the route network expansion implies, among other things, the quality improvement of aviation security systems, which, at present, acquire utter importance. All this stipulates the relevance of the presented scientific work. The degree of this issue development in scientific terms is not so high, since the problem of aviation security originated much later relative to other problems in the field of civil aviation, and does not have an appropriate scientific basis, which causes certain difficulties. Thus, the article explores the plan staging for the task of airport aviation security system improving based on integration of airport technical protection and air traffic control. The basic idea consists in the fact that at the present stage of their development the air traffic control (ATC) facilities possess strong scientific and technical capabilities of relevant objects detection and tracking, that is not always inherent in the means of aviation security in their area of responsibility. Hence, it is rather promising to explore the issue of joint application of technical means of both systems. Thus, it is necessary to understand herewith the historical incompatibility of these systems, which were created and developed to solve their local specific problems.

Hence, if a task of their aggregation to some extent, or joint application to solve the tasks of aviation security ensuring is being set, it is necessary to form a field of joint mutual interests, in which it will be possible to determine the identity of tasks and to formulate the requirements for shared facilities. Probably, information support for both systems may be their unifying foundation. Then the challenge of developing interface, solving the problem of the systems compatibility occurs. It is impossible herewith to get away from the problem of the compatibility criteria determining and solving many similar tasks. On the other hand , the problem solution of the aviation security systems and systems of air traffic control aggregation even in the first approximation may prod uce a significant effect, and not only economic. The article presents the setting of this complicated task and regards some approaches to its solution. The authors suggest herewith employing standard automated air trafic control systems as the basic structure of the complex system.

Thus, the author proposes to use the typical automated system of air traffic control as the basic structure of the integrated system.

At present, heat-resistant nickel-based alloys have found a very wide application in of energy and aerospace engineering products manufacturing. Their share in the total mass of modern aviation gas turbine engines is particularly large, since they are the preferred materials for production of disks, blades, combustion chambers, and turbine housings.

The article presents an overview of additive manufacturing methods actively employed in the aircraft and rocket engine parts manufacturing. Their classification is presented in dependence on the energy source employed and the source material shape. The advantages of additive technologies in comparison with conventional methods of forming parts and products are described, technology of the parts blanks manufacturing from heat-resistant alloys by direct laser fusion of metal powders is considered. Examples of the of additive technologies successful applicatioin in the aerospace industry in the production of various parts, both for the production of blanks, and in the hybrid, combined with subtractive methods, the technological process of manufacturing complex parts using multi-axis manipulators are presented.

The article considers the main components of the direct laser fusion (DLF) plant, affecting the quality of the resulting workpieces. It describes the existing nozzle designs emplloyed for feeding powder to the fusion zone in DLF installations. Their advantages and disadvantages, as well as conditions for their application are described. The article describes the principle of operation of modern powder feeders for the DLF technology. Parameters characterizing the DLF process and affecting the quality of workpieces forming are presented. Analysis of the defects accompanying of this process was performed, and possible causes of their occurrence were determined.

The advantages and disadvantages of the DLF process of metal powders are described. The main advantages of the DMD process are as follows:

– the laser beam is capable to perform melting and sintering of the material without overheating the substrate and deposited material, i.e., decrease the zone of thermal impact, and diminish changes in the microstructure of the material;

– the high focusing capacity of the laser source allows creating sufficiently accurate workpieces and parts with a wall of less than 0.5 mm;

– the ability to control the laser power, the heat flux density and, consequently, the microstructure of the deposited material allows the DLF process application for repairing complex parts made of a single-crystal nickel heat-resistant alloy.

The disadvantages of the DLF process include the following:

– a low level of mismatch of mechanical properties of the blanks made at different DLF plants from different powder batches under identical conditions of their forming;

– high cost of equipment, which prevents the widespread application of the DLF process in the industry;

– a limited list and low availability of powdery materials, as well as a large range of their quality spread;

– the relationship between the surfacing conditions of powder materials and the mechanical properties of the workpieces is not fully understood.

Modern aerial vehicles are dynamically developing both structurally and in the field of employing the newest materials, which is being associated with the basic requirements imposed on them, such as ensuring minimum weight and increased strength properties at high alternating loads. The most suitable metallic material meeting the above-said requirements is titanium alloys, which are being actively applied in the aerial vehicles framings. Since the 70s of the last century, the aircraft structural elements have been assembled by welding, while all-in-one joints herewith must meet the unified requirements developed for the industry. As a rule, three welding methods are being employed to form permanent joints in the aircraft building industry. They are welding with a non-melting electrode in a protective gas environment (both traditional and submerged tungsten electrode), and electron beam welding.

An immense experience has been accumulated on the these methods application in the aircraft building industry, nevertheless, each of the methods has a number of unrealized potential opportunities to improve the permanent joints quality in the field of warping reducing, crack and pore forming, and mechanical properties enhancing to the level of the base metal. The article presents the results of analysis of publications and the authors’ own research on the above-mentioned problems. The welding modes impact, the introduction of an additional heat source, and mixing intensification of a liquid-metal bath when applying the basic welding methods are considered.

The authors found that porosity elimination occurred with the life span increase of the welding bath, but, with this, the geometry of the weld seam changes dramatically, strength properties decrease up to 15% compared to the base metal.

With the additional heat source introduction, the bubbles degasification occurs, and the permanent joint properties similar to the base material are being obtained.

Currently, the development of electronic control systems and parameters tracking of the permanent joints forming process allows oscillating both the trajectory and welding modes, which allows in its turn introduction of pointed dosing of both energy and welding material into a specific point of the welding bath.

Due to the unique properties of metal melting, the possibility of oscillation allows causing the welding bath to overheating up to boiling temperatures, and cause its intensive mixing, which contributes also to obtaining satisfactory permanent joints with the properties similar to the base metal.

In recent years, the research is being conducted on hybrid or fully electric power plant application on aerial vehicles of various classes without fundamental changes in their layout. However, new trends of modifications in the layout of the power plant with an air propeller emerge at the same time. For example, on the X-57 experimental aircraft the distributed power plant, consisted of small diameter propeller, is being employed at takeoff and landing, while propulsors, located at the tip sections of the aircraft wings, are being employed at the cruising mode. A number of computational and experimental works are dedicated to studying propeller slipstreams interaction in this aerodynamic layout, including favorable interference evaluation. The presented work is devoted to the numerical study of the interference of two-bladed tractor propeller and straight wing with super high aspect ratio of the solar battery aircraft in the non-uniform flow. The work was executed in accordance with the experimental work.

The studies were conducted with the ANSYS FLUENT program, based on the of the Reynolds-averaged Navier-Stokes equations solution, on a structured computational grid (about 20 million cells) with the k-ε-realizable turbulence model, with improved turbulence parameters modelling near the wall and with account for the pressure gradient impact. Computations were performed at the flow velocity of 25 m/s and 50 m/s and Reynolds numbers Re = 0.17 and 0.35·10⁶. The angles of attack in the computation were being varied from 1° to 7° at the zero sideslip angle. Three aircraft configurations were considered: without propellers, as well as with running propellers with diameters of 0.22 m and 0.33 m. The rotation speed of the two-bladed pulling propeller as fixed for both options, and it was N = 15000 rpm. The presented work regarded symmetric rotation of the propellers at the wingtips in the fuselage direction.

Numerical studies of the interference between the propellers and the high aspect ratio wing revealed that the propeller diameter significantly affects the flow-around and aerodynamic characteristics of the aircraft of this configuration. Installation of the propeller leads to a decrease in the lift in the range of cruising angles of attack under study, the pitch moment herewith increases by nosing-up. The induced drag increases with the angle of attack increasing, while the propeller rotation enhances the nonlinearity of the С_xai (α) dependence at the incoming flow velocity of 25 m/s. The article demonstrates that the induced drag reduces depending on the propeller diameter, since the propeller rotation (in this case in the same direction, as the vortex behind the engine nacelle), introduces perturbation into flow-around, and straightens the flow behind the wing. With the propeller diameter increase, the dependence of the relative circulation over the wingspan moves away from the elliptical kind, and the incoming flow speed increasing only strengthens this difference.

The problem of community noise of propeller-driven unmanned aerial vehicles (UAVs) should be considered separately for civil and special-purpose vehicles.

Currently, there are no international standards regulating maximum permissible community noise levels of civil propeller-driven unmanned aerial vehicles (UAV), and low-noise levels are primarily a competitive advantage. The UAVs noise levels normalizing by the analogy with light propeller aircraft is possible in the future.

For the special-purpose propeller UAVs, the problem of acoustic signature is important. It is necessary to ensure the domestic aircraft invisibility when flying along a given trajectory, and to be able to acoustically localize the enemy’s UAVs identifying herewith the UAV type and determining the trajectory of its movement in real time.

In the framework of the propeller UAVs acoustic visibility estimation and while developing standards on the community noise the article suggests employing two units of measure, namely the A-weighted overall sound power level and the overall sound pressure level in dBA. The A-weighted overall total sound power level does not depend on the distance and cannot characterize the acoustic signature, which depends on the distance of the object from the radiation detection point and environmental conditions. At the same time, one may proceed from the spectrum of the acoustic power of the sound source, knowing its direction diagram or assuming it spherical, to the UAV noise level evaluation in the far acoustic field at the given atmospheric conditions and distance. Besides, the total level of acoustic power in dBA can be implemented for the comparative assessment of the degree of acoustic signature of various UAVs of the same class.

A technique for assessing the acoustic signature boundaries of the UAV is proposed. The following items became components of the technique: the noise models of the main sources or experimental data on the UAVs noise, data on the ambient noise, criteria for acoustic signature of various types of UAVs, as well as the software for assessing the aircraft community noise.

Environmental requirements, such as limits on community noise and emissions, will play an increasingly important role in the future of civil aviation. The possibilities of noise reduction in state-of-the-art layouts are limited, thus, it may be necessary to switch to radically new schemes to meet the goals declared by NASA, ACARE, the Ministry of Industry and Trade of Russia and other organizations for the next generation of aircraft.
Engine noise is one of the main factors in the overall aircraft noise. Although the current trend to increase the bypass ratio turbojet leads itself to the noise reduction, the possibility of placing large engines under the wing is limited. The upper position of the engines may help to eliminate this problem and additionally reduce the noise on the ground due to the shielding effect. Besides, the engines diameter increasing does not lead to the chassis struts elongation, i.e. there is a possibility of installing engines with ultra-high bypass ratio. Air intakes are better protected from foreign objects, especially on runways of poor quality. There is no gap in the slat spanwise, as in the layouts with engines under the wing. The jets of the engines do not fall on the flaps. The disadvantages include a significant risk of adverse aerodynamic interference, especially at transonic speeds, and increase in the cabin noise, which may require installation of additional sound-absorbing structures. Moreover, the thrust of the engines creates an undesirable negative dive moment at takeoff and in cruising flight. Many questions arise concerning rational design of the pylon-wing-nacelle assembly and its aero-elastic characteristics. Finally, the engine maintenance becomes noticeably complicated.
Intensive research on «quiet» layouts has been initiated in the US and Europe to meet the stringent environmental requirements of NASA and ACARE for the decades to come. TsAGI also conducts systematic research in this direction, trying to make allowances for the development of necessary technologies in various disciplines, especially in aerodynamics and power plants, since aerodynamics is the main bottleneck hindering introduction of the top-mounted engine layouts. This problem solution with a positive result is possible only with a powerful set of aerodynamic design tools. The set should include a detailed direct analysis method that accounts for all geometric features, an optimization procedure, and a reverse method, allowing create the aircraft surface element according to a given pressure distribution. The authors use in their practice the original version of the residual correction method, in which the upper level is represented by the RANS method, and the inverse method based on the full potential method is used as a corrector.
The article discusses the aerodynamic design features of various aircraft layouts with the engines location above the wing. In general case, their aerodynamics are more complex due to the possibility of adverse aerodynamic interference manifestation caused by the increased speeds over the wing. Thus, it is necessary to search for such configurations in which this risk is minimal, or even there is a chance of positive interference. Several aerodynamic models were designed, manufactured, and tested in TsAGI’s large transonic tubes. These included:
— the regional aircraft layout with natural flow-around laminarization of the wing of a small sweep (χ¼ = 15°)  with the cruising Mach number of M = 0.78. Aerodynamic tests in the T-128 WT (Wing Tunnel) demonstrated satisfactory transonic aerodynamic characteristics, including the possibility of obtaining extended laminar sections on the wing consoles, as well as excellent load-bearing characteristics at low speeds;
— the layout of business aircraft with a drop shape of the fuselage called a «tadpole», with a maximum cruise Mach number of M = 0.82 and a small wing sweep (χ¼ = 6°), with a normal distribution of the relative thickness (`с = 14–10% at the root and at the end respectively). Tests in the T-128 WT fully confirmed the speed properties of the layout;
— the layout of the «flying wing» with the engine nacelles located above the wing center section, designed with account for the unfavorable aerodynamic interference of the wing-pylon-nacelle assembly.

The article deals with the project of a heavy helicopter, being one of the transport system elements of the Arctic zone of the Russian Federation. The helicopter is being created based on the PD-12V prospective domestic gas turbine engine.

The software for helicopter appearance forming, which represents a set of jointly operating modules of weight and aerodynamic calculation, was employed for the carrier system parameters selection.

The dependences of rafts, emergency water touchdown, and thermal and sound insulation weight on the helicopter weight were obtained in this work. Various combinations of the main rotor diameter values and blade aspect ratio for the selected transport operations were analyzed. Optimal values of the helicopter main rotor parameters have been selected using the reduced criterion of the helicopter efficiency.

The project helicopter outdoes the Mi-8AMTSh-VA Arctic helicopter and Mi-26 helicopter by its performance characteristics by either loading capacity and flight range, or flight hour cost. The proposed methods for the helicopter, performing the specified set of transport operations, appearance forming can be employed hereafter while other prospective rotary-winged aircraft of vertical takeoff and landing design.

Measuring total forces and torques affecting the helicopter tail rotor became an up-to-date task of aerodynamics with the advent of the interest to studying «spontaneous» left rotation of single rotor helicopters.

A strain gauge balance is employed to measure the six components of the total aerodynamic force and moment. As far as the case in hand is the loads on the rotating propeller measuring, the strain-gauge balance should be a rotating one (RSB) to measure the six components. The article presents the results of the further development of the spoke-type RSB design with twelve measuring beams, which were presented in the earlier works of the authors. The article demonstrates that the structure consisted of the twelve measuring beams is scalable and applicable with various combinations of the expected loads, affecting the propeller in rotation. Besides, the anticipated places for the strain-gauge gluing are shown demonstrably, and the scheme of their connection into the Wheatstone measuring bridge is proposed.

Computations revealed that components interaction in such structure are minimal at maximum value of signal stresses in the supposed places of strain-gauge resistors gluing. Besides this, the strain-gauge balance design ensures high strength factor no less than four.

The expected errors of the six-component RSB proposed in the article are no worse than 1% of the measurement range. The further development of this work will be the RSB calibration, and the study of characteristics in rotation on a special test bench.

When considering practicality of works unfurling on one or another project implementation, the possibility of this project realization should be assessed mandatory along with its financial or other feasibility assessing.

The project realizability is understood as the capability of solving the necessary set of scientific and technical, planning and design, production and technological and organizational tasks to fulfill due-by-date the full scope of works, ensuring creation of a new or modernized aviation complex (AC).

A great variety of factors affects the AC realizability. The following basic factors can be outlined among them:

— Technical realizability;

— Scientific and technical capabilities of the design bureau (organizational and technical realizability of the project);

— Production and technological capabilities;

— Financial feasibility.

The realizability assessment of science-intensive projects is performed on the based on assessments of the main types of risks present while the projects implementation. Risk levels of a program implementation are the estimated value of the factors of various nature impact on the end result of the program in terms of the target indicators achieving. The main target indicator for the program implementation is of the selected version of the AС timely creation, meeting the requirements of the tactical and technical assignment (TTA).

The state-of-the-art techniques application for the complex comparison of the aircraft should be performed in conjunction with the aircraft flight performance (AP) realizability. The flight performance realizability is understood as the probability of achieving the flight performance characteristics declared in the design specifications. To determine the probability of the AP achieving, knowledge of a distribution law for each characteristic is necessary, and these laws are affected herewith by the distribution of the input parameters. The input parameters distribution can be obtained based on statistical data, mathematical modeling, as well as by the expert assessments method. As far as the highlighted risk factors are being affected by many random events, the distribution law of these factors is assumed to be normal. The main feature of the normal distribution law is that it is a limiting law, which is being approached by other distribution laws under rather frequently encountered typical conditions. The presented technique includes in its algorithm the first technique for the appearance forming, and accounting for the risks of the AP achieving specified in the design specfications is an additional module to the existing techniques. This module allows assessing the risk of flying performance realization and account for these risks directly while the aircraft appearance forming. The obtained formulas establish interrelation between the required flight performance changes and parameters of distribution laws of the risk factors.

The account for the risks of the AC creating is a necessary element when comparing the AC options, as well as while assessing the program implementation as a whole. The approach described in the article to the accounting for the risks of an aircraft creating at the early stages of development allows assessing the likelihood of the program implementation in terms of achieving flight performance by the aviation complex.

This study results application to supplement the general technique allows complex comparison of the AC options under the impact of the probabilistic (random) factors.

The presented work considers the passive part designing for the cooling system of the spacecraft onboard complex.

The equipment of cryogenic and helium temperatures level, necessary for ensuring standard operation conditions [3, 4] characteristic for the deep space, external solar radiation and instrument-hardware electromagnetic emissions, is frequently employed in thermal control systems, ensuring the thermal mode [2], for the state-of-the-art space platforms [1]. The telescope being designed will be capable of operating in both the single telescope mode and as a part of the interferometer between the “Earth-Space” bases (with ground-based telescopes). The telescope operation range is from 20 microns to 17 mm [5–7].

The observatory is planned to operate for three years with the reflector temperature of 4.5 K, and then for another 7–10 years with the total temperature of 50 K [8]. The term of the observatory active life is ten years. The reflector thermal mode sustaining is being implemented by the observatory cooling system, consisting of passive screens and Stirling and Joule-Thomson cryogenic machines.

The thermal model and the design scheme are being considered on the example of the passive cooling system of the onboard complex of the “Millimetron” observatory scientific equipment. The general cooling system includes both the active part, represented by the heat exchange units, removing heat from the cryoscreen and equipment to the Joule-Thomson and Stirling machines, and the passive part, represented by the protective screens system and reflective surfaces, removing the heat to outer space. The account of the joint operation of both parts is necessary for the characteristics analysis.

The main portion of the neat inflow from the solar radiation and instruments is being removed toe the space by the passive cooling system. The heat transfer computation while efficiency estimation of the telescope passive cooling system represents a complicated problem, primarily, through the necessity to account for the complex geometry, the possibility of heat inflows along the system elements, and thermo-physical properties of the screens. This problem solution can be obtained only by the numerical methods with the visibility coefficients determination of individual elements between themselves and with the outer space.

The cooling system computation is being complicated by the following factors:

- complex geometry of the passive screens and cryoscreen, their position in space and relative to each other;

- large temperature gradients from 320 K to 4.5 K between the elements, leading to the presence of temperature deformations of the structural elements;

- thermo-optical coefficients the thermo-physical characteristics of the elements are strongly dependent on temperature as well;

- the presence of three different thermal control mechanisms, namely, passive protection employing cryogenic screens and cooling by cryogenic machines of various temperature levels.

All these reasons stipulate the need for the expanded thermal analysis of the cooling system with a mathematical model developing to determine the cooling efficiency and temperature fields of the system elements.

Thermal bonds identification is necessary for correct developing of the mathematical model and obtaining numerical characteristics of the cooling system. The structure under study consists of individual elements such as screen lobes, cryoscreen, reflector, frame, etc. Each element of the system possesses the thermal bonds: radiation, internal thermal conductivity due to the presence of temperature gradients within the element itself, thermal conductivity through the frame or thermal bridges with neighboring elements.

The temperature values were obtained for each structural element. However, within the limits of one screen they differ by no more than 1 K, since the model is centrally symmetric. This difference is associated with the calculations error.

The spacecraft thermal control system, ,with the “Millimetron” observatory positioned on it ensuring the required reflector operating temperature of 4.5 K,  was developed. These temperatures values allow estimating the passive cooling system efficiency. However, more accurate forecasts require the computations correction by increasing the number of finite elements, and considering thermal conductivity of the passive screens materials and complex structure of the thermal bridges.

A wide variety of spatial-truss structures, including pyramidal and X-type trussed cores was developed at present in attempts to create multifunctional core materials of the three-layer structures of aerospace purpose. Computational and optimization methods of these typical trussed cores’ characteristics were considered in many scientific studies. However, very few comparative studies of such core materials mechanical characteristics were conducted. The presented article compares compressive mechanical characteristics of the X-type and pyramidal trussed cores by both analytical and experimental methods. In experimental phase of the study, the two samples of three-layer structures were produced: one with the pyramidal core and the other with the X-type core, to determine the ultimate compressive strength.

3D-models of the samples were designed with the SOLIDWORKS software for manufacturing. Sketches were obtained, and pattern cutting of flat elements was performed based on these models. Further manufacturing was being perpetrated by the flat figures cutting from the aluminum sheet on the laser-cutting machine. Samples for the experiment were assembled from the cut elements. The flat elements fixing with each other is being brought about by the «spike-groove» technique to simplify assembly operations. The assembled samples of the three-layer panels were tested alternately under similar conditions, on the same machine tool. Further, based on the results of compressive testing the «stress-deformation» diagram for both cores was obtained and analyzed. From these diagrams, critical compressive stress and stiffness of the cores were determined. The results of the conducted experiments are in good agreement with the results of analytical calculations. The obtained results demonstrate that with equal relative densities of the cores and similar slope angles of the cores the generalized critical stress of the X-type trussed core cannot be less that the generalized critical compressive stress of the pyramidal trussed core (and at the small relative densities it can be four times more). However, under the above said conditions their generalized compressive stiffness is the same in all cases.

The article presents a design procedure for aerodynamic load acting on the gliding parachute while its deployment and reloading to the airdrop object. The computational dependencies, which can be employed for quantitative estimation of these parameters, are presented. The average operational (aerodynamic) load and the upper confidence limit of the aerodynamic load acting on the gliding parachute while its deployment are the basic initial parameters when calculating the strength of gliding parachutes.

This information is utterly important while the parachute strength calculating and its appearance forming. The problem statement is as follows. To form, as a first approximation, methodological recommendations for calculating the aerodynamic load on the gliding parachute and the reloading on the airdrop object in the process of the parachute deployment, which can serve as a basis for further scientific research on the proposed method refining and adjusting. The article presents the main definitions and assumptions, as well as the method itself in the engineering statement. Maximum value computing of the axial overload acting on the landing object is based on a semi-empirical dependence that adequately reflects the integral average of the maximum overload value during the gliding parachute deployment.

While developing the engineering mathematical model of the dome (wing) filling of the gliding parachute, the theoretical part supposed that aerodynamic load on the dome (wing) is an additive function of three, practically simultaneously occurring processes. They are:

— impact loading of the lower wing shell, due to the jet of the incoming flow impact, its spreading and the lower wing generatrix straightening forming a local stretch of the lower shell;

— the air intakes filling in the of the stretched part zone of the lower shell; the local zone forming of the executed part of the upper shell and the wing profile;

— loading the completed part of the upper shell (the formed part of the wing) by the pressure drop while its flow around by the external flow.

The article presents computing dependences of the overload acting on the airdrop object on various parameters (the parachute area; the object mass; the height; and the speed of bringing the system into action) for both cargo and human parachute systems. While computing a number of empirical coefficients, the computations used the results of data processing of a vast number of flight experiments with both human and cargo parachutes.

A brief algorithm for the parachute strength computing a when forming the shape of a gliding parachute is given.

The results of the presented work may be useful for designers, testers, calculators, and scientists working in the field of parachute building and engaged in the gliding parachute systems design and testing.

The presented article considers the possibility of external airbags application on a helicopter to enhance the crews and passengers survival rate under conditions of the helicopter emergency landing.

The helicopter emergency landing modelling was performed by the finite element method using the scheme of explicit time integration. The analysis includes the helicopter hitting the hard landing surface at the speed of 17.2 m/s. The values of overloads at the helicopter center of mass and main gearbox, as well as the general impact of airbags on the helicopter fuselage deformation were determined by the crash test results.

Finite element modelling of the airbag curdling was performed to determine the time of the airbag gas filling. A mathematical model determining the gas source characteristics was developed in MATLAB Simulink. Mass flow rate and temperature of the gas were determined. Finite element modeling of the airbag filling with gas was performed.

The article cites the main disadvantages of the external airbags application on helicopters. It presents statistical data on aviation incidents of helicopters of various categories. Significant fuselage deformation reduction at the external airbags application is demonstrated by the results of the study. In conclusion, the inference is drawn on the positive impact of the external airbags on the survival rate of the humans onboard of the helicopter.

The main limitations of the external airbags application on a helicopter and statistical data of aviation incidents with various categories of helicopters are presented. According to the research results, a significant reduction in fuselage deformations when using external airbags has been shown. Finally, the conclusion is made that the positive effect of external airbags on the survival rate of people on board the helicopter.

The main issue while rocket engine design, particularly the solid propellant rocket engine (SPRE), is ensuring indispensable engine characteristics, during which operation the probability of acoustic instability occurrence at various modes cannot be excluded. Application of various shapes of the solid propellant channel, grooves as well as combustion products flow swirling inside the engine, which, in turn, may both reduce the probability of the acoustic instability occurrence and increase it, facilitates this. The presented article considers the SPRE, which distinctive feature consists in the presence of the controlled flow swirling inside the combustion chamber.

The purpose of the work was studying impact of the swirled flow and various shapes of the classical inlet subsonic sections of the nozzle on the flow coefficient and forming recommendations for their application.

The state-of-the-art techniques of computational aero dynamics were employed for studying the flow coefficient of classical subsonic nozzle sections under the swirled flow impact. Numerical modelling was being performed employing classical models based on averaged Reynolds Navier-Stokes equations (RANS), which ensure optimal relationship between the obtained results accuracy and resource intensiveness. The RNG k— turbulent model with typical set of model constants, able to ensure the required accuracy according to declared goal and adopted assumptions, namely quasi-stationary axisymmetric adiabatic approximation of the ideal-gas formulation was being employed in the presented work.

Geometry of the computational model supposed application of classical subsonic nozzle sectors (bottoms) with variable parameters of the subsonic jet narrowing, inlet section, from which the swirled flow boundary conditions were being set, unchanged geometry of the supersonic part of the nozzle and extra volume behind the nozzle cutoff. The grid quality was being maintained constant when the computational model geometry changing.

Classical bottoms with conical, elliptical and flat shapes of the nozzle subsonic part, as well as the contour designed with Vitoshinsky formula were being studied in this work. The swirled flow intensity, characterized by the Higher-Baer coefficient Sn, was the boundary condition for the combustion products flow at the nozzle subsonic part inlet. The dependencies of the flow coefficient on the swirled flow intensity at various shapes of the nozzle subsonic part were obtained.

The results of flow characteristics of the subsonic sectors contours under study are being compared with each other at the same swirled flow intensity. The article shows that the swirled flow intensity increasing at the nozzle subsonic part inlet up to Sn = 0.4 leads to the flow coefficient decrease by no more than 0.14%. The largest flow coefficient and more uniform velocity profile in the minimum section when the swirled flow feeding corresponds to the Vitoshinsky contour due to the smoother contour to the minimum nozzle section inlet. Recommendations on the parameters of the transitional sector from the cylindrical part of the chamber to the bottom contour and throat section of the inlet to the minimum section for various bottom shapes are presented. Radius of the inlet to the throat section minimum section has the greatest impact on the flow coefficient.

The article presents the results of the compact propulsion unit developing for correcting nanosatellites of the CubeSat format based on a low-thrust gas thruster with the weight of no more than 2 kg, the overall size of no more than 1,5U, and peak energy consumption of no more than 10 W. The correcting gas propulsion unit is accomplished in the form of a monoblock. The unit has diminished size and ensures herewith the total thrust impulse of no less than 65 N·s due to application of the compressed Nitrogen with the pressure 35.3–39.2 MPa (360–400 kgf/cm²), with the initial weight of 0.09 kg as a working medium, and composite tanks for its storage with total volume of 0.25 liters. With the satellite weight of about 5 kg the characteristic velocity changing will be 12.5 m/s. In the course of the work, the experimental studies of the unit’s constituent parts, namely newly developed low-thrust engine of the electrical storage type, consisting of the chamber with the gas-dynamic nozzle and a small-size low-pressure control valve, start valve and a high-pressure control valve. The thrust of the developed engine is a function of the working gas pressure at the engine inlet. It changes from 0.196 N (20.0 gf) at the pressure of 578.5 kPa (5.9 kgf/cm²) to 0.098 N (10 gf) at the pressure of 313.7 kPa; the thrust specific impulse in the continuous mode is of no less than 687 m/s (70 s) at the working gas temperature of 20°C. Instead of pyro valve A newly developed start valve with shut-off element from the shape memory effect material, which energy consumption is of no more than 5 W was applied in the unit instead of the pyro valve. To adjust the working media in the receiver, the control valve with flow limiter, which limits consumption at working pressures from 14.7 to 39.2 MPa (from 150 to 400 kgs/cm²) is applied. It allowed reducing the valve energy consumption by 3.1 W, and decreasing the unit peak energy consumption by 26%. Instead of large-size filling necks, a filling unit with the weight of no more than 48 g was developed.

Its main elements are a closure (metal-to-metal seal), a check ensuring safe operation of the device when propellant is being filled and vented, and a plug, which guarantees the the device tightness during operation phase after its tightening to the nominal torque at production phase. As the result of the presented work, a practical prototype of a small-sized gas propulsion system on compressed nitrogen was developed and designed to generate impulses to transfer a nanosatellite from the launching orbit to the target orbit, to maintain the required orbit during a specified nanosatellite lifetime and its exit from orbit.

Up to now, a problem of parameters’ accurate prediction at large Reynolds numbers is existing in gas dynamics science. The Navier-Stokes equation of motion is practically unsolvable with modern technology due to the lack of computational resources. With the Reynolds number increase, application of the finer mesh with small computational cells is necessary, which makes it almost impossible to calculate even elementary problems when employing direct numerical modeling.

Transition to solving simplified equations of motion is widespread. Reynolds-averaged Navier-Stokes (RANS) equations became the most popular. However, this approach is only a subterfuge containing inconsistencies while describing the true picture of the flow due to many assumptions. Besides, Reynolds equations are not substantiated experimentally. Nevertheless, practically all Russian and foreign electronic products of computational gas dynamics, such as: “Ansys”, “FlowVision”, “OpenFOAM”, etc., are based on the RANS equations.

Thus, an alternate approach to the turbulence description is being proposed. More understandable and physical like is the approach where turbulence is being characterized as a vortex flow, i.e. a flow in which rotational motion and torsion exist aside fr om the translational one. In other words, the flow will be laminar wh ere rotation and torsion do not present.

The article presents both computation and analysis of the gas-dynamic characteristics of a liquid-propellant rocket engine for laminar flow, with the purpose to realize a physically correct task, and significantly reduce the computational time by employing simpler equations. The studies were conducted in the laminar sublayer near the wall of the model chamber of a liquid-propellant rocket engine. The purpose of the work consisted also in writing a program code for obtaining the characteristics of the velocity field and its qualitative comparison with the computational results with the “Ansys” software package.

A system of equations for laminar flow consisted of the equations of continuity, motion and energy in the Poisson form is compiled and programmed in the Python programming language in the work being presented. Computation is performed for the chamber. The region of two by two cm and 41 by 41 mesh points is being set. The boundary conditions were being set in the form of the condition adhesion on the wall, tracking on the centerline, and artificial flow limiting at the outlet. Initial conditions are the longitudinal of u = 100 m/s and transverse of v = 0 m/s velocities, dynamic viscosity of μ= 10^–4 Pa·s, the initial densities field value of ρ= 6 kg/m³.

The computational results were analyzed with the “Ansys” program. For this purpose, the flow computation near the wall was performed for the combustion chamber using the default turbulence model. As the result, the hypothesis for the laminar sublayer existence near the wall was confirmed, which substantiated the statement on the laminar flows application correctness while this program developing. The presence of this fact is of great importance in many computations such as computations for friction, heat exchange, and carried-away wall destruction. The computation of the flow near the wall, using the laminar model, was performed as well.

To assess the adequacy of the results obtained by the developed program, computations were made using the Euler equation. The velocities of the ideal gas obtained with the Euler equations are 3% greater than for the laminar case.

The profile obtained for laminar flow by the “Ansys” program qualitatively repeats the profile calculated in the equation program code in the laminar formulation.

The current lines concentration near the wall can be observed in the velocities field, which confirms the presence of a boundary layer, and the lines parallelism indicates its laminarity.

Thus, the following conclusions can be drawn:

1. A method and a program for the gas-dynamic characteristics computing of the liquid-propellant rocket engine for laminar flow are developed;

2. Testing with the “Ansys” program revealed a qualitative match with the calculations by the developed program;

3. The linear dependence of the velocity profiles near the chamber wall (the presence of a laminar sublayer) is shown;

4. The difference in absolute velocities due to the viscoelastic term is estimated at ~3%, which corresponds to the gas-dynamics losses of the specific thrust momentum.

The presented work studied the impact of the stick type (side stick or central stick) and parameters (stiffness and time delay). The difference between the «command signal by the displacement» control, and the «command signal by the force» control was studied for each variable as well. Each study was being conducted on the stationary simulator, when the operator performed the task of pitch and tilt control. The main part of the studies is being conducted with account of the sensory system characteristics (the force gradient) and the gain of the controlled element (the control stick sensitivity), which is being selected according to the operator’s judgment. The study was emphasized enough on revealing the difference between the control signal transmission type to the flight control system for both control types, namely by the displacement and by the force. The major portion of the study related to the error dispersion dependence revealing associated with by the stick type (side stick or central stick) and command signal (DSC or FSC).

Switching from the command signal by the displacement to the signal proportional to the force reduces the error dispersion by 30–50%.

For the longitudinal channel, switching from the DSC stick to the FSC one leads to the three times error dispersion reduction, the throughput band increase by 60-70%, and cut-off frequency increase by 10-30%.

The operator evaluates the control object by the Cooper-Harper scale at the level of 3-3.5 PR employing the central DSC stick. When working with the DSC side stick, the estimation is 2.5-3.0 PR. Switching to the signal proportional to the force leads to the estimation improvement for the central stick by 0.5 points and by one point for the side control stick.

For the lateral channel, switching from the DSC stick to the FSC one leads to the two times error dispersion reduction, the throughput band increase by 25%, and cut-off frequency increase by 10%.

The operator evaluates the control object by the Cooper-Harper scale at the level of 4-4.5 PR, steering with the central DSC stick «control by the displacement». When steering with the DSC side stick, the estimation is 4.5-5.0 PR. Switching to the signal proportional to the force leads to the estimation improvement for the central stick by 0.5 points and by 2.5-3.0 point for the side control stick.

The article presents the description of the study, consisting in assessment of the aircraft motion control quality by mathematical models of pilots actions while simulation, and a pilot-operator while semi natural modelling. Simulation modelling includes the following:

1) mathematical model based on the fuzzy sets theory;

2) mathematical model based on the theory of fuzzy sets with optimized parameters by the Broyden-Fletcher-Golfarbd-Shanno method;

3) mathematical model in the form of transfer functions.

The purpose of the study consists in creating a method for assessing the aircraft flight control.

The result of the study is the values of the root-mean square deviation (RMSD) of the of the aircraft movement kinematic parameters of the reference sampling of parameters (with the ideal fulfillment of the target piloting task) from the results of simulation and semi natural experiments. The places ranged by the RMSD ascending were assigned to mathematical models and semi natural experiment of the parameters under study to determine the best implementation by the quality and nature of control. All places were being added up. The implementation with the lowest sum is the best by the control quality and nature, which is imitation simulation of mathematical model, based on the fuzzy sets theory with optimized parameters (the sum of places equals to five). It has minimum RMSD by the three parameters. It occupies the second place in the ascending order.

Thus, a mathematical model based on the fuzzy sets theory with optimized parameters possesses all advantages of the mathematical model, based on the fuzzy sets theory (logicality of control). In other words, the dependence of the input parameters on the output ones is expressed by the logic rules, which allows the nonlinear system control, while its implementation simplicity does not require complex mathematical apparatus. The optimization algorithm allows compensating the disadvantage, such as the low quality of control, of the mathematical model base on the fuzzy logic theory.

The presented method for assessing the aircraft of movement control quality may be used for selecting a mathematical model of the pilot’s control actions, employed for studying the kinematic parameters of the aircraft movement at a specific target piloting task

Keywords: mathematical model of the pilot’s control actions, root-mean-square deviation of kinematic flight parameters, motion dynamics model of modern maneuverable combat aircraft, piloting-modelling test bench of a modern maneuverable combat aircraft.

The article is devoted to a new gas nitridation method, which allows obtaining high-quality diffusion layers, meeting the requirements for operation of the products that running under severe conditions of sharp temperature changes and large sign-changing loads, particularly, for aircraft parts. The method consists in a combination of various temperature regimes at the ammonia and air concentration change in the furnace working part.

The authors propose the three-stage technology for the 03Cr11Ni10Mo2Ti steel nitridation. The first state ensures the surface restoration, oxides destruction, and guaranteed nitrided layer creation.

The high activity of the saturating atmosphere is being achieved by reducing the ammonia dissociation degree, as well as air oxygen binding with hydrogen while the ammonia decomposition. These processes ensure forming continuous nitrided layer on the surface The second stage ensures the passage of intense diffusion processes at a temperature of 550-600°C due to additional thermal cycling when concentration of the working mixture changing.

The second stage duration is being determined by the required thickness of the diffusion zone. In the atmosphere of the pure ammonia, the third stage allows resolving to a certain extent the hard and brittle high-nitrogen surface layer, which itself becomes the source of nitrogen at the low activity of the saturating atmosphere. Nitrogen reflux inward the metal and reduction of its content on the surface begins herewith. The stage of diffusion allows the phase content changing of the surface, and reduce its brittleness due to the certain hardness decrease and plasticity increase, which excludes micro-cracks appearing on the ready parts, i.e. fulfill the task set by the industry.

There are pre-eutectic (< 12 wt.% Si), eutectic (~12 wt.% Si), hypereutectic (> 12 wt.% Si) silumins. The structure of hypereutectic silumin consists of eutectic, primary grains of silicon, and intermetallic compounds based on iron, copper, etc. These elements are impurities getting into the alloy at the stage of melting from the charge.

Hypereutectic silumin is being employed in many branches of mechanical engineering as a material with good casting properties, which allows casting products of complex shapes. Low thermal expansion coefficient, high corrosion and wear resistance contribute to this alloy application as a material for plain bearings and pistons manufacturing.

Defects of macro and micro size pores and cracks emerge at the stage of casting. The size of the primary silicon grains reaches up to 100 microns while the castings cooling. The traditional methods application, such as alloying, changing the casting method, lead to the final product cost increasing, and restrictions on the casting shape appearing. Methods of materials’ high-energy processing ensure the surface recrystallization and of micro- and nano-crystalline structures forming.

The purpose of this work consists in analyzing the results obtained in mechanical tests performed under conditions of uniaxial tension of plane proportional hypereutectic silumin samples, subjected to a pulsed electron beam treatment.

The hypereutectic silumin alloy was prepared in a shaft type resistance laboratory electric furnace with silicon carbide heaters in a painted stainless steel crucible. The silicon content was 20 wt.%.

The obtained castings represented rectangular plates of the 55x120x20 mm size (without account for sprue), from which the samples of 15x15x5 mm size were being cut, as well as flat samples for the tensile tests.

Mechanical test of silumin were being brought about by the samples uniaxial stretching with the «INSTRON 3386» testing machine at a constant speed of 2.0 mm/min.

The studies of elemental and phase composition, the structure of the fracture surface were being performed by scanning electron microscopy («Philips SEM-515» and «LEO EVO 50» instruments) and transmission electron diffraction microscopy («JEOL JEM-2100F» instrument).

Due to the heating and cooling rates, the pulsed electron beam treatment allows for surface remelting, leading to the recrystallization of the layer up to 100–120 microns. The modified layer has a multiphase submicro-nanoscale structure, represented by high-speed crystallization cells separated by interlayers of the second phase, and globular silicon inclusions, which sizes vary from 1 µm to 2 µm.

The article presents the studies of the samples fracture. The main cause of destruction has been revealed. The processing mode, leading to a multiple increase in plastic properties, without loss of strength properties was determined.

Titanium (IV) oxide nanopowder / epoxy polymer (n-TiO₂/epoxy) nanocomposite films of 80-100 microns thickness were produced by adding n-TiO₂ to the mixture of epoxy resin ED-20 and 4,4’-diaminodiphenylmethane (DDM) used as a hardener with subsequent curing. Phase composition, structure, and microstructure of the obtained nanocomposites were being studied by X-ray phase analysis (XRD), scanning electron microscopy (SEM), infrared (IR) spectroscopy, and ultraviolet and visible spectroscopy (UV-vis). The phase composition of n-TiO₂ particles and n-TiO₂/epoxy resin composites, determined by the XRD, revealed the presence of two titanium (IV) oxide polymorphic modifications: anatase and rutile. The XRD patterns of the composites exhibit typical diffraction peaks for the cured ED-20. Based on the data obtained and using the Debye-Scherrer formula, the average nanocrystallite size was calculated to be 45 and 140 nm for the initial nanoparticles and those incorporated into polymer (4.2 wt.%), respectively. Apparently, aggregation of n-TiO₂ at this concentration leads to formation of microcomposite. XRD results agree with the data of scanning electron microscopy.

The particle size distribution histograms generated from the SEM data exhibit that while the n-TiO₂/epoxy resin formation, the diameter of the particles increases from 46 nm to 80 nm for the initial n-TiO₂ powder and the composite respectively, even at a relatively low nano-filler concentration of 0.5 wt. %. An increase in the n-TiO₂ size occurs possibly as the result of the nanoparticles aggregation processes.

The structure of the obtained n-TiO₂/epoxy resin nanocomposites was confirmed by the IR spectroscopy data as well.

Adding n-TiO₂ slightly changes the DSC profile of the pure epoxy resin, moving the peak maximum corresponding to the curing reaction towards lower temperatures. The reaction enthalpy increases from 98.8 kJ/mol to 119.3 kJ/mol.

The n-TiO₂ particles may have a twofold effect on the cure kinetics of the ED-20 resin. The presence of hydroxyl groups on their surface should accelerate the curing reaction. On the other hand, hydroxyl groups of the n-TiO₂ are capable of forming intermolecular bonds with epoxy resin, reducing the reactivity of epoxy groups in reaction with DDM and integrating into the forming network, possibly generating more complex structures. The detailed mechanism of such processes requires further studies.

Photo-activity of the n-TiO₂/epoxy resin nanocomposite under the UV irradiation was studied.

The article considers the issues associated with clarification of terms concerning thrust reverse, and requiring refinement in view of formulations and comprehension inaccuracy:

• factor of reversing;
• aircraft reverse capacity;
• optimal value of the engine reverse thrust; < li>reversing device efficiency.

The existing values of the factor of reversing R = = 0,4...0,5 do not indicate the degree of the reversing device (RD) structural perfection, as is commonly believed, but rather their gas-dynamic imperfection, since, significant losses of the total pressure of about 50% arise while the gas flow U-turn in the reversing devices.

The aircraft reverse capacity (Qrev = R/Glw), where R is the reverse thrust value and Glw is the aircraft landing weight, also cannot represent the factor, defining the thrust reversing effectiveness, since excessive reverse capacity leads to the reverse thrust excessiveness and run length increase.

A certain value of optimal reverse thrust, depending on external aerodynamics of the power plant, exists for each airplane type. There should be a possibility of the engine reverse thrust control value over wide range to employ a certain engine for various types of aircraft. Thus, the reverse thrust value depends on the aircraft layout, and it is a belonging to not only the engine, but to the aircraft as well.

Reverse thrust application effectiveness on the aircraft is higher at the reverse jets fluxion optimization, than at the reverse thrust optimization. Efficiency improving of application of the thrust reverse means fulfilling the following three indicators:

• reducing the aircraft run length;
• minimizing the reverse thrust value;
• ensuring engines protectiveness from the entry of reverse jets and foreign objects, thrown-into from the runway surface by the reverse jets.

Theoretical analysis of alternative layout option application feasibility of the radiative heat exchanger (RHX) for lunar landing module (LM) was performed. Being a part of the landing module working option, the RHX consists of two parts. Both parts are installed above the unpressurized instrument bay and oriented towards the zenith by their working surfaces. Controlled removal of the excessive heat fr om the LM is being performed by the said RHX. The selected RHX size and configuration lim it the working spaces of the equipment installed on the LM, in particular, cameras, antennae, navigation instruments and manipulators. One part of the already exited RHS remains on the LM top, reducing slightly its size. The authors suggest placing the other part of the RHX near the LM side edge, instead of the solar panel, which stays at the shade for the most part of the lunar day. Placed in a like manner, the RHS vertical part will be less dependable on the temperature changes on the lunar surface, but the RHX total area increasing should compensate the expected cooling capacity losses of the LM thermal control system (TCS). The authors performed comparison of characteristics of the state-of-the-art RHX and the RHX in the configuration proposed within the framework of the presented work by the specially developed mathematical program employing computational experiment. The results confirm that application of the alternative RHX layout allows preserving the RHX integral cooling capacity, and opens new possibilities for the equipment installing at the expense of the space releasing at the LM upper part. A zone in the replaceable solar battery area can be considered as one of the options for the LM’s TCS cooling capacity increasing as a place for the third RHX placing.

Recently, the tendency towards International Regulatory Requirements on civil aircraft community noise toughening is being observed. Modern manned aerial vehicles under design should be less noisy than the aircraft being operated at present. Modern aircraft design is being performed with regard to current and prospective International regulations on the community noise. Thus, the urgency of the acoustic design issue provision in the framework of the civil aircraft lifetime is beyond any doubt.

At the same time, information on what works should be performed at various stages of the new light propeller-driven airplane creation to ensure its successful certification on the community noise and competitiveness at the world market is not presented in published works. The purpose of the presented work consists in concept forming of light propeller-driven airplane design in the framework of the product lifecycle, as well as analysis of the EASA (European Aviation Safety Agency) certification test database to determine requirements to the aircraft being designed and the effect of various factors on certification noise levels

The article demonstrates the role and place of aero-acoustic studies in the new aircraft design. Based on the EASA acoustic certification test database analysis, the article revealed that the value of noise level margin, average for all light propeller-driven airplanes, being certified according to the clause 10.4b of the ICAO Standard, was 6 dBA. The impact of blades number and propeller diameter, as well as apparent power of the power plant and presence of exhaust noise silencers of the internal combustion engine on the airplanes community noise was considered.

The presented structure of works in the field of aero-acoustics while the a light propeller-driven aircrafts design can be employed in the design of propeller-driven unmanned aerial vehicles of an airplane type as well. Requirements to the unmanned aerial vehicles should additionally account for the degree of its audibility and acoustic signature, and flight tests in this case will be preliminary (developmental) test.

Severe requirements on energy and operation parameters are placed to the gas turbines’ air-gas channels designing.

Velocities distribution along the length of the interblade channel affects significantly the working body heat transfer to the structural elements, and velocity and pressure distribution profiles affect, in the first place, the temperature boundary layer profile distribution. It is essential to account for the specifics of the flow in the inter-blade channel, which represents a radial channel. Convoluted, non-closed lines of the flow with transverse pressure gradient, which significantly affect the slope of the flow bottom lines, and, correspondingly, the temperature boundary layer formation and transformation, are being realized in this radial channel.

Joint solution of the momentum and energy equations of the spatial boundary layer for the considered radial cavities of the inter-blade channel is necessary, which represents up-to-date scientific and engineering problem.

In [1, 2-4] the authors proposed analytical approach to hydrodynamic and thermal parameters determining in gas turbines’ rotation cavities with closed circular lines and transverse pressure gradient. However, the flow line is non-closed in the interchannel cavities, and solution of dynamics and energy equations is being significantly complicated.

The article considered the analytical approach to integrating momentum equations of the dynamic and spatial boundary layer for the flow-around surfaces of the curvilinear shape in the natural curvilinear system of coordinates with the presence of the transversal pressure gradient. The initial system of differential equations for the dynamic spatial boundary layer was integrated on the boundary layer thickness. As the result, a system of momentum equations in projections to the directions of natural coordinates was obtained.

The system of equations is presented in a more General form, in contrast to the already known solutions of G.Yu. Stepanov [6] and S.N. Shkarbul [7, 8], performed with account for the flow characteristics in the inter-blade channel of an axial turbine and along the cover disk of the impeller of a centrifugal pump, respectively. The suggested notation of the equation allows integrating in the case of the non-potential external flow over the surface of an arbitrary shape.

To solve the problem of the surface flow-around with account for the heat exchange, the joint solution of the obtained momentum equations and integral relation of energy of the temperature spatial boundary layer written in the natural curvilinear system of coordinates [5].

The resulting equations represent the parabolic type equations and require the finite-difference schemes application to solve them. To verify the obtained results, numerical studies of equations for the radial sector were performed.

Theoretical and experimental studies of the flow were performed in the radial sector (without accounting for the heat exchange) in the range of radii of R_max = 0.169 m and R_min = 0.031 m, at the flow angle of rotation from 0 to 90°. The flow velocity at the maximum radius varied within 5 ... 50 m/s, which corresponded to a change in the Reynolds number of Re_U = 5.6•10⁴...5.6•10⁵.

Computational results are in satisfactory agreement with the results of these current lines visualization for the flow in the rectangular channel with cylindrical side walls along the circumferential guides.

The primary trend in effectiveness improving of gas turbine engines consists in coordinated increase of the working process parameters, such as turbine inlet temperature (TIT) and overall pressure ratio (OPR), bypass ratio (BPR) together with efficiency increasing of engine subassemblies. Alongside with that, the requirements on the engine reliability and life enhancement are being put forward.

Ensuring the required engine life at high gas temperatures prior to the turbine is possible only by turbine blades and vanes cooling, or switching to the blades materials, which do not require cooling, such as ceramics. The turbine cooling strongly affects the engine efficiency, comparable to the turbine aerodynamic characteristics, and should be accounted for while the gas turbine engine working process optimization.

The turbine blades’ design and materials permanent improvement leads to decreasing the air flow volume required for the turbines cooling. Thus, the experimental and theoretical data on the aircraft gas turbine engine turbines cooling require regular analysis and generalization.

One of the first models for predicting the required air flow rate for cooling was developed by Holland and Thake in 1980. Ever since these models are permanently developing and become more and more detailed.

It is well-known that the increased air flow rate for turbines cooling always entails the specific fuel consumption increase and the engine specific thrust (power) decrease. The engine specific parameters exert determinative affect the engine efficiency figures and, hence, its parameters optimization criteria at the conceptual design stage.

In this respect, the necessity to analyze and generalize the well-known dependencies of relative air flow rate on the turbine cooling aroused.

As consequence of the performed studies, the published theoretical and experimental data on the aviation gas turbine engines’ turbines cooling was analyzed. The generalized graphical dependencies allowed obtaining the models, on which basis the algorithms for determining the required air flow rate of the aviation gas turbine engines’ turbines cooling dependence on the gas temperature prior to the turbine. These dependencies can be employed while various tasks solving at the engine conceptual design stage. Particularly, the universal model, allowing determine the required air flow rate for cooling depending on the cooling depth in the wide range of gas temperatures prior to the turbine, ensuring goal functions unimodelity while solving optimization problems.

The studies continuation will consist in developing more accurate models of the aviation gas turbine engines’ turbines being cooled for conceptual design stage, in particular by accounting for the new structural solutions.

The article presents an overview and current development status at the EDB Fakel of prospective PlaS-10 and PlaS-10S very low-power plasma thrusters to be applied as a part of small spacecraft.

The study of the world technical level of plasma thruster development was performed. General requirements defining competiveness and high commercialization potential of the thrusters, being developed at the EDB Fakel on the world space market were set forth. The article recounts a brief chronology of the design stages, demonstrates experimental results of the thruster laboratory prototype testing, and recounts further tasks to be fulfilled on this project.

Perspective spaceflight tasks require from small spacecraft an autonomous execution of orbit maneuvers both in the near-Earth and in interplanetary space, for which a low power propulsion system, capable of functioning under conditions of the small spacecraft onboard power supply deficit (up to 100 W) is necessary. The super low power plasma thrusters can fill the empty niche [1] of the small spacecraft movement control systems, and provide the small spacecraft of potential customer with high values of the total thrust impulse for orbital maneuvers performing.

To secure the EDB Fakel leading position at the small spacecraft world market, scientific and research works on developing PlaS-10 and PlaS-10S competitive plasma thrusters of very low-power and enhanced thrust efficiency, based on brand new technical solutions, were initiated. PlaS-10 and PlaS-10S thrusters are the result of the previously developed PlaS-type thrusters concept adaptation at EDB Fakel for very low-power applications [2]. While the PlaS-10 and PlaS-10S thrusters developing the primary efforts are aimed at ensuring the key parameters of these products such as a very low discharge power and high thrust efficiency. The standard size type of the products being developed is the mean diameter of their discharge chambers, which is equal to 10 mm. The PlaS-10 thruster is based on an inner cylindrical anode, and contains a low flow rate hollow cathode-compensator previously developed by EDB Fakel, characterized by relatively high (as applied to a small spacecraft) energetic and mass and size parameters. With the purpose to further improving integral and mass and size parameters of the product, an option of the PlaS-10S structure, employing newly developed thermo-emission cathode-compensator with directly heated filament emitter, requiring less electric power for its functioning, was developed. Besides, the external cylindrical anode was implemented to determine experimentally the best anode configuration in the PlaS-10S thruster.

The small spacecraft of the nearest future based on PlaS-10 and PlaS-10S super low power plasma thrusters will be able to accomplish all types of potential flight tasks, requiring high values of the total thrust impulse available onboard a small spacecraft. These tasks may range from maintaining relative position of a small spacecraft as a part of strict formation of low-orbit multi-satellite systems to accomplishing the exploratory small spacecraft flights into deep space. The high potential of modernization herewith, encumbered into the thruster structure at the stage of development, defines the possibility of thrusters’ thrust and energy characteristics enhancing with the course of time, which is the key factor capable of ensuring the high level of the PlaS-10 and PlaS-10S competiveness supporting in the future.

Fuel burning in the combustion chamber is being accompanied by toxic substances formation. Carbon oxides, having deleterious effect on both human and environment, represent a particular danger among them. In this regard, the article solves an actual problem of determining the optimal combustion chamber gaseous fuel supply to ensure low carbon oxide emission.

The article presents the experimental solution of the emission reduction of the deleterious and polluting substances at the combustion chamber outlet, and the test bench equipment description. It considers three options of burners, differing by the nozzle extension design. The atomizer geometry remains unchanged. The article presents the results of firing test of the three burners with different nozzle extensions. The flame structure comparison of the three burners was performed. Parameters estimation of the burners was carried out, and the burner with minimum value of nitrogen oxide and carbon oxide in the combustion products samples was selected. Temperature field at the outlet of the combustion chamber bay with three types of burners was studied. The article presents the results of deleterious and polluting substances emissions measurements from the bay with the burners of various design. Combustion efficiency was determined as well.

Inferences on the burner option most acceptable for implementation with the engine were drawn by the results of the performed work.

The article presents mathematical model of the liquid propellant rocket engine (LPRE) flow regulator and the study of its static characteristics, such as fuel component consumption dependence on the pressure difference, and dynamic characteristics, such as regulator amplitude-frequency response. The study was performed by the developed mathematical model, which unlike the well-known domestic and foreign counterparts ensures the most complete description of the fuel consumption regulation processes. It demonstrates that dynamic characteristics in technical systems are being determined by the areas of its movable part (slide-valve) and differential orifices.

The liquid flow regulator is one of the main units of any LPRE. These regulators are designate for maintaining the fuel components consumption keeping with the specified accuracy, or its varying according to the certain law under conditions of internal and external disturbing factors varying.

They are being employed in the modern multimode engines such as RD-253, RD-120, RD-170, RD-180, SSME, RL-10 as actuating elements.

The flow regulators employed in the LPRE are being separated into the two groups: direct- and indirect-acting regulators. The direct-acting regulators found wide application in modern LPRE. The direct-acting regulators are being applied as a rule at a flow rate m*g ≤0.2 kg/s, though they can be employed at greater flow rates, if high performance ensuring is necessary.

A feature of all flow regulators is their ability to control the flow rate and maintain the flow rate only at relatively slow changes of control and disturbing impacts in time.

The article presents a system of equations, describing working processes at the fuel components regulator normal functioning. Mathematical model of the improved direct-acting thrust regulator design for the LPRE with oxidizing gaz afterburning, allowing substantially increase effectiveness of automated for engine control and diagnostics systems. As the result of modelling, the dependencies of flow rate through the regulator on the angular position of the actuator and pressure difference at the regulator were obtained.

Recommendations on flow rate regulations modernization for the engines of the RD-170 family were given based on the obtained results. The results can be used while flow regulators designing and their state diagnostics while testing.

The article presents a description of the unsteady turbulent separated incompressible 3D flows of products in solid propellant rocket engines by the Reynolds-averaged Navier-Stokes equations intended for incompressible fluids. It is shown herewith that the differential one-parameter model, proposed by Spalart-Allmaras, as well as the SARC and SALSA models can be employed to perform turbulence simulation of the 3D flow of products in solid propellant rocket engine. These models can be applied for the averaged Navier-Stokes equations closing and simulating the unsteady turbulent separated incompressible 3D product flows in solid propellant rocket motors.

It is necessary to perform calculation of the processes, occurring inside the solid propellant rocket engine, with physical and technical characteristics determining of this engine, associated with the thrust, fuel consumption combustion chamber operation parameters etc., based on the numerical modelling methods application, in the course of the solid propellant engines development and design. Mathematical models were proposed herewith for describing transients with igniter actuation; with warming-up, further ignition and solid propellant burning transients. They describe as well the non-stationary transients from the simple to heterogenic flow, originating due to the movement of air and solid propellant products formed in the combustion chamber of the rocket engine; and those associated with the process of the solid propellant rocket engine plug movement.

Of all types of rocket engines employed as propulsion systems for various purpose aircraft, solid propellant rocket engines, along with the liquid propellant rocket engines, are the most widespread ones. This fact is being confirmed by the widespread application of solid fuel rocket engines as cruising propulsion systems in the objects from operational tactical missiles to launch vehicles of various classes; the solid fuel rocket engines application for braking wasted stages of launch vehicles; as well as for the spacecraft extra acceleration while transitions from transfer orbits to the required final orbits. Besides, the propulsion systems based on solid propellant rocket engines have found wide application as boosters with the purpose of increasing the energy capabilities of launch vehicles and expand the range of target tasks they are solving. The foregoing determines the relevance of the research. This research associates with the modern methodological support development, which includes the problems formulation; creation of mathematical models, algorithms and programs for solving the problems of the initial stage of the objects designing and, in particular, creation of a method for calculating the 3D flow of combustion products in solid fuel rocket engines of promising aircraft devices.

The presented article deals with the quantitative assessment of the flight experiment informativity content in the course of flight tests, and the issues of decision-making by the results of parachute systems (PS) tests. It states the main goal and objectives of the PS flight tests. High-grade and effective solution of the main tasks of the PS flight tests necessarily requires high level of the flight experiment results informativity. The article considers in detail the flight experiment informativity as the local criterion for the experiment effectiveness evaluating. The concept of informativity includes the quantity and quality of results; informative content sufficient for making a competent (correct) decision when determining the purpose of further research; the methodology correctness for organizing (preparing and conducting) a flight experiment. The authors formulated the concept of informative content of the experiment. The article considers a number of methods for various-level evaluation of the informative content of the flight experiment results. In the most simplest case, i.e. at the lowest level of the hierarchy, the informative content of the experiment is being quantified by a coefficient equal to the ratio of the volume of information obtained in the experiment to the planned volume. The next higher level in the hierarchical structure of the informative content of the flight experiment is associated with probabilistic approach to the problem. The informative content of the experiment can also be quantified by the probability of obtaining an unequivocal answer to the question posed by the experimenter, which allows making the only correct decision on further research trends selection. The next much higher level in the hierarchy structure of the flight experiment information content is associated with the quantitative assessment of the information by the Hartley, Shannon formulas as is being done in information theory and coding, as without regard and with account for the jamming impact. Obtaining sufficient amount of reliable information from the flight experiment allows directly proceed to the next important stage, namely making a decision on the results of the PS flight tests.

The article presents the optimal variant of a decision-making process typical block diagram based on the results of informative content experiments. The flight experiment results of the PS flight tests is of fundamental importance for the decision-making processes on the further research trends, since both testing terms and their cost significantly depend on it.

The International Space Station Russian Segment (the ISS RS) development along with the increasing number of scientific and applied research and experiments performed by cosmonauts onboard the space station actualize the issue of ensuring high-quality training for the scientific program implementation. Visual-instrumental observations of the Earth from space (VIOs) are one of the most informative methods of Earth’s remote probing, employed in manned space exploration. They are intended for observing natural and anthropogenic objects, phenomena occurring in outer space, atmosphere, on ocean and land surface (cyclones formation and typhoons origination, volcanic activity, thunderstorms, forest fires, bio-productive areas in the oceans, and processes in the upper atmosphere).

The experience of domestic cosmonauts training for the VIOs performing is indicative of the importance of cosmonauts training process at all of its stages. Cosmonauts training in this line should represent educational and training process oriented on cosmonauts’ mastering theoretical basics of experimental research on topical problems of earth sciences, studying physiographic specifics of territories and acquiring necessary skills and abilities on searching and identifying the objects under study, as well as practical application of the onboard equipment for remote geosystems’ probing.

Selection of research trends onboard the ISS is based on the basic principles of the Federal Space Program of Russia, foreseeing studying of the Earth surface, Moon studying and exploration, observing various processes and phenomena on both Earth and Lunar surface. This puts forward the requirements to cosmonauts’ training on this trend of their professional activities at all stages of their training for the space flight. These requirements consist, in the first place, in the necessity for the theoretical training, as well as conducting practicum and training using informational resources of specialized simulators that simulate visual situation under conditions of the ISS flight, and flights for aero-visual observations of test sections of land and sea.

Creation of simulator for cosmonauts’ training to perform VIO based on employing digital Earth surface model allows enhancing effectiveness and quality of cosmonauts training to perform the spaceflight onboard the ISS. In the course of design and development of the simulator stand for cosmonauts’ training to perform VIO a comprehensive analysis of specific features and conditions for the VIO performing, characteristics of the scientific equipment in use, as well as available experience of cosmonauts’ training on prospective space programs, including flights to the Moon and near-Lunar space, was performed.

Currently, aerial refueling is being employed to increase the aircraft flight range and duration. Refueling an aircraft in manual actuation through all control channels is one of the most difficult and stressful modes of piloting for a pilot, and requires high qualification and long training.

This is being especially complicated by negative factors such as:

The tanker aircraft trail line impact on the aircraft being fueled;
The airstream turbulence, etc. Automation allows increasing the probability of successful contact compared to manual actuation (for example, about twofold for a light aircraft). One of the trends unburdening a pilot, and simplifying this process may be automation of the speed control channel.

The article considers the speed control algorithm at all stages of the aircraft aerial refueling mode:

The aircraft’s approach to the tanker;
Directly the process of a drogue and a cone contacting;
Taking working position for the fuel pumping;
Separation from the tanker after refueling completion;
Re-entry for contacting when the hose and cone contact performing failed.

The purpose of the article consists in the speed control algorithm development at all stages of the aircraft aerial refueling mode.

The main objectives of the article are as follows:

Increasing the flight duration;
Reducing the burden on the pilot, and lowering the requirements for his qualification;
Increasing the probability of successful aircraft refueling from the first approach;
Refueling performing in conditions of air-turbulence;
Improving flight safety.

Speed control automation while aerial refueling should be performed through auto-throttle. Its algorithm should include the law of the specified relative speed of the aircraft and tanker, based on their mutual position. To be more exact, it means the mutual position of drogue and cone, as well as drogue and a certain element on the trailing edge in the area of the unit installation after the contact and while fuel pumping.

While the algorithm developing, classical approaches to flying vehiles’ control systems design, mathematical modelling methods and simulation on the flight simulator were employed.

Simulation results on the flight simulator revealed the operability of the algorithm ensuring speed control of the aircraft being fueled relative to the tanker.

A system of technical vision, operating in real-time scale onboard the aircraft being fuelled, can be employed to ensure the aircraft refueling autonomy.

The proposed algorithm for the auto-throttle signal generating can be considered hereafter as an element of ensuring automated aerial refueling of the aircraft.

The subject of this research is ballistic schemes optimization for the spacecraft with solar electric propulsion system. The article considers the problem of the initial conditions search for a spacecraft launch, at which the total time of its staying in the shadow at the insertion phase would be minimal.

The total duration of shadow sections during interorbital flight will depend on the relative position of the Sun and the spacecraft’s orbital plane. To solve the problem of the initial launch conditions selection, the dependence of the shadow section duration on the set of ballistic parameters, such as the ascending node longitude, the perigee argument, and the launch date of the flight, is being considered.

A ballistic scheme for leading out, at which elliptica transfer orbit forming is being performed by the upper stage of the rocket-carrier is selected, and a spacecraft finishing up to the working orbit is being performed by its own electric propulsion unit.

The article proposes a model for duration computing of the orbit shadow sections. Equations of motion in osculating elements are assumed as a mathematical model of the spacecraft controlled motion under the impact of the electric propulsion. An algorithm for solving the problem of optimal initial flight conditions search has been developed. The total duration of a spacecraft with the solar propulsion unit staying in the Earth shadow along the whole trajectory of the multi-turn flight was accepted as an optimality criterion. The following parameters, namely the launch date — perigee argument — the ascending node longitude, were selected as the optimized parameters of the elliptical orbit.

Computations of the spacecraft flight trajectories from high-elliptical orbit to the geostationary one for three initial orbit inclinations, performed with variation of the parameters being optimized, were carried out. The spacecraft launch windows and corresponding initial conditions of the orbit, rational in terms of the flight duration reduction, were found based on the simulation results. Analysis of the simulation results array revealed that launching date selection did not affect significantly the flight time at optimal combinations of the perigee argument and the ascending node longitude, and the time difference for the flights in 2020 lies within the limits of 1%.

The combination of the initial ascending node longitude and the perigee argument has a much greater impact than the launch date selection. The worst combinations of these parameters may increase the maneuver time by 12% of the minimum value, which gives their optimization the highest priority. Thus, the flight initial conditions selecting is an important problem of the low-thrust interorbital flights optimizing.

It may be noted as well that while flights with three initial values of the orbital inclinations simulating, a tendency for the increase in the relative difference in flight time between the optimal and non-optimal initial flight conditions with a decrease in the initial orbit inclination was found. As the result, the orbits with lower initial inclinations are more demanding in the initial parameters selection.

The article demonstrates the possibility of the approximate optimal control method and the «NEOS» software application for the flight tasks with account for shadow sections, including those with multiple simulation.

The obtained results can be applied for evaluating the design ballistic parameters of a spacecraft with electric propulsion unit flight, as well as determining the optimal initial launch conditions.

Escalating requirements to the new products characteristics are associated with improvements in design, which in its turn leads to the need of new materials and technologies developing for parts manufacturing. The present-day materials allow substantial improvement of the products functional properties and required service life, but very often due to drastic increase in their cost. Thus, their properties would be employed most effectively while developing material-saving technologies for their preparation and processing. Selective laser melting (SLM) technology is one of the most effective technologies for metal products manufacturing without machining. A layer-by-layer application of metal powder of the specified grain-size composition on the forming-up platform and laser hatching of the current section according to the pre-developed CAD-model are performed while the installation operation. The process is being cyclically repeated until completion of the part forming process. To prevent oxidation, the synthesis process is performed in the sealed chamber in the inert gas medium.

The 3D-printing technology has a defect such as the structure porosity and unattainability of the required level of mechanical and operational properties. Anisotropy of properties is being observed in the products manufactured by the SLM technology. The key factor affecting the properties of the synthesized material is the presence of porosity, cracks and unmelted granules. With this regard, additive technologies application for the critical parts manufacturing is being complicated, and their full-scale implementation in high-tech industries is being retarded.

While products shaping the whole layer (current section) of the part is being divided into separate square-shaped fragments called «islets», each of which is fused by the laser. The fragments are being fused according to a predetermined algorithm, developed in such a way as to localize the internal stresses of the metal in a small area, which allows obtaining homogeneous and dense structure with minimum porosity. Argon was used as an inert medium. From the viewpoint of the process parameters optimization, it is necessary to achieve density of the part being synthesized close to 100% with maximum printing speed. Pores of the alloys obtained by the synthesis employing the SLM technology are of different nature, such as shrinkage pores formed due to incomplete cavities filling with liquid metal; gas, spherical pores, caused by the capture of gas in the bath melt at the excessive overmelting; as well as non-melted areas formed due to lack of energy for their fusion. The unmelted areas may have the shape of the structure discontinuities due to the laser power deficiency and irregular structural formations due to excessive scanning speed. The presence of large pores in the material herewith leads to degradation of the material strength characteristics.

The alloys were being subjected to the cold isostatic pressing on the specially developed installation for the porosity reduction.

The article presents the results of the studies of the impact on the size, pores number and alloys structure of the cold isostatic pressing of the samples fabricated from the heat-resistant alloys, obtained by the selective laser melting technique of metal powders. It demonstrates that cold isostatic pressing application with the SLM-alloys allows substantial (about twice) reduction in pores size and number. The effect of the 316L SLM-alloy hardening manifesting in the hardness increase of the surface layer at the room temperature was revealed.

The article addresses the issue of determining the nominal value of roughness and its dispersion as the result of the outer surface of the «Rotor shaft of a gas turbine engine» part turning, being an element of the rotor part of an aircraft gas turbine engine.

The article describes a technique for establishing interrelation between the parameters of technological environments with quality indicators obtained as the result of processing in these technological environments. The technique is illustrated by the example roughness evaluating of the part outer surface as the result of turning.

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

The introduction performs the analysis of literature related to the problem of establishing interrelations between the technological environments parameters and operational and technical characteristics of products. The rationale for the need to establish such dependencies is being presented.

The main part provides a technique for assessing the value and dispersion of parts’ quality indicators depending on the values of the of technological environments parameters. Based on the results of this evaluation, a conclusion is being made on the probability of finding the value of the considered quality indicator within the specified limits. The technique is being illustrated by the example of roughness forming on the outer surface of the «Rotor shaft of a gas turbine engine» part while fine turning. The required roughness value is no more than Ra0.4. Based on computational results, probability evaluation of obtaining roughness of no more than Ra0.4 is being performed for the two different groups of technological environment parameters. The probability was 0.55 for the option A, and 0.71 for the option B.

It is noted in the conclusions that despite the fact that the probability value is greater for the option B than for the option A, in some cases the option A will be preferable, since the roughness values obtained while processing in a technological environment with these parameter values are of lower dispersion, i.e. more stable. The article indicates that the obtained roughness values will affect the operational and technical characteristics of the product, including reliability.

The studies on estimation of the external ultrasonic field impact on the surface quality of the obtained small diameter orifices in corrosion-resistant steels and electric discharge machining productivity were performed within the framework of the presented work.

The purpose of the performed studies consists in determining quantitative characteristic of the roughness indicator when small-diameter orifices processing by electrical discharge machining with ultrasonic oscillation superposition the part under treatment or EDM tool.

The combined machining method is based on superposition of thermal action of the electric current impulses, fed continuously to the section of the workpiece being machined, with forced impact of ultrasonic oscillations for erosion products evacuation from the inter-electrode gap.

The 12Х18Н10Т-grade austenitic stainless steel was selected as the material to be machined for experimental studies for accuracy increasing,while the small-diameter orifices through-piercing, the presented work employs the guide alignment bushing, made of wearproof dielectric material, trough which the electrode-tool is delivered and fixed.

Based on preliminary studies on the process fluid selection, preference was given to the IonoPlus IME-MH synthetic dielectric fluid for axial drilling machines, which is applied for finishing and semifinishing. Process fluid is forcefully fed through the guide sleeve.

Prior to the experiments commence, a study was performed to select the ultrasonic field sources. Piezoceramic and magnetostrictive ultrasonic field sources were being considered. Based on the previous experiments, a magnetostrictive transducer was selected, which has a wider range of oscillations amplitude adjustment.

The machining time was recorded with a calibrated stopwatch; and the tool wear was recorded by touching the surface of the part before and after machining.

The article considers methods and technological solutions on the effective small-size orifices machining aimed at quality enhancement of the machined surface and electrical discharge technology productivity.

In the process of experimental studies, various options for the ultrasonic head installing and the electrolyte supply direction to the treatment zone were applied

The modes and schemes for the parts samples treatment were obtained based on the materials selection for the electrode-tool and operation modes of electrical discharge and ultrasonic equipment.

Experimental results allow comparing electrical discharge machining methods by technological indicators of machining time and the obtained surface quality. Thereby, they give notion on ultrasonic oscillations impact on the productivity, accuracy and quality of electro-erosion piercing of the small-size diameter orifices.

The experimental studies revealed that the high-frequency oscillations transmitting to the electrodetool lead to productivity increasing due to h short-circuit prevention between the EDM-tool and part being processes.

Graphical interpretations of the obtained numerical values allow quantifying the relationship between the processing time and the EDM tool wear, with account for various schemes of the ultrasonic application while piercing orifices in the samples of plates and nozzles.

The studies of the orifices’ treated surfaces roughness, obtained by the electrical discharge machining with the ultrasonic oscillations superposition and working fluid flowing into the processed zone were performed.

The superposition of ultrasonic oscillations to the EDM tool facilitates obtaining a low roughness in comparison with the roughness obtained by traditional EDM machining by 15-25% due to a decrease in the number of burns and short-circuits.

In recent years, one of the prospective and highly competitive trends in the field of anti-icing materials creation is the development of passive ice-phobic coatings, oriented not only at the ice accumulation reduction on the surface while contacting with the hitting atmospheric water droplets, but being able to completely suppress ice formation under certain weather conditions.

The ice-phobic coating should demonstrate the following properties to achieve stable anti-icing characteristics:

Supercooled water accumulation reduction;
Low adhesion of liquid water or any form of solid water, including various kinds of ice, frost and snow, to the surface of the ice-phobic material;
Long delay time of the supercooled water droplets crystallization on the surface of the material, and finally
Low heat transfer between the droplet and ice-phobic material, which decreases the probability of the water droplet supercooling while its impingement with the cool surface.

For application in aviation industry, the ice-phobic coating should display firmness to the extended abrasive loadings and cyclic temperature difference.

A TSAGI-831 aviation profile and a flat plate were selected as tested aircraft aerodynamic elements. Both samples were made of the D16 aluminum. To impart water- and ice-repellent properties on the material surface of the samples being tested, super-hydrophobic coatings were being created. The method for super-hydrophobic cooatings processing on the aluminum alloys was developed at the RAS Institute of Physical Chemistry.

The tests on checking the effectiveness of the ice forming prevention and ice removal were performed on the EU-1 FSUE «TSAGI» artificial icing test bench under artificial icing conditions by the Appendix C, AP-25.

The tests results confirm their high anti-icing ability: the time before appearance of the first ice deposits on the surface of the super-hydrophobic coating after the aerosol flow starting was four minutes. Reduced ice accumulation and spontaneous ice removal phenomenon form the super-hydrophobic coatings surface were registered. Ice accumulation was being observed on reference sample without coating right after the flow commencing. All above said indicates the high potential of the developed super-hydrophobic coatings for the aircraft aerodynamic surfaces icing counteracting.

Commercial air transportation growth and environmental requirements toughening encourage designers of prospective aviation to develop and research innovative technical solutions and technologies to improve performance while conjoined emissions reduction. In recent years, increased attention has been paid to the study of the Distributed Electric Propulsion (DEP) application, which implementation onboard aircraft, according to researchers, will allow fuel costs cutting by more than 50% with conjoined carbon dioxide emissions reduction by approximately 50%. Many scientific and engineering problems should be solved while the aircraft with DER development. One of such problems, to which solution a great number of today’s studies is devoted, consists in ensuring high takeoff-landing performances. The presented work considers the possibility of employing combined lift force increasing power system (CLFIPS) for the wing lift force improving at the takeoff-landing modes. Evaluation of various factors impact, such as the propeller diameter and thrust; its position along the length and height relative to the airfoil chord at various angles of the flap deflection and blowout intensity on it, on the CLFIPS effectiveness. Along with the basic calculation option, the slipstream effect of the propeller on the aerodynamic characteristics of the airfoil with slotted flap, as well as with the system of circulation control by tangential blowout of the jet on the rounded rear edge of the airfoil are considered.

Computational study of the airfoils flow-around by the viscous gas flow was performed at the numbers of M = 0.13 Re = 7.2·10⁶ employing the FLUENT software based on the numerical solution of the Reynolds-averaged Navier–Stokes equations. The blow-off calculations at various values of the propeller active section diameter and its position were performed at the zero angle of attack.

Parametric studies of the high-lift airfoil flow-around were performed at various values of the propeller relative diameter, being modelled by the “active” disk, and its position relative to the airfoil. The studies confirmed the effectiveness of the combined lift force increasing system conjoining boundary layer control (BLC) system and propeller blow-off (PBO), compared to the speed circulation control by tangential blowout of the jet on the rounded rear edge of the airfoil, as well as the blow-off of the airfoil with the Fowler flap type.

It is advisable to go on with the studies on parameters optimization of the combined BLC/PBO system as well as the type and parameters development of the wing slot mechanics, which ensures effective jet deflection from the wing for the purpose of significant lift force increase.

At the preliminary design stage of the aircraft (up to the detailed design stage and performing full-scale fatigue tests of airplane glider units), it is necessary to ensure the fulfilling requirements for fatigue and survivability of composite aircraft structural components. To start with, a computational evaluation of safe life span and damages non-progression in structural elements from polymer composite materials (PCM) should be performed.

The following evaluations should be performed to this end:

1. Computational and experimental evaluation of the safe resource of elements of composite aircraft structures.

2. Computational and experimental evaluation of non-progression of the first category of damage on the elements of composite aircraft structures over the entire period of the aircraft operation (up to reaching the operating time equal to the design service life of the aircraft).

3. Computational and experimental evaluation of non-progression of the second category of damage on the elements of composite aircraft structures over the period between scheduled or targeted inspections, conducted through the certain intervals.

This article presents the basic regulatory requirements, methods and procedures for computational and experimental evaluations of the main fatigue life characteristics of composite wing panels at the outline design stage of a transport category aircraft. The example of computational and experimental evaluations of the safe resource and the frequency of inspections of the upper composite wing panel of a transport aircraft made of the AS4-PW carbon fiber laminate is presented. A number of important inferences was drawn.

The obtained results of computational and experimental evaluations of the life span characteristics of the upper composite panel of a wing from the AS4-PW carbon fiber laminate at the stage of outline design of the aircraft allow making the following conclusions:

1. The expected safe resource of the upper panel is being actually determined by the computed safe resource of the panel in the zone of impact damage of the BVID type, which the value is 6.7 times less than the calculated safe resource of the upper panel in the free holes zone.

2. The frequency of necessary inspections of the upper panel is determined, first of all, by the frequency of inspections of the panel in the impact damage zone of the VID type. The frequency of inspections is 5,300 flights and it actually determines the frequency of inspections according to the C-check maintenance form.

The obtained values of the safe resource and the frequency of inspections are within the range of real values of the life fatigue characteristics of the real aircraft, which allows concluding on the acceptability of such evaluations.

The article deals with multidisciplinary comparison of the twin-engine transport aircraft concepts with various types and layout of the power plant.

The main purpose of the study consists in the transport efficiency increasing of the wide-body aircraft. The key condition of the presented study is observance of the same operational requirements and a single level of technical excellence. All the concepts of a transport aircraft discussed in this article belong to the 16–23 tons load capacity class.

The article considered four technical concepts of a transport aircraft with two engines:

– the aircraft of traditional layout with turbofan engine (MTS-0);

– the aircraft of traditional layout with turbojet engine (MTS-1);

– the aircraft of integrated layout with turbojet engines positioned in the center wing section (MTS-2);

– the aircraft of integrated layout with turbojet engine above the stern of oval fuselage (MTS-3).

The authors performed analysis of the power plants efficiency; defined aerodynamic, weight and takeoff-landing characteristics, and perform comparison of both transport and economic efficiency of the concepts being considered.

The article showed that the aircraft with turbofan engine (MTS-0) demonstrated minimum fuel consumption, and it required minimum runway length at maximum flight range with the 20 tons load. The price and direct operating costs herewith of the aircraft with turbofan are the highest.

When performing average in the park transportation work with the 14 tons load, the integrated layout engines positioned in the center wing section (MTS-2) is being distinguished by the lowest price and operating cost value. Thus, it can be recommended for commercial application.

The article performs a comparative analysis of the known methods of the of the solar system planets exploring by automatic interplanetary stations (AMS). These are exploration by the flyby trajectories, from near-planet orbit, and planets exploration by the probes (stationary or mobile) with direct landing on the planet surface. The following conditions ensuring global planet exploration were selected as comparison criteria. They are contact studies (soil analysis, etc.); the possibility for visiting several regions of the planet; maximum routs length for detailed exploration of the planet; applicability while pioneer flights realization, and the possibility of reusable application of the one-type spacecraft for various space objects studying.

In the process of analysis, conclusion is being drawn that none of the applied methods solves scientific problems concurrently and comprehensively (on a global scale of the studied planet) and in detail (at the level of contact probes). It was proposed herewith to consider the fourth – practically unexplored method of research – by employing orbital refueling tankers (ORT) and reusable takeoff and landing complexes (RTLC). The article demonstrates the possibility of high-tech scenarios realization of scientific missions, combining both scales (such as exploration of several remote regions of the planet, or even several satellite planets near the giant planets) within the framework of a single mission, as well as contact studies (soil sampling, drilling, etc.). On the example of the flight to the giant planet system (Jupiter, Saturn, Uranus, Neptune) the author demonstrates the possibility of realizing scenario with multiple landing on the giant planet satellite, as well as with flight continuation to the next satellite of this planet, and its exploring with the same scenario. The RTLC fueling from the fueling tanker side (the ORT, besides the tanker function, performs the function of the scientific mission navigation and communication provision device), ensures the possibility of multiple contact and route studies on the planet (satellite) surface, which is yet impossible with conventional exploration techniques. The RTLC fueling from the fueling tanker side (the ORT, besides the tanker function, performs the function of the scientific mission navigation and communication provision device), ensures the possibility of multiple contact and route studies on the planet (satellite) surface, which is yet impossible with traditional exploration techniques.

At present, means of technological equipment with digital control prevail in technical objects production, which predetermines digital methods for both technical objects and technological processes representation, digital workflow and robotic production. It requires new approaches and methods for integration of designing and manufacturing. Organizational separation of technical preproduction into design and technological ones is characteristic for various branches of science-intensive mechanical engineering, including aviation and space-rocket industries. Complexity and functional completeness of the problems being solved by various automated systems separate designing, manufacturability adjustment and preproduction into separate stages of the science-intensive products’ life cycle. Primacy of design as the process of the new or being upgraded object (products, technological processes, production systems, information systems) description creation, necessary and sufficient for the object being designed realization under the specified conditions, is common to all stages. The main constraints for technical objects design are the specified quality indicators, and rational options selection criteria are both functional performance indicators and technical and economic indicators of realization at all stages of the life cycle. The «Designing» stage includes the following phases: development of technical specifications; technical proposal; draft design; technical project; working draft. Preproduction planning of aerospace enterprises includes the following stages: grouping or shop-to-shop routing of the product, ensuring manufacturability of the product design, technological processes developing, technological equipment design, material and information flows design and production system functioning adjustment. The results of each stage are being formalized in the form of project documentation. Design and technological models for the same design objects differ not only by the form of representation, but by the volume of the features and parameters being described as well, employed for the design and process design systems developing, which significantly complicates their integration. It is recommended to employ the following system-wide principles, ensuring information support of the objects for designing and technological design integration: the principle of inclusion; the principle of completeness; the principle of information unity; the principle of compatibility and the principle of invariance while automated systems creation and development. With account for the requirements on consistency, independence and completeness of the parallel design system based on representations and interpretations of the design automation methodology in the subject areas of designing and technological design the basic functions of the design systems were formulated.

The structure of the design process models were determined with separation of models of various objects, being formed and interacted in the design process, as well as the structural-parametric modeling process were developed.

It was recommended to apply a unified mathematical description of science-intensive products, technological systems and technological processes in designing and technological design to ensure effective integration of automated systems for all stages of the life cycle employing the PDM and PLM systems.

Most of the existing launch vehicles are being equipped with booster blocks, performing sequential spacecraft deployment into a specified orbit. However, a scheme with individual spacecraft leading-out by the last, modular, launch vehicle stage is possible as well.

As experience shows, when creating a launch vehicle with solid propellant rocket engines, borrowing of a number of elements is the case.

The problem statement can be formulated as follows: find such a vector of the basic design parameters so that the launch vehicle launch mass will be minimal, and a number of restrictions herewith, namely by the payload mass, size, the borrowed elements parameters will be met.

The task of a launch vehicle with modular stage III booster block (BB III) designing is:

– multi-criteria;

– multi-parametric.

The method of constraints is used to solve a multi-criteria problem.

The problem feature consists in the fact that while searching for the rational design solution, concurrently changes the vector of the determining parameters (mass and geometric ratios coefficients, which values depend on the design solutions for the BB III modules). Various approaches to the problem solution are possible.

The article presents a two-level coordinated optimization method.

When implementing the two-level coordinated optimization method, the upper-level model is being refined according to the lower-level data, which allows increasing the calculations accuracy without resorting to the excessive expansion of design models. The control parameters (design parameters) at the (i + 1)- th level are being selected so as to ensure a more detailed description of the object compared with the i-th level of detailing, the vectors of the parameters, being selected at different levels, at that should not contain the same elements. The great attention herewith is paid to the agreement assessing of the design solutions at both i-th and (i + 1)-th levels of the development management.

A study on the model example was performed for the launch vehicle with a solid propellant engine of bout 50 tons launch mass, with every module weight of 250 kg.

The presented graphs demonstrate the process of design solutions coordination at the i-th and (i + 1)- th levels of development management.

The two-level matched optimization method allows finding a rational solution without significant expansion of the design models.

Air fleet developing prospects all over the world are closely associated with creation of highly efficient methods for maintaining the aircraft airworthiness. One of the tasks, being solved while such methods developing, is cost reduction during the aircraft operation. A reliable and rather effective periodic inspections system can be replaced by the structure status monitoring, which consists in continuous data collection and analysis of airframe integrity throughout the aircraft entire life span.

Status monitoring is performed by the onboard system, which basic elements are recording and analyzing unit, and sensors. The sensors are fixing the structure response at its integrity violation during operation. The damages detection effectiveness and possibility of reliable determination of the operation conditions depends in many ways on the algorithms realization, in which accordance the analyzing unit operates.

Currently, a large number of sensors types, based on various physical principles, have been developed. Strain gauges, which change of readings may indicate the presence of the structure damage, were widely employed while the experiment and approbation of the onboard monitoring systems.

The article proposes a method for determining the sensors installation scheme while fatigue damage detecting in the fuselage joints with account for the local nature of changes in the stress-strain state near the cracks and the allowable size of cracks that can be considered safe under certain conditions. The multi-site damage parameters, at which the residual strength of the joints does not decrease below the permissible level, were selected by studying the fractures of the joint samples by fractography. The optimal sensors installation scheme determining was performed based on the analysis of relation between of the measurement system readings and damages. This relation is presented herewith in the form of the neural network approximation.

The neural network training to obtain the necessary relation was performed based on the results of local deformations determining by the finite element method for various options of the of cracks location in the critical section of the joint. Various factors affecting strain measurements were accounted for while determining the places of sensors installation.

The article presents the result of the developed methodology application for the optimal sensors installation scheme determining in one of the types of longitudinal fuselage joints when detecting multi-point fatigue cracks during fatigue tests.

A feature of polymer fiber-reinforced composites (PFRC) destruction is multi-focal point damages formation of microstructure under external impacts, their growth and coalescence, resulting in macro-damages formation and sudden destruction of a product. One of the factors impeding creation of the multi-level prognostic models of the PFRC destruction consists in limitation in non-destructive means, allowing study mechanisms of their internal structure damaging from micro- to macro-level.

A combination of two non-destructive acoustic methods was employed to study the multilevel damage

of a thick laminated carbon fiber-reinforced plastic (CFRP) with the [0°/ 90°]_4S stacking under loading on uniaxial stretching and three-point bending. The acoustic emission (AE) method allows continuous monitoring of the internal structure integrity of the tested material and loading parameters recording while the damage initiation and growth. Multilevel visualization of the material internal structure, acoustic microscopy (AM), allows identification and characterization of its damages from micro- to macro-level.

The samples being tested were cut out of a panel of the 4.32 mm thickness, fabricated by the prepreg of a thick laminated carbon fiber-reinforced plastic (CFRP) with the [0°/ 90°]_4S stacking under loading on uniaxial stretching and three-point bending. The acoustic emission (AE) method allows continuous monitoring of the internal structure integrity of the tested material and loading parameters recording while the damage initiation and growth. Multilevel visualization of the material internal structure, acoustic microscopy (AM), allows identification and characterization of its damages from micro- to macro-level.

The samples being tested were cut out of a panel of the 4.32 mm thickness, fabricated by the prepreg the harbingers of the full destruction of the material, namely:

– zones with high (critical) density of transverse matrix cracks in [90°] layers,

– the adhesion weakening/damaging along the «fiber-matrix» interfaces in [0°] layers,

– local fibers fractures.

Strike unmanned aerial vehicle (UAV) more than once proved their efficiency while performing special missions in various local conflicts. For this reason, Military Forces of large foreign countries pass the UAVs of this kind into service already for several years. In Russian Federation, similar UAVs are only at the stage of development. The problem of the power plant creating for any kind of aerial vehicle at this stage is one of the basic, and the problem of developing aviation engine for it relates to the most complex ones.

The presented work set and solved the task on determining optimal parameters of the operating procedure, control program for the bypass turbofan engine (TFE) and the power plant dimensionality, ensuring the best values of the selected efficiency criteria of “Scat” type strike UAV, while its performing characteristic mission tasks with account for its aerodynamic, mass-volume and flight performances.

To conduct this study the authors developed a technique, in which «Aircraft and Engine» instrumental-software complex and IOSO_NM 2.0 optimization pack are the basic program tools.

Parameters matching based on the statistical data on the power plant, aerial vehicle and their aggregate while the mission task modelling was performed for the purpose of forming the “base option” of the objet under study, relative to which the effectiveness of the appearance options being formed was estimated. Aviation engine RD-33 as a power plant engine prototype, and the “Skat” strike UAV breadboard model as an airframe were selected, while mission program was trained based on the typical combat assignments for the fighters.

Range parameters for the two mission programs, characterizing its functional purpose were accepted as the effectiveness criteria of the UAV under study.

Parametric studies of the “base option” were performed to determine regularities of the effect of the TFE and power plant working process parameters, the UAV airframe and parameters of their matching on both altitude-velocity and throttle performance of the engine, as well as on the UAV’s integral parameters and selected efficiency criteria. Analysis of the obtained results was performed, and boundary values of the parameters, at which physical existence of the studied object was observed, which was necessary for the varied parameters values range selection, were revealed.

As the result of the optimization problem solving, the UAV and its power plant parameters were determined from the condition of achieving the flight ranges maximum by the two formed mission programs while fulfilling all design specifications, imposed on the strike UAV under study. The flight range according to the first program herewith increased by 13-20% compared to the “base” variant, and 9-10% according to the secondo one.

The authors plan hereafter to perform the power plant efficiency estimation of “Skat” type strike UAV comparison with the other engine schemes.

The practical value of the presented work, consisting in the fact that its results may be employed by the scientific and design organizations preoccupied with prospective UAV and its power plant development, in ordering Air Force and industry organizations while requirements substantiating to the new samples of aviation engineering, as well as aviationand engineering universities while educational process improving.

The article presents calculation technique, which allows defining temperature fields in the machining zone while workpices shaping at the belt grinding operations by abrasive belts of various types, such as the ones:

– with the solid working area;

– intermittent, containing areas with abrasive grains and without them;

– composite, containing areas with abrasive grains, solid lubricant and without abrasive grains.

The technique includes analytical dependences for the temperature fields calculating, as well as equations for the thermo-physical parameters defining, which are necessary for these calculations, and a table with the values of the coefficient, determining what share of the thermal power, released while grinding, enters the workpiece while various groups of materials machining.

The article presents the results of numerical experiment on temperature fields calculation, performed relating to the belt grinding operations of gas turbine engine blades from VT9 and VT20 titanium alloys by abrasive belts of various types, namely, solid, intermittent and composite. It follows from the results of the experiment that at grinding the blades workpieces of the gas turbine engine inlet guide vane from the VT20 titanium alloy, application of intermittent belt instead of the solid one allowed temperature reduction in the contact zone of about 17.5%. At the same time, composite belt application instead of the solid one while grinding blades of the low-pressure compressor of the gas turbine engine allowed average contact temperature reduction by 38%. It was found that, depending on the machining mode, application of abrasive belts with intermittent working surface, i.e. with the sections without grains, as well as ones without grains and with solid lubricant allowed significant reduction, or total elimination of the burn marks on the machined surfaces of the work pieces.

Application of the foregoing technique allows predicting both structural and phase states of the surface layer of the workpieces being machined while belt-grinding operations in the presence of the metastable phase diagrams of the materials being machined.

According to the forecasts of the International Energy Agency, by the year 2040 the demand for liquefied natural gas (LNG) in the European Union will increase four times and twice in China. The LNG can become a greener substitute for oil and coal in the fast-growing urban areas of the developing world.

The Soviet Union was the first in the world to test a liquid hydrogen airplane in 1988, and in 1989 began equipment testing and research into the cryo-aircraft possibilities with the LNG utilization. Subsequently, several LNG-powered aircraft projects were developed, but they could not be realized for objective reasons.

One of the main problems of creating aviation cryogenic fuel system is the development of aviation cryogenic turbo-pump unit (TPU) capable of operating in the range of fuel consumption larger than the TPU for the space-rocket technology.

The article presents simulation of the aircraft turbo pump unit modelling, with account for the joint operation with the other units of the cryogenic fuel system.

Two TPU structures are possible in the aviation cryogenic system: the so-called “open scheme” and closed scheme. In the close scheme the pump driving is realized by the turbine, which working body is a cryogenic fuel warmed in the heat exchange unit. The pump driving in the open scheme is brought about from the external power source, i.e. electric motor. The closed scheme is more energy efficient, though it requires joint operation of the fuel system aggregates. The open scheme was selected as the object of research.

A mathematical model of the TPU, which has two modes of operation, has been developed for conducting computational and theoretical studies. The rated mode allows defining the TPU geometrical sizes. The non-rated mode allows defining the TPU basic parameters and plotting consumption-head-flow characteristic based on geometrical sizes, mass fuel consumption and input pressure. It should be noted that the TPU mathematical model operates in aggregate with mathematical model of the cryogenic fuel tank.

As the result of the calculation, the required power, pressure at the TPU outlet, as well as the flow and pressure characteristics of the pump are being determined by the aircraft flight cycle.

One of the trends for gas turbine engines cycle improving, allowing enhancing their efficiency, reducing specific fuel consumption and nitrogen oxides discharge, is exhaust gases regeneration through installing recuperator at the turbine outlet, in which a part of heat is being transferred to the air behind the compressor.

Comprehensive parameters optimization of the thermodynamic cycle of gas turbines, such as gas temperature T*₄ and compressor pressure ratior r*_cΣ, as well as parameters, defining the workflow of additional units like heat exchanger recovery factor, play an important role in its efficiency improving. Computer models of the bypass two-shaft turbojet engines with heat regeneration (TJER) developed in ASTRA CAE-system allowed realizing the problem solution of nonlinear multi-criteria optimization of their working process, and defining the most rational schemes depending of designated purpose and TJER operation conditions.

Based on the developed method of multi-criteria optimization numerical modelling was performed. The article presents the results of parameters optimization of the TJER working process in the system of Airbus A310 passenger plane by suc criteria as total mass of the power plant, and fuel consumed for the flight, as well as fuel consumption intensity per ton-kilometer and specific fuel consumption. The developed mathematical model for compact heat exchanger mass computing intended for solving optimization problems at the stage of conceptual design of the engine. The developed methods and models were realized in ASTRA CAE system.

The gas temperature increasing prior to a gas turbine engine (GTE) turbine is one of the key ways to its efficiency increasing. Operating temperatures in the turbine are limited by the heat resistance of the material, which the parts, interacting with hot gases are made from. In this regard, the task of developing and improving complex cooled blades that use compressed compressor air as a cooler becomes urgent.

Improvement of front cavity cooling system of the GTE turbine blade was performed in the course of the presented work. Analysis of thermo-hydraulic characteristics of various cooling systems options was performed to determine the most suitable structure.

The best option is the structure of the “Frankel packing” type, which represents the aggregate of channels crossing at a certain angle.

The study of the turbine blade cooled front cavity module was being realized according to the developed technique for computer aided design and calculation of heat exchangers. The technique for computer aided design and calculation of the plate-type heat exchanger may be applied for solving the wide range of tasks, including gas turbine engine design.

The proposed technique allows evaluating thermal and hydraulic characteristics of the cooling system with minimal costs, as well as optimizing the geometry of the heat exchange surface. Thermal characteristics for the three pen cross-sections under consideration were obtained by the results of the computational study according to the proposed technique.

Experimental study of the blade, being considered, was conducted according to the modular finishing technology by the calorimetric measurement in a liquid metal thermostat. Modular finishing technology envisages experimental studies of simplified blade modules.

Thermal characteristics for the three pen cross-sections under consideration were obtained by the results of experimental study of the blade front cavity.

Comparative analysis results of the calculated and experimental thermal characteristics of the cooling system of the front cavity module revealed the following:

- the most significant discrepancy of thermal characteristics occurs in the area of the entry edge of the front cavity;

- the less activity of heat removal is observed at the entry edge section, which indicates the fact that the structure under consideration has a potential for the heat removal increasing in the entry edge;

- the characteristics discrepancy over all sections is no more than 10%, which fits into the error of the experiment.

Application of the computer-aided design and calculation of thermal and hydraulic characteristics technique allows evaluating the thermal state of the designed blade with minimal costs and sufficient accuracy. It is advisable to use the coefficient of heat transfer from the blade outer surface to the cooling air as an evaluating criterion of the blade cooling system efficiency.

Despite the variety of the existing approaches, as of today, no universal technique, allowing accounting for the set of complex chemical and gas dynamic process while developing and modeling low-emission combustion chambers of gas turbine engines (GTE) accomplished in the framework of the LPP (Lean Prevaporized Premixed) concept has been developed. The LPP-chamber operation is based on low-temperature combustion of a pre-prepared “poor” air-fuel mixture with excess-air factor of 1.8-2.0.

The presented article proposes a method for the multilevel modelling implementation in the GTE low-emission combustion chamber design process. Combustion chamber accomplished in the framework of the LPP concept was selected as the object of the study. This concept is based on the combustion of pre-prepared “poor” air-fuel mixture.

Multilevel modeling includes three stages of computing: designing calculation, one-dimensional modelling, and gas dynamic processes modeling. The article presents the formed appearance of the combustion chamber and its elements in accordance with the proposed technique. Parameters computing along the flame tube length of the three chambers, where burner devices with different swirl angles of the swirl vanes were installed, was performed.

The calculations were being performed in the ideally gas approximation of the incompressible homogeneous environment in the adiabatic statement of the stationary problem.

The two-parameter RNG k- ε model with standard wall functions was used as the turbulence model.

Combustion was being modelled by the aggregate of laminar flamelets in the turbulent flow of unmixed components. The Kee58 mechanism, including eighteen mixture components and fifty-eight chemicalreactions was considered as a set of methane oxidation chemical reaction.

The NOx content computing in combustion products was based on thermal and super equilibrium mechanisms of NOx formation.

Analysis of the obtained results revealed that increasing of the twist angle in the blade swirl of the burner device leads to fundamental changes in the flow structure in the primary zone of the combustion chamber, which affects the change in emission characteristics as well. The chamber with the burner device with the twist angle of 45° ensures the best optimal emission characteristics on nitrogen oxides.

An important task in the thruster design is determining its basic geometrical dimension, which will define its thrust and specific characteristics. By specifying the main standard size of the thruster, we lay the foundation of the design and therethrough directly determine its operating range. Thus, it is especially important to understand what parameters can be obtained from the thruster at the initial stage of its design.

To solve the set problem, it was necessary to switch to dimensionless parameters that would characterize the thrust and specific characteristics of the thruster. The presented work derives the basic dimensionless parameters, characterizing the thruster operation from the viewpoint of energy consumption and working fluid utilization. The obtained coefficients allow characterizing the thruster operation regardless of its geometric dimension, and comparing operation parameters of thrusters of different standard sizes operating in different power ranges among themselves.

Thus, analysis of stationary plasma thrusters, developed by the “Fakel” Design Buro, was performed by the newly presented dimensionless parameters. The analysis was conducted for a single working liquid, namely Xenon, and a single discharging voltage of 300 V. As the result, the dependencies of the working liquid utilization factor and consumption ratio on the discharge current density were obtained.

It should be noted that, despite the differences in the thrusters’ standard sizes and the sizes of the discharge channel, the curves with characteristic working zones were obtained for the entire family of thrusters. The optimal operating range for stationary plasma thrusters, which corresponds to the discharge current density from 0.07 to (0.015–0.02) A/cm2, depending on their design features, was determined in the course of the analysis.

Eventually, with known operating power range, necessary for set task accomplishing, it is possible to determine geometric dimension of the thruster based on the optimal operation area of the engine, as well as define approximated thrust and specific characteristics of the thruster being developed by simple transformations, obtained dependencies of working liquid utilization factor and energy consumption ratio.

The presented work proves possible power plant reequipping options of the Su-27 type fourth generation fighter with new engines.

The research scientific task was formulated for this purpose. The set task consists in effectiveness assessing of the Su-27 type multifunctional fighter with the power plant based on the operational bypass turbojet with flows mixing and Al-31F afterburner, and the four options of its re-motorization while typical flight task performing using methods of mathematical modeling.

The aircraft re-deployment from airfield No. 1 to airfield No. 2 was assumed as a flight task, which was stipulated by sufficient technical substantiation for the decisions made, with relative simplicity of the engineoperation mode modeling

Technical parameters, characterizing the aircraft under study on the assumption of its assignation, namely the total flight range and climbing capacity, were assumed as the performance criteria. These criteria are controversial since the climbing capacity relates directly to the thrust-to-weight ratio, while the flight range relates to it inversely, having herewith a certain local optimum, which means that the effectiveness assessment can be soundly performed by these technical criterions.

The research technique was developed by the authors based on the multi-disciplinary analysis methodology and development of “Aircraft – Power plant” system technical profile at the preliminary design stages. The ThermoGTE and “Aircraft-Engine”instrumental-software systems, being more than once approved in aviation industry and demonstrated high efficiency, were employed as the basic tools for performing computational-theoretical studies.

Parameters and characteristics computing of the power plant was being performed in ThermoGTE. The data arrays on altitude-airspeed performance were being imported hereafter to the «Aircraft-Engine» software for subsequent trajectory parameters computing. Aerodynamic scheme of the object under study, by which aerodynamic and specific-weight characteristics of the aircraft, the flight program and profile, consisting of fifteen sections, were computed, was formed as well. The engine operation modes and conditions of execution were defined for each segment of this flight program.

As the result of the performed studies, values of trajectory parameters of the studied aircraft motion with five options of the power plant layouts being studied while the flight task performing. Efficiency assessment of the aircraft under study by the assumed criteria, which demonstrated the possibility of its efficiency improvement compared to the power plant based on the AL-31F engine, was performed.

This work practical value consists in the fact that its results can be employed in scientific and design organizations, engaged in development and modernization of serial and prospective aircraft and their power plants; Air Force and aviation industry ordering organizations while substantiating requirements to aviation engineering prototypes; as well as aviation engineering universities while educational process improving.

At present, more and more attention is being paid to the idea of geostationary satellites servicing with automatic spacecraft. This idea realization requires creation of service spacecraft, high precision and stable algorithms for autonomous navigation and spacecraft motion control. To ensure accuracy while such algorithms developing, it is necessary to account for deterministic and random disturbances, caused by natural factors, errors in control system elements operation, as well as navigation errors. A software- mathematical complex, which allows performing a spacecraft motion simulation in both deterministic and stochastic statements, was developed for algorithms testing and effectiveness evaluation.

To perform the basic task, the software-mathematical complex should ensure compatibility with mathematical programming libraries, the ability of quick modification of the designed algorithm structure, and convenient intuitive user interface. For meeting the above said requirements, the complex is being designed and implemented employing object-oriented programming of both the software complex itself and control algorithms. The complex structure is modular, in which control algorithms’ module, module of the spacecraft onboard systems model and module of the external environment model were elaborated independently from the kernel. Such complex architecture allows studying various options

Such architecture of the complex allows exploring various options for the control block building. The current version implements algorithms for the service module control when bringing it to the vicinity of the target module working position and holding it relative to the target module for inspection.

The service module control algorithms in the software-mathematical complex were developed based on linearized models of motion of the service and target modules in the vicinity of a circular orbit with the specified radius. These models account for the disturbance from the Earth, Moon and Sun gravitational fields, as well as the error of direction and value of the thrust of the correction engine. Combined optimization method is used while the problem of optimal control solving. Control algorithm for the service module at the stage of its being held relative to the target module was developed using the model of relative motion with the assumption of the steady-state mode existence.

The software-mathematical complex operability is being confirmed by the simulation results of the service module motion control algorithms at various stages of its functioning in both classical and stochastic statements.

The article regards the problem of the coordinates measuring system state assessing of the short range and near-in operating radius unmanned aerial vehicle (UAV) in conditions of abnormality (distortions) of measurement results obtained from the satellite navigation system (SNS). Optoelectronic system, incorporating both TV and thermal imaging information channels, as well as laser rangefinder of the target indicator is being considered as an extra information source.

This article urgency is stipulated by the necessity of positioning the short range and near-in operating radius UAV with restricted mass and size parameters without employing additional or high-accuracy measurement instruments onboard with full (partial) absence of satellite signals in autonomous flight mode.

The purpose of the article consists in preserving the UAV current position determining accuracy in conditions of partial or complete absence of the signals from the SNS.

The object of research is the UAV navigation system.

The subject of the research is navigation information processing processes in conditions of partial or complete absence of the satellite signal.

The scientific novelty of the research is stipulated by the development and scientific substantiation of a comprehensive technique for optimal estimation of the UAV current position by visual navigation method, allowing correction amendments forming to refine the UAV spatial position in the presence of the extra information source.

Theoretical significance of the results consists in supplementing of visual navigation methods by forming coefficients, characterizing the sparseness of the terrain exceptional points and actual share of the reference image generality from the current one, allowing determine the UAV’s sufficient altitude over exceptional points of the underlying terrain. Computation of the correction image period forming, with the regard to the instrumental errors of the strapdown inertial navigation system (SINS) based on micro-electrical and micromechanical systems was performed as well.

Practical significance of the research lies in application of integrated technique in the small-sized vehicles positioning problems in the absence of signals from the SNS, as well as substantiating intelligent image processing employing high-performance, small-sized equipment on board the UAV.

The experiment demonstrated that in the absence of the SINS correction, the UAV accumulates the maximum positioning RMS error on an average of 150 m during the first minutes of flight. With regard to this and the maximum possible UAV speed of the of 120 km / h, at a distance of 5 km from the launch point the limiting RMS error of positioning, during the return flight, will be about 300 m, which can lead to the UAV loss. The UAV correction according to the formed correction areas allows to reducing the RMS error to 200 m.

Thin glass elements made of K-208 brand of radiation-resistant optical glass are employed as protective coatings for solar cells and thermo-optical coatings for radiators-heat exchangers of spacecraft thermal control systems.

The glass elements manufacturing technology is based on heating polished glass blocks fr om K-208 glass to highly viscous state with subsequent glass tape extrusion through the stainless steel die.

The glass tape size-cutting and blanks obtaining of the required size is performed with diamond tools for scribing, or by the laser thermosplitting technique.

The presented article studies strength characteristics and heat resistance of glass elements fabricated by various techniques after the long-term storage process, which partially models operation process of such elements in space.

The test results reveal that samples fabricated by the laser thermosplitting method have the same strength after long-term storage, as samples tested after their manufacturing in 2007. This can be explained by the fact that this technology does not produce edge effects, which define the end strength of glass elements. The strength of the samples obtained by the diamond scribbling deteriorated after such a long-term storage period, which is stipulated by the temporal evolution of edge defects.

Thermal resistance of the K-208 ultra-thin glass with the edge obtained as the result of its laying-out by laser is at least 20-30% higher than with the edge obtained by the laser scribing which is of prime importance for the products employed in space engineering, wh ere large temperature drops occur.

The obtained results of experiments confirm high efficiency of the controlled laser thermosplitting while glass elements manufacturing from the K-208 thin glass for the spacecraft temperature-controlling coatings.

Mechanical strength and thermal resistance of glass elements after long-term storage are sufficient for their application in space-rocket engineering products.

Thermal-cycle tests of uncooled working blades of the second stage of the new generation helicopter gas turbine engine turbine were conducted, and changes in the composition, structure and micromechanical properties of the heat-resistant coating were studied.

The blades are made of the new VZhL-21 poly-crystalline casting alloy. The heat-resistant coating was applied employing the MAP-2 installation according to the serial technology by successive applying of the condensed layer of the Ni-20Co-20Cr-12Al-Ti-Y composition (inner layer) and diffusion layer of the Al-5Si-B composition (outer layer).

Both condensed and diffusion layers were being applied in vacuum at the specified parameters of the arc current and bias voltage at the products for 200-220 and 60-65 minutes respectively. After this, vacuum thermal processing of the blades was performed at the temperature of 1000 °C for 240 min to complete the coating structure and phase composition formation.

Comparative tests of blades with and without coating were conducted under identical conditions on a special test bench by a technique that ensures the thermal cycle reproducibility while multiple repetitions. The principle of operation of the experimental setup consisted in the ohmic heating of the test blade with direct electric current, varying according to a given algorithm. The thermal cycle selected for the blades testing was calculated based on an engine test: heating to 480 °С (120 s exposure at this temperature), temperature raising to 770 °С (150 s exposure). Further, cooling to 480 °С (120 s exposure), and cooling to room temperature. After the predefined running time, the blades were being removed from the test and subjected to microstructural and micro-chemical studies of the coating state on the JSM6460-LV scanning electron microscope with the INCA ENERGY 300 energy dispersive attachment, as well as micromechanical measurements on the Shimadzu DUH-211 dynamic ultramicrotester (Japan) using Berkovich indenter. The results of the studies revealed that the coating microstructure on all tested blades had not undergone significant changes compared to the initial one.

In the process of the thermal running time of 500-800 cycles, there is an aluminum diffusion from the coating surface to the contact bound of both coating zones and further to the blade surface. With the running time increase up to 1350 thermal cycles, aluminum diffuses deeper into the blade metal. The character of chromium diffusion seems to be more complicated. Chromium concentration changes insignificantly on the coating surface. However, in the place of the contact of both zones the chrome concentration reduces drastically at running time of 500 cycles and stays at the attained level up to the maximum running time of 1350 cycles. Finally, the “coating-blade” contact zone significantly enriches with chrome.

The creep of the coating material remains at approximately the same level up to 800 thermal cycles, and then increases sharply, while the share of the plastic component of the mechanical work on deforming the coating material starts increasing sharply somewhat earlier, beginning from 500 cycles.

Thus, the performed comprehensive study allows predicting the coating protective functions preserving for no less than 500 thermal cycles.

The article presents the studies of the process kinetics of obtaining nanopowder of metallic cobalt by hydrogen-reduction method under non-isothermal conditions, as well as properties analysis of the initial material and obtained products. Cobalt nanopowder was being obtained by hydrogen reduction of cobalt hydroxide nanopowder in the linear heating mode at a rate of 15°C/min within the temperature range from 25 °C to 500 °C. The Co(OH)₂ nanopowder was synthesized in advance by chemical precipitation from aqueous solutions of cobalt nitrate Co(NO₃)₂ (10 wt. %) and alkali NaOH (10 wt. %) under conditions of continuous stirring, control of the T = 20 °C temperature and pH = 9 acidity. Kinetic parameters of the hydrogen reduction process under non- isothermal conditions were calculated by the differential-difference method using the data of thermo-gravimetric analysis and non-isothermal kinetic equation. The phase composition and structure of the samples were analyzed by the X-ray method. The specific surface area and average particle size of the powder samples were determined using the Brunauer–Emmett–Teller (BET) method by the low-temperature adsorption of nitrogen. The morphology and size of the nanoparticles were studied by scanning and transmission electron microscopy. It has been established that the process of non-isothermal hydrogen reduction of Co(OH)₂ nanopowder occurs within the temperature range from 180 °C to 310 °C with a maximum rate 222.34·10^-5 s^-1 at a temperature of 280 °C. The dependence of the degree of conversion on еру temperature during the Co(OH)₂ reduction process has been determined in the form of a mathematical function y = 0,0756·e^0,0248x. The value of activation energy for the Co(OH)₂ nanopowder reduction process was found to be ~45 kJ/mol, which indicates a mixed reaction mode. It was revealed that the Co(OH)₂ hydroxide reduction at a temperature of 280 °C allowed to accelerating the process while ensuring the required properties of the product. The obtained metallic cobalt nanoparticles have a spherical shape with a nanometer size (about tens of nanometers) and are in a sintered state. Each of them herewith is connected to several neighboring particles by isthmuses.

The main technical characteristics of automatic landing systems (ALS) of manned and unmanned aerial vehicles (AV) are derivate of the characteristics of automatic control systems. The performed analysis of literary sources devoted to the study of the AV automatic control issues at the landing stage revealed a deficit of survey and analytical work, considering comprehensively the problem of the AV automatic control forming during landing process.

The purpose of this work consists in studying the AV spatial position control issues, relevant for the ALS of both manned and unmanned AVs, revealing the main problems getting in the way of AV ALS development and preferred technical solutions, which can be employed while the AV ALS creation.

To achieve the set goal, the following aggregate of systematic interrelated methodological approaches was applied to reveal the basic pros and contras of the objects being analyzed. These approaches are based on:

- search and analysis of scientific and technical literature, and its systematic review;

- analysis of trends to reveal the dominating ones in the ALS development with regard to the AV information support and control;

- SWOT-analysis.

The performed information search on the issues of AV control forming while automatic landing (AL) in scientific and technical literature and other open sources, its analysis and systematic review allowed outline the two groups of techniques for the AV control forming while the AL process:

- control actions forming based on the object state vector, being formed by the information support means;

- control actions forming based on the preprocessed information, being formed by information support means.

The techniques for the automatic control forming related to the second group are of practical interest, thus the subject matter of the article is limited by them.

The works, being analyzed, devoted to the AV control in the process of the AL performing are classified in accordance with to the following problematic areas, to which studying they are dedicated:

- the ALS architecture;

- synthesis of automatic control algorithms;

- fuzzy control;

- the AL process optimization;

- the AL process mathematical modelling.

The technical solutions proposed in the framework of the outlined problematic areas were analyzed, their advantages and disadvantages were revealed.

The authors proposed to employ multi-level architecture, Kalman filter, Luenberger observer, and model-oriented method for designing automatic control systems as the ALS technological base. The inertial navigation system, being corrected by the iformation obtained from the satellite navigation system with functional add-ons (differential navigation), and radio navigation system as a stand-by information source can be proposed as the AL information support core. The article presents a functional diagram of the ALS built on the proposed principles.

The automatic control system for the AV during the AL execution can be recommended to be built based on stabilization of the set flight path using linear deviations from it and, possibly, changing of the rates of these deviations. This approach will allow employing the constant gains in contrast to the variable coefficients used in the case of the of angular deviations application. Besides, the ALS should ensure the lack of necessity of the crew intervening in control process at low altitudes even in the case of control resources degradation, and preserve its operability in conditions of external information sources loss.

The article considers braking modes control of the small spacecraft (SS) of the CubeSat type by aerodynamic braking units. The controllability area for hitting any atmospheric entry point employing boundary condition for the range angle and angle of entrance is under consideration. When employing aerodynamic braking system, it is necessary to tend to obtain the range angle value ensuring hitting the specified region of the Earth surface for safe fall of SS fragments and the angle of entrance guaranteeing the SS burning out in the dense atmosphere.

The problem of finding optimal control of the SS with IAD can be solved stage-by-stage. Initially the problem of minimizing the flight time from the initial orbit to the atmospheric boundary is being solved. Then the requirements for the final values of the trajectory parameters of the aerodynamic braking section are being determined. Finally, the control law σ_x(t) should be found, which ensures the SS hitting the specified region of the phase coordinates.

As the result of the proposed approach, the complex task of optimizing the trajectory of SS is reduced to solving two problems: first, at the interorbital transfer section prior to atmospheric entry, and then at the section of main aerodynamic deceleration in the atmosphere. This allows eliminating the jumps of the right-hand parts in the formulated problem and simplifying it significantly without breaking the generality.

The study of the effect of perturbing factors acting on the SS of a CubeSat type with IAD was conducted, and the impact of variations in the atmospheric density was demonstrated. Ballistic analysis was performed using various control laws of the SS using IAD with 1–2 balloons, in condition of hitting the specified area at the boundary of the atmosphere with account for the levels of solar activity. Analysis of the possibility of control by the control function changing (ballistic coefficient) was conducted. A comparative assessment of the considered control programs was performed, depending on a number of basic conditions for the restrictions of the motion control problem of the SS with IAD.

One of the main stages in the design and modernization of the aircraft is a wind tunnel experiment. For this reason, further development and improvement of the wind tunnel test technique is necessary. A number of fundamental problems have to be solved to improve the accuracy of the experimental studies, one of them is the implementation of an interference free flow over the model. Existing approaches, such as permeable walls (perforated or slotted), adaptive walls or jet boundaries, do not allow us to close the issue of test section walls influence on aerodynamic characteristics of the model due to some disadvantages. In the framework of this analysis, a prospective boundary condition is studied, which is a combination of perforated boundaries and a controlled boundary layer.

The efficiency of using combined jet-perforated boundaries was investigated in test series with the models of aircraft and missile layouts at high subsonic and transonic speeds. Models were tested in solid and perforated walls, as well as in combined jet-perforated boundaries in TsAGI T-112 trisonic facility.

Models of civil aircraft were geometrically-similar schematized models. An approach based on the use of geometrically-similar models allows us to obtain useful estimates of the effectiveness of applying certain boundaries. It is assumed that proper choice of boundary conditions should ensure the coincidence of the obtained aerodynamic characteristics of various scales models. As a result, the basic aerodynamic characteristics of the models were obtained, as well as in the model location zone the boundary layer parameters were measured. The obtained experimental results show that the use of combined jet-perforated boundaries causes a noticeable increase in the boundary layer and its integral characteristics (the displacement thickness and the moment thickness). Thus, the curves corresponding to the lift and pitch moment coefficients in the combined jet-perforated boundaries coincided almost completely that indicates the least influence of the walls of the WT test section.

To analyze the obtained experimental results, numerical modeling of the flow around three-dimensional models was carried out. Numerical research at various boundaries makes it possible to significantly reduce the required amount of experimental studies. Simulation of the unbounded flow around the model allows obtaining the interference-free aerodynamic characteristics of the model, which must also be obtained with the correct selection of the boundary conditions in the wind tunnel test section. Their complete coincidence means solving the wall interference problem.

As a result, a comparison was made of the obtained experimental data in a wind tunnel and a numerical study for the missile layout model. The comparison was carried out for the lift and pitch moment coefficients depending on the angle of attack. Finally, it can be concluded that a new type of boundary condition that is a combination of perforated walls and a controlled boundary layer can effectively eliminate the influence of the WT test section walls on the aerodynamic characteristics of the model. Thus, new type of boundary condition has great prospects for implementation in new aerodynamic installations, as well as in the modernization of existing ones.

It is customary to assume that a multi-dome parachute system is a system with the number of domes in the bundle of two and more [1–20]. Efficiency of the multi-dome parachute system is understood in this work as the ability of the object – multi-dome parachute system ability to perform its functions within the framework of the specified values of its critical (most important parameters).

The presented work considers some critical parameters liable to be determined while the flight-testing of the system, comprising an air drop object and multi-dome parachute system, such as landing speed and non-simultaneity of the domes filling process.

The article presents the dependence of the vertical component of the landing speed, being determined while the multi-dome parachute system design computations, which is assumed as an average valued (mathematical expectation) of the real value of the vertical component of the landing speed, as it is a random value in reality. The most probable random error of the landing speed function was determined with account for inaccuracy of measurements of all arguments included in the function structure, which allows evaluating contribution of each error component to the speed determining error, as well as find the largest one and minimize it.

Further, alongside with accounting for the atmospheric parameters, the possible active impact of near-Earth atmospheric turbulence on the value of real vertical component of the landing speed was being reckoned in.

The experimental results on determining the average value of real vertical component of the landing speed, reduce to the standard atmospheric conditions at the sea level and regular weight according to the data of a series of flight-testing, are presented.

The article presents the dependence of distribution density and probability of not exceedance of assigned value of landing speed’s vertical component for a special case.

The authors marked the possibility of appearance of insignificant number of “jumping-out” measurements under the impact of intensive, powerful surface atmospheric turbulence on the multi-dome system.

The article presents the detailed analysis of the phenomenon of non-simultaneity of the domes filling process in the bundle. Substantiation of the non-uniformity parameter importance for the multi-dome parachute system operation effectiveness is being brought forward.

The authors introduced a parameter named the coefficient of domes in the bundle filling simultaneity. The notions of leaders and outsiders for the domes in the bundle were introduced as well. The analysis of their role in the domes filling in the bundle was performed. The article presents physical explanation of the domes filling non-uniformity phenomenon. Some important effects, associated with the non-uniformity phenomenon, as well as factors affecting the non-uniformity of the domes filling in the bundle were considered.

Certain experimental data on the non-simultaneity of domes filling in the bundle is presented for the possible theoretical studies in the future of the non-simultaneity of domes filling phenomenon. The article presents the experimental data by the time intervals of the domes filling process in the three-domed corrugated parachute system with the area of the single dome of F_S = 600 m², while the airdrop of the object of m ≈ 3 tons in a wide range of ram air of the system implementation according to the data of forty four flight experiments. The experimental data on the time intervals of the four-dome parachute system with the area of the single dome of F_S = 760 m², while the airdrop of the object of m ≈ 6 tons according to the data of eleven flight experiments.

The above-mentioned data can be used effectively for checking the adequacy of the mathematical models under development of simultaneity (non-simultaneity) of the four-dome parachute systems filling.

The above data may be effectively used for the test for goodness of developed mathematical models of simultaneity (or non-simultaneity) of canopies filling in the four-dome parachute system.

The relevance of the problem of enhanced acoustic comfort ensuring for passengers and cockpit personnel is beyond doubt. In particular, at present, there is a problem of professional diminished hearing among the aircrew members of civil aviation aircraft of Russia. The risk factor of this malady development is the noise inside the cockpit.

The problem solution of acoustic comfort ensuring in the cabin is impossible without fulfilling a complex of engineering and fundamental studies at all stages of creation of new samples of aerotechnics. One of the trends of the studies is identification, localization and ranging by intensity the main noise sources in the cabin of the aircraft-prototype. The results of this study are necessary to ensure optimal placement of sound proof, sound absorbing and vibration-damping materials in the onboard structure, and issue recommendations on noise reduction of the air conditioning and ventilation system.

The article presents the results of localization and ranging by the intensity of the noise sources in the RRJ-95 aircraft cockpit, employing the 3DCAM54 spherical array.

Acoustic measurements were performed on the RRJ-95 experimental aircraft No 95005 with the cockpit, updated from the viewpoint of noise reduction and reverberation disturbance. The tests were performed at the cruise speed mode at the altitude of 11 km, determined by the flight Mach number of 0.8.

Measurements were performed at the routine operation mode of the air conditioning and ventilation system and at its turn-off.

As the result of the conducted studies, the noise sources localization maps in the one-third-octave frequency bands of 630-3150 Hz were obtained. The main noise sources in the cabin are the air conditioning and ventilation system (ACVS) and the noise of the turbulent boundary layer. As far as the air feeding is being terminated after the ACVS turn-off, but the fans are not turned-off, the ACVS impact manifests itself while its turn-off from the side of ducts feeding air to the cockpit. The two basic mechanisms can be outlined in the ACSV noise. In particular, in the noise of the one-third-octave frequency band of 1000 Hz, the ACVS turbulent flow dominates the noise caused by the "rotor-stator’ interaction in the ACVS fans. In the one-third-octave frequency band of 1250-2500 Hz the noise of "rotor-stator’ interaction prevails while fans operation.

When designing modern highly maneuverable unmanned aerial vehicles (UAVs), one of the most urgent tasks is studying aeroelastic stability of the rudder-drive system, since the stability loss in the above-appointed system can lead to the general instability of the UAV stabilization system, which is unallowable. To ensure stability of the “aeroelastic UAV–stabilization system” circuit, the requirements on bandwidth and gain level, as well as necessary phase lag in the strictly defined frequency band are being imposed on the rudder drive. All this, in its turn, complicates the problem of ensuring stability of both the UAV stabilization system and the rudder-drive system.

The article presents the results of studying the aeroelastic stability of the rudder-drive system of the highly maneuverable UAV studying. They are based on the frequency characteristics and processed signals comparison at the output of the isolated drive with constant load, and at the output of the drive loaded with the rudder that oscillates within the frequency range of the structure elastic vibrations. The electric drive with digital microcontroller regulator, being employed at present as a part of stabilization system of the highly maneuverable UAV was considered as a drive. A hinge moment gradient, characterizing the drive loading by the rudder performing flexural-and-torsional vibrations in the supersonic aerodynamic flow, was obtained. Nonlinear mathematical model of the rudder drive with digital microcontroller regulator was used as a research tool.

The main results of the study are the transfer function coefficients of the dynamic hinge moment, and obtained frequency responses of the “rudder-drive” system for the UAV flight mode under consideration. The results of the “rudder-drive” system studying allow concluding that that the considered drive, being loaded by the rudder, vibrating within the range of the structure elastic vibrations, can be used as a part of the UAV stabilizing system.

The considered in the article technique for the transfer function of the dynamic hinge moment forming is invariant relative to the drive type and aerodynamic flow kind (sub- or supersonic). In this regard, the results of the studies obtained by its application can be employed while solving the variety of the problems on the stability ensuring of the stabilization systems of various UAV classes with regard for aeroelasticity.

The article analyzes a possible design layout of a promising small spacecraft for technological purposes. Specific requirements for such devices are requirements for micro-accelerations, which, on the one hand, determine the possibility and feasibility of implementing a particular gravity-sensitive technological process onboard the spacecraft, and, on the other hand, impose requirements for the orientation and motion control system of the spacecraft.

Since there are no fully implemented projects of small spacecraft for technological purposes at this stage of space technology development, the experience of designing and operating medium-class spacecraft and orbiting space stations is under discussion. However, small spacecraft have their own specifics in terms of the super-dense layout. Thus, while designing small spacecraft this experience should be significantly reworked with account for this feature.

The design requirements for the small spacecraft and its orientation and motion control system are formed in view of meeting the requirements for micro-accelerations that contribute to the favorable implementation of gravitationally sensitive processes, and with account for other features of small spacecraft. This feature consists in a significantly higher ratio of the mass of elastic elements to the spacecraft total mass for a small spacecraft than for spacecraft of other classes. This feature affects the actuating devices selection of the orientation and motion control system of a small spacecraft, as well as the characteristics of these actuating devices.

The results of this work can be used in the development of small spacecraft for technological purposes.

Laser welding technology application in the aerospace industry will significantly reduce the weight of the aircraft structure, material consumption and production time for parts and accessory manufacturing.

The thermal cycle of laser welding ensures minimum time of the area staying in the overheated state, eliminating thereby the possibility of grain growth and mechanical properties reduction of steels.

The article presents the studies of structural changes in the weld metal obtained by the unity effect of laser radiation on the steel surface.

The performed microstructural analysis allows establishing the weld metal formation staging, and its components, including the microhardness defining in each particular zone, which contributes to understanding and predicting the behavior of the weld metal while parts or structures operation.

The three most pronounced zones were defined while the unit laser impulse effect. They are:

1 – the arc-like zone of the dendrite structure.

2 – the recrystallization zone, located symmetrically to the zone 1. The structure of this zone is distributed randomly, the tempering bainite mainly prevails.

3 – the tempered perlite zone with uniformly sized grains of an average diameter of 40–70 microns. Zone 3 adjoins zone 2 and the welding spot surface.

One more zone with extremely insignificantly distorted structure of the basic metal is being observed under the weld-fusion line towards the basic metal.

Analysis of the average area of the zones revealed the following: zone 1 has a predominant area of 51.2% of the total weld metal area, and 47.5% along the computed volume.

High crystallization rates contribute to the dendritic structure development of zone 1, and the heat-affected heating zone therewith contributes to the uniform tempering of zones 2 and 3 and formation ofstructures of bainite and tempering sorbite respectively.

It was established as well that in the process of exposure, temperature conditions are being created for recrystallization and tempering of quenching structures. Thus, to ensure equal strength of the welded joint with the base metal, it is necessary to recommend tempering to relieve residual stresses and partial recrystallization of zone 2 even for low carbon steels.

Agricultural aviation is aviation employed for agricultural work. Agricultural aviation is applied most often for spraying fertilizers (pesticides, herbicides, insecticides) on agricultural crops, as well as for crops fertilizing, defoliation, desiccation, and somewhat less often for air seeding (hydro-seeding, i.e. seeds sowing with water flows under pressure).

The agricultural airplane developing is a necessity since it ensures the most effective work, associated with watering and visual surveillance of the acreage planted.

Besides, the agricultural land cultivation is being performed at the best agrotechnical terms, such as early spring, when the ground machinery is not yet able to operate due to the impassability.

The study consists in analyzing the most important problems of agricultural aircraft designing, using modern CAD, CAE systems. The authors considered several small Russian airplanes, on which basis the primary technical characteristics of the future product, as well as the most successful solutions of the airframe were selected. A detailed justification of the aircraft airframe layout is presented. The main problem of this project consists in the lack of competitive small aircraft from the domestic manufacturers, meeting modern requirements and economic capabilities of the potential consumers.

A 3D model of a piston-engined single-engine monoplane with a low-lying wing, which shell is made of composite materials, was designed as an object of research. Composite materials application for the aircraft airframe allowed solving plenty of the problems associated with the corrosion resistance, as well as enhance the landing gear struts reliability, which strength is especially important for the takeoff from the unprepared runway. The article presents solutions on structural appearance of the airframe elements and aggregates from modern composite materials, ensuring the possibility of developing and manufacturing of competitive aircraft of the “small aviation”. Digital modelling techniques were employed while this airplane creation, which allowed developing reasonable aerodynamic scheme.

The presented study is focused on determining the impact resistance and survivability characteristics of panel samples with the honeycomb filler and fragments of helicopter blades. The problems, associated with developing and producing the experimental samples impact tests performing, as well as studying the character and geometry characteristics of damages were being solved while these works execution. The authors developed a technique for impact resistance determining of aircraft sandwiched composite parts with honeycomb filler. The composite sandwiched structures in the form of the helicopter steering and main rotor fragments, and standard samples of the sandwiched panels with the honeycomb filler were the objects of the study. Carbon composite skins and honeycombed filler from aramid paper were employed for the panels manufacturing. The blade fragments represented the structures composed of T-25 fiberglass plastic layers with honeycomb or foam filler placed between them.

A technique for inflicting impact damages by vertically falling load, and registering such parameters as impact energy, maximum loading and impactor penetration depth was developed while laboratory studies. Application of piezometric transducers while impact tests allowed registering diagrams of the impact damage, which, besides the general energy-force assessment, allow step-by-step studying of the impact loading. The impact energy for the samples of sandwich-panels was being selected from the condition of incomplete destruction ensuring (2 J), and initiating significant damages of the skin and filler (10 J). The damages character studies of the helicopter steering and main rotor blades fragments were conducted within the energies range of 5–50 J. The depths of dents and cracks were determined by the digital indicator head. Computer tomography was employed for internal diagnostics of the damaged samples. Tomograms of the blades sections allowed studying stage-by-stage growth of damages in dependence of the impact loading increasing.

It can be declared by the results of this work that already small impact energies lead to dent on the skin forming, and crumpling of the honeycomb filler with partial destruction. At the impact energy of 10 J, significant destruction of skins and filler under them is being observed. The breakdown and cleavage of the skin material along the panel length are being observed on the external side of the sandwich-panel subjected to the impact. The tomographic images of the tail rotor blade show fractures of the fiberglass plastic layer and crumpling of the foam filler. Analysis of the main rotor blade sections also revealed the fracture of the skin upper layer and subsequent compression of the honeycomb filler.

Fuel resources provision is a key problem of the industrial and post-industrial world economies development. In this regard, science and technology are facing the problems of developing new alternative types of fuels in return of the conventional oil fuel or liquefied hydrocarbon gas. One of these fuels is cryogenic fuel, which is currently widely used in rocket and space technology. It is customary to assign the liquid hydrogen, liquefied natural gas (LNG) and cryogenic propane to the cryogenic fuels. These fuels are more environmentally friendly than traditional aviation kerosene, as well as possess better thermal properties, such as greater calorific value, cooling resource and the value of the gas constant, which determines the workability of the gasified cryofuel. This provides a potential opportunity to obtain high flight characteristics of promising aircraft.

The Russian Federation ranks the first in terms of proven LNG reserves in the world as of 2018. In this regard, the LNG is the most optimal choice of cryogenic fuel for Russia. However, to get the maximum benefit from the LNG application, the properly designed cryogenic fuel tanks (CFT) for the cryogenic fuel storing onboard an aircraft, and accounting for the thermo-physical and hydrodynamic processes in the CFT are necessary. For example, disturbances on the surface of the cryogenic liquid in the tank can affect the main CFT parameters (heat flows, temperatures, and pressure), which can lead to the early response of the safety valve (SV), and, consequently, to a greater loss of fuel through the SV.

The article presents a comparison of the CFT thermal state in the presence of vacillations on the liquid surface and in their absence. The LNG in the tank herewith is at the saturation line. It was found in the course of the study that the presence of disturbances on the liquid surface led to the increase of thermal flow between the gas in the above-the-fuel area and the liquid fuel by 69.85 W.

In the presence of fluctuations, the gas temperature in the above-the-fuel area is less by 18.47 K than in their absence at the accepted initial data. However, the presence of disturbances on the liquid surface does not practically affect the mass of the fuel discharged through the SV, since the LNG in the tank is at the saturation line. With the presence of vacillations, the thermal flow between the gas and liquid in the tank, evaporation rate (gas mass) and pressure in the above-the-fuel area are increasing, but the LNG boiling temperature rising herewith as well.

At present, the LPP (Lean-Premixed and Pre-vaporised) concept is one of the most effective concepts of the low-emission fuel combustion, which is based on the low-temperature (T_flame =1800—1900 K) combustion of pre-mixed “poor” fuel-air mixture (FAM). This concept foresees thorough mixing of the fuel with air in the burner prior to feeding to the combustion zone. It is well-known that technical perfection of these burners ensures successful problem solving of nitrogen oxides and carbon monoxide release reduction while maintaining high efficiency and stability of the combustion process. Thus, the efforts aimed at studying these burners design impact on the emission characteristics of the flame are necessary while development and adjustment of combustion chambers of gas turbine engines, accomplished within the framework of the LPP concept.

The presented article considers the structure of the dual-circuit burner of the low-emission combustion chamber of the gas turbine engine, operating on the natural gas. The results of the studies of the three burners differing by the size of the outlet part of the developed swirler hub are presented. The article presents also the results of the components concentration measuring of the final gas mixture, in particular carbon monoxide CO, nitrogen oxides NO and unburned hydrocarbons CH in the combustion products. Computation of the fuel combustion efficiency was performed. Selection of a burner, which demonstrated minimum of value of nitrogen concentration and maximum combustion efficiency level and carbon monoxide in the samples being drawn was conducted. The best appeared to be a burner having a structure with the central body diameter to the outlet nozzle diameter ratio of A/B = 0.62.

The presented article deals with the new specific parameters elaboration necessary for more qualitative analysis of a jet engine operating on liquid hydrocarbon fuels. The purpose of the article consists in elaborating specific parameters, which would be able to account for the degree of carbonization and failure of the jet engine nozzles with the time of operation.

Theoretical work on the sources of information reviewing and analysis of various existing specific criteria was performed. Earlier, experimental studies with hydrocarbon fuel were also conducted, which proved one more time that thermal precipitation formation in the fuel supplying ducts was one of the main factors of the jet engine operation effectiveness reduction and its thrust characteristics.

The results of this research consist in – developing and subsequent pending of the novel inventions with the methods of prevention and control of thermal precipitation formation:

– creating the plot of the thrust decay of the jet engine depending on the degree of nozzles carbonization;

– obtaining new specific parameters of the jet engines qualitative analysis in dependence of nozzles operability.

The scope of the research findings application includes diagnostics of both military and civil aviation jet engines; broadening the technique for complex and qualitative analysis of jet engines with the best engine scheme selection; scientific research for the purpose of creating effective monitoring system for the nozzles failure both on the ground and in the air and space.

At present, the problem of thermal deposits occurring on the walls of the fuel-feeding ducts, nozzles and sprayers is still staying unsolved. There is no complete theory of the thermal precipitations formation. The same relates to the complete theory of the thrust reduction of the jet engine due to the thermal deposits and failure of nozzles, filters and sprayers. It is worth mentioning that the existing parameters, characterizing the quality and perfection of jet engines, such as specific thrust, specific mass etc. do not account for the degree of nozzles carbonization with their possible failure. Application of new specific parameters, such as parameters presented in the article, is necessary for the purpose of more qualitative analysis of the jet engines characteristics.

The article outlines the ways of further theoretical and experimental studies.

As a rule, the state parameters, which changing allows detecting the engine failures, change directly neither while operation, nor while bench testing. Usually, the other combination of parameters, called the status signs, is being measured. These are temperature, pressure, fuel and air consumption, rotor rotation frequency etc. A well-defined combination of state parameters corresponds to each combination status signs. The structural diagram of the developed algorithm for the gas turbine engine monitoring and state diagnostics by thermo-gas-dynamic parameters is being performed by the two stages:

1. Determining the engine gas-air channel serviceability.

The results of the engine bench tests are being loaded to the control unit, which main purpose consists in making decision on the engine state in the «serviceable — non-serviceable» form. In the case when the control unit operation results indicate the serviceable condition of the engine gas-air channel control is being transferred to the algorithm input.

2. Determining the serviceable node of the engine gas-air channel.

The main task of the diagnostics unit consists in identifying the non-serviceable assemblies of the engine with the specified probability and computing the state parameters corresponding to them. After printing the diagnostics message, control is being transferred again to the algorithm input, and the monitoring and diagnostics process can be continued.

The measured parameters undergo pre-processing according to the technique being employed at the enterprise. After that, computations according to the mathematical model on the same modes are being performed. The algorithm for monitoring and diagnosing of a gas turbine engine state is based on the assumption of the existence of the adequate non-linear mathematical model of the engine under testing, as well as known values of the state parameters and signs of the reference engine in the diagnostics mode.

In the course of tests of the diagnosed engine, the status signs are being determined, while the state parameters are unknown. In the general case, the dependence of the state signs on the state parameters is nonlinear. Thus, the linear models have to be obtained on a number of basic modes, bearing in mind that deviations from the given mode when using such models are possible within 10%.

The sections of dynamically non-stabilized flows characteristic for the elements of air-gas channels of the turbopump units of the liquid rocket engines are being under consideration. Sector of variable cylindrical and rectangular cross-section, rotational flows in the cavities with immovable walls, and with immovable and rotating walls are studied. Inlet and outlet devices, sidelong cavities between rotor and stator, the cavities of hydrodynamic seals, as well as elements of inter-blade channel of the centrifugal pumps and gas turbines relate to the characteristic elements.

Due to the characteristic features of the operating and structural parameters, the initial sections of dynamically non-stabilized flows are predominant in the air-gas channels of the supply units. These sections affect significantly the energy parameters of the unit and thermal exchange processes and, as a consequence, the reliability of structural elements. Both, laminar and turbulent flow modes of the working fluid are being realized in the characteristic elements of the supply systems.

Using methods of the of the spatial boundary layer theory, the characteristic thicknesses of the boundary layer such as dynamic boundary layer thickness, the displacement thickness and momentum loss thickness were determined. Dependencies for determining the flow core velocity, which are necessary for estimating losses depending on the length of characteristic sections were obtained. To determine correctly the energy parameters, the right choice of the friction laws and velocity profiles in the boundary layer, as well as accounting for the initial section are necessary. The obtained dependencies are accounting for the velocity distribution profile in the boundary layer at the characteristic sections for the cases of both laminar and turbulent modes.

The article presents the analysis results of the ions median-energy ions angular and power distribution in the jets of stationary plasma thrusters. The data on the BHT-1500 thruster at the 700 V mode were used for the analysis. The article demonstrates that content of the median-energy ions is about 35% of total ion flow of the jet, and its contribution to the thrust is 25%. Energy specters of the median-energy ions differ greatly at the small and large escape angles. At the small escape angles the number of median-energy increases, and decreases at the large ones.

It is revealed that median-energy ions are being formed in the discharge area, and in the nearest part of the jet. Particles of the background gas do not participate in the processes of their generation, and, therefore, it may be considered that the median-energy ions are ions of the jet, rather than secondary ions being formed under conditions of the test bench. The background pressure effect on the median-energy ions content is insignificant.

Three mechanisms of median-energy ions generation occurring due to collision such as late ionization and further acceleration in the discharge area; charge-exchange and further acceleration in the discharge area, and elastic scattering in the discharge area and in the nearest part of the jet were examined. It was revealed that the median-energy ions formation according to any of the above-mentioned mechanisms was possible only in the areas of local non-uniformity of the electric field and of neutral particles flows. Such non-uniformities can appear near discharge channel walls or due to the cathode asymmetrical position.

The article presents the model of median-energy ions generation due to accelerated ions elastic scattering. Good qualitative agreement with experiments on both angular distribution and ion power spectra was obtained. However, the obtained scattering coefficient of about 40% cannot be substantiated within the framework of this model. In this regard, the presented model can be examined so long only as the working hypothesis. For clarifying the true mechanisms of median-energy ions generation the 3D kinetic model describing processes in the accelerating ducts of the thruster and in the nearest area of the jet, accounting for the cathode position and effect of the residual atmosphere particles of the vacuum chamber, is required. Much more detailed measurements of the fields of the particles and electric field in the direct vicinity to the outlet cross-section of the duct are required as well.

The increased lateral clearance of the toothed gearing leads to shock interaction of the wheels’ teeth, resonance vibrations excitation, tooth harmonics intensity growth and accelerated wear of the teeth lateral surfaces. The conducted studies allowed proposing a number of new diagnostic signs of the lateral gap value. The work was performed based on the analysis of vibration state of the differential gearbox of the NK-12MP turboprop engine. Fourteen engines undergone the refurbishment at the manufacturing plant were being considered. The performed analysis revealed that the following signs could be used as diagnostic signs:

– a series of harmonics, the frequency of the first of which is defined as the product of the rotation speed of the sun gear in reduced motion by the number of satellites and n-dimensional vector from them;

– the RMS deviation of the rotor rotation frequency of the turbocharger and the shaft of the rear air screw (gear box driven shaft), obtained from the corresponding signals of the “standard” tachometric rotor speed sensors;

– subharmonic components with the multiplicities of 0.5 and 1.5 of the sun pinion speed;

– the amplitude modulation depth of tooth harmonic at the intermodulation component;

– frequency modulation index at the frequencies of the first harmonic in absolute motion, the second and the third harmonics in relative motion of the sun pinion and intermodulation components.

The appropriate approximating dependences have been obtained for all diagnostic features, and norms, using the maximum allowable value of the lateral clearance of 0.43 mm, have been set. It was demonstrated on both vibration parameters and signals from the tachometric sensors of the shafts rotation frequency that lateral clearance increasing “sun pinion-satellites” pair led to its decreasing in the “epicycle-satellites” pair. The obtained dependencies are of both linear and highly nonlinear character with the lateral clearance value growth.

All above said allows drawing the following inferences.

1. The performed analysis allowed revealing a number of new diagnostic signs of a lateral gap of a “sun gear-satellites” gearing pair of the differential gearbox of the of turboprop engine.

2. Diagnostic signs from both the signal of vibration transducer and signals from the “standard” tachometric sensors of rear screw shaft and turbine compressor were revealed, which allows performing diagnostics of the lateral clearance value without installing extra sensors on the engine and ensuring this parameter monitoring while operation process.

The article considers the problem of predicting conflicts between aircraft and vehicles at the airfield. According to the ICAO data, the share of moving out of the runway limits and unauthorized entering the runway is about 30% of the total number of aviation accidents, each tenth accident herewith is associated with human casualties.

The existing surveillance aids (surface movement radar, MLAT, ADS-B) and automation systems A-SMGCS of levels 1 and 2 are not capable of ensuring the appropriate prediction of objects movement paths at the aerodrome. To solve this problem, the authors propose equipping all vehicles with special terminals to inform the air traffic controller on the supposed movement path and the movement commence. Using these terminals the drivers indicate the route and time of the movement commence, creating thereby the database on the transport traffic parameters along the aerodrome. The flight management system will perform the function of this terminal onboard the aircraft. On entering the prohibited area, or deviation from movement path a warning signal is issued for both the driver and air traffic controller. If the driver ignored it, the air traffic controller takes actions to prevent the conflict. The movement paths entered in advance allow analyzing the current situation in automatic mode and identifying potential conflicts during the required time interval. Thus, the proposed system will allow ensuring the A-SMGCS automation levels of 3 and 4. The authors suggest employing the MeSH networks for the data transfer, which allow transferring data, video, images, realizing voice communication and the possibility of the network subscribers’ position location. In addition, subscribers will be able to exchange information about their location, which will increase the awareness of drivers and pilots, and allow them taking decisions independently in case of an unexpected situation.

The forecast of the nanosatellite launches in the near future reveals steady growth. The development of technologies for removing spacecraft, exhausted their resources, from the working orbit is an urgent task. Equipping the Cubesat nanosatellites with a retraction device increases launch costs by up to 50%. The structural elements expenses are up to 25%. Thus, the works on studying new materials for the hulls and technologies for their manufacturing to reduce labor intensity are underway. Design of space structural systems is a balance between the weight, strength and rigidity. The standardized housing of the CubeSat module is being developed in accordance with the CubeSat Design Specification rev.13 and has mass-and-size limitations and rigidity requirements. The most common housing materials are Al 6082 or Al 7075 alloys. The UPSat composite structure from T300-5208 Carbon Hexcel unidirectional epoxy for the first Greek CubeSat is also known. Our work employs selective laser melting technology to manufacture the housing of the 1U module of CubeSats nanosatellites. When comparing the the three housings of the 1U volume, manufactured from these three materials, the lightest one is the housing made of composite material T300-5208. Its weight is 104.5 g versus 155 g obtained from an aluminum alloy 7075. The housing fabricated by the laser sintering is the heaviest, 216 g. However, the mass can be comparable with the composite version by reducing the wall thickness or growing a «mesh» structure. Parts from the ASP-40 AlSi10Mg powder alloy will be two times worse by the mechanical strength than aluminum ones. The specific strength of the unidirectional carbon fiber, compared with aluminum, is six times higher along the fibers. In the transverse directions, the properties of carbon fiber are lower by the order of magnitude.

The advantage of the SLM technology consists in the possibility of structural formation of housing and its fasteners for the servicing equipment, which cannot be fabricated by conventional machining. Besides, when developing a housing part, the effect of space radiation can be computed, to increase the wall thickness in the area of its maximum impact. The closed structure with the walls thickness of 1.8 mm enhances many times protection from the space radiation, which will increase electronic elements resource and the term of the nano-satellite active life.

The presented work deals with the problem of designing the aircraft electromechanical actuators of a translational type. The object of the study is a ball-and-screw transmission with separator, which presence in the structure ensures advanced reliability and stipulates less production costs due to the absence of the internal thread and a unit for balls spillover inside the nut.

It is well-known that in the ball-and-screw transmissions with recirculation of rolling bodies the unevenness of load distribution among the rolling bodies takes place, and the value of the load distribution unevenness ratio depends on the thread parameters.

The presented work proposes analytical determining of the load distribution unevenness ratio in the ball-and-screw transmissions with separator. The equation of the screw and separator turns deformation compatibility was compiled, which solution allowed obtaining analytical dependencies of the load distribution function along the screw helical centerline of the ball-and-screw transmission with separator.

The effect of such design parameters of the transmission as the number of turns and the width of the separator wall on the unevenness of the load distribution was studied. It was established that this transmission had the largest value of the load distribution unevenness ratio at the maximum possible thickness of the separator, and the load distribution unevenness increased with the number of turns increasing.

Based on the results obtained, the technique for calculating design parameters of the ball-and – crew transmission with separator was refined.

Application of ln parameter, characterizing the number of working turns and accounting for the load distribution unevenness ratio was proposed for engineering calculations.

It was demonstrated that while design parameters selecting of the ball-and-screw transmission with separator, the required loading capacity is achieved by both the nut length increasing at the small diameter of the screw and increasing diameter of the screw with the short nut.

In the simplest case, any solid or liquid substance consists of atoms of the same type. A surface atom can have fr om three to nine nearest neighbors, and accordingly its energy increases by the amount, proportional to the number of missing bonds, compared to an atom inside the lattice. By virtue of this, the energy of the atoms on the surface is greater than the energy of the atoms inside the lattice. Thus, a certain excess of energy must be associated with the crystal surface, depending on the structure of this surface and called a surface tension, or surface energy.

According to L.D. Landau and A.Ya. Hohstein opinion, the surface tension is a tangential force applied to a unit length of the contour, limiting a certain area of the interface, and tended to deform a solid. Thus, the surface tension should affect the mechanical properties of the material.

The presented article proposes a dimensionless criterion Χ, characterizing the surface tension contribution to the strength of a solid:

where σ_s is the surface tension, N/m, σ_y is the conventional volumetric yield stress of the solid material, and MPa; h is the thickness of a sample in the form of a strip (foil) of a solid. The value Χ = 1 determines the critical thickness h_cr of the material sample at which contribution of the surface tension to the tensilestrength of the sample becomes equal to the contribution of the bulk yield strength.

The CEM of the samples were also being compared in this work with and without accounting for the surface tension. Mechanical properties of aluminum alloy were studied with the MSC.Marc software based on the finite element method. The total number of elements was 20 thousand pieces. The finite elements represented identical parallelepipeds with eight nodes and eight integration points, which allowed solve volumetric problem with small plastic deformations. The properties of the ADT aluminum in the annealed state were being set to the models.

The obtained series of CEM samples with various thicknesses, with constant length and width, were subjected to the uniaxial tension with forces causing a stress of 50 MPa, which exceeded the bulk yield stress for this alloy, but did not exceed its tensile strength. Thus, the surface tension impact on the mechanical properties of sample models was determined, which confirmed the fact that a significant contribution of surface tension forces was observed only on samples of small thickness, comparable to the critical one.

As the result of the study, simplified equations, accounting for the surface tension forces acting only in the direction, opposing tension, for determining geometric parameters of the samples at which the influence of surface tension forces was comparable with the bulk yield strength of the material, were derived. Based on the derived dependencies for the aluminum alloy, the critical thickness of the sample was determined equal to 73.5 nm.

The results of this study allow accounting for other factors impact, such as temperature, pressure, surfactant, etc., by accounting for their effect on the surface tension magnitude.

Effectiveness increasing is one the basic trends of small-size gas-turbine engines (SGTE) refinement. One of the most affordable and effective techniques for SGTE gain performance is heat regeneration application [13, 19]. In this case, heat exchanger affects significantly the engine effectiveness.

In the event of a plate heat exchanger application in the SGTE of a complex cycle, the heat exchanging surface geometry, ensuring the best heat exchange efficiency, is being selected individually for each task [7, 17]. In this respect, the heat exchanger design technique, allowing obtaining the best thermal and hydraulic characteristics [15], is of primary importance.

Despite heat exchangers designing and calculating complexity, manufacturing stage causes most difficulties while the product creation. The key stages of heat exchanger manufacturing imply dealing with thin-walled and various-thickness parts made of heat-resistant steels.

Analysis of the existing manufacturing technologies has shown that the most effective way of working with such parts is laser cutting and welding on a low-power installation [8]. To perform individual operations on a laser installation, a set of special technological equipment that allows the parts positioning is required [16].

Parts cutting and welding operations in the heat exchanger manufacturing process were performed with low-power “Bulat HTS Portal-300” laser plant with numerical control [18]. The installation low power (up to 300 Watts) allows working with thin details Experimental study of the of the cutting mode effect on the parts edges quality, performed at a low-power laser installation with numerical control, revealed a number of features. The factors exercising the maximal impact on the cut quality are the air supply pressure, pulse energy, frequency, and cutting speed. Modes, ensuring the high quality of laser cutting, were obtained while the experimental heat exchanger manufacturing process.

Specifics of laser welding application on a low-power machine tool with numerical control while thin-walled and various-thickness parts connecting were studied. The pulse shape and spot size are the most important factors while welding modes selection to obtain qualitative joint. The pulse shape variation allows the most rational distribution of energy flow over the time of the thermal exposure. Laser welding modes, ensuring the qualitative pressure-proof weld seam, were obtained in the process of thin-walled and various-thickness parts connection.

While an experimental heat exchanger fabrication it was found that for laser cutting and high-level welding operations performing ensuring, special technological rigging application, ensuring positioning of the machined parts, was necessary.

Experimental heat exchanger was manufactured employing laser technology on a low-power laser installation with numerical control. The heat exchanger experimental studies confirmed the strength and tightness of the welded joints, as well as demonstrated a reliable match of the calculated and experimental characteristics.

The modern approach to the aircraft engine analysis as a part of an aircraft requires the presence of a perfect technique for the thermo-gas dynamic calculation (and such techniques do exist) and a mathematical model of GTE mass (based on the parametrical dependences, based on statistics of the already created GTEs). Considering the last circumstance, the assertion that such models need periodic updating is possible.

It is expedient for the turboprop engine mass models to present the equation of mass in the form of the sum of the gas turbine engine and the gearbox masses. The gas turbine engine mass should be expressed in the form dependency on the working process parameters (G_air,π_c,T_g).

This is explained by the fact that the gearbox mass does not depend on the working process parameters, and it is better to consider it by separate dependencies.

For the MGTE dependency actualization, the basic specification data on twenty three turboprops, such as G_air,π_c ,T_g, and a certification year were used.

Coefficients B, m₁ and m₂ were refined and corrected with the algorithm, proposed in the article.

Linear dependencies of m₁ on the airflow rate, and m₂ on the of pressure increase degree were obtained. To refine the k_Tg coefficient, which accounts for the temperature T_g impact on the engine mass, turbine models were developed, in which the structure being changed with temperature T_g. The corresponding elements of the turbine cooling system were being added, and the mass changed accordingly. This change was expressed by an approximating expression for kT_g.                                                                             

By approximation of integral quantitative values of  and assuming 1999 as a basic year, the expression for k_imp was obtained. This coefficient characterizes the of engine mass improving by the structural and technological solutions introduction.

The performed improvement of the parametric model equation of the turboprop mass allowed reducing its calculation error by 10%.

Icing is one of the most dangerous environmental impacts on an aircraft. Ice bodies on the wing surfaces and empennage change their shape and contours, worsen aerodynamic characteristics, as well as increase aircraft weight. In case of icing not only the aircraft drag increases but the value of the maximum lift coefficient significantly decreases. Various anti-icing systems are employed to remove the ice that builds up in flight. However, practically all these systems have their drawbacks. Application of the wing boundary layer control (BLC) by tangential air jet blow-out on the wing upper surface is known to be one of the most effective techniques for the wing lifting properties at the takeoff-landing modes. The wing lifting properties enhancement occurs due to elimination of the flow separation on the deflected flap by the tangential blow-out of the compressed air jet and flow circulation enhancement on the wing. The hot compressed air for the BLC is drawn from the engine and then piped to the slot nozzles system to be blown-out on the wing surface.

These pipelines are similar to those of the thermal ice-protection system, usually placed along the leading edge of the wing. Thus, the BLC can be employed also to protect against the wing icing. A significant drawback of the above said technical solutions is the jet blowing slot location in an ice sticking area. It is assumed that the hot air from the engine would melt this ice at a certain time instant, but until this moment, the aerodynamic characteristics of the aircraft will degrade. In addition, water evolved while the ice melting on the leading edge, flowing down along the flow is stiffens again out of the BLC coverage forming the so-called “barrier ice”, which also deteriorates the aircraft characteristics. The presented article explores the possibility of the tangential jet blow-out on the leading edge of the wing section to reduce deleterious effect of icing. Calculations were performed employing the program based on numerical solution of Reynolds–averaged Navier-Stokes equations. A case with the horn-like ice on the wing leading edge was under consideration. Comparison of the obtained results with experimental data was performed. The article emonstrates that tangential jet blow-out under of icing conditions allows restoring aerodynamic characteristics level to prior-to-icing state, including coefficients of lift and pitching moment. Specifics of spatial flow-around of the wing section in icing conditions when employing tangential jet blowing-out are presented.

The article discusses the distinctive features of helicopter airfoils flow-around generating integral criteria of their aerodynamic perfection. It demonstrates the importance of the concept of helicopter aerodynamic airfoils and its role in the system, including all cycles of aerodynamic configuration development of rotor blades from the objective function definition up to the elaboration (based on calculation and experimental studies) of recommendations for industrial application. The authors suggest a new approach to comparing experimental helicopter airfoils performance by to the three integral criteria.

The article describes a systematic approach to the development of TsAGI helicopter airfoils for aerodynamic configuration of rotor blades based on the calculation and experimental system. This system empooys the qualitative relationship between the objective vector of main rotor aerodynamic performance and the set of objective vectors of airfoil aerodynamic performance, which allows developing prospective helicopter airfoils for main and tail rotor blades for multipurpose helicopter based on the aerodynamic design procedure. The features of the complex procedure of aerodynamic design of helicopter airfoils used in TsAGI, and its main structural elements are under discussion. Quantitative relationship establishing of the main rotor performance vector and the airfoils performance vectors is performed at the stage of experimental studies of new aerodynamic configurations on large-scale models of the main rotor in wind tunnels. Some results of such

kind of studies are presented on the example of comparing conventional and perspective rotor configurations.

Experiments in the wind tunnel and flight tests confirm the effectiveness of the application and the need to further developing the new series of TsAGI airfoils designed to create aerodynamic configurations of the main rotor blades of modified and prospective helicopters with improved aerodynamic performance.

Based on the TsAGI calculation and experimental system, the article suggests new aerodynamic airfoil configurations of modified and perspective main and tail rotors of domestic helicopters. In particular, the TsAGI developments and their implementation in the design of the blades of the experimental main rotor at Mil Moscow Helicopter Plant allowed reaching record flight speeds of the helicopter — the flying laboratory of the classic single-rotor scheme (without wing and additional propulsive devices).

Flight safety in drastic meteorological conditions remains an extremely important task to this day. With the advent of high-performance computing software, allowing perform simulation of complex physical phenomena with plausible degree of accuracy, a wide spectrum of research trends, helping specialists all over the world study in most detail those phenomena, which could studied earlier by performing the full-scale experiment, is being opened.

The topic of the presented work is the surface wrapping method (SWM method) adaptation to increase modeling quality of the aircraft icing processes to predict more accurately the places, shape and size of ice deposits for further activities on the anti-icing systems design and testing techniques, including certification ones, development.

The essence of this method consists in transforming created mesh surface to the area of the target object. The original mesh may be of a uniform structure with the same distances between nodes, or an adaptive one with dimensions that are a function of the curvature and characteristic dimensions of the object body. The SWM method mathematical model can be based on nodes displacement along the normal to the target object, or on minimizing the function of the node displacement energy. The resulting offset nodes are used for the object surface mesh restructuring, and building volume elements in the entire area in totality In the framework of icing numerical modeling, elements elongation due to the large curvature of the ice, often inherent in the “glassy” type, may lead at a certain moment to the mesh zone overlapping, formation of closed volumes, elements “degeneration” and other defects. Thus, this method algorithm is supplemented by modifying the separation of the low-quality mesh element into several ones, and preliminary diagnostics of the sharp “peaks” presence, point contact of cells and nodes and determination of macro cavities with their coordinates derivation As the result of the suggested method application, the authors managed to obtain complex shapes of the ice buildups much more closer to the experimental data compared to the conventional smoothing techniques, employed in the majority of computing software.

The above described approach application brings prediction quality of the shape and size of ice deposits to the new level, especially on the thin elements of blades profiles and guide vanes, as well as under icing conditions, when buildups of rather complex shape might occur, including air inclusions inside as well.

Acoustic comfort ensuring for passengers and cockpit personnel is one of the most important tasks while civil aircraft design. Particularly, at present there is a problem of the Russian civil aviation flight crewmembers diminished hearing. The risk factor for this disease developing is the noise in the cockpit.

The problem solution of ensuring acoustic comfort in the aircraft cabin is impossible without performing a complex of engineering and fundamental studies at all stages of creating a new sample of aeronautical engineering. One of the research trends is identification, localization and ranking the main noise sources in the aircraft-prototype cabin. The results of this study are necessary for ensuring optimal placement of sound insulation, sound absorbing and vibration damping materials in the onboard structure and issuing recommendations for noise reduction of the air conditioning system (ACS).

The article presents the results of noise sources localization and ranking by intensity in the cockpit of the RRJ-95 aircraft employing the Simcenter Solid Sphere 3DCAM54 spherical array.

Acoustic measurements were performed on the RRJ-95 experimental aircraft No. 95005 with a cockpit modified from the viewpoint of noise reduction and reverberation interference. The tests were carried out at a cruising flight mode at the altitude of 11 km with a flight speed determined by the Mach number of 0.8. The signal recording time was no less than 60 seconds. The measurements were performed while normal ACS operation, and when it was switched off.

As the result of the study, noise sources localization charts in the one-third octave frequency bands of 630-3150 Hz were obtained. The main noise sources in the cockpit are the ACS and the turbulent boundary layer noise. As far as the air-feeding ceases with the ACS turning-off, but the system fans do not, the ACS effect manifests itself with its turning-off from the side of the air supplying pipelines to the cockpit as well. Two basic mechanisms in the ACS noise can be outlined. They are turbulent flow noise in the air ducts, and the noise caused by the “rotor-with-stator” interaction in the fans. In the one-third octave frequency bands of 1000 Hz, in particular, the noise of turbulent flow dominates the noise caused by the “rotor-with-stator” interaction in the ACS fans, while the noise of the “rotor-with-stator” interaction is dominating in the one-third octave frequency bands of 2500 Hz.

In the last twenty years, durability computing techniques with account for local elastic-plastic strain-stress state have achieved a status of the Industry Standard while producing aviation, automotive, cargo and earth moving equipment all over the world. Although the fundamental concepts of this approach are quite simple, the large-scale automation and this technique application for strength calculation of both large dynamically loaded structures and machines driving gears led, in one hand, to the new possibilities emergence for engineers, but, on the other hand, they created extra challenges for the designers of the durability evaluation software. Presently, there is a possibility of dynamic models application for aviation structure loading computing, finite element models, allowing compute local strains by the applied loads, and techniques for more accurate plasticity computing for damageability estimation.

The article considers one of the methods for solving the elastic-plastic problem at cycle-by-cycle calculation, which can be applied for the durability evaluation with account for non-linear effects of interaction of loads of various values, especially after rare loads of high values. The need for analytical methods for elastic-plastic stresses computation developing and improving is caused by high labor intensity and low computing speed through numerical methods, such as finite element method.

The article proposes a new approximate formula for determining elastic plastic stresses and strains at the point of failure. The proposed approach is based on the solution of the elastic-plastic problem by the finite element method for the static case, as well as the method developed by the authors for fitting the static and cyclic stress-strain curves based on standard constants and the Masing principle. The suggested formula for determining the dependencies of local stresses and strains on nominal stresses for typical concentrators provides the necessary dependencies, close in accuracy to the results determined by the finite element method. This formula application will allow developing effective methods for durability computing based on local elastic-plastic stresses and strains under multi-cycle loading, being typical for aircraft structures.

The article presents comparisons of local stresses dependencies at the most stressed points on nominal stresses, obtained with the proposed formula and the finite element method for typical stress concentrators of the aircraft structure such as strips with free hole, fillet, and stringer runout.

This article presents the results of a thermal model developing and application of a cathode with Barium emitter for the temperature field computing, determining internal and external conductive and radiative heat fluxes, gradients and velocities of temperature changing in the cathode stationary and dynamic operation modes, as well as heat release computing on the cathode emitter. Based on the computational results of the thermal state of the cathode design elements in functioning modes, the analysis of the cathode design and start parameters, which ensure meeting the thermal requirements to its main elements, was performed.

The objectives of the above said thermal computations were:

– determining a minimum power for the cathode pre-start heating, which ensures conditions of reaching emitter temperature within 160 sec (the level sufficient to ignite and maintain the discharge),

– estimating temperature distribution by the cathode elements at various boundary conditions and verifying the thermal model based on the thermal vacuum tests results to employ the model for determining the cathode structure thermal state at various boundary conditions.

The task of the thermal calculation was elements thermal state estimation of the cathode with Barium thermo-emitter in the start heating mode and in the automatic mode (which means the cathode operation when thermo-emitter temperature is maintained by bombarding by the ions of the working body. The discharge circuit between the anode and cathode herewith is closed, and the source of the external heating (heater) is turned-off by way of determining the estimated range and thermal flows over the cathode elements A 3D thermal model of a cathode with Barium emitter was developed with SolidWorks Flow Simulation 2014, which employs the finite volume method, i.e. a numerical method for integration of differential equation systems in partial derivatives. Boundary conditions for the thermal design were being set identical to the thermal vacuum test conditions.

The following elements were set in the model: parts geometrical sizes (with insignificant simplifications not affecting the temperature distribution), structural materials properties and contact thermal resistances between the model areas. The calculation accounted for only conductive and radiative heat exchange, since cathode operation conditions as a part of the thruster represent a deep vacuum. A power, corresponding to the operation mode, was set on the heat releasing elements of the cathode thermal model depending on time and operation mode. When calculating a radiative component of heat exchange, integral emissivity factor was assigned to each surface, depending on material and surface treatment class.

Anisotropic thermal conductivity was set in the ceramic parts properties, i.e. thermal conductivity of Aluminum oxide ceramics is two-directional. Direction of axial (transversal) and radial thermal conductivity ws determined along the corresponding axis of the coordinate system. A temperature dependence between the thermal conductivity coefficient and thermal capacity was accounted for in structural materials properties.

Experimental data obtained at EDB Fakel facility from thermal vacuum tests of a cathode with Barium emitter was employed for the thermal design model verification. The thermal model verification consisted in heaters power selection and heat release on the emitter from the condition that the temperature calculated values in the checkpoints coincide with the measured ones.

Based on the thermal design results, a minimum heater power for guaranteed start of the cathode with Barium emitter was selected.

Cathode thermal model verification with the thermal vacuum test results was carried out. This allows the cathode thermal model application for predicting a thermal state of the cathode structure while numerical reproduction of situations, which were not verified while physical experiment, as well as compare the temperature predictions with the temperatures registered in flight.

Turbofan engines (TFE) with common afterburning chamber are the basic ones applied in power plants of maneuverable aircraft both in our and foreign countries. Recently, the TFE with low bypass ratio (no more than 0.3 ... 0.5) that has a certain feature in the scheme of mixing and burning processes in the afterburning chamber are most widely spread.

The absence of special mixing devices at the afterburning chamber inlet in a number of TFEs structures may lead to the situation when a certain air portion of the second duct would not admix to the main flow and participate in combustion process even at the full speed-up.

In this case, rather high values of total excess air factor ( α_Σ ≥1,2 ... 1,3) realize at the afterburning chamber outlet, which may eventually reduce the engine speed-up degree at these modes.

With a view to the specifics of TFE interaction units in the engine system, the share of the air not participating in the air burning process at various operating modes may change in a rather wide range.

The estimation inaccuracy of this value can eventually lead to essential errors in determining the main TFE parameters sucha as its thrust and specific fuel consumption.

A specially developed model of the stage-by-stage air of the bypass duct admixing to the main flow at the afterburning chamber inlet was integrated into the general mathematical model of the engine and allowed refine both working process in the TFE and its characteristics at the speed-up modes.

The following scheme of the afterburning chamber of the two-stage successive air admixing of the bypass duct air to the main loop flow is assumed while mixing-afterburning chamber modelling. At the first stage, the entire gas of the internal loop and a fraction of air of the second loop participate in mixing. At the second stage, the remaining airflow, being flown through the subscreen duct, is being admixed to the gas at the afterburning chamber outlet.

Equality of static pressures herewith is assumed in the mixing section, as well as fulfillment of conditions of conservation of mass, energy an impulse for the mixing flows, peculiar to the conditionally full mixing in the conditional cylindrical duct.

Estimation, performed on the example of technical appearance analysis of the fifth generation Pratt & Whitney F119-PW-100 TFE analysis was performed. Its altitude-speed and throttle performances, among all, as a part of the F-22A Raptor aircraft power plant, revealed telling impact of the factor under consideration on both TFE characteristics and the aircraft as a whole.

The presented article studied the aircraft target purpose impact on working process optimal parameters and structural schemes of small-sized gas turbine engines (GTE).

The engine optimization was performed as a part of the aircraft system. Total weight of the fuel and power plant and the fuel, required for flight, as well as specific fuel consumption of the aircraft per ton-kilometer were being used as functions of the GTE efficiency. The aircraft of light, administrative and regional types was considered. Commercial loading weight (the number of passengers), flight range and trajectory were set for each of the aircraft under consideration.

The database of possible structural schemes of the engines was formed based on the initial data. Further, the engine evaluation criteria in the aircraft system were being computed. Minimax method of optimization was employed for rational solution obtaining. With this, functional limitations for the engine of each scheme were accounted for while optimization. Optimization of small-sized gas turbine engine in the aircraft system was performed with “ASTRA” CAE system.

The optimization results are presented in the form of dependencies of optimal of working process parameters of a small-size GTE on flight range for the aircraft under consideration. The studies revealed that with the flight range increase, the degree of bypass ratio and total degree of pressure ratio increased, the degree of pressure ratio in the fan decreased, and the gas temperature prior to the turbine changes insignificantly. It was found that with the engine size increase, the flight range exerted relatively slight impact on the working process optimal parameters. With the flight range increase, optimal parameters values by various criteria tend to minimax solution for any engine scheme.

The presented study demonstrated that the target purpose of the aircraft significantly affects the optimal parameters of the the power plant working process with the small-size GTE. In return, the working process parameters and the engine size determine its most rational design scheme.

The flying vehicle with ducted rocket engine (DRE) developing at the preliminary design stage begins with forming the volume-weight layout of the product. Then, determining geometry of both engine characteristic sections and aerodynamic surfaces is required. These issues can be solved by tuning and optimizing with special software. The result of these studies represents the entire range of various technical information, such as characteristics of the DRE elements and, consequently, the engine altitude and speed characteristics, the airframe aerodynamic characteristics, flight dynamics parameters according to the technical specifications and, surely, preliminary size of the airframe and DRF basic elements. This allows making a drawing of all three views. However, further studies of the thermal state, aerodynamic and strength characteristics require a 3D-model.

To solve such problem, it is effective to employ automated design systems, since their capabilities are noticeably superior to human ones. Analysis of the software products available on the market (KOMPAS-3D; SolidWorks; Autodesk Inventor and others) revealed that practically there were no holistic tools for solving these problems at the moment.

At present, automated design, systems are employed for converting drawings into electronic form. Initially, a 3D-model is created manually according to the paper drawings, and the original drawings are already being recreated from it, but in the electronic form. Reducing the time interval from appearing the drawing of three views of the preliminary studies to the 3D-model is required for the studies simplifying and conceptual flaws revealing. Thus, creation of the unified program for real objects modelling presents great scientific and practical interest.

Such program can be obtained, combining the initial software package with one of the automated design systems. Thus, the possibility of immediate transition from the drawing of three views to the 3D- model will appear. Such program advent will significantly accelerate the process of 3D-models creation, which, in its turn will allow immediate conceptual flaws revealing and accelerate various kinds of studies.

Rocket engine energy performance improvement is an actual task for modern researchers. The article considers rocket solid propellant engines, which distinctive feature consists in the recessed nozzle.

Recommendations on designing the inlet sections geometry of the recessed nozzles are few and inconsistent. The purpose of the presented work is studying the nozzle inlet shape effect on the flow-rate characteristics and developing appropriate recommendations on nozzle designing.

The flow coefficient is one of perfection indicators of the flow processes. Advanced methods of computational fluid dynamics (CFD) were employed for studying the flow coefficient of the recessed nozzles. The problem was being considered in quasi-stationary axisymmetric adiabatic approximation of the ideal gaseous setting. The RNG k- å two-parameter turbulence model with standard set of model constants, being passed verification while computing classic nozzles consumption and the specific impulse losses of the recessed nozzle, was employed for the flow structure modelling.

The computational geometrical model contained:

– combustion chamber;

– charging duct, from which surface the working medium was being supplied;

– various options of the nozzle recessed part shapes;

– the conical expanding part;

– as well as extra volume behind the nozzle cut.

The grid quality maintained constant while varying the recessed part sizes, and the nozzle degree of submergence.

The gas dynamic component of the flow coefficient was being studied. Nozzle inlet geometry formed by ellipse and by Vitoshinsky curve were being examined. The dependences of the flow coefficient on the nozzle inlet shape and degree of submergence coefficient were obtained.

The results of the flow characteristics of the inlet sections under study are being compared with the previously obtained results for the radius inlet. It was demonstrated that the best values of the flow coefficient corresponded to the inlet section formed by the Vitoshinsky curve. The distinctive feature of the inlet section designed by the Vitoshinsky equation is high stability of the gas-dynamic losses irrespective of its geometrical parameters changes.

Elliptical inlet nozzles allow improving flow coefficients indicators even for the worst option of the radius nozzles by 7%. The article presents recommendations on designing the inlet section of the recessed nozzle.

In the last few years the problems emerge in electrically driven pump unit (EDPU), which disrupt operation of the spacecraft thermo-regulating system (TRS) and disabling EDPU. The EDPU service life and operability depend greatly on the operability of rotor supports, sealing system efficiency, and required lubricating and cooling mode. As a rule, seals and supports are connected with the pump flowing part, and they are connected between each other by the hydraulic path, necessary for the unit normal operation. Large axial loading occurrence is considered the most probable cause of the EDPU failure. Thus, studying hydrodynamics of such auxiliary hydraulic paths is the paramount objective for the enterprises working in this field. For these problems solving, a 3D mathematical model of the working fluid flow in the impeller cavity of the EDPU being studied was developed.

To validate the computational model, hydraulic test bench was assembled, and special hydrodynamic tests of the EDPU under study were performed. The pressure changing behavior in various areas obtained by the tests coincides with the CFX computation, and the error does not exceed 3%.

The pressure force change in the axial clearance along the radius submits to the parabolic law, in which the liquid flux coefficient in the axial clearance φ plays an important part. It depends upon the structural and operating parameters of the pump and changes from 0.5 for the lossless flow to 0.76 with expendable flow from periphery to the center in the form of the working fluid leakages. The force acting from the axial clearance side depends on the φ coefficient, though the suggested recommendations are not enough for correct axial force determining.

To determine the fluid flow rate in the axial clearance, the axial force, obtained with software complex, was being used. The values of the φ coefficient were obtained this way for all modes, tested with the hydraulic test bench. Additional calculations of the EDPU various working modes were performed for the livid illustration of the way the coefficient φ depends on the structural and operating parameters, but without test bench testing since the computational model convergence has been already proved.

The obtained dependencies demonstrate that the φ coefficient depends weakly on the operating parameters, and to the greater extent it depends on structural ones, and more specifically, on the discharge openings diameter. In addition, the range of this parameter changes is wider than it is pointed in the source based on the experimental data, which cannot be always determined precisely due to the structure complexity, and, as a consequence, complex access of measuring devices to the EDPU areas of interest.

Pressure losses computing in the fuel system of the stationary gas turbine engine is an integral part for solving a number of engineering and operational tasks. For example, such calculation is necessary to determine a minimum required gas pressure at the inlet of the engine to ensure the engine reaching its operational modes. Likewise, this calculation may come in handy at the fuel gas composition changing, since gas properties change, which means the pressure loss change too that can require to make changes in control equipment. It is well known that fuel nozzles are carbonized while a combustion chamber operation process. Very often, it leads to the resistance increasing of the fuel system, and therefore the of pressure losses rising. Besides, any discrepancies in the dosing equipment can be detected by a hydraulic calculation.

The article considers a fuel system of a stationary converted aircraft engine intended for driving the gas pumping unit supercharger. The pressure losses computing technique for the fuel system of such engine is presented in the article. A relevance of the topic and the necessity of such techniques forming are disclosed. To check the adequacy of the developed technique the NK-16ST engine rig test was performed with pressure measuring in the fuel supply pipelines to the nozzles and in the gas doser. The results of the studies revealed that the gas fuel pressure level measured in the eight gas-extraction from collector to nozzles pipelines differed insignificantly, which confirmed the fuel distribution uniformity along the pipelines. Experimental results comparison with the computational studies confirms that their discrepancy does not exceed 6%.

The article considers artificial neural networks employed for sporting aircraft maneuvers computing method developing. System approach, describing it in the form of the MPL neuron network, is used for representation of such network. As long as initial training data represent complex functional dependencies with the number of variables greater than two, conventional approximation methods application is complicated. Thus, neural network modelling was employed for the problem solution. The concept of neuron represents the basis of neuron representation of aircraft flight trajectories (in the context of movement determining for an AIRCRAFT, and in the context of detecting and tracking devices). Correction of the MPL network architecture structure means the number of hidden layers and neurons (nodes) in each layer. Activation functions for each level are selected at this stage as well, i.e. they are assumed to be known. Weights and deflections are the unknown parameters with should be evaluated. Whereas excitations from the other neurons are fed to the input. For practical implementation of this approach a mathematical model of the Yak-55M sporting aircraft kind was developed on the X-Plane flight simulator using an algorithm of the training cycle of the network of multi-layer perceptron. The article presents also simulation results of the set problem on computing the safe parameters of a sporting aircraft maneuver starting. The study demonstrates that the neural network properties, such as non-linearity and good generalization ability, enhance its ability for complex tasks learning and can produce correct result for new initial data. The aircraft under analysis is out of effective system for collisions with ground prevention based on the predicted course of evasive manoeuver. However, the problem can be solved by developing relationship between the piloting errors and flight safety, and employing neuron network modelling for a number of maneuvers, which associate velocity and altitude parameters and automatically compared with the preset values. The model demonstrated the results of the sporting aircraft maneuvering starting parameters computing. With this, the probability of reliability of a great number of maneuvers should correspond to the reality. The results obtained while mathematical modelling should be loaded to the warning system to warn the pilot on the maneuver performing at the inappropriate altitude, and offer the recovery from the manoeuver allowing secure the flight and minimize hazardous situation.

This article is devoted to the topical issue of applying the autorotation phenomenon in emergencies while helicopter engine malfunctioning to ensure safe landing. In the beginning of the article, the basic theoretical data on the physics of the helicopter rotor autorotation process is presented, and the conditions for the occurrence of a stable autorotation mode are considered. The objective of the overrunning clutch is described on the example of the MI-8 helicopter. Further, the characteristic sets of initial conditions and spatial zones of the autorotation commence are considered, staying in which ensures or does not ensure a safe landing. It was emphasized that the key for the correct entry performing into autorotation is maintaining the rotor rotations. Two techniques for the rotor speed drop terminating in emergencies are presented. Besides, the article considers the pilot’s actions in case of an emergency associated with engine malfunctions in Mi-8, 24, 28 helicopters, ensuring stable autorotation mode and a safe landing. Based on the results of a series of field tests, a scientific substantiation was also presented for the main parameters selection, allowing the helicopter landing with idle engines, as well as recommended landing profile for the rotor self-rotation was elaborated. By the results of processing of video recording of ten landings, the values of the height of the helicopter pitch increasing commence are presented. The pitch angle value and height, at which this pitch angle was reached, as well as vertical and horizontal components of landing velocities are presented as well. In conclusion, the landing technique while autorotation mode performing, formed as the result of flight test data analysis with the listing numeric parameters of the flight is presented.

This work presents a spherical mechanism with four links, interconnected by rotational kinematic pairs. Based on the spherical mechanism by the type of a crank-and-rocker mechanism, a simulator, shown on the figures of the article, is being developed. While the mechanism kinematics analysis we create the structural scheme with notations of its sides and links. For the convenience, and simplified and advantageous computation it was determined that the opposite situated links would have equal angles of crossing. Two techniques are employed while spherical mechanism computing. The first technique consists in developing a mechanism mathematical model. Additional angles and hinges points of a mechanism, which will be employed in subsequent computations, are accounted for while the mathematical model developing. Since we use two techniques of comparison, we equate both techniques through the same speed and rotation time. Having performed kinematic computation, we specify the complete revolution of rotation of the mechanism. During equations analysis we make to the conclusion that with complete revolution of the leading link the driven links manage performing one-half revolution. While graphs plotting of angular speeds and accelerations maximum and minimum points can be observed. Likewise, the angular acceleration increases three-fold from the angular acceleration. For the complete pattern of computations, we perform analysis of moment of inertia of the mechanism connecting rod, which will be the capsule of the simulator. The centrifugal moment of inertia through the point, located at the center of the connecting rod lengthwise the direction cosines, was obtained for the mathematical model. The moment of inertia of leading and driven links is determined by the simpler technique. For precise computations, the displacement angle of the connecting rod relative to the driven crank in hinges are obtained. The angle of rotation of both connecting rod and its center point on each axis separately is being determined. For convenience, the values of mass and radius are set as constants. In the future, we shall set these values from the definite task of the mechanism. Having plotted the graph of the connecting rod moment of inertia by the two techniques, we obtain several maximum points of loads.

Thermal protective coatings are the type of coatings employed to insulate components operating at elevated temperatures. Application area of such coatings is the gas turbine engine blades, combustion chamber, nozzle guide apparatus and pipelines. Thermal protective coatings allow increase gas turbines temperature, enhancing thereby the turbine efficiency.

In conditions of high-temperature operation, special requirements are imposed on components of gas turbine engines. In this regard, thermal barrier coatings (TBC) were developed to protect the gas turbine elements, representing a system of the two or more layers applied on a substrate in a special way.

Coatings, obtained by the electric arc technique of physical vapor deposition (EAPVD), were selected for studying in this work. Three types of alloys were employed for the TBS system, such as SDP-4, representing a coating of NiCoCrAlY alloy; VSDP-16, a diffusion coating of a AlNiY type; and, finally ceramic layer from Zirconium oxide, stabilized by the Yttrium oxide (ZrO₂ + 8% Y₂O₃). Chemical composition of the thermal protective coating was determined by the X-ray micro-analyzer of the Inca Energy OXFORD instruments system. It was determined that after long-term operation the coating layer formed by the SDP-4 and VSDP-16 alloys had two clearly defined zones, such as β-NiAl phase and an inter-diffusion zone, while the NiCoCrAlY alloy did not exhibit phase separation, and the coating structure represents the β-NiAl and γ -phase mixture. It was established that oxygen diffusion occurs outside ceramic upper layer to its boundary with the heat-proof underlayer, which contributes to thermally grown oxide α-Al₂O₃ forming. It was noticed that the VSDP-16 alloy deposited on the SDP-4 layer increases the amount of aluminum in the binder coating layer, compensating its consumption for α-Al₂O₃ forming from the β-NiAl phase.

An important part of aircraft fuselage panels testing for fatigue and survivability is the study of normal deformation fields of buckling and warping of the skin. The article describes optical videogrammetry technique and its application for non-contact measurements of distributed normal fuselage panels’ deformations of a passenger aircraft under testing for internal overpressure.

Strength, reliability and resource of modern aircraft are ensured by multilateral experimental studies of the structural elements behavior under regular and extreme external impacts. One of the types of such tests is the study of aircraft fuselage panels deformation under the impact of internal overpressure. Significant component of deformation herewith refers to the displacement of points in direction of a normal to the surface, i.e. normal deformation. These deformations are of a complex character distributed over the surface. To obtain the full pattern of the distributed normal deformation measurement in a large number of points are required. Strain gauging is a traditional technique for deformation measuring. However, complex deformation fields studying with pattern, which complicates sensors placement and strain gage measurements results interpretation.

At present, contactless optical video-grammetry technique (VGT) manifested itself as a prospective trend for distributed deformations measuring. The results of measurements represent not relative deformation, but normal displacements of the surface points directly. It gives an additional advantage when interpreting the results and comparing them with the calculation or mathematical model of warping and buckling of the skin.

The goal of the presented work consisted in improving contactless optical video-grammetry technique for distributed normal deformations measuring at a large number of the surface points, and this technique application for testing aircraft fuselage panels under internal overpressure.

Video-grammetry technique with one digital camera was chosen (mono-grammetry method) for these measurements. This choice was stipulated by lack of space around the panel being tested on the experimental setup.

When in orbit, the spacecraft is affected by natural sources of radiation (solar and galactic cosmic rays, radiation of the Earth radiation belts) and artificial sources (the onboard radiation sources), which affect the spacecraft in a wide range of energies, penetrate through the eigen external atmosphere (EEA) deep into the structural elements, where particles of energies conversion occurs.

The following energies, affecting a spacecraft, relate to the energies of natural origin:

– solar cosmic rays, including electromagnetic radiation (solar radiation) and corpuscular radiation (solar wind);

– galactic cosmic rays, i.e. isotropic cosmic radiation coming from the interior of the galaxy;

– radiation belts of the Earth, namelly radiation of natural origin, formed by the solar wind and the Earth magnetosphere.

The spacecraft onboard equipment is affected not only by sources of natural origin, but there are also artificial ones situated onboard the spacecraft. Nuclear power plant (NPP) is an example of an artificial source that generates a flow of energy that exceeds all natural impacts by its intensity.

Radiation from natural and artificial sources affects the spaccraft through the medium of its eigen external atmosphere (EEA). Since the EEA is not static, but is constantly mixing as the result of the existing of pressure gradients, temperature, and concentration of activated nuclei and ionized particles of atmospheric substances, the induced radioactivity is being carried over the entire surface of the spacecraft with NPP. Gradients of atmospheric parameters also contribute to medium flows formation that transfer activated nuclei to the shadow area created by the radiation protection unit. The exited nuclei are splitting and their transition to new stable states is accompanied by radiation, which leads to the occurrence of induced radiation on the protected spacecraft structure.

The article deals with the main types of radiation that affect spacecraft with nuclear power plants, and gives their classification. Radiation impact of the onboard reactor, which surpasses solar and galaxy radiation by the intensity, forming basic contribution to the radiation doses, being accumulated by the equipment and structural elements, is the most dangerous for a spacecraft with NPP. The rate of the induced radioactivity propagation in the EEA volume and accumulation of critical dose of radiation in both onboard equipment and structural elements from activated and ionized EEA substance has not been determined at present.

In the existing economic conditions, the service life of a spacecraft with nuclear power plant is set within the range of seven years or more, which requires a complex of works to study and account for the intensity of radiation dose accumulation from the EEA.

The article regards the possibility of regional aircraft takeoff-landing characteristics improvement by employing blow-off from propellers of the auxiliary retractable multi-propeller distributed electro-power installation (DEPI). Its motors operate only during takeoff-landing modes being retracted into the wing while cruising flight. The DEPI motors small-size, commensurable with the flaps chord size, allow deflect the jets from propellers at substantial angles, ensuring herewith significant lift force increase. A large number of the DEPI motors reduces negative impact of any of these engines failure, which leads to the flight safety enhancement. Aerodynamic layout of an aircraft with DEPI as applied to the L410 class aircraft was formed, and calculations of takeoff-landing characteristics with account for the blowing effect were performed. The article demonstrates aerodynamic characteristics dependence on thrust-to-weight ratio, the wing geometric size and propeller diameter. It considers various options of cruise engines total thrust and DEPI motors relationship. It was shown that increasing in the DEPI thrust-to-weight ratio share leads to reduction of the runway length required for the takeoff. Thus, with typical total thrust-to-weight ratio being equal to 0.50, the increase in DEPS thrust from 0 to 25% results in runway length reduction from 780 m to 580 m, i.e. approximately by 25%.

An approach to compliance of Cp_{landing approach} and Cl_{landing approach} values, being realized with account for blowing, with flight-path angle at landing approach was suggested. The article demonstrates the presence of unique dependence between the flight-path angle, required Cp_{landing approach} value and re alized Cl_{landing approach} value.The possibility of realizing higher (approximately twofold) Cl_{landing approach} values due to the blow-off is shown. With typical wing load of 250 kg/m², the blow-off implementation allows required runway length reduction approximately by 20%.

The article presents the numerical research of the interference effect of fuselage and duct of the propeller-duct thruster, and performs evaluation of their impact on the maximum thrust value. It presents the results of the numerical research by means of the program based on numerical solution of averaged by Reynolds Navier-Stokes equations. It demonstrates the pressure and field of velocities change depending on the shape of the fuselage tailpiece and duct-type profile, and their effect on the maximum thrust value. Numerical studies revealed the necessity of such parameters selection as the profile thickness, chord and installation angle of the duct with affect for the flow conditions and interference while a flying vehicle design.

Aerodynamic designing of the optimized duct shape was being performed without changing the external fuselage lines. According to the marked, noted limitations, a new duct-type profile was designed for numerical studies. The opening angle of the duct was being selected based on flow velocities distribution analysis in the duct setting area in such a way that the flow would direct the duct at the angle corresponding to the mode of the maximum quality of the duct profile. The article shows that with the selected velocity of the air flow, the duct profiling changing insignificantly effects it thrust of the propeller itself, but it drastically effects the duct thrust. At this present velocity of the air flow, the rarefication is being observed along the entire internal surface of the duct. The highest rarefication zone occupies up to the 60% of the duct-type profile chord, while it is only 30% with the initial profile.

Thinning-down of a boundary layer and increase in speed in it due to the change of the fuselage shape allows reducing the drag of the fuselage itself. Analysis of the numerical results revealed that at low flight speeds the shape of the fuselage fodder part rather than the duct profile affects the maximum thrust value.

Data analysis of the pressure profile along the internal surface of the duct revealed that rarefication at the internal surface of the duct took the shape of the half-internal distribution, which corresponds to maximum thrust of the propeller-duct thruster.

It is necessary to solve the inverse problem of ensuring half-internal pressure profile along the internal surface of the duct for the defined flight speed while the screw-duct thruster design.

At present, there is an undeniable demand for developing new prospective layouts of various cooling systems and lifting complexes for air-cushion units, in which a flat barrier of substantial size (a screen, air duct, radiator) is being placed behind the axial fan. This problem can be solved by the effect of kinetic energy conversion of the swirling flow behind the impeller into the static pressure, observed in axial- radial diffusors, formed by the fan outlet manifold and a flat barrier, upon which the flow is ingoing. Implementation of such structures of fan installations allows not only preserving high energy-efficiency of the fan installations but as a whole, but significantly reduce their axial size as well.

The main parameter affecting the efficiency of the swirled flow dynamic pressure into static pressure conversion is the flow swirl intensity, characterized by the Rossby number, since with its increase, the total pressure loss in the axial radial diffuser decreases. The article demonstrates that namely fans with the said circulation distribution along the blades length implementation, whereby the flow is swirled by the law of the solid body, is expedient in such kind of fan installations. These fans swirling intensity can reach much higher intensity compared to those, for which classical methodology for the constant circulation is used while aerodynamic design.

Based on the available experimental data on the swirling flow total pressure loss in axial radial diffusers, calculation was performed for aerodynamic parameters of compact fan installations with variable circulation in the wide range of calculated parameters such as flow rate and hub-to-shroud rate, which finally determine the blade shape geometry. According to the obtained results, the installations under consideration can develop rather high for axial fans static pressure rate at a minimum axial size.

An additional analysis of fans with variable circulation revealed two limitations that significantly narrowed the range of design parameters.

The first limitation is stipulated by the criterion of the aerodynamic load limit of blades system, characterized by the value of equivalent diffusion cascade D_eq. Exceeding the D_eq maximum value for peripheral cascades may lead to the high intensive separated flow path of the rotor. Unlike the classical fans with constant circulation, the diffusion cascade criterion for the fans under consideration does not depend on the design parameters, and, eventually, determines the minimum value of the axial velocity, at which this limitation is fulfilled.

The second restriction is determined by the energy balance condition: the total kinetic energy of the flow should not exceed the energy transferred to the flow by the rotor blades. This problem manifests especially pointedly in the near-hub sections, since unlike the fans with the constant circulation, the quantity of energy transferred to the flow by the blades in the fans, which swirl the flow by the solid body law, reduces from shroud to