Calculation of the turbine of a turbojet bypass engine based on AL-31F. Design of an axial turbine for an aircraft engine JT9D20 Optimum number of CV blades

03.03.2020

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1. Design description

turbine engine strength power

1.1 AL-31F

AL-31F is a dual-circuit twin-shaft turbojet engine with mixing flows of the internal and external circuits behind the turbine, an afterburner common for both circuits and an adjustable supersonic all-mode jet nozzle. Low-pressure axial 3-stage compressor with adjustable inlet guide vane (VNA), high-pressure axial 7-stage compressor with adjustable VNA and guide vanes of the first two stages. Turbines of high and low pressure - axial single-stage; blades of turbines and nozzle devices are cooled. The main combustion chamber is annular. Titanium alloys (up to 35% of the mass) and heat-resistant steels are widely used in the engine design.

1.2 Turbine

General characteristics

The engine turbine is axial, jet, two-stage, two-shaft. The first stage is a high pressure turbine. The second stage is low pressure. All turbine blades and disks are cooled.

The main parameters (H=0, M=0, "Maximum" mode) and materials of the turbine parts are given in Tables 1.1 and 1.2.

Table 1.1

Parameter
The degree of reduction of the total gas pressure
Turbine efficiency in terms of stagnant flow parameters
Circumferential speed at the periphery of the blades, m/s
Rotor speed, rpm
Sleeve ratio
Gas temperature at the turbine inlet
Gas consumption, kg/s
Load parameter, m/s

Table 1.2

High pressure turbine design

The high-pressure turbine is designed to drive the high-pressure compressor, as well as propulsion and aircraft units mounted on the gearboxes. The turbine structurally consists of a rotor and a stator.

High pressure turbine rotor

The turbine rotor consists of rotor blades, disk and trunnion.

The working blade is cast, hollow with a semi-loop flow of cooling air.

In the inner cavity, in order to organize the flow of cooling air, ribs, partitions and turbulators are provided.

In subsequent series, the blade with a half-loop cooling scheme is replaced by a blade with a cyclone-vortex cooling scheme.

A channel is made in the inner cavity along the leading edge, in which, as in a cyclone, an air flow with a swirl is formed. The swirling of air occurs due to its tangential supply to the channel through the openings of the baffle.

From the channel, air is ejected through the holes (perforation) of the blade wall onto the back of the blade. This air creates a protective film on the surface.

In the central part of the blade on the inner surfaces there are channels, the axes of which intersect. A turbulent air flow is formed in the channels. Air jet turbulence and an increase in the contact area provide an increase in heat transfer efficiency.

Turbulators (bridges) of various shapes are made in the region of the trailing edge. These turbulators intensify heat transfer and increase blade strength.

The profile part of the blade is separated from the lock by a shelf and an elongated leg. The shelves of the blades, docking, form a conical shell that protects the locking part of the blade from overheating.

An elongated leg, ensuring the distance of the high-temperature gas flow from the lock and disk, leads to a decrease in the amount of heat transferred from the profile part to the lock and disk. In addition, the elongated stem, having a relatively low bending stiffness, reduces the level of vibration stresses in the profile part of the blade.

A three-prong herringbone lock ensures the transfer of radial loads from the blades to the disk.

The tooth, made in the left part of the lock, fixes the blade from moving it along the flow, and the groove, together with the fixation elements, ensures that the blade is kept from moving against the flow.

On the peripheral part of the blade, in order to facilitate running-in when touching the stator and, consequently, to prevent the destruction of the blade, a sample was made on its end

To reduce the level of vibration stresses in the working blades, dampers with a box-shaped design are placed between them under the shelves. When the rotor rotates under the action of centrifugal forces, the dampers are pressed against the inner surfaces of the shelves of the vibrating blades. Due to friction at the points of contact of two adjacent flanges on one damper, the energy of blade vibrations will be dissipated, which ensures a decrease in the level of vibration stresses in the blades.

The turbine disc is stamped, followed by machining. In the peripheral part of the disc there are grooves of the “Herringbone” type for fastening 90 rotor blades, grooves for accommodating plate locks for axial fixation of the blades and inclined holes for supplying air that cools the rotor blades.

The air is taken from the receiver formed by two flanges, the left side surface of the disk and the swirler. Balancing weights are placed under the lower shoulder. On the right plane of the disk web there is a shoulder of the labyrinth seal and a shoulder used when dismantling the disk. Cylindrical holes are made on the stepped part of the disk for fitting bolts connecting the shaft, disk and turbine rotor pin.

Axial fixation of the working blade is carried out by a tooth with a lamellar lock. A lamellar lock (one for two blades) is inserted into the grooves of the blades in three places of the disk, where cutouts are made, and accelerates around the entire circumference of the blade ring. Lamellar locks, installed at the location of the cutouts in the disk, have a special shape. These locks are mounted in a deformed state, and after straightening they enter the grooves of the blades. When straightening the lamellar lock, the blades are supported from opposite ends.

The rotor is balanced by weights fixed in the groove of the disk shoulder and fixed in the lock. The tail of the lock is bent on a balancing weight. The place of the bend is controlled for the absence of cracks by inspection through a magnifying glass. The rotor can be balanced by moving the blades, trimming the ends of the weights is allowed. Residual imbalance is not more than 25 gcm.

The disk with the trunnion and the HPC shaft is connected by fitting bolts. The heads of the bolts are fixed against rotation by plates bent on the cuts of the heads. The bolts are kept from longitudinal movement by the protruding parts of the heads included in the annular groove of the shaft.

The trunnion provides support for the rotor on a roller bearing (inter-rotor bearing).

The trunnion flange is centered and connected to the turbine disk. On the outer cylindrical grooves of the trunnion, the sleeve of the labyrinth seals is placed. Axial and circumferential fixation of labyrinths is carried out by radial pins. To prevent the pins from falling out under the influence of centrifugal forces, after they are pressed in, the holes in the bushings are flared.

On the outer part of the pin shank, below the labyrinths, there is a contact seal fixed with a castellated nut. The nut is locked with a plate lock.

Inside the trunnion in cylindrical belts, the bushings of the contact and labyrinth seals are centered. The bushings are held by a castellated nut screwed into the trunnion threads. The nut is locked by bending the antennae of the crown into the end slots of the trunnion.

In the right part of the internal cavity of the trunnion, the outer ring of the roller bearing is located, which is held by a castellated nut screwed into the trunnion thread, which is locked in the same way.

The contact seal is a pair of steel bushings and graphite rings. Flat springs are placed between graphite rings for guaranteed contact of pairs. Between the steel bushings, a spacer bushing is placed to prevent pinching of the mechanical contact seal.

High pressure turbine stator

The high-pressure turbine stator consists of an outer ring, nozzle vane blocks, an inner ring, a swirling device, and a seal with HPT inserts.

The outer ring is a cylindrical shell with a flange. The ring is located between the combustion chamber housing and the LPT housing.

A groove is made in the middle part of the outer ring, along which the dividing wall of the heat exchanger is centered.

In the left part of the outer ring, the upper ring is attached to the screws, which is the support of the flame tube of the combustion chamber and provides the supply of cooling air to blow the outer shelves of the blades of the nozzle apparatus.

A seal is installed on the right side of the outer ring. The seal consists of an annular spacer with screens, 36 HPT sector inserts and sectors for attaching HPT inserts to the spacer.

Annular threading is made on the inner diameter of the HPT inserts to reduce the surface area when the HPT rotor blades touch to prevent overheating of the peripheral part of the rotor blades.

The seal is attached to the outer ring with drilled pins. Through these drillings, cooling air is supplied to the HPT inserts.

Through the holes in the inserts, the cooling air is ejected into the radial gap between the inserts and the rotor blades.

Plates are installed between the inserts to reduce the flow of hot gas.

When assembling the seal, the HPT inserts are attached to the spacer in sectors using pins. This fastening allows the HPT inserts to move relative to each other and spacers when heated during operation.

The blades of the nozzle apparatus are combined into 14 three-bladed blocks. The blade blocks are cast, with deflectors plugged in and soldered in two places with a soldered bottom cover with a trunnion. The cast construction of blocks, having high rigidity, ensures the stability of the angles of installation of the blades, the reduction of air leakage and, consequently, the increase in the efficiency of the turbine, in addition, such a design is more technologically advanced.

The internal cavity of the scapula is divided into two compartments by a partition. In each compartment there are deflectors with holes that provide a jet flow of cooling air onto the inner walls of the blade. The leading edges of the blades are perforated.

In the upper shelf of the block, there are 6 threaded holes, into which the screws for fastening the blocks of nozzle devices to the outer ring are screwed.

The lower shelf of each block of blades has a trunnion, along which the inner ring is centered through the bushing.

The profile of the pen with adjacent surfaces of the shelves is aluminosilicated. Coating thickness 0.02-0.08 mm.

To reduce the flow of gas between the blocks, their joints are sealed with plates inserted into the slots of the ends of the blocks. The grooves in the ends of the blocks are made by electroerosive method.

The inner ring is made in the form of a shell with bushings and flanges, to which a conical diaphragm is welded.

On the left flange of the inner ring, a ring is attached with screws, on which the flame tube rests and through which air is supplied, blowing the inner shelves of the blades of the nozzle apparatus.

In the right flange, the swirling apparatus is fixed with screws, which is a welded shell structure. The swirling device is designed to supply and cool the air going to the rotor blades due to acceleration and swirling in the direction of turbine rotation. To increase the rigidity of the inner shell, three reinforcing profiles are welded to it.

The acceleration and swirl of the cooling air take place in the converging part of the swirl apparatus.

Air acceleration provides a decrease in the temperature of the air used to cool the rotor blades.

The swirl of the air ensures the alignment of the circumferential component of the air velocity and the circumferential speed of the disk.

Low pressure turbine design

The low-pressure turbine (LPT) is designed to drive the low-pressure compressor (LPC). Structurally, it consists of a LPT rotor, LPT stator and LPT support.

Low pressure turbine rotor

The low-pressure turbine rotor consists of a LPT disk with working blades fixed on the disk, a pressure disk, a trunnion and a shaft.

The working blade is cast, cooled with a radial flow of cooling air.

In the inner cavity there are 11 rows of 5 pieces each of cylindrical pins - turbulators connecting the back and trough of the blade.

The peripheral shroud reduces the radial clearance, which leads to an increase in the efficiency of the turbine.

Due to the friction of the contact surfaces of the shroud shelves of adjacent rotor blades, the level of vibration stresses decreases.

The profile part of the blade is separated from the locking part by a shelf that forms the boundary of the gas flow and protects the disk from overheating.

The blade has a herringbone-type lock.

The casting of the blade is carried out according to investment models with surface modification with cobalt aluminate, which improves the structure of the material by grinding grains due to the formation of crystallization centers on the surface of the blade.

In order to increase heat resistance, the outer surfaces of the feather, shroud and lock shelves are subjected to slip aluminosicillation with a coating thickness of 0.02-0.04.

For axial fixation of the blades from moving against the flow, a tooth is made on it, abutting against the disk rim.

For axial fixation of the blade from moving along the flow, a groove is made in the locking part of the blade in the region of the flange, into which a split ring with a lock is inserted, which is kept from axial movement by the disc shoulder. During installation, the ring, due to the presence of a cutout, is crimped and inserted into the grooves of the blades, and the shoulder of the disk enters the groove of the ring.

The fastening of the split ring in working condition is made by a lock with clamps that are bent onto the lock and pass through the holes in the lock and the slots in the shoulder of the disk.

Turbine disk - stamped, with subsequent machining. In the peripheral zone for placing the blades there are grooves of the "Herringbone" type and inclined holes for supplying cooling air.

Annular flanges are made on the disc web, on which labyrinth covers and a pressure labyrinth disc are placed. The fixation of these parts is carried out with pins. To prevent the pins from falling out, the holes are flared.

A pressure disk having blades is needed to compress the air supplied to cool the turbine blades. To balance the rotor, balancing weights are fixed on the pressure disk with lamellar clamps.

Annular collars are also made on the disc hub. Labyrinth covers are installed on the left shoulder, a trunnion is installed on the right shoulder.

The trunnion is designed to support the low-pressure rotor on a roller bearing and transmit torque from the disk to the shaft.

To connect the disk to the trunnion, a forked flange is made on it in the peripheral part, along which centering is carried out. In addition, the centering and transfer of loads go through radial pins, which are kept from falling out by the labyrinth.

A labyrinth seal ring is also fixed on the LPT trunnion.

On the peripheral cylindrical part of the trunnion, a mechanical contact seal is placed on the right, and a sleeve of a radial-face contact seal is placed on the left. The bushing is centered along the cylindrical part of the trunnion and is fixed in the axial direction by the bending of the comb.

In the left part of the trunnion on the cylindrical surface there are bushings for supplying oil to the bearing, the inner ring of the bearing and sealing parts. The package of these parts is tightened with a castellated nut, locked with a lamellar lock. Splines are made on the inner surface of the trunnion to ensure the transmission of torque from the trunnion to the shaft. In the body of the trunnion there are holes for supplying oil to the bearings.

In the right part of the trunnion, on the outer groove, the inner ring of the roller bearing of the turbine support is fixed with a nut. The castellated nut is locked with a plate lock.

The low pressure turbine shaft consists of 3 parts connected to each other by radial pins. The right part of the shaft with its splines enters the reciprocal splines of the trunnion, receiving torque from it.

Axial forces from the pin to the shaft are transmitted by a nut screwed onto the threaded shaft shank. The nut is secured against loosening by a splined bushing. The end splines of the bushing fit into the end slots of the shaft, and the splines on the cylindrical part of the bushing fit into the longitudinal splines of the nut. In the axial direction, the splined bushing is fixed by adjusting and split rings.

On the outer surface of the right side of the shaft, a labyrinth is fixed with radial pins. On the inner surface of the shaft, a splined bushing of the drive of the oil pumping pump from the turbine support is fixed with radial pins.

On the left side of the shaft, splines are made that transmit torque to the spring and then to the low-pressure compressor rotor. On the inner surface of the left side of the shaft, a thread is cut into which a nut is screwed, locked with an axial pin. A bolt is screwed into the nut, which tightens the low-pressure compressor rotor and the low-pressure turbine rotor.

On the outer surface of the left side of the shaft there is a radial-face contact seal, a spacer bushing and a bevel gear roller bearing. All these parts are tightened with a castellated nut.

The composite design of the shaft allows to increase its rigidity due to the increased diameter of the middle part, as well as to reduce weight - the middle part of the shaft is made of titanium alloy.

Low pressure turbine stator

The stator consists of an outer housing, blocks of nozzle blades, and an inner housing.

The outer housing is a welded structure consisting of a conical shell and flanges, along which the housing is joined to the high-pressure turbine housing and the support housing. Outside, a screen is welded to the body, forming a channel for supplying cooling air. Inside there are flanges along which the nozzle apparatus is centered.

In the area of the right flange there is a bead on which LPT inserts with honeycombs are installed and fixed with radial pins.

The blades of the nozzle apparatus in order to increase the rigidity in eleven three-blade blocks.

Each blade is cast, hollow, cooled with internal deflectors. Feather, outer and inner shelves form the flow part. The outer shelves of the blades have flanges, with which they are centered along the grooves of the outer casing.

Axial fixation of blocks of nozzle blades is carried out by a split ring. The peripheral fixation of the blades is carried out by the protrusions of the body, which are included in the slots made in the outer shelves.

The outer surface of the shelves and the profile part of the blades is aluminosicillated in order to increase the heat resistance. The thickness of the protective layer is 0.02-0.08 mm.

To reduce the flow of gas between the blocks of blades, sealing plates are installed in the slots.

The inner shelves of the blades end with spherical pins, along which the inner casing is centered, representing a welded structure.

Grooves are made in the ribs of the inner body, which enter the scallops of the inner shelves of the nozzle blades with a radial clearance. This radial clearance provides freedom for the thermal expansion of the blades.

Turbine support ND

The turbine support consists of a support housing and bearing housing.

The support body is a welded structure consisting of shells connected by posts. Racks and shells are protected from the gas flow by riveted screens. On the flanges of the inner shell of the support, conical diaphragms are fixed, supporting the bearing housing. On these flanges, a labyrinth seal bushing is fixed on the left, and a screen protecting the support from the gas flow is fixed on the right.

On the flanges of the bearing housing, a contact seal bushing is fixed on the left. On the right, the oil cavity cover and the heat shield are fixed with screws.

A roller bearing is placed in the inner bore of the housing. Between the housing and the outer ring of the bearing there are an elastic ring and bushings. Radial holes are made in the ring, through which oil is pumped during vibrations of the rotors, to which energy is dissipated.

Axial fixation of the rings is carried out by a cover, attracted to the bearing support by screws. In the cavity under the heat shield there is an oil extraction pump and oil nozzles with pipelines. The bearing housing has holes that supply oil to the damper and nozzles.

Turbine cooling

Turbine cooling system - air, open, regulated by discrete changes in air flow through the air-to-air heat exchanger.

The leading edges of the blades of the nozzle apparatus of the high-pressure turbine have convective-film cooling with secondary air. The shelves of this nozzle apparatus are cooled by secondary air.

The rear strips of the SA blades, the disk and rotor blades of the LPT, the turbine housings, the SA blades of the fan turbine and its disk on the left side are cooled by air passing through the air-to-air heat exchanger (VHT).

The secondary air enters the heat exchanger through the holes in the combustion chamber housing, where it is cooled by - 150-220 K and goes through the valve apparatus to cool the turbine parts.

The air of the secondary circuit through the support legs and holes is supplied to the pressure disk, which, by increasing the pressure, ensures its supply to the working blades of the LPT.

The turbine housing is cooled from the outside by the secondary air, and from the inside by the air from the IWT.

Turbine cooling is carried out in all engine operating modes. The turbine cooling circuit is shown in Figure 1.1.

Power flows in the turbine

Inertial forces from rotor blades through locks of the "Herringbone" type are transferred to the disk and load it. The unbalanced inertial forces of the bladed discs are transmitted through the fit bolts on the HPT rotor and through the centering collars and radial pins on the HPT rotor to the shaft and pins supported by bearings. Radial loads are transferred from the bearings to the stator parts.

The axial components of the gas forces arising on the working blades of the HPT are transferred to the disk due to the friction forces on the contact surfaces in the lock and the “tooth” of the blade against the disk. On the disc, these forces are summed up with the axial forces arising from the pressure drop across it and are transferred to the shaft through tight bolts. Fitted bolts from this force work in tension. The axial force of the turbine rotor is added to the axial force.

Outer contour

The outer circuit is designed to bypass a part of the air flow compressed in the LPC behind the LPC.

Structurally, the outer contour consists of two (front and rear) profiled cases, which are the outer shell of the product and are also used for fastening communications and units. The shells of the outer case are made of titanium alloy. The case is included in the power circuit of the product, perceives the torque of the rotors and partly the weight of the internal circuit, as well as the overload forces during the evolution of the object.

The front casing of the outer circuit has a horizontal connector to provide access to HPC, CS and turbine.

The profiling of the flow path of the outer contour is ensured by the installation in the front casing of the outer contour of the inner screen, connected with it by radial stringers, which are also stiffening ribs of the front casing.

The rear casing of the outer contour is a cylindrical shell bounded by the front and rear flanges. On the rear housing from the outside there are stiffening stringers. Flanges are located on the housings of the outer housing:

· To take air from the internal circuit of the product after 4 and 7 stages of HPC, as well as from the channel of the external circuit for the needs of the facility;

· For igniters KS;

· For HPC blades inspection windows, CS inspection windows and turbine inspection windows;

· For communications of a supply and removal of oil to a support of the turbine, venting of an air and oil cavity of a back support;

· Air bleed into jet nozzle (RS) pneumatic cylinders;

· For fixing the feedback lever of the control system ON HPC;

· For communications for supplying fuel to the CS, as well as for communications for bleeding air after HPC into the fuel system of the product.

Bosses for fastening are also designed on the body of the outer contour:

· Fuel distributor; fuel-oil heat exchangers of the oil tank;

· Fuel filter;

· KND automation reducer;

· Drain tank;

· Ignition unit, communications of systems of start of FC;

· Frames with attachment points for the nozzle and afterburner regulator (RSF).

In the flow part of the outer circuit, two-hinged communication elements of the product system are installed, which compensate for thermal expansion in the axial direction of the bodies of the outer and inner circuits during the operation of the product. The expansion of the housings in the radial direction is compensated by the mixing of two-hinged elements, structurally made according to the "piston-cylinder" scheme.

2. Calculation of the strength of the turbine impeller disk

2.1 Calculation scheme and initial data

A graphical representation of the HPT impeller disk and the calculation model of the disk are shown in Fig. 2.1. The geometric dimensions are presented in Table 2.1. A detailed calculation is presented in Appendix 1.

Table 2.1

Section i

n - the number of revolutions of the disk in the design mode is 12430 rpm. The disc is made of EP742-ID material. The temperature along the radius of the disk is not constant. - blade (contour) load, simulating the action of the centrifugal forces of the blades and their interlocks (blade roots and disk protrusions) on the disk in the design mode.

Characteristics of the disc material (density, modulus of elasticity, Poisson's ratio, coefficient of linear expansion, long-term strength). When entering the characteristics of materials, it is recommended to use ready-made data from the archive of materials included in the program.

The contour load is calculated according to the formula:

The sum of the centrifugal forces of the feathers of the blades,

The sum of the centrifugal forces of the interlocks (blade roots and disk protrusions),

The area of the peripheral cylindrical surface of the disk through which centrifugal forces are transmitted to the disk and:

Forces calculated by the formulas

z- number of blades,

The area of the root section of the blade feather,

Stress in the root section of the blade feather, created by centrifugal forces. The calculation of this voltage was made in Section 2.

The mass of the ring formed by the locking connections of the blades with the disk,

The radius of inertia of the locking ring,

u - angular speed of rotation of the disk in the design mode, calculated through revolutions as follows: ,

The mass of the ring and the radius are calculated by the formulas:

The area of the peripheral cylindrical surface of the disk is calculated by formula 4.2.

Substituting the initial data into the formula for the above parameters, we get:

Calculation of the disc strength is made by the program DI.EXE, available in the computer class 203 of the department.

It should be borne in mind that the geometric dimensions of the disk (radii and thicknesses) are entered into the DI.EXE program in centimeters, and the contour load - in (translation).

2.2 Calculation results

The calculation results are presented in Table 2.2.

Table 2.2

The first columns of Table 2.2 present the initial data on the disk geometry and temperature distribution along the disk radius. Columns 5-9 present the results of the calculation: radial (radial) and circumferential (circumferential) stresses, reserves for equivalent stress (ex. equiv.) and breaking revolutions (cyl. sec.), as well as disk elongation under the action of centrifugal forces and thermal expansions at different radii.

The smallest margin of safety in terms of equivalent stress was obtained at the base of the disc. Permissible value . The strength condition is met.

The smallest margin of safety for breaking revolutions was also obtained at the base of the disc. Allowed value . The strength condition is met.

Rice. 2.2 Stress distribution (radius and ambient) along the disk radius

Rice. 2.3 Distribution of margin of safety (equivalent voltage margins) along the disc radius

Rice. 2.4 Distribution of safety margin over breaking revolutions

Rice. 2.5 Distribution of temperature, stress (rad. and ambient) along the radius of the disk

Literature

1. Khronin D.V., Vyunov S.A. etc. "Design and design of aircraft gas turbine engines". - M, Mechanical Engineering, 1989.

2. "Gas turbine engines", A.A. Inozemtsev, V.L. Sandratsky, OJSC Aviadvigatel, Perm, 2006

3. Lebedev S.G. Course project on the discipline "Theory and calculation of aircraft blade machines", - M, MAI, 2009.

4. Perel L.Ya., Filatov A.A. Rolling bearings. Directory. - M, Mechanical Engineering, 1992.

5. Program DISK-MAI, developed at the department 203 MAI, 1993.

6. Inozemtsev A.A., Nikhhamkin M.A., Sandratsky V.L. “Gas turbine engines. Dynamics and strength of aircraft engines and power plants. - M, Mechanical engineering, 2007.

7. GOST 2.105 - 95.

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Principle of operation gas turbine engine.

A gas turbine engine, like a piston engine on any car, refers to internal combustion engines. Both of them convert the chemical energy of the fuel into heat, by burning, and then into useful, mechanical. However, how this happens is somewhat different. In both engines, 4 main processes take place - these are: intake, compression, expansion, exhaust. Those. in any case, air (from the atmosphere) and fuel (from tanks) first enter the engine, then the air is compressed and fuel is injected into it, after which the mixture ignites, due to which it expands significantly, and is eventually released into the atmosphere. Of all these actions, only expansion gives energy, all the rest are necessary to ensure this action.

Now what's the difference. In gas turbine engines, all these processes occur constantly and simultaneously, but in different parts of the engine, and in a piston engine, in one place, but at different times and in turn. In addition, the more compressed the air, the more energy can be obtained during combustion, and today the compression ratio of gas turbine engines has already reached 35-40:1, i.e. in the process of passing through the engine, the air decreases in volume, and accordingly increases its pressure by 35-40 times. For comparison, in piston engines, this figure does not exceed 8-9: 1, in the most modern and advanced models. Accordingly, having equal weight and dimensions, a gas turbine engine is much more powerful, and its efficiency is higher. This is the reason for such a widespread use of gas turbine engines in aviation today.

And now more about the design. The four processes listed above take place in the engine, which is shown in the simplified diagram under the numbers:

air intake - 1 (air intake)
compression - 2 (compressor)
mixing and ignition - 3 (combustion chamber)
exhaust - 5 (exhaust nozzle)
The mysterious section at number 4 is called the turbine. This is an integral part of any gas turbine engine, its purpose is to obtain energy from gases that exit the combustion chamber at high speeds, and it is located on the same shaft as the compressor (2), which drives it.

Thus, a closed cycle is obtained. Air enters the engine, is compressed, mixed with fuel, ignited, directed to the turbine blades, which remove up to 80% of the gas power to rotate the compressor, all that is left determines the final engine power, which can be used in many ways.

Depending on the method of further use of this energy, gas turbine engines are divided into:

turbojet
turboprop
turbofan
turboshaft

The engine shown in the diagram above is turbojet. It can be said to be “clean” gas turbine, because after passing through the turbine, which rotates the compressor, the gases exit the engine through the exhaust nozzle at great speed and thus push the aircraft forward. Such engines are now used mainly in high-speed combat aircraft.

Turboprop engines differ from turbojet engines in that they have an additional turbine section, which is also called a low-pressure turbine, consisting of one or more rows of blades that take the energy left after the compressor turbine from the gases and thus rotate the propeller, which can be located both in front and behind the engine. After the second section of the turbine, the exhaust gases actually exit by gravity, having practically no energy, so just exhaust pipes are used to remove them. Similar engines are used in low-speed, low-altitude aircraft.

Turbofans engines have a similar scheme with turboprops, only the second section of the turbine does not take all the energy from the exhaust gases, so these engines also have an exhaust nozzle. But the main difference is that the low-pressure turbine drives the fan, which is enclosed in a casing. Therefore, such an engine is also called a dual-circuit engine, because the air passes through the internal circuit (the engine itself) and the external one, which is necessary only to direct the air stream that pushes the engine forward. That's why they have a rather "chubby" shape. It is these engines that are used on most modern airliners, since they are the most economical at speeds approaching the speed of sound and efficient when flying at altitudes above 7000-8000m and up to 12000-13000m.

Turboshaft the engines are almost identical in design to turboprops, except that the shaft that is connected to the low-pressure turbine comes out of the engine and can power absolutely anything. Such engines are used in helicopters, where two or three engines drive a single main rotor and a compensating tail propeller. Even tanks, the T-80 and the American Abrams, now have similar power plants.

Gas turbine engines are also classified according to other signs:

by input device type (adjustable, unregulated)
by compressor type (axial, centrifugal, axial-centrifugal)
according to the type of air-gas path (straight-through, loop)
by turbine type (number of stages, number of rotors, etc.)
by type of jet nozzle (adjustable, unregulated), etc.

Turbojet engine with axial compressor received wide application. With the engine running, the process is continuous. The air passes through the diffuser, slows down and enters the compressor. Then it enters the combustion chamber. Fuel is also supplied to the chamber through the nozzles, the mixture is burned, the combustion products move through the turbine. The products of combustion in the turbine blades expand and cause it to rotate. Further, gases from the turbine with reduced pressure enter the jet nozzle and break out at great speed, creating thrust. The maximum temperature also occurs in the water of the combustion chamber.

The compressor and turbine are located on the same shaft. Cold air is supplied to cool the combustion products. In modern jet engines, the operating temperature can exceed the melting point of rotor blade alloys by about 1000 °C. The cooling system for turbine parts and the choice of heat-resistant and heat-resistant engine parts are one of the main problems in the design of jet engines of all types, including turbojet ones.

A feature of turbojet engines with a centrifugal compressor is the design of the compressors. The principle of operation of such engines is similar to engines with an axial compressor.

Gas turbine engine. Video.

Useful related articles.

TO aircraft engines include all types of heat engines used as propulsion devices for aviation-type aircraft, i.e. devices that use aerodynamic quality to move, maneuver, etc. within the atmosphere (airplanes, helicopters, cruise missiles of classes "B-B", "V-3", "3-V", "3-3", aerospace systems, etc.). This implies a wide variety of used engines - from piston to rocket.

Aircraft engines (Fig. 1) are divided into three broad classes:

piston (PD);
air-jet (WFD including GTD);
missile (RD or RKD).

The last two classes are subject to a more detailed classification, in particular the class WFD.

By principle of air compression WRDs are divided into:

compressor , i.e., including a compressor for mechanical compression of air;
compressorless :
- once-through WFD ( SPVRD) with air compression only from velocity pressure;
- pulsating WFD ( PUVRD) with additional air compression in special intermittent gas-dynamic devices.

Rocket engine class LRE also refers to the compressor type of heat engines, since in these engines the working fluid (fuel) is compressed in a liquid state in turbopump units.

Solid propellant rocket engine (RDTT) does not have a special device for compressing the working fluid. It is carried out at the beginning of fuel combustion in the semi-enclosed space of the combustion chamber, where the fuel charge is located.

By operating principle there is a division: PD And PUVRD work in a cycle periodical actions, while WFD, GTD And RKD cycle is carried out continuous actions. This gives them advantages in terms of relative power, thrust, weight, etc., which determined, in particular, the expediency of their use in aviation.

By principle of jet thrust WRDs are divided into:

direct reaction engines;
indirect reaction engines.

Engines of the first type create tractive force (thrust P) directly - that's all rocket engines (RKD), turbojet without afterburner and with afterburner chambers ( TRD And TRDF), turbojet bypass (turbofan And TRDDF), once-through supersonic and hypersonic ( SPVRD And scramjet), pulsating (PUVRD) and numerous combined engines.

Indirect reaction gas turbine engines (GTD) transfer the power generated by them to a special propeller (propeller, propfan, helicopter main rotor, etc.), which creates tractive effort using the same air-jet principle ( turboprop , turbopropfan , turboshaft engines - TVD, TVVD, TVGTD). In this sense, the class WFD combines all engines that create thrust according to the air-jet principle.

Based on the considered types of engines of simple circuits, a number of combined engines , connecting the features and advantages of engines of various types, for example, classes:

turbo-jet engines - TRDP (TRD or turbofan + SPVRD);
rocket-ramjet - RPD (LRE or RDTT + SPVRD or scramjet);
rocket-turbine - RTD (TRD + LRE);

and many other combinations of engines of more complex schemes.

Piston engines (PD)

Two-row radial 14-cylinder air-cooled piston engine. General form.

piston engine (English) piston engine ) -

Classification of piston engines. Aircraft piston engines can be classified according to various criteria:

Depending on the type of fuel used- for light or heavy fuel engines.
According to the method of mixing- on engines with external mixture formation (carburetor) and engines with internal mixture formation (direct fuel injection into cylinders).
Depending on the method of ignition of the mixture- for positive ignition and compression ignition engines.
Depending on the number of strokes- for two-stroke and four-stroke engines.
Depending on the cooling method- for liquid and air-cooled engines.
By number of cylinders- for four-cylinder, five-cylinder, twelve-cylinder engines, etc.
Depending on the location of the cylinders- in-line (with cylinders arranged in a row) and star-shaped (with cylinders arranged in a circle).

In-line engines, in turn, are divided into single-row, two-row V-shaped, three-row W-shaped, four-row H-shaped or X-shaped engines. Axial engines are also divided into single-row, double-row and multi-row.

By the nature of the change in power depending on the change in altitude- for high-altitude, i.e. engines that retain power as the aircraft rises to altitude, and low-altitude engines whose power decreases with increasing flight altitude.
Propeller drive method- for motors with direct transmission to the propeller and gear motors.

Modern aircraft piston engines are four-stroke radial engines that run on gasoline. The cylinders of reciprocating engines are usually cooled by air. Previously, piston engines with water-cooled cylinders were also used in aviation.

The combustion of fuel in a piston engine is carried out in cylinders, while thermal energy is converted into mechanical energy, since under the pressure of the resulting gases, the piston moves forward. The translational movement of the piston, in turn, is converted into rotational movement of the engine crankshaft through the connecting rod, which is the connecting link between the cylinder with the piston and the crankshaft.

Gas turbine engines (GTE)

Gas turbine engine - a heat engine designed to convert the energy of fuel combustion into the kinetic energy of a jet stream and (or) into mechanical work on the engine shaft, the main elements of which are a compressor, a combustion chamber and a gas turbine.

Single-shaft and multi-shaft engines

The simplest gas turbine engine has only one turbine, which drives the compressor and at the same time is a source of useful power. This imposes a restriction on the operating modes of the engine.

Sometimes the engine is multi-shaft. In this case, there are several turbines in series, each of which drives its own shaft. The high-pressure turbine (the first after the combustion chamber) always drives the engine compressor, and the subsequent ones can drive both an external load (helicopter or ship propellers, powerful electric generators, etc.) and additional compressors of the engine itself, located in front of the main one.

The advantage of a multi-shaft engine is that each turbine operates at optimum speed and load. With a load driven from the shaft of a single-shaft engine, the throttle response of the engine, that is, the ability to quickly spin up, would be very poor, since the turbine needs to supply power both to provide the engine with a large amount of air (power is limited by the amount of air) and to accelerate the load. With a two-shaft scheme, a light high-pressure rotor quickly enters the regime, providing the engine with air, and the low-pressure turbine with a large amount of gases for acceleration. It is also possible to use a less powerful starter for acceleration when starting only the high pressure rotor.

Turbojet engine (TRD)

Turbojet engine (English) turbojet engine ) - a heat engine that uses a gas turbine, and jet thrust is formed when combustion products flow out of a jet nozzle. Part of the work of the turbine is spent on compressing and heating the air (in the compressor).

Scheme of a turbojet engine:
1. input device;
2. axial compressor;
3. combustion chamber;
4. turbine blades;
5. nozzle.

In a turbojet engine, the compression of the working fluid at the inlet to the combustion chamber and the high value of air flow through the engine are achieved due to the combined action of the oncoming air flow and the compressor located in the TRD tract immediately after the inlet device, in front of the combustion chamber. The compressor is driven by a turbine mounted on the same shaft with it, and running on the same working fluid, heated in the combustion chamber, from which a jet stream is formed. In the inlet device, the static air pressure increases due to the deceleration of the air flow. In the compressor, the total air pressure increases due to the mechanical work performed by the compressor.

Pressure ratio in the compressor is one of the most important parameters of the turbojet engine, since the effective efficiency of the engine depends on it. If for the first samples of turbojet engines this indicator was 3, then for modern ones it reaches 40. To increase the gas-dynamic stability of compressors, they are made in two stages. Each of the cascades operates at its own speed and is driven by its own turbine. In this case, the shaft of the 1st stage of the compressor (low pressure), rotated by the last (lowest speed) turbine, passes inside the hollow shaft of the compressor of the second stage (high pressure). Engine stages are also called low and high pressure rotors.

The combustion chamber of most turbojet engines has an annular shape and the turbine-compressor shaft passes inside the chamber ring. Upon entering the combustion chamber, the air is divided into 3 streams:

primary air- enters through the front openings in the combustion chamber, slows down in front of the injectors and takes a direct part in the formation of the fuel-air mixture. Directly involved in the combustion of fuel. The fuel-air mixture in the fuel combustion zone in the WFD is close to stoichiometric in composition.
secondary air- enters through the side openings in the middle part of the combustion chamber walls and serves to cool them by creating an air flow with a much lower temperature than in the combustion zone.
tertiary air- enters through special air channels in the outlet part of the combustion chamber walls and serves to equalize the temperature field of the working fluid in front of the turbine.

The gas-air mixture expands and part of its energy is converted in the turbine through the rotor blades into the mechanical energy of the rotation of the main shaft. This energy is spent primarily on the operation of the compressor, and is also used to drive engine units (fuel booster pumps, oil pumps, etc.) and drive electric generators that provide energy to various on-board systems.

The main part of the energy of the expanding gas-air mixture is used to accelerate the gas flow in the nozzle, which flows out of it, creating jet thrust.

The higher the combustion temperature, the higher the efficiency of the engine. To prevent the destruction of engine parts, heat-resistant alloys equipped with cooling systems and thermal barrier coatings are used.

Turbojet engine with afterburner (TRDF)

Turbojet engine with afterburner - modification of the turbojet engine, used mainly on supersonic aircraft. It differs from the turbojet engine by the presence of an afterburner between the turbine and the jet nozzle. An additional amount of fuel is supplied to this chamber through special nozzles, which is burned. The combustion process is organized and stabilized with the help of a front-end device that provides mixing of the evaporated fuel and the main flow. The increase in temperature associated with the heat input in the afterburner increases the available energy of the combustion products and, consequently, the speed of the exhaust from the jet nozzle. Accordingly, jet thrust (afterburner) also increases up to 50%, but fuel consumption increases sharply. Afterburner engines are generally not used in commercial aviation due to their low fuel economy.

Double-circuit turbojet engine (TRDD)

The first to propose the concept of a turbofan engine in the domestic aircraft engine industry was A. M. Lyulka (Based on research conducted since 1937, A. M. Lyulka submitted an application for the invention of a bypass turbojet engine. The copyright certificate was awarded on April 22, 1941.)

It can be said that from the 1960s to this day, in the aircraft engine industry, the era of turbofan engines. Turbofan engines of various types are the most common class of turbofan engines used on aircraft, from high-speed fighter-interceptors with low bypass turbofans to giant commercial and military transport aircraft with high bypass turbofans.

Scheme of a turbojet bypass engine:
1. low pressure compressor;
2. inner contour;
3. the output stream of the internal circuit;
4. output stream of the outer circuit.

The basis bypass turbojet engines the principle of attaching an additional mass of air to the turbojet engine passing through the external circuit of the engine was established, which makes it possible to obtain engines with a higher flight efficiency compared to conventional turbojet engines.

After passing through the inlet, the air enters the low pressure compressor, called the fan. After the fan, the air is divided into 2 streams. Part of the air enters the outer circuit and, bypassing the combustion chamber, forms a jet stream in the nozzle. The other part of the air passes through an internal circuit completely identical to the turbofan engine mentioned above, with the difference that the last stages of the turbine in the turbofan engine are the fan drive.

One of the most important parameters of a turbofan engine is the bypass ratio (m), that is, the ratio of air flow through the external circuit to the air flow through the internal circuit. (m \u003d G 2 / G 1, where G 1 and G 2 are the air flow through the internal and external circuits, respectively.)

When the bypass ratio is less than 4 (m<4) потоки контуров на выходе, как правило, смешиваются и выбрасываются через общее сопло, если m>4 - streams are ejected separately, since mixing is difficult due to a significant difference in pressures and velocities.

The turbofan engine is based on the principle of increasing the flight efficiency of the engine, by reducing the difference between the speed of the expiration of the working fluid from the nozzle and the flight speed. The reduction in thrust, which will cause a decrease in this difference between speeds, is compensated by an increase in air flow through the engine. The consequence of an increase in air flow through the engine is an increase in the area of the front section of the engine inlet, which results in an increase in the diameter of the engine inlet, which leads to an increase in its drag and mass. In other words, the higher the bypass ratio, the larger the diameter of the engine, all other things being equal.

All turbofan engines can be divided into 2 groups:

with mixing flows behind the turbine;
without mixing.

In a turbofan engine with a mixture of flows ( TRDDsm) air flows from the external and internal circuits enter a single mixing chamber. In the mixing chamber, these flows are mixed and leave the engine through a single nozzle with a single temperature. TRDSM are more efficient, however, the presence of a mixing chamber leads to an increase in the dimensions and weight of the engine

Turbofan engines, like turbofan engines, can be equipped with adjustable nozzles and afterburners. As a rule, these are turbofan engines with low bypass ratios for supersonic military aircraft.

Military turbofan EJ200 (m=0.4)

Bypass turbojet engine with afterburner (TRDDF)

Dual-circuit turbojet engine with afterburner - modification of the turbofan engine. Differs in the presence of an afterburner chamber. Has found wide application.

The combustion products leaving the turbine are mixed with the air coming from the external circuit, and then heat is supplied to the general flow in the afterburner, which operates on the same principle as in TRDF. The products of combustion in this engine flow from one common jet nozzle. Such an engine is called dual-circuit engine with a common afterburner.

TRDDF with deflectable thrust vector (OVT).

Thrust vector control (VCT) / Thrust vector deviation (VVT)

Special rotary nozzles, on some turbofan engines (F), allow you to deflect the flow of the working fluid flowing from the nozzle relative to the engine axis. OVT leads to additional losses of engine thrust due to the additional work on turning the flow and complicates the control of the aircraft. But these shortcomings are fully compensated by a significant increase in maneuverability and a reduction in the takeoff run of the aircraft and landing run, up to and including vertical takeoff and landing. OVT is used exclusively in military aviation.

High Bypass Turbofan / Turbofan Engine

Scheme of a turbofan engine:
1. fan;
2. protective fairing;
3. turbocharger;
4. the output stream of the internal circuit;
5. output stream of the outer circuit.

turbofan engine (English) turbofan engine ) is a turbofan engine with a high bypass ratio (m>2). Here, the low-pressure compressor is converted into a fan, which differs from the compressor in a smaller number of steps and a larger diameter, and the hot jet practically does not mix with the cold one.

This type of engine uses a single-stage, large-diameter fan that provides high airflow through the engine at all flight speeds, including low takeoff and landing speeds. Due to the large diameter of the fan, the nozzle of the outer contour of such turbofan engines becomes quite heavy and is often shortened, with straighteners (fixed blades that turn the air flow in the axial direction). Accordingly, most turbofan engines with a high bypass ratio - no mixing.

Device inner contour such engines are similar to the turbojet engine, the last stages of the turbine of which are the fan drive.

Outer loop Such a turbofan engine, as a rule, is a single-stage large-diameter fan, behind which there is a directing vane made of fixed blades, which accelerate the air flow behind the fan and turn it, leading to an axial direction, the outer contour ends with a nozzle.

Due to the fact that the fan of such engines, as a rule, has a large diameter, and the degree of air pressure increase in the fan is not high, the nozzle of the external circuit of such engines is rather short. The distance from the engine inlet to the outer contour nozzle exit can be much less than the distance from the engine inlet to the inner contour nozzle exit. For this reason, quite often the nozzle of the outer contour is mistaken for a fan fairing.

Turbofan engines with a high bypass ratio have a two- or three-shaft design.

Advantages and disadvantages.

The main advantage of such engines is their high efficiency.

Disadvantages - large weight and dimensions. Especially - the large diameter of the fan, which leads to significant air resistance in flight.

The scope of such engines is long- and medium-haul commercial airliners, military transport aviation.

Turbopropfan engine (TVVD)

Turbopropfan engine (English) turbo propfan engine ) -

Air-jet engines according to the method of pre-compression of air before entering the combustion chamber are divided into compressor and non-compressor. In compressorless air-jet engines, the velocity head of the air flow is used. In compressor engines, air is compressed by a compressor. The compressor air-jet engine is a turbojet engine (TRD). The group, called mixed or combined engines, includes turboprop engines (TVD) and bypass turbojet engines (DTRD). However, the design and operation of these engines are largely similar to turbojet engines. Often, all types of these engines are combined under the general name of gas turbine engines (GTE). Gas turbine engines use kerosene as fuel.

Turbojet engines

Structural schemes. A turbojet engine (Fig. 100) consists of an inlet, a compressor, a combustion chamber, a gas turbine, and an outlet.

The inlet device is designed to supply air to the engine compressor. Depending on the location of the engine on the aircraft, it may be part of the aircraft design or the engine design. The inlet device increases the air pressure in front of the compressor.

A further increase in air pressure occurs in the compressor. In turbojet engines, centrifugal compressors (Fig. 101) and axial compressors (see Fig. 100) are used.

In an axial compressor, when the rotor rotates, the blades, acting on the air, twist it and force it to move along the axis towards the outlet of the compressor.

In a centrifugal compressor, when the impeller rotates, the air is entrained by the blades and moves to the periphery under the action of centrifugal forces. Engines with an axial compressor have found the widest application in modern aviation.

The axial compressor includes a rotor (rotating part) and a stator (stationary part) to which the input device is attached. Protective screens are sometimes installed in the inlet devices to prevent foreign objects from entering the compressor, which can cause damage to the blades.

The compressor rotor consists of several rows of profiled rotor blades arranged in a circle and successively alternating along the axis of rotation. Rotors are divided into drum (Fig. 102, a), disk (Fig. 102, b) and drum-disk (Fig. 102, c).

The compressor stator consists of an annular set of profiled blades fixed in the housing. The row of fixed blades, called the straightener, together with the row of working blades, is called the compressor stage.

Modern aircraft turbojet engines use multi-stage compressors to increase the efficiency of the air compression process. The compressor stages are coordinated with each other so that the air at the outlet of one stage smoothly flows around the blades of the next stage.

The necessary air direction to the next stage is provided by the straightener. For the same purpose, the guide vane, installed in front of the compressor, also serves. In some engine designs, the guide vane may be absent.

One of the main elements of a turbojet engine is the combustion chamber located behind the compressor. Structurally, the combustion chambers are tubular (Fig. 103), annular (Fig. 104), tubular-annular (Fig. 105).

The tubular (individual) combustion chamber consists of a flame tube and an outer casing, interconnected by suspension cups. In front of the combustion chamber, fuel injectors and a swirler are installed to stabilize the flame. The flame tube has holes for air supply, which prevents overheating of the flame tube. Ignition of the fuel-air mixture in the flame tubes is carried out by special ignition devices installed on separate chambers. Between themselves, the flame tubes are connected by branch pipes, which provide ignition of the mixture in all chambers.

The annular combustion chamber is made in the form of an annular cavity formed by the outer and inner casings of the chamber. An annular flame tube is installed in the front part of the annular channel, and swirlers and nozzles are installed in the nose of the flame tube.

The tubular-annular combustion chamber consists of outer and inner casings forming an annular space inside which individual flame tubes are placed.

A gas turbine is used to drive the TRD compressor. In modern engines, gas turbines are axial. Gas turbines can be single-stage or multi-stage (up to six stages). The main components of the turbine include nozzle (guide) devices and impellers, consisting of disks and rotor blades located on their rims. The impellers are attached to the turbine shaft and form a rotor together with it (Fig. 106). Nozzle devices are located in front of the working blades of each disk. The combination of a fixed nozzle apparatus and a disk with working blades is called a turbine stage. The rotor blades are attached to the turbine disk with a Christmas tree lock (Fig. 107).

The exhaust device (Fig. 108) consists of an exhaust pipe, an inner cone, a rack and a jet nozzle. In some cases, due to the layout of the engine on the aircraft, an extension pipe is installed between the exhaust pipe and the jet nozzle. Jet nozzles can be with adjustable and unregulated output section.

Principle of operation. Unlike a piston engine, the working process in gas turbine engines is not divided into separate cycles, but proceeds continuously.

The principle of operation of a turbojet engine is as follows. In flight, the air flow against the engine passes through the inlet to the compressor. In the input device, the air is pre-compressed and the kinetic energy of the moving air flow is partially converted into potential pressure energy. Air is subjected to more significant compression in the compressor. In turbojet engines with an axial compressor, with the rapid rotation of the rotor, the compressor blades, like fan blades, drive air towards the combustion chamber. In the straighteners installed behind the impellers of each stage of the compressor, due to the diffuser shape of the interblade channels, the kinetic energy of the flow acquired in the wheel is converted into potential pressure energy.

In engines with a centrifugal compressor, air is compressed by centrifugal force. Air entering the compressor is picked up by the blades of a rapidly rotating impeller and, under the action of centrifugal force, is thrown from the center to the circumference of the compressor wheel. The faster the impeller rotates, the more pressure is generated by the compressor.

Thanks to the compressor, turbojet engines can create thrust when working on site. The efficiency of the air compression process in the compressor

characterized by the degree of pressure increase π to, which is the ratio of the air pressure at the outlet of the compressor p 2 to the pressure of atmospheric air p H

The air compressed in the inlet and compressor then enters the combustion chamber, splitting into two streams. One part of the air (primary air), which is 25-35% of the total air flow, is directed directly to the flame tube, where the main combustion process takes place. Another part of the air (secondary air) flows around the outer cavities of the combustion chamber, cooling the latter, and at the outlet of the chamber it mixes with combustion products, reducing the temperature of the gas-air flow to a value determined by the heat resistance of the turbine blades. A small part of the secondary air enters the combustion zone through the side openings of the flame tube.

Thus, a fuel-air mixture is formed in the combustion chamber by spraying fuel through the nozzles and mixing it with primary air, burning the mixture and mixing combustion products with secondary air. When the engine is started, the mixture is ignited by a special ignition device, and during further operation of the engine, the fuel-air mixture is ignited by the already existing flame.

The gas flow formed in the combustion chamber, which has a high temperature and pressure, rushes to the turbine through a narrowing nozzle apparatus. In the channels of the nozzle apparatus, the gas velocity increases sharply to 450-500 m/s and a partial conversion of thermal (potential) energy into kinetic energy takes place. The gases from the nozzle apparatus enter the turbine blades, where the kinetic energy of the gas is converted into the mechanical work of the turbine rotation. The turbine blades, rotating together with the disks, rotate the motor shaft and thereby ensure the operation of the compressor.

In the working blades of the turbine, either only the process of converting the kinetic energy of the gas into mechanical work of the rotation of the turbine can occur, or further expansion of the gas with an increase in its speed. In the first case, the gas turbine is called active, in the second - reactive. In the second case, the turbine blades, in addition to the active effect of the oncoming gas jet, also experience a reactive effect due to the acceleration of the gas flow.

The final expansion of the gas occurs in the engine outlet (jet nozzle). Here, the pressure of the gas flow decreases, and the speed increases to 550-650 m/sec (in terrestrial conditions).

Thus, the potential energy of the combustion products in the engine is converted into kinetic energy during the expansion process (in the turbine and outlet nozzle). Part of the kinetic energy in this case goes to the rotation of the turbine, which in turn rotates the compressor, the other part - to accelerate the gas flow (to create jet thrust).

Turboprop engines

Device and principle of operation. For modern aircraft

having a large carrying capacity and flight range, engines are needed that could develop the necessary thrust with a minimum specific weight. These requirements are met by turbojet engines. However, they are uneconomical compared to propeller-driven installations at low flight speeds. In this regard, some types of aircraft intended for flights at relatively low speeds and with a long range require the installation of engines that would combine the advantages of a turbojet engine with the advantages of a propeller-driven installation at low flight speeds. These engines include turboprop engines (TVD).

A turboprop is a gas turbine aircraft engine in which the turbine develops more power than is required to turn the compressor, and this excess power is used to turn the propeller. A schematic diagram of a TVD is shown in fig. 109.

As can be seen from the diagram, the turboprop engine consists of the same components and assemblies as the turbojet. However, unlike a turbojet engine, a propeller and a gearbox are additionally mounted on a turboprop engine. To obtain maximum engine power, the turbine must develop high speeds (up to 20,000 rpm). If the propeller rotates at the same speed, then the efficiency of the latter will be extremely low, since the propeller reaches its maximum efficiency in the design flight modes at 750-1,500 rpm.

To reduce the speed of the propeller compared to the speed of the gas turbine, a gearbox is installed in the turboprop engine. On high-power engines, two counter-rotating propellers are sometimes used, with one gearbox providing the operation of both propellers.

In some turboprop engines, the compressor is driven by one turbine and the propeller by another. This creates favorable conditions for engine regulation.

The thrust at the theater is created mainly by the propeller (up to 90%) and only slightly due to the reaction of the gas jet.

In turboprop engines, multistage turbines are used (the number of stages is from 2 to 6), which is dictated by the need to operate large heat drops on a turboprop turbine than on a turbojet turbine. In addition, the use of a multistage turbine makes it possible to reduce its speed and, consequently, the dimensions and weight of the gearbox.

The purpose of the main elements of the theater is no different from the purpose of the same elements of the turbojet engine. The workflow of a theater is also similar to that of a turbojet. Just as in a turbojet engine, the air flow pre-compressed in the inlet device is subjected to the main compression in the compressor and then enters the combustion chamber, into which fuel is simultaneously injected through the injectors. The gases formed as a result of the combustion of the air-fuel mixture have a high potential energy. They rush into the gas turbine, where, almost completely expanding, they produce work, which is then transferred to the compressor, propeller and unit drives. Behind the turbine, the gas pressure is almost equal to atmospheric pressure.

In modern turboprop engines, the thrust force obtained only due to the reaction of the gas jet flowing from the engine is 10-20% of the total thrust force.

Bypass turbojet engines

The desire to increase the thrust efficiency of turbojet engines at high subsonic flight speeds led to the creation of bypass turbojet engines (DTJE).

In contrast to the conventional turbojet engine, in a gas turbine engine a gas turbine drives (in addition to the compressor and a number of auxiliary units) a low-pressure compressor, otherwise called a secondary circuit fan. The fan of the second circuit of the DTRD can also be driven from a separate turbine located behind the compressor turbine. The simplest DTRD scheme is shown in fig. 110.

The first (internal) circuit of the DTRD is a circuit of a conventional turbojet. The second (external) circuit is an annular channel with a fan located in it. Therefore, bypass turbojet engines are sometimes called turbofans.

The work of DTRD is as follows. The air flow on the engine enters the air intake and then one part of the air passes through the high-pressure compressor of the primary circuit, the other part - through the fan blades (low-pressure compressor) of the secondary circuit. Since the circuit of the first circuit is the usual circuit of a turbojet engine, the workflow in this circuit is similar to the workflow in a turbojet engine. The action of the secondary circuit fan is similar to the action of a multi-bladed propeller rotating in an annular duct.

DTRD can also be used on supersonic aircraft, but in this case, to increase their thrust, it is necessary to provide for fuel combustion in the secondary circuit. To quickly increase (boost) the thrust of the DTRD, additional fuel is sometimes burned either in the air flow of the secondary circuit or behind the turbine of the primary circuit.

When additional fuel is burned in the secondary circuit, it is necessary to increase the area of its jet nozzle to keep the operating modes of both circuits unchanged. If this condition is not met, the air flow through the secondary circuit fan will decrease due to an increase in the gas temperature between the fan and the secondary circuit jet nozzle. This will entail a reduction in the power required to rotate the fan. Then, in order to maintain the previous engine speed, it will be necessary to reduce the temperature of the gas in front of the turbine in the primary circuit, and this will lead to a decrease in thrust in the primary circuit. The increase in total thrust will be insufficient, and in some cases the total thrust of the boosted engine may be less than the total thrust of a conventional diesel engine. In addition, boosting thrust is associated with high specific fuel consumption. All these circumstances limit the application of this method of increasing thrust. However, boosting the thrust of a DTRD can be widely used at supersonic flight speeds.

Used literature: "Fundamentals of Aviation" authors: G.A. Nikitin, E.A. Bakanov

In 2006, the management of the Perm Engine Building Complex and OAO Territorial Generating Company No. 9 (Perm Branch) signed an agreement for the manufacture and supply of a GTES-16PA gas turbine power plant based on a GTE-16PA with a PS-90EU-16A engine.

We asked Daniil SULIMOV, Deputy General Designer-Chief Designer of Aviadvigatel JSC, to tell us about the main differences between the new engine and the existing PS-90AGP-2.

The main difference between the GTE-16PA plant and the existing GTU-16PER is the use of a power turbine with a rotation speed of 3000 rpm (instead of 5300 rpm). Reducing the rotational speed makes it possible to abandon the expensive gearbox and increase the reliability of the gas turbine plant as a whole.

Specifications of GTU-16PER and GTE-16PA engines (under ISO conditions)

Optimization of the main parameters of the power turbine

The basic parameters of a free turbine (ST): diameter, flow path, number of stages, aerodynamic efficiency are optimized to minimize direct operating costs.

Operating costs include the cost of purchasing ST and costs for a certain (acceptable for the customer as a payback period) period of operation. The choice of a payback period that is quite visible for the customer (no more than 3 years) made it possible to implement an economically sound design.

The choice of the optimal variant of a free turbine for a specific application as part of the GTE-16PA was made in the engine system as a whole based on a comparison of direct operating costs for each variant.

Using one-dimensional modeling of the ST, the achievable level of aerodynamic efficiency of the ST was determined by the average diameter for a discretely given number of stages. The optimal flow part for this variant was chosen. The number of blades, taking into account their significant impact on the cost, was chosen from the condition of ensuring the Zweifel aerodynamic load factor equal to one.

Based on the selected flow path, the weight of the SP and the production cost were estimated. The turbine options in the engine system were then compared in terms of direct operating costs.

When choosing the number of stages for ST, the change in efficiency, acquisition and operation costs (fuel cost) are taken into account.

The cost of acquisition increases evenly with the growth of the cost price with an increase in the number of steps. In a similar way, the realized efficiency also grows - as a result of a decrease in the aerodynamic load on the stage. Operating costs (fuel component) fall with increasing efficiency. However, the total costs have a clear minimum at four stages in the power turbine.

The calculations took into account both the experience of our own developments and the experience of other companies (implemented in specific designs), which made it possible to ensure the objectivity of the estimates.

In the final design, by increasing the load per stage and reducing the efficiency of the ST from the maximum achievable value by about 1%, it was possible to reduce the total costs of the customer by almost 20%. This was achieved by reducing the cost and price of the turbine by 26% relative to the variant with maximum efficiency.

Aerodynamic design ST

The high aerodynamic efficiency of the new ST at a sufficiently high load was achieved by using the experience of JSC Aviadvigatel in the development of low-pressure turbines and power turbines, as well as the use of multi-stage spatial aerodynamic models using the Euler equations (without viscosity) and Navier-Stokes (taking into account viscosity ).

Comparison of the parameters of power turbines GTE-16PA and HPP Rolls-Royce

Comparison of the parameters of the ST GTE-16PA and the most modern Rolls-Royce TRD family TRD (Smith diagram) shows that in terms of the angle of rotation of the flow in the blades (approximately 1050), the new ST is at the level of Rolls-Royce turbines. The absence of a strict weight limit inherent in aircraft structures made it possible to somewhat reduce the load factor dH/U2 by increasing the diameter and circumferential speed. The value of the output speed (typical of ground structures) made it possible to reduce the relative axial speed. In general, the potential of the designed ST to realize efficiency is at the level characteristic of the stages of the Trent family.

The peculiarity of the aerodynamics of the designed ST is also to ensure the optimal value of the turbine efficiency at partial power modes, which are typical for operation in the base mode.

While maintaining the rotational speed, a change (decrease) in the load on the ST leads to an increase in the angles of attack (deviation of the direction of the gas flow at the inlet to the blades from the calculated value) at the inlet to the blade rims. Negative angles of attack appear, the most significant in the last stages of the turbine.

The design of ST blade rows with high resistance to changes in angles of attack is ensured by special profiling of the rows with additional verification of the stability of aerodynamic losses (according to 2D/3D Navier-Stokes aerodynamic models) at high inlet flow angles.

As a result, the analytical characteristics of the new ST showed significant resistance to negative angles of attack, as well as the possibility of using the ST to drive generators that produce current at a frequency of 60 Hz (with a rotation speed of 3600 rpm), that is, the possibility of increasing the rotational speed by 20 % without noticeable loss of efficiency. However, in this case, loss of efficiency is practically inevitable at low power modes (leading to an additional increase in negative angles of attack).
ST design features
To reduce the material consumption and weight of the ST, proven aviation approaches to turbine design were used. As a result, the mass of the rotor, despite the increase in the diameter and number of stages, turned out to be equal to the mass of the rotor of the GTU-16PER power turbine. This ensured a significant unification of the transmissions, the oil system, the pressurization system of the supports and the cooling system of the ST were also unified.
The amount and quality of air used to pressurize transmission bearings has been increased, including its cleaning and cooling. The quality of lubrication of transmission bearings has also been improved by using filter elements with a filtration fineness of up to 6 microns.
In order to increase the operational attractiveness of the new GTE, a specially developed control system has been introduced, which allows the customer to use turbo-expander (air and gas) and hydraulic launch types.
The weight and size characteristics of the engine make it possible to use serial designs of the GTES-16P packaged power plant for its placement.
The noise and heat insulating casing (when placed in capital premises) ensures the acoustic characteristics of the GTPP at the level provided for by sanitary standards.
The first engine is currently undergoing a series of special tests. The gas generator of the engine has already passed the first stage of equivalent-cyclic tests and has started the second stage after the revision of the technical condition, which will be completed in the spring of 2007.

The power turbine as part of a full-size engine passed the first special test, during which 7 throttle characteristics and other experimental data were taken.
According to the test results, a conclusion was made about the operability of the ST and its compliance with the declared parameters.
In addition, according to the test results, some adjustments were made to the design of the ST, including a change in the cooling system of the hulls to reduce heat release into the station room and ensure fire safety, as well as to optimize the radial clearances to increase efficiency, adjust the axial force.
The next test of the power turbine is scheduled for summer 2007.

Gas turbine plant GTE-16PA
on the eve of special tests