Calculation of the centrifugal force of the carrying blade of the helicopter. Fundamentals of main rotor aerodynamics. Propeller design

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1 UDC: V.A. Grayvoronsky, A.G. Grebenikov I.N. Shepel, T.A. Gamanukha An approximate method for calculating the normal aerodynamic forces distributed along the rotor blade of a helicopter. NOT. Zhukovsky "KhAI" On the basis of the hypothesis of oblique sections, the questions of determining the efforts distributed along the rotor blade are considered, taking into account the compressibility and unsteadiness. Key words: blade, rotor, helicopter. A feature of the flow around the main rotor in horizontal flight is the presence of variable velocities, slip angles and angles of attack of the main rotor blade (HB) elements. The use of the carrier line scheme, as well as the decomposition of the flow into transverse and longitudinal in order to use the hypothesis of flat sections, is possible for a horizontal flight speed not exceeding 8 m / s. In fig. shows the spectrum of the flow around the blade located in the rear part of the disk at µ =, 46, from which it follows that the sliding angles along the blade change significantly. Fig. Spectrum of the flow around the rotor blade The nature of the flow around the rotor blade along the radius and azimuth at a low flight speed is shown in Fig. A, with a large one in Fig. B. The sliding angles of the blade sections differ by more than 5 times. a Fig .. Fields of velocities of the flow around the main rotor blade b 78

2 Table the values ​​of the slip angles of the flow at the blade at relative radii, 5 and, 9 for various flight speeds at azimuths and 8 are presented. Table. Slipping angles of the flow at relative radii V, km / h r =, 5 r =, With an increase in the horizontal flight speed, the influence of the reverse flow zone, where slip is also significant, also increases. If up to velocities µ =, 4 the reverse flow zone does not significantly change the magnitude of the forces and moments, then at high speeds its influence must be taken into account. The largest value of the radius of the backflow zone without taking into account o control of the blade corresponds to the azimuth ψ = 7 and is equal to r µ. Thus, the section of the blade is flowed around by a flow that is constantly changing in direction and magnitude. This circumstance leads to the need to calculate the characteristics of the blade sections by the total speed at the corresponding radius, taking into account the compressibility and non-stationarity. The total speed in the section is determined by the rotation of the blade, the movement of the helicopter, the swing motion of the blade, the inductive flow on the propeller, as well as the longitudinal centrifugal movement along the blade. Centrifugal flow occurs due to the boundary layer. As shown by numerical calculations, this flow has no significant effect on the flow around the blade. In fig. 3 shows the diagrams of the laminar and turbulent boundary layers. With a turbulent boundary layer, radial flow is practically absent due to significant tangential forces. The x coordinate defines a point along the chord in related system coordinates. For example, with a value of x =, 5 m and ω in = 5 rad / s, the highest speed from the centrifugal force in the laminar mode is Vr = .4 m / s, and in turbulent mode, which is more likely, it is ten times less, i.e. this flow can be ignored. Rice. 3. Distribution of radial velocities in the boundary layer: turbulent PS, laminar PS 79

3 The reason for the radial flow in the boundary layer can also be the pressure distribution along the blade. This can lead to a redistribution of aerodynamic load for heavily loaded propellers. The base plane for determining the kinematic parameters is the design plane of rotation of the screw (Fig. 4). Rice. 4. Kinematics of the flow around the blade in the design plane of the rotor rotation The kinematic diagram of the velocities in the cross section of the blade is shown in Fig. 5. Fig. 5. Velocity triangle of the blade section The relative velocity in the design plane of rotation at the radius r is determined by the expression W W (µ + υ) + r + (µ + υ) r sin (ψ) =. () The vertical component of the relative velocity V y = λ r β. () Then the total relative velocity in the section (µ + υ) + r + (µ + υ) r sin (ψ) + λ + r β λ β = r In these expressions, the known relative parameters are accepted: µ = V cos (α) ; λ = V sin (α) + υ; β = a sin (ψ) b cos (ψ). in in y. (3) In level flight, relative inductive velocities (4) 8

4 υ>; υ<. Определение этих скоростей может проводиться численными y методами, например методом дискретных вихрей, либо на основании дисковых теорий. Индуктивные скорости изменяются по диску НВ. Наиболее простой закономерностью является II гипотеза Глауэрта, согласно которой υ y = υ i ср (+ k cos ψ); где k коэффициент, учитывающий влияние относительного радиуса; 4 µ r k = 3 ; (5) µ, + λ υ i ср средняя по диску индуктивная скорость. Значения υ i ср и υ можно определить по дисковой теории В.И. Шайдакова . Для больших скоростей полета среднюю по диску индуктивную скорость можно определить по формуле CТ υi =, (6) ср 4 ξ µ где ξ коэффициент, учитывающий перетекание: ξ =,9,94. Параметры a,b,α в определяют в процессе аэродинамического расчета . Угол отклонения от оси х набегающего на сечение потока можно определить в зависимости от ψ согласно табл.. Угол атаки в текущем сечении это угол между хордой сечения лопасти и вектором скорости на бесконечности: () λ r β α e = ϕe cos δ + arctg (µ + υ) + r + (µ + υ) r sin(ψ). (7) Угол установки сечения ϕ e зависит в общем случае от крутки лопасти и управления АП и РВ. Его можно определить по конструктивным и балансировочным параметрам: где ϕσ ϕe = ϕ,7 + B sin r k, D коэффициенты РВ и АП; (7, r) k a + k a cos(ψ) D δ (ψ) δ балансировочный угол отклонения АП в горизонтальном полете. B, (8) Расчет усилий на лопасти с учетом пространственного характера обтекания будем проводить по гипотезе "косых" сечений, т.е. несущим профилем лопасти считается сечение по местной скорости подходящего к лопасти потока. Определение геометрии таких сечений весьма затруднительно из-за крутки, 8

5 deformation of the blade and especially in the areas of profile change and in the reverse flow zone. The section of the blade is determined by the local streamlines, which are considered rectilinear in the section of the blade and deviated from the normal section to one side or the other by an angle δ (table). Change in χ and δ depending on azimuth ψ, rad Expression for χ, rad δ, rad r cos (ψ) arctan µ + υ + r sin (ψ), χ< Направление потока на лопасти К концу ψ χ лопасти Таблица r cos(ψ) arctg + + µ υ r sin(ψ), χ < ψ + χ К комлю лопасти 3 r cos(ψ) arctg + + µ υ r sin(ψ), ψ + χ К комлю лопасти <χ< r cos(ψ) 3 arctg + + µ υ r sin(ψ), 5 К концу ψ χ лопасти <χ< При значении δ < профиль в косом сечении обтекается с носка, а при δ >from the tail. For modern helicopters, changes in speed and angle of attack in sections over time reach large values: V & ma> ± m / s, & α ma> ± o / s. This leads to a non-stationary change in all aerodynamic parameters; there is a delay in the breakdown. The movement of the helicopter is significantly different from the predicted stationary characteristics. The aerodynamic coefficients at a fixed moment of time will be determined not only by the values ​​of the speed and angle of attack at a given moment of time, but also by the process of their change in the previous time. Naturally, more distant moments in time will have a weaker effect on this process. The nature of the dependences α & = f (t) and V & = f (t) also has a significant impact. Reliable enough 8

6 there are no dependencies on this issue, but there are some experimental dependencies that allow taking this phenomenon into account. In particular, the paper describes a method for approximating experimental data by three parameters that determine the nature of the change in the angle of attack, which makes it possible to translate the results obtained to other conditions. The data of this work were used to determine the coefficient of normal force of the profile in normal sections and sections along the streamline. In addition, the normal force coefficient was corrected depending on the relative section thickness and compressibility. In the process of preliminary calculation, the kinematic parameters in the sections of the blade were determined in accordance with the above dependences. As the initial geometric, kinematic and balancing parameters of the helicopter Mi- are taken: C = ,; ω = 5.8 / s; a = 4.7; a = 5.7; in =,; T V =, 35; D =, 7; k =, 4; ϕ 7 = 4. In fig. 6 shows the kinematic parameters in azimuth W and W P in the seventh section, as well as the angles of attack α and α and the angles of conventionally undisturbed flow δ and χ. w w P α ep 5 α e 6 e HB ep 3 8 w α e 8 w P α ep Ψ Fig. 6. Kinematic parameters of the blade section in section "7" on the hypothesis of oblique sections; the subscript "p" marks the parameters according to the hypothesis of normal sections. The total velocities in the section W and W P practically change according to the 1st harmonic. Naturally, at all azimuths, the total velocity W is greater than the velocity W P, and the angle of attack along the streamline is less than the angle of attack in the normal section. The angles of orientation of the total flow δ and χ, which are more sensitive to the flapping motion of the blades, differ significantly from a simple harmonic change. In fig. 7 shows the change in angular and linear accelerations in section "7". For the specific case of calculation, α & practically varies in the range 83

7 + - / s. This change is close to the 1st harmonic. Linear acceleration W & in the range of + - m / s. The indicated circumstances of a significant change in both the angle of attack and the total speed are the reason for the non-stationarity of the aerodynamic characteristics. Unfortunately, the separate influence of these two factors on aerodynamic performance has not been studied. In fig. 7 shows the change in the normal flow load according to the hypothesis of oblique sections and normal 5 ẇ p α. P. ẇ α p Fig. 7. Change of normal force in azimuth in section "7"; the subscript "n" marks the parameters according to the hypothesis W & and α & angular and linear acceleration Ψ These data were obtained taking into account the non-stationarity in the angle of attack. The load according to the hypothesis of oblique sections is somewhat higher than according to the hypothesis of normal sections, especially in the zone of the retreating blade n ψ = ψ = 3 ψ = n ψ = Fig. 8. Change in linear load along the radius for azimuth ψ = 3 and 84

8 Change in linear load along the radius for azimuth ψ = 3 and shown in Fig. 8. For azimuth ψ = 3, the normal load for both calculation options is practically the same. At the azimuth ψ = the normal load according to the hypothesis of "oblique" sections is higher than according to the hypothesis of normal sections. This is due to the simultaneous effect of changes in speed and angle of attack on the linear load. Bibliography. Main rotor theory. [Text] Ed. A.K. Martynova, M .: Mechanical Engineering, 973. p .. Mikheev S.V., Anikin V.Kh., Sviridenko Yu.N., Kolomensky D.S. The direction of development of methods for modeling the aerodynamic characteristics of rotors. [Text] // Proceedings of the VI Forum Ros VO. M., 4.5 p. 3. Shaidakov, V.I. Disk vortex theory of a rotor with a constant load on the disk. [Text] / V.I. Shaidakov // Designing helicopters: tech. Sat. scientific. tr. // MAI, no. 38, M., p. 4. TsAGI main stages of scientific activity, / M., Fizmatlit, p. 5. Baskin, V.E. Normal force of the main rotor blade section during dynamic stall. [Text] / V.E. Baskin, V.R. Lipatov // Proceedings of TsAGI, vol. 865, p. 6. Graivoronskiy, V.A. The dynamics of a helicopter flight. [Text]: Textbook. Manual / V.A. Grayvoronsky, V.A. Zakharenko, V.V. Chmovzh. X .: Nat. aerospace. un-t them. NOT. Zhukovsky KhAI, 4. 8 p. 7. Fogarty, L.E. The laminar boundary layer on a rotating blade. / J. aeronaut Sei., Vol. 8, no. 3, 95. Received by the editors of the Editors the method of developing normal aerodynamic worms, spade worms, spadeless gunta helicopters An approimate method of calculation of normal aerodynemic effort distributed over the rotor blades of the helicopter On the basis of the hypothesis of oblique cross-sections are considered questions of definition effort distributed over the rotor blades with the compressibility and unsteadiness. Keywords: blade, rotor, helicopter. 85

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SUBSTANCE: invention relates to a method for determining in flight bending stresses on a rotor shaft of a helicopter with a main rotor torsion bushing. To determine the stresses, the flight performance characteristics are measured by standard means during the entire flight time, from which significant parameters are selected and systematized, their approximating functions are determined in order to obtain the final function of the dependence of the stresses in the rotor shaft on the selected flight parameters. technical characteristics, calculate the loads on the rotor shaft using mathematical model, signal if they are exceeded. Determination of the residual resource and control of the permissible level of loads is provided. 2 c.p. f-ly, 7 ill.

The invention relates to the field of aviation, in particular to systems for monitoring the technical condition of aircraft, namely, monitoring the level of bending stresses of the main rotor shaft of a helicopter in flight, in particular for a light multipurpose helicopter with hinged blades, for example, helicopters: ANSAT, VK-117, EC -145.

The transmission is the most complex element of the helicopter design. It is known that the largest percentage of helicopter accidents (up to 39%), according to statistics, is associated precisely with the failure of helicopter transmission units.

At the stage of developing monitoring systems, the most important thing is to determine and establish diagnostic signs of the technical condition of the helicopter transmission units. The main task in the development of a monitoring system is to establish threshold values ​​of diagnostic indicators, upon reaching which appropriate decisions on further flight safety must be made in operation. If any diagnostic sign has reached its threshold value, then a decision is made to limit the resource, to replace an extraordinary part, or to remove the transmission unit from operation. As a rule, the vast majority of diagnostic signs are not displayed in the cockpit during flight. Their analysis is carried out after the completion of the flight. However, some particularly critical diagnostic signs can be displayed during flight, if safety conditions so require.

In recent decades, promising helicopters began to use the so-called hingeless main rotors equipped with a hinged bushing, in which the functions of the horizontal, vertical and axial hinges are performed by an elastic element of an extended type - a torsion bar. The main part of the torsion bar design is an elastically deformable section. The presence of plywood of layers and slots provides the torsion streams with loading predominantly in a uniaxial stress-strain state with transverse shear and bending when the blade is swinging in the plane of rotation. This makes it possible to reduce the cost of operating the helicopter, but at the same time, the initial costs for the design and manufacture of such structures increase. Therefore, the accuracy of loading prediction and, accordingly, the estimation of the resource of the carrying system of the helicopter is today one of the key tasks of the helicopter industry.

The rotor shaft is loaded by forces and moments from its hub and the torque generated at the output of the main gearbox. Main rotor shaft length is determined by layout, aerodynamic and operational considerations.

Since the semi-rigid hub has a higher bending moment compared to the articulated hub, the control of the bending stresses of the main rotor shaft of a helicopter with a jointless hub in flight is an urgent problem.

A known system for monitoring the loading of the rotor shaft (US patent No. 2010219987, SIKORSKY AIRCRAFT, publication date 09/02/2010, IPC G06F 15/00, G08B 21/00).

A method for virtual monitoring of a load on a helicopter main rotor system in accordance with one embodiment of the present invention includes selecting at least one parameter of the aircraft during one complete rotation of the main rotor. Calculation of coefficients for obtaining a set of high-frequency signals from a parameter of at least one aircraft. Multiplying each of the plurality of high frequency signals by a factor to obtain a plurality of analyzed signals. Estimation of the rotor load based on the analyzed signals.

A real-time rotor health system in accordance with one embodiment of the present invention includes a sensor system for measuring loads to obtain data. The module is made with the possibility of virtual monitoring of loads for obtaining calculated data and detecting malfunctions in real time and obtaining an algorithm for subtracting the calculated signals from the measured signals to obtain values, which are then compared with standard values ​​to give the final result on the rotor state.

Sensors read parameters such as aircraft takeoff weight, density altitude, rotor speed, airflow speed, normal acceleration, climb rate, engine torque, pitch angle, roll angle, yaw rate, pitch rate , angular rate of roll, deflection in the longitudinal direction, lateral position, pedal position and a set of positions per revolution of the main rotor. The vectors of the given sixteen parameters are multiplied by the given values ​​of the matrix, which includes 10 rows and 16 columns, to obtain ten coefficients (c1, c2, c3, c4, c5, c6, c7, c8, c9, and c10) to determine ten values ​​of the oscillations ... The oscillation values ​​are multiplied by a factor to obtain amplified oscillations. If the vibration vectors are denoted as w1, w2, w3, w4, w5, w6, w7, w8, w9, and w10, and the coefficients are c1, c2, c3, c4, c5, c6, c7, c8, c9, and c10, then the calculated signal of the main rotor shaft shear force will be written in the form:

L = c1 * w1 + c2 * w2 + c3 * w3 + c4 * w4 + c5 * w5 + c6 * w6 + c7 * w7 + c8 * w8 + c9 * w9 + c10 * w10

The amplitude and phase of the shear force are calculated through the Fourier transform.

A known system for collecting data, monitoring and diagnosing the technical condition of helicopter propeller drive units (RF patent for invention No. 2519583, publ. 02/27/2014, IPC B64D 45/00), including piezoelectric vibration sensors, which are installed on the body, at least , one of the rotor drive units of the helicopter and are located so that they receive data with completeness sufficient to diagnose the technical condition of parts, units of at least one rotor drive unit of a working helicopter, and an on-board electronic unit. The electronic unit is connected to the outputs of the vibration sensors and is made with the possibility of digital processing of vibration signals, control and implementation of collection, primary processing and evaluation of the parameters of signals from individual sensors and / or their combinations, accumulation of sensor data and saving them on external and / or removable media suitable for computer readout, and secondary processing in terrestrial conditions. The efficiency of data collection, the information content of monitoring and diagnostics of the technical condition of the propeller drive units of an operating helicopter is increasing.

The disadvantage of this control system is the impossibility of making an unambiguous conclusion about the level of fatigue stresses in the helicopter assemblies, including the rotor shaft, based on the vibrations measured in flight. Also, the disadvantage is the need to install sensors and electronic units on helicopters, the time required for secondary data processing in ground conditions.

There is a known method of operating a helicopter (RF patent No. 2543111, publ. 02/27/2015, IPC В64С 27/04, B64F 5/00, G01L 3/24), which consists in the fact that during each flight, the actual thrust of the main rotor of the helicopter is monitored, and Before the start of helicopter operation, initial data are collected on the characteristics of the engines of the power plant in accordance with the forms and initial data are collected on the magnitude of the main rotor thrust during control hovering of the helicopter. During the entire operation of the helicopter, the actual data on the magnitude of the main rotor thrust in the hovering modes of the helicopter are collected and recorded, the statistical data on the main rotor thrust are compared with the initial values ​​using an on-board computer value, a signal is generated to the monitor with the help of an on-board computer about the need to adjust the engine parameters to values ​​that provide a deviation of the rotor thrust within 0.5% of the initial value. The regulation of the engine parameters is carried out either in automatic mode, or by service personnel on the ground. EFFECT: increased efficiency of helicopter application.

The disadvantage of this method of operation is the impossibility of determining the level of fatigue stresses on the rotor shaft based on the results obtained, because the fatigue stresses on it are determined by bending stresses. Also, the disadvantage is the need to install sensors and electronic units on helicopters, the time required for secondary data processing in ground conditions. Also, a disadvantage is the need to collect initial data on the characteristics of the engines of the power plant in accordance with the forms and collect initial data on the magnitude of the main rotor thrust during control hovering of the helicopter before starting the operation of the helicopter.

As the closest analogue, US patent No. 2011112806, publ. 2011.05.12, IPC G06F 17/10. SUBSTANCE: invention relates to a method of providing information about a critical state of a component of a rotorcraft, including at least one engine driving a rotor, including a fairing, a shaft and a plurality of blades. A sensor for measuring bending and cyclic loads acting on an aircraft rotor includes a computing unit designed to calculate (a) the current bearing temperature of the main rotor assembly using the first computational model, (b) predicting the bearing temperature using the first computational model, and (c) applying a load to a selected component of the rotor assembly using a second computational model, the first and second computational models are configured to calculate, respectively, the predicted and current value of the bearing temperature and the load acting on the selected component based on the control flight parameters; and a display unit for displaying on a single scale a movable indicator that is driven by the highest projected bearing temperature and load on the selected component. The display shows another movable indicator driven by the current bearing temperature.

The disadvantage of the prototype is the need to install external sensors, which presents certain difficulties, since the design of serial helicopters is not adapted to the installation of external sensors, in addition, in the procedures Maintenance and field repair, external sensors are not fully integrated with the rest of the aviation equipment; they require additional manuals and manuals for technical operation and additional trained specialists.

The objective of the proposed technical solution is to create a method for monitoring bending stresses on the rotor shaft during the entire flight time (from takeoff to landing) to identify shaft fatigue damage and to prevent emergencies.

The technical result is the determination of the residual resource and the control of the permissible level of loads.

The technical result is achieved by the fact that the method for determining in flight the bending stresses on the main rotor shaft of a helicopter with a torsion bushing of the main rotor includes measuring during the entire flight time by standard means of monitoring the flight performance of the helicopter, calculating the loads on the main rotor shaft using a mathematical model and signaling if they are exceeded, from the number of measured performance characteristics, significant parameters are selected and systematized, which have a direct effect on the level of loading of the rotor shaft, approximate functions of significant parameters are determined in order to determine the final function of the dependence of stresses in the rotor shaft σ (t) from the selected performance parameters, the absolute values ​​of the rates of change of the swashplate rotation angles in the longitudinal and transverse directions are added to the final function:

The proposed method makes it possible to assess the level of loading of the rotor shaft at any time of its flight operation. Based on the use of standard means for monitoring the parameters of a helicopter flight, it allows one to determine the level of bending stresses during the entire duration of the flight, use it to register flight restrictions and inform the crew about exceeding the permissible load level, as well as determine the residual life.

In the claimed invention, an analysis is made of the conditions for justifying the establishment of limit values ​​for particularly critical diagnostic features on the example of indicating the actual bending stresses of the main rotor shaft of a single-rotor helicopter in flight, in particular for ANSAT helicopters.

The essence of the invention lies in the fact that from the number of parameters monitored in flight, those parameters are selected and systematized that have a direct impact on the level of loading of the NV shaft. The approximating functions of significant parameters are determined in order to determine the final function of the dependence of the stresses in the NV shaft on the selected parameters of the LTH. The absolute values ​​of the rates of change of the swashplate rotation angles in the longitudinal and transverse directions are added to the final function.

A flight experiment is being conducted. The choice of the critical parameter is determined from the current values ​​of the helicopter's performance characteristics (LTH). To do this, a strain gauge is installed on the shaft of the helicopter and in real flight the values ​​of stresses σ ist (t), as well as the values ​​of trajectory parameters measured by standard means of monitoring the parameters of the helicopter flight, for example: the longitudinal and transverse tilt angle of the swashplate, the total pitch of the main rotor , helicopter speed, helicopter pitch angle, helicopter roll angle, rate of change of the swashplate tilt angle in the longitudinal and transverse directions, etc.

The preliminary analysis selects the parameters of the performance characteristics that have the greatest effect on the stresses on the NV shaft, for which graphs of changes in the voltage on the shaft are plotted depending on the values ​​of the parameters recorded by the standard means of control, and the correlation coefficients are found and estimated in order to filter the parameters of the performance characteristics.

Trajectory parameters of LTC with a correlation coefficient of more than 0.2 are chosen as significant ones.

Approximating curves (dependences of stresses on the rotor shaft on the selected parameters of the flight characteristics) are constructed and a system of equations is drawn up in order to determine the approximation of the function for bending stress in time σ calc (t):

and the corresponding weight coefficients A1, A2, A3, ..., An are found.

Coefficients A1, A2, A3 are found by polynomial approximation using the least squares method (for a specific helicopter with specific flight characteristics).

The final formula takes the form:

where Dprod is the angle of inclination of the swashplate in the longitudinal direction,

Dpop - the angle of inclination of the swashplate in the transverse direction,

Dosh - the common pitch of the main rotor,

X n - other significant flight performance parameters,

- the absolute value of the rate of change of the angle of rotation of the swashplate in the longitudinal direction,

- the absolute value of the rate of change of the angle of rotation of the swashplate in the transverse direction.

The calculation of the bending stress of the rotor shaft of the helicopter is carried out in real time during the entire flight time in the computing unit of the on-board computer based on the programmed program. When the safe voltage level is exceeded, the pilot is alerted and the calculation of the consumed resource in hours begins according to the formula:

where Pr is the damage caused by the voltage level exceeding the safe one;

Fri. - damage per hour of a typical flight, taken when calculating the resource for normal operating conditions.

Damage introduced by the voltage level exceeding the safe Pr is determined by the following method:

For each load level exceeding the safe one, using the fatigue curve (the curve is taken from the fatigue test results of the main rotor shaft), the corresponding number of cycles to failure (Ni) is determined;

Damage introduced by the stress level exceeding the safe Pd is defined as the ratio of the number of cycles at this level to the number of cycles to failure (Ni).

Thus, after each flight, the consumed resource of the main rotor shaft is calculated. If there was no excess of the maximum loading level, then the consumed resource of the rotor shaft is equal to the actual flight time, if the safety level of loading was exceeded, then the time determined by the method described above is added to the actual flight time.

Since there is always a measurement procedure necessary to obtain reliable information for each diagnostic feature, then, accordingly, it is also required to take into account the inevitable measurement errors for each diagnostic feature. Then the decision to exceed or not to exceed its limit values ​​should also be made taking into account the upper (or lower) tolerance of the region of limit states.

A certain limiting value of σ CR should be set, exceeding which entails a rapid exhaustion of the rotor shaft fatigue life and its possible destruction in the subsequent flight time. Since this parameter, or diagnostic feature, is especially critical, it is necessary to display its current value in the cockpit. Let us denote as - permissible by the indicator value of the current measured value σf.

The actual current value of σph can be represented as a sum:

where mσ - expected value bending stresses in the most loaded section of the rotor shaft in the considered flight mode, Δσ is the deviation of the actual value of σf from its mathematical expectation.

Description of the implementation of the invention

Practical determination of parameters affecting the level of shaft loading.

1. A flight experiment was carried out on a helicopter with a single-rotor ANSAT scheme, during which the values ​​of bending loads were measured at a specific time interval using a strain gauge mounted on the main rotor shaft. The experimental dependence σ ist (t) is shown in Fig. 1 (curve 1). This dependence was obtained for a typical flight mode, which includes the following modes:

a) Hovering (including hover turns)

b) Overclocking

c) Low speeds at the ground

d) Climb

e) Horizontal flight at different speeds

f) Bends

g) Motor planning

h) Braking

During the flight, the following trajectory parameters were measured in time using the helicopter's standard control facilities.

1. Speed, unit of measure km / h.

It was measured by the device "Speed ​​indicator USVITs-350 with digital output". The error in the digital signal output of the current indicated speed under normal climatic conditions at the nominal values ​​of the input signals does not exceed ± 6 km / h.

2. Height, unit of measure m.

Measured by devices:

- "Height indicator VMC-10" - mechanical altimeter with digital output. The error in the digital signal of the relative flight altitude, the variation of the readings with the atmospheric pressure of 760 mm Hg set on the meter. (1013 hPa) in normal climatic conditions, depending on the altitude, is: from ± 10 m (at an altitude of Ohm) to ± 30 m (at an altitude of 6000 m);

- "Radio altimeter A-053-05.02" - an airborne radar station with continuous emission of frequency-modulated radio waves. Height measurement error when flying over any smooth surface (runway type) with horizontal speed up to 120 m / s and vertical speed no more than 8 m / s at roll and pitch angles up to ± 20 ° in the altitude range from 0 to 1500 m in 95% height measurements, m: by digital output 0.45 or ± 0.02N (whichever is greater).

3. Roll angle and pitch angle of the helicopter, degrees.

Measured by the device "Aviogorizont AGB-96D" - gives the signals of the roll and pitch of the helicopter. The attitude indicator error in roll and pitch on a vibrating base is no more than ± 2.5 °.

4. The position of the controls, the unit of measure is degrees.

It is measured by the device "Potentiometric two-channel position sensors of the controls DP-M". Measurement error ± 30 ".

5. Position of the output links (rods) of the steering drives (tilt angles of the swashplate in the longitudinal and transverse directions) RP-14, mm.

It is measured by the device "Potentiometric sensors MU-615A series 1". Angle measurement error under normal conditions: ± 2% of the nominal measurement range.

6. Angular velocities, rad / s.

Measured by the device "Block of primary information sensors BDPI-09" - provides information about the projections of the vectors of angular velocity and linear acceleration.

Figures 2-7 show the dependence of the stresses on the rotor shaft on the measured parameters. The list of the given parameters is not limited to the given parameters and depends on the specific helicopter.

During the experiment, the following parameters were measured over time:

σ (t) is the value of the bending stress over time, measured by a strain gauge on the shaft,

Dprod (t) - the angle of inclination of the swash plate in the longitudinal direction,

Dpop (t) - the angle of inclination of the swashplate in the transverse direction,

Dosh (t) - common pitch of the main rotor,

V (t) - helicopter speed,

f t (t) is the pitch angle of the helicopter,

f to (t) - the roll angle of the helicopter.

The correlation coefficients are determined for each parameter

All parameters (correlation coefficient> 0.2) were chosen significant and for them approximating curves were constructed and equations were drawn up for each moment in time and for each parameter:

According to the selected significant parameters, the final formula takes the form:

Coefficients A1, A2, A3, A4, A5, A6 are found by solving the matrix equation:

The calculated values ​​of the bending stress are shown in figure 1 (curve σ calc (t)).

The proposed method makes it possible to assess the level of loading of the NV shaft at any moment of its flight operation. Based on the use of standard means for monitoring the parameters of a helicopter flight, it allows one to determine the level of bending stresses during the entire duration of the flight, use it to register flight restrictions and inform the crew about exceeding the permissible load level, as well as determine the residual life.

1. A method for determining in flight bending stresses on the main rotor shaft of a helicopter with a main rotor torsion bushing, including measuring during the entire flight time by standard means of monitoring the flight performance of the helicopter, calculating the loads on the main rotor shaft using a mathematical model and signaling if excess, characterized in that from the number of measured performance characteristics, significant parameters are selected and systematized that directly affect the level of loading of the main rotor shaft, approximating functions of significant parameters are determined in order to determine the final function of the dependence of stresses in the main rotor shaft σ (t) on of the selected flight performance parameters, the absolute values ​​of the rates of change of the swashplate rotation angles in the longitudinal and transverse directions are added to the final function:

2. The method for determining in flight the bending stresses on the main rotor shaft of a helicopter with a torsion bushing of the main rotor according to claim 1, characterized in that, to determine the significance of the performance parameters, the dependences of the stresses on the main rotor shaft on the selected parameters are constructed and the coefficients are calculated and evaluated correlation.

3. A method for determining in flight bending stresses on the main rotor shaft of a helicopter with a main rotor torsion bushing according to claim 2, characterized in that the significance of the parameters is determined by the value of the correlation coefficient> 0.2.

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Introduction

Helicopter design is a complex, evolving process that is divided into interrelated design stages and stages. The aircraft being created must meet the technical requirements and meet the technical and economic characteristics specified in the design specification. The terms of reference contains the initial description of the helicopter and its performance characteristics that provide high economic efficiency and the competitiveness of the designed machine, namely: carrying capacity, flight speed, range, static and dynamic ceiling, resource, durability and cost.

The terms of reference are specified at the stage of pre-design studies, during which patent search, analysis of existing technical solutions, research and development work are carried out. The main task of pre-design research is the search and experimental verification of new principles of functioning of the designed object and its elements.

At the stage of preliminary design, the aerodynamic scheme is selected, the appearance of the helicopter is formed and the main parameters are calculated to ensure the achievement of the specified flight performance characteristics. These parameters include: the mass of the helicopter, the power of the propulsion system, the dimensions of the main and tail rotor, the mass of fuel, the mass of instrumental and special equipment. The calculation results are used in the development of the layout of the helicopter and the compilation of the alignment sheet to determine the position of the center of mass.

The design of individual units and assemblies of the helicopter, taking into account the selected technical solutions, is carried out at the development stage technical project... In this case, the parameters of the designed units must satisfy the values ​​corresponding to the draft design. Some of the parameters can be refined in order to optimize the design. During technical design, aerodynamic strength and kinematic calculations of units, selection of structural materials and structural schemes are performed.

At the stage of the working project, the design of working and assembly drawings of the helicopter, specifications, picking lists and other technical documentation is carried out in accordance with the accepted standards

This paper presents a methodology for calculating the parameters of a helicopter at the stage of preliminary design, which is used to complete a course project in the discipline "Design of helicopters".

1. Calculation of the take-off mass of the first approximation helicopter

2. Calculation of the parameters of the main rotor of the helicopter

2.1Radius R, m, the main rotor of a single-rotor helicopter is calculated by the formula:

g- acceleration of gravity equal to 9.81 m / s 2;

p- specific load on the area swept by the rotor,

p =3,14.

Specific load value p the area swept away by the screw is selected according to the recommendations presented in the work / 1 /: where p = 280

We take the radius of the rotor equal to R = 7.9

Angular velocity w, s -1, the rotation of the main rotor is limited by the value of the peripheral speed w R the ends of the blades, which depends on the take-off weight

2.2 Relative air densities on static and dynamic ceilings

2.3 Calculation of the economic speed at the ground and at the dynamic ceiling

The relative area is determined

The value of the economic speed at the ground is calculated V s, km / h:

where I

The value of the economic speed at the dynamic ceiling is calculated V dean, km / h:

where I= 1.09 ... 1.10 is the induction coefficient.

2.4 The relative values ​​of the maximum and economic at the dynamic ceiling of the horizontal flight speeds are calculated:

where V max= 250 km / h and V dean= 182.298 km / h - flight speed;

w R= 232 m / s - the peripheral speed of the blades.

2.5 Calculation of the permissible ratio of thrust to rotor filling for maximum speed at the ground and for economic speed at a dynamic ceiling:

2.6 Main rotor thrust coefficients at the ground and at the dynamic ceiling:

2.7 Calculation of the rotor filling:

Main rotor filling s calculated for cases of flight at maximum and economic speeds:

As a calculated filling value s the main rotor is the largest value of s Vmax and s V dean .

G.V. Makhotkin

Propeller design

Air propeller has gained a reputation as an irreplaceable propulsion device for high-speed floating craft operating in shallow and overgrown waters, as well as for amphibious snowmobiles, which have to work on snow, ice and water. We have already accumulated considerable experience both in our country and abroad. propeller applications on high-speed small craft and amphibians... So, since 1964 in our country, amphibious snowmobiles (Fig. 1) KB im. A. N. Tupolev. In the United States, several tens of thousands of airboats, as the Americans call them, are operated in Florida.

The problem of creating a high-speed shallow-draft motor boat with a propeller continues to interest our amateur shipbuilders. The most accessible power for them is 20-30 liters. with. Therefore, we will consider the main issues of designing an air propulsion unit with the expectation of just such a power.

Thorough determination of geometric dimensions propeller will allow you to fully use the engine power and get a thrust close to the maximum with the available power. In this case, the correct choice of the screw diameter will be of particular importance, on which not only the efficiency of the propeller depends in many respects, but also the noise level, which is directly determined by the magnitude of the peripheral speeds.

Studies of the dependence of thrust on travel speed have established that to realize the capabilities of the propeller with a power of 25 liters. with. it must have a diameter of about 2 m. To ensure the lowest energy consumption, air must be thrown back by a jet with a larger cross-sectional area; in our particular case, the area swept by the screw will be about 3 m². Reducing the diameter of the propeller to 1 m to reduce the noise level will reduce the area swept by the propeller by 4 times, and this, despite the increase in speed in the jet, will cause a drop in thrust at mooring lines by 37%. Unfortunately, it is not possible to compensate for this decrease in thrust either by step, or by the number of blades, or by their width.

With an increase in the speed of movement, the loss in traction from a decrease in the diameter decreases; thus, increasing the speeds allows smaller propellers to be used. For propellers with a diameter of 1 and 2 m, providing maximum thrust at the mooring, at a speed of 90 km / h, the thrust values ​​become equal. Increasing the diameter to 2.5 m, increasing the thrust at the mooring, gives only a slight increase in thrust at speeds over 50 km / h. In general, each range of operating speeds (at a certain engine power) has its own optimal screw diameter. With an increase in power at a constant speed, the diameter optimal in terms of efficiency increases.

As follows from what is shown in Fig. 2 graphs, the thrust of the propeller with a diameter of 1 m is greater than the thrust of the water propeller (standard) of the outboard motor "Neptune-23" or "Privet-22" at speeds over 55 km / h, and the propeller with a diameter of 2 m - already at speeds over 30 -35 km / h. Calculations show that at a speed of 50 km / h, the kilometer fuel consumption of an engine with a propeller with a diameter of 2 m will be 20-25% less than the most economical outboard motor "Privet-22".

The sequence of selection of propeller elements according to the given graphs is as follows. The diameter of the propeller is determined depending on the required thrust at the mooring at given power on the screw shaft. If the operation of the motorboat is supposed to be in populated areas or areas where there are noise restrictions, the acceptable (for today) noise level will correspond to the peripheral speed - 160-180 m / s. Having determined, based on this conditional norm and the screw diameter, the maximum number of its revolutions, we will establish the gear ratio from the engine shaft to the screw shaft.

For a diameter of 2 m, the permissible noise level will be about 1500 rpm (for a diameter of 1 m - about 3000 rpm); thus, the gear ratio at an engine speed of 4500 rpm will be about 3 (for a diameter of 1 m - about 1.5).

Using the graph in Fig. 3, you will be able to determine the amount of thrust of the propeller if the propeller diameter and engine power have already been selected. For our example, the engine of the most available power is selected - 25 hp. with., and the diameter of the propeller - 2 m. For this particular case, the magnitude of the thrust is 110 kg.

The lack of reliable gearboxes is perhaps the biggest obstacle to overcome. As a rule, chain and belt drives made by amateurs in artisanal conditions are unreliable and have low efficiency. Forced installation directly on the motor shaft leads to the need to reduce the diameter and, consequently, reduce the efficiency of the propeller.

To determine the blade width and pitch, use the nomogram shown in Fig. 4. On the horizontal right scale, from the point corresponding to the power on the screw shaft, draw a vertical line until it intersects with the curve corresponding to the previously found screw diameter. From the point of intersection, draw a horizontal line to the intersection with the vertical drawn from a point on the left scale of the number of revolutions. The resulting value determines the coverage of the propeller being designed (aircraft manufacturers call the ratio of the sum of the widths of the blades to the diameter).

For two-blade propellers, the coverage is equal to the ratio of the blade width to the propeller radius R. Above the coverage values, the values ​​of the optimal propeller pitches are indicated. For our example, the following are obtained: coverage σ = 0.165 and relative pitch (ratio of pitch to diameter) h = 0.52. For a screw with a diameter of 1 m σ = 0.50 m and h = 0.65. A propeller with a diameter of 2 m should be 2-bladed with a blade width of 16.5% R, since the coverage is small; a propeller with a diameter of 1 m can be 6-bladed with a blade width of 50: 3 = 16.6% R or 4-bladed with a blade width of 50: 2 = 25% R. An increase in the number of blades will give an additional reduction in noise level.

With a sufficient degree of accuracy, it can be assumed that the propeller pitch does not depend on the number of blades. We give the geometric dimensions of a wooden blade with a width of 16.5% R. All dimensions in the drawing fig. 5 are given as a percentage of the radius. For example, section D is 16.4% R, located at 60% R. The chord of the section is divided into 10 equal parts, that is, 1.64% R each; the sock is broken through 0.82% R. The profile ordinates in millimeters are determined by multiplying the radius by the percentage value corresponding to each ordinate, that is, by 1.278; 1,690; 2.046 ... 0.548.

We started a conversation yesterday with, in the light disputes and discussions of the Indian tender... Now let's take a quick look at the competitor, our Mi-26, and then compare the two helicopters.

Designing a heavy rotorcraft at M.L. The mile began with a search for the most optimal layout and layout. As with the creation of the V-12, three schemes were considered: single-screw and two twin-screw - transverse and longitudinal. Initially, it was decided to use the main units from the Mi-6 and V-12 for the new machines: blades - for a single-rotor helicopter; blades, main gearboxes and control system boosters - for twin-rotor helicopters; and from the Mi-8: blades - for a transverse helicopter with 23 m main rotors. The following options were studied: a single-rotor helicopter with a 35 m main rotor; twin-screw transverse scheme with screws with a diameter of 23 and 35 m; longitudinal twin-screw configuration with 35 m rotors.However, they all had the same drawbacks - inconsistency of parameters terms of reference, low return weight and high take-off weight and, therefore, low performance characteristics.

The firm's analysts came to the conclusion that to solve the problem it is not enough to limit ourselves to the choice of optimal parameters - unconventional design methods are needed. At the same time, it was necessary to abandon both the use of serial units and the use of generally accepted design solutions.

The heavy helicopter project was given a new designation Mi-26 or "product 90". Having received a positive opinion from the NII MAP, the staff of the MVZ im. M.L. Mil "" in August 1971 began to develop a preliminary design, which was completed three months later. By this time, the military customer made changes to the technical requirements for the helicopter - increased the mass of the maximum payload from 15 to 18 tons. The project was redesigned. The Mi-26 helicopter, like its predecessor Mi-6, was intended for transportation different types military equipment, delivery of ammunition, food, equipment and other materiel, intra-front transfer of military units with military equipment and weapons, evacuation of the sick and wounded and, in individual cases, for the landing of tactical assault forces.

The Mi-26 was the first Russian helicopter of the new third generation. Such rotorcraft were developed in the late 60s - early 70s. by many foreign firms and differed from their predecessors in improved technical and economic indicators, primarily in transport efficiency. But the parameters of the Mi-26 significantly exceeded both domestic and foreign indicators of helicopters with a cargo compartment. The weight efficiency was 50% (instead of 34% for the Mi-6), the fuel efficiency was 0.62 kg / (t * km). With practically the same geometrical dimensions as the Mi-6, new apparatus had twice the payload and significantly better flight performance. The doubling of the payload had almost no effect on the takeoff weight of the helicopter.

The Scientific and Technical Council of the Ministry of Aviation Industry approved the preliminary design of the Mi-26 in December 1971. The design of the air giant involved a large amount of research, design and technological work, as well as the development of new equipment. V short time it was envisaged to create and build units and systems with low relative masses and high resources, a bench base, test components and assemblies, study the properties of structures made of new materials, study new blade profiles, aerodynamic characteristics of a helicopter, stability of lightweight blades, etc. In this regard, "" MVZ im. M.L. Mil "" attracted to close cooperation TsAGI, LII, VIAM, NIAT, TsIAM and other organizations.

In 1972 "" MVZ im. M.L. Mil "" received positive opinions from the institutes of the aviation industry and the customer. Of the two proposals submitted to the Air Force command: the Mi-26 and the rotorcraft developed by the Ukhtomsk Helicopter Plant, the military chose the Milev aircraft. An important stage in the design of the helicopter was the competent preparation of the technical task. The customer initially required the installation of a wheel drive, heavy weapons, sealing the cargo compartment on the helicopter, ensuring the operation of engines on automotive fuels and other similar improvements, entailing a significant weighting of the structure. The engineers found a reasonable compromise - minor requirements were rejected, and the main ones were met. As a result, a new cockpit layout was made, which made it possible to increase the crew from four to five people; the height of the cargo compartment, in contrast to the original project, has become the same along its entire length. The design of some other parts of the helicopter has also undergone modifications.

In 1974, the appearance of the heavy Mi-26 helicopter was almost completely formed. It had a classic layout for Mil transport helicopters: almost all the power plant systems were located above the cargo compartment; the engines put forward relative to the main gearbox and the cockpit located in the bow balanced the tail section. When designing a helicopter, for the first time, the fuselage contours were calculated by specifying surfaces with second-order curves, due to which the all-metal semi-monocoque fuselage of the Mi-26 received its characteristic streamlined "dolphin-like" shapes. In its design, it was initially envisaged to use panel assembly and glued joints of the frame.

In the forward fuselage of the Mi-26, sealed and equipped with an air conditioning system, there was a spacious and comfortable cockpit with seats for the commander (left pilot), right pilot, navigator and flight equipment, as well as a cockpit for four people accompanying the cargo and the fifth crew member. - flight mechanic. On the sides of the cabins, there were blister hatches for emergency escape from the helicopter, as well as armor plates. Under the floor of the cabins there were compartments for navigation and radio communication equipment, life support systems and auxiliary power point- gas turbine unit TA-8A, providing autonomous starting of engines, power supply of loading and unloading mechanisms and other systems. A navigation radar was located under the radio-transparent fairing in the bow.

The central part of the fuselage was occupied by a capacious cargo compartment with a rear compartment passing into the tail boom. The length of the cabin was 12.1 m (with a gangway - 15 m), the width was 3.2 m, and the height varied from 2.95 to 3.17 m. 20 tons, designed to equip a motorized rifle division, such as an infantry fighting vehicle, self-propelled howitzer, armored reconnaissance vehicle, etc. Loading of equipment was carried out on its own through the cargo hatch in the rear of the fuselage, equipped with two drop-down side flaps and a descending ladder with podrapnikov. The gangway and sash control was hydraulic. For the mechanization of loading and unloading operations, the cargo compartment was equipped with two LG-1500 electric winches and a telpher device providing loading, unloading and transportation along the cabin of loads up to 5 tons, as well as tightening wheeled non-self-propelled equipment. In addition, loading passengers or light cargo could be carried out through three gangway doors along the sides of the fuselage. In the landing version, the Mi-26 carried 82 soldiers or 68 paratroopers. Special equipment made it possible to turn the helicopter into an ambulance for transporting 60 wounded on stretchers and three accompanying paramedics within a few hours. Oversized cargo weighing up to 20 tons could be transported on an external sling. Its units were located in the structure of the load-bearing floor, so that the dismantling of the system was not required when transporting goods inside the fuselage. Behind the cargo hatch, the fuselage smoothly passed into the tail boom with a profiled end boom-keel and stabilizer.

Eight main fuel tanks with a total capacity of 12,000 liters were placed under the cargo floor of the fuselage. In the ferry version, four additional tanks with a total capacity of 14800 liters could be installed in the Mi-26 cargo compartment. Above, above the cargo compartment, there were compartments for the engines, the main gearbox and two fuel tanks. Mushroom-shaped dust protection devices were installed at the entrances to the engine air intakes. Consumable fuel tanks and engines were protected by armor.

To ensure the planned small values ​​of the mass of units and parts of the Mi-26 operating at high loads, and the required level of strength and reliability, the OKB designed, and the pilot production "" MVZ im. M.L. Milya "" built over 70 test stands, including such unique ones as a stand for re-static tests of the fuselage and chassis by the method of "dropping" a full-scale product, a closed stand for testing the main gearbox, a full-scale stand for testing the power and load-bearing systems of a helicopter, a stand preliminary static tests and fine-tuning of the fuselage compartments, a static test bench for the rear of the fuselage. When testing the fuselage, the required strength was achieved by consistently identifying weak points and strengthening them. As a result, the Mi-26 surpassed its predecessor in terms of cargo compartment volume and payload mass by almost two times, while the fuselage mass remained unchanged. Stands were also created for testing the gearboxes and shafts of the tail transmission and individual parts of the main gearbox, dynamic tests of the blades, combined tests of the articulations of the bushings and butt parts of the main and tail rotor blades, etc. were carried out. The results of bench tests were immediately taken into account when designing units and systems.

The primary task in the design of the Mi-26, like all other rotary-wing aircraft, was the creation of a modern main rotor with a low mass and high aerodynamic and strength characteristics. When developing the Mi-26 blades, the OKB engineers relied on a wealth of experience in the design and operation of blades with a steel spar and an aluminum alloy spar. Little experience of using fiberglass in blades of this size led to the designers' decision not to use it as the main material for such a large propeller. The steel spar provided a much higher fatigue strength. In addition, by this time, a unique technology for the production of steel spars with lugs for fastening to the sleeve, made in one piece with the pipe, had been developed. The main rotor blade of the heavy helicopter was designed on the basis of a steel spar and a fiberglass shaping structure. Between the inner fiberglass layer and the outer fiberglass sheathing were fiberglass power belts and lightweight foam. The rear compartment with fiberglass sheathing and honeycomb filler made of nomex paper was glued to the outer skin. Each blade was equipped with a pneumatic system for detecting through microcracks in the spar at the stage of their formation. Research conducted jointly with TsAGI to optimize the aerodynamic layout of the blades has significantly increased the efficiency of the propeller. An experimental set of five dynamically similar Mi-26 blades passed preliminary tests in 1975 at the Mi-6 flying laboratory.

For the first time in the history of helicopter engineering, the highly loaded Mi-26 main rotor was designed with eight blades. In order to assemble such a screw, the sleeve sleeves had to be made removable. The attachment of the blades to the hub was traditional, by means of three hinges, however, in the design of the axial hinge, the engineers of MVZ im. ML Mil "" introduced a torsion bar that perceives centrifugal loads. A number of joint assemblies were made using metal-fluoroplastic bearings. The vertical joints were fitted with hydraulic spring dampers. To reduce the mass of the rotor hub, titanium was used in its design instead of steel. All this made it possible to create an eight-bladed rotor with a thrust of 30% more and a mass of 2 tons less than that of the five-bladed Mi-6 propeller. The preliminary tests of the Mi-26 main rotor carried out in 1977 at the Mi-6 flying laboratory confirmed the correctness of the choice of parameters, showed high aerodynamic characteristics, the absence of various kinds of instability, a low level of vibrations, moderate stresses in the blade spars and the level of loads in the units of the carrying system. not exceeding the calculated one.

On the Mi-26 helicopter, a tail rotor was installed with the direction of rotation, in which the lower blade went against the flow. All-glass blades of a five-bladed semi-rigid tail rotor were attached to the hub by means of horizontal and axial hinges with a torsion bar. The spars of its blades were first made by hand-laying fabric, and then by a new method of machine spiral winding. Despite the twofold increase in the tail rotor thrust, its mass remained the same as that of the Mi-6 propeller. The main and tail rotor blades were equipped with an electrothermal anti-icing system. An experienced tail rotor has passed preliminary tests at the Mi-6 flying laboratory. In addition to the blades, fiberglass was used as a structural material in the manufacture of the stabilizer spar and some non-force elements of the fuselage structure.

One of the most difficult tasks was the creation of the main gearbox, which was supposed to transmit power above 20 thousand hp. For all Mil helicopters, with the exception of the Mi-1, the main gearboxes were designed by engine designers, and the Mil Design Bureau performed only a draft layout. When working on the Mi-26, the propulsion design bureaus were unable to create a main gearbox designed for the Mi-26 mass set by the project managers. The unique main gearbox was developed in-house at the cost center. Two kinematic schemes were considered: the traditional planetary and a fundamentally new multi-threaded, previously not used in the domestic helicopter industry. Studies have shown that the second scheme will provide significant gains in mass. As a result, the three-stage main gearbox VR-26, which surpasses the R-7 gearbox used on the Mi-6 in terms of transmitted power almost twice, and in terms of output torque - more than one and a half times, turned out to be heavier than its predecessor by only 8.5%. The gear ratio of the main gearbox was 62.5: 1.

The Mi-26 chassis is a tricycle, including a front and two main supports, with two-chamber shock-absorbing struts. A retractable tail support was installed under the end beam. For the convenience of loading and unloading operations, the main landing gear was equipped with a system for changing the ground clearance.

During the development of the Mi-26, special attention was paid to ensuring the autonomy of the basing, increasing the reliability and ease of operation. The presence of special ladders, hoods, manholes and hatches made it possible to carry out ground handling of the helicopter and its assemblies without the use of special airfield facilities.

Design bureau designers completed the design of most of the units and systems in 1975. By the same time, the state commission adopted the final model of the helicopter and, in accordance with a government decree, the assembly shop of the cost center began building full-scale models of the Mi-26. V.V.Shutov was appointed the new responsible leading designer. The first copy of the helicopter, assembled the following year, entered repeated static and vibration tests. In October 1977, the assembly of the first flight model was completed ahead of schedule, and on the last day of the same month, the tractor rolled out the first Mi-26 from the workshop to the development site. The finalization of the ballast-laden helicopter and its systems on the ground continued for a month and a half. Mounted on the blades, special loading flaps-moulinets made it possible to check the operation of the engines in all modes without a helicopter tether. On December 14, 1977, test pilot G.R. Karapetian for the first time tore off the helicopter from the ground and carried out a three-minute testing of systems and assemblies in the air. In February of the following year, the Mi-26 flew from the factory site to the MVZ flight research station, where it was soon demonstrated to the command of the USSR Air Force.

Together with the firm's pilot G.R. Karapetian, the factory test pilots G.V. Alferov and Yu.F. Chapaev took an active part in fine-tuning the new helicopter. The duties of the lead engineer for flight tests were performed by V.A. Izakson-Elizarov. In mid-1979, the factory test program was successfully completed. Representatives of the customer who took part in them gave a preliminary positive conclusion on the compliance of the obtained flight performance characteristics with the specified parameters. The Rostov Helicopter Production Association (RVPO) began to master the serial production of the Mi-26, and the first prototype after flaw detection and replacement of some parts at the end of October of the same year was presented to the customer for stage "A" of joint state tests.

The state tests of the Mi-26 passed in record time. This was due to the large preliminary research and experimental work carried out at the plant. At stage "A", the testers faced only one problem - the lateral low-frequency oscillations of the helicopter in some flight modes.

The flaw was eliminated after changing the rear of the cowl fairings. In addition, the designers installed a new set of blades with an improved aerodynamic layout on the prototype. In May 1979, the second flight prototype assembled at the pilot plant of the MVZ entered the state tests, on which the operation of the external suspension system, airborne transport, rigging and mooring and sanitary equipment was checked, as well as the "fitting" of placement in the cargo compartment of various combat units was carried out. technology. In April 1980, the second Mi-26 entered the Air Force Research Institute for the final second stage "B" of state joint tests, and the first device was used to practice landings in autorotation mode. The non-motorized descent and landing mode caused some concern among the testers due to the relatively low weight of the main rotor and the high load on it, but the helicopter demonstrated a guaranteed landing capability with inoperative engines.

There were no unpleasant surprises during stage "B", except for a tire that once burst. During the state tests, both helicopters made one and a half hundred flights and "scored" over 104 flight hours.

State tests ended by August 26, 1980. The final act, signed by the customer in October of the same year, stated: “Experienced medium (according to the military classification of that time, the Mi-26 was considered“ average. ”- Ed.) Military transport helicopter Mi- 26 state joint stage "B" tests passed ... Flight technical, combat and operational characteristics basically correspond to the characteristics specified by the Resolution. The static ceiling and the maximum load mass exceed those specified by the TTT ... The Mi-26 experienced military transport helicopter and its components, which received a positive assessment according to the test results, should be recommended for launch into serial production and adoption by the Soviet Army. " An attempt by American specialists of the Boeing-Vertol company, undertaken simultaneously with the Soviet helicopter builders, to create a rotary-wing giant similar in parameters to the Mi-26 under the HLH program, ended in failure.

Thus, the experience of the development and testing of the Mi-26 helicopter has shown that, firstly, the development of the theory and practice of helicopter construction makes it possible to expand the limits that limit the maximum mass of the helicopter; secondly, the greater the amount of work performed at the early stages of design, the more successful the final stage of the helicopter is; and, thirdly, the testing of units, individual elements and systems at stands and flying laboratories before the start of flights of a new helicopter can significantly reduce the time for its fine-tuning and flight tests, as well as increase safety. It should be noted that this was an example of the most successful and fruitful cooperation "" MVZ im. ML Mila "" with the Research Institute and the leadership of the Air Force.

In the mid 80s. the experienced Mi-26 was retrofitted, in accordance with the results of the combat use of helicopters in Afghanistan, with ejector exhaust devices, as well as a passive anti-aircraft defense system missile systems... The first serial Mi-26, built at the Rostov Helicopter Production Association, took off on October 25, 1980. The new helicopter was replaced on the stocks of the Mi-6. In total, about 310 Mi-26 helicopters were built in Rostov.

Deliveries of Mi-26 helicopters to separate transport and combat regiments of the Ground Forces aviation, to regiments and squadrons of border troops began in 1983. After several years of fine-tuning, they became reliable and beloved machines in the army. Combat use of the helicopter began in Afghanistan. Helicopters that were part of the 23rd air regiment of the border troops were used to transport goods, deliver reinforcements and evacuate the wounded. There were no combat losses. The Mi-26 took part in almost all armed conflicts in the Caucasus, including two "Chechen" wars. In particular, it was on the Mi-26 that the operational delivery of troops and their redeployment during the battles in Dagestan in 1999 was carried out. In addition to the army aviation and aviation, the Mi-26 border troops entered the air units of the Russian Ministry of Internal Affairs at that time. Everywhere the helicopter has proved itself to be an extremely reliable and often irreplaceable machine.

Found the use of the Mi-26 in the fight against fires and during natural disasters. In 1986, helicopters were used in the liquidation of the consequences of the accident at the Chernobolsk nuclear power plant. Given the seriousness of the situation, the designers developed and equipped the corresponding modification in just three days. The pilots of the Mi-26 dropped tens of thousands of tons of special liquid and other protective materials from their heavy trucks onto the death-breathing reactor and contaminated area.

Aeroflot began to receive Mi-26s in 1986. The Tyumen Aviation Enterprise was the first to receive them. It was during the development of gas and oil fields in Western Siberia that Rostov heavy trucks were especially useful. The unique crane-assembly capabilities of the machine were in particular demand. Only on it can cargo weighing up to 20 tons be transported and installed directly at the site of operation.

The Russian and Ukrainian Mi-26s had a chance to participate in the UN peacekeeping missions. They worked in the territory of the former Yugoslavia, Somalia, Cambodia, Indonesia, etc. Due to their unique carrying capacity, Rostov heavy trucks are in great demand abroad. There, for the last ten years, they have been operated both by domestic airlines and as part of foreign airlines that have hired helicopters for rent or leasing. One of the companies leasing the Mi-26T is the Cypriot company Nutshell. The air giant belonging to it extinguished fires, transported goods, acted under the auspices of the UN as a peacekeeper in East Timor. Mi-26T performed in Germany and other European countries transportation of heavy bulky cargo, construction and installation work during the construction of power lines, antenna mast structures, reconstruction and construction of industrial facilities, extinguishing forest and city fires.

In 2002, the Mi-26 of the Russian airline "Vertical-T" provided assistance even to the US military. A heavy-duty carrier took a downed Boeing-Vertol CH-47 Chinook helicopter, the heaviest rotary-wing aircraft of the US Army Aviation, from hard-to-reach regions of Afghanistan to the American base in Bagram. Wealthy Americans are very sensitive about saving and saving their rotorcraft.

Heavy rotary-wing aircraft are currently successfully operated for civil and military purposes both in our country and abroad. They are used for the delivery of humanitarian aid, the evacuation of refugees, the transportation of goods and equipment, for crane and assembly works, during the construction of bridges, at assembly heavy equipment industrial enterprises, during the construction of drilling rigs, power lines, unloading ships in the outer roadstead and many other types of work, both in ordinary and hard-to-reach areas.

After the demonstration of the Mi-26 at the air show in Le Bourget in 1981, foreign customers became interested in the world's most cargo-lifting helicopter. The first four copies of the air giant were purchased by India. After the collapse of the Soviet Union, heavy vehicles ended up, in addition to the Russian Armed Forces, in the armies of the CIS countries. They are also operated by North Korea (two helicopters), South Korea (one), Malaysia (two), Peru (three), Mexico (two), Greece and Cyprus. In 2005, Venezuela placed an order for the Mi-26. The further expansion of the use of the Mi-26, both in our country and abroad, is facilitated by the receipt for it in 1995. domestic certificate of airworthiness.

Well, now let's move on directly to the analysis of the Indian tender participants.

Not so long ago, news came from India about the result of a tender for the purchase of an attack helicopter. That tender was won by the American Boeing AH-64D, which surpassed the Russian Mi-28N in a number of characteristics. Now there is new information about the course of another tender concerning the supply of helicopters, and again the situation may be unpleasant for Russia. But first things first.

Last Sunday, the Indian edition of the Times Of India published information about the upcoming completion of the competition, the purpose of which is to buy a dozen heavy transport helicopters by the Indian Air Force. The main competitors during these "competitions" were the Boeing CH-47 Chinook and Mi-26T2 helicopters. Despite belonging to the same class, these machines differ significantly in their characteristics. First of all, it is worth remembering the payload of these rotorcraft. The American CH-47 helicopter of the latest modifications can lift cargo with a total weight of over twelve tons into the air, and for the Russian Mi-26T2 this parameter is 20 thousand kilograms. Thus, the characteristics of both helicopters can transparently hint at the result of the competition.

However, the Times Of India came up with a completely unexpected piece of news. With reference to a source in the Indian Ministry of Defense, the publication writes that the winner has already been chosen, and this is not a Russian car. The source named the lower cost of the American helicopter as the main reason for this choice. In addition, Indian journalists mentioned some technical superiority of the Chinook. Such a message looks at least strange. Until now, all competitions with the participation of Mi-26 helicopters of different modifications ended in the same way: the signing of a contract with Russia. Now it is argued that Russian helicopter not only did not win the competition, but for some reason it became worse than the American rotorcraft cars, which is markedly different from him. Let's try to understand the current situation.

First of all, it is worth touching on the technical characteristics. As already mentioned, the Russian helicopter has a large payload. Moreover, according to this parameter, no helicopter in the world can compete with the Mi-26. The record high carrying capacity is supported by the size of the cargo compartment: 12x3.25x3 meters (approximately 117 cubic meters). The CH-47 cargo bay, in turn, is noticeably smaller: 9.2x2.5x2 meters (about 45 cubic meters). It is not hard to guess which helicopter will be able to carry more weight and volumetric cargo. In terms of carrying capacity, we can recall two cases when Russian Mi-26 helicopters took out damaged CH-47s from Afghanistan. In addition, the normal take-off weight of American helicopters is only a couple of tons higher than the maximum payload of the Russian Mi-26. With regard to flight data, then speed and the range of the Mi-26 and CH-47 are approximately equal. Thus, in technical terms, the Russian helicopter clearly wins. Naturally, provided that the customer needs a vehicle with a carrying capacity of two dozen tons. Judging by the initial terms of reference of the competition, the Indian Air Force wants to get just such helicopters.

Let's move on to the financial side of the matter. According to open sources, late-modification CH-47 helicopters cost foreign customers about $ 30 million apiece. There is no such information regarding the Mi-26T2, but the previous helicopters of this model cost no more than 25 million. In other words, even with a significant change in the composition of the equipment, engines etc. the Russian helicopter of the new modification turns out to be, at least, not more expensive than the American one. Perhaps, when calculating the economic nuances, the Indian tender committee took into account not only the price of the helicopters, but also the cost of maintenance. However, this argument does not seem entirely correct due to the better carrying capacity of the Mi-26T2. It is quite obvious that a large payload will cost the operator an appropriate amount. Here, the reasoning again returns to the technical conditions of the competition, in which the carrying capacity of 20 tons was spelled out. Why, one wonders, include such a requirement, if you simply feel sorry for the money to buy the helicopters that meet it?

However, the most interesting information that can shed light on the results of the Indian competition came from RIA Novosti. The Russian news agency also refers to an anonymous source, this time close to our defense industry. Despite his anonymity, this person shared quite obvious and expected information. The Novosti source claims that Russian helicopter manufacturers have not yet received any official notification of the outcome of the Indian competition. Perhaps the source of RIA Novosti, for some reason, does not have the proper information, but a number of things allow us to recognize the correctness of his words. The decision of the competition committee, as always happens, will immediately be announced and distributed by means mass media... And at the moment we have information only from unofficial anonymous sources. First of all, an unnamed person from the Indian Ministry of Defense is suspicious. The fact is that the accepted as true statement about winning CH-47 raises too many doubts and questions, both technical and economic. The source of the Russian RIA Novosti, in turn, shared information that does not obviously contradict the logic and a number of other facts.

Thus, currently news about the results of the tender for the supply of heavy transport helicopter for the Indian Air Force should be recognized as a rumor, at least not having official confirmation. At the same time, until the announcement of the results of the tender by the commission of the Indian Ministry of Defense, the question of the winner remains open. In such a situation, it is worth waiting for the end of the work of the competition committee and reconcile with reality your suspicions regarding one or another anonymous source.