Gas turbine engine. Photo

03.03.2020

A bypass turbojet engine (TEF) is an “improved” turbojet engine, the design of which makes it possible to reduce fuel consumption, which is the main disadvantage of a turbofan engine, due to improved compressor operation and, accordingly, an increase in the volume of air masses passing through a turbofan engine.

For the first time, the design and principle of operation of the turbofan engine was developed by the aircraft designer A.M. The cradle back in 1939, but then they did not pay much attention to its development. Only in the 50s, when turbojet engines began to be massively used in aviation, and their "gluttony" became a real problem, his work was noticed and appreciated. Since then, the turbofan engine has been constantly improved and successfully used in all areas of aviation.

In fact, a bypass turbojet engine is the same turbojet engine, the body of which “envelops” another, external, body. The gap between these bodies forms the second circuit, while the first one is the internal cavity of the turbojet engine. Of course, the weight and dimensions increase at the same time, but the positive result from the use of such a design justifies all the difficulties and additional costs.

Device

The first circuit contains high and low pressure compressors, a combustion chamber, high and low pressure turbines and a nozzle. The second circuit consists of a guide vane and a nozzle. This design is basic, but some deviations are possible, for example, the flows of the inner and outer circuits can mix and exit through a common nozzle, or the engine can be equipped with an afterburner.

Now briefly about each constituent element of the turbofan engine. A high pressure compressor (HPC) is a shaft on which movable and fixed blades are fixed, forming a stage. Movable blades during rotation capture the air flow, compress it and direct it into the housing. The air enters the fixed blades, slows down and is additionally compressed, which increases its pressure and gives it an axial motion vector. There are several such stages in the compressor, and the compression ratio of the engine directly depends on their number. The same design is for the low pressure compressor (LPC), which is located in front of the HPC. The difference between them is only in size: LPC blades have a larger diameter, covering the cross section of both the primary and secondary circuits, and a smaller number of steps (from 1 to 5).

In the combustion chamber, compressed and heated air is mixed with fuel, which is injected by injectors, and the resulting fuel charge ignites and burns, forming gases with a large amount of energy. The combustion chamber can be one, annular, or made of several pipes.

The turbine in its design resembles an axial compressor: the same fixed and movable blades on the shaft, only their sequence is changed. First, the expanded gases fall on the fixed blades, which equalize their movement, and then on the movable ones, which rotate the turbine shaft. There are two turbines in the turbofan engine: one drives the high-pressure compressor, and the second drives the low-pressure compressor. They work independently and are not mechanically connected to each other. The LPC drive shaft is usually located inside the HPC drive shaft.

A nozzle is a converging pipe through which exhaust gases exit in the form of a jet stream. Usually, each circuit has its own nozzle, but it also happens that the jet streams at the outlet enter a common mixing chamber.

The outer, or second, circuit is a hollow annular structure with a guide vane, through which air passes, pre-compressed by a low-pressure compressor, bypassing the combustion chamber and turbines. This air flow, falling on the fixed blades of the guide vane, is leveled and moves to the nozzle, creating additional thrust due to compression of the LPC alone without burning fuel.

The afterburner is a pipe placed between the low pressure turbine and the nozzle. Inside it has swirlers and fuel injectors with igniters. The afterburner makes it possible to create additional thrust by burning fuel not in the combustion chamber, but at the turbine outlet. Exhaust gases after passing LPT and HPT have a high temperature and pressure, as well as a significant amount of unburned oxygen coming from the secondary circuit. Through the nozzles installed in the chamber, fuel is supplied, which mixes with gases and ignites. As a result, the output thrust sometimes doubles, however, the fuel consumption also increases. Turbofan engines equipped with an afterburner are easily recognizable by the flame that escapes from their nozzle during flight or upon launch.

cross section of the afterburner, swirlers are visible in the figure.

The most important parameter of a turbofan engine is the bypass ratio (k) - the ratio of the amount of air that has passed through the second circuit to the amount of air that has passed through the first. The higher this figure, the more economical the engine will be. Depending on the degree of bypass, the main types of bypass turbojet engines can be distinguished. If its value is<2, это обычный ТРДД, если же к>2, then such engines are called turbofan engines (TVRD). There are also turboprop-fan engines, in which the value reaches 50 or even more.

Depending on the type of exhaust gas discharge, turbofan engines are distinguished without mixing flows and with it. In the first case, each circuit has its own nozzle, in the second, the gases at the outlet enter the common mixing chamber and only then go outside, forming a jet thrust. Mixed-flow engines, which are installed on supersonic aircraft, can be equipped with an afterburner, which allows you to increase thrust even at supersonic speeds, when secondary thrust plays little role.

Principle of operation

The principle of operation of the TVRD is as follows. The air flow is captured by the fan and, partially compressed, is directed in two directions: to the first circuit to the compressor and to the second circuit to the fixed blades. In this case, the fan does not play the role of a screw that creates thrust, but a low-pressure compressor that increases the amount of air passing through the engine. In the primary circuit, the flow is compressed and heated as it passes through the high pressure compressor and enters the combustion chamber. Here it mixes with the injected fuel and ignites, resulting in the formation of gases with a large supply of energy. The flow of expanding hot gases is directed to the high-pressure turbine and rotates its blades. This turbine rotates the high-pressure compressor, which is mounted with it on the same shaft. Next, the gases rotate the low-pressure turbine, which drives the fan, after which they enter the nozzle and break out, creating jet thrust.

At the same time, in the second circuit, the air flow captured and compressed by the fan hits the fixed blades, which straighten the direction of its movement so that it moves in the axial direction. In this case, the air is additionally compressed in the second circuit and goes outside, creating additional traction. Also, the thrust is affected by the combustion of oxygen in the secondary air in the afterburner.

Application

The scope of application of bypass turbojet engines is very wide. They were able to cover almost the entire aviation, displacing the turbojet and theater engines. The main disadvantage of jet engines - their inefficiency - was partially overcome, so that now most civilian and almost all military aircraft are equipped with turbofan engines. For military aviation, where compactness, power and lightness of engines are important, turbofan engines with a low bypass ratio (to<1) и форсажными камерами. На пассажирских и грузовых самолетах устанавливаются ТРДД со степенью двухконтурности к>2, which saves a lot of fuel at subsonic speeds and reduces the cost of flights.

Low bypass bypass turbojet engines in military aircraft.

SU-35 with 2 AL-41F1S engines installed on it

Advantages and disadvantages

Bypass turbojet engines have a huge advantage over turbojet engines in the form of a significant reduction in fuel consumption without loss of power. But at the same time, their design is more complex, and the weight is much greater. It is clear that the greater the bypass ratio, the more economical the motor, but this value can be increased in only one way - by increasing the diameter of the second circuit, which will make it possible to pass more air through it. This is the main disadvantage of the turbofan. It is enough to look at some turbojet engines installed on large civil aircraft to understand how they make the overall structure heavier. The diameter of their second circuit can reach several meters, and in order to save materials and reduce their weight, it is shorter than the first circuit. Another disadvantage of large structures is the high drag during flight, which to some extent reduces the flight speed. The use of turbofan engines in order to save fuel is justified at subsonic speeds, when the sound barrier is overcome, the secondary jet thrust becomes ineffective.

Various designs and the use of additional structural elements in each individual case makes it possible to obtain the desired version of the turbofan engine. If economy is important, turbofan engines with a large diameter and a high bypass ratio are installed. If you need a compact and powerful engine, conventional turbofan engines with or without an afterburner are used. The main thing here is to find a compromise and understand what priorities a particular model should have. Military fighters and bombers cannot be equipped with engines with a three-meter diameter, and they don’t need it, because in their case, not so much economy is a priority, but speed and maneuverability. Here, turbofan engines with afterburners (TRDDF) are also more often used to increase traction at supersonic speeds or during launch. And for civil aviation, where the aircraft themselves are large, large and heavy engines with a high bypass ratio are quite acceptable.

The invention relates to low-pressure turbines of gas turbine engines for aviation applications. The low-pressure turbine of a gas turbine engine includes a rotor, a stator with a rear support, a labyrinth seal with internal and external flanges on the rear support of the stator. The labyrinth seal of the turbine is made in two levels. The inner tier is formed by two labyrinth sealing combs directed towards the turbine axis, and the working surface of the labyrinth seal inner flange directed towards the turbine flow path. The outer tier is formed by the sealing combs of the labyrinth directed towards the flow path of the turbine, and the working surface of the outer flange of the labyrinth seal directed towards the axis of the turbine. The sealing combs of the labyrinth of the inner tier of the labyrinth seal are made with parallel inner walls, between which a damping ring is installed. The outer flange of the labyrinth seal is made with an outer closed annular air cavity. Between the flow path of the turbine and the outer flange of the labyrinth seal there is an annular barrier wall mounted on the rear support of the stator. The working surface of the inner flange of the labyrinth seal is located in such a way that the ratio of the inner diameter at the outlet of the flow path of the turbine to the diameter of the working surface of the inner flange of the labyrinth seal is 1.05 1.5. The invention improves the reliability of the low-pressure turbine of a gas turbine engine. 3 ill.

Drawings to the RF patent 2507401

The invention relates to low-pressure turbines of gas turbine engines for aviation applications.

A low-pressure turbine of a gas turbine engine with a rear support is known, in which the labyrinth seal separating the rear discharge cavity of the turbine from the flow path at the outlet of the turbine is made in the form of a single tier. (S.A. Vyunov, "Design and design of aircraft gas turbine engines", Moscow, "Engineering", 1981, p. 209).

The disadvantage of the known design is the low stability of the pressure in the unloading cavity of the turbine due to the unstable value of the radial gaps in the labyrinth seal, especially at variable engine operating modes.

Closest to the claimed design is a low-pressure turbine of a gas turbine engine, including a rotor, a stator with a rear support, a labyrinth seal with inner and outer labyrinth flanges mounted on the rear support of the stator (US patent No. 7905083, F02K 3/02, 03/15/2011).

The disadvantage of the known design, taken as a prototype, is the increased value of the axial force of the turbine rotor, which reduces the reliability of the turbine and the engine as a whole due to the low reliability of the angular contact bearing, which perceives the increased axial force of the turbine rotor.

The technical result of the claimed invention is to increase the reliability of the low-pressure turbine of a gas turbine engine by reducing the magnitude of the axial force of the turbine rotor and ensuring the stability of the axial force when operating in transient conditions.

The specified technical result is achieved by the fact that in a low-pressure turbine of a gas turbine engine, including a rotor, a stator with a rear support, a labyrinth seal made with inner and outer flanges mounted on the rear support of the stator, the labyrinth seal of the turbine is made in two tiers, while the inner tier of the labyrinth seal formed by two sealing combs of the labyrinth directed to the axis of the turbine, and the working surface of the inner flange of the labyrinth seal directed to the flow path of the turbine, and the outer tier of the labyrinth seal is formed by sealing combs of the labyrinth directed to the flow path of the turbine, and the working surface of the outer flange of the labyrinth seal directed to the axis of the turbine, and the sealing combs of the labyrinth of the inner tier of the labyrinth seal are made with parallel inner walls, between which a damping ring is installed, and the outer flange of the labyrinth seal is made with an external closed annular air cavity, while between the flow path of the turbine and the outer flange of the labyrinth seal there is an annular barrier wall mounted on the rear support of the stator, and the working surface of the inner flange of the labyrinth seal is located in such a way that the following condition is met:

where D is the inner diameter at the outlet of the flow path of the turbine,

The labyrinth seal at the outlet of the low-pressure turbine is two-tier, arranging the seal tiers in such a way that the inner tier is formed by two labyrinth sealing scallops directed towards the turbine axis and the working surface of the labyrinth seal inner flange directed towards the flow path of the turbine, and the outer tier is formed directed to the flow path turbine sealing combs of the labyrinth and directed to the axis of the turbine working surfaces of the outer flange of the labyrinth seal, allows you to ensure reliable operation of the labyrinth seal in transient modes of operation of the turbine, which ensures the stability of the axial force acting on the turbine rotor, and increases its reliability.

The implementation of the sealing scallops of the labyrinth of the inner seal tier with parallel inner walls, between which a damping ring is installed, reduces vibration stresses in the labyrinth and reduces the radial gaps between the scallops of the labyrinth and the flanges of the labyrinth seal.

The execution of the outer flange of the labyrinth seal with an external closed air cavity, as well as the placement of an annular barrier wall installed on the rear stator support between the flow path of the turbine and the outer flange of the labyrinth seal, can significantly reduce the rate of heating and cooling of the outer flange of the labyrinth seal in transient modes, bringing it closer thus to the rate of heating and cooling of the outer tier of the labyrinth seal, which ensures the stability of the radial clearances between the stator and the rotor in the seal and increases the reliability of the low-pressure turbine by maintaining a stable pressure in the unloading after-turbine cavity.

The choice of the ratio D/d=1.05 1.5 is due to the fact that at D/d<1,05 снижается надежность работы лабиринтного уплотнения из-за воздействия на уплотнение высокотемпературного газа, выходящего из турбины низкого давления.

When D/d>1.5 reduces the reliability of the gas turbine engine by reducing the axial unloading force acting on the rotor of the low-pressure turbine.

Figure 1 shows a longitudinal section of a low-pressure turbine of a gas turbine engine.

Figure 2 - element I in figure 1 in an enlarged view.

Figure 3 - element II in figure 2 in an enlarged view.

The low-pressure turbine 1 of the gas turbine engine consists of a rotor 2 and a stator 3 with a rear support 4. To reduce the axial forces from gas forces acting on the rotor 2 at its outlet, an unloading cavity 6 of increased pressure, which is inflated with air due to the intermediate stage of the compressor (not shown) and is separated from the flow path 7 of the turbine 1 by a two-tier labyrinth seal, and the labyrinth 8 of the seal is fixed by a threaded connection 9 on the disk of the last stage 5 of the rotor 2, and the inner flange 10 and the outer flange 11 of the labyrinth seal are fixed on the rear support 4 of the stator 3. The inner tier of the labyrinth seal is formed by the working surface 12 of the inner flange 10, directed (facing) towards the flow path 7 of the turbine 1, and two sealing combs 13, 14 of the labyrinth 8 directed towards the axis 15 of the turbine 1. The inner walls 16,17 respectively of the scallops 13, 14 are made parallel to each other. A damping ring 18 is installed between the inner walls 16 and 17, which helps to reduce vibration stresses in the labyrinth 8 and reduce the radial gaps 19 and 20, respectively, between the labyrinth 8 of the rotor 2 and the flanges 10, 11. The outer tier of the labyrinth seal is formed by the working surface 21 of the outer flange 11, directed (facing) towards the axis 15 of the turbine 1, and the sealing scallops 22 of the labyrinth 8 directed towards the flow path 7 of the turbine 1. The outer flange 11 of the labyrinth seal is made with an outer closed annular air cavity 23 bounded from the outside by the wall 24 of the outer flange 11. Between the wall 24 of the outer flange 11 of the labyrinth seal and the flow path 7 of the turbine 1 there is an annular barrier wall 25 mounted on the rear support 4 of the stator 3 and protecting the outer flange 11 from the high-temperature gas flow 26 flowing in the flow path 7 of the turbine 1.

The working surface 12 of the inner flange 10 of the labyrinth seal is located in such a way that the condition is met:

where D is the inner diameter of the flow part 7 of the turbine 1 (at the outlet of the flow part 7);

d is the diameter of the working surface 12 of the inner flange 10 of the labyrinth seal.

The device works as follows.

During operation of the low-pressure turbine 1, the temperature state of the outer flange 11 of the labyrinth seal can be affected by a change in the temperature of the gas flow 26 in the flow path 7 of the turbine 1, which could significantly change the radial clearance 19 and the axial force acting on the rotor 2 due to a change in air pressure in the unloading cavity 6. However, this does not happen, since the inner flange 10 of the inner tier of the labyrinth seal is inaccessible to the influence of the gas flow 26, which contributes to the stability of the radial clearance 20 between the inner flange 10 and the labyrinth combs 13, 14, as well as the stability of the pressure in the cavity 6 and the stability of the axial force acting on rotor 2 of turbine 1.

CLAIM

A low-pressure turbine of a gas turbine engine, including a rotor, a stator with a rear support, a labyrinth seal with inner and outer flanges mounted on the rear support of the stator, characterized in that the labyrinth seal of the turbine is made in two tiers, while the inner tier of the labyrinth seal is formed by two labyrinth seal combs, directed to the axis of the turbine, and the working surface of the inner flange of the labyrinth seal directed to the flow path of the turbine, and the outer tier of the labyrinth seal is formed by the sealing combs of the labyrinth directed to the flow path of the turbine, and the working surface of the outer flange of the labyrinth seal directed to the axis of the turbine, and the sealing the scallops of the labyrinth of the inner tier of the labyrinth seal are made with parallel inner walls, between which a damping ring is installed, and the outer flange of the labyrinth seal is made with an outer closed annular air cavity, while between the flow path of the turbine and the outer flange of the labyrinth seal there is an annular barrier wall installed on the rear stator support, and the working surface of the inner flange of the labyrinth seal is located in such a way that the following condition is met:

D/d=1.05 1.5, where

D is the inner diameter at the outlet of the flow path of the turbine,

d is the diameter of the working surface of the inner flange of the labyrinth seal.

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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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Ministry of Education and Science of the Russian Federation

Federal Agency for Education

Samara State Aerospace University

named after Academician S.P. Queen

Department of the Theory of Aircraft Engines

Course work

on the course: "Theory and calculation of blade machines"

Axial turbine designaviationengineJT9 D20

Samara 2008

Exercise

Perform a design calculation of the main parameters of the high-pressure turbocharger and construct a meridional section of the high-pressure turbine of the JT9D-70A turbofan engine, perform a thermodynamic calculation of the turbine, a kinematic calculation of the second stage of the turbine, and profile the impeller blade in three sections: sleeve, middle and peripheral sections.

The initial parameters of the turbine are known from the thermodynamic calculation of the engine in takeoff mode (H P =0 and M P =0).

Table 1. Initial data for turbine design

high pressure turbine
Parameter	Numerical value	Dimension















TTND = TT
RTND = RT

Essay

Coursework on thermogasdynamic design of the JT9D20 axial turbine.

Explanatory note: 32 pages, 1 figure, 2 tables, 3 appendices, 4 sources.

TURBINE, COMPRESSOR, FLOW PART, WORKING WHEEL, NOZZLE DEVICE, STAGE, FLOW OUTPUT ANGLE, EFFECTIVE ANGLE, PROFILE SETTING ANGLE, GRID PIT, GRID WIDTH

In this course work, the diametrical dimensions of the high-pressure turbine were calculated, the meridional section of the flow path was constructed, the kinematic calculation of the stage at the average diameter and the calculation of the parameters for the blade height with the swirl law b = const were performed with the construction of velocity triangles at the inlet at the outlet of the RC in three sections (sleeve, peripheral and section on the average diameter). The profile of the blade of the impeller of the second stage is calculated, followed by the construction of the contour of the profile in the lattice in three sections.

Conventions

D - diameter, m;

Relative bushing diameter;

h - blade height, m;

F - cross-sectional area, m 2;

G - mass flow rate of gas (air), kg/s;

H - flight altitude, km; compressor head, kJ/kg;

i - specific enthalpy, kJ/kg;

k is the isentropic index;

l - length, m;

M - Mach number;

n - speed, 1/min;

Р - pressure, kPa;

Reduced speed;

s - flow velocity, m/s;

q(), (), () - gas dynamic functions of;

R - gas constant, kJ/kggrad;

L * k(t) - specific work of the compressor (turbine);

k(t) - efficiency of the compressor (turbine);

S - axial width of the crown, m;

T - temperature, K;

Assigned resource, h;

V - flight speed, m/s;

z - number of steps;

k, t - the degree of increase (decrease) of the total pressure;

The coefficient of restoration of the total pressure of air (gas) in the engine elements; tensile stresses, MPa;

Mass flow change factor;

U - circumferential speed, m/s;

Y t * =U t cf /C * t s - turbine load parameter;

Gap size, m;

U 2 t cf h t out /D cf out - stress parameter in the turbine blades, m 2 /s 2;

K tk, K tv - matching parameters of the gas generator, turbofan.

Indices

a - axial component;

c - air section at the compressor inlet

vent - fan

vzl - takeoff;

w - bushing section;

d - gases section at the outlet of the turbine

k - compressor section at the outlet of the compressor

kr - critical

ks - combustion chamber

n - cross section of the undisturbed flow

on - guide apparatus;

cool - cooling;

n - flight parameter, peripheral diameter;

pr - given parameters;

ps - retaining stage

s - isentropic parameters;

c - second section at the exit of the nozzle

cp - average parameter;

st - step parameter;

t - turbine fuel section at the turbine inlet

h - hourly

* - braking parameters.

Abbreviations

HP - high pressure;

LP - low pressure;

VNA - input guide vane;

GDF - gas dynamic functions

GTE - gas turbine engine

Efficiency - efficiency factor;

ON - guide vane;

RK - impeller;

SA - turbine nozzle apparatus;

SAU - standard atmospheric conditions

Turbofan engine - turbojet bypass engine.

Introduction

1. Design calculation of the main parameters of the high pressure turbine

1.1 Calculation of the geometrical and operating parameters of the HP turbine

1.2 Construction of the meridional section of the HP turbine flow path

2. Gas-dynamic calculation of the HP turbine

2.1 Distribution of heat drop by steps

2.2 Calculation of the step by average diameter

2.3 Calculation of the effective operation of the stage, taking into account friction losses of the disk and in the radial clearance

2.4 Calculation of flow parameters at different radii

Conclusion

List of sources used

Introduction

This work contains a simplified version of the gas-dynamic calculation of an axial turbine, in which the variant search for optimal (compromise) parameters is replaced by reliable statistical recommendations obtained by systematizing materials for the calculation of turbines of modern gas turbine engines. The design is carried out according to the initial parameters obtained in the thermogasdynamic calculation of the engine.

The purpose of designing an axial aircraft turbine is to determine the main geometric, kinematic and thermodynamic parameters as a whole and its individual stages, which provide the calculated values of the specific and general parameters of the engine. In this regard, the design tasks involve: selection of the main geometric parameters of the turbine being designed for given parameters of the working fluid, taking into account the intended purpose of the gas turbine engine; distribution of heat drop over the steps, calculation of flow parameters in the gaps between the steps; calculation of flow parameters in the elements of the flow path of the second stage of the turbine at the average diameter; selection of the swirl law and calculation of changes in flow parameters along the radius (blade height) of the designed stage; performing profiling of working blades of the designed stage.

1. Design calculation of the main parameters of the turbine of high

pressure

1.1 Calculation geometric and regime parameters HP turbines

The geometric parameters of the turbine to be determined are shown in Figure 1.

Figure 1. - Geometric model of an axial turbine

1. The value of the ratio D cf / h 2 (h 2 - the height of the rotor blades at the outlet of the HP turbine) is determined by the formula

where e t is the stress parameter, the value of which is usually within (13 ... 18) 10 3 m 2 / s 2.

We accept e t \u003d 15 10 3 m 2 / s 2. Then:

In order to obtain high efficiency, it is desirable to have. Therefore, a new value is chosen. Then,

2. Given the value of the axial gas velocity at the turbine inlet (C 0 =150 m / s), determine the reduced axial velocity l 0 (l 0 = 0.20 ... 0.25)

Annular area at the inlet to the SA of the HP turbine:

3. Calculate the annular area at the outlet of the turbine. To do this, the magnitude of the axial velocity component at the outlet of the turbine is preliminarily estimated. We accept that /= 1.5; . Then

4. According to the selected value, the height of the working blade at the outlet of the HP turbine is determined:

5. Average diameter at the HP turbine outlet

6. Peripheral diameter at the outlet of the valve:

7. Sleeve diameter at the outlet of the valve:

8. The shape of the flow part looks like: Therefore:

The height of the nozzle vane at the turbine inlet is estimated as follows:

9. Peripheral diameter of the nozzle apparatus at the HP turbine inlet:

10. Sleeve diameter at the HP turbine inlet:

11. HP turbine rotor speed:

1.2 Construction of the meridional section of the flowparts

HP turbines

The presence of the meridional shape of the flow path is necessary to determine the characteristic diameters Di in any control section of the step, and not only in sections "0" and "2". These diameters serve as the basis for performing, for example, the calculation of flow parameters at various radii of the flow path, as well as the design of control sections of the blade airfoil.

1. The width of the crown of the nozzle apparatus of the first stage:

accept kSA = 0.06

2. First stage impeller ring width:

accept kRK = 0.045

3. Width of the crown of the nozzle apparatus of the second stage:

4. Second stage impeller ring width:

5. The axial clearance between the nozzle apparatus and the impeller is usually determined from the ratio:

Axial clearance between the nozzle apparatus and the impeller of the first stage:

6. Axial clearance between the impeller of the first stage and the nozzle apparatus of the second stage:

7. Axial clearance between the nozzle apparatus and the impeller of the second stage:

8. The radial clearance between the ends of the blade feathers and the body is usually taken in the range of 0.8 ... 1.5 mm. In our case, we take:

2 . G azodynamic calculation of the turbine VD

2.1 Distributionheat drop reduction by steps

Thermodynamic parameters of the working fluid at the inlet andexiting the stairs.

1. Find the average value of the heat drop per step

.

The heat drop of the last stage is taken equal to:

Accept:

kJ/kg

Then: kJ/kg

2. Determine the degree of reactivity (for the second stage)

m

; ; .

3. Let us determine the parameters of the thermodynamic state of the gas at the inlet to the second stage

; ;

; ; .

4. Calculate the value of isentropic work in the stage when the gas expands to pressure.

Accept:

.

5. Let us determine the parameters of the thermodynamic state of the gas at the outlet of the stage under the condition of isentropic expansion from pressure to:

; .

6. Calculate the degree of gas reduction in the stage:

.

7. Determine the total pressure at the stage inlet:

,

8. We accept the angle of flow exit from the RC.

9. Gas-dynamic functions at the exit from the stage

; .

10. Static pressure downstream

.

11. Thermodynamic parameters of the flow at the outlet of the stage under the condition of isentropic expansion from pressure to

; .

12. The value of isentropic work in the stage when the gas expands from pressure to

.

2.2 Step calculation according to average at diameter at

Flow parameters behind the nozzle

1. Let us determine the isentropic velocity of gas outflow from the SA:

.

2. Determine the reduced isentropic flow velocity at the outlet of the SA:

;

3. Speed coefficient CA is accepted:

.

4. Gas-dynamic functions of the flow at the outlet of the SA:

; .

5. Determine the total pressure recovery coefficient from the table:

.

6. The angle of the flow exit from the nozzle blades:

;

Where.

7. Angle of flow deflection in an oblique section of SA:

.

8. Effective angle at the outlet of the nozzle array

.

9. The installation angle of the profile in the lattice is found according to the graph, depending on.

Accept: ;

;

.

10. Blade profile chord SA

.

11. The value of the optimal relative step is determined from the graph depending on and:

12. Optimal SA lattice spacing in the first approximation

.

13. Optimum number of SA blades

.

We accept.

14. The final value of the optimal pitch of the SA blades

.

15. The size of the throat of the SA channel

.

16. Parameters of the thermodynamic state of the gas at the outlet of the SA under the condition of isentropic expansion in the nozzle array

; .

17. Static pressure in the gap between SA and RK

.

18. Actual gas velocity at the outlet of the SA

.

19. Thermodynamic parameters of the flow at the outlet of the SA

;

; .

20. Density of the gas at the outlet of the SA

.

21. Axial and circumferential components of the absolute flow velocity at the outlet of the SA

;

.

22. Circumferential component of the relative flow velocity at the entrance to the AC

.

23. The angle of entry of the flow into the RC in relative motion

.

24. Relative flow velocity at the inlet to the AC

.

25. Thermodynamic parameters of the gas at the entrance to the AC

;

; .

26. Reduced flow velocity in relative motion

.

27. Total pressure in relative air movement

.

Flow parameters at the outlet of the RC

28. Thermodynamic flow parameters

;

;.

29. Isentropic flow velocity in relative motion

.

30. Reduced isentropic flow velocity in relative motion:

.

We accept, because relative motion is energy-isolated motion.

31. Reduced flow velocity in relative motion

Let's accept:

,

Then:

; .

32. Using the graph, we determine the total pressure recovery factor:

.

33. The angle of the flow exit from the RC in relative motion (15º<в 2 <45є)

Let's calculate:

;

.

34. Let's determine from the table the angle of flow deviation in the oblique section of the rotor blades:

.

35. Effective angle at the outlet of the DC

.

36. Let's determine from the table the angle of installation of the profile in the working blade:

Let's calculate:;

.

37. Blade profile chord RK

.

38. The value of the optimal relative lattice spacing of the Republic of Kazakhstan is determined from the tables:

.

39. Relative pitch of the RK lattice in the first approximation

.

40. Optimum number of blades RK

.

We accept.

41. The final value of the optimal pitch of the blades of the Republic of Kazakhstan

.

42. The size of the throat of the channel of the working blades

.

43. Relative speed at the exit from the Republic of Kazakhstan

44. Enthalpy and temperature of the gas at the outlet of the RC

; .

45. Density of gas at the outlet of the RC

46. Axial and circumferential components of the relative velocity at the exit from the RC

;

.

47. Circumferential component of the absolute flow velocity behind the RC

48. Absolute gas velocity behind the RK

.

49. The angle of the flow exit from the RC in absolute motion

50. Total enthalpy of gas behind the RC

.

2.3 Calculation of the effective operation of the stage, taking into account friction losses

disk and in the radial clearance

To determine the effective operation of the stage, it is necessary to take into account the energy losses associated with leakage of the working fluid into the radial clearance and friction of the stage disk against the gas. For this we define:

51. Specific work of gas on the blades of the Republic of Kazakhstan

52. Leakage losses, which depend on the design features of the stage.

In the designs of modern GTE turbines, bandages with labyrinth seals are usually used on impellers to reduce leakage. Leakage through such seals is calculated by the formula:

We accept the flow coefficient of the labyrinth seal:

The gap area is determined from the expression:

To determine the pressure first, the isentropic reduced flow velocity at the outlet to the RC at the peripheral diameter and the corresponding gas-dynamic function are found:

; .

Peripheral pressure

Seal pressure ratio

We accept the number of scallops:

Leakage loss

53. Energy loss due to friction of the stage disc on the gas

where D 1w is taken according to the drawing of the flow part

54. Total energy loss due to leakage and disk friction

55. The total enthalpy of the gas at the outlet of the RC, taking into account losses due to leakage and friction of the disk

;

56. Gas enthalpy according to static parameters at the outlet of the RC, taking into account losses due to leakage and friction of the disk

57. Total gas pressure at the outlet of the RC, taking into account losses due to leakage and disk friction

58. Actual effective operation of a stage

59. Actual efficiency steps

60. The difference between the actual effective work and the given one

which is 0.78%.

2.4 Calculation of parameters flow at different radii

turbine pressure blade wheel

At values D cf / h l< 12 по высоте лопатки возникает переменность параметров потока, определяемая влиянием центробежных сил и изменением окружной скорости. В этом случае для снижения потерь энергии лопатки необходимо выполнять закрученными. Применение закона закрутки dб/dr = 0 позволяет повысить технологическое качество лопаток. Применение закона б 1 =const позволяет выполнять сопловые венцы с б 1л =const, а закон б 2 =const позволяет улучшить технологичность лопаток соплового венца последующей ступени.

Determination of parameters for the spigot section of the blade

1. Relative bushing diameter

2. Flow exit angle in absolute motion

3. Speed ratio

4. Absolute flow rate at the outlet of the SA

5. Circumferential component of absolute speed

6. Axial component of absolute velocity

7. Isentropic velocity of gas outflow from SA

8. Thermodynamic parameters at the outlet of the SA

; ;

;

; .

9. Static pressure

.

10. Gas density

11. Circumferential speed in the sleeve section at the entrance to the RC

12. Circumferential component of the relative velocity at the entrance to the DC

13. The angle of entry of the flow into the RC in relative motion

.

14. Relative speed at the hub

15. Thermodynamic parameters at the entrance to the RC in relative motion

,

16. Total pressure at the inlet to the valve in relative motion

17. Reduced relative speed at the entrance to the RC

Parameters in peripheral section

18. Relates. peripheral section diameter

19. Angle of flow exit from SA in absolute motion

20. Speed ratio

21. Absolute speed at the exit from the SA

22. Circumferential and axial components of absolute speed

23. Isentropic velocity of gas outflow from SA

24. Thermodynamic parameters of the flow at the outlet of the SA

;

, ; .

25. Static pressure

26. Gas Density

27. The circumferential speed of rotation of the wheel on the periphery

28. Circumferential component of the relative velocity at the entrance to the RC

29. The angle of entry of the flow into the RC in relative motion

.

30. Relative flow velocity at the periphery

31. Thermodynamic parameters of the flow in relative motion at the entrance to the AC

,

32. Total pressure at the inlet to the CV in relative motion

.

33. Reduced relative velocity at the entrance to the RC

Calculation of flow parameters at the outlet of the RC

34. Relative bushing diameter

35. Flow angle in absolute motion

36. Peripheral speed in the sleeve section at the outlet of the valve

37. Static pressure at the outlet of the valve

38. Thermodynamic parameters in RK

,

39. Isentropic flow velocity at the outlet of the RC

40. Reduced isentropic speed

41. Flow velocity behind the RK in relative motion.

, Where

speed factor.

42. Thermodynamic parameters of the flow at the outlet of the RC

;

43. Gas density behind the working crown

44. Flow exit angle in relative motion

45. Circumferential and axial components of the relative flow velocity

46. Absolute speed at the output of the working crown

47. Circumferential component of absolute speed

48. Total enthalpy and temperature of the flow at the outlet of the AC

49. Gas-dynamic functions at the outlet of the RC

;

50. Total flow pressure in absolute motion at the outlet of the valve

Calculation of parameters in the peripheral section at the outlet of the RC

51. Relative diameter of the peripheral section

52. Flow angle in absolute motion

53. Peripheral speed in the peripheral section at the outlet of the RC

54. Static pressure at the outlet of the valve

55. Thermodynamic parameters during isentropic expansion in the Republic of Kazakhstan

;

56. Isentropic flow velocity at the outlet of the RC

57. Reduced isentropic speed

58. Flow velocity behind RK in relative motion

Speed ratio;

59. Thermodynamic parameters of the flow at the outlet of the RC

;

60. Gas density behind the working crown

61. Flow outlet angle in relative motion

62. Circumferential and axial components of the relative flow velocity

63. Absolute exit speed from RK

64. Circumferential component of absolute speed

65. Total enthalpy and temperature of the flow at the outlet of the AC

66. Gas-dynamic functions at the outlet of the RC

;

67. Total flow pressure in absolute motion at the outlet of the valve

3. Profiling of the impeller blade

Table 2. - Initial data for profiling of RV blades

Initial parameter and calculation formula	Dimension	Control sections

D (according to the drawing of the flow part of the stage)

Table 3. - Calculated values for the profiling of the blades RK

Value	Average diameter	Periphery

Conclusion

In the course work, the flow path of the high-pressure turbine was calculated and built, a kinematic calculation of the second stage of the high-pressure turbine at an average diameter was made, the calculation of effective operation, taking into account friction losses of the disk and in the radial clearance, the calculation of the parameters for the height of the blade with the swirl law b = const with the construction of triangles of speeds. Profiling of the impeller blade in three sections was performed.