Podguiko N. A., Marakhtanov M. K., Semenkin A. V., Khokhlov Y. A. Studying cold hollow magnetron cathode for electric thruster. Aerospace MAI Journal, 2022, vol. 29, no 1, pp. 109-117.

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.

Komov A. A. Aircraft landing gear scheme and engine protection. Aerospace MAI Journal, 2022, vol. 29, no 1, pp. 7-18.

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.

Baklanov A. V. Fuel combustion efficiency ensuring in low-emission combustion chamber of gas turbine engine under various climate conditions. Aerospace MAI Journal, 2022, vol. 29, no 1, pp. 144-155.

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.

Dunyashev D. A., Goldovskii A. A., Pravidlo M. N. Design problems of a small-size unmanned aerial vehicle launching system by free fall method. Aerospace MAI Journal, 2022, vol. 29, no 1, pp. 27-35.

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.

Mitrofanov O. V., Osman M. . Smooth metallic panels designing while stability and strength ensuring at postbuckling behavior. Aerospace MAI Journal, 2022, vol. 29, no 1, pp. 36-47.

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.

Golovchenko E. V., Mistrov L. E., Dum'yak S. G. A thechnique for flight check-up of ground-based radio-technical support facilities for flight support with unmanned aerial vehicle application. Aerospace MAI Journal, 2022, vol. 29, no 1, pp. 156-170.

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.

Dolgov O. S., Safoklov B. B. Developing maintenance and refurbishment model of aerial vehicles with artificial neural network applicaion. Aerospace MAI Journal, 2022, vol. 29, no 1, pp. 19-26.

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.

Yurtaev A. A., Badykov R. R., Benedyuk M. A., Senchev M. N. Determining radial gaps values of centrifugal compressor and turbine of a small-sized gas turbine engine at maximum operation mode. Aerospace MAI Journal, 2022, vol. 29, no 1, pp. .

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.

Ageev A. G., Zhdanov A. V., Galanova A. P. The residual fuel flow-over in the wing tanks while aircraft maneuvering. Aerospace MAI Journal, 2022, vol. 29, no 1, pp. 48-56.

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.

Balyk V. M., Borodin I. D. Selection of stable design solutions for unmanned aerial vehicle under conditions of uncertainty factors action. Aerospace MAI Journal, 2022, vol. 29, no 1, pp. .

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.

Shilkin O. V., Kolesnikov A. P., Kishkin A. A., Zuev A. A., Delkov A. V. Designing passive thermal control system with a capacity of up to 3 kW by heat pipes and active heating elements for a spacecraft. Aerospace MAI Journal, 2022, vol. 29, no 1, pp. 67-80.

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.

Malinovskii I. M., Nesterenko V. G., Starodumov A. V., Yusipov B. H., Ivanov I. G. Analysis and constructive methods for axial forces distribution optimization in turbojet engine to enhance the high-pressure rotor bearing sevice life. Aerospace MAI Journal, 2022, vol. 29, no 1, pp. 81-94.

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.

Kalenskii S. M., Morzeeva T. A., Ezrokhi Y. A., Pankov S. V. Selection of rational parameters of distributed propulsion system in structure of the long range aircraft. Aerospace MAI Journal, 2022, vol. 29, no 1, pp. 95-108.

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.

Bogomolov M. A., Gras'ko T. V., Zinenkov Y. V., Lukovnikov A. V. Optimal engine parameters searching for the short-haul passenger aircraft. Aerospace MAI Journal, 2022, vol. 29, no 1, pp. 118-130.

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.

Ivanov P. I. Weight model rescue system at parachute systems flight tests conducting. Aerospace MAI Journal, 2022, vol. 29, no 1, pp. 171-183.

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].

Lupanchuk V. Y. Optical surveillance system of unmanned aerial vehicle and a method of its stabilization. Aerospace MAI Journal, 2022, vol. 29, no 1, pp. 184-200.

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.

Efremov A. V., Shcherbakov A. I., Korzun F. A., Prodanik V. A. Prospective means for the aircraft pilot induced oscillation suppression. Aerospace MAI Journal, 2022, vol. 29, no 1, pp. 201-210.

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(δmax). However, when δ exceeds δmax, 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.

Terekhov R. I. Estimation of fly-by-wire emergency servo-control of regional aircraft with account for nonlinear specifics of control surfaces dynamics. Aerospace MAI Journal, 2022, vol. 29, no 1, pp. 211-225.

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 Wz = 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.

Petrov M. A., Matveev A. G., Petrov P. A., Saprykin B. Y. Computation and analyzing bulk forming processes with a rotating tool using FE simulation. Aerospace MAI Journal, 2022, vol. 29, no 1, pp. 226-244.

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.

Petrova L. G., Belashova I. S. Assessment of solid-solution hardening of austenitic alloys at nitrogen alloying. Aerospace MAI Journal, 2022, vol. 29, no 1, pp. 245-252.

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.

Novogorodtsev E. V., Karpov E. V., Koltok N. G. Characteristics improvement of spatial fixed-geometry air intakes of external compression based on boundary layer control systems application. Aerospace MAI Journal, 2021, vol. 28, no 4, pp. 7-27.

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 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 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.

Volkova A. O., Jet-perforated boundaries as an effective method to reduce wall interference for airfoil tests in a transonic wind tunnel. Aerospace MAI Journal, 2021, vol. 28, no 4, pp. 28-38.

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.

Pigusov E. A., Experimental study on wing adaptive high-lift devices of transport aircraft on takeoff-landing mode. Aerospace MAI Journal, 2021, vol. 28, no 4, pp. 39-47.

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: δF = 30°, δF = 40° and δF = 30°/20°. The spoiler droop (adaptive element) in the tests deflected by the angles δSD = 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.

Petronevich V. V., Lyutov V. V., Manvelyan V. S., Kulikov A. A., Zimogorov S. V. Studies on six-component rotating strain-gauge balance calibration for aircraft propellers testing. Aerospace MAI Journal, 2021, vol. 28, no 4, pp. 48-61.

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.

Lamzin V. V., Lamzin V. A. Integrated assessment technique for the earth remote probing spacecraft rational parameters and development program. Aerospace MAI Journal, 2021, vol. 28, no 4, pp. 62-77.

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.

Bezzametnov O. N., Mitryaikin V. I., Khaliulin V. I., Markovtsev V. A., Shanygin A. N. Impact damages effect assessment on compressive strength of integral panels from polymer composite materials. Aerospace MAI Journal, 2021, vol. 28, no 4, pp. 78-91.

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.

Kudryavtsev I. V. Ensuring dynamic state of straight waveguide paths at heating by supports arrangement. Aerospace MAI Journal, 2021, vol. 28, no 4, pp. 92-105.

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.

Podruzhin E. G., Zagidulin A. R., Shinkarev D. A. Drop testing simulation of the mainline aircraft landing gear. Aerospace MAI Journal, 2021, vol. 28, no 4, pp. 106-117.

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.

Maskaykin V. A., Makhrov V. P. Thermal conductivity research of the aircraft heat-insulating skin under flight conditions. Aerospace MAI Journal, 2021, vol. 28, no 4, pp. 118-130.

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.

Sirotin N. N., Nguyen T. S. Numerical simulation technique for working blades operational damages of turbojet low-pressure compressor rotor. Aerospace MAI Journal, 2021, vol. 28, no 4, pp. 131-150.

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 npr = 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 npr = 80%, and npr = 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.

Balakin D. A., Zubko A. I., Zubko A. A., Shtykov V. V. Vibration diagnostics of gas turbine engines bearing assemblies technical condition with rhythmograms and scatterograms. Aerospace MAI Journal, 2021, vol. 28, no 4, pp. 151-162.

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.

Klinskii B. M. Determining test-bench box aerodynamics impact on the force from the gas turbine engine thrust by layout changing of the inlet lemniscate mouth piece. Aerospace MAI Journal, 2021, vol. 28, no 4, pp. 163-179.

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 nr 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 Reng.cor is being determined with no account for the values of the input impulse ΔRinlet and aerodynamic drag ΔRwindage), 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 R0eng.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 ΔRwindage with connected pipeline at the inlet, applied while the test-bench box calibration, as the difference between the thrust values ΔRair_force cor = R0eng.corr Reng.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 σin in the inlet device difference from 1.0, as well as total pressure at the inlet Pin*, damped temperature Tin* and the moisture content d difference from the standard values.

The aerodynamic force value (ΔRAF) 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 (ΔRair.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 nfan.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.

Tkachenko A. Y. Working fluid mathematical model for the gas turbine engine thermo-gas-dynamic design. Aerospace MAI Journal, 2021, vol. 28, no 4, pp. 180-191.

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.

Bernikov A. S., Bogachev V. A., Mikhailov D. N., Petrov Y. A., Sergeev D. V. The study of martian dust impact on “ExoMars” spacecraft structures unfurling elements after touchdown. Aerospace MAI Journal, 2021, vol. 28, no 4, pp. 192-203.

«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.

Ilyukhin S. N. Trajectory estimation procedure of small-sized aerial vehicles at the studies on a ballistic track. Aerospace MAI Journal, 2021, vol. 28, no 4, pp. 204-218.

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.

Tikhonov V. N. Analysis of accuracy characteristics, probabilistic characteristics and expert evaluations of aircraft by the pilots while in-flight refueling. Aerospace MAI Journal, 2021, vol. 28, no 4, pp. 219-231.

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.

Shevchenko I. V., Sokolov V. P., Rogalev A. N., Vegera A. N., Osipov S. K. Study of cyclonic cooling system geometry parameters impact of gas turbine blade leading edge on its thermo-hydraulic characteristics. Aerospace MAI Journal, 2021, vol. 28, no 4, pp. 232-244.

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.

Zelenskii A. A., Ilyukhin Y. V., Gribkov A. A. Memory-centric models of industrial robots control systems. Aerospace MAI Journal, 2021, vol. 28, no 4, pp. 245-256.

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.

Kovalev A. A., Krasko A. S., Sidorov P. A. Shock interaction simulation of sprayed particles with the part surface while plasma coatings forming. Aerospace MAI Journal, 2021, vol. 28, no 4, pp. 257-266.

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.

Zaharov E. N., Usachev D. V. An approach to the assessment of military-oriented aircraft engineering based on neural-like networks. Aerospace MAI Journal, 2021, vol. 28, no 4, pp. 267-280.

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. 
  1. 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.
  2. 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.
  3. If necessary, factor analysis is performed to check correctness and reliability of the resulting operating model of the neural-like network.
  4. 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>
  1. 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.

Ovsyannikova E. B., Timushev S. F. ON THE 100th ANNIVERSARY OF THE PROMINENT SCHOLAR PROFESSOR B.V. OVSYANNIKOV. Aerospace MAI Journal, 2021, vol. 28, no 3, pp. 7-16.

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.

Ankudinov A. A., Vashchenko A. V. Axial-vortex stage application prospects in turbo-pumps of liquid propellant rocket engines. Aerospace MAI Journal, 2021, vol. 28, no 3, pp. 17-23.

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.

Gemranova E. A. State diagnosing of automatic relief valve circuit and parkiing seal of liquid rocket engine turbo pump. Aerospace MAI Journal, 2021, vol. 28, no 3, pp. 24-32.

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.

Ivanov A. V. Analysis of contacless seal type impact on the pump characteristics of а rocket engine turbo-pump unit while operating mode changing. Aerospace MAI Journal, 2021, vol. 28, no 3, pp. 33-45.

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.

Kamensky K. V., Martirosov D. S. А method for current state monitoring of liquid rocket engine in stationary and transient modes. Aerospace MAI Journal, 2021, vol. 28, no 3, pp. 46-53.

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.

Kochetkov Y. M., Burova A. Y. Gas-dynamic reasons for vibrations origination in turbopump units. Aerospace MAI Journal, 2021, vol. 28, no 3, pp. 54-62.

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.

Filin N. A., Mkrtchyan M. K. Little-known facts of turbopump unit creation history in ijquid rocket engine. Aerospace MAI Journal, 2021, vol. 28, no 3, pp. 63-72.

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.

Trulev A. V., Shmidt E. M. Bench tests methodological specifics of submersible electric centrifugal pumps gas separating installations for oil extraction. Aerospace MAI Journal, 2021, vol. 28, no 3, pp. 73-80.

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.

Ivanov P. I. Filling the double-shell wing of a gliding parachute. Aerospace MAI Journal, 2021, vol. 28, no 3, pp. 81-94.

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.

Lamzin V. A., Lamzin V. V. Method for characteristics predicting of prospective earth probing spacecraft with optoelectronic imaging hafdware. Aerospace MAI Journal, 2021, vol. 28, no 3, pp. 95-112.

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.

Kaurov I. V., Tkachenko I. S., Salmin V. V. Design technique for small spacecraft thermal control system and mathematical models verificatioin based on telelmetry data. Aerospace MAI Journal, 2021, vol. 28, no 3, pp. 113-129.

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.

Nikitin I. S., Magdin A. G., Pripadchev A. D., Gorbunov A. A. Turbojet engine power increasing by air-cooling at the inlet device. Aerospace MAI Journal, 2021, vol. 28, no 3, pp. 130-138.

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.

Koval' S. N., Badernikov A. V., Shmotin Y. N., Pyatunin K. R. Digital twin technology application while gas turbine engines development. Aerospace MAI Journal, 2021, vol. 28, no 3, pp. 139-145.

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.

Aleksandrov Y. B., Nguyen T. D., Mingazov B. G., Korol'kova E. V., Sharafutdinov R. R. Swirler vanes installation angle impact on flow mixing efficiency behind the flame tube head of gas turbine engine combustion chamber. Aerospace MAI Journal, 2021, vol. 28, no 3, pp. 146-158.

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 RAV is the radius of the annular source RCS;

  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.

Filinov E. P., Bezborodova K. V. Double bypass turbojet engine structure analysis. Aerospace MAI Journal, 2021, vol. 28, no 3, pp. 159-170.

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 — Msu+t, as well as the fuel consumption in kilogram per one ton-kilometer of transported cargo —

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.

Baturin O. V., Nikolalde P. .., Popov G. M., Korneeva A. I., Kudryashov I. A. Mathematical model identification of gas turbine engine with account for initial data uncertainty. Aerospace MAI Journal, 2021, vol. 28, no 3, pp. 171-185.

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.

Shvetsova S. V., Shvetsov A. V. Unmanned aerial vehicles integration into modern infrastructure systems operation. Aerospace MAI Journal, 2021, vol. 28, no 3, pp. 186-193.

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.

Vlasova A. V. Interaction capabilities of air traffic control systems with structures ensuring airport aviation security. Aerospace MAI Journal, 2021, vol. 28, no 3, pp. 194-201.

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.

Balyakin A. V., Skuratov D. L., Khaimovich A. I., Oleinik M. A. Direct laser fusion application for powders from heat resistant allows in engine building. Aerospace MAI Journal, 2021, vol. 28, no 3, pp. 202-217.

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.

Murav’ev V. I., Bakhmatov P. V., Grigor’ev V. V. Properties ensuring of aircraft titanium structures joints obtained by fuse welding identical to the basic metal properties. Aerospace MAI Journal, 2021, vol. 28, no 3, pp. 218-227.

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.

Vinogradov O. N., Kornushenko A. V., Pavlenko O. V., Petrov A. V., Pigusov E. A., Trinh T. N. Specifics of propeller and super-high aspect ratio wing interference in non-uniform flow. Aerospace MAI Journal, 2021, vol. 28, no 2, pp. .

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 Сxai (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.

Moshkov P. A., Samokhin V. F. Calculated estimation technique for audibility boundaries of propeller unmanned aerial vehicles. Aerospace MAI Journal, 2021, vol. 28, no 2, pp. 20-36.

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.

Bolsunovskii A. L., Buzoverya N. P., Bragin N. N., Gerasimov S. V., Pushchin N. A., Chernyshev I. L. Numerical and experimental studies on the over-the-wing-engine configurations aerodynamics. Aerospace MAI Journal, 2021, vol. 28, no 2, pp. 37-49.

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.

Artamonov B. L., Zagranichnov A. S., Lisovinov A. V. Heavy helicopter for arctic transport system. Aerospace MAI Journal, 2021, vol. 28, no 2, pp. 52-66.

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.

Petronevich V. V., Lyutov V. V., Manvelyan V. S., Kulikov A. A., Zimogorov S. V. Study on six-component rotating strain-gauge balance development for helicopter tail rotor testing. Aerospace MAI Journal, 2021, vol. 28, no 2, pp. 69-84.

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.

Klyagin V. A., Laushin D. A. An approach to the probability determining of the specified flight performance achieving, and account for risk factors while an aircraft appearance forming. Aerospace MAI Journal, 2021, vol. 28, no 2, pp. 85-95.

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.

Shilkin O. V., Kishkin A. A., Zuev A. A., Delkov A. V., Lavrov N. A. Passive cooling system designing for a spacecraft onboard complex. Aerospace MAI Journal, 2021, vol. 28, no 2, pp. 96-106.

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.

Mousavi Safavi S. M., Garipov L. A., Kluev S. V., Yusupov I. R. Comparative study on compressive mechanical characteristics of X-shape and pyramidal trussed fillers. Aerospace MAI Journal, 2021, vol. 28, no 2, pp. 107-114.

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.

Ivanov P. I. Computation of aerodynamic load on gliding parachute while its deploying and overloading, acting on the airdrop object. Aerospace MAI Journal, 2021, vol. 28, no 2, pp. 115-126.

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.

Nikolaev E. I., Yugai P. V. Analysis of the external airbags application expediency on a helicopter. Aerospace MAI Journal, 2021, vol. 28, no 2, pp. 127-139.

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.

Shaidullin R. A., Bekerov A. R., Sabirzyanov A. N. Flow swirl impact at the rocket engine nozzle inlet on the flow coefficient. Aerospace MAI Journal, 2021, vol. 28, no 2, pp. 142-151.

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.

Prokhorenko I. S., Katashov A. V., Katashova M. I. Gas propulsion correcting unit for nanosatellites. Aerospace MAI Journal, 2021, vol. 28, no 2, pp. 152-165.

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.

Mkrtchyan M. K., Kochetkov Y. M. . Aerospace MAI Journal, 2021, vol. 28, no 2, pp. 166-174.

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.

Ragulin I. A., Aleksandrov V. V. Lag effect impact in the control system channel of highly automated aircraft on the control lever type selection and its command signal. Aerospace MAI Journal, 2021, vol. 28, no 2, pp. 177-188.

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.

Vereshchikov D. V., Zhuravskii K. A., Kostin P. S. Motion control quality assessment of maneuverable aircraft. Aerospace MAI Journal, 2021, vol. 28, no 2, pp. 191-205.

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.

Bibikov P. S., Belashova I. S., Prokof'ev M. V. Nitridation technology specifics of high-alloy corrosion-resistant steels of aviation purposes. Aerospace MAI Journal, 2021, vol. 28, no 2, pp. 206-215.

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.

Ivanov Y. F., Rygina M. E., Petrikova E. A., Teresov A. D. Structure and mechanical properties of hypereutectic silumin irradiated by a pulsed electron beam. Aerospace MAI Journal, 2021, vol. 28, no 2, pp. 216-222.

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.

Bukichev Y. S., Bogdanova L. M., Spirin M. G., Shershnev V. A., Shilov G. V., Dzhardimalieva G. I. Composite materials based on epoxy matrix and titanium dioxide (IV) nanoparticles: synthesis, microstructure and properties. Aerospace MAI Journal, 2021, vol. 28, no 2, pp. 224-237.

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.

Komov A. A., Echevskii V. V. Reverse capacity and aircraft thrust reverse application efficiency. Aerospace MAI Journal, 2021, vol. 28, no 1, pp. 7-18.

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.
Neruchek A. O., Kotlyarov E. Y. Alternative layout of lunar landing module radiative heat exchanger and its thermal analysis based on computational experiment. Aerospace MAI Journal, 2021, vol. 28, no 1, pp. 35-44.

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.

Moshkov P. A., Samokhin V. F. Problems of light propeller-driven airplane design with regard to community noise requirements. Aerospace MAI Journal, 2021, vol. 28, no 1, pp. 19-34.

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.

Kishkin A. A., Zuev A. A., Delkov A. V., Shevchenko Y. N. Analytical approach while studying equations of boundary layer impulses at the flow in the inter-blade channel of gas turbines. Aerospace MAI Journal, 2021, vol. 28, no 1, pp. 45-60.

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 Rmax = 0.169 m and Rmin = 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 ReU = 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.

Filinov E. P., Kuz'michev V. S., Tkachenko A. Y., Ostapyuk Y. A. Determining required turbine cooling air flow rate at the conceptual design stage of gas turbine engine. Aerospace MAI Journal, 2021, vol. 28, no 1, pp. 61-73.

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.

Kaplin M. A., Mitrofanova O. A., Bernikova M. Y. Development of very low-power PlaS-type plasma thrusters. Aerospace MAI Journal, 2021, vol. 28, no 1, pp. 74-85.

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.

Baklanov A. V. Burner geometry impact on gas turbine engine combustion chamber characteristics. Aerospace MAI Journal, 2021, vol. 28, no 1, pp. 86-95.

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.

Aung K. M., Kolomentsev A. I., Martirosov D. S. Mathematical modelling of liquid rocket engine flow regulator in frequency and time domains. Aerospace MAI Journal, 2021, vol. 28, no 1, pp. 96-106.

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.

Sotskov I. A. Developing mathematical model of the 3d turbulent flow of combustion products in solid propellant rocket engines. Aerospace MAI Journal, 2021, vol. 28, no 1, pp. 107-114.

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.

Ivanov P. I., Krivorotov M. M., Kurinnyi S. M. Experiment informativity in flight tests of parachute systems. Decision making. Aerospace MAI Journal, 2021, vol. 28, no 1, pp. 126-136.

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.

Vasil’eva N. V., Dedkova E. V., Kutnik I. V., Fokin V. E., Chub N. A., Yurchenko E. S. Simulator stand designing for cosmonauts training to perform visual-instrumental observations. Aerospace MAI Journal, 2021, vol. 28, no 1, pp. 115-125.

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.

Chebakova A. A., Ganyak O. I., Tkachenko O. I. Speed control channel automation while aircraft aerial refueling. Aerospace MAI Journal, 2021, vol. 28, no 1, pp. 137-146.

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.

Salmin V. V., Petrukhina K. V., Kvetkin A. A. Approximate calculation of initial conditions of a spacecraft with solar electric-rocket propulsion plant starting while transferring from highly elliptic orbit to geostationary one. Aerospace MAI Journal, 2021, vol. 28, no 1, pp. 147-160.

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.

Rasulov Z. N., Kalugina M. S., Remshev E. Y., Afim’in G. O., Avetisyan A. R., Elfimov P. V. Studying isostatic pressing of samples being produced by the slm method for new components manufacturing of the combustion chamber housing. Aerospace MAI Journal, 2021, vol. 28, no 1, pp. 161-174.

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.

Kovalev A. A., Rogov N. V. Evaluation of quality indicator dispersion depending on technological process parameters. Aerospace MAI Journal, 2021, vol. 28, no 1, pp. 175-186.

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.

Bogdanov K. A. Studying ultrasonic oscillations impact on the surface roughness at the electrical discharge machining. Aerospace MAI Journal, 2021, vol. 28, no 1, pp. 187-199.

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.

Zhigulin I. E., Emel’yanenko K. A., Sataeva N. E. Studying ultrasonic oscillations impact on the surface roughness at the electrical discharge machining. Aerospace MAI Journal, 2021, vol. 28, no 1, pp. 200-212.

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.

Pavlenko O. V., Petrov A. V., Pigusov E. A. Studies of flow-around of high-lift wing airfoil with combined energy system for the wing lifting force increasing. Aerospace MAI Journal, 2020, vol. 27, no 4, pp. 7-20.

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.

Tudupova A. N., Strizhius V. E., Bobrovich A. V. Computational and experimental evaluation of fatigue life characteristics of the transport category aircraft composite wing panels. Aerospace MAI Journal, 2020, vol. 27, no 4, pp. 21-29.

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.

Chanov M. N., Skvortsov E. B., Shelekhova A. S., Bondarev A. V., Ovchinnikov V. G., Semenov A. A., Chernavskikh Y. N. Technical concepts analysis of transport aircraft with various power plant types and layout. Aerospace MAI Journal, 2020, vol. 27, no 4, pp. 30-47.

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.

Saprykin O. A. Planets exploration with reusable takeoff and landing complexes. Aerospace MAI Journal, 2020, vol. 27, no 4, pp. 48-58.

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.

Milyukov I. A., Rogalev A. N., Sokolov V. P. Approaches to design engineering and technological designing integration. Aerospace MAI Journal, 2020, vol. 27, no 4, pp. 59-70.

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.

Kryuchkov M. D. Parameters optimization technique for the carrier rocket with modular booster block modification. Aerospace MAI Journal, 2020, vol. 27, no 4, pp. 71-80.

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.

Bautin A. A., Svirskiy Y. A. Neural networks technologies application in problems of critical places status monitoring of transport aircraft structure. Aerospace MAI Journal, 2020, vol. 27, no 4, pp. 81-91.

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.

Ryzhova T. B., Petronyuk Y. S., Morokov E. S., Gulevskii I. V., Levin V. M., Shanygin A. N. Application of acoustic methods for identification and characterization of full destruction harbingers of carbon fiber-reinforced polymers while strength experimental study. Aerospace MAI Journal, 2020, vol. 27, no 4, pp. 92-104.

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.

Agaverdyev S. V., Zinenkov Y. V., Lukovnikov A. V. Optimal parameters selection of the strike unmanned aerial vehicle power plant. Aerospace MAI Journal, 2020, vol. 27, no 4, pp. 105-116.

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.

Balyakin A. V., Skuratov D. L. Calculation results of temperature fields while grinding workpieces from titanium alloys by abrasive belts of various types. Aerospace MAI Journal, 2020, vol. 27, no 4, pp. 117-123.

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.

Aslanov A. R., Raznoschikov V. V., Stol’nikov A. M. Studying parameters of aircraft cryogenic turbo-pump unit by the aircraft flight cycle. Aerospace MAI Journal, 2020, vol. 27, no 4, pp. 124-132.

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.

Omar H. H., Kuz'michev V. S., Tkachenko A. Y. Efficiency improving of aviation bypass turbojet engines through recuperator application. Aerospace MAI Journal, 2020, vol. 27, no 4, pp. 133-146.

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*, 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.

Remchukov S. S., Yaroslavtsev N. L., Lepeshkin A. R. Computer-aided design and calculation of the blade front cavity cooling system of the gas turbine engine. Aerospace MAI Journal, 2020, vol. 27, no 4, pp. 147-158.

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.

Baklanov A. V. Multilevel modelling application in the gas turbine engine low-emission combustion chamber design process. Aerospace MAI Journal, 2020, vol. 27, no 4, pp. 159-172.

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.

Semenenko D. A., Saevets P. A., Komarov A. A., Rumyantsev . V. Characteristics analysis of stationary plasma thruster. Aerospace MAI Journal, 2020, vol. 27, no 4, pp. 173-180.

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.

Sklyarova A. P., Gorbunov A. A., Zinenkov Y. V., Agul'nik A. ., Vovk M. Y. Search for optimal power plant to improve maneuverable aircraft efficiency. Aerospace MAI Journal, 2020, vol. 27, no 4, pp. 181-191.

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.

Fedorov A. V., Hoang V. T. Software package for motion control algorithms design of service module in geostationary orbit. Aerospace MAI Journal, 2020, vol. 27, no 4, pp. 192-205.

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.

Goncharov V. M., Zaitsev A. V., Lupanchuk V. Y. Coordinates measuring techniques improving of unmanned aerial vehicle in conditions of abnormality (distortion). Aerospace MAI Journal, 2020, vol. 27, no 4, pp. 206-221.

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.

Vyatlev P. A., Sergeev D. V., Sysoev A. K., Sysoev V. K. Long-term storage impact on spacecraft temperature-regulating coating elements characteristics. Aerospace MAI Journal, 2020, vol. 27, no 4, pp. 222-228.

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.

Il’inkova T. A., Il'inkov A. V., Klimkin Y. O., Zhivushkin A. A., Budinovskii S. A. Structure and properties transformation of heat-resistant coating in the process of high-temperature cyclic tests of the turbine blade. Aerospace MAI Journal, 2020, vol. 27, no 4, pp. 229-240.

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.

Nguyen T. H., Nguyen V. M., Le H. N., Nguyen H. . Kinetics of cobalt nanopowder obtaining process by hydrogen-reduction method under non-isothermal conditions. Aerospace MAI Journal, 2020, vol. 27, no 4, pp. 241-249.

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·e0,0248x. 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.

Pogosyan M. A., Vereikin A. A. Position and motion control of aerial vehicles in automatic landing systems: analytical review. Aerospace MAI Journal, 2020, vol. 27, no 3, pp. 7-22.

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.

Kul'kov V. M., Yoon S. W., Firsyuk S. O. A small spacecraft motion control method employing inflatable braking units for deceleration while orbital flight prior to the atmospheric entry. Aerospace MAI Journal, 2020, vol. 27, no 3, pp. 23-36.

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 σx(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.

Volkova A. O., Ivanov A. I., Streltsov E. V. Application of combined jet-perforated boundaries to solve the problem of the wind tunnel wall interference at transonic speeds. Aerospace MAI Journal, 2020, vol. 27, no 3, pp. 37-48.

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.

Ivanov P. I., Kurinnyi S. M., Krivorotov M. M. Parameters liable to be defined while a multi-dome parachute system flight-testing for its efficiency estimation. Aerospace MAI Journal, 2020, vol. 27, no 3, pp. 49-59.

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 FS = 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 FS = 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.

Moshkov P. A., Vasilenkov D. A., Rubanovskii V. V., Stroganov A. I. Noise sources localization in the rrj-95 aircraft pressure cabin by spherical microphone array. Part 2. Passenger cabin. Aerospace MAI Journal, 2020, vol. 27, no 3, pp. 60-72.

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.

Akimov V. N., Gryzin S. V., Parafes S. G. Studying the “rudder-drive” system with accounting for the rudder flexural-and-torsional vibrations. Aerospace MAI Journal, 2020, vol. 27, no 3, pp. 73-83.

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.

Sedel'nikov A. V., Taneeva A. S., Orlov D. I. Forming design layout of a technological purpose small spacecraft based on other class of technological spacecraft design and operation experience. Aerospace MAI Journal, 2020, vol. 27, no 3, pp. 84-93.

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.

Bakhmatov P. V., Pletnev N. O. Studying specifics of a permanent joint welding spot forming while the unit laser impulse effect on a low-carbon steel surface. Aerospace MAI Journal, 2020, vol. 27, no 3, pp. 94-102.

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.

Shevchenko M. O., Pasichnaya M. M. Developing airframe structure of a modern airplane for agricultural work performing. Aerospace MAI Journal, 2020, vol. 27, no 3, pp. 103-110.

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.

Bezzametnov O. N., Mitryaikin V. I., Khaliulin V. I., Krotova E. V. Developing technique for impact action resistance determining of the aircraft parts from composites with honeycomb filler. Aerospace MAI Journal, 2020, vol. 27, no 3, pp. 111-125.

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.

Aslanov A. R., Stol’nikov A. M., Raznoschikov V. V. Studying thermal state of the cryogenic fuel tank at the liquid fuel “mirror” vacillations. Aerospace MAI Journal, 2020, vol. 27, no 3, pp. 126-138.

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.

Baklanov A. V. The effect of the central body diameter of a dual-circuit burner on the hazardous substances release. Aerospace MAI Journal, 2020, vol. 27, no 3, pp. 139-145.

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 (Tflame =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.

Altunin K. V. Elaborating new specific parameters of a jet engine. Aerospace MAI Journal, 2020, vol. 27, no 3, pp. 146-154.

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.

Ahmed H. S., Osipov B. M. Diagnostics algorithm with gas turbine engine mathematical model application. Aerospace MAI Journal, 2020, vol. 27, no 3, pp. 155-166.

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%.

Zuev A. A., Arngol’d A. A., Nazarov V. P. Sections of dynamically non-stabilized flows in characteristic channels of the air-gas channels of liquid rocket engines turbopump units. Aerospace MAI Journal, 2020, vol. 27, no 3, pp. 167-185.

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.

Nadiradze A. B., Frolova Y. L. Mechanisms for forming median-energy ions in the jets of stationary plasma thrusters. Aerospace MAI Journal, 2020, vol. 27, no 3, pp. 186-197.

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.

Kryuchkov A. N., Plotnikov S. M., Sundukov A. E., Sundukov E. V. Vibration diagnostics of lateral clearance value in the toothed gearing of differential gearbox of a turboprop engine. Aerospace MAI Journal, 2020, vol. 27, no 3, pp. 198-208.

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.

D’yachenkova M. V., Anyutochkina A. S., Rubtsov E. A. Aircraft and vehicle motion path registering and analyzing system for conflicts prediction at the aerodrome movement area. Aerospace MAI Journal, 2020, vol. 27, no 3, pp. 209-2018.

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.

Borshchev Y. P., Sysoev V. K., Yudin A. D. Analysis of selective laser fusion technology application for the CubeSat nano-satellites skeleton structures manufacturing. Aerospace MAI Journal, 2020, vol. 27, no 3, pp. 219-228.

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.

Abdulin R. R., Podshibnev V. A., Samsonovich S. L. Determining load distribution unevenness ratio in ball-and-screw transmission with separator. Aerospace MAI Journal, 2020, vol. 27, no 3, pp. 229-239.

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.

Amosov A. P., Voronin S. V., Loboda P. S., Ledyaev M. E., Chaplygin K. K. Determining surface tension effect on aluminum alloy mechanical properties by computer simulation tecnhique. Aerospace MAI Journal, 2020, vol. 27, no 2, pp. 214-222.

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 σs is the surface tension, N/m, σy 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 hcr 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.

Remchukov S. S., Lebedinskii R. N. Laser technologies application specifics while plate heat exchangers developing for small-size gas turbine engines. Aerospace MAI Journal, 2020, vol. 27, no 2, pp. 90-98.

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.

Grigor'ev V. A., Ryzyvanov I. P., Zagrebel'nyi A. O. Improving parametric model of aircraft turboprop engine mass. Aerospace MAI Journal, 2020, vol. 27, no 2, pp. 81-89.

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 (Gairc,Tg).

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 Gairc ,Tg, 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 kTg coefficient, which accounts for the temperature Tg impact on the engine mass, turbine models were developed, in which the structure being changed with temperature Tg. 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 kTg.                                                                             

By approximation of integral quantitative values of  and assuming 1999 as a basic year, the expression for kimp 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%.

Pavlenko O. V., Pigusov E. A. Application specifics of tangential jet blow-out on the aircraft wing surface in icing conditions. Aerospace MAI Journal, 2020, vol. 27, no 2, pp. 7-15.

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.

Animitsa V. A., Golovkin V. A., Nikol'skii A. A. Aerodynamic design of tsagi helicopter airfoils. Aerospace MAI Journal, 2020, vol. 27, no 2, pp. 16-28.

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).

Gulimovskii I. A., Greben’kov S. A. Applying a modified surface mesh wrapping method for numerical simulation of icing processes. Aerospace MAI Journal, 2020, vol. 27, no 2, pp. 29-36.

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.

Moshkov P. A., Vasilenkov D. A., Rubanovskii V. V., Stroganov A. I. Noise sources localization in the rrj-95 aircraft pressure cabin by spherical microphone array. Part 1. Cockpit. Aerospace MAI Journal, 2020, vol. 27, no 2, pp. 37-51.

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.

Svirskiy Y. A., Bautin A. A., Luk’yanchuk A. A., Basov V. N. Approximate method for local elastic-plastic problems solving. Aerospace MAI Journal, 2020, vol. 27, no 2, pp. 61-70.

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.

Sinitsin A. P., Parakhin G. A., Rumyantsev . V. Thermal design of cathode with barium thermo-emitter. Aerospace MAI Journal, 2020, vol. 27, no 2, pp. 71-80.

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.

Ezrokhi Y. A., Fokin D. B., Nyagin P. V. Mathematical modelling application for characteristics estimation of bypass turbojet with common afterburner. Aerospace MAI Journal, 2020, vol. 27, no 2, pp. 99-111.

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 ( αΣ ≥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.

Tkachenko A. Y., Kuz'michev V. S., Filinov E. P., Avdeev S. V. Aircraft target purpose impact on working process optimal parameters and power plant configuration. Aerospace MAI Journal, 2020, vol. 27, no 2, pp. 112-122.

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.

Lokhtin O. I., Raznoschikov V. V., Aver’kov I. S. A technique for 3D-model developing of a flying vehicle with ducted rocket engine. Aerospace MAI Journal, 2020, vol. 27, no 2, pp. 131-139.

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.

Sabirzyanov A. N., Kirillova A. N., Khamatnurova C. B. Geometrical parameters effect of recessed nozzle inlet section on the flow coefficient. Aerospace MAI Journal, 2020, vol. 27, no 2, pp. 140-148.

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.

Malov D. V., Shablii L. S. The study of liquid flux coefficient dependence in axial clearance of electrically driven pump unit on operating and structural parameters. Aerospace MAI Journal, 2020, vol. 27, no 2, pp. 149-156.

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.

Baklanov A. V. Pressure losses in combustion chamber fuel system of the natural gas running gas turbine engine. Aerospace MAI Journal, 2020, vol. 27, no 2, pp. 157-168.

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%.

Ied K. ., Maslennikova G. E., Tiumentsev Y. V. Computing safe parameters of maneuver commencing of aerobatics aircraft using artificial neural network. Aerospace MAI Journal, 2020, vol. 27, no 2, pp. 169-184.

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.

Morozov A. A., Ilyukhin S. N., Khlupnov A. I. Autorotation application analysis for the safe-landing field-tests. Aerospace MAI Journal, 2020, vol. 27, no 2, pp. 185-195.

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.

Faizov M. R., Khabibullin F. F. Computations analysis of a four-link spherical mechanism for a spatial simulator. Aerospace MAI Journal, 2020, vol. 27, no 2, pp. 196-206.

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.

Golovach A. M., Dmitrieva M. O., Bondareva O. S. Structural degradation of electric arc thermal-barrier coating on gas turbine engine blades after operation. Aerospace MAI Journal, 2020, vol. 27, no 2, pp. 207-213.

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.

Busarova M. V., Zhelonkin S. V., Kulesh V. P., Kuruliuk K. A. Application of optical videogrammetry technique for normal deformation fields of aircraft fuselage panel meausring. Aerospace MAI Journal, 2020, vol. 27, no 2, pp. 52-60.

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.

Kolbasin I. V. Main sources and radiation composition affecting eigen external atmosphere of a spacecraft with nuclear power plant. Aerospace MAI Journal, 2020, vol. 27, no 2, pp. 123-130.

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.

Dunaevskii A. I., Perchenkov E. S., Chernavskikh Y. N. Takeoff-landing characteristics of regional aircraft with auxiliary retractable distributed electric power installation. Aerospace MAI Journal, 2020, vol. 27, no 1, pp. 19-29.

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 Cplanding approach and Cllanding 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 Cplanding approach value and re alized Cllanding approach value.The possibility of realizing higher (approximately twofold) Cllanding 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%.

Alesin V. S., Gubsky V. V., Pavlenko O. V. Fuselage and duct interference effect on maximum thrust of the air pushing propeller-duct thruster. Aerospace MAI Journal, 2020, vol. 27, no 1, pp. 7-18.

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.

Glushkov T. D. The study of compact fan installations with variable circulation distribution along blade length. Aerospace MAI Journal, 2020, vol. 27, no 1, pp. 30-42.

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 Deq. Exceeding the Deq 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 the hub. Overall, this limitation determines the maximum value of the axial velocity coefficient and the range of optimal design parameters of considered fans.

With account for the analysis being done, the aerodynamic designing of the experimental fan was performed and studied experimentally. The obtained results reflected the main concepts used in aerodynamic design. Significantly higher values of pressure ratio and flow rate were obtained on the experimental fan installation compared to the similar compact fan units, designed employing the classical technique for constant circulation.

Ivanov P. I., Berislavskii N. Y. Problematic issues of functioning of multi-dome parachute systems. Aerospace MAI Journal, 2020, vol. 27, no 1, pp. 43-52.

Multi-dome parachute system (MPS) represents a bundle (connected together) of single-dome parachutes. The main advantage of the MPS over single-dome parachute systems (PS) consists in the possibility of their effective employing when heavy and super-heavy loads airdrop, such as military equipment, rocket stages, etc.

Replacing one parachute with an MPS bundle allows:

-reduce the average filling time and height loss when filling the bundle compared with a single parachute of the same area;

- eliminate manufacturing and operation complexity of a large area parachute system (PS), i.e. simplify of manufacturing and operation technology of the PS; significantly simplify parachute packing and PS installing;

- increase domes stability in the bundle and stability of the load descent. A bunch of parachutes composed of unstable domes could become stable in certain cases;

- increase the PS reliability due to the redundancy;

- bring about wide unification while of serial PS development;

- conveniently place (distribute) the PS in the laid state on the airdrop delivery object.

With a view to MPS advent, a number of incompletely explored and poorly studied issues arises, such as:

1.   Why do some MPS domes adjoin each other at the steady descent, while the other do not?

2.   Why the domes are stable in some MPS, while in the other they are unstable and tend to twisting?

3.   Why in some cases the resistance coefficient of a bunch of domes is less than the one of an individual dome, and in the other is greater.

The above said, as well as a number of other issues induce performing thorough studies of multi-dome parachute systems. It was also revealed that a system stable at small perturbations of motion parameters could be unstable at large perturbations.

The experiment shows that the nature of the domes operation can change in a bundle. Stable domes in a bundle can turn out to be unstable. There were cases when unstable domes in a bundle became stable, both in the process of filling and steady descent. The system stability increases with the number of domes increasing in the bundle. It was found that employing MPS was more preferable from the stability viewpoint of descent of the object-parachute system.

With an increase in the number of parachutes in a bundle from three and more, the maximum angle of the object’s pitching practically did not change.

Fluctuations of the object-parachute system with more than three parachutes in a bundle practically independent from the parachute design.

With the number of parachutes in a bundle from one to three, the parachute design significantly affects the system fluctuations.

The article pays certain attention to the main quality indicator of the object-parachute system, namely its reliability.

To sum up, we note the following. The article briefly presents some important results of the study on multi-dome parachute systems. The following main issues were considered:

- the advantages and disadvantages of the MPS; problematic issues, which solving the MPS require;

- the problem of the leader dome and interference interaction of domes in a bundle;

- resistance coefficient and dynamic coefficient of the bundle domes;

- techniques for reducing dynamic coefficient value and aerodynamic load on the MPS due to the domes corrugation and the brake parachutes employing;

- the problem of non-simultaneous of domes filling in a bundle;

- design factors effect (extension and connecting links), as well as the number of domes in a bundle on some MPS characteristics;

- loss of height while the filling the MPS parachutes bundle;

- issues of the object-MPS system stability;

- the issue of the object-MPS system reliability.

Manvelyan V. S. Six-component rotating strain-gauge balance for coaxial rotors testing. Aerospace MAI Journal, 2020, vol. 27, no 1, pp. 53-64.

Aerodynamic strain-gauge balance is employed to study the total loads on an object streamlined by the airflow in aerodynamic experiment. As a rule, the total loads are being represented by six components, namely by three forces along the orthogonal axes and three moments around the vectors of these forces. The strain-gauge balance is a special measuring device, which operation principle is based on the strain-gauge effect. Rotating strain-gauge balance is employed to measure loads affecting rotating object.

Coaxial rotor is a system with two airscrews rotating in opposite directions. To analyze the processes while coaxial rotor operation and of airscrews interaction, it is necessary to measure loads on each airscrew, i.e. both on the one rotating clockwise and the other rotating counter-clockwise. To solve the set task two rotating strain-gauge balances were developed in Central Aero-hydrodynamic Institute named after professor Zhukovsky (TsAGI) – one for each airscrew.

All over the world, companies such as RUAG (Switzerland), NLR (Netherlands), ONERA (France), etc. are engaged in rotating strain-gauge balance development. The most common design of rotating strain-gauge balance is a monoblock of a cylindrical shape. The external rigid rim is fixed to the internal cylindrical support by the beams used to be measure the loads. The external rigid rim is coupled with the internal hub by the beams, serving to loads measuring. The external rim is coupled with the screws hub, and internal hub is coupled with the shaft of the installation, which rotates the screws. Thus, the beams, on which the strain-gauge resistors, forming the measuring bridge, are glued, are deformed, and measuring strain-gauge resistor bridges convert the beams deformation into electric signal.

One of the most significant aspects of the design is the number and shape of the beams and the scheme of strain gauge gluing. The most widespread structure includes trapezoidal shape beams at the front view, and eight beams, connecting the rim and the hub, namely, a two beams in each of four packs. The main disadvantage of such structure is low value of the signaling stress under the strain-gauge resistor, pasted for lateral force measuring, and high mutual effect of the components, which leads inevitably to higher error value (more than 1,5% of measuring range).

To avoid the above-mentioned issues, the new structure of the strain-gauge balance was developed in TsAGI. The design is similar to the one described above, but it is based on the unique shape and increased number of beams from eight to twelve, i.e. three beams in each of four packs. Computations confirmed that the signaling stress under the strain- gauge resistors pasted for lateral force measuring increased, while mutual effect of the components decreased. Alongside with other solutions, increasing the number of beams and their unique shape ensures lower value of the expected error (less than 1% of measuring range). The expected error will be confirmed by future studies on the results of static and dynamic calibration.

Bokhoeva L. A., Baldanov A. B., Chermoshentseva A. S. Optimal structure of multi-layer wing console of unmanned aerial vehidle with experimental validation. Aerospace MAI Journal, 2020, vol. 27, no 1, pp. 65-75.

The article explores stress-strain state of a composite layered wing console of an unmanned aerial vehicle (UAV). An optimal structure of the multilayer skin, ensuring maximum strength and stiffness at the specified loads was determined with the ANSYS system. The wing structure consists of two complete and two incomplete layers. Automated procedure for fiber laying angle selection in a layer was developed. Seventeen options of fiber laying angle were obtained, out of which three options of optimal reinforcing were selected. The second supplementary layer was added over the entire wing surface for deformation reduction. Thirty three options of fibers laying were considered while computing the wing model of two layers. When conputing three layers, forty seven options of fibers laying in a layer were considered. Sixty four options of fibers laying were regarded while computing a wing of two complete and two incomplete layers. According to the performed calculations, a four layer wing console was produced from layered fiberglass. It was produced by the cold forming method. Workshop drawings of tooling were developed. New tooling from phenol-impregnated modified wood was obtained for the hollow wing console fabrication, for which a Patent No 19273 was received. The weight of the hollow console is 1.46 kg, which is 3% greater than that for the computational model. The designed and fabricated wing console of the two complete and two incomplete layers weight is 43% less than that of the console of the two complete layers. Fabrication of the designed console requires 25-30% less material. The presented approach can be widely employed while structural elements and products from composite materials design and fabrication.

Aruvelli S. V. Optimal appearance determining technique of cargo parachute system at early design stages. Aerospace MAI Journal, 2020, vol. 27, no 1, pp. 76-87.

The purpose of the presented article consists in technique developing for optimal appearance determining of the gliding cargo parachute system at the early design stages according to the two optimality criteria, namely, lift-to-drag ratio and cost of the parachute system materials. These criteria reflect the facts that maximum flight range depends on the lift- to-drag ratio, and cost of materials minimization reflects the cost-effectiveness of the system. The lift- to-drag ratio to cost relationship forms the existence domain of the gliding parachute system, which facilitates the decision-making based on operation requirements and relative cost of the systems.

The problem of the optimal appearance determining is set as multidisciplinary multi-objective optimization problem based on MDF architecture and genetic algorithm. The algorithm is classified as a stochastic global search method in a mixed integer statement of the optimization problem.

As the result of the work, a technique for the optimal appearance determining of a gliding cargo parachute system at the early design stages according to the two performance criteria, namely, the lift-to- drag ratio and the cost of the parachute system materials, but with the possibility of changing and increasing the number of performance criteria, was developed.

The results of this work can be used in the parachute making industry when developing integrated computer-aided design (CAD) systems for gliding cargo parachute systems. The developed technique for the optimal appearance determining of gliding cargo parachute system can be used both in the design process of new parachute systems with improved characteristics, as well as for old structures modernization by redesigning individual elements of the system.

The technique was tested on the task of the appearance determining of the system for a payload weight of 135 kg. A comparison was made with one of several existing typical gliding cargo parachute systems of this class, which revealed that the optimized configuration of the parachute system was more cost- effective than those existing ones.

Dolgov O. S., Zotov A. A., Kolpakov A. M., Volkov A. V. Basic aspects of flap technological design with boundary layer control. Aerospace MAI Journal, 2020, vol. 27, no 1, pp. 88-99.

The article studies aerodynamic, structural, strength and technological considerations while developing a flap design with boundary layer blowing. As the result of the interdisciplinary approach, the principles of functionality and reliability ensuring of the structure were considered together with the principles that ensure its manufacturability, which allowed to highlighting the main of the technological design aspects of the flap with boundary layer blowing.

Introduction considers statistics on the number of domestic airfields and airports and performs a comparative analysis with the number of airfields and airports in the United States of America.

According to the strategy approved by the Government of the Russian Federation for the period up to 2030, the task was set to create a single transport environment for implementing high-quality competitive services for passengers and goods transportation. Given this strategy, it is obvious that regional aviation should play a leading role. Its revival is non-alternative, fastest and, eventually, the least costly way of ensuring the livelihood of the population in the regions, which corresponds to the geopolitical tasks of ensuring the integrity of Russia.

On the assumption of current situation, employing short unpaved grounds as runways may become the set problem solution.

Ensuring the feasibility of short unpaved grounds operation without their additional equipping may be possible with employing the flaps with controlled boundary layer on the aircraft.

Further, analysis of the limitations at the approach to forming the flap appearance with the possibility of the boundary layer blowing was performed.

Various design solutions implementing the impact on the boundary layer were analyzed.

The key principles for the structure manufacturability ensuring of the flap with the core in the form of regular discrete elements arranged chequer-wise have been elaborated.

Technological design aspects discussed in the article will allow the aircraft designer to design a flap with the boundary layer control, without significant increase in weight and internal stresses. Its application will allow the aircraft takeoff and landing employing ultra-short runways. It is especially relevant within the context of solving the problem of creating a single transport environment up to 2030 to ensure high- quality competitive services for passengers and goods transportation in the Russian Federation, by reviving regional aviation and re-creating local air routes in a situation of widespread reduction of the airfield and airport network.

Thus, following the above said principles, together with the requirements to the technical specifications for the product, the aircraft designer will be able to create the best technological design, which meets herewith the requirements of operational reliability and functionality.

Lashin V. S. Asymmetry parameters assessment technique while descent spacecraft design. Aerospace MAI Journal, 2020, vol. 27, no 1, pp. 100-107.

Over the past decade, the interest in Mars exploration has increased, as evidenced by the number of modern missions, both domestic and foreign, for the “Red Planet” reclamation and studying. All in all, 44 missions of spacecraft from different countries were sent to Mars. The following well-known missions can be presented as an example:

-                the interplanetary station of the European Space Agency ESA (European Space Agency), as well as the Beagle-2 lander;

-                ExoMars, which is a joint program of the European Space Agency (ESA) and the Russian state-owned corporation Roscosmos, consisting of orbital and descent (Schiaparelli) vehicles;

-                Mars Science Laboratory, which is NASA program, under which the third-generation Curiosity Mars rover was successfully delivered and operated to Mars;

-                InSigh, whicht is NASA program for the delivery of a research lander with a seismometer to Mars.

As a part of these missions, the uncontrolled descent of the spacecraft in the atmosphere of Mars was considered. The majprity of such descents ends in failure, which may indirectly indicate errors at the design stage of the spacecraft.

The presented article considers the problem of a small descent spacecraft designing that performs uncontrolled motion in the atmosphere of Mars. The task of a small descent spacecraft designing begins with selection of this spacecraft shape. It is well-known that most of the descent vehicles involved in the of the of Mars surface exploration are of a segmental-conical shape.

The purpose of this work consists in obtaining a technique for assessing permissible deviations of the spacecraft parameters, which affect the secondary resonance effects origination during descent. It is well- known that the presence of various types of asymmetry may be the cause of a long-continued resonance realization, or resonance effects. Resonant phenomena can lead to a significant increase in the angle of attack or angular velocity of the descent vehicle.

It is worth noting that the authors consider a design technique for a spacecraft with a small initial angular velocity, which it apparatus acquires due to non-ideal conditions while separation from the orbital complex. The angular velocity herewith can increase and enter a long-continued resonance under the impact of the secondary resonance.

The gist of the method consists in finding maximum values of the asymmetry parameters at which the angular velocity does not reach resonance values.

Given that at small angles of attack the derivative of the angular velocity is proportional to the generalized asymmetry parameter, we find the range of acceptable values.

It follows from the obtained scheme for the admissible values area determining that until the symmetry parameter satisfies this inequality the angular speed does not reach its resonant values by the secondary resonant effect. As a consequence to this fact, there is no realization of the long-continued resonance, which can lead to disturbances in the parachute system operation.

By applying this technique for determining the region of permissible deviations of the descent vehicle asymmetry parameters, the effects of the long- continued resonance on both angular velocity and angle of attack values can be eliminated.

Kuz’mina S. I., Ishmuratov F. Z., Popovskii V. N., Karas’ O. V. Analysis of dynamic response and flutter suppression system effectiveness of a long-haul aircraft in transonic flight mode. Aerospace MAI Journal, 2020, vol. 27, no 1, pp. 108-121.

The work is devoted to the study of aircraft aeroservoelasticity problems in transonic flight mode. Review of the state-of-the-art methods and computational algorithms used to obtain aeroservoelasticity characteristics was performed.

An agreed usage of the following approaches for the set problems solving is applied in the presented article:

– a method for unsteady aerodynamic forces computation in transonic flow using Euler equations with account for the flow viscosity,

– an algorithm for aircraft aeroelasticity characteristics computing based on the Ritz polynomial method,

– mathematical models of control systems and techniques for aeroservoelasticity problems solving in the frequency, time and root domains.

The developed methodology application has been demonstrated while the developing and studying the flutter suppression system (FSS) for medium-range aircraft with transonic cruise flight mode M=0.82 Numerical results were obtained for the airplane of conventional layout with a high aspect ratio wing and two engines located on pylons under the wing. The results of computational studies of the aircraft dynamic response were obtained employing various aerodynamic models, i.e. transonic and linear ones. The numerical studies revealed that the aircraft does not possess sufficient margins on flutter speed in transonic flight mode. For the given aircraft version the possibilities for flutter speed increase by active control system, which employed symmetrical ailerons deflection were studied. Signals from deflection sensors, located on the wingtips, were are used while FSS developing.

Gain dependence on the speed for optimal flutter 6. suppression was performed based on the frequency characteristics analysis of the open loop in the form of Nyquist locus. For each speed, the gain was selected in in such a way as ensure approximately double stability margin on amplitude. Comparison of damping and frequencies of elastic vibrations dependence on the flight speed for both open and closed loop was performed. Parametric calculations revealed that the developed FSS ensured the flutter speed increase by 45% for the first flutter form, and by 10% for the second one. Stability problem studies of the “aircraft + FSS” closed loop under the external impact. The problem was being solved in time domain.

It was demonstrated that for ensuring the closed loop stability sufficiently higher speed of aileron deflection is required.

The obtained results of the study allowed conclude that two important factors, affecting aero elasticity characteristics, exist at the transonic flow-around:

– basic stationary flow field effeect on the aerodynamic derivatives. Besides the Mach number and density, the basic flow field is determined by the angle of attack, profiles curvature and sections twisting.

– viscosity effect on the aerodynamic derivatives. These two factors are missing from the linear

methods for aerodynamic forces determining, but their regard affects significantly dynamic response of modern aircraft. Application experience of the developed approach demonstrates the possibility for effective solution of the aeroelasticity problems at transonic flight modes.

Aleksandrov Y. B., Nguyen T. D., Mingazov B. G., Sulaiman A. I. Computational grid impact on numerical computing results of three-dimensional non-stationary swirl flow behind the vane swirler. Aerospace MAI Journal, 2020, vol. 27, no 1, pp. 122-132.

The balanced design of the front-mounted device ensures combustion chamber efficiency and gas turbine engine at large. In the majority of modern gas turbine engines for ground and aviation purposes, a vane swirler is being installed concentrically with the fuel nozzle at the flame tube inlet. The swirler forms a swirl of air, and facilitates the best mixing conditions for air-fuel mixture. Besides, while the flow swirls in a low-pressure zone, its core is formed, which allows return gases fr om the flow periphery to the core of the swirled jet, forming thereby a reverse flow zone, and stabilize the fuel combustion by the stall characteristics. Increasing the swirler blades installation angle leads to intensification of the air- fuel mixture mixing, and a reverse flow zone boundaries expansion. However, hydraulic losses at the front-end device are increasing herewith, which, in its turn, contributes to the engine power or thrust reduction.

The fuel-air mixture mixing quality characterizes the efficiency of the front-end device. The majority of works by Lefebvre A., Kosterin V.A., Gupta A., Akhmedov R.B., and others suggest evaluating mixing process by the mixing coefficient, which represents the ejected air consumption to the swirled jet consumption ratio:

where m is mixing coefficient; Ge is the flow rate of the ejected air; Gsw is the flow rate of the swirling jet.

In our work, an experimental setup was developed to study the swirler mixing coefficient. Using the FMD (Fused Deposition Modeling) method of printing, various designs of the swirl with different blade swirler installation angles were created, which were blown into the open space. The flow visualization was realized by smoke pollution of the air supplied to the swirler. In the course of the experiment, both temperature and total pressure fields of the flow were measured in axial and radial directions. Temperature distributions were employed for mixing coefficient (m) computing. Bases on these measurements the coefficient was computed by the expression:

where Tsw, T0 , Ty are the temperatures in front of the swirl, in the jet and in the ambient air respectively.

A spatial computational domain, simulating the volume of the combustion chamber flame tube, was developed for numerical studies of the vane swirler. It is well known that computational grid strongly affects computation results. It is characterized by the type and number of elements; characteristic size, and the presence of near-wall thickening. The grids of three basic elements types, such as tetrahedral, hexahedral, and polyhedral, were employed. The polyhedral grids were obtained the tetrahedral grid converting. The number of elements herewith decreased by six times, and the number of nodes increased about five times, which allows compute gradients of parameters variation more accurately compared to tetrahedral due to the fact that one finite element has more nodal points. However, such a transformation does not allow precisely control the characteristic size of the elements, and a deterioration of the result due to the increase in the characteristic size of the grid element can occur.

A combined DES turbulence model (Detached Eddy Simulation) in a non-stationary setting was used for computing. The calculation was performed in the ANSYS Fluent 19.2 software with an academic license.

The performed experimental studies of flow mixing behind the scapular swirler were compared with numerical calculations using various grid models. The best results in numerical simulations were obtained when using the DES viscosity model in an non­stationary set of calculations, and hexahedral mesh elements. The polyhedral mesh obtained by converting from tetrahedral elements did not demonstrate good results, as the original tetrahedral mesh had. An increase in characteristic size of the elements led to a greater deviation of the calculated data from the experimental ones. The results obtained are valid for swirlers with various blade angles.

Ahmed H. S., Osipov B. M. Multimode identification to obtain an adequate gas turbine engine model for its diagnosing by thermal-gas dynamic parameters. Aerospace MAI Journal, 2020, vol. 27, no 1, pp. 133-143.

Modern aircraft engines are the most cost intensive, energy consuming and heavily loaded elements of an aircraft, which operate in conditions of both high thermal and power loading to ensure high economic indicators. All this requires special attention to reliability provision in flight. Aircraft engines operation as of assumes organizing technical diagnostics system at the maintenance organization, which is defined as an aggregate of means and an object of diagnostics, and performers, if necessary. This system is prepared to diagnose, or perform it according to the regulation, set by the appropriate documentation. Technical diagnostics (TD) is a division of knowledge studying technical conditions of units under test and revealing technical states, developing techniques of their determining, as well as principles of elaboration and organization of the systems application. The following tasks are related to the main tasks of technical diagnostics:

– technical condition control, which means defining the type of technical condition;

– searching for a place and determining causes of failure and malfunction;

– predicting technical condition, in which an object will appear to be at some future instant in time;

– genesis, i.e. definition of the state condition in which the object was at some point in the past;

– recognition of technical objects states in conditions of limited information to increase reliability and service life of these objects.

The engine mathematical model is of most importance in the technical diagnostics system. Its development presents a problem, since, as a rule, technical documentation does not hold characteristics of the engine units. In this regard, obtaining complete mathematical models of engines for diagnostic purposes is an urgent task. This article proposes an algorithm developed by the authors, and implemented as a computer program.

Komarov A. A., Semenenko D. A., Pridannikov S. Y., Rumyantsev . V. Magnet current impact on start-up processes of stationary plasma thruster. Aerospace MAI Journal, 2020, vol. 27, no 1, pp. 144-151.

An important characteristic of the electro-jet thruster is its start-up time. The thruster start-up time reducing requires optimization of parameters, affecting the start-up process. Cathode heater power, the value of the flow rate into cathode at start-up, the ignition pulses magnitude and duration, and the magnetic field magnitude in the acceleration channel are related to these parameters. One of the parameters that affecting the thruster start-up process is the starting level of the magnet current. The magnet current reducing facilitates the thruster start-up. However, the magnet current reduction is accompanied by the adverse factors, such as discharge current oscillations building- up upon the startup, and increasing of the inrush discharge current. The root mean square value of the discharge current oscillations herewith can reach up to 70% of the discharge current level. The article presents the results of tests on determining the magnet current impact on the processes occurring while the thruster start-up. The test objective was to define a minimum level of a magnet current, at which a thruster start-up would be accompanied by transition to a stable operating mode without the discharge current oscillations evolution. The tests were performed with the SPT-140 thruster. A special attention during the tests was paid to the changes of the discharge current oscillations and inrush discharge current surge. Oscilloscope patterns, giving an idea on the magnet current impact on these parameters, were obtained in accordance with the results of these tests. Minimum level of the magnet current at the startup, which did not lead to the discharge current oscillation evolution, was obtained in accordance with the results of these tests. The effect of the magnet current on the discharge current inrush surge level and oscillations while startup was demonstrated. It was determined that the SPT-140 thruster was proceeding to unstable operation mode at the startup with the magnet current less than 3 A. At the same time, the magnet current magnitude practically does not affect the value of the inrush discharge current surge.

Ezrokhi Y. A., Morzeeva T. A. Estimated and analytical study of the possibility to develop a bypass turboprop with afterburning chamber based on baseline gas generator. Aerospace MAI Journal, 2020, vol. 27, no 1, pp. 152-163.

Analysis of development of engines for any type of aircraft, including those with high maneuverability, reveals that all engine-building enterprises of a world level both domestic and foreign permanently perform intensive development of their engines modifications directed to improving their thrust and economic characteristics, as well as service life and reliability. The exigency for such modifications development is dictated by the necessity to support the aircraft efficiency during its life span. The main tendency for the BTAC family development is associated with employing on the basic gas generator new fans of higher productivity and pressure rate. As practice shows, such an approach may allow drastic thrust increase (more than 20%) of the upgraded engine with concurrent reduction of its specific weight. To perform evaluation, a bypass turbofan with afterburning chamber, which basic parameters are typical to multimode engine of a fourth generation maneuverable aircraft. It was believed that the upgraded engine was developed based on the basic gas generator by installing a new fan with the specified values of air consumption Ga and pressure ratio n*a .

The dependencies of the takeoff thrust, gas temperature level in front of the turbine, bypass ratio, as well as total value of pressure ratio in compressors and HPC on the new fan parameters were obtained by the results generalization of parametric computational studies. They allowed evaluate probable characteristics of the upgraded engine, being developed based on basic gas generator and a fan of higher pressure rate and productivity. Representation of the obtained dependencies in the form of nomograms allow elucidate the most probable data while analyzing information available in the open press on parameters and characteristics of foreign engines, discarding erroneous values.

The results obtained in article allow also solving the problem often occurring while the engine modernization, i.e. define parameters of the new fan, which should be installed on the original basic gas generator to obtain a preset value of takeoff thrust of the upgraded engine, as well as temperature level increase at the turbine inlet necessary for its operation ensuring.

It was demonstrated in particular that for the thrust increase by 10% under impossibility to increase air consumption through the engine (for example due to the restriction from the air intake side) the pressure rate growth in the fan should be about 30%. The required temperature rise herewith should be no less than 120-130 K. However, if the throughput margin of the air intake, which can be employed, will be at least 5%, the similar engine thrust value can be obtained at significantly lower fan pressure ratio (of about 14%) and gas turbine inlet temperature (of no more than 85 K).

The capabilities of the obtained nomograms allowing revealing a set of data discrepancies on engines available in publicly-accessible information are demonstrated on the example of the afterburning turbofan parameters and characteristics analysis of General Electric engines family, developed on the basis of the F-404-GE-400 core


Aver’kov I. S., Vlasov S. O., Raznoschikov V. V. Artificial neural networks application for experimental data analysis of composite solid propellants combustion. Aerospace MAI Journal, 2020, vol. 27, no 1, pp. 152-163.

While studying and solving the problems associated with a ramjet mathematical model developing, situations occur when a process model contains a complex mathematical formulation or a large number of assumptions. A number of experimental studies is being conducted in such cases, based on which corrections are being introduced to the model to increase accuracy of the obtained results.

The presented article regards the process of creating an electronic database of experimental studies on determination of the multicomponent combined solid propellant combustion rate, with their subsequent processing and analyzing with artificial neural networks. For gas generator and propellant consumption regulator of a ramjet operation modelling, information on combustion rate of a solid propellant is required.

Mass fractions of solid propellant components are included in the alterable variables vector. It is unreasonable to conduct experiments for all analyzed propellant compositions due to the complexity, expensiveness and long duration of their implementation. The authors suggest conducting experimental studies of particular compositions in the area under study and performing approximation by the obtained points. As the result, a function, reflecting the combustion rate behavior in dependence of the solid propellant composition and pressure is obtained.

There is a three-component propellant being a mixture of C6H2N8O4, ammonium perchlorate NH4ClO4 and a binder (rubber). The predicted parameter is the burning rate at various compositions and pressures.

The obtained topologies are built based on experimental research, and can be used later in formation of appearances of new ramjet engines.

When processing the obtained results, it is necessary to account for the fact that all experiments have certain error. The surfaces, obtained by neural networks allow identify the points at which random errors could reach high values, which is become noticeable by the function behavior.

  1. Experimental data processing using neural networks allows forming a matrix of combustion rates database in specified intervals of alterable variables.

  2. The burning rate topology analysis give grounds for analyzing the results obtained during the experiments, and, thus, to determine the experiments in which errors could be made.

Semenova A. S., Kuz’min M. V. Finite element grid discreteness selecting for rotating parts of inter-rotor bearing of a gas turbine engine considering surface roughness. Aerospace MAI Journal, 2020, vol. 27, no 1, pp. 171-179.

The presented work is devoted to the development of a technique for selecting the finite element grid size of the bearing rotating parts, contacting among themselves, with account for the surface roughness for strength calculation. It is customary in static calculation to thicken finite elements in the area of contact to ensure its accuracy. For the dynamic calculation, where parts are rotating, this technique does not work.

It is well known that reliability of machines and mechanisms operation depends substantially on their bearing blocks operability. This is especially important for aircraft engineering products as bearing blocks for aircraft engines, reducers and other products are one of the most critical components and, as a rule, limiting their resources. The inter-rotor bearing is one of the most problematic parts of the aircraft engine. While revealing signs of defect of the inter-rotor bearing the engine is removed fr om operation since this can lead to rotors jamming and the engine failure. The main cause of the rolling bearings failure under normal conditions is occurrence of contact stresses and, consequently, the rolling surfaces wear-out.

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

– no friction;

– the contact area is smaller compared to the curvature radius;

– the contacting bodies materials are homogeneous, isotropic and perfectly elastic.

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

– friction simulation;

– accounting for nonlinear properties of the material;

– accounting for the contacting surfaces roughness by selecting finite element grid size.

The developed technique allows estimating stresses and deformations of the rotating parts of rolling bearings of any shape.

The purpose of the presented work consists in determining the optimum size of finite elements for dynamic calculation wh ere the contacting parts are rotatting.

Comparative evaluation of stresses and strains in contact of rollers with raceways of the 5AV1002926R4 bearing in 2D statement of the two options was performed:

– the size of a grid was selected with account for the surface roughness of the contacting bodies;

– the grid size was reduce by half compared to the first option.

The grid discreteness evaluation was performed with the LS-DYNA software package.

The developed technique is suitable for all types of planar and solid-state finite elements.

Biruykov V. I., Kurguzov A. V. Forming cyclogram of energy-propulsion system for prospective inter-orbital space transportation vehicle with electric propulsion and liquid stages. Aerospace MAI Journal, 2020, vol. 27, no 1, pp. 180-190.

At present, liquid rocket-thrusters are employed mainly as cruise engines for inter-orbital transportation means. These engines efficiency is limited by the energy capability of the fuels being used. Electric propulsion application, in which reactive mass and energy source are separated, is seemed promising. Due to the high exhaust velocity of the reactive mass, the electric propulsion employs reactive mass an order of magnitude higher efficiently than the chemical one.

The available limitations of the power source energy and high specific impulse allows the electric propulsion ensure only insignificant thrust, which limits the scope of its application. That is why more often chemical and electric rocket engines are used conjointly. Transportation is performed firstly by the chemical stage, then it is separated, and finishing is executed by the electric propulsion stage.

It is necessary to validate scientifically parameters selection for the energy-propulsion system and electric propulsion stage of the prospective inter-orbital transportation vehicle. To do this, criteria, characterizing the effectiveness of transportation operation performing, obtaining at the specified input parameters of the energy-propulsion system is required. Some of these criteria can be obtained analytically, while the other by the simulation results only. Thus, a technique allowing planning cyclogram of the transfer with specified input parameters, this planning validation, and obtaining trajectory information, based on the cyclogram, which allows evaluate space factors impact, depending on location, and the effect of radiation of the Van Allen belts.

The article proposes analytical dependencies, on which basis cyclogram of the transfer from the low near-Earth orbit to a geostationary orbit can be formed. The flight is performed by the super–synchronous highly elliptical orbit. The energy- propulsion system of the vehicle consists of chemical and electric propulsion stages. The liquid stage puts the payload, consisting of electric propulsion stage and target spacecraft, on the super-synchronous geot–ransition orbit, and separates. Further finishing is performed by the electric propulsion. The power source are solar batteries with the preset power.

To verify correctness of the cyclogram analytical construction, a random set of points is formed in the studied space of the input parameters. For each point, a propulsion system cyclogram is generated, and numerical simulation is performed. Deviation of the last trajectory point from the radius, specified while the cyclogram construction, is evaluated. Dependencies of the volume of trajectory information on the input parameters are formed. Based on the results of the study, a conclusion was made that the proposed technique for cyclogram generating of the transfer can be employed when selecting design parameters of the energy-propulsion system of a perspective inter-orbital transportation vehicle.

Varsegov V. L., Abdullah B. N., Axial turbine blades geometry impact of small-sized turbojet engines on the turbine efficiency. Aerospace MAI Journal, 2020, vol. 27, no 1, pp. 191-200.

Small-sized turbojet engines are employed for unmanned aerial vehicles (UAV). Due to low efficiency and thrust-to-weight ratio, they are limited to short range applications. However, transition from rated idle mode to MAXIMAL mode at high altitude takes time, which requires further development to improve efficiency of these gas turbines.

When creating promising small-sized turbojet engines, the problem of turbines gas-dynamic efficiency increasing inevitably arises, as it directly affects the fuel efficiency of the engine, and ultimately determines its competitiveness.

The presented article considers profile losses, i.e. the flow separation from the surface of the rotor blade profile. The issue of the setting angle βset and the angle at the rotor blade inlet βx effect on the turbine efficiency is under consideration.

The main task of the calculation consists in determining optimal shape of the axial turbine rotor blades to ensure the required parameters and characteristics of the turbine at continuum flow and minimum energy losses with specified values of the angles at the inlet and setting angles.

The article presents also the results of a numerical study of the turbine air-gas channel, i.e. the joint operation of the turbine guide blades and the rotor blades, to assess the quality of the rotor blades geometry to improve the turbine efficiency.

In this work, the 3D computational model was constructed in the SolidWorks program with subsequent computational grid applying with Turbo Grid program. The flow was simulated by the SST turbulent viscosity model.

Kochubei A. A., Vernigorov Y. M., Demin G. V. Physico-technological basics of aircraft long parts hardening in the devices with rotating magnetic field. Aerospace MAI Journal, 2020, vol. 27, no 1, pp. .

The article gives an account of the studies of hardening treatment of long thin-walled parts employing imposition of magneto-dynamic effect. It presents characteristics of movement of the ferromagnetic indenters moving freely in rotating magnetic field (REMF) and thermodynamic model, which determines energy characteristics of ferromagnetic indenters moving freely in REMF. The article describes characteristics of its impulse function on the processed surface, as well as the degree of their effective loading. It presents analytical dependencies, allowing objectively ensure prediction of the surface layer parameters of quality while its forming, and productivity of magneto-dynamic hardening treatment. A technique for technological process developing of parts treatment operation with magnetodunamic effect imposition. Recommendations on the design of devices with REMF, as well as technological outfit means, allowing enhancing efficiency of their employing in the parts hardening treatment technology, are given.

The purpose of the study consists in developing a hardening treatment technology by surface plastic deformation of long thin-walled parts with magneto­dynamic effect imposition and practical recommendations on its application.

The following conclusions were made by the results of the conducted study:

1.   The rotating electromagnetic field application as an energy source of the freely moving ferromagnetic indenters is the basis for developing and improving of a new method for parts hardening treatment, called magneto-dynamic processing.

2.   Magnetohydrodynamic treatment enhances technological capabilities of hardening treatment by freely moving indenters, and ensures efficiency increasing of finishing-strengthening treatment of the inner cavities of long thin parts.

3.   Technological effect of the magneto-dynamic processing is stipulated by the motion of a large number of ferromagnetic freely moving indenters, placed into the REMF, forming in gross amount a magneto-liquefied moving layer. This layer interacts with the surface layer of the processed parts, being the result of the effect on each ferromagnetic freely moving indenter of the whole row of forces and moments.

4.   It was proved that for stable magneto­liquefaction process of the rotating layer both input and dissipated energies should be set equal in such a way that the magneto-liquefied moving layer would transfer from liquefied phase to a hard one under condition when the REMF induction would be less than 0.08 Tl.

5.   Based on the energy balance modelling the dependency for energy characteristics evaluation of ferromagnetic indenters freely moving in the REMF was obtained. It allows substantiate the force conditions of the shock-pulse impact, which ensure plastic deformation in contact zone of indenter with the processed surface and, as a consequence, the hardening effect development.

6.   The nature of the energy-force action of indenters on the processed surface layer depends on the degree of their constricted state in the MRF layer. It was confirmed experimentally that the loading quantity of freely moving ferromagnetic indenters, which formed the MRF layer, into the processing chamber of the device should not exceed three concentrically arranged layers, commensurable with the indenter length.

7.   Based on theoretical and probabilistic representations, the dependence allowing predicting duration of the magneto-dynamic hardening treatment and correspondingly evaluate the process productivity was obtained.

8.   The presented analytical dependencies for determining quality parameters of the surface hardened while magneto-dynamic method processing determine with adequate fidelity the effect of energy condition and size of ferromagnetic indenters, the initial state of the surface geometry, as well as mechanical properties of the material, subjected to the treatment, on their formation. The results of the studies demonstrate that the presented analytical dependencies can be employed while developing magneto-dynamic parts hardening treatment technology with with an accuracy to 10-15%.

9.   An algorithm, determining technological conditions of treatment was developed. Recommendations are given on embodiment of the devices with REFM , on which basis formalization of operations for hardening procession by magneto­dynamic method are possible. It contributes to effectiveness enhancing of the production planning process employing CAD TP

10. Application of feed-through type installations, realizing magneto-dynamic processing method compared to the existing hardening technology with UPD-2.5 allows significantly decrease both energy and materials consumption of the equipment, reduce technological processing time, decrease auxiliary time on parts setting and, thus, increase the productivity of hardening process, ensuring herewith quality parameters of the surface layer, regulated by technical requirements.

Sergeev S. V., Al-Bdieri M. S., Dubrovina N. A. Surface modification of the AK12MMGH aluminum alloy by micro-oxidation technique to improve operating characteristics. Aerospace MAI Journal, 2020, vol. 27, no 1, pp. 217-223.

Coatings formed by micro-arc oxidation on aluminum alloys have a unique combination of properties such as high heat resistance, wear resistance, adhesive strength and corrosion resistance. This combination of properties is largely stipulated by the nanocrystalline structure, which, according to a number of studies, is represented in the MAO-layers by small-scale pores and crystallites with sizes not more than 100 nm.

For modifications employing MAO the AK12MMGH aluminum alloy was selected. Oxidation was performed in an alkaline electrolyte with addition of liquid glass. Capacitors capacity of MAO installation, was-78 pF, except for the mode of the sample No 3, when MAO was being performed at 100 pF. This was done to significantly reduce the processing time and increase the coating thickness. The processing time т was determined by the process intensity decrease (arc discharge occurrence on the ribs).

Samples No. 1 and No. 2 have the thinnest coating. This is associated with the lower concentration of liquid glass. The thickest coating was formed on the sample No. 4, due to the increase in the electrolyte concentration. Despite this, being compared with the sample No. 5, it has a more porous technological layer. The same as samples No. 4 and No. 5, sample No. 3 has a thick coating. In this case, it is stipulated by the fact that capacitor capacitance increase of the MAO installation led to the increase of micro-arc discharges, and, as a consequence, the volume of reaction products, formed per unit time, increases.

The surface modification of the AK12MMGH aluminum alloy by micro-arc oxidation method allowed that formed coatings had a layered structure intrinsic to MAO-coatings of aluminum alloys. The installation capacitor capacitance increasing steps up the MAO process intensity, which leads, in its turn, to the number of electrochemical reaction products build-up, and, as a consequence, to the thicker coatings forming.

The cross-section study revealed that porosity is characteristic only for the outer technological layer. The MAO installation capacitors capacitance increasing helps the porosity reduction. Hardness measurement revealed heterogeneity of mechanical properties of MAO coatings in thickness depending on the phase composition and the presence of defects.

Dmitrieva M. O., Golovach A. M., Sotov A. V. Hot isostatic pressing impact on samples structure grown of Inconel 738 super alloy by selective laser melting technique. Aerospace MAI Journal, 2020, vol. 27, no 1, pp. 224-232.

Selective Laser Melting (SLM) is an additive manufacturing technology intended for metal powders fusion by the high-power laser. Powder materials application ensures in this case more steady chemical composition over the product cross-section and zonal segregation absence.

One of the most important and complex trends in this technology consists in heat-resisting alloys powders application, since this particular is employed for the most critical parts manufacturing. Among the SLM technology benefits are the following:

– the possibility of manufacturing parts of any configuration complexity;

– the possibility of simultaneous growth of several samples;

– high materials utilization ratio, and products prototyping simplification

Disadvantages of the technology under consideration include the presence of residual porosity, restrictions on the employed materials and laser radiation sources s, as well as sizes of the products being fabricated.

The hot isostatic pressing (HIP) technique is applied to eliminate residual porosity. It consists in processing a part, set in a special capsule, by the gas pressure about 100-200 MPa at elevated temperatures. The purpose of the presented research is studying the HIP impact on the samples structure, grown of heat resisting Inconel 738 alloy by the SLM technique.

The samples being studied were fabricated on the SLM 280L installation for selective laser fusion of metal powder. They were synthesized both perpendicularly and at the angle of 45 degrees to the substrate at the laser radiation power of 325 W. The samples were being subjected to the HIP in the gas thermostat. After etching, the studies of microstructure were conducted with METAM LV-31 metallographic microscope. Electron-microscopic analysis of the samples and original powder material was performed with TESCAN Vega SB electron-scan microscope.

Chemical composition of the original powder material was being determined by INCAx-Act energy dispersive X-ray spectroscope. The microstructure analysis was performed with NEXSYS ImageExpert Pro 3 image analysis program. X-ray microanalysis revealed that chemical composition of the original powder of the heat resisting alloy complies with the Q/AMC 4-2-10­2018 certificate.

Original powder substance chemistry researched on an INCAx-Act energy dispersive X-ray spectroscope. Microstructure analysis was carried out using the NEXSYS ImageExpert Pro 3 image analysis program. X-ray microanalysis showed that the original powder substance chemistry corresponds to the Q/AMC 4-2-10-2018 certificate.

The results of electron-microscopic analysis of the original material allowed revealing that the powder particles were spherically shaped, characteristic to the technique for molten dispersing. Metallographic analysis of the sample grown vertically to the substrate at the laser radiation power of 325 W allowed establishing that microstructure represents an aggregate of fused powder particles, which were micro-ingots of the dendrite structure. After the SLM process, the microstructure of the sample cross-section is characterized by the defects such as micro-cracks. The microstructure of the sample cross-section, grown at 45 degrees to the substrate, is characterized by the presence of the same defects, but differs by their larger outstretch.

Metallographic analysis of the samples after HIP revealed that the structure defectiveness after the post processing decreased. Since the products were subjected to HIP without setting into the special capsule, healing of defects could not be attained. All surface defects remained in full, and internal ones reduced by the cross-section. The ineffectiveness of HIP application in this case is explained by the presence of chrome dioxide on the surface of powder particles, having formed under the impact of high temperatures while fusing.

Thus, the HIP technique application allowed decrease the structure defectiveness, due to micro cracks size reduction along the cross-section, but the full healing of defects was not attained. HIP effectiveness increase in this case is possible by placing the samples into the special airtight shell, and excluding chrome oxides forming on the powder particles by excluding metal-with-oxygen contact during the entire technological process.

Bodunov N. M., Khaliulin V. I., Sidorov I. N., Kostin V. A. On preform impregnation process simulation while transfer molding of composite products. Aerospace MAI Journal, 2020, vol. 27, no 1, pp. 233-245.

This article envisages an analytical approach to transfer molding simulation as applied to production of articles from composite materials. Navier-Stokes equations, modified by Brinkman, with corresponding initial and boundary conditions are used to describe the flow of incompressible liquid through porous media for two-dimensional unsteady and steady flows. The authors suggest a numerical-analytical method based on the sought solution approximation by linear combination of polynomial basic functions for the flow velocity components. This method novelty consists in selection of generalized variables and finding concrete basic functions, which in some cases allow obtaining analytical solutions, identically satisfying the initial equations, and reducing non-linear boundary problems in other cases. The unknown coefficients contained in the found solutions are determined from the corresponding initial and boundary conditions by the collocation method, or weighted residuals method while solving concrete applied problem.

Partial analytical solutions of Navier-Stokes equations, describing a slow flat flow of a viscous liquid, which basis is formed by the polynomial solution of the linear bi-harmonic equation, were found without accounting for the inertial forces. The external parameters included into solutions are being determined from boundary conditions by the collocation method, or weighted residuals method, while internal parameters, expanding the class of solutions, are selected from mathematical and physical reasons, as well as comparing theory with experimental data and other exact solutions. These solutions can be employed for describing slow flow of a viscous liquid through the porous medium. Approbation of the obtained partial analytical solutions was performed on the examples of solving two test problems, i.e. the problem of a plate flow-around, and Couette problem on liquid flow movement located between two planes under the impact of the pressure difference, whereas one plane is immovable, and the other moves at constant speed. Computational results demonstrated acceptable accuracy of the obtained solutions.

Gorbovskoi V. S., Kazhan A. V., Kazhan V. G., Shenkin A. V. Numerical studies of nozzle thrust characteristics of supersonic civil aircraft by computational gas dynamics method. Aerospace MAI Journal, 2019, vol. 26, no 4, pp. 7-16.

One of the most urgent problem while developing a new generation supersonic civil aircraft is ecologic requirements ensuring, including the community noise level near the airport. It requires developing and studying new technical solutions ensuring both low nozzle thrust losses level at all flight modes and reduction of jet flow velocity to decrease its noise level at the take-off/landing modes. One of the possible trends for this problem solving is mixer-ejector type nozzle application on the supersonic civil aircraft. Its operation principle consists in the fact that at the take­off mode with sound absorption, the hot jet is split into smaller jets by the multi-lobe nozzle. The increased surface area of the ruffled jet intensifies its mixing with atmospheric air, and reduces the length of the mixing layer initial section. The mixed jet velocity in the nozzle outlet section reduces, and thus the effect of acoustic suppression is achieved. Mixing zone shielding by the tail part elements of the airframe allows additional enhancing of acoustic suppression. At the flight modes without acoustic suppression the mixer- ejector type nozzle transforms into conventional supersonic nozzle with much higher thrust characteristics.

To reduce time and financial costs at the preliminary design stages, it is expedient to employ computational methods, ensuring high level of confidence. Modern software for fluid numerical modelling are applicable for solving a wide class of problems. However, refinement of flow numerical modelling technique is necessary for solving concrete problem. In the presented work, rational parameters for computing and computational grids were selected.

Modern Computational Fluid Dynamics (CFD) software allows solving a wide class of problems. However, refinement of flow numerical modelling technique is necessary for solving concrete problem. In the presented work, rational parameters for computing and computational grids were selected to study physics of the flow and obtain integral characteristics of the nozzle, such as mixer-ejector nozzle, at the take-off, landing, transonic and supersonic flight regimes. This method is employed to predict the nozzle thrust losses with ANSYS CFX commercial CFD code of Reynolds- averaged Navier-Stokes equation numerical solution. The numerical study of losses in mixer-ejector nozzle with active system of acoustic suppression at the take­off and landing modes are performed, and obtained results are validated by the experimental data. The accuracy of validation does not exceed 0.5% of the ideal thrust losses at all flight modes.

Murav’ev V. I., Bakhmatov P. V., Grigor’ev V. V. Specific defects forming features while aircraft bulky titanium structures assembling. Aerospace MAI Journal, 2019, vol. 26, no 4, pp. 17-27.

This article presents the results of the study of specific defects forming while VT20 and VT23 titanium alloys electron-beam welding. It was established that the presence of capillary-condensed moisture, resided in the defects of the edges’ surface, impacts dominantly on the submicropores formation. Other conditions electron-beam welding conditions, which may lead to specific defects forming, were revealed. These conditions may include:

  • Improper assembling and preparation of the abutting edges for welding;

  • Electron-beam welding modes;

  • A solid-phase joint formation prior to the front of the molten bath;

  • Oscillatory processes of the electron beam (~0.5 mm), which may lead to uneven melting (due to insufficient temperature of the edges’ overall melting) over the grains boundaries with submicropores forming (less than 0.00025 mm), which cannot be detected by modern X-ray machines;

  • Hydrodynamic collapse of the crater leading to the root defect generation as peak-shaped formations.

It was revealed by radiographic control and scanning microscopy that defects in the form of dark stripes represented the chains of submicropores projected onto each other. It was established also that specific defects formed while electron-beam welding impacts significantly on the strength properties of welded joints, as well as on their destruction stadiality. The performed studies allowed make a conclusion on the necessity of monitoring such basic factors as the surface quality of the abutting edges for welding; electron beam focusing conditions, its power and oscillatory processes; and hydrodynamic instability in the weld penetration channel.

Moshkov P. A. Problems of civil aircraft design with regard to cabin noise requirements. Aerospace MAI Journal, 2019, vol. 26, no 4, pp. 28-41.

The presented work is devoted to the problem of modern aircraft design with classical power plant layout, i.e. two turbofan engines on pylons under the wing, with account for the cabin noise requirements. The objective of the work consists in developing the list of scientific research and development activities, which execution is necessary for an aircraft design by the specified parameters of acoustic comfort.

The article considers the problem of noise level normalization in the aircraft cabin and cockpit. The main sources noise in the cabin were determined based on SSJ-100 aircraft testing. To minimize their sound pressure levels in the cabin a list of works while civil aircraft design was developed.

Determining relative contribution of various sources to the total sound pressure level along the cabin length, measured with the A-weighted scale of a standard noise level meter, is necessary for the right selection of methods and means for its reduction. The main sources of noise in the cabin and cockpit are the systems for air conditioning and ventilation, as well as pressure pulsation fields in the boundary layer on the aircraft fuselage surface.

Noise from the engines vibrational impact does not appear to be significant while evaluating total noise level in dBA. Acoustic radiation of the power plant, such as ventilator and jet noise, does not affect total levels of sound pressure weighted by A scale of a standard noise level meter in the cabin and cock pit at the cruise flight mode. The sound of aircraft avionics is not a significant source. But it can be said in general that placement of aircraft equipment systems aggregates should be executed with account for their acoustic characteristics.

The noise level they create in the cabin should be 10-15 dBA lower than the calculated sound pressure level in the cabin of the aircraft under development, determined at the control point of the cabin as the energy sum of noises from air conditioning system and turbulent boundary layer.

The results of this work can be used in the design of modern civil aircraft, with regard for the requirements to acoustic comfort.

The cabin noise problems of civil aircraft was considered. It was shown, based on the SSJ-100 flight tests that the dominant sources of noise in the cabin were the turbulent boundary layer and air conditioning system. The main directions of scientific and research activities, necessary for the aircraft design according to the specified parameters of acoustic comfort were formulated for these two main sources. Basic methods for noise reduction in the cabin were considered.

Valitova N. L., Kostin V. A. On probabilistic methods application to solving aircraft strength inverse coefficient problems. Aerospace MAI Journal, 2019, vol. 26, no 4, pp. 42-50.

Solving problems of static strength, fatigue resistance, and aeroelasticity can be performed in both deterministic and probabilistic formulation. Deterministic approach for aircraft strength computing is adopted as the basic one both in this country and abroad. Aircraft safety requirements increasing leads to the necessity of considering probabilistic safety criteria and development of normative standards for them.

The article deals with solving the inverse strength problems in a probabilistic setting in a general form. In the most general case, the elements of the “output”, as well as parameters of the structure under study, characterized by a certain operator, are stochastic. It is assumed that the probabilistic measure of the “output” is known and can be defined in the form of theoretical distribution law. In this case, the inverse strength problem in probabilistic setting is reduced to either determining the probabilistic measure of parameters of the “input” (at the determined parameters of the “object”), or to determining the probabilistic measure of the “object” parameters. It is assumed initially, that the problems under consideration are quasi-static, and unique deterministic dependence between the “input” and the “output” is known.

Examples of linear transformations for random variables are given when determining probability characteristics of load restoration and identification of structures for the two models, namely a beam and a thin-walled Odinokov’s structure.

Further, the article presents methods for analyzing static systems with random parameters. The real structural elements parameters randomness is being caused by the external environment disturbing effects, unavoidable technological production errors etc. It manifests in the form of cracks, starved spots, initial irregularities and other factors, which may affect the structure behavior in various ways. In particular, destruction may be associated with a large number of dislocations and stresses redistributions. This allows expecting non-linear manifestations in the structure material behavior in the form of hysteresis loops, leading in general case to non-Gaussian distribution of random values.

When considering static systems hereafter, an internal random value (e.g. crack) is being interpreted as an additional random impact at the deterministic system input. This affects the system behavior and leads to natural mixing of random output processes while their transformation in the system, i.e. the effect of natural formation of mixture of distributions.

The examples of determining the probability density for the potential energy dissipation of the rod deformation at random thermal effects, as well as functions of the mixture density in the presence of the internal defect in the beam were considered.

The obtained material can be recommended for developing a base of standards on mixtures’ references necessary for the purposes of structures diagnostics.

Amir'yants G. A., Malyutin V. A., Soudakov V. G., Chedrik A. V. On strength and aeroelastic characteristics of a large-scale model of an airplane wing section. Aerospace MAI Journal, 2019, vol. 26, no 4, pp. 51-65.

The article presents the computational and experimental results of aeroelasticity issues studies accompanying design and testing in wind-tunnel of a large-scale model of a passenger aircraft-demonstrator wing element the 7-th European framework program AFLoNext. The goal of the project consists in developing advanced flow control technologies for new aircraft configurations to achieve a quality leap in improving their aerodynamic performance.

Design, manufacture and assembly of a large-scale model, which serves for visual presentation of typical phenomena of flow separation in the fixation area of the wing with engine with high degree of bypass, were performed. However, such engines application on arrowhead wings causes undesired phenomenon of flow separation on the wing at low speeds and high angles of attack, which may lead to deterioration of the aircraft overall aerodynamic characteristics. To avoid these phenomena, the two newest types of technologies for active flow control are studied within the framework of the project. The pipe tests of the model were performed on the aerodynamic balance of the ADT-101 TSAGI pipe.

Based on the developed demonstrator CAD-model, detailed mathematical model of a demonstrator was built to compute the strength and safety of the pipe tests. Preliminary calculations of the structure stress- strain state indicated the need to strengthen the attachment area of the caisson spar to the beam of the supporting device. Comparison of natural frequencies and shapes of the first tones of mathematical model oscillations with the results of ground frequency tests was performed prior to testing. The difference between experimental and computed natural frequencies of the first oscillation tones did not exceed 10%.

Analysis of the structure behavior in the flow revealed the most loaded elements, in which minimum safety margin was η = 3, which corresponds to the ADT-101 TSAGI requirements. To control the nacelle and slat oscillations at the start-ups, computation of overloads limit values on nacelle and slat for understated strength margin of η = 2 with reference of the “stall” phenomena and turbulence was performed.

Critical flutter and divergence speeds were determined for ensuring safety of the demonstrator mathematical model tests performance in the pipe. The obtained values were out of the bounds of the velocities realized during the tests.

High measurements accuracy of the wing flow control systems efficiency was ensured by a comparative analysis of the local angles of attack of the structure under the impact of the ADT flow.

Kolyshev E. S., Krapivko A. V. Experimental methods for determining dynamic characteristics of aircraft landing gear. Aerospace MAI Journal, 2019, vol. 26, no 4, pp. 66-80.

The article describes methods and algorithms for determining the fundamental eigen modes of landing gear, such as torsion, lateral and longitudinal bending of support, according to the amplitude-phase frequency characteristics measured at characteristic points of the structure. Resonant frequencies, shapes and decrements of vibrations are determined using transfer functions (dynamic compliance and dynamic stiffness). A typical accelerometers arrangement of a system for oscillations registering and arrangement of vibration exciters are given. The described methods for obtaining dynamic characteristics were developed based on the long experience in landing gears GVT of various aircraft.

The novelty in landing gear GVT is marked:

  1. Moveable carriages with vibration exciter mounted on them, which are equipped with special connecting devices for attaching rods to the axis of wheels. The rods are equipped with forces sensors transmitted to the structure, in order to eliminate the excitation system effect.

  2. The GVT is performed for the landing gear both in a free state and at various vertical loads on supports created from action of the aircraft mass by hydraulic lifts.

  3. The applied shock method application on landing gear to obtain amplitude-phase frequency characteristics at the selected points of structure according to the results of response functions processing. This method allows giving an operational evaluation of the landing gear resonant characteristics and speed up the ground frequency testing procedure.

  4. The GVT results processing is performed using transfer functions of dynamic compliance and dynamic stiffness of landing gear strut for bending and torsion and their cross links.

  5. To determine hydraulic lifts effect on landing gear dynamic characteristics, the GVT in a free state is performed in cases when the aircraft is installed on the standard hydraulic lifts and when the aircraft is installed on pneumatic supports.

Parakhin G. A., Rumyantsev . V., Pankov B. B., Katashova M. I. Low-current cathode designing for small stationary plasma thruster. Aerospace MAI Journal, 2019, vol. 26, no 4, pp. 81-89.

At present, the interest of spacecraft producers to low-power electric propulsions and propulsion installations on their basis is growing. The above mentioned fact imparts topicality to the task of expanding the family of cathodes for such thrusters towards decreasing discharge current maintained by the cathode.

It is well known, that effective cathode of the electric propulsion does not require any additional heat source in a steady-state operation, and thermoemitter operating temperature maintaining is ensured by the ion current on its surface. This article describes two complementary trends of works aimed at such cathode designing.

The first trend consists in the cathode thermal scheme optimization and thermal losses reduction. Some of design solutions, related to this field of work, were employed in the cathode experimental design and demonstrated their efficiency. On the other hand, the optimized design appeared to be sensitive to the smallest changes in the thermal scheme and, thus, needed a retrofit.

The second trend is a development and application of new thermal emissive materials with a lower operating temperature. The article presents the results of the works which have been in progress with some intermittences since 2013. The article demonstrates the results of Barium oxide-based thermoemitter samples developed and tested at EDB Fakel. The issues of thermoemitter manufacturing procedure; raw materials (powders) purity and dispersity; sintering temperature, and tool set, developed in the course of the works, are tackled.

As the result of handling of work, the authors came to a conclusion that for a higher efficiency of the new cathode design being developed it is necessary to consolidate the results of works in both trends. Further additional measures for the design optimization are planned.

Gol'berg F. D., Gurevich O. S., Zuev S. A., Petukhov A. . The onboard mathematical model application to control gas turbine engine with extra combustion chamber. Aerospace MAI Journal, 2019, vol. 26, no 4, pp. 90-97.

Modern gas turbine engines control is performed by the parameters accessible for measuring, which for the most part characterize indirectly the engine critical parameters such as thrust value R, specific fuel consumption CR, as well as parameters, affecting directly operational safety and reliability, such as gas temperature  in the combustion chamber (CC), stall margin (ΔSm) etc.

Employing the all-modes self-identified thermo­gas-dynamic model of the above said engine in modern digital automatic control systems (ACS) offer scopes for new opportunities of substantial control quality enhancing. This model allows computing with high precision the engine critical parameters in real-time scale, and realize the engine control directly by these parameters.

The article presents the results of studying such methods for controlling the fuel consumption GFE into extra combustion chamber, and nozzle throat area FT of the multi-mode engine.

The scheme of structural and algorithmic construction of such system is introduced.

Implementation of the three control programs, such as thrust changing RΣ depending on throttle position, and minimum  and maximum  values limiting of the air-to-fuel ratio αECC in the extra combustion chamber is being accomplished by affecting the fuel consumption (GFE).

Ensuring the minimum possible value of the specific total fuel consumption C = (GFM + GFA )/RΣ) , as well as restriction of fan stall margin, are implemented by affecting nozzle throat area by the extremal controller.

The effectiveness evaluation of the control methods under consideration was brought about by the integrated mathematical models “Engine – ACS – Onboard Mathematical Model” employed in CIAM.

It was shown, that direct engine thrust control by the impact on fuel consumption into the extra combustion chamber allowed ensuring the thrust value invariance to the engine components degradation while in operation.

The impact on the nozzle throat area herewith minimizes specific fuel consumption and limits the fan stall margin.

Baryshnikov S. I., Kostyuchenkov A. N., Zelentsov A. A., Lukovnikov A. V. Effectiveness estimation of turbo-compound scheme application on purpose of indicators increasing of aircraft piston diesel engine of 300 H.P.. Aerospace MAI Journal, 2019, vol. 26, no 4, pp. 98-107.

The goal of the presented work consisted in improvement of the engine basic indicators — specific power and effective specific fuel consumption (ESFC). This goal achieving is possible though three methods, based on a heat balance equation, namely, effective power increasing, as well as heat emission decreasing into cooling system and exhaust energy utilization. Effective power increase seems to be a conservative method that ensures relatively low performance increase, and is the main research trendHeat removal limitation to the cooling system was actively studied in 90-s, and currently considered unworkable. Thus, the best way to increase the engine indicators radically is the exhaust gases energy utilization.

There are many ways realization, including mechanical and electric compounding, the Renkin cycle application, thermoelectrical generators. However, the most efficient way from the niewpoint of specific parameters is mechanical compound.

Historically, turbo-compounding is a logical continuation of turbocharging. Turbo-compound engines are the pinnacle of aviation piston engines. VD-4K and Napier Nomad engines represent the examples of such engines, demonstrating at that time the unsurpassed fuel efficiency levels.

A six-cylinder boxer four-stroke turbocharged CI water-cooled engine was selected for the purpose of this study. The key factor for the diesel engine selection was the high air to fuel ratio, which was about two times higher than this for the gasoline engine. Owing to this, other things being equal the compound turbine will ensure twice as much power.

In this work, identification of the basic engine was being performed with the AVL BOOST software. The Patton, Nitschke, Heywood friction model, allowing determine friction losses based on the engine arrangement; Vibe combustion model, and Woschni 1978 heat exchange model were employed. Based on the obtained model a turbo-compound modification was developed. Optimization of basic parameters, such as charge pressure, pressure drop on both power and compressor turbines, gas distribution phases and ignition advance angle.

Based on the obtained results, a comparison of three variants of the engine, such as basic one; with the Garret turbine, which roughly corresponds to domestic prospective turbines; and the one with reference turbine was performed.

As a whole, the achieved results fit theoretical estimations with high degree of precision, with the exception of the exhaust gases temperature: contrary to the initial expectations, the temperature decreased. However, this result fits the pattern, established in other authors’ works.

The results of the comparison revealed that the power increment in the turbo-compound engine could achieve 10%, and ESFC reduction could achieve 11%.

Kiselev F. D. Fracture diagnostics and operational workability evaluation of working turbine blades of aircraft engine. Aerospace MAI Journal, 2019, vol. 26, no 4, pp. 108-122.

The topmost constituent part of the study on determining the cause parts of destruction of the aircraft in operation is fracture diagnostics employing the methods of physics-of-metals analysis of the fracture structure, material structure and composition determining, defect detection control, mechanical properties characterization, parts strength and survivability analysis.

Diagnostics of aircraft turbine blades operational fractures was performed, factors contributing to destruction were revealed, and causes of blades destruction were established. The article considers operational damageability specifics, on frequent occasions differing from the test bench ones, the systematization results of loading types, fracture mechanism, and operational fractures of gas turbine engine blades.

Methodical aspects were developed and new techniques were elaborated for fracture diagnostics were developed. The article systematizes external, fractographic and metallographic signs of diagnostics characteristic to anomalous (abnormal) modes of the engine functioning and a blade fracture at normal aircraft engine functioning (operating parameters did not outrun the operational limitations). The suggested classification allows determining blades fractures while operative diagnostics with account for joint action of static, vibration and thermal stresses in the blade material. It helps identifying blades fractures by the operational fractures types and revealing thermo­loading factors, determining the fracture mechanism, outlining it from all set of mechanical and thermal loadings acting on the blade.

The article presents the results of experimental studies of cyclic crack resistance of the blade made of VZHL12U (equiaxial crystallization) and ZHS26, ZHS32 (directional crystallization and single-crystal version correspondingly) alloys. Characterization of the blades material resistance to fatigue destruction with kinetic diagrams plotting (dependence of the crack growth rate on the stress intensity factor) was performed at the temperature of 850°C with samples loading on the vibro-bench. Eigen oscillations frequencies of the samples were of 70-120 Hz. Pulsating stretching scheme with the frequency of 50 Hz was used as well. The values of the cycle asymmetry coefficient in both cases were 0.15 and 0.35.

According to the results of high-temperature test and fatigue crack growth rate measuring on the samples from the above said alloys, kinematic diagrams of fatigue destruction, i.e. dependence of fatigue crack growth rate on stresses intensity coefficient values were plotted.

Based on the conducted fractographic studies and their results comparison with experimentally obtained ones and schematic kinetic diagram of fatigue destruction the schemes are developed; fractographically illustrated stages of fatigue crack growth and various fracture micromechanisms at different sites of the kinetic diagram of fatigue fracture in the material of the samples and blades.

The results of the work can be applied for developing more advanced modifications of turbine blades of high reliability.

Ezrokhi Y. A., Kadzharduzov P. A. Working process mathematical modelling of aircraft gas turbine engine in condition of elements icing of its air-gas channel. Aerospace MAI Journal, 2019, vol. 26, no 4, pp. 123-133.

The article presents general approaches to of aviation gas turbine engine operation modelling in icing conditions.

Component-level engine model is considered, in which the parameters, determining each component operation mode, represent a set of independent variables. These variables values are computed as the result of solving a system of nonlinear equations that determine conditions for the engine system components concurrent operation and its control laws Airflow continuity with account for its bleed and leaks, compressor and turbine power balance for the shaft of each engine are related to the concurrent work conditions, while fuel feeding conditions to the main combustion chamber and afterburner, as well as conditions, determining position of the nozzle actuator inlet guide vanes are related to the control laws.

It is assumed, that the ice formation in air-gas channel of this or that compressor stage, which leads to its airflow capacity reduction due to reduction of its conditional cylinder area of the inlet cross-section. The losses level the of inlet total pressure increase in the compression duct in consequence of inevitably occurring deterioration of compressor elements flow-around due to icing. Quantitative values of these impacts are determined from the engine gas-flow channel sizes, rate of ice growth, as well as the results of well-known generalizations on the unevenness effect of gas-flow channel on the total pressure losses in it.

Ice accretion rate may be set as data of engine testing results in icing conditions, or as a variable allowing evaluating its effect on the main engine performance parameters (thrust, rotation frequency, fuel consumption etc.). The other way to identify the ice accretion rate is solving of complicated thermodynamic problem of ice accretion on this of that part of engine duct surfaces.

The possibilities of the developed mathematical model were demonstrated based on data of test results of the ALF502R turbofan engine tested in ice crystal conditions in NASA Glenn Research Center. Good calculated and tests results matching herewith was demonstrated, which indicates the principal and proved approaches of turbofan operation modeling under the influence of this external factor.

Varsegov V. L., Abdullah B. N. Gas dynamic optimization of wedge-shape vaned diffuser of a centrifugal compressor of small-sized turbojet engines based on numerical modelling. Aerospace MAI Journal, 2019, vol. 26, no 4, pp. 134-143.

A competitive small-sized turbojet engine development under modern conditions of aviation engines building requires high efficiency values of parts with high degree of pressure ratio. Centrifugal compressors find extensive application while developing small-sized gas turbine engines employed for unmanned aerial vehicles and gas turbine power plants.

To ensure high efficiency and compressor pressure ratio, a numerical gas-dynamic calculation is performed with Ansys Workbench (Fluid flow CFX) program, which allows studying the air flow in the diffuser channels.

The presented article considers the flow in a wedge­shaped diffuser and optimize geometry optimization of the wedge-shaped diffusers blades of a centrifugal compressor, as well as geometry impact on the total pressure loss coefficient ξ, and the coefficient of static pressure recovery in the diffuser Cp at different entry angles α3l .

The main task of the calculation consists in determining the optimal shape of the wedge-shaped diffuser blades, insuring required parameters and characteristics of the diffuser, with an uninterrupted flow and a minimum of energy loss at given input angles.

The article presents also the results of the compressor stage numerical study, i.e. joint operation of the impeller with a diffuser to assess the quality of the geometry and operation of the diffuser to increase the compressor efficiency.

In the presented work, the calculation model is built with the SolidWorks program, and then, using the Turbo Grid program, the computational grid was applied. The flow simulation was performed using the SST turbulent viscosity model.

Nadiradze A. B., Frolova Y. L., Zuyev Y. V. Conical plume model calibration of the stationary plasma thruster by the thruster integral parameters. Aerospace MAI Journal, 2019, vol. 26, no 4, pp. 144-155.

The article presents the analysis of possible reasons for divergence of parameters measured under laboratory conditions and realized in space, based on application of multi-fractional conical model of the stationary plasma thruster jet. Three possible methods for the jet model calibration by the thruster integral parameters, such as discharge current, flow-rate and engine thrust were considered. The study of measuring conditions impact on the jet integral parameters was conducted. The need for calibration is stipulated by the fact that jet measured parameters may incorporate essential errors associated with the effect of experiment conditions and vacuum chamber walls. Calibration coefficients, linking measured and integral parameters of the jet, such as total ion current, flow-rate by ions and the jet axial impulse, are being introduced to minimize errors. Inasmuch as the jet integral parameters are being measured with high precision, the thruster jet model accuracy may be significantly increased after calibration.

The calibration methods regarded in the article allow obtain concurrence either by current density or by the flow-rate, or by the thrust (axial pulse). Jet calibration by the ion current and ions flow-rate gives the undervalued thrust value. Calibration by the thrust gives the jet parameters estimation for the worst case (overvalued parameters by the ion current and flow rate) necessary for analyzing the jet impact on a spacecraft. However, it is impossible to obtain the exact concurrence for parameters due to the effect of jet «disintegration» caused by interaction between the accelerated ions and neutral particles. Besides, the particles of residual atmosphere in vacuum chamber may affect the processes of jet formation in the acceleration channel of the thruster. To obtain more accurate jet model, it is necessary to account for the above-mentioned factors, and to use more complicate correction methods.

Marchukov E. Y., Vovk M. Y., Kulalaev V. V. Technical appearance analysis of energy systems by mathematical statistics techniques. Aerospace MAI Journal, 2019, vol. 26, no 4, pp. 156-165.

Aerospace industry development is impossible without implementation of up-to-date samples of high-efficiency new generation energy systems (ES). The term “technical appearance” implies the aggregate of parametric, structural and technological solutions, reflecting most substantial specifics of the system appearance [5]. It is well-known that designing and production of new technology, inclusive of ES in aerospace industry, leads to the necessity of taking compromise optimal or rational engineering and technological decisions. Besides, designer always faces the requirement for conformity of technical appearance forecast of the ES being developed to its real-life prototype. An engineering approach based on statistical analog technique for decision-making while developing new technology may be of help for the appointed tasks solution and meeting the above said requirements [10, 15]. This technique foundation consists in the fact, that deep analysis and synthesis of static structural and energy data of the ES, selected analogs and prototypes according to the parameters of technical requirements to the design according to [15-17] are performed while prospective equipment development. The article regards the energy system (ES) in general form as a mechanical machine for input energy conversion into useful work. Methodological basics of the new generation ES optimal appearance forecasting by mathematical statistics techniques [15-24]. The article demonstrates that development and introduction of the special statistical criterion, integrating all operational parameters in the form of multi-parametrical function, is urgent for solving scientific and engineering problems of new ESs development with specified properties of enhanced effectiveness. This criterion may be named forecast criterion. The introduced special forecast criterion is based on ES statistical analog data fields processing (already created and successfully operated) by mathematical statistics techniques [15-17]. The criterion of the analytical form analysis by independent parameters-arguments leads to formulation and solution of the extreme problem of a multi-parameter function optimizing by known mathematical methods [18, 20, 24], where obtained optimal parameters determine the forecast of the newly created ES optimal technical appearance. Algorithm for compiling and special forecast criterion computing in general is presented. To demonstrate the legitimacy of the criterion introduction, an example of computing the forecast of the ES technical appearance in general is given. The scientific results of the article may be used to develop a comprehensive software product for modeling technical optimal concept of a new generation ES with increased output energy operational parameters and optimal mass-dimensional (volumetric) characteristics.

Kartas S. S., Panchenko V. I., Aleksandrov Y. B. Geometric parameters effect of ejector with curvilinear section of mixing chamber on its characteristic. Aerospace MAI Journal, 2019, vol. 26, no 4, pp. 166-173.

Ejector is the simplest device without moving parts for liquids, gas, and other media moving. Power transfer from one stream to the other proceeds by their turbulent mixing. Very often, injector is employed to maintain continuous airflow in a duct, or a premise, thus performing a fan role. It is used also for jet engines testing. The exhaust stream flowing from the jet nozzle draws in the air from the shaft into the ejector, ensuring thereby the premise ventilation and engine cooling.

Over the past 60 years, plenty of studies has been performed on ejectors as a part of jet engines, which purpose consisted in increasing engine thrust, and reducing fuel consumption, jet noise and output temperature.

In modern conditions, these devices are used in various fields, such as aircraft and machine building, firefighting equipment, and as pumps, compressors, and mixers at oil tank farms.

In general, the described ejector structures include straight-line mixing chambers. Employing a curvilinear section of mixing chamber, which allows improve the ejector parameters, may be suggested as an option of such ejectors. An option of the ejector of this kind consists of a high-pressure flow nozzle, a low-pressure flow nozzle, mixing chamber, and diffusor. With this, the initial section of the mixing chamber is curvilinear.

The disadvantage of this ejector is certain difficulties in manufacturing curvilinear surfaces of nozzles and initial section of the mixing chamber. The advantage of this ejector consists in average velocity reduction of the active jet at the mixing chamber inlet, and, as a consequence, mixing losses reduction.

The article presents the results of numerical calculation of the  characteristics of curvilinear ejectors with F1/F2 = 1 geometric parameter (elbows and bends) at relative sizes of R/a = 1; 2. These results revealed that with the same ejection coefficients, the relative pressure drop is greater for a curvilinear ejector with a relative radius of R/a = 2. The numerical calculation was performed in a stationary setting using the Fluent program and the k-e RNG turbulent viscosity model. Based on preliminary calculations and the grid independence analysis of the obtained results, the grid models were selected.

Volkov S. S. Assessment techniques for psychophysiological state of special purpose systems operators. Aerospace MAI Journal, 2019, vol. 26, no 4, pp. 174-183.

The article deals with assessing techniques for psychophysiological state (PPhS) of a flight crew, cosmonauts, test pilots and other representatives of the aerospace industry. An approach, involving gas discharge visualization method in conjunction with fuzzy logic system for psychophysiological state monitoring is being offered for consideration. Prospectives of automation system for psychophysiological state assessment techniques implementation in the interests of aerospace comples are demonstrated.

The purpose of the work consists in demonstrating the increase of the PPhS assessment quality of special purpose systems operators of the aerospace industry. Special purpose systems operators are both civil and military aviation flight crew, cosmonauts, test pilots, and specialists dealing with robotic systems.

This work novelty lies in the intelligent tools application for operators’ PPhS determining. The interest to this method application is caused by the fact that human ability to perform professional duties is characterized by his psychophysiological state. Psychophysiological state monitoring of operators of special purpose systems (SPS) of aerospace industry allows increasing efficiency of their decisions and raise their readiness to perform special duties. Eventually, the ability to perform special duties unconditionally may and must be controlled and monitored to enhance readiness to perform the assigned task during the periods of flying vehicles flights and testing.

In this respect, the necessity for performing control of SPS operators of aerospace industry at the stage of their preparation for flights and tests performing, as well as during special assignments performing with automation tools application is imminent. It would allow assess with certain fidelity their readiness to perform the assigned tasks during flights and tests, and point out to particular official the necessity to pay attention to this or that pilot, cosmonaut or technician. However, such control implementation is not possible without methodological tools and means for assessing flight crews, cosmonauts and other aerospace industry prepresentatives fitness for their functional assignment.

As the result of the studies, an algorithm of the decision-making support system with fuzzy logic system for automated assesment system of PPhS operators was developed.The fuzzy logic system operation is based on the Mamdani algorithm.

The PPhS assessment techniques implementation, described in the article, in the aerospace industry will allow monitoring the health of the flight crew, cosmonauts, test pilots and operators of robotic systems, as well as reducing the risk of injury and mortality factor while equipment operation.

Boyarskii G. G., Sorokin A. E., Khaustov A. I. Experimental pressure-flow characteristics determining of micropumps for orbital station biotechnical system. Aerospace MAI Journal, 2019, vol. 26, no 4, pp. 184-190.

While conducting research at the space stations, great attention is paid to revealing the weightlessness effect on the cells, which allows the results transferring to the other objects and models in various areas of biology and medicine. For such studies performing, the authors suggest to apply a biotechnical system for cell culture (BTS CC) in conditions of spaceflight, which main element is a micropump, meeting the following requirements:

– to possess minimum size: diameter of not more than 10 mm, and length of not more than 50 mm,

– to ensure a liquid supply with viscosity of 1 cSt from zero to 0.1 liters per minute,

– to ensure pressure of up to 3 J/kg.

The existing techniques for axial pumps design do not allow correctly determine the micropump geometric size and its pressure-flow characteristic, since with a pump size reduction compared to the full- size pump, relative size of gaps and roughness increase, which changes significantly redistribution of the velocities fields and volume leakages, as well as disk and friction losses. A micropump designing with such specifics requires new structural and designing concepts.

Based on the full-size pump designing experience and with account for the BTS CC pump operation specifics, a new micropump of 6.5 mm diameter and 45 mm length was developed. Its control block allow changing rotation speed and the electric motor and impeller of the micropump by setting the current frequency and value, varying hereby the pump delivery and pressure.

Any pump characteristic is its head dependence H on delivery Q at various rotation frequencies of the pump shaft, i.e. H = f (Q, n). Thus, to determine the micropump pressure-flow characteristics, experimental studies are necessary to examine the effect of geometric size and mode parameters on its characteristics.

The main difficulties in the pressure-flow characteristics determining of micro-pumps, i.e. the dependence of the pump head on its supply and shaft speed, is their small size, commensurable with the sensors size.

Analysis of publications related to the study of fluid micro-flows in micro-pumps revealed that they use tracers were employed for this purpose, which introduction disrupts the micro-pump operation. Thus, to determine micro -pumps characteristics, a test

bench was designed and manufactured. It includes non­inertial micro-sensors (for the pressure drop-head registration and measurement). The flow rate was measured by weight, with account for the liquid evaporability. The micropump pressure-flow

characteristics are modeled by changing hydraulic resistance at the pump outlet by varying the flow section of the throttle. The measurements were repeated for different speeds of the impeller shaft from 2000 to 20,000 rpm.

The results of the tests revealed that the micropump pressure-flow characteristic represent a falling dependence typical for the full-sized axial pumps. However, stratification of dynamic characteristics is being observed at various impeller rotation frequencies. Thus, for the range of n1 > 8000 rpm the pressure-flow characteristic goes higher, than for n2 < 8000 rpm. The obtained pressure-flow characteristic of the developed micro-pump allows estimating the effect of the micropump micro-sizes on its efficiency.

Kirsanov A. P. Stealthy movement of aerial object along rectilinear paths in the onboard doppler radar station detection zone. Aerospace MAI Journal, 2019, vol. 26, no 4, pp. 191-199.

Onboard radar stations operating in the pulse–Doppler mode show the characteristic feature in the detection zone. This feature consists in the fact that in every point of the detection zone the aircraft has a sector of directions moving along wich it not detected by the onboard Doppler radar. This sector is called the sector of invisible motion directions of the aircraft. Due to these features, there are stealthy paths allowing an aircraft stays non-detected by Doppler radar station, such as radar station of an airborne early warning aircraft, while moving along them. The majority of stealthy trajectories is curvilinear with variable curvature. The article deals with the study of the rectilinear paths of the aircraft stealthy movement in the onboard Doppler radar station detection zone. It was established that any aerial object position relative to the early warning aircraft might be the start of the rectilinear stealthy path at the appropriate selection of direction of movement. An equation to determine the stealthy movement duration along the rectilinear path depending on the aircraft initial position and its direction of movement was obtained. Areas in the detection zone of the pulse-Doppler radar station to which the aerial object may enter, moving along the rectilinear stealth paths, were plotted. Their shapes and sizes depending on the aerial object position and motion parameters relative to the radar station were studied. Conditions of the unlimited time duration of movement along the stealthy paths, and conditions of the rectilinear stealthy paths for the aerial object outgoing to the onboard Doppler radar station location were found.

Bezzametnov O. N., Mitryaikin V. I., Khaliulin V. I. Low-speed impact testing of various composites. Aerospace MAI Journal, 2019, vol. 26, no 4, pp. 216-229.

The purpose of the study consists in technique development for detecting impact damages character of composites with various nature of reinforcing material and interlacement type. A series of experiments on the presence of internal defects after impact damages inflicting was conducted while this work performing. The samples based fourteen fabric types were selected as the subject of the study, including fiberglass cloth, hybrid materials, Kevlar® and high molecular polythene. Temperature mode was developed, and technology for plates manufacturing by the compression molding technique was worked out.

The experiment technique was being developed with regard for the international Standards recommendations for damage resistance testing while the falling load impact. Evaluation of criteria on impact resistance was performed within the energy range of 10, 20 and 30 J. Initially the dent depth was determined with digital detecting head. The internal damages areas were being estimated by the semi-automated ultrasonic NDT complex with phased array. This technology allows obtaining scanning results in the form of projections onto three planes, namely C-scan (top view), S-scan (end view) and B-scan (side view). To analyze the damages areas of samples after the impact, the C-scan, depicting the scanned area below the sensor, was registered. The layer-by-layer studying of the samples damages character was performed by the X-ray computer tomography method. This method allows visualize the sample internal structure by processing shadow projections obtained while the object X-raying.

The obtained results allow determine optimum characteristics of the composite material pack content while developing the structure with the set requirements to the impact resistance. The nose part elements and high lift devices of an aircraft, helicopter blades and transmission shafts, moving parts of jet engines may be among these structures.

Based on these works results graphs of the damages areas dependence on the impact energy of each material were plotted. The less damage area was demonstrated by the fiberglass samples, while the greatest one belonged to the fabrics of hybrid content. To evaluate the impact resistance criteria the energy of the damage initiation, maximum load of impact and absorbed energy were registered. Maximum value of the damage initiation energy was demonstrated by the samples from hybrid fabric material, and the least one by the fiberglass samples. This criterion reflects the limiting value of the impact energy which a material can sustain without being damaged.

Savel’eva L. V., Vendin I. O. Cutting conditions effect on tool front surface wear rate while workpieces machining. Aerospace MAI Journal, 2019, vol. 26, no 4, pp. 209-215.

The article tackles the issue of determining the degree of various cutting modes effect (cutting speed, cutting thickness, cutting width, feed, cutting depth, temperature, front angle, vibration) on the front surface wear of the cutting tool.

The authors describe the nature of cutting modes effect on the front surface wear of the tool, and suggest recommendations on optimal cutting modes, which ensure maximum life span of the tool.

The article consists of three main sections: introduction, the bulk section, conclusions.

The introduction considers causes of the tool wear. As a rule, cutting tools wear occurs under the impact of molecular adhesion forces of the treated metal surface with the cutting tool, or under abrasive action of solid particles existed in the structure of the machined material.

The main section regards the tool wear process over the front surface. It analyzes an experimental dependence of the cutting speed impact on the tool wear intensity. As the result of the analysis conclusion was made that the wear increased with the cutting speed increase. According to professor A.M. Danielyan’s studies, with the cutting speed, feed and cutting depth 20% increase the cutter surface wears out correspondingly 3.5, 1.7 and 1.05 times faster. This research data demonstrates that the largest effect may be achieved not by the cutting speed increase, but by the cut width and thickness increase. The effect of the cut thickness and feed on the wear intensity of the cutting tool is analyzed. With large cut thickness (more than 0.5 mm), a misgrowth of significant height is formed, eliminating the contact of the rear surface with the cutting surface. Only the front surface of the tool thereby wears out. With the cut thickness reduction, the wear occurs on both back and front surfaces simultaneously. At very small cut thickness (less than 0., the misgrouth is of rather insignificant height, and the wear occurs only on the back surface. With feed increase, the cut thickness increases either, and, thus, the wear on the front surface increases. The experimental dependence of the cut depth impact on the tool wear intensity is analyzed. As the result, the optimal cutting depth is determined, at which the front surface wear is minimal. The experimental dependence of the tool temperature influence on the tool wear intensity is analyzed. The optimal tool temperature, at which the wear of the front surface is minimal, is determined. The effect of the tool front angle and vibrations on tool wear is analyzed.

Recommendations on selection of optimal cutting modes, ensuring maximum tool life are presented in conclusions.

Ushakov I. V., Simonov Y. V. Experimental detection of micro-destructions viscosity in central and boundary areas of brittle samples while loading on the substrate by vickers pyramid. Aerospace MAI Journal, 2019, vol. 26, no 4, pp. 230-239.

The main purpose of the work consists in developing the earlier proposed technique for viscosity detection of micro-fracture of thin brittle amorphous nano-crystalline samples.

The regularities of deformation and fracture under local loading of solid thin samples of nano-crystalline material by Vickers pyramid are determined experimentally. The main studies were performed on amorphous metallic alloy Co71,66Si17,09B4,73Fe3,38Cr3,14, converted into the nano-crystalline state by controlled isothermal annealing.

The dependency of the symmetry of micro-patterns of destruction from the load value and a distance to the sample boundary was established. It is established that with the load growth occurrence of symmetry elements starts to be observed in the initially asymmetric fracture patterns. Statistical analysis of symmetric cracks, as well as the distances between them, allows find the micro-destruction viscosity of the material. At a certain optimal load, the probability of symmetrical micro-patterns formation is maximal. A further load increase leads to the symmetry reduction, and, accordingly, to the decrease of micro-destruction viscosity calculation accuracy.

For the first time, a technique for determining the minimum allowable distance to the boundary of a thin sample, on which the micro-destruction viscosity determining was possible, was proposed. It was established that the optimal load value while determining the micro-fracture viscosity near the sample boundary coincides with the value of such for the central areas.

For the first time, mechanical testing modes, which allow obtain symmetrical and analyzable micro­patterns of destruction were determined. These conditions include the following: using the optimal load on the indenter; accounting for the allowable distance between the adjacent loading areas and a distance from the loaded area to the sample boundaries. Based on the experimental results analysis, algorithms for to determining the optimal load on the indenter and the allowable distance to the sample boundary have been developed. The obtained results allowed improve the earlier proposed technique for micro-fracture viscosity detection by local loading of thin, hard and brittle samples.

Sedel'nikov A. V., Belousova D. A., Orlov D. I., Filippov A. S. Assessment of temperature shock impact on orbital motion dynamics of a spacecraft for technological purposes. Aerospace MAI Journal, 2019, vol. 26, no 4, pp. 200-208.

The main objective of the work is assessing the of temperature shock impact on the orbital motion dynamics of the spacecraft for technological purposes.

The problem consists in the uncertainty of center of mass displacement due to the impact of temperature shock and, thus, the motion control error. This problem is particularly relevant for the spacecraft for technological purposes, and products sensitive to the experimental conditions.

The importance of assessing the impact of temperature shock is determined by the need to ensure the spacecraft functioning with the specified parameters of motion, as well as maintaining controllability of the spacecraft in the presence of orbital eclipse periods.

Analysis of the studies by the scientists from various countries reveals that control of a small spacecraft with no large elastic elements in the design-layout scheme often reduces to the target values active control of the angular velocity of its rotation.

In this case, the orbital eclipse periods are not highlighted separately, and no changes in spacecraft movement control law are made while its immersing in and out of Earth shadow.

The article deals with the issues related to the temperature shock impact on the orbital motion change of a spacecraft for technological purposes, and modeling the scale and nature of the effect.

The temperature shock impact assessment is based on the 3D modeling of the processes occurring at the spacecraft entering and exiting the orbital eclipse period.

For a small “Return— MKA” type spacecraft the three-fold excess of admissible micro-accelerations was obtained.

As the result of the conducted study, a conclusion was made that control algorithms development, levelling the temperature shock from the viewpoint of occurring micro-accelerations compensation, was required for successful implementation of gravity- sensitive processes onboard the spacecraft for technological purposes with the orbital eclipse period.

A three-dimensional heat conduction problem was solved to determine the target parameters of control algorithms. The following simplifying assumptions were introduced to solve the problem:

– the elastic element model was a frame structure;

– the elastic element was rigidly fixed in the small spacecraft body;

– the elastic element properties satisfied the conditions of homogeneity;

– the heat flow was uniform;

– operating temperature range was −170... + 110 °C;

– the properties of the elastic element material were considered constant throughout the operating temperature range;

– orientation changing of normal to the elastic element surface due to its own oscillations was neglected.

Ivanov P. I., Kurinnyi S. M., Krivorotov M. M. Asymmetry in the parachute canopy filling process. Aerospace MAI Journal, 2019, vol. 26, no 3, pp. 7-16.

The main purpose of the work consists in studying dynamics and specificity of filling the large area parachutes of the main class employed for rescuing re­entry spacecraft as well as large weight cargoes airdrop of civil and military hardware. The problematic issues here are these associated with the occurrence of large aerodynamic load values while parachute dynamic filling, which may lead to premature loss of its strength. The issues of long delay in the filling process, which increases the path and height loss and is very dangerous while low-height airdrop, are of no less importance. The article tackles the issues associated with the filling process deviation from the rated value, such as asymmetry occurring while the parachute canopy filling.

The dependence between the filling time and aerodynamic load on the parachute, i.e. maximum drag force value, was established experimentally. The article demonstrates that with the parachute filling time increasing the aerodynamic loads on the parachute and overloads on the object decreased, while the filling path increased.

The relationship between the edge contour of the canopy inlet orifice shaping, filling time and aerodynamic loads on the parachute was established. One of the possible causes of both deceleration and intensive canopy filling dynamics, consisting in substantial asymmetry of the shaping process of the edge contour of the parachute canopy inlet orifice, was revealed.

The authors introduced the notion of the canopy contour shaping asymmetry coefficient at the intensive dynamics of the canopy filling process, as an effective tool for studying the processes of canopy edge shaping processes and their quantitative evaluation.

Setting the rated boundary value for the asymmetry coefficient it is possible to make judgments on the tendency of the canopy shaping by the degree of distance from this boundary. Thus, it will show the propensity of the specified parachute for the asymmetric filling and the ensuing negative consequences, associated with intensive dynamics of the filling process and load-carrying capacity loss. Practically, the asymmetry coefficient represents the square root of the ratio of impact pulses from the air- velocity pressure (which form local pressure drops along the carrying surface) for the canopy with asymmetry, and a canopy being filled symmetrically, under the same initial conditions on speed.

The larger the coefficient of asymmetry, the larger the dome is predisposed to asymmetric filling, the more shock impulses differ. In this association, the probability of the canopy and shrouds destruction increases in the local loading area from the pressure drop at the loads being measured by the strain sensor in the parachute thimble, which are substantially lower than its load-bearing capacity.

Novogorodtsev E. V. Numerical study of total pressure in the air intake with sharp edges applying eddy-resolving sbes-method. Aerospace MAI Journal, 2019, vol. 26, no 3, pp. 17-31.

The values of the total pressure oscillations intensity root mean square parameter ε in the channel of isolated air intake with sharp edges were determined as applied to industrial aerodynamics problems based on numerical solution of Navier-Stockes system of equations. Numerical solutions of Navier-Stockes system of equations were obtained using eddy­resolving Stress Bkended Eddy Simulation (SBES) approach employing ANSYS CFX solver. Simulation of the 3D flowing of the viscous compressible gas around and inside the object was performed employing spatial regular multi-shell grid. The procedure of computational grid generation was being performed in manual mode employing ICEM CFD software.

To evaluate fidelity of the computational study based on SBES method application, comparison of the obtained values of the root-mean square parameter of pulsations intensity with experimental data was performed. The data processing procedure herewith was conducted in concordance with the standard experimental technique approved in TsAGI.

Numerical simulation results are presented in the form of plots of parameter e values in the engine section as a function of the specific reduced air flow q(λen) through the engine cross section. The air intake duct throttling was modelled by cross-clamping of the auxiliary duct in the form Laval nozzle. The auxiliary duct wall profile in the longitudinal section herewith was constructed using the Vitoshinsky formula.

The article performed a comparison of total pressure oscillations obtained while computational study in monitored points of the metering cross­section with oscillograms obtained while experimental study according to readings of the total pressure pulsations sensors, installed on the model at the same points of the reference cross-section.

The parameter ε values obtained in the framework of this work in the engine cross-section for the air intake and engine synchronization mode in all regarded range of of the incoming flow Mach numbers M = 0-1.8 (at zero angles of attack and sideslip) are in good agreement with the experimental data. Maximum discrepancy between computational and experimental results was Δε max = 1% in absolute units of the ε parameter.

The ε parameter values were obtained for both the air intake configuration without a boundary layer control system, and the one with a boundary layer control system.

Bragazin V. F., Gusarova N. A., Dement’ev A. A., Skvortsov E. B., Chernavskikh Y. N. On practicality of deflectable thrust vector application for civil aircraft. Aerospace MAI Journal, 2019, vol. 26, no 3, pp. 32-42.

The study focuses on the engine deflectable thrust vector (DTV) application on the civil aircraft to improve its controllability, as well as take-off and cruising-flight characteristics.

Thrust vector deflection is achieved through the movable nozzles. Three options of the engines location in the aircraft layout, namely, on the pylons under the wing, as well as on the pylons of the fuselage nose and tail parts were considered. Esteems of the DVT application as an additional element to the aerodynamic control elements were obtained.

The DVT application as an additional balancing element of pitch and/or yow control leads to the possible reduction of the horizontal tail (HT) and/or vertical tail (VT). Thus, for the aircraft layout with the engines under the wing, the HT area reduction may be of 11%, and VT area reduction of 8%. For the aircraft layout with the engines in the fuselage tail part, the VT area reduction may be of 13–20%. The DVT application along with the aircraft aerodynamic control elements allows increase the effectiveness of the lateral, pitch and yow control, as well as reduce the aircraft response time to the steady-state overload.

The aircraft cruising aerodynamic quality changing depending on the engines position on the aircraft and thrust vector deflection was considered. The largest increase in maximum quality was realized with the engines location in the front part of the fuselage and upward thrust vector deflection. It was revealed, that aerodynamic quality increases about 2% within the angles range of 0° to ±10°. According to the preliminary estimates, the aggregate impact of several factors may ensure the fuel consumption reduction in the cruising flight by approximately 3–4%.

While studying the takeoff trajectory, it was found that the largest trajectory slope angle at the safe takeoff speed was possible with the DVT engines application in the taili part of the fuselage.

According to the preliminary data, the DVT application bears a potential to improve a civil aircraft operational characteristics. The DVT significant useful effects are the possibility of aircraft control dynamics improvement and flight safety enhancement at the takeoff/landing and climbing modes.

Levin V. I., Karasev D. Y., Sitnikov M. S. Aircraft break wheels designing using 1D thermodynamic models. Aerospace MAI Journal, 2019, vol. 26, no 3, pp. 43-61.

The OEM, EASA and ICAO requirements to aircraft systems and equipment force manufacturers to conduct more verification calculations and tests to confirm the announced characteristics, as well as analysis of various modes of operation. Currently, there are already new methods of design, as well as automation of calculations and tests. Thus, it is necessary to develop both theoretical and practical basis for their implementation.

The objectives of this work consist in determining a convenient method for thermal processes computing in the the aircraft wheel structure, as well as describing a method for developing a 1D model for the wheel thermodynamic calculations, performing computations by this model, and comparing the obtained results with the results of test modes.

The article provides a summary of the research and work conducted at the enterprise of the brake wheels manufacturing company. The approach to computing the thermal energy distribution dynamics over the friction disk volume and the wheel structure while braking process is being substantiated. The adequate accuracy while using the reduced model of the disks temperature computing is demonstrated. The article presents the processes and methodology issues of developing architecture and parameterization of the wheel structure model for computing the points of the monitored temperature. The model additionally accounts for the convective thermal exchange with the pneumatic network of the air cooling from the brake wheel. Speed, direction and successive air heating are also being accounted for. The results of computing and testing at three test modes are presented. The adequate accuracy of the computational results compared to the testing data is being determined.

Eventually, all declared goals were achieved. A convenient method for thermodynamics computing of the wheel based on the 1D model was determined. Virtual testing was performed on both a model and a test bench. Analysis of the results allows stating the expediency of the 1D models while brake wheels designing.

Virtual tests were performed on the developed and validated model, which allowed determine more optimal modes of the test bench equipment application. This, in its turn, allowed the time reduction of the field tests and the number of test launches.

Currently, a set of documentation has been developed to justify changes in the regulations for the design and conduct of accelerated life tests of the wheels. The prospects for the used computing method development for solving the related tasks of the break wheels design.

Boldyrev A. V., Pavel’chuk M. V., Sinel’nikova R. N. Enhancement of the fuselage structure topological optimization technique in the large cutout zone. Aerospace MAI Journal, 2019, vol. 26, no 3, pp. 62-71.

Topological optimization techniques play an important role while selecting a structural layout of aggregates for a flying vehicle of minimal mass. The goal of the presented work consists in increasing weigh efficiency of the aircraft structure in the stresses concentration zones. The article proposes a of topological optimization method for edging of the cutout for the hatch in the fuselage, based on the full- stress concept with regard for the functional limitations on the generalized hull skin displacements at the cutout contour.

For the design object synthesis, a method, based on Komarov’s mathematical model of a deformable solid body with variable density is being applied. An artificial material with variable density and rigidity, called a “filler", in which the strength and elastic properties linearly depend on density, is being employed.

Finite element models, integrating the manifold of the load-bearing elements of the structure and continuous medium of variable density are being developed while topological design. Earlier, such combined model was employed in [25, 26]. The material distribution in the filler allows revealing theoretically optimal structure and, using the strategy [8], developing the structural layout closest to the theoretical solution from the viewpoint of its stressed operation. The topological optimization process is based on stage-by-stage substitution of the filler by structural elements, realizing the technical decisions being taken

The article presents a numerical example of the fuselage compartment design with rectangular cutout, demonstrating the operability of the suggested technique. Conventional layout with well-known prototypes technical solutions is adopted as an initial structure. The topological optimization resulted in obtaining new technical solution allowing 16,7% reduction in the mass of the strengthening members of the cutout relative to the initial structure. The parts of the internal panel are shifted inward the fuselage from its theoretical contour and duplicate the hull skin at the cutout portion. The internal panel is fixed to the hull skin by the longitudinal and sloped walls, reinforced and ordinary bulkheads. The manifold of stressed elements forms closed and hollow contours in the cutout corners, enhancing the structure rigidity in the hatch cutout zone in radial and longitudinal directions.

Mamedov I. E., Sharifova B. A. UAV functioning mode optimization while seawater sampling. Aerospace MAI Journal, 2019, vol. 26, no 3, pp. 72-79.

Water is a necessary factor for the humankind survival. For this reason, the quality of water resources should be protected. Thus, it is necessary to organize permanent monitoring of water resources. Industrial and agricultural wastes are the main sources representing danger for water basins. Water quality of rivers and lakes may be evaluated by monitoring such indices as quantity of dissolved oxygen, pH., temperature, and electric conductance. Low concentration of oxygen dissolved in the water, undesirable temperature and abnormal salt content lead to water quality degradation. The article is dedicated to the issues of UAV application for the seawater salinity and conductance determining. The UAV application for this purpose allows increasing space-time resolution of the results of the studies being performed. The task of forming the UAV empirical model in water sampling mode was formulated. Electric conductance sensors while corresponding UAV flight altitude control are being immersed into the water and taken out after conduction measuring. Thermal sensors are applied herewith, installed on the other UAV flying 30-40 meters higher than the first one. Temperature survey is performed to reveal undercurrents of the incoming external water, which temperature and salinity differ greatly from those of the basic water body. The studies employing heuristic procedure of collating the values of the searched indicator, computed by different representations in the form of one graphics data, and checking the obtained results by the data represented by the other graphics data were performed. The article suggests an empirical model of the UAV, employed for the water quality studying. The empirical model of the UAV in the mode of sampling for the samples analysis is presented as well. Specific issues of realizing the suggested empirical algorithm for the empirical model development were considered. Indirect validation of the developed empirical model demonstrated close agreement of experimental and modelled dependencies character obtained based on heuristic algorithm of the UAV functioning in the water quality studying mode.

Abdulin R. R., Podshibnev V. A., Samsonovich S. L. Determining planetary gearing optimal gear ratio allowing minimize its outer diameter at the specified load torque. Aerospace MAI Journal, 2019, vol. 26, no 3, pp. 80-90.

Mass and size parameters reduction is one of actual issues of aircraft electromechanical drives design. It concerns especially mechanical transmissions employed in drive systems of mechanical transmission. Harmonic and planetary gears are most compact. They allow obtaining large gear ratio for a single stage. Their application as the output stage of a multi-stage reduction gearbox of an electro-mechanical drive, as a unit transmitting the largest moment, allows mass and size parameters reduction of a drive system.

The goal of this article consists in determining the optimal values of gear ratios at which the outer diameter of planetary transmissions has its minimum size for the specified load moment.

It was demonstrated, that the main parameter affecting the outer diameter of planetary transmissions for the specified load moment was the carrier radius. For a single-row planetary transmission this radius was expressed through the gear tooth module value, the number of teeth of the central sun-gear and gear ratio between the sun-gear and satellites. The article presents substantiation of the above said parameters selection. Minimum acceptable carrier radius was found. It was established, that optimal gear ratio value of the single­row planetary transmission equaled four.

The carrier radius planetary gear with double-row planets was expressed by gear tooth module and two gear ratios, namely between the central sun-gear of the planet gear and first-row satellites, and between planet gear of the second row and the crown-wheel. The dependence of the carrier radius on these gear ratios, which is represented by a surface with «ravine», was plotted. A unified optimal gear ratio value was not obtained for the planetary transmission with double­row satellites was not found. However, a set of quasi­optimal values do exist. The “ravine” direction, along which the quasi-optimal values were located, was determined. The optimal relationship of gear ratios between the central sun-gear and the first-row satellites, and between the second-row satellites and the crown wheel was derived. This relationship allows ensure minimum outer diameter of the planetary transmission with double-row satellites. An example of the minimum outer diameter of the planetary gear with double-row satellites computing is given.

The obtained optimal gear ratio values expand the knowledge on planetary transmissions and allow minimize overall dimensions of aircraft drive systems while developing multi-stage reduction gearboxes for electromechanical drives with output planetary transmission.

Nikolaev E. I., Nedelko D. V., Shuvalov V. A., Yugai P. V. External airbags application onboard a helicopter. Aerospace MAI Journal, 2019, vol. 26, no 3, pp. 91-101.

The subject of the presented article is an energy absorption system in the form of external airbags, fixed under a helicopter fuselage. The external airbags are meant for reducing the risk of injury of the passengers and helicopter damage in case of a crash landing.

The study of the external airbags impact while crash landing was performed by the finite elements method. The airbags mathematical model, accepted in the computations, assumes gas simulation by the thermodynamic parameters (pressure, temperature) averaged by the airbags volume. The article presents the airbags initial characteristics for the case of the gaseous nitrogen application. Gas leakage from the airbags is determined by the area of the vent hole and the value of relative pressure for initiation of the gas outflow from the vent hole. The initial pressures values and the holes areas were selected by the condition of overloads minimizing and the strength of airbags material ensuring.

The purpose of this work consists in analyzing the helicopter fuselage loading with the external airbag, and identifying the time dependencies of main thermodynamic parameters of the gas work. The study of a helicopter collision encompasses the moment of time of the airbags contact with the ground to the moment of the fuselage gaining a stable position on the ground. The process visualization of the helicopter fuselage spatial position changing so far as the airbag crimping is demonstrated. Velocities and overloads in the helicopter fuselage center of mass are presented according to the results of computations. The obtained dependencies of pressure, temperature and mass flow rate may be employed for technical requirements forming to the external airbags and gas generating elements structures. Computational results considered in the article allows drawing inference on the possibility of the external airbags application for the helicopter energy shock absorbing and increasing the rate of passengers and a crew survivability. The presented values of loads acting on the fuselage from the airbags side may be employed for the detailed designing of the airbags fixing to the fuselage. The conclusion presents the issues which may become a further development of the research topic.

Chernovolov R. A., Garifullin M. F., Kozlov S. I. Validation of designing and manufacturing procedures of aircraft dynamically similar models with polymer composite materials application. Aerospace MAI Journal, 2019, vol. 26, no 3, pp. 102-112.

Drained dynamically scaled models have been designed for studying unsteady aerodynamic characteristics in wind tunnels. At present, such models testing is of the greatest interest both from the viewpoint of their application for studying safety of the prospective aircraft from the flutter and buffeting, and for verification of calculated aerodynamics with account for the structure elasticity.

The article presents an algorithm for design parameters selecting of a dynamically scaled model and its tuning by test results. The proposed procedure for implementing this algorithm is demonstrated on a simple example (a beam of constant cross section, reinforced by layers of a polymer composite material). Issues of technology for design and manufacturing of a typical element of the dynamically scaled aircraft model applying polymer composite materials are considered. Frequency tests conducting technique is presented, as well as the results of computational and experimental studies of the shapes and frequencies of natural oscillations with account for the additional loads placement. Computed shapes and frequencies of natural oscillations obtained by the finite element method using several successively condensed grids are given. The research findings comparison indicates that calculated values of the cross-section bending stiffness obtained using theoretical relationships and characteristics of the material, accounting for epy specifics of dynamically similar model manufacturing technology, are close enough to those obtained by the experiments at static loading and resonant tests conducting. Setting-up such model does not require special efforts. It allows considering, that the accepted calculating and design technique ensures obtaining required characteristics of the dynamically similar model.

Matiukhin L. M. The fuel molar weight impact on filling, and indicator indices of a piston combustion engine. Aerospace MAI Journal, 2019, vol. 26, no 3, pp. 113-123.

The problems arising while improvement of any type of the internal combustion engine (ICE), such as reciprocating, rotary-piston, gas turbine or jet engines, are common for all of them.

The notions of the volumetric efficiency (nv) and residual gases (γr) traditionally used in the theory of piston internal-combustion engines do not allow characterize the air-fuel mixture composition, which defines the all power, economic and ecological indices of the engines. All the above-mentioned coefficients are applied only while the reciprocating ICE design. With this, the main indicator of pistons filling, namely volumetric efficiency, characterizes not so much the cylinders’ filling as its downgrade due to the presence of hydraulic resistances and incoming charge warming up. The essential drawback of all known equations for the volumetric efficiency determination is ignoring the impact of the fuel type, excess-air coefficient and recirculation’s degree on the cylinders filling. The general-technical concepts of (volume) fractions are far more informative. The aggregate of air-fuel mixture fractions determines its composition and thermodynamic characteristics values. The incoming charge (air) fraction allows unambiguous judgment on the degree of filling the whole cylinder volume, i.e. on the existing reserves of filling. Using the air or mixture volumetric fraction as the main filling indicator while piston ICE cycle computing allows accounting for the fuel molar weight and recirculation impact on the engine indices. As the result of the analysis, in order to account for the fuel impact on the filling the so-called “displacement coefficient” was proposed. Power and economic indices of the engine depend on this coefficient value. The value of this coefficient determines the degree of qualitative power regulation efficiency. Together with the recirculation degree, this coefficient determines the value of stoichiometric relationships and, thus, affects the indicator and effective indices of the engine.

As the sum of the fractions equals to the one, there is no necessity with the suggested approach in separate determining the fraction of the residue gases, since this fraction is equal to the difference between the one and the incoming charge fraction. The suggested approach is of prime importance while analyzing operating cycles of the engines operating on gaseous fuels, and on hydrogen in particular. As a result, the structure of the main calculation dependencies is simplified, and their analysis becomes more clearly evident and easy- to-understand. The possibility of the computing results visualization facilitates their analysis and is a great advantage of the suggested approach in terms of didactics.

Employing the ICE computation as a base of the air-fuel mixture fractions in modern applied programs might have led to the labor intensity reduction and execution time cutting due to the number of variables reduction.

Zubrilin I. A., Didenko A. A., Dmitriev D. N., Gurakov N. I., Hernandez M. M. Combustion process effect on the swirled flow structure behind a burner of the gas turbine engine combustion chamber. Aerospace MAI Journal, 2019, vol. 26, no 3, pp. 124-136.

The article presents the results of computational and experimental study of the swirling flow structure of a swirling jet behind the burner unit of an industrial gas turbine installation. The burner unit being studied in this work is intended for burning poor pre-prepared mixtures. The burner consists of an axial vane swirler with hollow blades through which the main part of the fuel enters, and a “central body”, functioning as a stabilizer with a pilot flame. Natural gas is employed as a fuel. The studies were performed by applied methods of computational gas dynamics and experimental methods. Experimental velocity measurements were performed with a laser Doppler particle velocity meter LAD-056S. Combustion products composition measurements were performed by sampling with subsequent chromatographic analysis. Experimental studies were conducted under the following conditions:

- The inlet temperature Тк = 330 К;

-  Differential pressure ΔP* ≈ 3,3%;

-  Reynolds number at the burner outlet Re ≈ 12000;

-  The proportion of fuel consumption in the standby zone is 11.5% of the total fuel consumption;

-  The excess-air factor for the case of mixing fuel without combustion was α = 2.08, and for the case without combustion α = 1.8.

The flow and combustion processes modelling was performed in three-dimensional unsteady formulation using Large Eddy Simulation (LES) method. Combustion processes were being described with the Flamelet Generated Manifold model. The GRI 3.0 mechanism was selected as the kinetic mechanism of chemical reactions. As a result, a comparison of time- averaged velocity fields and turbulence characteristics was being performed for the case of fuel combustion and without combustion. The obtained simulation results are well agreed with the experimental data on the flow velocity, its fluctuation components, as well as chemical composition. Thus, the employed approach may be applied for calculation study of the combustion processes of the gaseous fuel in swirling flows. An exception is carbon monoxide, which needs to be modeled using approaches accounting for non­equilibrium chemical combustion processes, such as a network of ideal reactors. The flow structure behind the burner was studied in detail, and the characteristics of the recirculation mixing zone were obtained. It was shown, that the fuel supply does not significantly affect the flow structure. It was found, that the combustion process changes the shape of the reverse streams, increasing it in diameter. Mass flow while combustion is significantly lower than in the so-called “cold” case. Due to the air-fuel mixture low consumption through the recirculation mixing zone for the given burner unit, the combustion process characteristics are mainly affected by the interaction between the recirculation mixing zone and the main flow. Pressure fluctuations associated with the vortex core precession, detected while cold purges, were not found during combustion.

Grigor'ev V. A., Zagrebel'nyi A. O., Kalabuhov D. S. Updating parametric gas turbine engine model with free turbine for helicopters. Aerospace MAI Journal, 2019, vol. 26, no 3, pp. 137-143.

A priori estimation of an aircraft engine mass takes on an important role while its creation, especially at the initial designing stage, when conceptual basics of the engine are being established. At this stage, when the design working out of the engine is not done yet, its weight estimation together with fuel economy indicators allows making valid selection of the engine working process parameters values. The presented work refines the parametric model of a gas turbine engine with the free turbine (GTE FT), used in the problem of the helicopter engine working process parameters optimization at the conceptual design stage. With this, while performing parametric studies the design mass of the power plant should be estimated according to the GTE parameters, though, up to now these dependencies are not studied quite well. Thus, the estimation of the engine mass dependencies on its parameters is being performed at present based on the generalized statistic data on the already accomplished structures or parametric mass models, since there is no more precise information at this stage. In fairness, it should be noted that they are all related to the aircraft engines. A rather smaller number of works is oriented of the mass estimation of the helicopter GTE FT. This is primarily due to the fact, that these engines belong to the class of the small-size and have thereupon a number of specifics.

At the same time, as new versions of gas turbine engines appear the periodical refinement the parametric model coefficients values is required. he article considers the mass model of the gas turbine engine with free turbine for several options for the reduction gear mass accounting for, namely, both as a part of the engine, and the power plant. The authors suggest representing the coefficients used in the above said GTE FT models in the form of dependencies on the working process parameters. It allowed perform parametric studies and obtain predictive solutions corresponding to the achieved current design level of gas turbine engines.

Mil’kovskii A. G., Atamasov V. D., Kolbasin I. V., Ustinov A. N., Kalinina A. M. New phenomena in the space experiment on creating an artificial solar eclipse while the spaceships “APOLLO”-”SOYUZ” joint flight. Aerospace MAI Journal, 2019, vol. 26, no 3, pp. 144-151.

The presence of gas-and-dust plasma atmosphere is discovered in every spacecraft, which is confirmed by many domestic and foreign researchers. Due to the medium mixing under the impact of parameters gradients, the radionuclides of plasma atmosphere formed with the intensive impact of gamma and neutron radiation of the reactor would migrate to the outboard space area, surrounding protected part of the spacecraft structure and instrument bay with electronic equipment. These elements would be exposed to radiation due to the induced radiation. In this case, the deterioration of the spacecraft radiation protection against the onboard reactor occurs, which would lead to fluences excess of radiation fluxes on the instrument bay and sensitive structural elements relative to the acceptable levels. Formation of the flows of the eigen external atmosphere (EEA) substance irradiated by the reactor from the operating reactor into the area of the instrument bay and back is stipulated by the presence of parameters gradient of the EEA substance between the specified areas. These parameters are the volume plasma potential and, correspondingly, concentration of charges, pressure and temperature of the gas-and- dust plasma medium. This plasma migration got physical substantiations, published in many scientific works on nuclear physics, performed under I.V. Kurchatov guidance, which attaches authenticity and meaningfulness to the outlined concept, as well as determines the necessity to developing measures for the spacecraft extra radiation protection.

In 1975, an international experiment was conducted in the outer space under the “EPAS” program, during which the artificial Eclipse of the Sun and the solar corona was photographed during the Apollo and Soyuz spaceships joint flight. The spacecraft EEA was repeatedly registered while this experiment. We employed the said photos to analyze the properties of the spacecraft outboard atmosphere. It allowed comprehending the similar processes in the atmosphere of the spacecraft with nuclear reactor.

The physical phenomenon of the “identic luminosity” was recorded by the experimental method in conditions of the space flight under the EPAS program. This phenomenon is a confirmation of the induced radiation phenomenon from the EEA area being under the direct impact of the radiation source due to the various processes of the radiant energy transfer between the particles of the atmospheric environment, varying in weight, shape, chemical content etc., to the shadowed area, protected from direct radiation of the nuclear source, into the atmosphere area. The “identic luminosity” of atmospheric matter can only be explained by the fact that the energy losses while the radiation migration between the described areas are minute. This phenomenon is reliably rendered on all published EEA photos employing high-sensitivity photo film. Such film employing was predetermined by the weak luminosities of the phenomena studied in the experiment such as solar corona and the spacecraft Apollo EEA. They are approximately millions of times smaller weaker than the Sun radiation. Thus, they are being detected only during its full eclipse. This was artificially created in the “Apollo”-“Soyuz” spaceships joint flight (EPAS).

It is necessary to add justification for the necessity for measures to clean the spacecraft outboard space from the EEA caused not by the induced radiation phenomenon only, but also by other non-traditional processes that lead to disturbances in the spacecraft onboard systems functioning.

Lepeshinskii I. A., Tsipenko A. V., Reshetnikov V. A., Kucherov N. A., Sya S. . Joint measurement of gas-dynamic parameters of two-phase highly concentrated flows by laser-optical and probe methods. Aerospace MAI Journal, 2019, vol. 26, no 3, pp. 152-160.

The article considers the problems of joint application of the laser-optical technique for measuring parameters of the two-phase highly concentrated gas-drop flow. Each technique does not allow measuring all necessary parameters. The probe method allows adequate measuring of the local values of the phase flow rates and determine concentration, while measuring phase velocities and drops dispersivity requires suggestion of various hypotheses, requiring experimental verification.

Laser methods allow measure the drops velocities and their sizes in the two-phase flow. However, earlier they could not be applied for studying the flows with large concentration of dispersed phase, as well as determining the gas phase parameters in the two-phase flow. The laser engineering evolution resulted in developing lasers with high spatial and temporal definition, allowing their operation in the area of high concentration of the condensed phase. Combining these two techniques for the two-phase flow study allows go ahead in the area of measuring the parameters, which were either impossible to be measured, or determined with significant error. Particularly, to measure the gas phase velocity and improve measurement accuracy.

Laser-optical methods and Probe methods have long been employed to measure two-phase flow parameters. They are the ones of the few, by which local phase flow rate can be measured. However, their application arouses a number of problems. This is isokinetic problem while sampling and the impact elasticity coefficient selection. Certain design improvements and the probe technique application in compilation with PIV-method allows solving these problems and determining all parameters of the two- phase flow at high concentrations.

The probe represents a cylindrical channel employed in two modes: sampling and measuring the stagnation pressure of a two-phase flow. The problem of isokinetic sampling and selecting the elastic coefficients values of the impact of drops, determining the kinetic energy transfer in the two-phase flow during its braking (the stagnation pressure measurement), were analyzed. To ensure isokineticity, a structural solution was proposed for the probe, which ensures significant error reduction. Application of laser with high temporal and spatial resolution for measuring (PIV-system) allowed determine the drops velocity in a highly concentrated two-phase flow, and, based on the joint measurement with a probe, the coefficient of impact elasticity. The proposed techniques allowed measuring for the first time all the necessary parameters of the two-phase flow. Particularly, we managed to measure the gas phase velocities, and to perform a qualitative comparison with the flow rate of the gas phase at the two-phase flow outlet from the nozzles of the engine combustion chamber mixer.

Katashova M. I., Parakhin G. A., Rumyantsev . V. Multiple mode cathode-compensator developing for the stationary plasma thruster. Aerospace MAI Journal, 2019, vol. 26, no 3, pp. 161-166.

There is a need today in creating a highly efficient multi-mode stationary plasma thruster capable of both inter-orbital transfer and spacecraft position keeping in a set point. A multi-mode cathode-compensator capable of operating at a discharge current up to 15 A is needed for this purpose. The cathode operates on the principle of a gas-electric source of electrons based on a hollow cathode, and it is the most thermally and energy intensive element of the thruster. The K-3/15 cathode structure was designed and studied experimentally on the possibility of flame operation in at least two modes within the discharge current ranges fr om 3 to 5 A and from to 15 A at the experimental design bureau “Fakel” base. The main purpose of the К-3/15 tests was verifying the cathode operability at various start-up powers, propellant flow rates and discharge currents to determine optimal start-up modes. In the process of stand-alone testing, it was determined that the optimal start-up mode for the cathode is a start lies within (160±5) sec at the heating power of 130-139 W and at the cathode flow rate from 30 to 0.60 mg/s. A special attention was paid to determining the current-voltage and voltage-flow rate characteristics in the discharge current range from 3 to 15 A at propellant flow rates to the cathode in the range from 0.30 to 0.60 mg/s. A comparative analysis of the main characteristics of the КН-3В cathode and К-3/15 cathode was performed as well. It was revealed, that compared to the KH-3B cathode the cathode K-3/14 current effectiveness value would manifest itself at the high-current modes (above 10 A), wh ere this parameter value was three times lower. It was determined that the K-3/15 cathode ensured the multi-mode operation with respect to the discharge current and had much higher resource parametrics compared to the KH-3B cathode. It is being forecasted, that parameter changing of the thermo­emitter from mono-crystal lanthanum hexaboride will allow three times increase of the flame operation.

Artyushenko V. M., Kucherov B. A. Analysing the system of restrictions on spacecraft control means application, accounted for while their scheduling. Aerospace MAI Journal, 2019, vol. 26, no 3, pp. 178-189.

A number of tasks of various resources scheduling should be solved to ensure spacecrafts mission control. One of such tasks is tracking, telemetry and command (TT&C) ground stations scheduling. That task is performed under strict resource restrictions. These restrictions include both restrictions on the resource being scheduled and temporal restrictions being imposed on the operativeness of the ground stations distribution plan developing. To ensure operative and qualitative TT&C, accounting for all these restrictions is required.

The restrictions on employing ground stations include the ones on applying separate ground stations as well as restrictions on various ones simultaneous employing. Restrictions stipulated by mission control centers capabilities to perform communication sessions with spacecraft are also a part of the restrictions on TT&C ground stations application.

The restrictions on employing a separate ground station include radio-visibility zones, a set of ground stations network for each spacecraft, a set of service operations to be done for ground station (during which it cannot be used to perform communication sessions with spacecraft) and a set of operation modes supported by each ground station. The restrictions on simultaneous application of different ground stations include ones caused by electromagnetic compatibility and restrictions caused by necessity of employing same resources. The restrictions caused by electromagnetic compatibility can be defined through the sets of two communication sessions characteristics, which cannot be performed simultaneously. These definitions can be used to identify conflict situations while TT&C ground stations scheduling. The resources which simultaneous application may be limited can be sharable or non- sharable. Demands for such resources can be associated with ground stations or their models. It will allow, in is turn, identify conflict situations while ground stations scheduling. Another restriction, which should be regarded while identifying conflict situations during ground stations scheduling, is the maximal number of communication sessions, which each mission control center can perform concurrently. The presented restrictions can be considered as the system of resource restrictions to be accounted for while TT&C ground stations scheduling. The proposed mathematical task formulation of accounting for the system of restrictions can be employed in future development of methodical support for ground stations scheduling.

Maron A. I., Maron M. A., Lipatnikov A. Y. Defining the number of employees for project realization of ground-based radio engineering flight support means upgrade. Aerospace MAI Journal, 2019, vol. 26, no 3, pp. 190-200.

The study relevance is stipulated by the fact that at present the number of projects for ground-based flight support radio engineering means (REFSM) is increasing. The REFSM upgrade represents a project. Such project is associated with a large number of works to be performed. Thus, just one division of the St. Petersburg Center for the of Air Traffic Organization performs technical operation of retranslation stations equipment in the area from Priozersk to Nizhni Novgorod. It is required defining the number of employees for the project completion in the specified time. It should be noted herewith that the same employees ensure operative runability restoring of equipment. The error-free running time of modern REFSM means is tens of thousands hours. It is ensured by both redundancy and technical servicing. A the same time, the defects causing the unit transfer from the operation condition to the fault operable state occur more frequently than the defects leading to inoperability. Such defects require operative elimination since they increase the failure occurrence probability. This problem has not been resolved up to now. Classical methods for queuing systems computing are based on computing probabilities of the system being in various states. They are practically inapplicable due to the dimensionality of the problem under consideration. Simulation methods describe special cases only. They do not guarantee the solution of the problem without analytically found initial approximations to the required number of personnel. The presented article solves the problem by the mean dynamic method. It presents the program for performing computations of the required number of employees in MathCAD Prime. The example of the number of employees computation is given. The proposed method gives practically exact results when the number of units to be upgraded is a couple of dozen or more. In case they are less in number, the obtained number of employees should be refined by simulation. The values obtained by the proposed method herewith will be the initial approximations. The materials of the article are of practical value for the managers of the flight support and communication REFSM services while the upgrading projects planning.

Ied K. . Developing a technique for hazardous situations warning system design while piloting errors occurrence. Aerospace MAI Journal, 2019, vol. 26, no 3, pp. 201-209.

Studying the accident rate of sports aircraft indicates a large number of accidents associated with control loss etc., due to piloting errors and piloting at unacceptable speeds, altitudes and overloads. The current situation requires a flight test methodology developing and specifying airworthiness standards for aerobatic aircraft to improve flight safety.

To define the safe altitude of the maneuver commence, it is also necessary to identify the probabilistic characteristics of piloting errors. Obtaining a functional relationship, based on studying altitude changes in the presence of piloting errors with the regard to the probability of these errors, will allow determine the safe altitude of the maneuver commence with a specified degree of probability.

A mathematical model was developed for studying the impact of pilot’s errors on the changes of trajectory parameters when performing maneuvers on an aircraft.

As a rule, control system of a light sports aircraft is characterized by the extreme simplicity, and is not supplemented with the capability of automated control (autopilot system). Thus, a task arises to develop a warning system, which is not based on automated control (automatic withdrawal from the dangerous altitude), but produces a warning signal only. It requires developing a technique for the warning system developing, which level should be associated directly the probability of the emergency occurrence to prevent this situation transfer to catastrophic one.

The article suggests this problem solving by the technique, according to which it is necessary to supplement the aircraft system with a unit, which would receive velocity and altitude parameters and compare them with the preset values of the acceptable velocities. This is important for warning the pilot on a possible situation to withdraw straightway from the maneuver being performed.

Korobeinikova E. S. Evolvement of quality management systems effectiveness assessment mechanism in aerospace industry. Aerospace MAI Journal, 2019, vol. 26, no 3, pp. 210-219.

Two significant disadvantages are inherent to the procedures of aerospace industry suppliers’ quality management systems (QMS) certification for compliance with whether the universal standard ISO 9001:2015 “Quality Management Systems – Requirements” or industry-specific AS/EN 9100:2016 “Quality management systems – Requirements for aviation, space and defense organizations” have two significant disadvantages. These disadvantages do not let the interested parties (primarily, customer companies and the State) to obtain maximum value added fr om external audits.

Firstly, only the inference on the compliance / non-compliance of QMS with the requirements of the declared standard is the result of certification, without quantitative estimation of the QMS maturity level of the monitored enterprise. Secondly, within the audit the QMS effectiveness is assessed in terms of achieving the results determined by each particular enterprise, whereas, there are quite specific indicators in the aviation industry, characterizing the effectiveness of the implemented systems and the competitiveness of the enterprise.

The aim of the article was to develop recommendations for improving the methodology of the QMS effectiveness assessing. Two trends of improvement were proposed, namely, creating a mechanism for quantitative assessment of the QMS effectiveness level, based on the AS9101 Standard for effectiveness assessing of separate processes, as well as detecting competitiveness rates of the enterprises critical to the specified industry (and, accordingly, clarifying the term “competiveness” for an aviation enterprise).

The first is the development of a mechanism for quantitative assessment of the QMS effectiveness level. The mechanism is based on the one used for assessment of the individual processes effectiveness in the standard AS 9101. The second direction is determining the competitiveness indicators that are critical for organizations of the aerospace industry (and, accordingly, clarifying the term “competitiveness” for aviation enterprises).

A quantitative assessment of the system effectiveness can be performed using the QMS assessment matrix (based on the PEM – process evaluation matrix – used in AS 9101). It is proposed to mark one of its axis with the level of the planned results of the activities

It is proposed to mark the level of planned performance results achievement on one of the matrix axes, and the level of implementation of the QMS standard requirements on the other. The final quantitative assessment of the QMS effectiveness is a score fr om one to four, obtained at the intersection of grades on both axes.

The planned performance results herewith, indicated on the second axis of the QMS assessment matrix, are computed as a complex indicator of the enterprise competitiveness.

This indicator will be computed by the formula:

where αi is the weight of the indicator i, determined by experts;

ci is the parametric index of the parameter i, computed by the differential method (the values of relative indicators determined by the industry are assumed as the base). Individual and group indicators, evaluated while computing the complex indicator, can be derived from the definition of the aerospace enterprise competitiveness specified by the author. Thus, the competitiveness is the ability of an enterprise to meet the consumer needs in terms of the competitive production. This means the qualitative production, corresponding to the consumers’ expectations on acquisition costs operation. It implies also the servicing quality, and related products and services in the necessary quantity and within the required terms, as well as demonstrating to the parties concerned (both direct customers and integrators of various levels, primes) the steady development in conditions of changing external medium, characterized by the costs cutting and profit rising. It should demonstrate also, the effective management, flexibility and ability to optimize their activities, including implementation of new management technologies, peculiar to the industry, namely increase labor productivity, maintain labor, scientific potential and cooperation expressed in the number of customers and partners increasing

С = f (C ; P; R; P; V; V; K; Q; N; m),

Cp — product competitiveness;

P — profit;

R — profitability;

PT — labor productivity;

Vp — the volume of production;

Vr — sales volume;

K — human resources;

Qcoop — an indicator of cooperation activity (increase in customers, suppliers and partners while maintaining the existing ones);

N — scientific and technical potential (includes such indicators as growth in new technologies applicaton (including IT technologies), the volume of in-house development, R&D costs);

M — effective management (increase in use of new management technologies - for example, risk management, lean production and others).

Thus, due to the new methodology application, the QMS effectiveness esteems and the set of competitiveness indicators while QMS analysis of the existing aerospace industry enterprises, the audit emphasis are shifting from the system correspondence to the Standards requirements to the system effectiveness in terms of achieving specific indicators, important to the customers of the aviation industry. Besides, the audits results a cquire quantitative character and allow comparing various suppliers.

Liu L. ., Shi J. ., Bao H. . A metal-composite joint and its mechanical performance. Aerospace MAI Journal, 2019, vol. 26, no 3, pp. 220-227.

A jointing technique, which can be employed in metal-composite joints and may enhance the ability to non-admission of joints disbond, is proposed in this article. This type of joints will contain a certain number of thin pins running though the substrates in the overlap region of the metal-composite adhesive bonded joints. There is adhesive on the surface of the pins and thus, the pins are bonded together with the substrates. And thus, the pins running through the joint plates not only arrest the cracks in the adhesive layer of the bonded joints, also transfer some load between the metallic and composite components. Comparative test results show that the proposed joint method can increase the strength, the failure strain of the metal-composite joints comparing with the traditional adhesive joints, moreover, the joint method can decrease the suddenness of the joint significantly and therefore, improve the damage tolerance performance of the bonded joints. Secondly, the effects of the number and arrangement of the pins on the mechanical performance of the joint will be analyzed in accordance to the test results also. And finally, an optimized method which can improve the load capacity and fracture toughness of the joints will be obtained.

Nasonov F. A., Gavrilov G. A., Babaitsev A. V., Nazyrova O. R. Target modification of constructional epoxy-carbon plastics as a materials science approach to the effect of mechanical joints orifices on bearing capacity. Aerospace MAI Journal, 2019, vol. 26, no 3, pp. 228-242.

The materials science approach to polymer matrices physic-mechanical properties management requires the assessment of modifying additives impact on technological and main operational properties of compositions. Works on studying and intercomparing the main technological properties of the initial epoxy composition and the one modified by technological Zinc Stearate (ZC) technological addition were conducted by viscosimetry and thermo-analytic methods. The developed kinetic model of the compositions hardening process revealed the trifling impact of the composition modification on the hardening process. Pilot samples from the plastics filled with carbon long-fibered fillers (impregnating under pressure and autoclave molding) were fabricated, and their non-destructive control and standard samples testing were performed for mechanical properties measuring.

Estimation by the computer tomography method revealed the stability augmentation of material structure along the edge of the orifice contour after machining for carbon plastics modified by ZC within the interval of 0.1-2% of mass. Thermal effects measuring of machining processes with various tools were performed by IR-thermography method combined with recording function at the specified intervals. The dependence of thermal effects from the modifier concentration was established. The article demonstrates that while this parameter measuring as an integral characteristic, temperatures reduction (temperatures maximums) is observed at the modifier content in matrix samples of 0.1–0.3% by weight, and at the content of 0.2–0.5% by weight in the carbon plastic samples (depending on the applied tool).

Podguiko N. A., Marakhtanov M. K., Khokhlov Y. A. Magnetron discharge application prospects as an electrons emitter in cathode-compensator for electric propulsion thrusters. Aerospace MAI Journal, 2019, vol. 26, no 3, pp. 167-177.

The subject of the presented article consists in assessing the prospects of magnetron discharge application as an electrons emitter for electric propulsion thruster cathode-compensator. This theme relevance is associated with the development of new stationary plasma thrusters (SPT) for the spacecraft operating on iodine, as well as low-orbit spacecraft employing outboard air as a working substance.

The paper assesses the energy aspect of magnetron cathode-neutralizer application for modern stationary thrusters. The highest operating voltages of the prospective dual-mode SPTs are 500-800 V. If a ten percent sacrifice of the propulsion system efficiency is possible with the view of increasing the service life and chemical resistance of the cathode-neutralizer, then the operating voltage of the magnetron cathode should be reduced to 120-180 V.

The article proposes a mathematical model of a magnetron discharge, on which basis a theoretical estimation of the magnetron minimum operation voltage and its dependence on the secondary ion- electron emission coefficient is presented. For a magnetron discharge with a copper cathode in the argon atmosphere, the minimum operating voltage equaled to 126 V. Besides, the minimum magnetic flux necessary for the discharge existence was computed.

An experimental study of plasma-forming gas pressure impact on the operating voltage value of the magnetron discharge was conducted for several options of the cathode material-working gas combination. These combinations were copper - argon, stannum - argon, stannum - argon-air mixture and aluminum - argon-air mixture. Minimum discharge voltage of 160-170 V was obtained when operating on an argon- air mixture and employing an aluminum cathode.

The performed studies allowed making the following inferences and recommendations:

  1. Cathode design should ensure optimal values of both the magnetic flux above the cathode surface and working gas pressure in the discharge area for the effective operation (minimum voltage).

  2. One of the ways to the electron cost in the magnetron cathode is the optimal.

Anisimov K. S., Kazhan E. V., Kursakov I. A., Lysenkov A. V., Podaruev V. Y., Savel’ev A. A. Aircraft layout design employing high-precision methods of computational aerodynamics and optimization. Aerospace MAI Journal, 2019, vol. 26, no 2, pp. 7-19.

Nacelle shape and engine position optimization was performed for Blended Wing Body aircraft (BWB). Aerodynamic characteristic computing method, used in the optimization procedure, is based on numerical calculations of the Reynolds-averaged Navier-Stokes equations. The EWT-TsAGI software, used for the flow computation, is based on the finite volume method of the second approximation order for all variables and includes monotonic modified Godunov scheme. The engine is simulated by the “active disks” method. Computations were performed on multi­block structured meshes with hexahedral cells. The power plant was designed with account for the initial requirements to the aircraft formulated in the AGILE project.

The developed optimization procedure consists of the two steps. At the first step, the isolated nacelle for the high bypass ratio engine is being developed and optimized for the cruise regime. Geometry of axially symmetric nozzle is described by the 11 parameters Parametric geometry of the inlet is specified by 7 control geometric parameters: 6 parameters specify the axially symmetric inlet, and one parameter (incidence angle) is employed for the air intake 3D design. The engine effective thrust is an objective function of optimization at the specified engine flow-rate constrains. To find the optimum solution, the Efficient Global Optimization method, based of simulation models, is used. It was shown, that SEGOMOE optimization method decreases the number of computed geometries.

At the second step, installation angles and the engines position over the airframe are optimized. A total of nine parameters is varied. The objective function is the effective thrust of the total layout (thrust minus layout drag) with the specified lift force constraint. An automatic structural mesh rebuilding is realized for the effective optimization procedure. The EGO based optimization algorithms require the initial points set calculating for the simulation model creation. It is shown, employing the large set of initial points (DOE) is more effective for the optimization process parallelization. Aerodynamic characteristics of the final layout with optimally installed engines were calculated. The main source of aerodynamic losses for the obtained configuration at the cruise flight’s Mach number of 0.85 is the compression shocks occurring due to the interference of the airframe with engine nacelle and between the neighboring engine nacelles. The subsequent studies should pay special attention to the aerodynamic interaction of the airframe and engine nacelles.

The described procedure was performed in the context of the third generation multidisciplinary optimization techniques, developed within the AGILE project. During the project, the new technologies were implemented for the novel aircraft configurations, selected as test cases for the AGILE technologies application.

Galkin N. A., Kondratenko A. N., Gaponenko O. V., Chiryukin E. V., Sviridova E. S. Methodical approach to aggregating computing of spacecraft manufacturing labor intensity. Aerospace MAI Journal, 2019, vol. 26, no 2, pp. 20-33.

For the purpose of the aerospace industry (AI) enterprises readiness to the implementation of State and commercial programs, it is necessary to perform an assessment of the production capabilities loading with regard to the labor costs for development efforts (DE) and spacecraft (SC) production.

The set task was being solved by the product capabilities conformity evaluation of the aerospace equipment (AE) head manufacturer with the federal target and government programs determining the required nomenclature and number of products, as well as the due dates of their production.

The spacecraft production is of a unit character with irregular repetition in the course of the years of production, where the products after the flight development tests (FDT) of the SC No 1 may have changes in the composition of the onboard equipment and design. The SC of manned programs production is individual and depends on the crew list and mission objectives.

Nowadays, based on the experience of the previous works and the prospective trends of development, engineers worked upon a number of unified space platforms (USP), which can significantly reduce the labor intensity of the SC manufacture. Development of the unified space platforms significantly reduces the volume and design cycles. In connection to the tried- and-true structural elements application the share of testing per one product set, which allows reduce the number of manufactured experimental installations.

The algorithm of SC manufacturing labor cost determining describes the sequence of labors costs computing of classification groups, containing tactical and technical characteristics of the products. The initial data on the actual and planned labor intensity of the SC production at the manufacturing enterprises were the products, both being manufactured and under development.

The first article of the stock-produced item manufactured for the flight development tests (FDT), at both single and several SC launch is assumed as a calculated labor intensity. The labor intensity calculation does not account for labor costs for the product manufacturing for performing inspection­sampling and periodical test.

The algorithm for the aggregating assessment of the SC production labor intensity is based on the layout solutions classification (constructive-technological schemes) of various types of SC. This algorithm has successfully proved itself within the framework of the “The SC Investments” research effort (RE) implementation, significantly increasing the accuracy of the loading prediction per product.

Calculation by the proposed algorithm is determined by a sufficient degree of technical solutions study at the stages of technical, draft and working projects, when analogous products, novelty factors or structural complexity of a new product can be determined.

Based on the obtained calculations, it is possible to evaluate and analyze the loading of the production capabilities of the main enterprise, specializing in the SC manufacturing. This will ensure the authenticity, completeness and estimation efficiency of the similar enterprises potential production.

Further development of this aggregating calculation algorithm of the DE and SV production labor intensity within the framework of assessing the feasibility measures of strategic plans for the technological development of the AI, the authors see in its automation. Besides, a coefficient characterizing technical level and industrial organization at the main manufacturing enterprises of the AE should be added to the algorithm. The proposed algorithm for the labor costs of SC production calculating was used by the center of integrated planning specialists of NPO “Technomash” in assessing the feasibility of the Russian Federal Space Program policy and the tasks of the Defense Procurement and Acquisition in 2017–2018, which confirmed its practical significance.

Calculated evaluation of labor costs for the SV production are recommended for employing as a basis for conducting technical and economic analysis, comparing alternative projects and developing perspective plans and programs. This labor input intensity algorithm will increase the accuracy of the enterprise predicted loading, resulting in the balance of the production program.

Kargaev M. V. Stresses computing in the main rotor blade based on the nonlinear loading model under static wind impact. Aerospace MAI Journal, 2019, vol. 26, no 2, pp. 34-42.

Wind is an important factor collateral to the helicopters operation. Due to a number of aeroelastic characteristics specifics, the non-rotating helicopter blades are sensitive enough to the wind impact. With this, the level of loads, acting on the blade, is commeasurable with the loads acting in flight. Traditionally, with high wind speeds mooring is employed to ensure the blades safety in parking position. It represents a flexible wire rope, which one end is fixed to the blade mooring unit, located as a rule at the blade tip section, and the other end is attached to the mooring node at the fuselage, or chassis of the helicopter. It represents a flexible wire rope, which one end is fixed to the blade-mooring unit, located as a rule at the blade tip section, and the other end is attached to the mooring node at the fuselage, or chassis of the helicopter.

The non-rotating main rotor blade according to its characteristics relates to flexible rods with deflections within the elastic deformations of the material commensurable with their length. This stipulates the necessity to consider the problem of the moored blade wind loading in a nonlinear formulation.

In this article, the parameters of the stress-strain state of the blade required for the mooring efficiency analysis are obtained based on a nonlinear model, which accounts for both geometric and aerodynamic nonlinearities. Computational algorithm for the initial nonlinear equation solution of the blade loading, developed based on the V.V. Petrov’s method of successive perturbation of parameters of was realized. The static loading is being considered as a process, developing at monotonous increasing of the loading parameter. The interval of load changing via its step- by-step application with small increments is split by steps, and for each step the linearized boundary value problem is being solved.

The blade deformed state, obtained in this manner at the current step, is assumed as the initial state for the next loading step. For error correction at each loading step, an iterative process is used, which allows performing calculations with a given accuracy.

The mooring effectiveness analysis was realized based on the computations performed for the moored and non-moored main rotor blades of the Mi-8 helicopter. The article presents the dependencies of critical gliding angles and limiting, under the strength condition, wind velocities values corresponding to them.

The article presents the dependencies of critical gliding angles and corresponding to them limiting, under the strength condition, wind velocities values. It also presents the dependencies of limiting velocities at the condition of a swaying absence condition on the characteristic section installation angle for the modes of blowing from both front and rear edges. The optimum installation angle, at which the range of safe wind speeds for the main rotor as a whole was the largest, was determined. This allows recommending to set the angle of the total step equal to the optimum one while a helicopter parking.

Alekseev V. V., Bobrov A. N., Kalugin K. S. Study of complex strength characteristics of gas turbine odels fabricated by additive methods. Aerospace MAI Journal, 2019, vol. 26, no 2, pp. 43-50.

Recently, the studies related to the additive technologies application in various industries, including aviation and space-rocket mechanical engineering, are considered promising. An indisputable advantage of additive technologies is minimization, and, in some cases, complete elimination of the need for parts machining, which significantly reduces both the time consumption and the finished part cost.

There are several basic 3D-printing methods, differing in the source material and technology of the parts formation. Recently, the parts production by selective laser sintering of metal polymer compositions powders (SLM-printing) has become topical.

The SLM-printing technology consists in layer-by layer deposition and sintering of powder on a special substrate. However, application of the selective laser powders sintering method is associated with problems of the porosity formation and a decrease in the strength of the parts produced. Thus, the issue of practical application for parts of the space-rocket and aviation equipment, created by the 3D-printing, still remains open.

To substantiate the possibility of 3D-printing application in turbines production for laboratory test benches on compressed air, the strength calculation of the turbine from PLA-plastic printed on the 3D printer were performed. The tests were performed to confirm the calculations results.

When developing a turbine 3D-model the rotor wheel geometry was selected, based on the prototype, which was used in the turbine structure employed in the laboratory test bench installation at the BMSTU for the laboratory works for studying the energy characteristics of active turbines.

Besides the external loads, the gas turbines rotor wheels load-bearing capacity is affected by loading conditions, such as gas temperature. However, the gas turbines employed in laboratory work benches on the compressed air are operating, as a rule, at low operating temperature of 30-50°C. Thus, the temperature stresses may be neglected while strength calculations of the turbine disk.

A 3D-model of the turbine under test was built with the Autodesk Inventor program. A finite-element model containing about 4.15 million elements was built for the above said model. Its strength analysis was performed with the Autodesk Simulation Mechanical 2019 module. The mesh thickening was reduced to the base of one blade only, since the load distribution is symmetrical. It can be seen from the safety factor distribution fields that minimum safety factor corresponds to the root sections of the blades, and it is no less than 3.3.

While theoretical calculations the modified safety factor n1, accounting for the effect of the part material porosity (for the case of its manufacture by 3D­prototyping) through coefficient k, was 3.28.

For tests performing, an axial active supersonic gas turbine was manufactured from PLA-plastic according to the SLM-printing technology.

For tests performing, a test bench, consisting of an electric motor, a voltage regulator, a tachometer, a video camera, as well as a turbine under study was assembled.

The methodology of the experiment conducting is as follows: the turbine is fixed on the motor shaft by the keyed and glue joints. When the motor is connected to the mains (220 VAC), the shaft and the turbine begin rotating. The rotational speed is changed by a voltage regulator connected to the motor circuit, and can aquire values from 0 to 24000 rpm, which corresponds to the voltage range in the motor network from 0 to 220 V. The data on the motor rotational speed are read from the digital optical tachometer. The experiment is being shot by the video camera.

The strength calculations of the axial supersonic gas turbine fabricated from the PLA-plastic by the SLM-printing additive technology revealed that the safety factor in operation conditions of laboratory test benches with compressed air was higher than the maximum allowable one for the considered unit.

As a confirmation for calculations, the turbine rotational speed during the test reached 24,000 revolutions per minute, which is the maximum possible value for the engine used in the tests. With this, visible defects were not detected in the turbine itself.

On the assumption of the performed studies it was established that the turbine manufactured using additive technologies can be employed for the laboratory text benches operating on compressed air.

Pronin M. A., Ryabykina R. V., Smyslov V. I. Experimental study of the aircraft forced vibrations while the engine blade break-away. Aerospace MAI Journal, 2019, vol. 26, no 2, pp. 51-60.

The presented article is a generalization of works relating to the ground reproduction of the force impacts on the aircraft structure, on the part of the engine with imbalance in case of the blade loss.

While ground testing the engine rotor does not rotate, and rotating force is formed by the fixedly installed vibration exciters. The immediate purpose of the experiment consists in frequency characteristics measuring, which associate the aircraft vibrations with the excitation force from the engine rotor imbalance. These characteristics are necessary for the computational dynamic scheme correction of the structure employed in loads computing in flight, possibly prolonged, while the blade break-away over the water surface. These computations are used for the aircraft safety evaluation while the blade loss.

The article presents the testing technique and facilities. The estimates of the modelling method applicability and its trustworthiness are given for the first time. The text is supplemented by the examples of real data of the tests.

The quantitative confirmation for the case of the ground experiment is given in the applicability esteems of the rotating inertial force reproduction by the harmonic forces stationary in space. At the same time, it was noted that the loads calculation while flight fluctuations, with a high level of the engine overloading, can not be based on either use of only relative acceleration of the blade, or the approximate theory of the gyroscope.

The circumstance of the experiment performing while the compulsory routine tests prior to its first flight was considered separately as practically the only possible for the experiment under consideration. The domestic tests on the aircraft with the engine blade loss modelling performed for the first time revealed the feasibility and possibility of their realization in conditions of dire time deficit prior to the first flight.

The presented details and features of the technique allow apply them in the future in the practice of such tests by the design bureau itself.

The main result is substantiation and practical confirmation of the possibility of reproducing on the ground the forced oscillations of an airplane after the blade loss, and while the mandatory regular modal tests.

Avdeev A. V., Katorgin B. I., Metel'nikov A. A. Energy characteristics computing technique for mobile multifunctional laser power plants based on fiber lasers. Aerospace MAI Journal, 2019, vol. 26, no 2, pp. 61-69.

Multifunctional Laser Power Plant (MLPP) should simultaneously solve the tasks of energy generation (Power Supply System (PSS)), radiation conversion and transmission (Laser System (LS)), and heat removal (Thermal Mode Supporting System (TMSS)). Meanwhile, the above said tasks are duly elaborated in modern projects. Thus, it is necessary to develop the MLPP design methodology, which accounts for the above listed subsystems interaction.

The article presents the developed technique for parameters analysis of the LS, TMSS and PSS subsystems of a multifunctional laser power plant, and results of its approbation while solving the task of space debris removal.

Computing was performed for the initial data Xtask based on the analysis presented in [1–5, 8]:

  1. acting on the Space Debris Fragment (SDF) with the orbit of HSDF = 1000 km by the ΔhSDF value required to its descent to [50; 900] km;

  2. the FSD velocity change per one pulse ΔFpulse of [0,1; 1,6] m/s;

  3. the impact distances range of RySDF [10; 150] km;

  4. the height difference of the SDF and spacecraft (SC) orbits of Horb [0; 150] km;

  5. relative FSD and SC closing-in velocity of Vrel [10,8; 12] km/s.

The following requirements to the MLPP operation mode (Υmode) were obtained for the initial data presented above: the energy density of [2,5⋅104; 2,5⋅105] J/m2 at the SDF; pulse duration of [2,7⋅10-9; 2,7⋅10-7] s; FSD exposure time of [2; 28] s; pulse frequency of [1; 1250] Hz.

The requirements to the sub-systems performance for this mode are as follows:

  1. LS (XLS): the output aperture dimensions of [0,5; 3] m; M2 and λ LS are assumed equal to 1 for calculations simplification; efficiency is [0.31, 0.59]; the laser pulse energy of [3⋅105] J; the threshold pulse power for one channel of 4,2⋅106 W; the beam strength of fiber of [0,01; 0,08] J.

  2. Requirement to the PSS generated energy is NPSS = [0,87; 5,7⋅108] W.

  3. The energy removed by TMSS is NTMSS = = [0,5; 4,5⋅108] W.

As a result, the inference cam be made that the data obtained while the technique application allow perform the MLPP parameters analysis for selecting the types of PSS, TMSS and their parameters, necessary for the MLPP required operation mode. Besides, this technique allows determining the limitations imposed by the PSS and TMSS subsystems on the LS pulse energy. The presented technique may be employed for the integrated assessment of the subsystems parameters and recommendations development of the MLPP application.

Vetrov V. V., Morozov V. V., Kostyanoi E. M., Os’kin A. S., Fedorov A. C. Caliber air-intake device for a flying vehicle with rocket-ramjet engine. Aerospace MAI Journal, 2019, vol. 26, no 2, pp. 70-80.

The work is devoted to the caliber air-intake device development for an aircraft with a rocket-ramjet engine moving in the dense layers of the atmosphere.

Analysis of the trends in the near-range aircraft with active start development demonstrates that one of the main directions of their improvement is the flight range increase The mass-size characteristics of the aircraft herewith remain at the same level, which does not allow employ the extensional development trends. Under these conditions, an important place is ranked by the trend related to the rational onboard energy utilization, within which framework the already classical solution are employed. However, the potential of these solutions is currently close to its limit.

In this regard, special attention is paid to propulsion systems (PS), which energy capabilities can be improved through the atmospheric air employing, and to a rocket-ramjet engine (RRE) in particular.

One of the key elements that largely determines the rocket-ramjet engine efficiency in total is the air-intake device (AID).

The proposed work novelty lies in the fact that the guided artillery shell (GAS) with its specific layout and functional features is considered as the object of study, and the search for a reasonable compromise between the requirements for the propulsion system and the shell as a whole is performed.

The problem of the AID rational configuration is being solved complexly based on the combination of numerical modelling methods and wind tunnel tests.

The initial variant of the twelve-nozzles caliber AID was developed for the pilot studies.

The works aimed at obtaining the throttle characteristics were performed.

One of the key features of the AID initial version was low efficiency of the boundary layer drainage system, which negatively affected its characteristics. In this regard, the initial model was modified to the second and later to the third option, characterized by an increased area of drain channels.

A positive result, manifested in an increase in the coefficient of the total pressure restoration by 14-20%, and the coefficient of air consumption by 11-27% for the third option, allowed form priorities for the subsequent AID configuration with a modified boundary layer discharge system and boxlike nozzles.

This solution allowed maintaining the aft location of the caliber non-regulated AID and the power plant with moderate total pressure losses and more stable air intake operation.

The performed studies allowed soundly obtain the most rational option of the caliber four-nozzle non­regulated AID for aft located RRE, integrated into the GAS structure. According to the preliminary estimates, this solution ensures provides a flight range increase by 25% compared to the GAS, equipped with the solid engine and bottom gas generator.

Osipov I. V., Remchukov S. S. Small-size gas turbine engine with free turbine and heat recovery system heat exchanger within the 200 HP power class. Aerospace MAI Journal, 2019, vol. 26, no 2, pp. 81-90.

The article presents a preliminary study of a small- size gas turbine engine (SGTE) of the 200 HP power class with a free turbine (FT) and a heat exchanger (HE) of the engine exhaust heat regeneration system. The presented engine is being developed primarily for unmanned aerial vehicles of various types and purposes (helicopters and airplanes).

The engine is available in two versions, namely, without a heat exchanger of the heat regeneration system, for the aircraft with short range and flight duration, and with a heat exchanger for the aircraft with long flight duration.

Characteristics calculations were performed for both the TSEr-200 engine with complex heat regeneration cycle and for the TSE-200 engine without heat regeneration [5].

Computational studies on sizes and type of the recuperative heat exchanger, rational for the given problem, were performed while the TSEr-200 engine development. A bundle of tubes was employed to determine basic dimensions of the heat exchanger matrix, on the assumption of the preliminary computation convenience (as the most worked out) [6].

The design arrangement of the heat exchanger and gas genera The structural layout of the heat exchanger and gas generator was developed based on the primary matrix computations.tor was developed based on the primary matrix computations. The heat exchanger includes 12 separate modules interconnected by the common manifold. Each matrix module is placed in individual casing.

Computational studies of various plate matrix types, as the most technologically worked-out at present and less expensive, were performed after the general layout developing. These computational studies were performed with the Ansys software package [11] using existing techniques for gas dynamic flows computing [12-15]. The computation results revealed significant hydraulic losses in the place of the flow turning inside the heat exchange matrix. Analysis of the results led to the necessity of studying the one- pass scheme of the coolant movement.

Computational studies of the heat exchanger option with the one-pass flow scheme revealed that total hydraulic losses for coolants did not exceed 3%. However, the layout of the heat exchanger with the engine was changed to organize the return of the air, preheated in the heat exchanger, to the combustion chamber. A distinctive feature of the proposed layout of the heat exchanger with SGTE is that the heat exchanger consists of 8 unified blocks, arranged in a circle among the three manifolds: the front one and two rear ones. All manifolds are cast and they are bearing elements of the engine.

For further work on the heat exchanger of the TSEr-200 engine, an option of the matrix with the “Frenkel packing” type plates of a single-pass scheme was adopted.

To confirm the feasibility of the heat exchanger project for the TSEr-200 engine, a matrix of the demonstration version of the heat exchanger with the “Frenkel packing” type heat exchange surface was developed. The module will be tested on the CIAM universal test bench as a part of the demo small gas turbine unit with the 4 kW capacity.

Ezrokhi Y. A., Khoreva E. A. Studying criterion parameters of the total pressure input non-uniformity impact on the thrust of a turbojet engine. Aerospace MAI Journal, 2019, vol. 26, no 2, pp. 91-98.

The presence of the total pressure non-uniformity may affect the basic engine parameters, and, in the first place, its gas-dynamic stability margin, as well as thrust-economic characteristics. Circumferential non­uniformity of the total pressure and its non-stationary component greatly affect the engine gas-dynamic stability. As for the engine thrust, the radial and circumferential effects are close enough, and non­stationary component does not affect the engine thrust at all. It allows employ one-dimensional approaches while this phenomenon modelling, and consider the impact of both stationary components of non–uniformity of the total pressure (both circumferential and radial) from the single methodological positions

In case of a non-uniform input flow, the flight-thrust decrease occurs for to the several reasons. Reduction of the general level of the total pressure along the engine passage, which leads to the pressure drop reduction in the jet nozzle pressure difference and, correspondingly, the decrease of the engine specific thrust may be assigned to the first cause. Besides, due to the general level of the input pressure reduction, physical air consumption reduction through the engine occurs as well.

The second reason of flight thrust reduction is associated with additional total pressure losses due to the “wash-out” of areas with various level of the total pressure in compression elements. It leads to the additional losses of the total pressure in compressor stages, which reduces the aircraft engine thrust to an even greater degree.

The authors suggested and justified criterion parameter Er for correct estimation of the thrust- economic parameters of the engine, operating in conditions of non-uniform input field of the total pressure. To the contrary of the W parameter, this parameter reflects additionally the relative values of the area, occupied by the zones with various total pressure values, being conditional indicator of the reduced pressure “concentration” per unit of the input area.

On a calculation example of the one-shaft turbojet with sufficiently conservative level of the design parameters the effect of the total pressure non–uniformity on its key parameters, such as thrust and gas-dynamic stability margin of the compression system was considered. This kind of engine selection is explained by the fact that to the contrary of the bypass jet engine, considered in the previous articles, the non-uniform field at the turbojet compressor inlet is considered as known, and its impact on a single compressor would be determinant for the whole turbo jet engine.

The performed calculation estimations revealed that the decrease in the engine thrust δR due to the non-uniform field of the total pressure at the inlet was completely defined by value of this parameter (dependence between δR and Er is almost linear), and also by the engine operating mode, such its shaft rotation frequency.

Zuev A. A., Nazarov V. P., Arngol’d A. A. Determining local heat transfer coefficient by a model of temperature boundary layer in gas turbine cavity of rotation. Aerospace MAI Journal, 2019, vol. 26, no 2, pp. 99-115.

Accounting for heat transfer specifics in flow­through parts of turbo-pump assemblies of liquid rocket engines (LRE) is a topical task. Currently, accounting for the specifics of the flow with heat transfer while realizing both potential and vortex rotary flow in the flow-through parts is implemented generally by the following methods: employing empirical equations, numerical and analytical methods for solving partial differential equations [1].

High temperatures of the working fluid lead to thermal deformations of components, including the turbine disks [18]. When designing the flow-through parts of the LRE turbo-pump units and assemblies, it is necessary to account for the temperature change of the working fluid flow along the working channel, since the viscosity parameter is a function of temperature and determines the flow regime and, as a result, losses, particularly disk friction and hydrodynamic losses in the flow-through part. The LRE turbo-pump energy parameters modelling is a topical scientific and technical task. The issues of the workflow parameters optimization, and the propulsion system mathematical model were reviewed in the V.A. Grigoriev’s treatise [19], where analysis of the models was performed, and merits and demerits for various design stages were disclosed.

A model for dynamic and thermal spatial boundary layers distribution with convective component for the combustion products turbulent flow in the LRE gas turbines rotation cavities is proposed. For combustion products, the Prandtl number is less then unity (Pr < 1), and dynamic boundary layer thickness is less than the thermal boundary layer one. It was assumed, that the temperature change and thickness of energy loss within the dynamic boundary layer border occurs due to the dynamic velocity transfer, and beyond the border – due to thermal conductivity only. This assumption complies well with the inferences of many authors [20, 21, 24]. Thermal resistance manifests itself over the entire thermal boundary layer thickness. Thermal resistance exists within the dynamic boundary layer borders due to the turbulent heat transfer, and beyond the border – due to thermal conductivity [24]. The distribution model of the dynamic and thermal spatial boundary layers with convective component is necessary for analytical determination of the local heat transfer coefficient in the LRE turbines rotation cavities.

The main objects of research, where the potential and vortex rotational flow is realized, are the flow­through components of LRE gas turbines such as inlet and outlet devices, as well as cavities between the stator and the working wheel [20].

An integral relation for the thermal spatial boundary layer energy equation, allowing integration over the surface of any shape, which is necessary for determining the thickness of energy loss, was obtained. The expressions for determining the energy loss thickness for thermal spatial boundary layer are necessary to determine the local heat transfer coefficients for the typical flow cases with account for the heat exchange.

Expressions for determining the local heat transfer coefficient in the Stanton number form for the straight linear uniform flow, rotational flow according to the rigid body law, and rotational flow of the free vortex of a power profile distribution for dynamic and thermal boundary layers parameters in case of Pr < 1 were obtained analytically.

Local heat transfer coefficient in the Stanton number form for straight linear uniform turbulent flow is

where m — is the turbulization degree of spatial boundary layer dynamic velocity profile,

– is the dynamic and thermal boundary layers ratio of the thickness, λ — is the coefficient of thermal conductivity,

 – the laminar sublayer coefficient of turbulent velocity distribution profile (obtained considering the two-layer turbulence model with a viscous laminar sublayer), Re — the Reynolds number.

Local heat transfer coefficient in the Stanton number form for rotational flow according to the rigid body law is

where ε — is the angle tangent of the bottom streamlines bevel, J — is the relative characteristic thickness.

Local heat transfer coefficient in the Stanton number form for rotational flow of a free vortex is

Analytical expressions for heat transfer coefficients agree well with the experimental data and dependencies of other authors [7–10].

The obtained analytical expressions well agree with the data of other authors and are necessary for engineering calculations while designing the LRE flow-through parts of turbo-pumps.

Baklanov A. V. Experimental study of the flame tube temperature state of a gas turbine engine multi-nozzle combustion chamber. Aerospace MAI Journal, 2019, vol. 26, no 2, pp. 116-125.

The flame tube walls cooling is one of the important components while organizing processes in the gas turbine combustion chamber. The combustion chamber operation reliability and engine endurance as a whole depend on the effective flame tube walls cooling. Convective-film cooling is one of the most widespread cooling systems. It includes the air film forming, which does not allow the hot gas interaction with metal and drawing heat from the backside of the wall due to the convection. The article presents the results of the studies on the flame tube walls temperature determining of the gas turbine engine operating on the gaseous fuel.

The article presents the combustion chamber structure of the converted aviation gas turbine engine serving as the gas pumping unit supercharger drive. The combustion chamber walls preparation and its testing as a part of a gas turbine engine were performed. The article presents the results on the flame tube walls temperature for the two operation modes of the gas turbine installation corresponding to 16 and 18 MW. The analysis of the obtained results allowed revealing that with the gas turbine installation power increase from 16 to 18 MW the temperature state of the wall did not drastically change. The walls temperature at the considered modes does not exceed 800°С, which indicates the flame tube sufficient cooling. However, the temperature distribution in various cross-sections was not of the similar nature. In some cross-sections maximum compared to the other cross-sections temperature was observed. It can be explained by the fact that the air passed through the conduit is split upon the hole flanging forming a vortex flow. As a result, the film-cooling loses its effectiveness, and the wall temperature behind the hole increases. The film-cooling effectiveness was determined at various sections on the flame tube walls. A technique for the wall temperature computing was developed, and comparison of computational and experimental results was performed.

Semenova A. S., Zubko A. I. Studying technical condition of the interrotor bearing with the SP180-M vibratory-diagnostic test bench after passing life tests. Aerospace MAI Journal, 2019, vol. 26, no 2, pp. 126-138.

The presented article deals with the studying of roller bearing after accelerated life test for the resource of 2000 hours.

To analyze the 5АВ1002926Р4 bearing vibration state a cpmprehensive analysis was being performed, including spectral analysis, RMS analysis in low-, medium- and high-frequency ranges, analysis of a pick-factor in low- and high-frequency ranges, and analysis of a “raw” signal of records.

The obtained test results allowed evaluate the bearing technical condition and transfer to further life tests with the test bench at “CIAM named after P.I. Baranov”.

It is well-known, that machines and mechanisms reliability depends essentially on their bearing assemblies working capacity. It is especially important for aviation engines as their bearing assemblies are one of the most responsible units often limiting an engine resource.

A reliable estimate of roller bearings technical condition, applied in gas turbine engines presents a problem at the aircraft building enterprise while both manufacturing and incoming inspection and fault detection. It concerns especially the indecomposable bearings since their technical condition estimation system currently in force is based mainly on the subjective methods such as checks on ease of rotation, or noise. Thus, the instrumental control methods implementation allowing not only estimate, but also forecast the working capacity during the operational process with more fidelity, is of current interest.

One of such instrumental methods is the quality monitoring of bearings vibration characteristics (a method of vibration diagnostics), operating with the specified loadings and frequencies of rotation. For vibrations measuring the vibrational converters, i.e. seismometers or accelerometer are used.

Methods of bearings vibrations measurement at control test benches are defined by the Standards [4, 5, 6]. The bearings condition is defined through the analysis of vibration signals [7].

Currently, various test benches, installations and diagnostics complexes, realizing this technique, have been developed, and being manufactured. One of them is the SP-180M test bench for roller bearings incoming inspection, being produced by LLC “Diamekh”. The test bench is meant for experimental studies for technical condition evaluation of separate bearings by vibration diagnostics method. These are the bearings of the first category (new), and bearings of the second category (being reinstalled), being installed in the engine while assembling.

The roller bearings, depending on the structure specifics of the product, where they are employed (parameters of inertia, stiffness and damping) may generate vibration of various intensity at various frequencies.

The vibration sensors mounting location and their characteristics significantly affects measuring results.

Thus, the SP-180 test bench has the single-type fixing of bearings, and fixed position of vibration sensors

The vibration signal amplitude, generated while interaction of working surfaces and external and internal rings of the bearing will depend on the rotational frequency of the test bench. Thus, its operating frequencies have the specified values.

Krylov A. A., Moskaev V. A. A technique for fluoroscopic control and analysis of technical condition of aircraft structural elements with honeycomb filler. Aerospace MAI Journal, 2019, vol. 26, no 2, pp. 139-146.

Application of various non-destructive testing (NDT) methods and means in conditions of operation is an effective method for sustaining the required reliability of aerotechnics. The structures with honeycomb filler from aluminum, steel and titanium alloys are employed in the modern aircraft airframes elements. Currently, x-ray method is the most effective one for such structures inspection. The article covers the non-destruction inspection technique performing of the aircraft structural elements with the honeycomb filler, and estimation of the images obtained by the fractal analysis.

The proposed technique consists of three main blocks:

1.   The block forming initial data, restrictions and assumptions:

a)   Variable parameters of the fluoroscopic installation (“Norka” X-ray TV unit);

b)   Invariable parameters characterizing design specifics of aircraft or control object (CO).

2.   A block of the fluoroscopic control methodology of aircraft design elements with honeycomb filler:

a)   A model for images base formation with account for the fluoroscopic installation parameters adjustment:

-     The CO X-raying schemes elaboration;

-    forming the images base when changing the anode voltage value at the emitter and the distance from the emitter to the CO. The best picture of the element with a honeycomb core was obtained in the framework of the experiment at U = 50 kV; F = 90 cm (F is a focal length, U is the anode voltage);

b)     A model for the image quality assessing:

-    Expert evaluation of the images database, with the concordance coefficient calculation [3];

c)     The CO fault detection performing:

-    Parameters adjustment of the “Norka” X-ray TV unit according to the image quality assessment model;

-    The CO fault detection according to the X-raying scheme;

-    The fault detection results decoding and analysis by fractal analysis.

3.   Recommendations formation on fault detection and repair of aircraft structural elements with honeycomb filler.

Fractal dimensions of the honeycomb filler without defects and the one with defects (the presence of moisture and geometry violation of the honeycomb filler structure boundaries) were obtained applying FracLab software.

The result of fractal dimension computing was obtained using the FracLab program by the direct geometric method of counting the cells of the honeycomb filler structure without defect and the one with defect.

The graph deviation of the structure with a defect from the linear dependence, characterizing the self­similarity of the structure under study, is twice as large as on the graph without a defect. It indicates the boundaries structure violation of the honeycomb filler. In addition, the graph with a defect in the double logarithmic coordinates has a kink, characterizing transition between different types of the structure (liquid presence in the honeycomb filler).

The additional information on the state of the system under study can be extracted by determining the self-similarity ranges limits.

Thus, employing the fluoroscopic control technique will allow performing the fault detection inspection of the aircraft structural elements with the honeycomb filler based on fractal analysis, as well as analyzing the obtained images base, and trace the dynamics of the honeycomb filler parameters changes, and defects of its internal structure, while the aircraft operation. However, it should be noted that the fractal analysis may be employed in the long term for automated parameters adjustment of the “Norka” X- ray TV unit, and the images base decoding without an operator.

Levochkin P. S., Martirosov D. S., Kamenskii S. S., Kozlov A. A., Borovik I. N., Belyaeva N. V., Rumyantsev D. S. Liquid rocket engines functional diagnostics system in real-time mode. Aerospace MAI Journal, 2019, vol. 26, no 2, pp. 147-154.

The hardware-software complex of the functional diagnostics system of the liquid jet engines operation during fire tests was developed. The system analyzes data in the real time mode. It deals with troubleshooting of units, structural elements or loops of a liquid rocket engine and determines the time instant of their occurrence.

Theoretical studies of the processes occurring in a rocket engine have been conducted since the 1930s. Differential equations reflect the dependencies between the engine parameters. The developed system employs the linearized equations of dynamics allowing accelerate computing and obtain numerical results in the real-time mode.

Each engine and each of its units are described by mathematical equations, on which basis the parameters values are calculated.

At each stationary mode, the averaged values of the operating engine measured parameters computed employing a mathematical model are compared.

If a calculated value deviates from the actual one, then there is a considerable probability of a defect presence in a unit, or in the entire engine. Functional diagnostics is based on this principle.

Modern measuring systems and high-speed computing systems are employed to diagnose engines in real-time mode.

The system consists of a hardware-software complex, an information system and a database, a telemetry signal emulator and an operator’s automated workplace.

The LRE functional diagnostics system solves the following tasks:

1. Increases the safety of the LRE fire tests conducting;

2. Determines the the engine functioning correctness in all stationary modes specified by the test profile;

3. Detects and localizes the malfunctions disrupting the proper functioning;

4. Identifies the engine “weak points”, such as elements or loops prone to structural or manufacturing failures.

5. Confirms the engine reliability before prior to its employing as a part of the launch vehicle.

The results of the emergency protection system and functional diagnostics system operation were compared. The proposed system has always found a failure before the emergency protection system did.

Petukhov V. G., Zhou R. . Computing the perturbed impulse trajectory of transferring between the near-earth and near-lunar orbits by the continuation method. Aerospace MAI Journal, 2019, vol. 26, no 2, pp. 155-165.

The problem of computing a two-impulse flight between circular near-Earth and near-lunar orbits with specified altitudes and inclinations over a specified time is considered. A mathematical model of motion, accounting for the Earth, Moon and Sun attractive forces as point masses and the second zonal harmonic of the Earth gravity potential at all spacecraft movement sections is used. The first velocity impulse is formed at the initial near-Earth orbit, and puts the spacecraft on the lunar flight trajectory. At the Moon passage instant at the minimum distance the second impulse is formed putting the spacecraft on the near-lunar orbit.

A numerical method for calculating two-impulse transfer between the circular orbits of the Earth and the Moon for a fixed time with account for the main perturbing accelerations has been developed. The method consists of the procedure for calculating the guess values, using the method of point-like spheres of impact, and the procedure for solving the boundary value problem for calculating the perturbed flight trajectory using the continuation method for reducing the boundary value problem to the Cauchy problem.

The advantage of the developed method is the procedure automation for selecting the initial guess values for solving the boundary value problem, and the computational stability of the solving process of the boundary value problem itself. The method revealed its efficiency and computational stability when calculating a series of transfers to a polar circular low lunar orbit of an artificial lunar satellite for various start dates and flight durations. The developed method may be applied for the design-ballistic analysis and operational planning of prospective lunar missions.

The article presents the numerical examples of trajectories computing for the flights between the low near-Earth and near-lunar orbits. Computing of the series of such trajectories allowed calculate the optimal start date and optimal flight duration, as well as dependencies of the required velocity impulses and longitude of the ascending node of the near-lunar orbit on start date and flight duration.

Nikolaeva E. A., Starinova O. L. Application of a heavy spacecraft with low-thrust engines for asteroid deviation from a dangerous trajectory. Aerospace MAI Journal, 2019, vol. 26, no 2, pp. 166-174.

The problem of asteroid danger for the Earth has long enough attracted the attention of scientists and society. Studying the traces of the space originated catastrophes on surface of the Earth and celestial bodies, as well as observing asteroids in the near-Earth space reveal the seriousness of asteroid hazard for the Earth civilization and the necessity of developing measures for its prevention.

The studies related to the issues of asteroid hazard encompass several trends.

Above all, detecting dangerous asteroids approaching the Earth (AAE) and their orbits determining. Currently, there are several national programs for optical observation of such bodies (NASA, LINEAR, ESA). It is assumed that these programs allowed detect great majority of such bodies with the size order of a kilometer or more. A whole number of such studies and projects envisage the countermeasures against these outlanders by their changing orbits or their destruction into small splinters, burning down in the atmosphere.

The urgency of the asteroid danger overcoming is beyond doubt at present, and the developing measures for its prevention should be one of the most important tasks to be solved by the humankind in the 21st century.

The goal of the presented work consists in developing a mathematical model, simulation and effectiveness analysis of the Earth protection systems to overcome the asteroid danger by the gravitational tractor.

To achieve the set goal, the following tasks were solved:

1)   Studying parameters asteroids approaching the Earth;

2)   Developing mathematical models of the joint motion of asteroid and all the bodies involved in the process of deviation from the dangerous trajectory (Sun, Earth, spacecraft, asteroid);

3)   Developing a software package, ensuring simulation and visualization of the proposed method of the asteroid danger counteracting;

4)   Analyzing the simulation results of the proposed method of the asteroid danger counteracting.

The main results obtained in the work are as follows:

-     a mathematical model of the motion of bodies, with perturbations from the gravitational tractor acting on them: a variable mass asteroid, spacecraft, the Earth and the Sun, with account for the gravity of all bodies;

-     based on a a mathematical model of the bodies motion system, the software package “Simulation of the Earth protection systems functioning to overcome the asteroid hazard” for the asteroid trajectory simulation by the selected method of the asteroid danger overcoming in heliocentric coordinate system was developed;

-     simulation of the potentially hazardous bodies deviation method (asteroid deviation by the gravitational tractor) for the 99942 Apophis asteroid was performed with the developed software complex “Simulation of the Earth protection systems functioning to overcome the asteroid hazard”;

-   the simulation resulted in obtaining the flight trajectories of all the bodies of the system under consideration (the Earth, the Sun, asteroid and a spacecraft) and heliocentric movement parameters;

-   the efficiency analysis of the selected method was performed.

Ermakov V. Y. Studying the effect of the beam aerial drive control algorithm on its vibration activity onboard a spacecraft. Aerospace MAI Journal, 2019, vol. 26, no 2, pp. 175-181.

Modern space vehicles (SV), as a rule, include bearing-out structures of slight rigidity. These are solar batteries, antenna-feeder devices, elements of thermal conditioning systems. Actuators and special purpose units, as well as units of technological and support systems are being placed inside the SV hull. SVs are exposed to vibrations from the external and internal perturbance sources both on Earth and in orbit. The feature of the SV loading in orbit is low-force spectrum of perturbances up to tens of Newtons with frequencies from fractions of hertz to hundreds of kilohertz. Vibrations may have deleterious effect upon both orientation and stabilization accuracy, and movement dynamics including various types of orbital maneuvering. These perturbances might be created, for example, by operation of the narrow-beam aerial (NBA) drive, which leads to occurrence of elastic vibrations of the structure and mounting faces of the precise equipment. While the observation session onboard an SV, mechanical disturbances, stipulated by operation of aggregates with non-balanced masses, may occur. This may affect both the orientation accuracy of the SV itself and equipment elements which may degrade the quality of the registered information, and introduce significant error to the SV angular position measurements, obtained by the orientation and stabilization control system. This, in turn, may make the SV mission target task performance impossible. To reduce these perturbances an algorithm for the NBA drive operation for the “Spectr-R” type SV was developed. Dynamic analysis of data obtained for the suggested algorithm and conventional was performed. Positive results of the suggested algorithm, tested on the “Spectr-R” type SV are demonstrated.

Silaev M. Y., Es’kova E. A., Gerus D. S., Remshev E. Y. Acoustic emission method application while determining mechanical characteristics of the brnicrsi-2,5-0,6-0,7 wire for elastic elements production. Aerospace MAI Journal, 2019, vol. 26, no 2, pp. 182-192.

A great number of electromechanical systems, an important part of which represents an elastic element from bronze, are applied in aerospace technology.

Severe requirements are placed to the physico- mechanical characteristics of these parts. The existing standard methods for mechanical properties determining are not sufficient for such products.

Acoustic emission method is one of the promising methods to solve this problem.

Acoustic emission is radiation of mechanical waves by the material, caused by local dynamic rearrangement of its structures This method is non destructive.

Beryllium bronze is used as a rule in special products. This project studies a cheaper substitute for Nickel-chromium-siliceous bronze.

Besides, mechanical tensile tests of the wire with parameters registration of acoustic emission were being conducted. Bronze was subjected to various heat treatment to select the optimal mode.

As the result of this work, the microstructure of the samples was studied for various thermal treatment modes. It was revealed that the acoustic emission parameters were the figures of strength and plasticity.

The strength and plastic characteristics are related to the grain size by the dependence proposed by Hall- Petch. This dependence modernization allowed adopt the stress at maximum value of the pulse amplitude up to the yield point achieving as the stress corresponding to the dislocations motion start.

The possibility of determining the microplastic deformation starting of wire samples by AE method was established. Based on the obtained regularities, it was revealed that the number of signals is a characteristic of strength, while the amplitude is a characteristic of plasticity. The Hall-Petch dependence modernization may allow developing a technique for operational control of microstructure in the release of special products.

Kovalev A. A., Zinova V. V. A tool-blank state monitoring while cutting process using kalman filter. Aerospace MAI Journal, 2019, vol. 26, no 2, pp. 193-204.

The article discusses the issue of the cutting process monitoring possibility using the acoustic emission method by processing the input signal using the Kalman filter. A filter was selected to solve the problem. The inference was drawn on the possibility of monitoring the gradual wear-out and chipping of the cutting edge by Kalman filter.

The article consists of three main parts: introduction, the main part, and conclusions.

The introduction considers the problems occurring while automating the technological process of blank parts machining. With this, a part of events is deterministic, while the other part is random. Thus, to ensure the required quality level in the process of automation the cutting zone continuous monitoring is required. It will allow making changes directly while blank parts machining technological processes executing.

The main part of the article presents operation principles of the monitoring systems, based on the

system harmonic oscillations analysis. Various filtering algorithms were considered in particular.

The Kalman filter was chosen as the object of study as one of the most common algorithms in the theory of automatic control. The goals were set and the tasks were formulated. Criteria are being set, which the desired filter should meet for continuous for the cutting area monitoring. The main approaches to solving filtering problems are being considered and compared with the Kalman filter. The inference is being drawn that this filter is the most suitable for solving the set problem. Measurements are being performed, the results, processed by the three Kalman filters versions are being analysed, and one of them, best meeting all the necessary requirements is being selected.

The conclusions formulated the possibilities for Kalman filter application for continuous monitoring of the tool blank state in the cutting process and gave recommendations to the future work, and filter coefficient selecting in particular.

Khryashchev I. I., Danilov D. V., Logunov A. V. Developing a sparingly doped high-temperature nickel alloy for gas turbine blades. Aerospace MAI Journal, 2019, vol. 26, no 2, pp. 205-218.

Development of mono-crystal high-temperature nickel alloys for gas turbine blades and vanes is one of the leading trends ensuring enhancement of parameters, efficiency and reliability of modern gas turbines.

Currently, one of the most widely used alloys for turbine blades manufacturing is the second-generation domestic ZhS32 alloy with Re content of about 4.5%. The goal of this work consists in the alloy creation with the equivalent level of heat resistance, but with no expensive elements, such as rhenium and ruthenium.

Besides, determination of the optimum heat treatment mode based on experimental works in production is a costly method.

Computing diffusive activity of doping elements may allow decrease development costs and optimize the regime for realizing the total potential of the alloy, embedded while it’s designing.

Analysis of nickel high-temperature alloys was performed while this work execution, and an optimal scheme of doping process to achieve maximum heat resistance was selected. With application of the computer aided method for high-temperature alloys optimization a new sparingly doped alloy for gas turbine blades, meant for operating at the temperatures up to 1050°C. The alloy is distinguished by high structural stability and economical use of doping elements. The new sparingly doped alloy relates to the first generation. With this, it complies with the third generation GS32 alloy by the level of heat resistance at 1000°C.

In the course of the works, development of nickel- based heat resistant alloys has been analyzed and an optimum alloying system has been selected to achieve the maximum heat resistance of the alloy. With the use of computerized optimization method of heat resistant alloys, a new lean alloy has been developed for gas turbine blades intended for operation at temperatures to 1050°C. The alloy exhibits high structural stability and efficient use of alloying elements. A new lean alloy is the first-generation alloy but its heat resistance at 1000°C corresponds to that of the third-generation alloy ZhS32.

A unique techniques for determining the diffusion coefficient of doping elements, and, based on the obtained data, for determining an optimal duration of the thermal treatment, were developed.

The microstructural studies of a new sparingly doped SLZhS32 alloy were conducted; a thermal treatment mode was tested with account for the diffusion processes kinetics; the samples were fabricated and strength tests were conducted.

The developed new sparingly doped alloy can be widely used for gas turbine blades manufacturing, ensuring the cost reduction without deterioration of the alloy operational properties.

Aydemir T. ., Golubeva N. D., Shershneva I. N., Kydralieva K. A., Dzhardimalieva G. I. Formation, structure and magnetic properties of nanocomposites obtained by Fe(III)Co(II) cocrystallized complexes thermal decomposition. Aerospace MAI Journal, 2019, vol. 26, no 2, pp. 219-228.

Considerable interest in d-elements nanoparticles well as the possibility of creating magnetic carrier with is stipulated by their magnetic properties specifics, as high information recording density on their basis.

Magnetic particles are widely used in biomedicine, and ferrous oxides (magnetite and maghemite), possessing high biocompatibility, play exceptionally significant role. Iron- and cobalt-containing particles are characterized by high values of coercive force and magnetic susceptibility. For example, for magnetite Fe3O4, the saturation magnetization (δs, Ms) is 92 eme⋅g-1, and for γ-Fe2O3-74 eme⋅g-1, the coercive force magnitude for anisotropic nanoparticles of the latter ranges from 200 to 400 Oe.

The structure and properties of metal-containing nanocomposites obtained while thermal transformations of Fe (III) Co (Il)-acrylate complexes were studied in this article.

It was shown that thermal transformations of the complexes under study included the stages of dehydration, solid phase polymerization and decarboxylation of the forming metal polymer. The solid phase product of the complexes thermal transformation are metal-containing nanoparticles, stabilized by carbonized polymer matrix. The crystalline nanostructured phases are Fe3O4, CoFe2O4 and CoO. The average crystallite size is 10 nm. Magnetic properties of the obtained nanocomposites also were studied. Hysteresis loops measured at temperatures below 200 K are open and displaced to a negative field. The coercive force and residual magnetization are 0.18 T and 15.5 mT, respectively.

An original approach consisting in combining nano-size metal particles synthesis with its stabilizing polymeric shell in situ was developed. The approach is based on metal containing monomers homo- and copolymerization in the solid phase with subsequent controlled thermolysis of the formed metall-polymers.

Accordingly, matrix-stabilized metal oxide nanoparticles were obtained by the method of polymer-mediated synthesis. In the nanocomposite obtained at 643 K and conversion of Δ m = 42%, the crystalline phase contains nanoparticles of ferromagnetic oxides Fe3O4 and CoFe2O4, and CoO antiferromagnetic nanoparticles. The nanocomposite microstructure includes polycrystalline agglomerates with sizes of 30 nm, consisting of individual nanocrystallites with an average size of 10 nm. The magnetic properties of the obtained products depend on the nature of the components, the temperature and the magnitude of the applied magnetic field. The coercive force and residual magnetization at room temperature are 0.18 T and 15.5 mT, respectively. The strong dependence of the magnetic characteristics on the phase composition, temperature, and magnetic field suggests that nanocomposites of this type are of interest for the sensor materials production for aerospace and biomedical applications.

Antipov V. V., Prokudin O. A., Lurie S. A., Serebrennikova N. Y., Solyaev Y. O. Sial interlaminar strength estimation based on the results of the samples’ three-point bending tests. Aerospace MAI Journal, 2019, vol. 26, no 2, pp. 229-238.

Laminated aluminum-glass plastics (GLARE, SIAL) are promising structural materials for application while aircraft structural elements manufacturing These composite materials represent layered panels formed by thin layers of fiberglass and aluminum alloy. Compared with metals, SIALs possess increased specific strength, long-term strength and fire resistance. Studying the dependence of SIALs mechanical properties on the parameters of their reinforcement is an important task, which solution is necessary for the structures’ design and strength computation. One of the important characteristic, determining the SIALs structural properties, as well as the other composite materials, is interlaminar strength.

The samples testing on the three-point bending by the “short beam” technique is one of the simplest techniques for determining the interlayer strength of composite materials. This method is widely used in composite structures research and development, since it does not require the application of complex experimental equipment and strain gauges. At the same time, the interlayer strength is an important parameter from the designing viewpoint, as it is used in formulating the strength criteria of composite materials The interlaminar cracks occurrence may lead to a decrease in the bearing capacity of structural elements, and further to their destruction, for example, by the local buckling mechanism.

However, such a simple method as testing on three-point bending holds certain disadvantages associated primarily with the fact that during such tests a complicated stress state is realized in the samples, that is, not only the interlaminar shear stresses occur, but the also tensile / compressive stresses arise as well, leading to errors in determining the materials characteristics. Besides the above mentioned errors associated with non-uniform tensed-state of the samples, the complexity, occurring while samples testing on the interlaminar shear, consists in the fact that the interlaminar strength being determined while testing proves to be not a constant of the material, but it depends on the distance between the supports. This problem is known both for conventional composite materials and for metal-polymer composites. It is explained by a decrease in the tangential stresses actually acting in short samples (according to the standards the samples relative elongation shoul be of 5 to 10), compared to the classic beams models. These models assume the constant value of the shearing force, and, correspondingly, constant values of tangential forces (up to sign) along the sample length. Thus, application of the traditional relation for estimating transversal shear stresses acting in a beam, according to the formula 3 P / (4 b h), leads to the increase in the apparent interlayer strength of the material. Besides, the sample length impact on the results of the tests on the interlaminar strength is explained by:

1)    Stress concentration nearby the supports;

2)   Statistical dependence of strength on the sample size;

3)   The interlaminar cracks occurrence not on the neutral axis of the sample and

4)   Special dependence of interlayer strength on the parameters of fracture mechanics.

The article proposes a scheme for SIAL testing on the interlaminar shear strength by the short beam technique. These tests employ the samples with the large number of layers and unidirectional reinforcement scheme, which allows reduce the error of experiments while employing the standard equipment. The samples apparent interlaminar strength, depending on the distance between the supports, was determined by the results of the tests. Based on the calculations, the accordance of the obtained experimental data and theoretical estimates is demonstrated. The calculated SIAL interlayer strength value was of ~ 60 MPa, which corresponds to the typical interlayer strength of polymer composites. However, while testing the destruction was being realized at the contact boundary of metal and composite layers, which allows affirm that the found interlayer strength value is a characteristic of the metal / composite contact.

Shved Y. V. Determining technique for optimal rigging angle and aspect ratio of the soft wing with sling support. Aerospace MAI Journal, 2019, vol. 26, no 1, pp. 7-18.

While developing paragliders and gliding parachutes many issues on the optimal selection of the airfoil, its relative thickness and twist over the span, the law of the wing arc distribution and its shape in the sweep, length and slinging arise. Selection criteria for of some of these parameters may be transferred practically without changing the methods, rather explicitly elaborated for the historically earlier appeared aerial vehicles with balancing by the payload weight (hang-gliders). However, the paraglider, also related to the flying vehicles balanced by the load, has some specifics, since it employs momentless carrying shell.

The parameters estimates of the aerial vehicles with the soft wing and sling support with various working-out degree are presented in [5-19]. However, the issue of working-out the simple and vivid analytical technique for obtaining optimal characteristics of the above said aerial vehicles, which does not employ iteration approximating and general empirical assumptions, still remains open. The article is devoted to the study of some aspects of this technique.

author proposes to perform the calculation in the following sequence:

  1. It is assumed, that in the assigned flight mode, the wing has the required angle of attack. Aerodynamic coefficients of the airfoil Cxp and Cya for the specified mode are being elected.

  2. Based of the obtained coefficients, the glading angle is calculated according to the expression proposed in the article. Then, with account for the obtained gliding angle, the gliding speed is calculated using the following expression.

  3. After selecting several options of the wing profiles and aspect ratio the comparative calculation of the flight quality is performed. With too small values of the wing lift coefficient, the main contribution to the resistance is brought by the air-dropped cargo and slings. If the Cya is too large, the inductive resistance becomes prevalent. Consequently, for each wing aspect ratio, the system slings and cargo type it is possible to determine the optimum carrying capacity of the designed wing profile. Conversely, it is possible to determine the optimal aspect ratio with given the remaining design characteristics.

  4. After the final selection of the profile, by the center of pressure on the wing MAC (middle aerodynamic chord) is determined. Further, with account for the obtained coordinates of the center of pressure on the MAC, the coordinate of the wing suspension relative to the load center of gravity is determined by the proposed formula.

The article demonstrates also the independence of the of self-balancing wings angle of attack from the thrust magnitude. This conclusion is based on the fact, that for the angle of the slant of the slings relative to the center of the pressure of the MAC in the horizontal flight mode under thrust and in the gliding mode, identical equations were obtained.

In [1] the algorithm for static parameters calculation of the motor flight vehicle with a soft wing is presented. In the presented article it was expanded for the gliding descent mode.

Brutyan M. A., Potapchik A. V., Razdobarin A. M., Slitinskaya A. Y. Jet-type vortex generators impact on take-offand landing characteristics of a wing with slats. Aerospace MAI Journal, 2019, vol. 26, no 1, pp. 19-26.

To increase maximum lifting force coefficient of the aircraft wing with fixed geometry, it is reasonable to use the flow control concept. For this purpose, the new way of flow control about a wing with deflected slat, suggested by authors, is being studied experimentally and numerically. A number of slanting holes (along the flow and longwise a wingspan), through which the air jets are blown-out, is made to create vortex cores at the nose section of the upper surface of the wing’s main part, which opens while the slat root section moving-out. The pilot experimental studies of the new method of the wing with slat flow- around at the take-off and landing modes were performed on a model of a modern long-range aircraft with mechanized wing with moving-out slats and flaps.

The slats are made along the wingspan with a gap along the motor-nacelle pylon. The aircraft model testing while the landing state of the high-lift device with jet-type vortex generators and without them were performed with ADT T-106 TsAGI, equipped with aerodynamic scales. Slats and flaps were in landing state; with corresponding deviation angles of δsl = 24° and δfl = 36°. Weight measurements of aerodynamic characteristics were performed at the Mach number of the incident flow М = 0.15. It corresponds to the Reynolds number value of Re = 3.1⋅106 at the pressure pumping up to 5 atm in the working section of the tube. The angle of attack was being changed from 4 to 26°.

Numerical simulations of jet-type vortex generators impact on the wing flow-around pattern in a take-off and landing configuration were performed. Numerical calculations were performed to compare the experiment and the expanded range of the studied parameters. The well-known ANSYS CFX software based on the numerical solution of averaged Navier-Stokes equations for the compressible perfect gas with two-parameter SST turbulence model was used. The flow was considered turbulent starting from leading edge. The surface of the model was assumed adiabatic; the viscosity-temperature relation was determined by Sutherland’s law with the constant C = 110.4 K. The number of computational nodes used for the flow-around modelling with streams increased approximately up to 68 million.

The performed studies of passive technique for streams forming by the air blow-by from low the wing underside to its upside at the numbers of Re = 3.1⋅106 and М =0.15 revealed the possibility of the maximum lifting force coefficient increase.

Artamonov B. L., Shydakov V. I. Algorithm of transient flight modes performance by convertiplane. Aerospace MAI Journal, 2019, vol. 26, no 1, pp. 27-40.

The article considers the Project Zero convertiplane implemented according to the structure with two rotary screws positioned in the fixed wing. The screws are driven by electric motors powered by batteries, and controlled by a common and cyclic step. Electric transmission of the Project Zero convertiplane allows smooth change of the propeller rotations while transient flight mode performing with minimum required power.

The article analyzes control laws of screws, which allow performing transient flight modes from helicopter to aircraft without losing altitude at minimal engine power consumption. The described algorithm uses the results of experimental studies of the convertiplane body model in the t-1 MAI wind tunnel by th angle of attack at various rotation angles of the screws axes of rotation relative to the fuselage longitudinal datum line. This allowed reduce the problem to a system of transcendental equations of the convertiplane motion, which was solved numerically by successive approximations method. The aerodynamic characteristics of the propellers located in the ring fairings are being computed based on the disk vortex theory.

It is shown that while the convertiplane transition from hover mode to flight mode the screw control laws are of a rather complex character, and may be realized only by employing automation. The obtained convertiplane control laws at the transient flight mode are effective from the energetic viewpoint. The power consumption in the transient process endpoint is three times less than in the hover mode, which allows further convertiplane flight speed increase.

Gubernatorov K. N., Kiselev M. A., Moroshkin Y. V., Chekin A. Y. Studying elements reliability impact on the aircraft functional systems architecture. Aerospace MAI Journal, 2019, vol. 26, no 1, pp. 41-50.

The reliability of the more electric aircraft and its systems must not be less than the reliability of the conventional aircraft and systems to meet the required safety level. The level of the system reliability is specified in the part 25 or 23 of FAR. Power system of the more electric aircraft is a very important system due to the approach of ensuring the dragging and operating systems, such as control system and landing gear system. The weight-size parameters and the reliability of the more electric aircraft power system are opposite and depend of the power and energy system architecture.

This article demonstrates an approach to the architecture design of the more electric aircraft power system, that follows modern trends and ensures the required safety level and minimum volume and mass using state-of-the-art technologies, such as permanent-magnet generator and power electronics.

The current reliability level of power supply system elements (generators, rectifiers) cannot provide an extremely improbable event of the functional failure of the power generation system. Thus, the power supply system designers are forced install emergency (alternative) power sources such as batteries, a RAT, and auxiliary power unit, providing power to important systems to complete the flight and perform a safe landeing. These systems for example represent to an engine and an aircraft control system. The emergency (alternative) power supplies and the associated cables and switching system possess a considerable mass and volume. For example, the modern aircraft such as Boeing-787 and Airbus-350 have a very complicated power system to meet the required level of reliability. So these systems employ additional power converters, batteries, ram-air turbines and complicated distribution system. All of these have mass and occupy the aircraft volume.

Here is another example. The MC-21 emergency energy system weight is about 85% of the main energy system weight.

Hence, we can conclude that in order to meet the safety requirements, the power supply system designers should install almost one more power generation system onboard.

It is worth adding, that besides generation function the emergency power sources perform some other functions such main engines on-board starting, voltage ripples smoothing in the DC power systems with batteries and other. However, these functions are not taken into account in the presented article. The main attention is paid to the electric power supply system architecture developing, which meets the safety requirements, and contains minimum set of components to reduce weight-size parameters at large.

Shustrov T. L. Simulation as a substantiation of the trace contaminants removal system selection. Aerospace MAI Journal, 2019, vol. 26, no 1, pp. 51-63.

The article is dedicated to one of the most important problems while preparing any potential long-term or interplanetary space missions, namely the inefficiency of the life support subsystems employed at the habitable spacecraft. The article focuses mainly on the trace contaminants removal system (TCRS) being an important element of the space object life support system. It purifies the atmosphere of an object from any contaminants, and keeps it at the predetermined chemical balance.

The main hazards requiring permanent system regeneration and its keeping at the maximal possible technical level are as follows:

  • Atypical habitability conditions at the space object;

  • The crew impact (chemicals secretion as a metabolism result), as well as the spacecraft itself (chemicals emission as a result of degradation of coverings, used for internal plating, ) on the artificial living space.

The artificial atmosphere of any isolated airtight object is affected by its inhabitants, which could lead to the sensitive equipment failures, destructive emergencies, and deaths among the crewmembers. The presented article suggests employing simulation model as an attempt to improve the design and production of the future trace contaminants removal systems. The model allows computing the resulting amount of trace contaminants formed by any number of potential sources. The model structure provides the designer with maximum flexibility while the process regulation, which might help while creating individual configuration of the trace contaminants removal systems with account for the space mission scenario.

The article presents mathematical/technical description, structure, and examples of the simulations results. Most subprocesses are at the final stage of testing. The simulation results correspond to the telemetry data from the space station. In the future, after the final testing the authors plan to create the “artificial helper” for the model that will perform automatic selection of the trace contaminants removal system based on the results obtained after the simulation.

Gaponenko O. V., Gavrin D. S., Sviridova E. S. Structure analysis of the strategic plans of the space-rocket industry development by method of space functional and industrial technologies R&D classification. Aerospace MAI Journal, 2019, vol. 26, no 1, pp. 64-81.

The task of a subject for study classification arises while information analysis support of strategic programs for space-rocket industry technological development and managerial decision making on a sectorial level. In this case, it is an aggregate of scheduled measures, namely research and development work (R&D) on the cosmonautics and aerospace industry technological development.

The existing R&D classification in Federal target and Government programs (FTP) “Military-industrial complex development of the Russian Federation” does not fUlly reflect the structure of program activities, i.e. an aggregate R&D technological development R&D, and is applicable only to industrial technologies. In the Federal Space Program (FSP) the R&D is classified according to the target purpose of finished products. The R&D classification employing is not applicable in other FTP and vice versa. In the authors’ opinion, classification according to technological trends is the most efficient.

In domestic practice of analytical studies associated with the space activities technological R&D are subdivided into the intrinsic cosmonautics technologies (the space functional technologies), and industrial technologies for the space engineering development (the space industrial technologies).

There are also the third system-wide studies in the programs of cosmonautics and rocket building development, besides the functional and industrial technologies. These include complex system analytical research.

The forecasting of the space technology development without accounting for the capabilities of aerospace industry risks turning into vain dreams and fiction, and vice versa, the development of industrial production with no strategic targets in the form of promising space technologies may lead (and already leads) to creation of inefficient and economically unviable production structures.

The same technology, depending on the stage of the product life cycle of aerospace technology, can be attributed both to the target technology and to the of industrial production technology.

The unified R&D classification system of aerospace functional and aerospace manufacturing technologies and system-wide research effort is advisable. There is a necessity of a unified classifier for the cosmonautics development strategic programs (FSP, state programs “Development of the MIC”, strategic programs and plans of other governments) in parts of R&D sections.

The article proposes a unified classifier of space- rocket and manufacturing technologies. It is based on the classification features of technologies used by NASA in the technological road maps of 2015.

The classifier was realized by the authors in the form of an object-relational database on PostgreSQL. The database is switched as an external data source to Excel, and further the analytical capabilities of the free Excel table mechanisms are used.

A comparative analysis of R&D technologies performed by NASA, the European Space Agency and State Space Corporation “Roscosmos” within the framework of long-term strategic programs of space activity was performed using the developed classifier. The classifier allows also compare the same technological trend in different programs.

Besides the number of works the developed classifier allows analyzing their financing, starting/ ending dates and starting/ending level of technological readiness by technological trends.

The classifier allows reveal the technological development trends, to which most attention is paid in the states participants of the space activities, and vice versa which are related to unessential, and their studies are not financed by strategic programs. The structural specifics of each of the considered programs of technological development can be analyzed.

Practical implementation of techniques, associated with program events classification forming and scientific-methodological support of the strategic programs of national space-rocket industry development (including application of the classifier suggested by the authors) with subsequent analysis of the obtained classes will contribute to the managerial decisions effectiveness in Russian space-rocket industry, and eventually in rational implementation of the State budgetary funds allotted for this purpose.

Nikitin S. O., Makeev P. V. A project of the “Synchropter” type high-speed helicopter with pushing air propeller. Aerospace MAI Journal, 2019, vol. 26, no 1, pp. 82-95.

Due to the helicopters ability to perform vertical take-off and landing, as well as effective operation while hover mode, they became indispensable practically in all regions of the world. With that, the requirements for the helicopters flight performance enhancement become ever more acute, primarily concerning the increase in speed and range.

Currently, a number of rotary-winged aircraft structures of vertical take-off and landing, realizing increase in speed and flight range, are under development and in some cases at the stage of testing and batch production in leading world countries. There is a number of concepts and technical solutions, mainly in the field of aerodynamics, allowing increase a helicopter cruising speed. In this regard, the exploratory research and these projects implementation development are highly relevant.

The presented work is devoted to creation of a project of a perspective passenger high-speed aircraft with vertical take-off and landing based on a helicopter with intermeshing rotors and a pushing air propeller.

The project employs a set of the following technical solutions:

-    The blades rotational speed reduction (from 220 to 180 mps) as the flight speed increase; special arrow­shaped tips setting on the blades to reduce to zero the probability of a wave crisis on the advancing blades with flight speed increasing;

-    Balancing the unbalanced lateral tilting moments on the two rotors of a “synchropter” scheme, rotating in opposite directions;

-     Application of rotors with elastic torsion sleeves;

-    Application of a system of the blades individual control to prevent the flow disruption on the retreating blades;

-    The aircraft fuselage layout with account for the specifics of the scheme with low frontal resistance at near-zero angles of attack;

-    Application of a propulsion propeller with maximum efficiency in operating conditions.

The capabilities of modern computer-aided design technologies were demonstrated while the project developing. The main emphasis is made on the aircraft dynamic designing with implementation of modern tendencies of the high-speed helicopters development. The main limitations and possible ways for the helicopter speed increase implementation were considered. The article presents the computational results of aerodynamic characteristics with account for the decisions made.

The developed project has the following characteristics: the take-off weight of 6500 kg, payload mass of 1000 kg, maximum speed of 420 km / h, static ceiling of 4700 m, dynamic ceiling of 5600 m, and flight range of 1228 km.

The obtained results indicate the achievement of indicators close to the modern world level, demonstrated on similar developed helicopters.

The developed project has prospects for further flight performance improvement by improving the‘ aerodynamic characteristics of the fuselage, propellers, as well as exploiting more fully the capabilities of the individual blade control system.

Erkov A. P. Buckling of stepped beams. Aerospace MAI Journal, 2019, vol. 26, no 1, pp. 96-110.

The article discusses the problems of stability of two types of beams of variable stiffness: with a stepped change in cross section with two zones and with a step change in section with three zones. Simply supported boundary conditions at two ends are considered, as well as with embedding at one end and with a free second end. Beams of isotropic material and of the laminated composites are discussed.

To study the stability of beams of variable stiffness, the Ritz method was used. Beams with the ratio of the maximum and minimum flexural rigidity in the zones does not exceed 8 are considered, since in practice the ratio greater than 8, as a rule, is not applied. Analytical expressions for determining the critical force are obtained. The calculation results and their verification are given.

The results of analytical calculations were compared with the results obtained by the finite element method (MSC.Nastran / MSC.Patran). Based on a comparative analysis, graphs of the error of analytical solutions (relative to the solution obtained by the finite element method) were constructed. To minimize the error of analytical equations, a correction factor was introduced.

The study showed that the equations applicable for calculating the critical force of isotropic beams are also applicable to composite beams. Correction factors obtained for isotropic beams are also applicable to composite beams.

In addition to assessing the accuracy of analytical equations for the critical force, the influence of local effects in the area of the junction of zones with different flexural rigidity is investigated. In practice, the Bernoulli hypothesis does not work in the junction area of the zones, which has some influence on the magnitude of the critical force.

Results of investigation:

- Analytical equations were obtained for determining the critical force for two types of beams of variable stiffness with two types of boundary conditions;

- The accuracy of analytical equations was investigated. A correction factor was introduced, which allows to obtain a more accurate result for the critical force;

- The technique can be applied to other types of beams of variable stiffness and other boundary conditions not considered in this paper;

- The resulting analytical expressions are easy to automate. For this suit, for example, Microsoft Excel can be used.

Baklanov A. V. The impact of the of fuel supplying method to the combustion chamber on carbon oxides formation in combustion products of the gas turbine engine. Aerospace MAI Journal, 2019, vol. 26, no 1, pp. 111-125.

The fuel burning in the combustion chamber of a gas turbine engine (GTE) is attended by toxic substances formation. Carbon oxides, having deleterious effect on human and environment, are of particular danger. In this regard, the article solves the actual problem of determining the optimal method of gaseous fuel supplying to the GTE combustion chamber to ensure low emission of carbon oxide.

The article considers the burner with two types of injectors, differing by the gas spray method. The first injector is a centrifugal gas injector (CGI), and the second one is a jet injector (JI).

A technique of target feeding of a jet, formed by the injector in the burner unit was developed.

The fire tests of nozzles were performed. While the tests performing, it was revealed that during the burner operation with the fuel feeding by the CGI, the flame front was being stabilized along the walls of the burner nozzle extension with visible hollow red colored core. Behind the main flame, the reddish “tail” which length corresponded to the length of the main flame was observed. This indicates that the fuel has no time to burn out in the primary zone, and flame front is stretching out.

In this regard, the quality determination of air-fuel mixture preparation in the swirled jet at the outlet of burners with two types of nozzles was performed. It was established, that the nozzle with the jet-like fuel atomization ensured the best mixing quality. The engine throttle characteristics were determined, and carbon oxides concentration in the combustion products measuring was performed by the results of the experiments. The results demonstrated that with the power increase the carbon oxide concentration level in the combustion products decreases. The 25% from the initial variant decrease in concentration was observed herewith for the combustion chamber with JI, which corresponds to the 28775-90 State Standard.

Khramin R. V., Slobodskoi D. A., Lebedev M. V., Sobul’ A. V. The test bench development improvement of the gas turbine engine due to the application of the new method for axial force determining impact on the radial-thrust bearing. Aerospace MAI Journal, 2019, vol. 26, no 1, pp. 126-133.

A radial-thrust bearing of rotor supports is one of the most critical elements of aviation gas-turbine engine, as its failure leads to the engine destruction. To ensure the required reliability of such bearings, experimental studies allowing increase the accuracy of design models employed for the bearing life determination under engine operating conditions were performed. One of the main factors affecting the endurance of radial-thrust bearings is the axial load.

The current quantitative method of axial load determination during engine tests employs the technological supports with dynamometric rings. Qualitative methods of axial load determination based on vibration sensor readings do not allow correct determining of the axial load.

This article presents the method used to measure the axial force applied to radial-thrust bearing. The method is based on dynamic strain gauging of bearing rings. Strain gauges are installed into special slots in the bearing rings. The slot width should be maximum possible but not exceeding the distance between the adjacent rolling elements. The slot depth should comply with the requirements for admissible deformation of raceways and sensitivity of the strain gauges.

The strain gauges readings are taken in the values of relative strain (mm/mm). For ease of use, these values are converted into stresses values (kgf/mm2) by multiplying them by the elasticity modulus of the bearing ring material.

To determine the dependency of the strain gauge readings on the axial load, calibration on a special installation is performed. During calibration, the strain gauges measure the variable stresses in the slot. The amplitude of variable stresses with flicker frequency of the rolling elements is proportional to the axial load, and is a key parameter. To determine it, the signal from the strain gauge, at any given moment, is represented as a Fourier series, and spectrum of the signal amplitude-frequency response is formed. This spectrum is being used to determine the amplitude on flicker frequency of the rolling element. Based on the test results, the calibration factor is determined which characterizes the dependency of axial load on the amplitude of the strain gauge reading signal. Then, by the measured dynamic stresses recalculation, the axial load applied to the bearing is determined.

The accuracy of axial load measurement by dynamic strain gauging of bearing rings does not exceed ±1% of the reference load. The above­described method has been applied during engine tests together with the current method with temporary supports and dynamometric rings.

Based on the test results, the accuracy of axial load determination has been increased and the number of the required engine tests has been reduced.

Gogaev G. P., Nemtsev D. V. The study of flight conditions impact on high-pressure turbine disk damaging of the highly maneuverable aircraft. Aerospace MAI Journal, 2019, vol. 26, no 1, pp. 134-142.

The increase in the GTE life cycle cost brings to the forefront the problem of the full safe use of the aviation engines lifetime, which can be achieved by the transition to operation on a technical condition. This transition is possible with the sufficient product testability ensuring obtaining the objective information required for the reliable technical condition estimating.

The crucial problem herewith consists in methods and algorithms developing for estimation the lifetime depletion, accounting for loading specifics of each engine.

Excessive conservatism is inherent to the currently employed methods for lifetime cycle depreciation control due to the lack of actual operation conditions record keeping. Premature engines exclusion fr om operation occurs thereby, which is unfavorable and has an adverse effect on supporting the required combat readiness level of the aircraft fleet.

Thus, the trend of control techniques improvement, analysis of loading and GTE lifetime deprecation control, fully accounting for the operation specifics of each engine is relevant enough.

The purpose of this work consists in studying the impact of flight conditions on the high-pressure turbine (HPT) disc damaging of highly maneuverable aircrafts.

The main contribution to the parts damage accumulation of the highly maneuverable aircraft engine is made by the damages, caused by intermittent operation modes (the low-cyclic fatigue mechanism), and operation at the maximum set modes (the mechanism of long-term strength depletion).

As the service experience of the 4th generation engines being a part of highly maneuverable aircraft of the task aircraft fleet shows, the contribution of a static component to the overall damage of the basic engine parts is significantly less than the cyclic one. Thus, the estimation of the residual engine life is made, as a rule, based only on accounting for the cyclic damages of its basic parts.

The main idea of the 4th generation engine life deprecation accounting for consists in comparing the actual value of the technical condition parameter (the accumulated damage) of the engine basic parts during the operation with its maximum permissible value, accumulated while the endurance tests, with subsequent determination of the residual resource of the engine basic parts according to this comparison.

Currently the number of cycles before the failure (Npi) and the single damage (Пi) for each cycle type are is detemined at the extreme loads (engine power rating, speed, and flight altitude) for the given GTE operation range.

However, the performed analysis of the highly maneuverable aircraft operation belonged to the task aircraft fleet, revealed that about 80% of the operation was performed at subsonic speeds and heights up to 10 km (without participation in combat operations), at which the basic part load was much lower than its maximum value. Thus, the existing methodology application leads to the excessive conservatism of the accumulated damage calculation.

To assess the effect of flight conditions on the single damage of the main parts, a complex of calculations for HPT disk of the 4th generation engine were performed. The obtained results demonstrated that the single damage of all cycle types of the HPT disk significantly depends on the flight conditions. Thus, the single damage of the loading cycles in the zone, wh ere 80% of operation time is performed in default of combat operations participation, is on average 25% below the values at the maximum loads for all cycle types.

In the context of the HPT disk of the 4th generation engine, the article shows that the existing technique for the lifetime deprecation monitoring by low-cycle fatigue of the 4th generation GTE basic parts includes assumptions leading to the accuracy reduction of determining the accumulated damage and the residual life of the engine and its main parts. This, in turn, leads to an early removal of a serviceable engine, and the life cycle cost increasing.

To avoid the excessive conservatism of the currently used technique, it is necessary to accumulate the cyclic damage of the engine basic parts with account for real flight conditions.

Tkachenko A. Y., Filinov E. P. Gas turbine unit efficiency upgrading for gas-turbine locomotive of a new generation. Aerospace MAI Journal, 2019, vol. 26, no 1, pp. 143-151.

Up to now, at least half of the railways are not electrified. Thus, it is necessary to employ heat engines to set a locomotive into motion. Employing a gas turbine unit (GTU) is one of the possible options. The GTU power is transferred to the generator, and electric motors set the locomotive into motion. It is worth mentioning that in the future aircraft engines of the civil aviation with worked-out lifetime, and updated for the railway application may be installed on a gas-turbine locomotive. Such an approach would significantly reduce the transportation cost value and gas-turbine locomotives implementation to the national economy.

This work was performed in several stages:

– Mathematical models verification used while performing design calculations and GTD operational characteristics computing to increase their identity to the mathematical models employed by PJSC Kuznetsov;

– Studying the number of stages of a low-pressure compressor (LP) effect on the of a gas turbine unit performance employed as a part of the gas-turbine locomotive;

– Proposals development on improving the units’ joint operation to reduce the air consumption through the gas turbine unit.

One of the ways to improve the operation efficiency of gas turbines for application as a part of the gas turbine locomotive consists in the air flow reduction through the unit, which would allow reduce the total pressure losses in the suction tract due to more rational operation conditions of the air filters. The possibility of air consumption reduction through the engine in condition of preserving the effective power of the gas turbine unit by eliminating one and two stages of the low pressure compressor will be discussed further.

The following main scientific results were obtained as a result of the study:

  1. Mathematical models verification used while performing design calculations and GTD operational characteristics computing to increase their identity to the mathematical models employed by PJSC Kuznetsov. Comparison of the results of GTU climatic characteristics computing, based on the initial gas generator, with data obtained at the PJSC Kuznetsov allows talking about the identity of mathematical models of thermo-gas-dynamic computation, performed by the PJSC Kuznetsov, and ACTPA mathematical models;

  2. A study of the low-pressure compressor number of stages impact on the operational characteristics of the GTU employed as a part of the gas-turbine locomotive. Based the obtained results, a conclusion can be made on the inexpediency of changing the number of stages of the low-pressure compressor without refinements (changing the joint operation conditions of the GTU units by throughput efficiency correction of nozzles assembly);

  3. Proposals on improving the joint operation conditions of the units to the effect of air consumption reduction through the GTU, and the most rational options of nozzles assembly of the low-pressure turbine and a free turbine were elaborated.

Dukhopel'nikov D. V., Vorob'ev E. V. Technique justification for erosion profile determining of the accelerating electrode of ions gas-discharge source. Aerospace MAI Journal, 2019, vol. 26, no 1, pp. 152-157.

Determining erosion rate of electric rocket engines elements and other gas-discharge devices is the most important stage of their design and testing. The simplest method to determine the surface erosion rate under the ion bombarding may be employing of optically contrasted multilayer coatings pre-applied to the surface under study. The pattern of alternating optically contrasted bands occurs while sputtering these coatings by the non-uniform ion beams. The boundaries between these bands are the lines of equal erosion depth.

The surface slope angle in the erosion zone while a massive material sputtering by a non-uniform ion beam is determined by the equation


where Ma is the atomic mass of the material, ρ is its density, Seff is effective sputtering rate, j is the ion current density, t is the ion beam exposure time, and q is the ion charge.

While selecting a multilayer coating structure computation of separate layers thickness δi is performed on the assumption of the required band width and the surface slope angle in the erosion zone

The layers thickness herewith should be selected so that the bands widths on the image repeatedly exceeded the registration resolution of the equipment employed for the sputtered patterns photo-registration. Thus, to obtain accurate results using the represented technique, the correct surface slopes angles a; determining is required.

At the same time, while sputtering multilayer coatings, different points of the layer, lying in depth, begin sputtering at different time moments, in contrast the massive material. Thus, the necessity occurred to confirm the correctness of application of the expressions, obtained for the massive material, to the layers thicknesses computing of the multilayer coating.

This article is dedicated to the analytical proof of the expressions usage appropriateness to calculate the erosion slope angle and the layers thickness in the depth of the multilayer coating. It shows that these expressions can be used for any layer of any material located in the depth of the multilayer coating of arbitrary structure.

Volkov S. S. Psychophysiological condition assessment of an operator of the ground complex’s ergative system. Aerospace MAI Journal, 2019, vol. 26, no 1, pp. 158-165.

The article considers an automated system for psychophysiological condition (PPhC) assessment of a flight crew, spacemen, test pilots and other representatives of the airspace industry. The PPhC operation is based on the gas discharge visualization (GDV) method.

The purpose of the work consists in demonstrating the effectiveness and necessity of the psychophysiological state monitoring of the ergative system operators. The ergative system operators are the flight crew of both military and civil aviation; astronauts; test pilots; robotic systems specialists.

This work novelty consists in the GDV method application in a new area. The interest to this method application is caused by the fact that operators are working in special conditions of professional activities. In this regard, they suffer fatigue, overtiredness, undersleeping, performance decrement, stress etc. The PPhC neglecting may lead to tragic aftermath. Thus, the authors suggest developing prospective automated system for operators’ psychophysiological condition estimation, which would allow monitor operators’ readiness to perform their service duties while their professional activities.

During the survey, the snapshots of ten fingers are made with the filter, and another ten without it. The obtained images are being separated into sectors. Further, the mathematical apparatus described in the article is applied to them. The stressed background and normalized glow area, necessary for the psychophysiological state determining, are being computed. After obtaining the information on the operator’s PPhC the official takes a decision on the given person’s readiness to perform his service duties.

The results of the studies allowed developing an algorithm for the software operation of the operatots PPhC estimation system. Neural network technologies are supposed to be the basis of this work. They will improve and expedite the information processing process.

The automated PPhS estimation system, described in the article, introduction into the aerospace industry, will allow monitoring the health of the flight crew, cosmonauts, test pilots and robotic complexes operators, as well as reduce the risk of injury and death while equipment operation.

Anan’ev A. V., Filatov S. V., Petrenko S. P., Rybalko A. G. Experimental approbation of free-falling uncontrolled containers application, employing short-range unmanned aerial vehicles. Aerospace MAI Journal, 2019, vol. 26, no 1, pp. 166-173.

Suppression of enemy’s air defense systems by employing small size striking unmanned aerial vehicles (UAV) to reduce the risk of the piloted aircraft fire damaging is a topical task. The world practice of the small-sized UAV application for striking with free- falling uncontrolled containers (FFUC) is a premise for their application. The majority of scientific publications, describing the UAV striking application, are based, in general, on mass media information, combat effectiveness estimation simplified to its lowest limit, expert esteems of the UAV combat effectiveness without their transformation into qualitative estimations. Highly in-depth academic studies are known also. However, they are based on the probabilistic apparatus, which application is impossible due to the lack of probabilistic laws and random values parameters required for the calculations.

Thus, by this time, the full-fledged computational ballistic algorithms for the small size striking UAVs cannot be realized in practice. With account for the above said, practical approbation of the UAV striking application as the most crucial stage of the aviation complexes lifetime is of first and foremost interest.

Thus, the first and most valid method for the UAV striking capabilities estimation is performing experimental ballistic tests. Their results can be employed for such UAVs efficiency estimation in striking variant, and forming tabulated data on FFUC hitting accuracy, parameters spread, according to which firing tables will be composed.

To reach the purpose set in this work, the following problems were defined and solved:

- The target environment was created for refinement of the FFUC practical application employing UAV;

- Estimation of FFUC with UAV application in striking embodiment on the land objects with application of the simplified deflection measuring technique and estimates of arguments of the FFUC dispersion was performed;

- Statistical data on experimental UAV application in striking mode while hitting ground objects and the enemy’s manpower, for subsequent determination s of FFUC dispersion were collected and processed.

A target, simulating the command center of the medium range surface-to-air missile system battery was employed while testing.

Systematized data on the FFUC dropping were obtained according to the results of the work. They can be utilized for mathematical support developing for the command post of the short range UAVs in striking configuration while developing aiming algorithms.

The obtained results confirm the hitting effectiveness of the FFUC equipped with ammunition of “tactical grenades” type of the enemy manpower and vulnerable (light armored) ground objects.

By results of the obtained statistical data and preliminary calculations, the accuracy of the FFUC application was from 8 m to 10 m.

Legaev V. P., Generalov L. K., Galkovskii O. A. An analytical review of existing hypotheses about the physics of friction. Aerospace MAI Journal, 2019, vol. 26, no 1, pp. 174-181.

Assigned the task to determine the laws of change in the coefficient of friction and determine the factors affecting it as part of the research work. The purpose of this work is to improve the performance parameters of precision machines.

Physics of the external friction process has the next form. When the contacting solids are shifted, the external friction force increases due to deformation of these solids, this phenomenon is called preliminary displacement. Static friction force Fs is the force of friction, corresponds the highest value of the preliminary displacement. One of the contacting solids moves irreversibly (slides) across the surface of another solid after a static friction force has been achieved, in

Relation of external friction force F to the movement x this case the external force is equal to the kinetic friction force Fk [2].

Friction interaction occurs in certain parts of the nominal contact only. Friction interaction is the third solid. The complete complex of frictional links forms a frictional interaction, which is discrete [1].

The preliminary displacement is caused by the redistribution of the contact irregularities in the support surface [1].

The total static friction force is the boundary point, under which preliminary displacement friction passes to kinetic friction force.

Kinetic friction is the friction of two solids that are in motion relative to each other [4].

Kinetic friction has a dual molecular-mechanical nature. It is caused by volumetric deformation of the material and overcoming intermolecular bonds

where Ffm - is the molecular component of the friction force; Ffd - is the mechanical (deformation) component of the friction force [2].

If the adhesion bond is less strong than the underlying layers, then there is a positive gradient of mechanical properties at depth:

Under normal friction process, the positive gradient rule is always come true.

The contact is always discrete and the area external friction depends on the applied load at external friction. Contact surface is continuous and independent on the applied load at internal friction.

The coefficient of friction depends on three factors almost equally: combination of materials; construction of friction pair; operating mode [1].

To execute the rule of positive gradient must be present lubrication film in the friction contact, or oxide film, soft components film [1].

The growth of the film slows down with increasing its thickness [1].

The growth of the film reduces the coefficient of friction to a known limit. Very thick films increase the coefficient of friction [1].

The relative sliding of two solids produces heat in a thin surface layer. The temperature rising can lead to local softening and melting of the material. The temperature field leads to a change in the mechanical properties of the material in a thin surface layer. The intensity of the heat flow depends on the friction work and the size of the area on which it is generated [1].

Important constructive characteristics of the friction units is the coefficient of mutual overlap, proposed by A.V. Chichinadze,

where Aa1 - the nominal friction area of the first element; Aa2 - the nominal friction area of the second element; Aa2 ≤Aa1 .

Wear products have a great impact on the strength and coefficient of friction [2].

Friction and wear characteristics and mechanical properties of friction pairs materials are in various nonlinear functional dependencies. At the same time, these dependencies can significantly change from the friction mode and from the thermal mode of friction pairs.

The construction of the friction unit significantly affects the force and coefficient of friction. In this regard, the nominal Aa, contour Ac and actual areas of friction Ar, the coefficient of mutual overlap Kov, the shape and size of the contact elements, their stiffness and elasticity is among the main parameters determining friction.

More rigid elements of surfaces intrusion into softer counterbody due to waviness, roughness, heterogeneity of mechanical properties and duality of molecular-mechanical nature of friction.

Accordingly, the speed ν of the indenter determines the friction force. At the same time, an increase in the load on the separately selected indenter leads to an increase in the friction force. However, the support reaction force N affects the area of the actual contact Ar in the actual operating conditions of the friction pair. The actual contact area depends on the load. Increasing the area of actual contact reduces specific pressure. Thus, the dependence of the friction force on the relative velocity of the friction pair and the load is not linear and differs for different materials.


  1. In the research of the friction of polymeric and metallic materials should be used adhesion- deformation theory of friction, which includes the definition of the molecular and mechanical components of the friction forces.

  2. The thermal and mechanical properties of materials should be determined by the known friction force of the friction pair.

  3. A positive gradient rule should be observed and lubrication films, oxide films or films of a soft component in the friction contact should be provided.

  4. It is necessary to determine the area of the contacting surfaces at the given micronutrients and friction forces.

  5. It is necessary to take into account the shape and size of the friction unit and the coefficient of mutual overlap.

Mudrov A. P., Faizov M. R. A spherical simulator motion study. Aerospace MAI Journal, 2019, vol. 26, no 1, pp. 182-191.

The article presents a spherical mechanism allowing perform spatial movements along a sphere. A 3D model of the mechanism was developed with the SolidWorks software. The model allows synthesize and examine the mechanism structure. Computing of the angular displacement, speed and acceleration of the connecting rod from the center point of the link and the slide of the mechanism itself was performed. The center point calculation was performed with account for the small and large thickness of the mechanism links Calculations were made for all angles between the links, which were employed for calculation of the spherical mechanism with two degrees of freedom. Based of the obtained mathematical model, computing of the moment of inertia from a given crank motion was performed. The motion parameters along the coordinate axes were determined, which would allow application of the direction cosines formulas. Additional angles calculation used when creating a mathematical model for the moment of inertia were obtained from the spatial sphere around the mechanism. The instant rotation moment of the mechanism was obtained. Using to the obtained data, a certain movement of the mechanism and the time interval of its movement were set, which are reflected by the obtained plots. These plots were obtained for comparing the two methods. The obtained plots reflect the movement of the connecting rod itself, and the slide mechanism. In addition, using Maple, the computation of motion with the moment of inertia of the mechanism itself, with a specified various masses, but with certain geometrical parameters of the mechanism links, was verified.

Rebrov S. G., Yanchur S. V., Drondin A. V., Zernov O. D. Developing the concept of solar energy units robotic assembly in orbit. Aerospace MAI Journal, 2019, vol. 26, no 1, pp. 201-211.

The recent foreign experience in the spacecraft development, including lengthy sup-porting structures developing, and giant space telescopes and solar energy systems con-struction indicates that further development in this field of engineering is impossible without transferring the structures manufacturing directly into the space.

As applied to power systems, this is motivated by the low packing efficiency (measured in percent of occupied volume) of the power system elements inside the launch vehicle, and, correspondingly, the small value of the “Stowed Volumetric Power”, or “Stowed Volume Power Density”, or “Stowed Volume Efficiency” (measured in kW/m3) parameter of the head fairing. This practically excludes the other ways of increasing the power systems total power, other than forming by independent delivery of the power systems parts by way of several launches. The latter leads to a multiple increase of the projects costs, which is not often acceptable.

The article proposes a solution to the describedabove said problem in the form of a concept of a robotized assembly of solar power arrays in space, which is based on the application of the Solar Arrays Assembly Machine (SAAM).

SAAM is a robot with which a solar cell of a large area is being assembled by attaching the mounting panels to each other. The mounting panel can be a honeycomb of high stiffness, allowing the SAAM to move along its plane. When moving, the SAAM “clings” to the reference holes made on the mounting plates in advance. SAAM has four telescopic supporting rods for moving around the mounting plates and two mounting arms for fixing the panels.

The concept demonstrates the scheme of the SAAM application. determines The optimal route for the SAAM movement and the order of the solar array assembly are determined. The scope of its possible application has been determined: for assembling a wing of a solar array with an area of less than 64 m2, the target (competitive) mass of the SAAM is of linear dependence on the area of the solar array. When assembling solar arrays with an area of more than 64 m2, traditional deployment systems cannot be employed. So the SAAM does not have competitive alternatives implemented.

The basic SAAM size are determined. A layout was made allowing develop the basic technological operations and algorithms of moving and assembling. The system weight and size parameters depend on the materials used, electromechanical assemblies, SAAM batteries, and will be refined further further work. The time of the solar array forming depends on the speed of SAAM electromechanical units and manipulators operation. But this is not a limiting factor, since the modular structure of the system should allow the SAAM to recharge from the assembled segments at several stages of the assembly.

Kovalev A. A., Konovalov D. P. Workpiece thermal deformations simulaiton occurring while holes drilling process. Aerospace MAI Journal, 2019, vol. 26, no 1, pp. 201-211.

The article tackles the issue of determining the error caused by the workpiece thermal deformations occurring in the holes drilling process in a part, being the main part of the unbraked wheel of the aircraft landing gear and is called the “Drum”.

The article describes the mechanism of these errors occurrence. A method for the treatment process simulation was developed, and proposed an algorithm for estimating the error in the workpiece size occurred due the thermal deformation while drilling.

The article consists of three main parts, namely introduction, body part and conclusions.

The introduction considers the mechanism of errors occurrence due to thermal deformations of the workpiece, which in turn presents one of the total machining error components. It presents the cases when this error may significantly affect the total machining error. Thus, it is relevant that this error component is estimated.

The basic part presents a method for computing the temperature in the cutting zone for further machining process simulation. It describes the object of simulation, i.e. the operation of drilling a through hole of 13.5 mm diameter with the tolerance range of 120 μm in a workpiece from the ML12 magnesium alloy with the cutting modes recommended by the cutting tools manufacturer, namely, the cutting speed of 264 m/min and feeding of 0.35mm/rev. An algorithm for the size error estimation is presented as a block diagram. The step-by-step description of the hole drilling simulation process is presented on the example of this operation. As a result, the temperature distribution, equivalent von Mises stresses, and displacements caused by thermal deformations over the part volume were obtained. Based on the diagram of displacements, caused by thermal deformations, the error was 191 μm at the specified cutting modes and machining conditions, which appeared greater than the tolerance range by the size of the hole.

The conclusions note that the cutting parameters recommended by the cutting tool manufacturer do not always provide the required machining accuracy. It was concluded that the required accuracy was not achieved for a specific hole drilling operation. The ways leading to the error reduction due to changes in cutting parameters, as well application of the other types of cutting fluid are presented in the conclusion.

Verchenko A. V., Kurskaya I. A., Chigrinets E. G., Maksimov D. V., Geiko Y. S. Water-jet cutting process optimization of work pieces from aircraft materials. Aerospace MAI Journal, 2019, vol. 26, no 1, pp. 212-229.

Single and small batch production is predominant in parts production for aerospace industry. Stamped blanks and castings manufacturing with small batches is not cost-effective due to the high cost of tooling. Thus, forgings or billet plates made of thick plates that are close to the part’s shape are used in increasing frequency as work pieces.

One of the most up-to-date and promising method of cutting and obtaining finished parts is the method of water-jet cutting. It ensures wide ranges of processed material thickness, the ability to cut almost any material, high performance, obtaining high quality cutting surface, the ability to process complex geometry. All this makes this method of processing the most popular in conditions of modern aircraft building, shipbuilding, etc. The absence of thermal impacts on the material, low cutting force, the erosional nature of the destruction do not contribute to the development of internal stresses in the cut zone.

The process of water-jet cutting is complex, poorly understood, which result is affected by many technological parameters such as cutting pressure, nozzle feed, grain, hardness, abrasive consumption, distance from nozzle to the surface being processed, physical and mechanical characteristics of the material being processed. The design complexity of the cutting technological process consists in selection of optimal cutting conditions, which will ensure the specified quality of the part surface layer at the minimum cost. The production technologist faces the difficulty of determining not only the cutting surface hardness, but the size of the smooth and wavy cut zone as well.

goal of the work consists in improving the efficiency of the waterjet cutting process by optimizing processing modes based on the development of an adequate theoretical model for the formation of surface roughness at different depths of the cut section.

To achieve this goal the following tasks were solved:

  1. Theoretical and experimental studies of the cut surface roughness profile formation depending on the processing parameters;

  2. Theoretical studies of the wavy cut zone formation depending on the technological parameters of the process;

  3. Development of methods for predicting the quality of the cut surface;

  4. Development of methods for optimizing the process of water-jet cutting.

The paper presents the results of theoretical and experimental studies of the surface roughness profile formation while water jet cutting of various materials, such as 30HGSA steel, hardened 30HGSA steel, D16T aluminum alloy, fiberglass-titanium composite material. A theoretical model for the roughness profile formation of the cut surface was obtained, which shows the dependence of roughness on the main technological parameters of the process (nozzle feed, particle radius, mixture pressure, etc.) depending on the depth measurement of the cut surface roughness. It reflects thereby the distribution of the ratio of the smooth and wavy cut zone. Statistical processing of the studies results was performed using MathCad. The experimental studies result was the obtained dependencies of the number of the particles’ useful encounters with a material on the magnitude of the nozzle feed, abrasive consumption, and section depth. One and two-factor regression equations describe the effect of abrasive consumption, nozzle feed, thickness of the material being processed, section depth on the cut surface roughness.

A two-factor regression model for the formation of the roughness profile of the cut from the nozzle feed rate and the roughness measurement depth while polymeric composite materials (PCM) processing of the fiberglass-titanium type was obtained. The material layering and shagging while cutting were not detected, the cut quality was high. To assess the water impact while cutting fiberglass-based PCM, an analysis was performed using differential scanning calorimetry, which resulted in the conclusion that the waterjet cutting technology can be used for PCM processing.

Based on the theoretical and experimental studies results, a for designing and optimizing the technological processes of water-jet cutting technique has been developed, with account for the specified cut surface roughness ensuring and obtaining the minimum cutting costs.

The optimization of the technological process of water-jet cutting of the “Bracket” part of the Mi-28 helicopter was performed, which resulted in a 2.5 times reduction in labor intensity, a cost cut of 843.51 rubles, which allowed the company to save 1286 rubles while each part production. The technique for the water-jet cutting technological process optimization application was undergone industrial tests at the Rostverol plant.

According to the technical requirements for rotor blade manufacturing, as well as the results obtained by the authors, the possibility of hydro-abrasive cutting application for removing the technological allowance in the basis part of Mi-28 helicopter main rotor spar as an alternative to the rough milling was demonstrated. Application of cutting feed within the 160-240 mm / min range min reduces the labor intensity by 80% with the required quality indicators.

At present, measures for the suggested technology introduction into batch production are under development at the PJSC Rosvertol Blade plant.

Zhemerdeev O. V., Kondratenko A. N. Methods for determining technical potential state of the enterprises based on a modified model of factors of production. Aerospace MAI Journal, 2019, vol. 26, no 1, pp. 230-235.

The core indicators, characterizing the enterprise technological capacity are technical level of production (TL), identified by the technical level of the leading elements of the fixed productive assets (FPA), and wear-out (W). The existing equipment classification is expanded with account for clean zones and premises (CZ&P). Each FPA element group in the classification is associated with the TL (l), adopting values fr om 1 to 7. Accounting for CZ&P is especially relevant while determining the TL of an instrument making enterprises, production of electronic components and optical elements, as well as some assembling industries of machine building.

Basic coefficient of the technical level of production at the enterprise is defined as a weighted average value (l). Weighting factors calculation is performed employing gross book (replacement) value of the group elements adduced to the prices of the current year, using deflator indexes of the Ministry of Economic Development “Fixed Investments”. The calculating formula is based on the effect of labor efforts decrease with the technical level (l) growth, and weighting factor considers the accomplished capital investments has been made. The TL coefficient for particular production method (technological lim it) is defined similarly.

The transitive coefficient of production TL is an extra tool for monitoring and prediction of the technical level of production. Its calculation is similar to the of the calculation of the basic coefficient technical level of production. With this, while weighting coefficients computing, besides the gross book (replacement) value of the group elements the real wear-out of FPA elements is considered. To determine the real wear-put of the elements it is most preferable to employ the probability models approach based on lognormal distribution. The TL transitive coefficient presents interest for the basic TL coefficient trend forming due to the FPA elements disposal. Actual wear-out (W) is determined as a weighted average value of actual wear of FPA elements.

The developed method is based on an accessible input data, and the proposed variables of technological capacity are “tied” with the capital investments.

Nochovnaya N. A., Nikitin Y. Y., Savushkin A. N. Exploring the properties changes of the titanium alloy blades surface after chemical cleaning from carbonaceous impurities. Aerospace MAI Journal, 2019, vol. 26, no 1, pp. 236-243.

It is important to understand how a cleaning technology can change the physico-chemical properties of the material being cleaned after removing carbonaceous impurities from the compressor parts surface of a gas turbine engine. In continuation of the previous work, the creep of VT20 titanium alloy samples was examined, and one of the selected chemical technologies that remove carbonaceous impurities was tested on compressor blades with subsequent determination of some surface properties.

To evaluate the creep of VT20 titanium alloy characteristics, the standard flat samples, some of which were coated with carbonaceous impurities that simulate exploitation, were fabricated. Two titanium compressor blades of a gas turbine engine were used in the research work: blade 1 (small) after operation with a small amount of contaminants on the surface, and blade 2 (large), on which carbonaceous impurities, imitating operation, were coated.

The creep tests results proved that the impurities removal by cleaning solution No. 1, alkaline and acid solutions (“loosening + etching”), and HDL 202 did not reduce the time of the samples destruction and degraded their plasticity, compared to the original samples.

Allowing for the results of the previous work, 9. cleaning solution No. 1 was selected for testing the of carbonaceous impurities removal from the surface of the blades. The results of blades processing revealed 10 that the surface was completely cleaned. In in X-ray microanalysis spectrograms the elements such as sulfur, oxygen and carbon, indicative of the presence of carbonaceous impurities, are missing. The values of surface roughness and micro-hardness did not sustain significant changes. Processing in the indicated solution leads to activity (potential) increase of the of blade No. 1 surface. The lower values of the blade No. 2 surface potential were observed (about 10%) compared to the initial state.

Kolesnikov A. V., Mikhailov I. V. Superplastic forming of aerospace facilities’ parts and multilayer structures from vt20 titanium alloy. Aerospace MAI Journal, 2019, vol. 26, no 1, pp. 244-250.

The structures from titanium allows are increasingly employed in aerospace structures. Labor intensity may be significantly reduced while titanium parts manufacturing by application of superplastic forming process (SPF) and combined process of SPF and diffusion welding (SPF/DW). Superplasticity manifests itself in alloys with a fine-grained structure under certain strain-rate conditions and maintaining a constant temperature during the formation process. Maintaining a constant strain rate in the process of shaping is ensured by a continuous change in the forming pressure over time. The computing of the plot of the forming pressure change with time is rather labor consuming. For this problem solving and the process visualization, modeling with the MSC “Marc” program was performed.

By the example of forming a cellular panel from VT20 titanium alloy, the possibility of manufacturing parts by the SPF method is demonstrated. The simulation result allowed obtain relative deformations distribution, which analysis revealed that maximum relative deformations constituted 97%. This is quite acceptable, and there will be no destruction while the forming process. The simulation results allowed also develop the control program according to which the cellular panel was produced by the superplastic forming press.

The article considers the form shaping modeling of multilayer wedge-shaped panels with transverse and longitudinal corrugation set. It follows from the relative deformation distribution analysis that maximum relative deformations in the structure constituted 126.6%, which is acceptable. The forming of the wedge-shape three-layer panels was performed by the SPF/DW method according to the computed plot of forming pressure change with time.

After the superplastic forming process, there are no both corrugation forming and springing effect, which eliminates the finishing work.

Thus, the SPF and SPF/DW technologies and modeling the process of production to obtain the forming parameters allow significantly enhance the production possibilities while producing complex parts from titanium alloys.

Klimov V. G., Nikitin V. I., Nikitin K. V., Zhatkin S. S., Kogteva A. V. Wear-resistant composites application in repair and modification technology of the gtd rotor blades. Aerospace MAI Journal, 2019, vol. 26, no 1, pp. 251-266.

The of production and operation costs of gas turbine engines employed in aviation, oil and gas or energy industries constitutes a significant portion of costs reducing the net marginal profit of operator- enterprises. These costs reduction is a natural desire of any holding. Against this background, the ability to maintain the resource of the gas turbine engine at the lowest cost to itself remains the main criterion of the competitiveness of the producer in the market.

It should be kept in mind, that the operation costs of gas turbine engines through their lifetime cycle often exceed their original cost. To be precise, the effective repair technologies often stops the loss risks in future orders.

A distinctive feature of domestic aviation gas turbine building is a low assigned and overhaul period of the engines operated according to the first performance strategy.

Often the causes of understated life cycles are the imperfections of the structures that occur at the development stage. Consequently, the presence of the extremely expensive parts and units with a relatively short lifetime requires their permanent replacement or recovery. These parts are the rotor blades, and the turbine stator. Many factors can lead to their failure, starting from structure changing due to the uneven temperature fields, to the loss of geometry due to burn­out or mechanical damage. The last one is the factor, most frequently occurred in the products.

From the viewpoint of repairing technologies, the turbine blades recovery is the most cost-effective among all other engine parts. The cost of the engine hot section (turbine) producing exceeds the cost of a cold section (compressor) by average of 400-700%. However, the repair complexity remains the main obstacle in its implementation.

This article proposes to employ the high-temperature nickel powders of the VPR type as wear-resistant surfacing materials applied by laser action. The structure formation peculiarity of the described materials is revealed. It is manifested at high cooling rates in the form of natural composites formation with dispersion eutectic hardening along the boundary of the dendritic framework. This structure has a non­directional arrangement of strengthening phases that increases the wear-resistant characteristics of the resulting composite.

The original method of restorative surfacing is described. It allows repairing and modifying rotor blades of gas turbine engines (GTE), with increasing the wear-resistant characteristics of the part contact surfaces. Based on the conducted comparative studies, including analysis on a scanning electron microscope; measurement of micro-hardness and the coefficient of the materials linear expansion; testing of abrasion resistance of cladding and their fatigue strength, the possibility of VPR type materials application of as an alternative to classical wear-resistant composites with mechanical admixture of various carbides was proved. It is shown that under conditions of pulsed laser action at high cooling rates, the average hardness and overall resistance to abrasive wear of certain VPr alloys grow due to the formation of a finely dispersed stable eutectic structure close to the initial powder material. The positive performance characteristics of alloys of VPr 11-40N and VPr 27 grades were obtained, which allows employ them when rebuilding the GTE rotor blades.

Kuznetsov E. N., Lunin V. Y., Panyushkin A. V., Chernyshev I. L. Boundaries of non-separation flow-around of bodies of rotation, with the nose part in the form of Riabouchinsky half-cavity. Aerospace MAI Journal, 2018, vol. 25, no 4, pp. 7-15.

Bodies that are optimal at the so-called low critical Mach number M*, at which at least one sonic point on the body flown-over surface occurs, were studied theoretically in Ref. [1]. It was confirmed that M* achieves its maximum value and, consequently, the wave drag minimum value occurred for the body identical to the Riabouchinsky finite cavity in the classical theory of incompressible fluids. It was experimentally studied in Ref. [12], which demonstrated that in the transonic velocities range the Riabouchinsky half- cavity had the smallest drag among the bodies of rotation with the same aspect ratio  λ=L/D=0.87(where L is the nose part length and D is the diameter of its mid-section). This conclusion is incorrect for aspect ratios λ>2 due to the friction impact the drag as it follows fr om Ref. [24]. The absence of turbulent boundary layer separation from the side surface of the body of rotation under study at zero angle of attack in the range of Mach numbers 0.8≤M≤0.95 was demonstrated in Ref [17].

The main objective of this work is determination of angles of attack αsep at which turbulent boundary layer separation from the side surface of the studied body occurs. The study was performed with NUMECA FINE/Open software based on Reynolds Averaged Navier-Stokes equations (RANS). The solution of the problem was performed in the framework of fully turbulent flow model without accounting for laminar-turbulent transition using Spalart-Allmaras (SA) and k-ω SST turbulence models. To determine the boundaries of the non-separated flow-around computation was performed in stationary problem setting at various angles of attack. With that, the flow separation indicator was the presence of the zone on the model surface wh ere the friction coefficient Cf < 0. The results obtained with two turbulence models are close to each other, and the difference between the two separation angles does not exceed 1°.

The results of the study obtained for αsep for the nose part with aspect ratio of are as follows:

αsep=15° for М=0.5, αsep=9° for М=0.65,

αsep=12° for М=0.8, αsep=13° for М=0.85,

αsep=5° for М=0.9, αsep=11° for М=0.95.

Computing results for the longer nose part with aspect ratio are:

αsep=20° for М=0.5, αsep=13° for М=0.7, αsep=21° for М=0.9, αsep=18° for М=0.95.

The angles of attack αsep which realize turbulent boundary layer separation from the side surface of the investigated body at free-stream Mach numbers 0.5≤M≤0.95 were obtained. Separation zones location is shown for various models and modes.

Bragin N. N., Kovalev V. E., Skomorokhov S. I., Slitinskaya A. Y. On evaluation of buffeting of a swept wing with high aspect ratio at transonic speeds. Aerospace MAI Journal, 2018, vol. 25, no 4, pp. 16-27.

The article presents to the development of a technique for buffet initialization boundary evaluation, occurring on a swept height aspect ratio wing at increasing angle of attack during cruising flight modes. The lift coefficient value of the buffet onset is one of the limitations that should be accounted for while designing the win aerodynamic layout of a subsonic aircraft. According to the norms, the margin between the cruise flight mode and the CLbuff value or the lift coefficient of the buffet onset should be at least 30%. Thus, knowing the CLbuff value through the entire operational speed range is a prerequisite for an aerodynamic wing configuration design beginning from the preliminary design stage. The problem of determining the CLbuff has become of special urgency at the transonic speeds due to the substantial aspect ratio increase of the (by 15–20%) of the long-range aircraft wings due to the composite materials application in load-carrying structure.

Despite the successes in CFD aerodynamics gained over the last years, non-stationary separation flow modes are studied, basically, using experimental tools, including wind tunnel tests of airplanes high scale models. Though the cost of such studies is high, they ensure required reliability of the obtained results. It is worth mentioning, that the time consuming computations on multiprocessor computers are costly as well. With this, the high accuracy and reliability of the obtained results are not guaranteed. Preparing mathematical model and building-up computational meshes with hundreds of millions of nodes are commensurable with costs of developing and manufacturing scaled models for tests in the wind tunnel. Thus, numerical methods do not always prove to be less labor consuming and costly than the experimental ones. Despite the fact that computer facilities and software develop rapidly, and the situation gradually changes, experimental methods remain as before the basic tool while performing the studies of complex flows.

The article presents the analysis of typical features of the wing flow-around at the angles of attack corresponding to the start of the buffet mode. The technology of application of the program for computing transonic flow-around based on the full potential method for the buffet initialization computing is demonstrated. Computational results comparison with the data of experimental studies obtained for the model of the wing with fuselage in the wind tunnel is presented.

Danilenko N. V., Kirenchev A. G. Vortex formation of gravity flows. Aerospace MAI Journal, 2018, vol. 25, no 4, pp. 28-36.

Vortex formation analysis of the air medium as a gas turbine engine (GTE) propellant allows extracting one of its specifics, namely the gravity character of the technogenic vortices formation. The fudamentals of this vortex formation are being subordinated to the natural vortexes formation of the Earth atmospheric environment. The theory of technogenic vortices of the GTE operating on the ground, the same as the theory of natural atmospheric vortexes is being in the state of its development. It confirms the state of the issues of the working process principle and classification of both technogenic and natural vortices stated in scientific and technical literature and textbooks on the theory of gas turbine engines and metrology. The most informative is the database on meteorological studies of natural vortices (tornados, cyclones, circulations and atmospheric fronts). Thus, due to the technogenic and natural gravity vortex forming similarity, the gist of technogenic vortices' work process should be searched for in the gist of the cyclonic type vortices of the environment. The work process study herewith of the cyclonic type vortices (tornados and others) may be the basis for creating a theory of natural vortex forming.

The problem is set to study the work process of small-sized technogenic vortices with their subsequent adaptation to the work process of natural vortex forming.

The above said problem should be solved relying on the basic equations of gas dynamics (gas flow energy preserving and other) with subsequent yield to the methodology and essence of the gravity type vortices' work process. The most accessible to learning the gravity vortex formation and its vorticity is the energy conservation equation, including its components in the form of internal and kinetic energy of a gas flow, kinetic energy of the environment angular rotation, and heat exchange elements in the form of external mechanic work and heat. Hence, extracting the master unit of vortex formation (angular rotation energy) allows establish functional dependence of the vortex formation under study from the sources of energy capable of generating gravity vortices of various types.

The article presents the methodology of studying, and analysis of the problems of work process cognition of vortex movement of the Earth's ambiences. Classification of vortices according to the gist of their work process. The article indicates the way of splitting the vortex formation into the vortices according to the gist of the work process included into its classification, and further cognition of their physical entity and exploration resulting in vortex characteristics, their consequences and application areas.

Yudintsev V. V. Rotating space debris objects net capture dynamics. Aerospace MAI Journal, 2018, vol. 25, no 4, pp. 37-48.

By now, several methods for near Earth orbits active cleaning from large-size space debris were suggested. The most difficult stage of such mission is the stage of space debris capture. Capturing method selection and subsequent orbital transportation of space debris depends on its type and angular motion. Rockets' orbital stages may rotate with high angular velocity, which aggravates their capture by manipulators and other means. One of prospective techniques of such object capture is application of a net connected with the space tug by a tether. The object capture by a net can be performed by the net separation with a certain relative speed in relation to the space tug and space debris, or by the net unrolling on the trajectory of the space debris object relative to the space tug. Elastic properties of the net and tether allow reduce the load acting on the space tug while an object capturing process and control the value of this impact.

The paper presents discrete mathematical model of the net movement as a system of material points' elastic interaction, as well as these components interaction with the space debris surface. The possibility of capturing an orbiter type object, rotating with significant angular velocity was demonstrated through this model. The article demonstrates that capturing the object, rotating with angular speed of 5 degrees per second, requires the speed of the net relative to the space debris from 2 to 5 m/s. To capture an object, rotating with angular speed of 30 degrees per second, the net speed should be no less than 10 m/s.

Bolsunovskii A. L., Bondarev A. V., Gurevich B. I., Skvortsov E. B., Chanov M. N., Shalashov V. V., Shelekhova A. S. Development and analysis of civil aircraft concepts employing integration principles. Aerospace MAI Journal, 2018, vol. 25, no 4, pp. 49-63.

The search for the technologies allowing significantly improve operational characteristics of the prospective civil aircraft was the goal of the work.

The study of innovative technologies, including those, providing the airframe and engine integration, was performed applying methods of alternatives analysis, based on the factorial analysis, and experiments planning assuming performing a series of computing experiments with subsequent comparison of their results.

Three possible innovations trends were considered:

– application of a turbojet engine with the increased bypass ratio for the fuel consumption and noise-at-terrain reduction;

– application of the so-called distributed power plant with the separated gas generator and the fan connected by mechanical transmission;

– airframe and power plant elements integration for obtaining useful effects in aerodynamics and structure, as well as and obtaining new operational properties.

According to these independent principles a number of the long-haul aircraft possible configurations, differing by various combinations of the bypass ratio, the turbojet schemes and technologies of elements integration into the “airframe + engine” system was developed. The number of possible strategies of the integral aircraft, including the base option, corresponds to the number of binomial coefficient of the three factors 23 = 8 according to performing the full-factorial experiment.

A multidisciplinary expert assessment of aircraft configuration options was performed, which turned into the basis for the most effective concepts selection.

Comparison of possible characteristics revealed that some options of airframe and engine integration under consideration had potential for considerable of fuel consumption reduction compared to the conventional long-haul aircraft configuration. The results of the study allow recommend two strategies for further studies. These strategies also possess potential for additional improvement of the other operational characteristics, such as noise-at-terrain and operating safety. Other configurations possess a number of useful elements that can find application while critical technologies development and reduction of technical risks.

Egoshin S. F. Impact evaluation of multi-propeller wing blow-over system on the stol aircraft characteristics. Aerospace MAI Journal, 2018, vol. 25, no 4, pp. 64-76.

The article undertook the attempt to obtain an analytical solution to compute the take-off length of an aircraft equipped with the multi-propeller wing blow-over installation, and estimate the benefits of such engineering decision through the transport operation evaluation of this aircraft.

The main difficulty of this problem lies in the fact that the wing and propeller interaction is an extremely complex and insufficiently studied task. Currently, only approximate semi-empirical formulas for calculating the aerodynamic characteristics of the wing at small relative diameters of the blowing airscrew exist. The exact calculation of the wing flow-around in this case is possible only for strictly specified, nonparametric configurations using the finite-difference method.

In addition, the current complicated situation in the sphere of local air transport in Russia (reduction of airlines and operating airfields) requires the search and evaluation of effective technical solutions for a prospective aircraft of local airlines. One of the ways is envisaged as developing a short takeoff and landing (STOL) aircraft capable of carrying out transport operations in conditions of an underdeveloped airfield infrastructure. It is considered that equipping such a STOL airplane with a multi-propeller electrically powered blow-over system will be an effective solution to the problem. However, the above said complex aerodynamic task does not allow a quick search for the optimal characteristics of this aircraft.

The developed mathematical model, under certain constraints, allows obtain an evaluative analytical solution for the take-off run length of such STOL aircraft, reveal the specifics of parameters interaction and evaluate possible advantages and disadvantages of the aircraft. Within the framework of the model, it was demonstrated that the maximum possible power consumption from the main engines is the optimal value of the corresponding parameter of the electric power plant. The amount of this power consumption determines the blown part of the wing area through the relationship with the critical rotation speed of the auxiliary propellers.

As for performance of a transport STOL aircraft based on L-410, it was shown that a blow-over system based on conventional electrotechnical materials can reduce the take-off run by 30% (up to 300 m), while reducing the payload by 1520% at flight ranges up to 300 km or up to 50% when flying to the maximum range. At the same time, if the electric power plant is designed based on high-temperature superconductors (HTSC), the payload reduction will be much less: negligible at flight distances up to 300 km or about 25% with flight to the maximum range. This allows conclude that the HTSC technology application for such STOL aircraft creation is rather promising.

Lopatin A. A., Nikolaeva D. V., Gabdullina R. A. Experimental data generalization on heat transfer in cooling system with axial sectional finning in conditions of free convection. Aerospace MAI Journal, 2018, vol. 25, no 4, pp. 77-85.

At present, power electronic components with high heat release have been widely applied in various fields of modern industry. The main problem the developers of the element base are facing consists in creating cooling and thermal stabilization systems capable of removing heat fluxes of high density, while working in a wide range of ambient temperatures. When creating such systems, it is necessary, alongside with the thermal ones, to account for the mass-size characteristics of the device as a whole. Thus, much depends on the heat transfer intensification method selection.

Quite enough attention is paid in modern literature to the issues of radio electronic equipment thermally loaded elements, as evidenced by a significant number of articles and monographs on this topic. Heat release is one of the main causes of the unstable operation of radio electronic equipment. Among the basic factors exerting maximum destructive effect, the increased temperature of the elements is one of the main ones. Thus, the devices operation in the field of the radio electronic equipment is closely associated with heat removal from the thermally loaded components. Depending on the structure and shape of the cooled components, various solutions are employed for organizing continuous heat removal.

Certain problems of large heat fluxes removing in the elements of industrial electronic equipment were considered in [1-3]. The correct choice of the cooling system type ensures trouble-free operation of all cooled components of the device.

A considerable amount of publications in modern scientific publications, both in Russia and abroad, attest to the considerable interest in the issues related to the heat transfer intensification for surfaces of various shapes and sizes as applied to cooling systems for electronic equipment. The issues of heat transfer intensification are set forth in [4-10]. In particular, the criterion equations of various authors for the Nusselt number computing for natural convection are presented in [2, 9]. Experimental studies of the convective heat transfer intensification in rectangular dissected channels and in channels with discrete turbulators were described in [1, 10]. In the studied dissected channels, a process of rational convective heat transfer intensification was implemented, reliably controlled by changing the values of the basic dimensionless geometric parameters. The generalizing dependences for discrete-rough channels were obtained in [7] for free convection conditions, and flow modes and mechanisms of intensification were studied. In [11-14], the authors experimentally studied one of the parameters characterizing the cooling systems both qualitatively and quantitatively, namely, the thermal resistance.

The fins application as a method of heat transfer intensifying leads to the increase in the heat transfer coefficient value by a factor of tens. This method of intensification implies a wide variety of vatious types of fins, such as: longitudinal, transverse, rolling, spiral and many others [15,16]. In [15] the analysis of the expediency of employing different types of fins from the viewpoint of the coolant aggregate state is presented. The optimal edges number selection is presented in [16]. The heat transfer intensification of the systems with a cut-off fin is also considered in [17-20].

The purpose of this work consists in studying the efficiency of the split finning under conditions of natural convection. A test bench was developed for performing the experiment on the study of heat transfer. While the experiments on heat transfer near the cut edge under conditions of natural convection, criterion dependencies were obtained.

Relying on the analysis of literature sources and accounting for the results obtained while experimental studies, the authors established that from the viewpoint of the of axial split finning practical applicability, there are a number of specifics, associated primarily with the fact that the “petals”, obtained as a result of dissection of the heat exchange surface, can be considered as independent ribs. The studies of heat transfer intensification under conditions of natural convection with the cut ribs application were conducted and presented as a result of the work. While the experiment, the effectiveness of the use of split finning is demonstrated, and the most optimal geometric parameters of the working area were revealed. The process of heat transfer was visualized. The boundary layer thickness near the cut edge was computed. Criterion dependencies for heat transfer computing of the systems with axial cut fins were obtained. The prospect of this study is of experimental data verification by numerical modeling programs.

Zichenkov M. C., Ishmuratov F. Z., Kuznetsov A. G. Studying the gyroscopic forces and structural damping joint impact on the wing flutter of the aeroelastic euram model. Aerospace MAI Journal, 2018, vol. 25, no 4, pp. 86-95.

The article deals with the structural damping role analysis while studying the gyroscopic forces impact on the flutter speed. The algorithm for accounting for gyroscopic forces in polynomial Ritz method while computing the aircraft aeroelasticity dynamic characteristics, developed earlier by the authors, was employed. The algorithm was realized in the KC-M software developed in TSAGI and validated while solving aeroelasticity problems in many practical applications.

The computations were performed on the example of the wing of the well-known aeroelastic research model of the four-engine long-distance aircraft EuRAM (European Research Aeroelastic Model), developed and studied in the framework of the European project 3AS (Active Aeroelastic Aircraft Structure). The model characteristic feature is the flutter form occurrence associated with the lateral vibrations of external engines. This form is affected by the gyroscopic forces due to the engines rotating rotors.

The flutter characteristics analysis at various levels of structural damping (characterized by logarithmic vibrations decrement δ ) revealed that the vibrations tones interaction character with account for gyroscopic effect was not principally changed. It was found herewith, that the gyroscopic forces impact on the speed of the considered flutter form might be of different sign depending on the level of the structural damping.

For example, at δ = 0.02 the flutter speed increases by 11.5%, with the maximum value of the engine rotor kinematic momentum (scaled to the model). While increasing the structural damping value, in the beginning, the gyroscopic forces' impact on the flutter speed decreases, it does not practically exist at δ = 0.04, and with further increase of the decrement the impact changes its sign, and the flutter speed decreases. The flutter absence was marked at δ > 0.046 in the range of small rotor rate speed, but while the rate speed increase the flutter may occur. Its speed decreases at that (about 10%) with the rate speed increase. This indicates the importance of accounting for the dynamics of rotor systems while the aircraft aeroelastic phenomena analysis.

The obtained the results were confirmed also by the finite element computing method in NASTRAN system using Rotordynamics module (accounting for the rotor systems dynamics). The computing results on gyroscopic forces impact on aeroelasticity characteristics at various structural damping values performed with KC-M software accord well with computations performed with NASTRAN software.

It was noted that while experimental validation of the gyroscopic impact on the flutter speed of the model in the wind tunnel various results might be obtained depending on the structural damping level. Thus, the detailed computational and experimental analysis of the model dynamic characteristics is required while such tests preparing and running.

Ezrokhi Y. A., Kalenskii S. M., Morzeeva T. A., Khoreva E. A. Analysis of a concept of the distributed power plant with mechanical fans drive while integration with a “flying wing” type flying vehicle. Aerospace MAI Journal, 2018, vol. 25, no 4, pp. 96-109.

The article presents analysis of a concept of the distributed power plant (DDP) while its integration with a “flying wing” type flying vehicle.

A modified airframe model of a prospective long-range aircraft (LRA) of PJSC “Tupolev” development with two power plants integrated into the tail-end was selected as a flying vehicle.

Those power plants represent a bypass turbojet engine where two taken-out fan modules are driven by mechanical transmission from fan turbine of this turbojet. The choice in favor of a mechanical way of power transfer for the aircraft of 2030 level is based on the results earlier performed studies on the engines of new schemes in CIAM named after P.I.Baranov.

The results obtained while numerical modeling of the flow on the upper surface of an LRA airframe were also employed. This modeling revealed that for a long-range flight the mean values of the full pressure's losses prior to the fans differed greatly and depended monotonously on the flow deceleration level in the air intake. According to the calculations, the average value the full pressure restoration coefficients was correspondingly ~0,923 for the first fan module, ~0,952 for the bypass turbojet and ~0,958 for the second fan module.

Refining of the earlier developed model of the distributed power plant was performed to evaluate the impact of the conditions at the inlet of each of fan modules. The performed of mathematical model refinement allowed implementing independent selection of parameters, dimensionality and gear-ratios of reducing gear for DPP fan modules drive, as well as performing independent regulation of output devices of these modules.

The article considers separately the impact of the two main factors on the engine thrust, namely, the fall of the full presure level at the inlet, and its proper heterogeneities.

Calculations revealed that for the earlier selected DPP option while its integration into the flying vehicle under consideration, regulation of nozzles of the turbojet bypass loop and fan modules was required at the cruising mode. With this, gas temperature increase prior to the turbine by ~70 К was required.

Three different variants of the engine which allow excluding the above said regulation were investigated while this work.

The first variant is a version with fans equal by dimensionality and pressure ratio at the designed cruising mode.

The second variant is a version with the first fan module with the pressure ratio increased by 5% relative to the BTJEs fan at the cruising mode.

The third variant is a version with first fan module air consumption decreased by 50% at the cruising mode.

Parametric studies performed employing the develop methodology allowed selecting the degree of bypassing and the degree of pressure increase in the fan optimal by the specific fuel consumption at the cruising mode for each DPP option. The dimensionality of fan modules and main DPP units was refined with account for various losses levels at the inlet.

Analysis of effects associated with the presence of a non-uniform field of the full pressure and leading to its average level decrease at the fan inlet revealed that impact of the presence of non-uniformity might be from 15 to 30% of total impact on the engine thrust.

At the same time, while the power plant parameters selection at the cruising mode with account for the degraded coefficients of the full pressure preservation prior to the fan, the fall of the thrust level due to the proper non-uniformity might be ~2,53% at the given mode. This should be accounted for while selecting an optimal DPP appearance of the configuration under consideration.

Panov S. Y., Kovalev A. V., Aisin A. K., Achekin A. A. Aircraft air intakes location impact on vortex formation intensity. Aerospace MAI Journal, 2018, vol. 25, no 4, pp. 110-119.

Kalugin K. S., Sukhov A. V. Methane application specifics as a fuel for liquid rocket engines. Aerospace MAI Journal, 2018, vol. 25, no 4, pp. 120-132.

At present, a significant part of the research aimed at increasing the energy and mass characteristics of liquid-propellant rocket engines (LPRE) is being performed in the field of new materials application and processing technologies. Other studies are aimed at modernizing the principle of organizing the work process. However, new fuels application is a more long-term and quickly realizable prospect notwithstanding the research costliness in the field of LPRE. Methane may appear one of the propellants, which application may become a new stage in rocket and space industry development. The article considers the historical process of methane formation as a liquid rocket fuel component since it was firstly mentioned by Tsiolkovsky in his works (“Exploration of the World Spaces with Jet Devices”, “Space Rocket”, “Jet Airplane”, “Achieving the Stratosphere”, etc.). A comparative analysis of methane with kerosene was performed in view of the similarity of the work process organization in the LPRE combustion chamber, as well as close hydrocarbon structure. A component close to methane, currently in use in rocket engines, is hydrogen due to the cryogenic nature of both components, which creates difficulties at the design stage of valves, pipelines and gas lines, as well as the organization of the work process in the combustion chamber. Additionally, analysis of the most common fuels based on kerosene, methane and hydrogen was performed. This is especially interesting, since methane fuel pair of oxygen-methane occupies an intermediate position between “oxygen-kerosene” and “oxygen-hydrogen” pairs with respect to the specific impulse and fuel mixture density. The analysis was performed based on physical-chemical, energy, operational, environmental, economic and some other characteristics. This allowed identify the main advantages and disadvantages of methane application as a LPRE fuel and determine its prospects both in Russian and foreign rocket and space industries.

A brief analysis of liquid rocket engines on methane, created or projected in NPO Energomash by V.P. Glushko, KBHA them. S.A. Kosberg, KB Himmash them. A.M. Isaev, the Research Center. M.V. Keldysh, and also to the American firm SpaceX.

Finally, it was concluded that the methane LPRE could replace oxygen-kerosene engines in the near future, since the fuel pair oxygen-methane outperformed the oxygen-kerosene pair by its energy, environmental and economic indicators. Interplanetary flights can become a special field of methane application, since a large amount of methane, the main element of natural gas, can be found almost everywhere in the solar system: on Mars, Titan, Jupiter and many other planets and satellites, which will allow refueling rockets in flight, significantly increasing them the flight range.

Kuz'michev V. S., Omar H. H., Tkachenko A. Y. Effectiveness improving technique for gas turbine engines of ground application by heat regeneration. Aerospace MAI Journal, 2018, vol. 25, no 4, pp. 133-141.

The requirements for the ground based gas turbine installations efficiency improvement are constantly increasing.

Heat conversion into work in gas turbine engines, operating by the Brayton cycle, is attended by significant losses, which depend on the cycle parameters and reach up to 60-70% or more. At present, high-tech aircraft engines and their modifications are widely used as ground-based gas turbine plants, making provision for the gas turbines efficiency improvement based on the of combined thermodynamic cycles application.

The article considers the schemes of gas turbine units (GTU) for ground application with combined cycles, allowing improve their efficiency. One of the ways for the gas turbine units cycle improving is heat regeneration of the exhaust gases by installing a heat exchanger at the turbine outlet where a part of the heat is transferred into the air behind the compressor. However, relative bulkiness and substantial weight of the heat exchanger (even of plate type) do not allow at present active application of this scheme in aviation, but it is widely employed in ground applications.

In the case of ground based gas turbine unit, heat exchangers are located in the exhaust chamber or tower behind a power turbine. Thermal ratio of the most widely used tubular heat exchangers is  θ = 0.8-0.9, and plate- type heat exchangers are characterized by the thermal ratio of θ = 0.5-0.8.

It is obvious that the main parameters of the thermodynamic cycle of gas turbines unite, such as the gas temperature T*g and the compressor pressure ratio (π* ), as well as the parameters determining the working process of additional units (heat regenerators, steam turbine, etc.) of the combined installation play an important role in its efficiency improving. Comprehensive optimization of these cycle parameters is the main goal of the gas-turbine combined unit thermodynamic design.

Computer models of a gas turbine unit with combined thermodynamic cycles developed in the ASTRA SAE-system allowed solve the problems of nonlinear multi-criteria optimization of their operating parameters, determine the most rational schemes depending on the intended purpose and operating conditions of the gas turbine unit.

Russian Turbofan engine TRDDF RD-33 was selected as the basis for studying the heat regeneration impact on efficiency effectiveness. Its low pressure compressor was cut off to eliminate the bypass duct while converting it into ground based installation.

The following variation values of the cycle basic thermodynamic parameters were selected (π*kΣ = 15, 30, 45, 60 и T*g = 1200, 1500, 1800, 2100 K). The GTE module without exhaust gases heat regeneration and a GTE with exhaust gases heat regeneration were developed employing ASTRA computer program. The paper presents some results of the study on GTE efficiency improvement.

Ogloblin D. V., Gorelov A. D., Voroshilin A. P., Zueva K. S. Automated testing system for technical diagnostics of spacecraft power supply systems. Aerospace MAI Journal, 2018, vol. 25, no 4, pp. 142-151.

One of the urgent tasks at present consists in time reduction of flying vehicle launch preparation through introduction of new technologies, equipment and various kinds of tests based on computational experiments.

One of the most expensive types of aircraft is a spacecraft for various missions and tasks in the near-earth space and interplanetary flights. The main system of a spacecraft is the power supply system. Stringent requirements on external impact stability, and operability maintenance in emergency situations, since its failure results in the spacecraft loss. The power supply system preparation and testing is a major part of the testing program and it is performed employing rather labor-intensive methods.

One of the most interesting objects for the test complex optimization are the spacecraft for astronomical observations, as they present a set of a large number of complex technical devices for various purposes and control systems. Such aircraft requires a special approach to ensuring the quality of various electrical systems tests, as well as control and monitoring systems.

It is important to ensure uninterrupted power supply of the onboard service and scientific equipment for the timely data obtaining during the mission. Thus, the task of rapid and high-quality electrical tests performing of such spacecraft is of paramount importance.

The following core systems are being subjected to comprehensive electrical testing: the onboard radio telemetry system, onboard control system, propulsion system, the solar panels orientation system, power system, electrification control system.

Besides scientific equipment, these systems form the basis of almost any spacecraft. Due to the large number of systems subjected to electrical checks, the issue of the electrical tests time reduction, while preserving their quality (guaranteed reliability level of the systems) is relevant. It is necessary to determine the level of reliability and the number of tests based on the system model.

This above said problem can be solved by optimizing the measurement devices' number and functional characteristics, as well as application of automated measurement systems (AIC) for processing a large number of parameters (with account for specifics of electrical tests). This solution allows optimize the testing process, while reducing herewith the number of employed measuring equipment and the test program cost.