Gas turbine engine. Photo

TO aircraft engines include all types of heat engines used as propulsion for aircraft aviation type, i.e. devices using aerodynamic quality for movement, maneuver, etc. within the atmosphere (airplanes, helicopters, cruise missiles of classes "B-B", "B-3", "3-B", "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).

There are two more detailed classifications. last class, especially 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 special device to compress 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 pulling 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 considered engine types simple circuits a number of combined engines connecting the features and benefits of engines 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)

Double-row radial 14-cylinder piston engine with air-cooled. 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- on lung engines or heavy fuel.
• According to the method of mixing- for engines with external mixture formation (carburetor) and engines with internal mixture formation ( direct injection fuel into the 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 radial four-stroke engines running 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. translational movement piston, in turn, is converted into rotational motion crankshaft engine 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 - 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 the compressor, combustion chamber and 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 consecutive standing turbines, each of which leads its 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 optimal number rpm 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 turbine low pressure plenty of gas 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 ) is a heat engine that uses a gas turbine, and jet thrust formed when combustion products flow out of the 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. The compressor is growing full pressure air due to the mechanical work done by the compressor.

Pressure ratio in the compressor is one of the most important parameters TRD, since the effective Engine efficiency. 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 holes in the combustion chamber, slows down in front of the injectors and takes a direct part in the formation fuel-air mixture. Directly involved in the combustion of fuel. Fuel-air mixture in the fuel combustion zone in the WFD is close to stoichiometric in its 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 the drive of 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 turbofans with a high degree bypass.

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 double-circuit turbo jet 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. OBT leads to additional losses of engine thrust due to the implementation additional work on the turn of the flow and complicate 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 motor uses a large diameter, single-stage fan that provides high flow air through the engine at all flight speeds, including low speeds during takeoff and landing. 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 directing vanes (fixed blades that turn 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 quite 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 a significant drag air 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 ) -

Today, aviation is almost 100% composed of machines that use a gas turbine type of power plant. In other words - gas turbine engines. However, despite the increasing popularity of air travel now, few people know how that buzzing and whistling container that hangs under the wing of an airliner works.

Principle of operation gas turbine engine.

A gas turbine engine, like a piston engine on any car, refers to engines internal combustion. 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 in different moment time and in sequence. 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 indicator does not exceed 8-9: 1, in the most modern and perfect samples. Accordingly, having equal weight and dimensions, the gas turbine engine is much more powerful, and the coefficient useful action he 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 additional section turbine, 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 gases and thus rotate air 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 they are simply used to remove them. exhaust pipes. 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. Similar power plants now they even have tanks - the T-80 and the American Abrams.

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. When running engine is coming continuous process. 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. To cool the products of combustion, cold air. In modern jet engines working 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. Principle of operation similar engines similar to engines with axial compressor.

Gas turbine engine. Video.

Useful related articles.

The utility model makes it possible to increase the efficiency of a bypass turbojet engine (TEF) by guaranteeing cooling of the last turbine stage at maximum modes(for example, in takeoff mode) and increase efficiency in cruising modes of operation. The cooling system of the last stage of the axial low-pressure turbine of the turbofan engine contains an air intake from the outer circuit of the engine and an additional air intake behind one of the intermediate compressor stages. The cooling system is equipped with a device for regulating the air supply to the cavity adjacent to the rear surface of the turbine disk of the last stage. The control device contains a rotary ring with a drive. The swivel ring contacts the end wall of the turbine support. Two holes are made in the end wall of the support. One hole is connected to the annular cavity of the turbine support of the last stage, and the other is connected to the cavity of the air collector located in the annular cavity of the turbine support. The swivel ring of the control device is provided with a through elliptical hole located with the possibility of alternate communication with one of the two through holes of the end wall of the turbine support.

The utility model relates to aircraft engine element cooling systems, and more specifically to the cooling system of a low-pressure turbine (LPT) of a bypass turbojet engine (TRDD).

Cooling air is used to cool the hot structural elements of turbojet engines.

A well-known turbine cooling system of a turbojet bypass engine, in which air is used to cool the turbine blades, which is taken from the intermediate or last stage of the high pressure compressor (HPC) (see, for example, "Design of the turbocharger TRDDF", MAI Publishing House, 1996, page .27-28). The cooling air taken from the HPC has a sufficiently high pressure (compared to the place of its release into the turbine flow path), which ensures its guaranteed supply to all cooling surfaces. In this regard, the efficiency of such a cooling system is very high.

The disadvantage of using such a cooling system is to reduce the specific thrust at maximum modes and efficiency in cruising modes. This decrease occurs due to the fact that part of the power of the high pressure turbine, which goes to compress the LPT cooling air, is lost and is not used either to rotate the high pressure compressor (HPC) or to create engine thrust. For example, when the flow rate of the LPT blade cooling air is ~5% of the air flow rate at the HPC inlet, and air is taken from its last stage, the power loss can be ~5%, which is equivalent to reducing the turbine efficiency by the same amount.

Closest to the claimed technical solution is the turbine cooling system of a bypass turbojet engine, in which air taken from the external circuit channel is used to cool the low-pressure turbine blades (see, for example, "Turbojet bypass engine with an afterburner AL-31F" Tutorial, publishing house of VVIA named after N.E. Zhukovsky, 1987, pp. 128-130). Turbine cooling is carried out in all engine operating modes. With this variant of cooling air extraction, additional turbine power is not consumed for its compression in HPC, therefore, a larger amount of potential energy of the gas flow behind the turbine can be converted in the jet nozzle into the kinetic energy of the exhaust jet, which, in turn, will lead to an increase in engine thrust and its economy.

The disadvantage of using such a cooling system is the reduction in cooling efficiency due to insufficient pressure air taken from the channel of the external cooling air circuit at engine operating modes close to maximum (for example, takeoff mode). In these operating modes, the optimal ratio for the efficiency of the engine (the maximum value of the specific thrust of the engine) is the ratio of pressures in the channel of the outer circuit and at the outlet of the low-pressure turbine is close to one. Such a pressure difference, taking into account losses in the supply channels and nozzles, is not enough to implement efficient cooling working blades of the LPT of the engine in these modes.

Known technical solutions have limited capabilities, as they lead to a decrease in the efficiency of the engine.

The utility model is based on the task of increasing the efficiency of the turbofan engine by guaranteeing cooling of the last turbine stage at maximum modes (for example, takeoff) and increasing efficiency in cruising modes of operation.

The technical result is an increase in the efficiency of the turbofan engine.

The problem is solved by the fact that the cooling system of the last stage of the axial low-pressure turbine of the bypass turbojet engine contains an air intake from the outer circuit of the engine. The air intake communicates through the cavities of the racks and the annular cavity of the last stage turbine support, provided with a front end wall, with the cavity adjacent to the rear surface of the turbine disk, and through the pressure disk with the internal cavities of the blades. The end wall of the turbine support has through holes, and the outer surface of the turbine housing of the last stage is made in the form of a part of the inner surface of the channel of the outer contour of the engine.

What is new in the utility model is that the cooling system is additionally provided at the inlet with an air intake behind one of the intermediate stages of the compressor, connected by a pipeline to a hollow air collector at the outlet. The cooling system is equipped with a device for regulating the air supply to the cavity adjacent to the rear surface of the turbine of the last stage. The control device contains a rotary ring with a drive. The swivel ring contacts the end wall of the turbine support. Two holes are made in the end wall of the support. One hole is connected to the annular cavity of the turbine support of the last stage, and the other is connected to the cavity of the air collector located in the annular cavity of the turbine support. The swivel ring of the control device is provided with a through elliptical hole located with the possibility of alternate communication with one of the two through holes of the end wall of the turbine support.

The implementation of the cooling system of the last stage of the axial low-pressure turbine of a bypass turbojet engine in accordance with the claimed utility model provides:

Additional supply of the cooling system at the inlet with an air intake behind one of the intermediate stages of the compressor, connected by a pipeline to a hollow air collector at the outlet, communicating with the cavity, of the rear surface of the disk of the last turbine stage, ensures guaranteed cooling at maximum modes, including takeoff mode;

The supply of the cooling system with a device for regulating the air supply to the cavity adjacent to the rear surface of the disk of the last turbine stage from the intermediate stage of the compressor or from the external circuit ensures efficient cooling of the LPT rotor blade in all engine operating modes. The control device allows you to combine positive traits both cooling systems, that is, by connecting in series various cooling air supply channels, it is most rational to ensure the operability and efficiency of the turbine cooling system in the entire range of engine operating modes and thereby improve the traction, economic and resource characteristics of the engine. Thus, in take-off mode, the control device is connected in such a way that cooling air from the intermediate stage of the compressor is supplied with a pressure sufficient to effectively cool the last stage of the turbine. This allows either, at a fixed cooling air flow rate, to increase the life of the turbine and the entire engine as a whole, or to reduce the cooling air flow rate and thereby increase traction characteristics engine. The air in the duct of the outer circuit does not have the overpressure necessary for efficient cooling. In cruising mode, the control device ensures the supply of cooling air from the channel of the external circuit, while the channel for air intake from the compressor is blocked (the position of the ring is switched by a signal depending on the speed of the low-pressure turbine shaft of the engine n nd and the stagnation temperature of the air at the engine inlet T * N). Due to the fact that the cooling air does not undergo compression in the compressor, the required HPC power decreases and increases free energy the working fluid behind the turbine; this leads to an increase in engine thrust and its efficiency. In addition, the air from the channel of the outer circuit has a large cooling resource, which will either increase the life of the turbine and the entire engine as a whole at a fixed flow rate of cooling air, or reduce the consumption of cooling air and thereby further increase the efficiency of the engine.

Thus, the problem posed in the utility model has been solved - increasing the efficiency of the turbofan engine by guaranteeing cooling of the last turbine stage at maximum modes (for example, takeoff) and increasing efficiency in cruising operating modes compared to known analogues.

The present utility model is explained in the following detailed description cooling system and its operation with reference to the drawings shown in figures 1-3, where

figure 1 schematically shows a longitudinal section of the last stage of the axial low-pressure turbine of a bypass turbojet engine and its cooling system;

figure 2 - view A in figure 1;

figure 3 - section B-B in Fig.2.

The cooling system of the last stage of the axial low-pressure turbine of a bypass turbojet engine contains (see figure 1) the air intake 1 from the outer circuit 2 of the engine. The air intake 1 communicates with the cavity 3 adjacent to the rear surface of the disk 4 of the turbine through the cavity 5 of the racks 6 and the annular cavity 7 of the turbine support of the last stage, provided with a front end wall 8 with through holes 9 (see Fig.2, 3) of the turbine, and through channels 10 in disk 4 with internal cavities of blades 11.

The cooling system of the last stage of the low-pressure axial turbine of the bypass turbojet engine additionally contains an air intake behind one of the intermediate compressor stages at the inlet (the air intake and the intermediate stages of the compressor are not shown in figure 1). This air intake is connected by a pipeline 12 with a hollow air collector 13 at the outlet adjacent to the end wall 8 of the turbine support with through holes 14 (see Fig.2, 3).

Moreover, the cooling system is equipped with a device for regulating the air supply to the cavity 3 adjacent to the rear surface of the disk 4 of the turbine of the last stage. The control device is made in the form of a rotary ring 15 (see Fig.1-3) with a drive (the drive is not shown) in contact with the end wall 8 of the turbine support, where the hole 9 provides communication cavity 3 with the annular cavity 7, and the hole 14 provides communication of the cavity 3 with the cavity 16 of the air collector 13 located in the annular cavity 7 of the turbine support. The drive of the rotary ring 15 can be made, for example, in the form of a pneumatic motor or a drive of a similar type. The swivel ring 15 of the control device has a through elliptical hole 17, which allows alternate communication with the through holes 9, 14 in the end wall 8 of the turbine support.

The proposed cooling system contains an air intake a (air intake not shown in figure 1) behind one of the intermediate stages of the compressor, air intake 1 b from the channel of the outer circuit 2. The operation of the cooling air supply system is described below.

The cooling system of the last stage of the axial low-pressure turbine of a bypass turbojet engine operates as follows. Ring 15 can be in two positions. When the ring 15 is turned to position I (see Fig.2) (take-off mode of the engine), air a flows through the pipe 12, under the action of a pressure difference, through the air collector 13, the hole 14 in the wall 8 and the hole 17 in the ring 15 into the cavity 3 , adjacent to the rear surface of the disk 4. In this case, the passage to the cavity 3 of the air b is blocked by the ring 15. When the ring 15 is turned to position II (not shown) (cruise mode), the hole 17 is rotated so that the hole 14 is blocked by the ring 15, and air b enters cavity 3 through hole 9 and hole 17 in ring 15. In this case, the air a, taken after the intermediate stage of the compressor, does not enter the cavity 3.

Switching ring 15 to position I or II is carried out by a signal depending on the speed n of the shaft of the low-pressure turbine of the engine and the stagnation temperature of the air at the engine inlet T* H. When high values parameter (take-off engine operation) ring 15 is in position I, at low values ​​of the parameter (cruising mode) - in position II.

Execution of the cooling system in accordance with the declared technical solution allows to provide the necessary cooling of the last stage of the low-pressure turbine in all engine operating modes, while simultaneously increasing the efficiency and economy of its operation.

The cooling system of the last stage of the axial low-pressure turbine of a bypass turbojet engine, containing an air intake from the outer contour of the engine, communicating through the cavities of the racks and the annular cavity of the turbine support of the last stage, equipped with a front end wall, with a cavity adjacent to the rear surface of the turbine disk, and through the pressure a disk with internal cavities of the blades, where the end wall of the turbine support has through holes, characterized in that the cooling system is additionally equipped at the inlet with an air intake behind one of the intermediate stages of the compressor, connected by a pipeline to a hollow air collector at the outlet, and a device for regulating the air supply to the cavity, adjacent to the rear surface of the turbine of the last stage, where the control device is made in the form of a rotary ring with a drive in contact with the end wall of the turbine support, two holes are made in the end wall of the support, where one hole is connected to the annular cavity of the turbine support of the last stage, and the other - to cavity of the air collector located in the annular cavity of the turbine support, the rotary ring of the control device is provided with a through elliptical hole located with the possibility of alternate communication with one of the two through holes of the end wall of the turbine support.

Turbine

The turbine is designed to drive the compressor and auxiliary units engine. Engine turbine - axial, jet, two-stage, cooled, two-rotor.

The turbine assembly includes sequentially arranged single-stage high and low pressure axial turbines, as well as a turbine support. Support - an element of the power circuit of the engine.

high pressure turbine

SA HPT consists of an outer ring, an inner ring, a cover, a swirling device, nozzle blade blocks, labyrinth seals, nozzle blade joint seals, spacers with honeycomb inserts and fasteners.

The outer ring has a flange for connection with the flange of the rim of the LPT nozzle apparatus and the VVT ​​body. The ring is telescopically connected to the VVT ​​body and has a cavity for supplying secondary air from the OCS to cool the outer shelves of the nozzle blades.

The inner ring has a flange for connection with the cover and the inner body of the OKS.

SA TVD has forty-five blades, combined into fifteen cast three-blade blocks. The block design of the SA blades makes it possible to reduce the number of joints and gas overflows.

Nozzle blade - hollow, cooled two-cavity. Each blade has a vane, outer and inner flanges, which together with the vane and flanges of adjacent blades form the flow path of the HPT SA.

The TVD rotor is designed to convert the energy of the gas flow into mechanical work on the rotor shaft. The rotor consists of a disk, pins with labyrinth and oil sealing rings. The disk has ninety-three slots for fastening the HPT rotor blades in “Christmas tree” locks, holes for tight-fitting bolts tightening the disc, pin and HPT shaft, as well as inclined holes for supplying cooling air to the rotor blades.

HPT working blade - cast, hollow, cooled. In the inner cavity of the blade to organize the cooling process, there is a longitudinal partition, turbulent pins and ribs. The shank of the blade has an elongated leg and a herringbone type lock. In the shank there are channels for supplying cooling air to the blade airfoil, and in the trailing edge there is a slot for air outlet.

The trunnion shank contains an oil seal and a radial roller bearing back support high pressure rotor.

Low pressure turbine

SA LPT consists of a rim, blocks of nozzle blades, an inner ring, a diaphragm, and honeycomb inserts.

The rim has a flange for connection with the VVT ​​housing and the outer ring of the HPT, as well as a flange for connection with the turbine support housing.

SA TND has fifty-one blades soldered into twelve four-blade blocks and one three-blade block. Nozzle blade - cast, hollow, cooled. The feather, outer and inner shelves form with the feather and shelves of neighboring blades the flow part of the SA.

A perforated deflector is placed in the inner part of the cavity of the blade airfoil. On the inner surface of the pen there are transverse ribs and turbulence pins.

The diaphragm is designed to separate the cavities between the HPT and LPT impellers.

The LPT rotor consists of a disk with working blades, a trunnion, a shaft and a pressure disk.

The LPT disc has fifty-nine grooves for fastening the working blades and inclined holes for supplying cooling air to them.

The working vane of TND - cast, hollow, cooled. On the peripheral part, the blade has a shroud with a labyrinth seal comb, which seals the radial gap between the stator and the rotor.

From axial movements in the disc, the blades are fixed by a split ring with an insert, which, in turn, is fixed by a pin on the disc rim.

The trunnion has internal splines in the front part to transmit torque to the LPT shaft. On the outer surface of the front part of the trunnion, there is an inner race of the roller bearing of the rear support of the HPT, a labyrinth and a set of sealing rings, which, together with the cover installed in the trunnion, form the front seal of the oil cavity of the HPT support.

A set of sealing rings is installed on the cylindrical belt in the rear part, which, together with the cover, form a seal for the oil cavity of the LPT support.

The TND shaft consists of three parts. The connection of the shaft parts to each other is forked. The torque at the joints is transmitted by radial pins. At the rear of the shaft there is an oil pump for the turbine support.

In front of the LPT there are splines that transmit torque to the low-pressure compressor rotor through the spring.

The pressure disk is designed to create an additional backwater and provides an increase in the pressure of the cooling air at the inlet to the working blades of the LPT.

The turbine support includes a support housing and a bearing housing. The support housing consists of an outer housing and an inner ring connected by power racks and forming a power circuit for the turbine support. The structure of the support also includes a screen with fairings, a defoaming mesh and fasteners. Inside the racks there are pipelines for supplying and pumping oil, venting oil cavities and draining oil. Air is supplied through the cavities of the racks to cool the LPT and air is removed from the pre-oil cavity of the support. Racks are covered with fairings. An oil sump pump and an oil collector are installed on the bearing housing. An elastic-oil damper is placed between the outer race of the LPT rotor roller bearing and the bearing housing.

A cone-fairing is fixed on the turbine support, the profile of which ensures gas entry into the afterburner combustion chamber with minimal losses.