This article is aimed at finding the causes of controllability variations of a helicopter while transporting sling load.The maximum angular acceleration taken by the helicopter at similar controller displacement at different flight speeds was taken as a quantity characteristic of controllability efficiency to study the load impact on the helicopter cont- rollability.This article offers research results obtained with the use of the НеliСargо software. This software has proven to be a great tool for integrated research of the impact of an underslung load on the parameters of a helicopter controllability, and allows carrying out an analysis of the impact of an underslung load on the parameters of controllability under its dyna- mic behavior.The performed computational experiments have shown that the helicopter maximum angular acceleration with an underslung load significantly rises, as compared to a helicopter without cargo or a helicopter carrying the same load inside the cargo compartment. The data obtained during computational experiments corresponds to the results of analytical computations, and is in line with the literature based on the experience of helicopter flight operations.The cause of the variation in the helicopter controllability parameters during transportation of an underslung load has been found, that is - the underslung load considerably increases the main rotor thrust, due to sling load, as compared to a helicopter without cargo. When compared to a helicopter carrying a load inside the cargo compartment, this increased efficiency is mainly attributed to the fact that a helicopter with an underslung load has lower rotational inertia, since the load is not inside the cargo compartment, but outside.The obtained results can be used to improve flight manuals and flight personnel training publications, which could play a significant part in ensuring flight safety and security, and increasing the operational efficiency of helicopters with external slung load systems.

When a helicopter performs an air operation and aerial work using external sling there may occur special situations, as the result of unfavourable conditions impact. The most difficult and dangerous stages of flight are the takeoff and landing.To study an aircraft flight dynamics in extreme flight conditions, including a helicopter with cargo on external sling, the most effective is to use the theoretical methods, mathematical modeling in particular, which is relatively cheap, able to model the abnormal situation including disasters, and able to set the necessary environ-mental effect.During the theoretical studies it is impractical and impossible to simulate all the abnormal situations that may occur at the given phases of flight. Therefore, at the initial stage of the research it is necessary to deal with making up a list of the most typical abnormal situations - the scheme of simulation cases. This issue is the subject of this article.The article describes the typical errors of piloting a helicopter with cargo on external sling at takeoff and landing, as well as the most dangerous equipment failures and noncalculated operating conditions, based on the analysis of which the scheme of simulation cases has been made. After modeling in accordance with the scheme of simulation cases, it will be possible to determine the effect of the cargo parameters, flight parameters, the impact of unfavourable factors on the dynamics of the system "helicopter - cargo on external sling" at takeoff and landing, which theefore will determine the boundaries of the system safe operation on these phases of flight, and to formulate recommendations for the crew to prevent abnomal situations or to get out of them.For further research it is expected to use the earlier developed HeliCargo software application, which allows to simulate the dynamics of helicopter with cargo on external sling.

Aircraft are high-tech engineering products which are characterized by a wide range of properties including the two most important groups that respectively characterize the efficiency and technical level.Improving the aircraft efficiency is an essential factor for air transport development, but the efficiency can not be fully describe the technical system, especially in forecasting and new technology requirements development. Aircraft designer must evaluate the prospects of a technical solution, but it’s not always possible to calculate the efficiency accuratelyat the design stage. The operator should be able to choose the most technically-advanced aircraft available in the market inorder not to let it grow obsolete quickly. This determines the need for non-economic evaluation of technical systems that can be done by assessment of their technical level.The technical level is a general index that includes a set of technical perfection indicators. Technical perfection is reflected in terms of material and energy intensity, in terms of ergonomics, safety, etc. and is achieved as a result of original design solutions, the use of new high-strength low-density materials, the introduction of advanced technological pro-cesses, calculation methods, verification, testing, etc.There is a tight connection between the product properties and its weight, because weight is the material reflection of these properties. Therefore, improvement of the product properties usually leads to an intense increase of its weight. To deal with this phenomenon is only possible with widely using scientific and technical progress results. In accordance with this, the technical perfection can be interpreted as a major component of quality that is created without the weight increase. This approach requires investment in research and testing new technical solutions.The method was developed to determine the technical level of civil long-haul aircraft which has been modified to incorporate features of general aviation aircraft operation. However, according to the authors of this article, this method requires some clarification. This is the subject of this article where the technical level generalized index equation is refinedand with the use of which the technical level of civil long-range aircraft is determined.

Based on the information from the Internet the authors decided to reproduce the possible conditions of landing the plane IL-76TD-90VD in Antarctica in their computer experiment using the mathematical modeling system of aircraft flight dynamics (MMS AFD), which is developed and continuously improved in MSTUCA. MMS AFD has already shown its ability to complete such tasks for different aircraft types in various conditions, including IL-76TD, IL-96-300, IL-96T and also slippery runways.For the computational experiments the information from the web sources was used. The authors did not aim to produce a final numerical results, which describe the possible operating conditions. It was expected only to get accurate qualitative results, which can adequately describe the behavior of the airplane IL-76TD-90VD when landing on the slippery runway.So, the authors were able to obtain a typical turn of plane at the end of run and show that the pilot style in these conditions has little effect on the qualitative behavior of the aircraft. The most important factor leading to such maneuvers, lies in the special connection of longitudinal and lateral friction of gear wheels with runway.The authors managed to identify the minimum allowed value of standard friction coefficient of the runway (0,42) which allows a safe landing under crosswind 15 m/s.The results that agree qualitatively with the known situations indicate the possibility for MMS AFD to properly reproduce various cases of aircraft IL-76TD-90VD landing to Antarctica ice airfield and identify any inaccurately knownparameters.

The controllability of military transport aircraft deteriorates at heavy single piece landing. To solve this problem and a specific methodology for pilotage of the pre-emption, and automation tools are being developed. Preliminary study ofpilotage technique and authomatic control algorythm demand a reliable mathematical model of aircraft dynamics at cargo item drop. Such model should take into account significant change in the position of the aircraft center of mass and aircraft inertia tensor. Simplified models were based on modeling the movement of the center of mass and rotation around the center of mass of the aircraft. Such models do not take into account the inertial forces and moments of moving a cargo item. This circumstance does not allow to obtain reliable results in the simulation. The article presents the description of the complete mathematical model of the movement of military transport aircraft in landing of a cargo item. Examines the complex material system of solids and a detailed description of the properties of its components. The equations of motion of the aircraft as a system carrier (aircraft without a cargo item) and wear (of moving a cargo item) bodies to reflect the changes in the inertia tensor. The functioning of the power plant, steering actuators, flight control system, an exhaust chute, the sensors of the primary information are taken into account. The equations of motion for systems of bodies projected on the aircraft reference plane are being recorded. This approach takes into account changes of the inertia tensor and the position of the main central axes of inertia in the process of landing of a cargo item. It allows us to simulate the condition of the aircraft at all speeds of the pitch, normal overload, and masses of single piece and placement, as evidenced by the high convergence of modeling results with data from flight tests.

Every year new aircraft emerge in civil aviation (HA). The wide-body A-380 aircraft with a take-off weight of up to 560 t has come to operation recently. The wake vortex behind such plane poses a real threat for other planes. Such wake is especially dangerous during weak cross wind at take off and landing.Vortex wake behind the A 380 plane characteristics research using the developed copmuting software has been executed in this article. Design-software complex includes two mathematical models: the mathematical model of the close Wake vortex and the mathematical model of the distant Wake vortex. These mathematical models are based on the vortex method. A mathematical model of the close Wake vortex is based on the analytical-experimental approach. At cruising flight regimes it is a four vortex system Wake vortex, and at takeoff and landing regimes it is - six-or eight-vortex system. A mathematical model of the far Wake vortex is based on the exact solution of the Helmholtz equations. This allows taking into account the vortex diffusion and dissipation over time. The influence of the axial velocity in the mathematical model of the distant Wake vortex is given by placing it in the center of the vortex flow. Its intensity is found from the experimental data. Calculated fields are per-turbed velocities for the A-380 aircraft.Fields of the indignant speeds at a light cross wind of 0.5 m/s ÷ 1.5 m/s in varioustime points are presented. The moments at which there is a wing vortex lag of the A-380 plane over very center are runwayare shown. Calculation of aerodynamic characteristics of the MC-21-400 plane in the vortex trace of the A-380 plane is executed. It is shown when the MC-21-400 plane gets in to the center of a wings vortex, the arising moments of the roll are not parried.

For single-rotor helicopters there are special flight modes, when tail rotor (TR) is under significant inductive influence of vortical wake of main rotor (MR). Inductive influence of vortical wake of MR can provoke essential changes in flowing of TR and its aerodynamic characteristics comparing to isolated rotor. In this case increase of tail rotor pitch, necessary for helicopter controlling, is possible.The article contains computational modelling of TR work with vortical wake of MR at the example of MIL Mi-171 helicopter. The modelling has been made on the base of non-linear blade (free wake) vortical model of rotor, produced at Helicopter Design Department of MAI.Helicopter hovering modes with crosswind of various intensity Vz was considered. Thrust-time lationship forisolated TR and TR with vortical wake of MR for equal flight modes was obtained. Flow around the rotors was analyzed, its vortical wake was considered.The results make it possible to clarify the peculiarities of TR work on considered modes and MR influence on its work. It was found out that vortical wake of MR has a more significant impact on TR work with crosswind on the right, when TR falls into vortex ring state mode. Inductive influence of vortical wake of MR leads to vortexring state mode for TR on lower speeds (Vz :: 5 m/s) than in case of isolated work of TR (Vz :: 12,5 m/s). In that case,the required tail rotor pitch has increased by 13% for Vz = 5 m/s. The results of modelling and flight tests led to good agreement.

The paper presents numerical results analysis of main rotor vibration due to helicopter main rotor thrust pulsation.The calculation method, the object of research and numerical research results with the aim to reduce the amplitude of the vibrations transmitted to the hub from the helicopters main rotor by the individual blade control in azimuth by the installation angle of blades cyclic changes are set out in the article. The individual blades control law for a five-blade main rotor based on the blade frequencies is made. It allows reducing the vibration from thrust. Research takes into account the main rotor including and excluding the blade flapping motion. The minimal vibrations regime is identified.Numerical study of variable loads caused by unsteady flow around the main rotor blades at high relative speeds of flight, which transmitted to the rotor hub, is made. The scheme of a thin lifting surface and the rotor vortex theory are used for simulation of the aerodynamic loads on blades. Non - uniform loads caused by the thrust, decomposed on the blade harmonic and its overtones. The largest values of deviation from the mean amplitude thrust are received. The analysis of variable loads with a traditional control system is made. Algorithms of higher harmonics individual blade control capable of reducing the thrust pulsation under the average value of thrust are developed.Numerical research shows that individual blade control of high harmonics reduces variable loads. The necessary change in the blade installation is about ± 0,2 degree that corresponds to the maximum displacement of the additional con- trol stick is about 1 mm.To receive the overall picture is necessary to consider all six components of forces and moments. Control law with own constants will obtained for each of them. It is supposed, that each of six individual blade control laws have an impact on other components. Thus, the problem reduces to the optimization issue. The individual blade control general law will be received as a result. It will meet a lot of conflicting requirements.

The main difficulty is that blade is a highly dynamic moving object.This work suggests two-channel measuring system of helicopter blades position. This measuring system consists of strain-gauge and optical measuring channels. The optical measuring system is based on high-speed camera and computer unit, which processes received video frames with blade image. The strain-gauge measuring system consists of strain sensors (located on the blade face) and of strain-gauge station, which records measurements and transmits them to the computer unit via radio channel for further analysis. The two-channel system concept was designed and device-algorithmic part of optical and strain-gauge channels was developed.Problems of the optical measuring channel, connected with low image contrast between blade and image background, blades identification and their position estimation were solved. For the strain-gauge measuring channel, a model of measurement translation from intrinsic frame of reference into helicopter frame of reference was suggested and measurement transmission system via radio channel to the computer unit was developed.The operational integrity and physical realizability of suggested two-channel system on research facility unit were inquired. The optical channel made measurements of two blades location, the strain-gauge channel made measurements only of that blade which had been equipped with sensors. The results of the research validated physical realizability of sug-gested two-channel measuring system.

Strong reduction of new aircraft design period using new technology based on artificial intelligence is the key problem mentioned in forecasts of leading aerospace industry research centers. This article covers the approach to development of quick aerodynamic design methods based on artificial intelligence neural system. The problem is being solved for the classical scheme of small sized unmanned aircraft vehicle (UAV). The principal parts of the method are the mathematical model of layout, layout generator of this type of aircraft is built on aircraft neural networks, automatic selection module for cleaning variety of layouts generated in automatic mode, robust direct computational fluid dynamics method, aerodynamic characteristics approximators on artificial neural networks.Methods based on artificial neural networks have intermediate position between computational fluid dynamics methods or experiments and simplified engineering approaches. The use of ANN for estimating aerodynamic characteris-tics put limitations on input data. For this task the layout must be presented as a vector with dimension not exceeding sev-eral hundred. Vector components must include all main parameters conventionally used for layouts description and completely replicate the most important aerodynamics and structural properties.The first stage of the work is presented in the paper. Simplified mathematical model of small sized UAV was developed. To estimate the range of geometrical parameters of layouts the review of existing vehicle was done. The result of the work is the algorithm and computer software for generating the layouts based on ANN technolo-gy. 10000 samples were generated and the dataset containig geometrical and aerodynamic characteristics of layoutwas created.

Developing quick aerodynamic design method of small unmanned aerial vehicle based on artificial neural webs technology requires a rich database used for creating, learning and testing algorythms. This database must be significantly larger than the number of existing vehicles, that's why creating a layout generator is important. On the first stage the database was increased by varying the parameters of the layout mathematical model in the accepted range. The layout generator is an artificial neural web trained on the rich database of unmanned aerial vehicles layouts in the frames of the simplified mathematical model. The important element in creating a new layout is the selection algorithm applied to the neural network output.The initial number of layouts was equal to 25. The simplified mathematical model describes the unmanned aerial vehicle layouts with 50 parameters. The layout generator forms an input file for the CFD code, which dimension is of the order of several thousands. The CFD codes BLWF and VISTRAN were chosen for this task. They add to each layout theinformation about its aerodynamic characteristics. Selection of the database was done into two levels. At first it was im-plemented on the layout generator output. Then on output vector after the CFD calculation. The selected databaser was used for creating an artificial neural networks number of which were equal to the number of aerodynamic coefficients. The algorithm was built in MATLAB and has convenient interface, which can be used for design process.

This paper presents the initial stage of work out of the helicopter body aerodynamic configuration. The main purpose of this work is to design the model of the fuselage and to minimize its drag.The analysis of experimental data obtained in TsAGI and other research centers was made at the first stage of the work. All features of flow around parts of the fuselage obtained from experimental data were taken into account. The dependencies of the fuselage component drag, such as the bow, fairings exhaust pipes of helicopter, sponsons, and tail sectionof the fuselage, on their form are described in this article.At the second stage the fuselage geometry was created in program SolidWorks. All the features of the flow around various fuselage components derived from the experimental data were considered in designing.The third stage is calculating of fuselage model aerodynamic characteristics. The calculations were made in the program ANSYS CFX (TsAGI License №501024). Boundary conditions were chosen so as to correspond to normal atmospheric conditions at 1,000 meters with velocity of flight is V = 85 m/s. The output of the hot jet from engines is takinginto account in computation. The purpose of this calculation is to find the optimal angle of the engine exhaust pipe whenthe hot spray does not intersect with the tail and stabilizer and creates the maximum of propulsive force. The volume of the grid in computational domain is approximately 13 million cells.Data analysis has shown that the fuselage has a 20% less drag at cruising flight (аf = -4 °) compared to the original model. The hot jets do not intersect with the tail and stabilizers at cruising flight so the fuselage is protected from overheating.

The problem of fuselage shape optimization of the wing-body configuration is considered in the following three formulations. In the first one, the angle of attack is fixed and equal to zero, the wing has a symmetric airfoil, and the fuselage is based on circular cross sections. In the second one, the fuselage cross sections are elliptical. In the third one, the angle of attack is varied, the lifting force coefficient is fixed, the wing is preliminary optimized, the fuselage is designed by the cross sections that consist of upper and lower half-ellipses with a possibility of a shift along vertical axis. The configuration volume, fuselage length, shape and position of the wing are fixed. The drag coefficient is the objective function. The optimization is carried out by the Indirect Optimization based on Self-Organization (IOSO) technology. Aerodynamic coef- ficients are obtained from the solution of the RANS equations with SST turbulence model by the ANSYS CFX software on the structured multiblock meshes. The results obtained by the optimization are compared with the configuration that is designed by traditional means. The fuselage of this configuration has a cylindrical part in the area of the wing-fuselage connection and nose part of the von Karman’s ogive shape. The solution of the optimization problem in the first formulation reduces drag coefficient at zero angle of attack by approximately 3 %. The use of the fuselage with elliptical cross sections makes it possible to reduce drag coefficient at zero angle of attack by 9 %. The solution of the optimization problem in first two formulations reduces drag coefficient at the wide range of angles of attack. When the lifting coefficient is selected for the third problem formulation as constraint the drag reduction is about 7 %. Additional drag reduction of about 2,5 % is obtained by the use of the fuselage asymmetric relative to the horizontal plane. The optimal fuselage design has a specific grotto in the lower part of the fuselage - the constriction from the sides with continuing height growth. The nose part of the optimal fuselage is widened, has a triangular shape in the top view and is deflected down.

The search of optimal variants for composite repair patches allows to increase the service life of a damaged air- plane structure. To sensibly choose the way of repair, it is necessary to have a computational complex to predict the stress- strain condition of "structure-adhesive-patch" system and to take into account the damage growth considering the material properties change. The variant of the computational complex based on inclusion method is proposed.For calculation purposes the repair bonded joint is divided into two areas: a metal plate with patch-shaped hole and a "patch-adhesive layer-skin" composite plate (inclusion).Calculation stages:Evaluation of the patch influence to the skin stress-strain condition, stress distribution between skin and patch in the case of no damage. Calculation of the stress-strain condition is performed separately for the skin with hole and for the inclusion; solutions are coupled based on strain compatibility.Definition of the damage growth parameters at new stress-strain condition due to bonded patch existence. Skincrack stress intensity factors are found to identify the crack growth velocity. Patch is modelled as a set of "springs" bridging the crack.Degradation analysis of elasticity properties for the patch material.Repair effectiveness is evaluated with respect to crack growth velocity reduction in the initial material in compari- son with the case of the patch absence.Calculation example for the crack repair effectiveness depending on number of loading cycles for the 7075-T6 aluminum skin is given. Repair patches are carbon-epoxy, glass-epoxy and boron-epoxy material systems with quasi- isotropic layup and GLARE hybrid metal-polymeric material.The analysis shows the high effectiveness of the carbon-epoxy patch. Due to low stiffness, the glass-epoxy patchdemonstrates the least effectiveness. GLARE patch containing the fiberglass plies oriented across the crack has the same effectiveness as the carbon and boron patches.Proposed bonded repair calculation method and corresponding computational model allow to analyze effectively the possible structural damage cases and to select optimal variant of patch installation subject to material durability undercyclic loads. Lack of this information may lead to establishing the inadequate inspection intervals of the damage locationand may reflect on economic factors of the airplane maintenance and flight safety.

The experience of construction of composite primary aircraft structures has approved that the weight decrease for composite aggregates of aircraft in comparison with metallic analogues cannot be obtained within the frames of conventional structures based on stiffened laminated composite skin. One of the main reasons is the low level of stress-strain characteris- tics of current polymer resins, which does not allow to realize high strength characteristics of carbon fibers in laminated composite packages. This consequence does not allow to use stiffened composite skin as a universal structure element, un- like metallic stiffened skin that sustains all main mechanical loads, including impact, loads from pressurizing and also sus- tains environmental factors. Hence, creation of lightweight and reliable composite primary structure elements of airframe, the novel types of structure layouts, allowing to realize high potential of current composite materials to the maximal extent, should be developed. In order to provide the development of such layots the novel approach is required, as the conventional stage-by-stage approach is based on a number of sufficient assumptions, most of which are correct only for the structures made of metallic alloys, but not correct as applied to the structures made of composite materials. The impossibility of the application of a stage-by-stage approach together with a significant increase of number of design parameters (in comparison with metallic structures), leads to the radical increase of labor input of the design task of composite structures.The novel approach to design of composite airframes, allowing to significantly decrease the extremely high labor input of the design process for composite structures is presented in this study. The approach presumes simultaneous solu- tions of design tasks on different levels of detailing of composite aircraft structure within the frames of the one integral design stage. The application of the novel approach allowed to obtain a number of lightweight solutions for the structure of cylindrical fuselage section of civil aircraft and for the hermetic cabin of the "Flying wing" aircraft.

Lattice composite fuselage structures are developed as an alternative to conventional composite structures based on laminated skin and stiffeners. Structure layout of lattice structures allows to realize advantages of current composite materials to a maximal extent, at the same time minimizing their main shortcomings, that allows to provide higher weight efficiency for these structures in comparison with conventional analogues.Development and creation of lattice composite structures requires development of novel methods of strength anal- ysis, as conventional methods, as a rule, are aiming to strength analysis of thin-walled elements and do not allow to get confident estimation of local strength of high-loaded unidirectional composite ribs.In the present work the method of operative strength analysis of lattice composite structure is presented, based onspecialized FE-models of unidirectional composite ribs and their intersections. In the frames of the method, every rib is modeled by a caisson structure, consisting of arbitrary number of flanges and webs, modeled by membrane finite elements. Parameters of flanges and webs are calculated automatically from the condition of stiffness characteristics equality of real rib and the model. This method allows to perform local strength analysis of high-loaded ribs of lattice structure without use of here-dimensional finite elements, that allows to shorten time of calculations and sufficiently simplify the procedure of analysis of results of calculations.For validation of the suggested method, the results of experimental investigations of full-scale prototype of shell of lattice composite fuselage section have been used. The prototype of the lattice section was manufactured in CRISM and tested in TsAGI within the frames of a number of Russian and International scientific projects. The results of validation have shown that the suggested method allows to provide high operability of strength analysis, keeping high accuracy of estimation of strength parameters, and can be used as a base method of strength analysis on the preliminary stage of design of lattice composite fuselage section structures.

This article presents the results of the hydrodynamic parameters of radio-controlled models (RCM) of the aircraft with the landing gear on an air cushion (REFERENCED) obtained during tests in the water tank of TSAGI NIMC on land- ing modes with varying alignment and pressure in air cushion chassis on calm and disturbed water surfaces.ACLG’s RCM is based on the Froude criterion. The experimental data of the real aircraft Dingo, LMS, An-26, C-130 Hercules (size, weight, thrust) parameters were processed. Tests were carried out, using the standard testbed, utiliz- ing the dynamically-corresponding models in the water tank. Drag best value rate while travelling along the smooth water was reached at the rear centering, with balloon pressure on the water of 700 Pa. In this case, the hump drag, at velocity of 2 mps, does not exceed 29 Newtons (hydrodynamic fineness on the ‘drag hump’ is, Кг=13,5), while at velocity of 10 mps, the drag is 30 Newtons, at Кг =13.The most unfavourable mode of motion is the one with configuration of lowered pressure in the balloons (400 Pa).In these cases, the Кг = 6,5. At motion with yaw angle of 10º, the drag rate meaning practically stands stable (Кг = 13,1), while at 20º it grows (Кг = 10,6).At motion along the waved surface, the critical wave length equals to two thirds of the ACU, while drag raises by the quarter, compared to other wave types. Such vehicles can be used in the hard-to-reach regions of the Russian Federation.

The recent experience of creating an unmanned combat aerial vehicle indicates that the main problems do not concern the development of an unmanned fighter as an aerial vehicle. The greatest challenge lies in creating the algorithms, data sensors, control hardware, communications hardware, etc. necessary for utilization of an unmanned aerial vehicle (UAV). In this context it is important to highlight the problem of replacing the pilot as a sensor and a flight operator on board of the UAV. This problem can be partially solved by introducing remote control, but there are some flight stages where it can only be executed under a fully independent control and data support due to various reasons, such as tight time, short duration, lack of robust communication, etc. These stages include combat deployment (surface attack or air attack) which make the highest demands on the fighter's design, that is why the promising UAV are currently considered to be "as autonomous as possible". It is obvious that the efficiency of an autonomous UAV will be determined mostly by the effectiveness of its automated control algorithms, and this dependence will increase together with the level of UAV autonomy. On the other hand, the optimal control algorithms can only be synthesized based on the control object characteristics. It means the development of UAV external design and the synthesis of its control algorithms should occur simultaneously and interdependently. This article presents the content and gives an example of the use of the method of maneuverable UAV external design, the distinctive feature of which lies in the interdependent processes of UAV external design developing and the synthesizing of its automated control algorithms.

The article deals with the wind-powered drive location and functioning pecularities of agricultural airplane special equipment. The data on the advantages and prospects of the use of the wind-powered drive on agricultural aircraft as well as the experimental and analytical data on discontinuity of the wind stream in the zone of the wind-powered drive installa- and parameters of its operation at aerial spraying on the example of specially equipped An-2 are presented. A mathematical model of a wind propeller in discontinuous wind stream is developed. Software package to comprehensively assess wind-powered drive conditions and performance indicators is generated and tested taking into account experimental data, as well as airplane performance and flight modes. The typical data on the wind propeller and the wind-powered drive in the spraying process are received; specific features of the dynamics at the wind-powered drive turning-on and the wind propeller rotation at zero load are determined and the data on the wind-propeller blades load imbalance at spraying and transient processes are defined. The author made a quantitative estimation of how the standard change of flight parameters and the operation of the aircraft power plant at spraying influence on the factors of wind-powered drive operation and working liquids suppey. The main tendencies, qualitative and quantitative aspects of a wind-powered drive use are defined when performing aerial works and the proposals and recommendations for improving and using it in the future are made. The findings of the work can be used in the wind-powered drive of agricultural airplane special equipment designing and modernization and drilling of the operational and technological procedures of aerial works performance in agriculture and forestry.

It is known that on the territory of the Russian Federation the construction of several large wind farms is planned. The tasks connected with design and efficiency evaluation of wind farm work are in demand today. One of the possible directions in design is connected with mathematical modeling. The method of large eddy simulation developed within the direction of computational hydrodynamics allows to reproduce unsteady structure of the flow in details and to determine various integrated values. The calculation of work for single wind turbine installation by means of large eddy simulation and Actuator Line Method along the turbine blade is given in this work. For problem definition the numerical method in the form of a box was considered and the adapted unstructured grid was used.The mathematical model included the main equations of continuity and momentum equations for incompressible fluid. The large-scale vortex structures were calculated by means of integration of the filtered equations. The calculation was carried out with Smagorinsky model for determination of subgrid scale turbulent viscosity. The geometrical parametersof wind turbine were set proceeding from open sources in the Internet.All physical values were defined at center of computational cell. The approximation of items in equations was executed with the second order of accuracy for time and space. The equations for coupling velocity and pressure were solved by means of iterative algorithm PIMPLE. The total quantity of the calculated physical values on each time step was equal to 18. So, the resources of a high performance cluster were required.As a result of flow calculation in wake for the three-bladed turbine average and instantaneous values of velocity, pressure, subgrid kinetic energy and turbulent viscosity, components of subgrid stress tensor were worked out. The received results matched the known results of experiments and numerical simulation, testify the opportunity to adequatelycalculate the flow parameters for a single wind turbine.

The article is devoted to the validation and application of CFD code for turbulent flows. Two-dimensional unsteady flows in the cavities and compartments and three-dimensional flow in the compartment of complex geometry have been considered. Two turbulence parameter oriented models are used.Numerical simulation of unsteady transonic flow (Mоо=0.74) in a narrow channel with a cavity inside has been conducted. The dependence of the static pressure on time at fixed points in space has been obtained. The fast Fourier transform has been applied for processing data of static pressure. The difference of 6-10% between the numerical and experi-mental data has been obtained.The computations of unsteady transonic cavity flow with Mach number Mоо=0.85 have been performed. Low frequency oscillations of the static pressure in several fixed points in space have been obtained. Power spectrum of oscillations at the center of the cavity is compared with experimental data and Rossiter modes. An acceptable agreement between experimental and computed data has been achieved. The influence of geometrical factors on the frequency characteristics of the flow has been investigated. For this purpose two round flaps have been added to the cavity. The most low-frequency oscillation modes changed by the presence of the flaps. The first mode was gone, the second mode amplitude decreased and the third mode amplitude significantly decreased. The changes in height of protruding part of the geometry to the external flow have led to changes in pressure pulsation amplitude without changing the frequency. The spectral functions obtained while using the two considered models of turbulence have been compared for this case. It is found that the frequency values are only slightly different; the main difference is present at the amplitude of pulsations.The effect of deflection of flat flap on the non-stationary subsonic flow parameters in a cylindrical body with an inner compartment has been investigated. The cases of deflection angles of the flap inside the compartment with values 26º and 41º above the horizontal plane, and also the case without flap have been considered. Low-frequency oscillations of the static pressure have been obtained. The presence of the flap did not change the frequency of static pressure pulsations. With the increase of the choke deflection angle, the oscillation amplitude increases at all considered points of the flow too.