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Calculation of the takeoff weight of a helicopter of the first approximation. To the calculation of the flight characteristics of a helicopter at the design stage Calculation of the main parameters and development of the layout of the helicopter

INTRODUCTION

Helicopter design is a complex process that develops over time, divided into interrelated design stages and stages. The created aircraft must meet technical requirements and comply with the technical and economic characteristics specified in the terms of reference for the design. The terms of reference contain the initial description of the helicopter and its performance characteristics, providing high economic efficiency and 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-project research, during which a 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, an aerodynamic scheme is selected, the appearance of the helicopter is formed, and the calculation of the main parameters is performed to ensure the achievement of the specified flight performance. These parameters include: helicopter mass, power propulsion system, dimensions of the main and tail rotors, mass of fuel, mass of instrumentation and special equipment. The calculation results are used in the development layout diagram helicopter and compiling a balance sheet to determine the position of the center of mass.

The design of individual units and components of the helicopter, taking into account the selected technical solutions, is carried out at the stage of developing a technical project. At the same time, 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. At technical design aerodynamic strength and kinematic calculations of units are carried out, as well as the choice of structural materials and design schemes.

At the stage of the detailed design, the execution of working and assembly drawings of the helicopter, specifications, picking lists and other technical documentation in accordance with 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 "Helicopter Design".

1. Calculation of the takeoff weight of a helicopter of the first approximation

where is the payload mass, kg;

Crew weight, kg.

Range of flight

2. Calculation of parameters rotor helicopter

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

where is the takeoff weight of the helicopter, kg;

g - free fall acceleration equal to 9.81 m/s 2;

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

=3,14.

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

We accept the radius of the main rotor equal to R= 7.9

Angular velocity , s -1, rotation of the main rotor is limited by the circumferential speed R the ends of the blades, which depends on the takeoff weight of the helicopter and amounted to R= 232 m/s.

C -1.

RPM

2.2 Relative air densities on static and dynamic ceilings

2.3 Calculation of the economic speed near the ground and on the dynamic ceiling

The relative area of the equivalent harmful plate is determined:

Where S uh= 2.5

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

where I = 1,09…1,10 - coefficient of induction.

km/h.

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

where I = 1,09…1,10 - coefficient of induction.

km/h.

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

where V max=250 km/h and V din\u003d 182.298 km / h - flight speed;

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

2.5 Calculation of the permissible ratios of the thrust coefficient to the filling of the main rotor for the maximum speed near the ground and for the economic speed on the dynamic ceiling:

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

2.7 Calculation of the filling of the main rotor:

Rotor filling calculated for cases of flight at maximum and economic speeds:

As an estimated filling value rotor, the largest value is taken from Vmax And V din:

chord length b and elongation rotor blades will be equal to:

Where zl is the number of rotor blades (zl = 3)

2.8 Relative increase in main rotor thrust to compensate for the aerodynamic drag of the fuselage and horizontal tail:

where Sf is the area of the horizontal projection of the fuselage;

S th - the area of the horizontal tail.

S f \u003d 10 m 2;

S go \u003d 1.5 m 2.

3. Calculation of the power of the helicopter propulsion system.

3.1 Calculation of power when hovering on a static ceiling:

Specific power required to drive the main rotor in hover mode on a statistical ceiling is calculated by the formula:

where N H st- required power, W;

m 0 - takeoff weight, kg;

g - free fall acceleration, m/s 2;

p - specific load on the area swept by the main rotor, N / m 2;

st - relative air density at the height of the static ceiling;

0 - relative efficiency main rotor in hover mode ( 0 =0.75);

The relative increase in main rotor thrust to balance the aerodynamic drag of the fuselage and horizontal tail:

3.2 Calculation of specific power in level flight at maximum speed

Specific power required to drive the main rotor in level flight at maximum speed is calculated by the formula:

where is the peripheral speed of the ends of the blades;

Relative equivalent harmful plate;

I uh- induction coefficient, determined depending on the flight speed according to the following formulas:

At km/h,

At km/h

3.3 Calculation of specific power in flight at a dynamic ceiling with economic speed

The specific power to drive the main rotor on a dynamic ceiling is:

where din- relative air density on the dynamic ceiling,

V din- economic speed of the helicopter on the dynamic ceiling,

3.4 Calculation of specific power in flight near the ground at economic speed in the event of one engine failure during takeoff

Specific power required to continue takeoff at economic speed in case of failure of one engine is calculated by the formula:

where is the economic speed near the ground,

3.5 Calculation of specific reduced powers for various flight cases

3.5.1 The specific reduced power when hovering on a static ceiling is:

where is the specific throttle characteristic, which depends on the height of the static ceiling H st and is calculated by the formula:

0 - power utilization factor of the propulsion system in the hover mode, the value of which depends on the takeoff weight of the helicopter m 0 :

At m 0 < 10 тонн

At 10 25 tons

At m 0 > 25 tons

3.5.2 The specific reduced power in level flight at maximum speed is:

where is the power utilization factor at maximum flight speed,

Throttle characteristics of engines, depending on airspeed V max :

3.5.3 Specific reduced power in flight at dynamic ceiling with economic speed V din is equal to:

where is the power utilization factor at the economic flight speed,

and - engine throttling levels depending on the height of the dynamic ceiling H and flight speed V din according to the following throttle characteristics:

3.5.4 The specific reduced power in flight near the ground with an economic speed in case of failure of one engine on takeoff is equal to:

where is the power utilization factor at the economic flight speed,

The degree of engine throttling in emergency mode,

n =2 - number of helicopter engines.

3.5.5 Calculation of the required power of the propulsion system

To calculate the required power of the propulsion system, the maximum value of the specific reduced power is selected:

Required power N helicopter propulsion system will be equal to:

where m 0 1 - helicopter takeoff weight,

g = 9.81 m 2 / s - free fall acceleration.

3.6 Choice of engines

We accept two turboshaft engines VK-2500 (TV3-117VMA-SB3) with a total power of each N\u003d 1.405 10 6 W

The VK-2500 (TV3-117VMA-SB3) engine is intended for installation on new generation helicopters, as well as for replacing engines on existing helicopters to improve their flight performance. It was created on the basis of a serial certified engine TV3-117VMA and is produced at the Federal State Unitary Enterprise “Plant named after V.Ya. Klimov".

4. Calculation of the mass of fuel

To calculate the mass of fuel that provides a given flight range, it is necessary to determine the cruising speed V kr. The calculation of cruising speed is carried out by the method of successive approximations in the following sequence:

a) the value of the cruising speed of the first approximation is taken:

km/h;

b) the induction coefficient is calculated I uh:

At km/h

c) the specific power required to drive the main rotor in flight in cruising mode is determined:

where is the maximum value of the specific reduced power of the propulsion system,

Coefficient of change in power depending on the flight speed V kr 1 , calculated by the formula:

d) The cruising speed of the second approximation is calculated:

e) The relative deviation of the speeds of the first and second approximation is determined:

When the cruising speed of the first approximation is refined V kr 1 , it is taken equal to the calculated speed of the second approximation . Then the calculation is repeated from point b) and ends under the condition .

Specific fuel consumption is calculated by the formula:

where is the coefficient of change in the specific fuel consumption depending on the mode of operation of the engines,

Coefficient of change in specific fuel consumption depending on flight speed,

Specific fuel consumption in takeoff mode.

In the case of flight in cruise mode, the following is accepted:

At kW;

At kW.

kg/Wh,

The mass of fuel spent on the flight m T will be equal to:

where is the specific power consumed at cruising speed,

Cruising speed,

L - range of flight.

5. Determination of the mass of components and assemblies of the helicopter.

5.1 The mass of the main rotor blades is determined by the formula:

where R - rotor radius,

- filling of the main rotor,

kg,

5.2 The mass of the main rotor hub is calculated by the formula:

where k Tue- weight coefficient of bushings of modern designs,

k l- coefficient of influence of the number of blades on the mass of the bushing.

You can take into account:

kg/kN,

therefore, as a result of the transformations, we get:

To determine the mass of the main rotor hub, it is necessary to calculate the centrifugal force acting on the blades N CB(in kN):

KN,

kg.

5.3 Mass of the booster control system, which includes the swashplate, hydraulic boosters, the main rotor control hydraulic system is calculated by the formula:

where b- blade chord,

k boo- weight coefficient of the booster control system, which can be taken equal to 13.2 kg/m3.

Kg.

5.4 Weight of the manual control system:

where k RU- weight coefficient of the manual control system, taken for single-rotor helicopters equal to 25 kg/m.

Kg.

5.5 The mass of the main gearbox depends on the torque on the main rotor shaft and is calculated by the formula:

where k ed- weight coefficient, the average value of which is 0.0748 kg/(Nm) 0.8.

The maximum torque on the main rotor shaft is determined through the reduced power of the propulsion system N and screw speed :

where 0 - power utilization factor of the propulsion system, the value of which is taken depending on the takeoff weight of the helicopter m 0 :

At m 0 < 10 тонн

At 10 25 tons

At m 0 > 25 tons

N m

Mass of the main gearbox:

Kg.

5.6 To determine the mass of the tail rotor drive units, its thrust is calculated T rv :

where M nv- torque on the rotor shaft,

L rv- the distance between the axes of the main and tail screws.

The distance between the axes of the main and tail screws is equal to the sum of their radii and clearance between the ends of their blades:

where - gap taken equal to 0.15 ... 0.2 m,

The radius of the tail rotor, which, depending on the takeoff weight of the helicopter, is:

At t,

At t.

Power N rv, spent on the rotation of the tail rotor, is calculated by the formula:

where 0 - relative efficiency of the tail rotor, which can be taken equal to 0.6 ... 0.65.

Torque M rv transmitted by the steering shaft is equal to:

N m

where is the frequency of rotation of the steering shaft,

with -1,

Torque transmitted by the transmission shaft, N m, at a speed of rotation n in= 3000 rpm equals:

N m

Weight m in transmission shaft:

wherek in- weighting factor for the transmission shaft, which is equal to 0.0318 kg / (Nm) 0.67.

Weight m etc intermediate gear is equal to:

where k etc- weighting factor for the intermediate gearbox, equal to 0.137 kg / (Nm) 0.8.

Weight of the tail gear that rotates the tail rotor:

where k xp- weighting factor for the tail gear, the value of which is 0.105 kg/(Nm) 0.8

kg.

5.7 The mass and main dimensions of the tail rotor are calculated depending on its thrust T rv .

Thrust coefficient C rv tail rotor is equal to:

Tail rotor blade filling rv calculated in the same way as for the main rotor:

where is the allowable value of the ratio of the thrust coefficient to the filling of the tail rotor.

chord length b rv and elongation rv tail rotor blades is calculated by the formulas:

where z rv- number of tail rotor blades.

Mass of tail rotor blades m LR calculated by the empirical formula:

The value of centrifugal force N cbr acting on the tail rotor blades and perceived by the hub hinges,

Tail rotor hub weight m tuesday calculated using the same formula as for the main rotor:

where N CB- centrifugal force acting on the blade,

k Tue- weight coefficient for the sleeve, taken equal to 0.0527 kg/kN 1.35

k z- weighting factor depending on the number of blades and calculated by the formula:

5.8 Calculation of the mass of the helicopter propulsion system

Specific gravity of the helicopter propulsion system dv calculated by the empirical formula:

where N- power of the propulsion system.

The mass of the propulsion system will be equal to:

kg.

5.9 Calculation of the mass of the fuselage and equipment of the helicopter

The mass of the helicopter fuselage is calculated by the formula:

where S ohm- the area of the washed surface of the fuselage, which is determined by the formula:

M 2,

m 0 - takeoff weight of the first approximation,

k f- coefficient equal to 1.7.

kg,

Fuel system weight:

where m T- the mass of fuel spent on the flight,

k ts- weighting factor taken for the fuel system equal to 0.09.

kg,

The mass of the helicopter landing gear is:

where k w- weighting factor depending on the chassis design:

For non-retractable landing gear,

For retractable landing gear.

kg,

The mass of the electrical equipment of the helicopter is calculated by the formula:

where L rv- the distance between the axes of the main and tail screws,

z l- number of rotor blades,

R - rotor radius,

l- relative elongation of the main rotor blades,

k etc And k email- weight coefficients for electrical wires and other electrical equipment, the values of which are equal to:

kg,

Mass of other helicopter equipment:

where k etc- weight coefficient, the value of which is equal to 2.

kg.

5.10 Calculation of the second approximation helicopter takeoff mass

The mass of an empty helicopter is equal to the sum of the masses of the main units:

Takeoff weight of the helicopter of the second approximation m 02 will be equal to the sum:

where m T - mass of fuel,

m gr- mass of payload,

m eq- mass of the crew.

kg,

6. Description of the layout of the helicopter

The designed helicopter is made according to a single-rotor scheme with a tail rotor, two gas turbine engines and two-bearing skis. The fuselage of the frame structure helicopter consists of the nose and central parts, tail and end beams. In the bow there is a two-seat crew cabin, consisting of two pilots. Cabin glazing provides good review, right and left sliding blisters are equipped with emergency release mechanisms. In the central part there is a cabin measuring 6.8 x 2.05 x 1.7m, and a central sliding door measuring 0.62 x 1.4m with an emergency drop mechanism. cargo cabin It is designed for transportation of goods weighing up to 2 tons and is equipped with folding seats for 12 passengers, as well as nodes for attaching 5 stretchers. In the passenger version, 12 seats are placed in the cabin, installed with a step of 0.5m and a passage of 0.25m; and in the back there is an opening for the rear entrance door, consisting of two wings.

The tail boom of riveted construction of beam-stringer type with a working skin is equipped with nodes for attaching a controlled stabilizer and a tail support.

Stabilizer with a size of 2.2 m and an area of 1.5 m 2 with a NACA 0012 profile of a single-spar design, with a set of ribs and duralumin and fabric sheathing.

Double-support, skis, self-orienting front support, dimensions 500 x 185mm, main support shaped type with liquid-gas two-chamber shock absorbers, dimensions 865 x 280mm. The tail support consists of two struts, a shock absorber and a support heel; ski track 2m, ski base 3.5m.

Main rotor with hinged blades, hydraulic dampers and pendulum vibration dampers, mounted with a forward inclination of 4° 30". The blades are rectangular in plan with a chord of 0.67 m and NACA 230 profiles and a geometric twist of 5%, the tip speed of the blades is 200 m/s, the blades are equipped with a visual spar damage alarm system and an electrothermal anti-icing device.

The tail rotor with a diameter of 1.44m is three-bladed, pusher, with a cardan-type sleeve and all-metal rectangular-shaped blades in plan, with a chord of 0.51m and a NACA 230M profile.

The power plant consists of two turboshaft gas turbine engines with a free turbine VK-2500 (TV3-117VMA-SB3) of the St. V.Ya.Klimov with a total power of each N = 1405 W, installed on top of the fuselage and closed by a common hood with opening doors. The engine has a nine-stage axial compressor, an annular-type combustion chamber and a two-stage turbine. The engines are equipped with dust protection devices.

The transmission consists of the main, intermediate and tail gearboxes, brake shafts, main rotor. The main gearbox VR-8A is three-stage, it provides power transmission from the engines to the main rotor, tail rotor and fan for cooling, engine oil coolers and the main gearbox; the total capacity of the oil system is 60kg.

The control is duplicated, with rigid and cable wiring and hydraulic boosters driven from the main and backup hydraulic systems. The AP-34B four-channel autopilot ensures the stabilization of the helicopter in flight in terms of roll, heading, pitch and altitude. The main hydraulic system provides power to all hydraulic units, and the backup one - only hydraulic boosters.

The heating and ventilation system provides the supply of heated or cold air to the crew and passenger cabins, the anti-icing system protects the main and tail rotor blades, the front windows of the crew cabin and engine air intakes from icing.

Equipment for instrument flight in difficult meteorological conditions day and night includes two artificial horizons, two speed indicators HB, combined exchange rate system GMK-1A, automatic radio compass, radio altimeter RV-3.

The communication equipment includes R-860 and R-828 VHF command radio stations, R-842 and Karat communication HF radio stations, SPU-7 aircraft intercom.

7. Helicopter balance calculation

Table 1. Empty Helicopter Balance Sheet

Unit name	unit weight, m i, kg	Coordinate x i center of mass of the unit, m	Static moment of the unit M xi	Coordinate y i center of mass of the unit, m	Static moment of the unit M yi
1 main rotor
1.1 Blades
1.2 Sleeve
2 Control system
2.1 Booster control system
2.2 Manual control system
3 Transmission
3.1 Main gearbox
3.2 Intermediate gearbox
3.3 Tail gear
3.4 Transmission shaft
4 Tail screw
4.1 Blades
4.2 Sleeve
5 Propulsion system
6 Fuel system
7 Fuselage
7.1 Bow (15%)
7.2 Middle part (50%)
7.3 Tail section (20%)
7.4 Fixing the gearbox (4%)
7.5 Hoods (11%)

8.1 Main (82%)
8.2 Front (16%)
8.3 Tail support (2%)
9 Electrical equipment
10 Equipment
10.1 Instruments in the cockpit (25%)
10.2 Radio equipment (27%)
10.3 Hydraulic equipment (20%)
10.4 Pneumatic equipment (6%)

Static moments are calculated M cx i And M su i relative to the coordinate axes:

The coordinates of the center of mass of the entire helicopter are calculated by the formulas :

Table 2. Centering list with maximum load

Table 3. Centering list with 5% remaining fuel and full commercial load

Center of mass coordinates empty helicopter: x0 = -0.003; y0 = -1.4524;

Center of mass coordinates with maximum load: x0 =0.0293; y0 = -2.0135;

Center of mass coordinates with 5% remaining fuel and full payload narrow: x 0 \u003d -0.0678; y 0 = -1,7709.

Conclusion

In this course project, calculations of the take-off weight of the helicopter, the mass of its components and assemblies, as well as the layout of the helicopter were carried out. During the layout process, the balance of the helicopter was clarified, the calculation of which is preceded by the preparation of a weight report based on the weight calculations of the units and the power plant, lists of equipment, equipment, cargo, etc. The purpose of the design is to determine the optimal combination of the main parameters of the helicopter and its systems that ensure the fulfillment of the specified requirements.

To the calculation of helicopter flight characteristics at the design stage

In his publications in 1999-2000. magazine "AON" has repeatedly raised the issue of the expediency of development and production of helicopters of various classes in Ukraine. After the scientific-practical conference "Perspective multi-purpose Ukrainian helicopter of the XXI century", organized on the basis of LLC "Aviaimpeks" in October 1999, there has been some progress in resolving this problem. Currently, a number of projects for the development and production of light helicopters are being implemented in Ukraine. Some samples and models of the designed helicopters were presented at the Aviamir-XXI air shows in 1999 and 2000.

We were especially impressed by a letter from V.N. Alekseev from Dnepropetrovsk ("AON" No. 12, 1999), in which he called for the creation of the necessary theoretical and scientific base necessary for the development of helicopter construction in our state. This must be done because specialized helicopter firms, research institutes and universities that would be deeply involved in theoretical and experimental research in the areas of aerodynamic and strength calculations, motion dynamics, control systems, etc. in relation to a helicopter, currently in Ukraine there is none. At the same time, foreign firms pay great attention creation of modeling centers and the development of effective mathematical models, investing considerable funds in this.

At the stage of the preliminary design (preliminary design), when the basic design solutions are laid down, the aerodynamic and weight parameters of the helicopter, its units and systems are determined, it is necessary to find the area of geometric and kinematic parameters of the main and tail rotors, under which the flight performance specified in the tactical and technical requirements is met. technical characteristics of the future helicopter. At the same time, it is necessary to make maximum use of statistical data on domestic (Soviet) and foreign analogues, as well as modern mathematical methods and calculation models.

In the process of designing helicopters, there are always several intermediate stages that must be achieved within a strictly defined time frame at a certain cost. Violation of calendar or budget constraints can lead to the most serious consequences for both the project and the design organization. Figure 1 shows the increase in the cost of making changes to the project aircraft at various stages of its creation, which indicates the importance and responsibility of the decisions made at the stage of preliminary design.

In this article, the authors propose a numerical method for calculating the main flight characteristics of a helicopter, based on the well-known approach to the aerodynamic calculation of a helicopter using the Mil-Yaroshenko method. Unlike the Mil-Yaroshenko graphic-analytical method, the proposed approach allows numerically solving the problem of aerodynamic calculation of a simplified layout consisting of a main and tail rotor, based on the equations of the Glauert-Locke impulse theory.

1. Statement of the problem. Basic ratios

We consider a steady straight flight of a helicopter with small trajectory inclination angles. At a given main rotor speed (HB), we consider that its thrust balances the weight of the helicopter. It is possible to change the projection of the resultant HB force on the direction of helicopter movement only by changing the angle of attack of the main rotor (Fig. 2). To maintain the balance of forces along the vertical, it is necessary to change the angle of the common pitch HB and the power transmitted to the propeller.

We write the equation of motion of a helicopter in steady horizontal flight as follows:

To equations (1) we add an equation expressing the equality of the powers on the NV shaft Nn and the power plant of the helicopter Nsu

where x is the power loss factor.

The angle between the direction of the resultant and the normal to the velocity vector can be determined from the relation

(N/T<< 1), и в горизонтальном полете выполняется условие R » T. Тогда уравнения движения вертолета (1) - (2) принимают вид

Helicopter detrimental drag coefficient, related to swept area HB;

Coefficient

filling HB;

Circumferential speed of the end of the blade HB.

The angle of inclination of the resultant HB force required for horizontal flight is found from the first equation of system (4)

The maximum angle of inclination of the trajectory with a steady climb is found from the relationship:

where is the value of the angle of inclination of the resultant when using the entire available power of the power plant in a given flight mode.

The task of the calculation is to determine the required angle of inclination of the resultant for each steady-state flight mode of the helicopter. The flight mode of the helicopter is set by the flight altitude H, the coefficient of the propeller mode m or the relative flight speed . The vertical velocities of a steady climb are found by the formula

The values of the coefficients of longitudinal force and torque NV included in formulas (3), (4) were determined by the formulas of works . These formulas are as follows:

Percolation coefficient

(8)

Angle of attack HB

Torque factor HB

Longitudinal force coefficient

Included in equations (10) and (11), the coefficients of the first harmonics of the flap motions of the blades were found using simplified formulas (12) - (14).

The value of the end loss coefficient B HB included in formulas (8) - (14) was determined according to the recommendations , and the inertial-mass characteristics of the blade can be calculated using approximate formulas .

When calculating the characteristics of the tail rotor (RV), it was considered that the condition of the helicopter's track balancing is fulfilled in all flight modes:

From this condition, the required value of the thrust coefficient RV was found:

where - the fill factor and peripheral speed of the end of the blade RV, respectively.

Then, according to formulas (8) - (14), the aerodynamic characteristics of the RV were calculated.

Of great practical interest are the characteristics of the helicopter's descent in the self-rotation mode. In this case, it is important to know the required values of the angles of the common pitch j 0.7 HB depending on the rate of descent in order to maintain a constant given speed of the HB.

Calculation of the helicopter descent characteristics in the HB self-rotation mode is carried out on the basis of the aerodynamic quality of the helicopter , (17).

t is the NV thrust coefficient at a given flight mode;

The coefficient of propulsive force HB in the mode of self-rotation.

The angle of descent of the helicopter in the self-rotation mode HB is equal to the reverse quality of the helicopter

The horizontal and vertical components of the helicopter descent rate are found from the relations

The proposed method makes it possible to calculate the main flight characteristics of a helicopter at the stages of preliminary design, when the blade profile is selected, the geometric, kinematic, inertial-mass parameters of the main and tail rotors, the characteristics of the power plant and the flight weight of the helicopter are known.

The calculation is performed for different heights in the range of flight values of the operating mode coefficient when the angles of the common blade pitch change from j 0.7 = 2° to 20° with a step of 2°.

2. Substantiation of the reliability of the results obtained

The substantiation of the reliability of the results obtained by the proposed method was carried out on the basis of solving test problems to determine the flight characteristics of known helicopters.

On fig. Figure 3 shows the altitude dependences of the characteristic flight speeds of the Mi-4 and Mi-34 helicopters. The calculation results are compared with the work data. For the Mi-4 helicopter, the calculation was performed for the flight weight m=7200 kg and the peripheral speed of the blade tip wR=196 m/s, the Mi-34 helicopter was calculated in the aerobatic version with m=1020 kg and wR=206 m/s.

Comparison of the calculated data on the required angles of the common pitch NV of the Mi-34 helicopter for level flight at the nominal engine operation mode (wR=180 m/s) for different altitudes is illustrated in Fig. 4.

On the graphs of Fig. Figure 5 shows the dependences of the vertical speed and descent angle of the Mi-4 helicopter in the HB self-rotation mode for a height of H=0 km.

The limited volume of the article does not allow us to provide all the calculated material for these helicopters.

Methodological studies have shown that the proposed method allows one to analyze the influence of numerous parameters that determine the helicopter flight mode on its flight characteristics with a sufficient degree of accuracy. Within the change of the operating mode coefficient m from 0.08 to 0.3, when the angles of attack of the blade sections along the HB disk do not exceed the maximum allowable, the assumptions made in the theory about the linearity of the dependence Cy(a) and Схрср=const are valid, this method provides an error calculations, not exceeding 8-10%. For light helicopters, this corresponds to a swept area load G/F of up to 25 kgf/m2 and maximum flight speeds of up to 220-230 km/h.

3. Calculation examples

The article presents some results of calculations of the flight characteristics of helicopters Robinson R22 (m=620 kg, wR=217 m/s) and Hughes 269В/300 (m=930 kg, wR=202 m/s). The geometric and kinematic parameters of the main and tail rotors, as well as helicopters as a whole, are taken from the work.

The R22 helicopter has a two-bladed HB with a diameter of 7.67 m (sn=0.03) and a NACA-63015 blade profile, the load on the swept area is 13.45 kgf/m2. As a power plant, one Lycoming U-320-B2C piston engine with a take-off power of N = 160 hp is used.

The helicopter model 269/300 uses a three-bladed propeller with a diameter of D = 8.18 m (sn = 0.04) and a blade profile NACA-0015, the load on the swept area is 17.7 kgf/m2. The Lycoming HIO-360D piston engine provides takeoff power equal to 190 hp.

Figure 6 shows the operational ranges of altitudes and steady-level flight speeds for the R22 and Hughes 269/300 helicopters. Maximum ground speeds are 190 km/h for the Robinson R22 and 175 km/h for the Hughes 269/300. It also shows the values of the economic speed Vek, which provides the mode of maximum steady climb.

The required values of the common pitch angle of the helicopter HB during descent in the self-rotation mode near the ground are shown in Fig.7. With these values of jc, the speed of rotation of the HB is kept constant.

5. Johnson W. Helicopter theory. Book 1. M.: Mir, 1983.

6. Braverman A.S. Helicopter quality and propulsion efficiency. Linearization of aerodynamic calculation // On the calculation of helicopter flight characteristics. Proceedings of TsAGI them. prof. N.E. Zhukovsky, issue 2448, 1989.

7. Statistical data of foreign helicopters / Reviews No. 678. TsAGI im. prof. N.E. Zhukovsky, M.: ONTI TsAGI, 1988.

8. Araslanov S. A. What helicopters does Ukraine need? // General Aviation, No. 10, 1999.

Introduction

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

At the stage of preliminary design, an aerodynamic scheme is selected, the appearance of the helicopter is formed, and the calculation of the main parameters is performed to ensure the achievement of the specified flight performance. These parameters include: the mass of the helicopter, the power of the propulsion system, the dimensions of the main and tail rotors, the mass of fuel, the mass of instrumentation and special equipment. The results of the calculations are used in the development of the layout scheme of the helicopter and the preparation of the balance sheet to determine the position of the center of mass.

The design of individual units and components of the helicopter, taking into account the selected technical solutions, is carried out at the stage of developing a technical project. At the same time, 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 are performed, as well as the choice of structural materials and design schemes.

At the detailed design stage, working and assembly drawings of the helicopter, specifications, packing lists and other technical documentation are prepared in accordance with 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 "Helicopter Design".

1. Calculation of the takeoff weight of a helicopter of the first approximation

where is the payload mass, kg;

Crew weight, kg.

Range of flight

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

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

where is the takeoff weight of the helicopter, kg;

g- free fall acceleration equal to 9.81 m/s 2 ;

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

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

We accept the radius of the main rotor equal to R= 7.9

Angular velocity w, s -1 , rotation of the main rotor is limited by the circumferential speed wR the ends of the blades, which depends on the takeoff weight of the helicopter and amounted to wR= 232 m/s.

2.2 Relative air densities on static and dynamic ceilings

2.3 Calculation of the economic speed near the ground and on the dynamic ceiling

The relative area of the equivalent harmful plate is determined:

Where Suh= 2.5

The value of the economic speed near the ground is calculated Vh, km/h:

where I

The value of the economic speed on the dynamic ceiling is calculated Vdin, km/h:

where I\u003d 1.09 ... 1.10 - induction coefficient.

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

where Vmax=250 km/h and Vdin\u003d 182.298 km / h - flight speed;

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

2.5 Calculation of the permissible ratios of the thrust coefficient to the filling of the main rotor for the maximum speed near the ground and for the economic speed on the dynamic ceiling:

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

2.7 Calculation of the filling of the main rotor:

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

As an estimated filling value s rotor, the largest value is taken from sVmax And sVdin:

chord length b and elongation l rotor blades will be equal to:

Where z l is the number of rotor blades (z l \u003d 3)

2.8 Relative increase in main rotor thrust to compensate for the aerodynamic drag of the fuselage and horizontal tail:

where S f is the area of the horizontal projection of the fuselage;

S th - the area of the horizontal tail.

S th \u003d 1.5 m 2.

3. Calculation of the power of the helicopter propulsion system.

3.1 Calculation of power when hovering on a static ceiling:

The specific power required to drive the main rotor in hover mode on a statistical ceiling is calculated by the formula:

where N Hst- required power, W;

m 0 - take-off weight, kg;

g- free fall acceleration, m/s 2 ;

p-specific load on the area swept by the main rotor, N/m 2 ;

D st- relative air density at the height of the static ceiling;

h 0 - relative efficiency. main rotor in hover mode ( h 0 =0.75);

Relative increase in main rotor thrust to balance the aerodynamic drag of the fuselage and horizontal tail:

3.2 Calculation of specific power in level flight at maximum speed

The specific power required to drive the main rotor in level flight at maximum speed is calculated by the formula:

where is the peripheral speed of the ends of the blades;

Relative equivalent harmful plate;

Iuh- induction coefficient, determined depending on the flight speed according to the following formulas:

At km/h,

At km/h

3.3 Calculation of specific power in flight at a dynamic ceiling with economic speed

The specific power to drive the main rotor on a dynamic ceiling is:

whereD din- relative air density on the dynamic ceiling,

Vdin- economic speed of the helicopter on the dynamic ceiling,

3.4 Calculation of specific power in flight near the ground at economic speed in the event of one engine failure during takeoff

The specific power required to continue takeoff at economic speed in the event of one engine failure is calculated by the formula:

where is the economic speed near the ground,

3.5 Calculation of specific reduced powers for various flight cases

3.5.1 The specific reduced power when hovering on a static ceiling is:

where is the specific throttle characteristic, which depends on the height of the static ceiling Hst and is calculated by the formula:

x 0 - power utilization factor of the propulsion system in the hover mode, the value of which depends on the takeoff weight of the helicopter m 0:

at m 0

at 10 25 tons

at m 0 > 25 tons

3.5.2 The specific reduced power in level flight at maximum speed is:

where is the power utilization factor at maximum flight speed,

Throttle characteristics of engines, depending on airspeed Vmax :

3.5.3 Specific reduced power in flight at dynamic ceiling at economic speed Vdin is equal to:

where is the power utilization factor at the economic flight speed,

and - engine throttling levels depending on the height of the dynamic ceiling H and flight speed Vdin according to the following throttle characteristics:

3.5.4 The specific reduced power in flight near the ground at economic speed with one engine failure on takeoff is equal to:

where is the power utilization factor at the economic flight speed,

The degree of engine throttling in emergency mode,

n=2 - number of helicopter engines.

3.5.5 Calculation of the required power of the propulsion system

To calculate the required power of the propulsion system, the maximum value of the specific reduced power is selected:

Required power N helicopter propulsion system will be equal to:

where m 01 - helicopter takeoff weight,

g\u003d 9.81 m 2 / s - free fall acceleration.

3.6 Choice of engines

We accept two turboshaft engines VK-2500 (TV3-117VMA-SB3) with a total power of each N\u003d 1.405 10 6 W

The VK-2500 (TV3-117VMA-SB3) engine is intended for installation on new generation helicopters, as well as for replacing engines on existing helicopters to improve their flight performance. It was created on the basis of a serial certified TV3-117VMA engine and is produced at the Federal State Unitary Enterprise "Plant named after V.Ya.Klimov".

4. Calculation of the mass of fuel

To calculate the mass of fuel that provides a given flight range, it is necessary to determine the cruising speed Vkr.The calculation of cruising speed is carried out by the method of successive approximations in the following sequence:

a) the value of the cruising speed of the first approximation is taken:

b) the induction coefficient is calculated Iuh:

at km/h

c) the specific power required to drive the main rotor in flight in cruising mode is determined:

where is the maximum value of the specific reduced power of the propulsion system,

Coefficient of change in power depending on the flight speed Vkr 1 , calculated by the formula:

d) The cruising speed of the second approximation is calculated:

e) The relative deviation of the speeds of the first and second approximation is determined:

The cruising speed of the first approximation is being refined Vkr 1 , it is taken equal to the calculated speed of the second approximation. Then the calculation is repeated from point b) and ends with the condition.

Specific fuel consumption is calculated by the formula:

where is the coefficient of change in the specific fuel consumption depending on the mode of operation of the engines,

Coefficient of change in specific fuel consumption depending on flight speed,

Specific fuel consumption in takeoff mode.

In the case of flight in cruise mode, the following is accepted:

kg/Wh,

The mass of fuel spent on the flight mT will be equal to:

where is the specific power consumed at cruising speed,

Cruising speed,

L- range of flight.

5. Determination of the mass of components and assemblies of the helicopter.

5.1 The mass of the main rotor blades is determined by the formula:

where R- rotor radius

s- filling of the main rotor,

5.2 The mass of the main rotor hub is calculated by the formula:

where kTue- weight coefficient of bushings of modern designs,

kl- coefficient of influence of the number of blades on the mass of the bushing.

You can take into account:

therefore, as a result of the transformations, we get:

To determine the mass of the main rotor hub, it is necessary to calculate the centrifugal force acting on the blades NCB(in kN):

5.3 The mass of the booster control system, which includes the swashplate, hydraulic boosters, the main rotor control hydraulic system, is calculated by the formula:

where b- blade chord,

kboo- weight factor of the booster control system, which can be taken equal to 13.2 kg/m 3 .

5.4 Mass of the manual control system:

where kRU-weight coefficient of the manual control system, taken for single-rotor helicopters equal to 25 kg/m.

5.5 The mass of the main gearbox depends on the torque on the main rotor shaft and is calculated by the formula:

where ked- weight coefficient, the average value of which is 0.0748 kg / (Nm) 0.8.

The maximum torque on the main rotor shaft is determined through the reduced power of the propulsion system N and screw speed w:

where x 0 - power utilization factor of the propulsion system, the value of which is taken depending on the takeoff weight of the helicopter m 0:

at m 0

at 10 25 tons

at m 0 > 25 tons

Mass of the main gearbox:

5.6 To determine the mass of the tail rotor drive units, its thrust is calculated Trv:

where Mnv- torque on the rotor shaft,

Lrv- the distance between the axes of the main and tail screws.

The distance between the axes of the main and tail screws is equal to the sum of their radii and clearance d between the ends of their blades:

where d- gap taken equal to 0.15 ... 0.2 m,

The radius of the tail rotor, which, depending on the takeoff weight of the helicopter, is:

Power Nrv, spent on the rotation of the tail rotor, is calculated by the formula:

where h 0 - relative efficiency of the tail rotor, which can be taken equal to 0.6 ... 0.65.

Torque Mrv transmitted by the steering shaft is equal to:

where is the frequency of rotation of the steering shaft,

Torque transmitted by the transmission shaft, N∙m, at a speed of rotation nin= 3000 rpm:

Weight min transmission shaft:

where kin- weighting factor for the transmission shaft, which is equal to 0.0318 kg / (Nm) 0.67.

Weight metc intermediate gear is equal to:

where ketc- weight factor for the intermediate gearbox, equal to 0.137 kg / (Nm) 0.8.

Weight of the tail gear that rotates the tail rotor:

where kxp- weighting factor for the tail gear, the value of which is 0.105 kg/(Nm) 0.8

5.7 The mass and main dimensions of the tail rotor are calculated depending on its thrust Trv.

Thrust coefficient Crv tail rotor is equal to:

Tail rotor blade filling srv calculated in the same way as for the main rotor:

where is the allowable value of the ratio of the thrust coefficient to the filling of the tail rotor.

chord length brv and elongation lrv tail rotor blades is calculated by the formulas:

where zrv- number of tail rotor blades.

Mass of tail rotor blades mLR

The value of centrifugal force Ncbr acting on the tail rotor blades and perceived by the bushing hinges,

Tail rotor hub weight mtuesday calculated using the same formula as for the main rotor:

where NCB- centrifugal force acting on the blade,

kTue- weight coefficient for the sleeve, taken equal to 0.0527 kg/kN 1.35

kz- weight coefficient depending on the number of blades and calculated by the formula:

5.8 Calculation of the mass of the helicopter propulsion system

Specific gravity of the helicopter propulsion system gdv calculated by the empirical formula:

where N- power of the propulsion system.

The mass of the propulsion system will be equal to:

5.9 Calculation of the mass of the fuselage and equipment of the helicopter

The mass of the helicopter fuselage is calculated by the formula:

where Sohm- the area of the washed surface of the fuselage, which is determined by the formula:

m 0 - takeoff weight of the first approximation,

kf-coefficient equal to 1.7.

Fuel system weight:

where mT- the mass of fuel used for the flight,

kts- weight coefficient taken for the fuel system equal to 0.09.

The mass of the helicopter landing gear is:

where kw-weighting factor depending on the chassis design:

For fixed landing gear,

For retractable landing gear.

The mass of the electrical equipment of the helicopter is calculated by the formula:

where Lrv- the distance between the axes of the main and tail screws,

zl- number of rotor blades,

R- rotor radius,

ll- relative elongation of the main rotor blades,

ketc And kemail- weight coefficients for electrical wires and other electrical equipment, the values of which are equal to:

Mass of other helicopter equipment:

where ketc-weight coefficient, the value of which is equal to 2.

5.10 Calculation of the second approximation helicopter takeoff mass

The mass of an empty helicopter is equal to the sum of the masses of the main units:

Takeoff weight of the helicopter of the second approximation m 02 will be equal to the sum:

where mT- mass of fuel,

mgr- mass of payload,

meq- mass of the crew.

6. Description of the layout of the helicopter

The designed helicopter is made according to a single-rotor scheme with a tail rotor, two gas turbine engines and two-bearing skis. The fuselage of the frame structure helicopter consists of the nose and central parts, tail and end beams. In the bow there is a two-seat crew cabin, consisting of two pilots. Cabin glazing provides good visibility, right and left sliding blisters are equipped with emergency release mechanisms. In the central part there is a cabin measuring 6.8 x 2.05 x 1.7m, and a central sliding door measuring 0.62 x 1.4m with an emergency drop mechanism. The cargo cabin is designed for the carriage of goods weighing up to 2 tons and is equipped with folding seats for 12 passengers, as well as nodes for attaching 5 stretchers. In the passenger version, there are 12 seats in the cabin, installed with a step of 0.5m and a passage of 0.25m; and in the back there is an opening for the rear entrance door, consisting of two wings.

The tail boom of riveted construction of beam-stringer type with a working skin is equipped with nodes for attaching a controlled stabilizer and a tail support.

Stabilizer with a size of 2.2 m and an area of 1.5 m 2 with a NACA 0012 profile of a single-spar design, with a set of ribs and duralumin and fabric sheathing.

Double-support, skis, self-orienting front support, dimensions 500 x 185 mm, main supports of a shaped type with liquid-gas two-chamber shock absorbers, dimensions 865 x 280 mm. The tail support consists of two struts, a shock absorber and a support heel; ski track 2m, ski base 3.5m.

The tail rotor with a diameter of 1.44m is three-bladed, pusher, with a cardan-type sleeve and all-metal rectangular-shaped blades in plan, with a chord of 0.51m and a NACA 230M profile.

Equipment for instrument flights in difficult meteorological conditions day and night includes two artificial horizons, two NV speed indicators, a GMK-1A combined heading system, an automatic radio compass, and a RV-3 radio altimeter.

The communication equipment includes R-860 and R-828 VHF command radio stations, R-842 and Karat communication HF radio stations, SPU-7 aircraft intercom.

7. Helicopter balance calculation

Table 1. Balancing list of an empty helicopter

Unit name	unit weight, m i, kg	Coordinate x i center of mass of the unit, m	Static moment of the unit M xi	Coordinate y i center of mass of the unit, m	Static moment of the unit M yi
1 Rotor
1.1 Blades
1.2 Sleeve
2 Control system
2.1 Booster control system
2.2 Manual control system
3 Transmission
3.1 Main gearbox
3.2 Intermediate gearbox
3.3 Tail gear
3.4 Transmission shaft
4 Tail screw
4.1 Blades
4.2 Sleeve
5 Propulsion system
6 Fuel system
7 Fuselage
7.1 Bow (15%)
7.2 Middle part (50%)
7.3 Tail section (20%)
7.4 Fixing the gearbox (4%)
7.5 Hoods (11%)

8.1 Main (82%)
8.2 Front (16%)
8.3 Tail support (2%)
9 Electrical equipment
10 Equipment
10.1 Instruments in the cockpit (25%)
10.2 Radio equipment (27%)
10.3 Hydraulic equipment (20%)
10.4 Pneumatic equipment (6%)

Static moments are calculated M cxi And M sui relative to the coordinate axes:

The coordinates of the center of mass of the entire helicopter are calculated by the formulas:

Table 2. Centering list with maximum load

Unit name	unit weight, m i, kg	Coordinate x i center of mass of the unit, m	Static moment of the unit M xi	Coordinate y i center of mass of the unit, m	Static moment of the unit M yi
Helicopter
Fuel tanks 1 and 2

Table 3. Centering list with 5% remaining fuel and full commercial load

Unit name	unit weight, m i, kg	Coordinate x i center of mass of the unit, m	Static moment of the unit M xi	Coordinate y i center of mass of the unit, m	Static moment of the unit M yi
Helicopter

Empty helicopter center of mass coordinates: x 0 =-0.003; y 0 =-1.4524;

The coordinates of the center of mass with the maximum load: x 0 \u003d 0.0293; y 0 \u003d -2.0135;

Center of mass coordinates with 5% remaining fuel and full commercial load: x 0 \u003d -0.0678; y 0 = -1,7709.

Conclusion

In this course project, calculations of the take-off weight of the helicopter, the mass of its components and assemblies, as well as the layout of the helicopter were carried out. During the layout process, the alignment of the helicopter was clarified, the calculation of which is preceded by the preparation of a weight report based on the weight calculations of the units and the power plant, lists of equipment, equipment, cargo, etc. The design goal is to determine the optimal combination of the main parameters of the helicopter and its systems that ensure the fulfillment of specified requirements.

Introduction

At the stage of preliminary design, an aerodynamic scheme is selected, the appearance of the helicopter is formed, and the calculation of the main parameters is performed to ensure the achievement of the specified flight performance. These parameters include: the mass of the helicopter, the power of the propulsion system, the dimensions of the main and tail rotors, the mass of fuel, the mass of instrumentation and special equipment. The results of the calculations are used in the development of the layout scheme of the helicopter and the preparation of the balance sheet to determine the position of the center of mass.

The design of individual units and components of the helicopter, taking into account the selected technical solutions, is carried out at the stage of developing a technical project. At the same time, 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 are performed, as well as the choice of structural materials and structural schemes.

At the detailed design stage, working and assembly drawings of the helicopter, specifications, packing lists and other technical documentation are prepared in accordance with 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 "Helicopter Design".

1. Calculation of the takeoff weight of a helicopter of the first approximation

where is the payload mass, kg;

Crew weight, kg.

Range of flight

kg.

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

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

where is the takeoff weight of the helicopter, kg;

g - free fall acceleration, equal to 9.81 m / s 2;

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

The value of the specific load p on the area swept by the propeller is selected according to the recommendations presented in the work /1/: where p=280

We take the rotor radius equal to R=7.9

The angular velocity w, s -1 , of rotation of the main rotor is limited by the circumferential velocity wR of the ends of the blades, which depends on the take-off mass of the helicopter and amounted to wR=232 m/s.

with -1 .

rpm

2.2 Relative air densities on static and dynamic ceilings

2.3 Calculation of the economic speed near the ground and on the dynamic ceiling

The relative area of the equivalent harmful plate is determined:

Where S e \u003d 2.5

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

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

where I \u003d 1.09 ... 1.10 is the induction coefficient.

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

where V max \u003d 250 km / h and V dyn \u003d 182.298 km / h - flight speeds;

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

2.5 Calculation of the allowable ratios of the thrust coefficient to the filling of the main rotor for the maximum speed near the ground and for the economic speed on the dynamic ceiling:

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

2.7 Calculation of the filling of the main rotor:

The main rotor filling s is calculated for the cases of flight at maximum and economic speeds:

;

As the calculated filling value s of the main rotor, the largest value of s Vmax and s V dyn is taken:

The chord length b and relative elongation l of the main rotor blades will be equal to:

Where z l is the number of rotor blades (z l \u003d 3)

2.8 Relative increase in main rotor thrust to compensate for the aerodynamic drag of the fuselage and horizontal tail:

where S f is the area of the horizontal projection of the fuselage;

S th - the area of the horizontal tail.

S th \u003d 1.5 m 2.

To perform a combat mission and ensure flight safety, the design of the helicopter must be sufficiently strong and rigid. By strength they mean the ability of a structure to perceive, without collapsing, the given external loads encountered during operation. Rigidity refers to the ability of a structure to resist deformation under load.

During operation, the helicopter is subjected to loads of various nature and magnitude: static (constant or slowly changing over time), dynamic (shock and vibration). Depending on the type of loading, the structure or its separate part must have the appropriate type of strength.

Combination of required values various kinds strength, providing normal work structures within the established limits and deadlines, is called operational strength.

During operation, the strength of the structure does not remain unchanged. Large loads, close to the limit ones, can cause permanent deformations in its elements. Small, but repeatedly repeated loads cause the development of fatigue cracks that weaken the structure. Wear and tear occur

rubbing parts, abrasive wear of HB blades, blades gas turbine engines under the influence of dust, sand. In addition, at maintenance damage is introduced in the form of dents, scratches, scratches, nicks, etc. All this leads to a gradual decrease in structural strength and forces the helicopter to limit the resource (flying hours) of the helicopter.

During operation, the structure is constantly affected by temperature changes, precipitation, dust, solar radiation, etc. The impact of these factors causes corrosion of structural elements, cracking of glazing and other non-metallic parts, and damage to protective coatings. As a result, it is necessary to limit the calendar time of equipment operation (service life).

Thus, all the above external factors that reduce the strength and degrade the performance of the structure, limit its durability. The durability of an aircraft is the ability to maintain operability, taking into account maintenance and repair, up to a certain limiting state, at which flight safety requirements are violated, and operational efficiency is reduced. The indicators of durability are the resource and service life.

One of the main tasks of the technical operation of aviation equipment is to maintain the required strength during the entire service life under real operating conditions.

General principles for calculating the strength of a helicopter

The Strength Standards also provide for: the effect of negative G = -0.5 when entering into planning, energetic turns of the helicopter in hover, the effect of vertical and lateral gusts of air, etc. Each of the design cases is decisive for the strength of one or another part or unit of the helicopter.

Landing design cases are considered various options landings: on all supports, only on the main ones, landing with a side impact, etc.

Ground design cases consider the effect of wind, helicopter towing on an unprepared site, etc.

The particular difficulty in calculating the strength of a helicopter lies in the fact that its main loads, for example, forces from the HB blades, are variable in magnitude and direction, which causes oscillations of the blades themselves and the helicopter structure as a whole. Such loading is called dynamic. With prolonged action of repeatedly repeated loads, the destruction of the structure occurs at stresses that are much lower than with a constant, static load. This is due to the phenomenon of material fatigue.

The Strength Standards also provide all the necessary data for calculating the rigidity of the structure, its dynamic strength and resource (service life).

The concept of calculating static strength

If the load of the structure is constant or changes slowly, then the deformations and stresses in it will also be constant or change gradually, in proportion to the load, without oscillatory processes. Such loading is called static.

For a helicopter, static loads can be considered: the thrust of the main and tail propellers; centrifugal forces of the blades; aerodynamic forces of the wing and tail.

The calculation for static strength includes:

- determination, in accordance with the Strength Standards, of the magnitude and nature of the distribution of design loads;
- construction of diagrams of transverse Q and longitudinal N forces, bending and torque moments for the considered part of the helicopter structure;
- identification of the most loaded sections of the structure, in which the greatest stresses are possible;
- determination of stresses in structural elements and their comparison with destructive ones.

The static strength of the structure is ensured if the stresses in its elements do not exceed destructive values.

However, static strength does not guarantee safe operation helicopter, since under the action of variable loads in its structure, corresponding alternating stresses arise. These stresses, superimposed on the constant ones, increase the total stresses, and can also lead to fatigue failure of the structure.

Helicopter Variable Load Sources

The main loads of the helicopter are variable in nature, they constantly change in magnitude and direction with certain frequencies.

The main sources of variable loads are the main and tail screws. The reason for the periodic change in the forces acting on the HB blades is the continuous change in the speed and direction of the flow on them in different azimuths and in different sections during the translational flight of the helicopter. When the blade, during its rotation, moves towards the incoming flow on the helicopter, the total speed of its flow increases, and when moving backward, on the contrary, it decreases. Since the aerodynamic forces are proportional to the square of the flow velocity, the lift force Ul and the drag Xl of the blade also change constantly. This causes the blades to flap in the vertical plane and oscillate in the plane of rotation.

During the flywheel motion, the centers of mass of the blades periodically approach and move away from the axis of the screw, which causes the appearance of variable Coriolis forces acting in the plane of rotation. These forces also cause the blades to oscillate in the plane of rotation.

All these variable forces are transmitted to the HB bushing and further through the propeller shaft and gearbox to the helicopter fuselage, causing it to oscillate in the vertical and horizontal planes. The amplitudes of the variable forces transmitted from the blades can be thousands of newtons, and for heavy helicopters, tens of thousands. The frequencies of these forces are a multiple of the product of the propeller speed and the number of blades.

Additional sources of variable forces can be poor balancing and misalignment of the blades. Poor balancing consists in unequal static moments of the blades, which causes an imbalance in their centrifugal forces. Misconicity manifests itself in different amplitudes of the flapping motion of the blades due to differences in their external shapes, torsional rigidity, or inaccurate adjustment of installation angles. For the same reasons, variable tail rotor forces arise.

Calculation of the takeoff weight of a helicopter of the first approximation. To the calculation of the flight characteristics of a helicopter at the design stage Calculation of the main parameters and development of the layout of the helicopter

General principles for calculating the strength of a helicopter

The concept of calculating static strength

Helicopter Variable Load Sources

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