Double-circuit turbojet engine. Low pressure turbine of a gas turbine engine Low pressure turbine

03.03.2020

Transmission

For the first time an aircraft with a turbojet engine ( TRD) took to the air in 1939. Since then, the design of aircraft engines has been improved, various types have appeared, but the principle of operation for all of them is approximately the same. To understand why an aircraft with such a large mass can take to the air so easily, you need to understand how an aircraft engine works. A turbojet engine propels an aircraft using jet propulsion. In turn, jet thrust is the recoil force of the gas jet that flies out of the nozzle. That is, it turns out that the turbojet installation pushes the plane and all the people in the cabin with the help of a gas jet. The jet stream, flying out of the nozzle, is repelled from the air and thus sets the aircraft in motion.

Turbofan engine device

Design

The device of the aircraft engine is quite complicated. The operating temperature in such installations reaches 1000 degrees or more. Accordingly, all the parts that make up the engine are made of materials that are resistant to high temperatures and fire. Due to the complexity of the device, there is a whole field of science about turbojet engines.

TRD consists of several main elements:

fan;
compressor;
the combustion chamber;
turbine;
nozzle.

A fan is installed in front of the turbine. With its help, air is drawn into the unit from the outside. In such installations, fans with a large number of blades of a certain shape are used. The size and shape of the blades provide the most efficient and fast air supply to the turbine. They are made from titanium. In addition to the main function (drawing in air), the fan solves another important task: it is used to pump air between the elements of the turbojet engine and its shell. Due to this pumping, the system is cooled and the destruction of the combustion chamber is prevented.

A high power compressor is located near the fan. With its help, air enters the combustion chamber under high pressure. In the chamber, air is mixed with fuel. The resulting mixture is ignited. After ignition, the mixture and all adjacent elements of the installation are heated. The combustion chamber is most often made of ceramic. This is due to the fact that the temperature inside the chamber reaches 2000 degrees or more. And ceramics is characterized by resistance to high temperatures. After ignition, the mixture enters the turbine.

View of the aircraft engine from the outside

A turbine is a device consisting of a large number of blades. The flow of the mixture exerts pressure on the blades, thereby setting the turbine in motion. The turbine, due to this rotation, causes the shaft on which the fan is mounted to rotate. It turns out a closed system, which for the operation of the engine requires only the supply of air and the presence of fuel.

Next, the mixture enters the nozzle. This is the final stage of the 1st engine cycle. This is where the jet stream is formed. This is how an airplane engine works. The fan forces cold air into the nozzle, preventing it from being destroyed by an excessively hot mixture. The cold air flow prevents the nozzle collar from melting.

Various nozzles can be installed in aircraft engines. The most perfect are considered mobile. The movable nozzle is able to expand and contract, as well as adjust the angle, setting the correct direction of the jet stream. Aircraft with such engines are characterized by excellent maneuverability.

Types of engines

Aircraft engines are of various types:

classic;
turboprop;
turbofan;
straight-through.

Classic installations work according to the principle described above. Such engines are installed on aircraft of various modifications. Turboprop function somewhat differently. In them, the gas turbine has no mechanical connection with the transmission. These installations drive the aircraft with the help of jet thrust only partially. This type of installation uses the main part of the energy of the hot mixture to drive the propeller through the gearbox. In such an installation, instead of one, there are 2 turbines. One of them drives the compressor, and the second - the screw. Unlike classic turbojet, screw installations are more economical. But they do not allow aircraft to develop high speeds. They are installed on low-speed aircraft. TRDs allow you to develop much greater speed during the flight.

Turbofans engines are combined units that combine elements of turbojet and turboprop engines. They differ from the classic ones in the large size of the fan blades. Both the fan and propeller operate at subsonic speeds. The speed of air movement is reduced due to the presence of a special fairing in which the fan is placed. Such engines consume fuel more economically than classic ones. In addition, they are characterized by higher efficiency. Most often they are installed on liners and large-capacity aircraft.

Aircraft engine size relative to human height

Direct-flow air-jet installations do not involve the use of moving elements. Air is drawn in naturally thanks to a fairing mounted on the inlet. After the intake of air, the engine works similarly to the classic one.

Some aircraft fly on turboprop engines, which are much simpler than turbojet engines. Therefore, many people have a question: why use more complex installations, if you can limit yourself to a screw one? The answer is simple: turbojet engines are superior in power to screw engines. They are ten times more powerful. Accordingly, the turbojet engine produces much more thrust. This makes it possible to lift large aircraft into the air and fly at high speed.

In contact with

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

Turbojet engines

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

Turboprop engines

Device and principle of operation. For modern aircraft

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

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

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

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

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

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

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

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

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

Bypass turbojet engines

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

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

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

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

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

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

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

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 increasing 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, if the flow rate of the HPP cooling blades 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 to reduce the cooling efficiency due to insufficient air pressure 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 effective cooling of the LPT engine working blade 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 the positive qualities of 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 makes it possible either to increase the service life of the turbine and the entire engine at a fixed cooling air flow rate, or to reduce the cooling air flow rate and thereby increase the traction characteristics of the 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 the free energy of the working fluid behind the turbine increases; 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 by the following detailed description of the 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 figure 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) (takeoff 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* N. At high values of the parameter (take-off engine operation), ring 15 is in position I , at low values of the parameter (cruising mode) - in position II.

The implementation of the cooling system in accordance with the claimed technical solution allows you to provide the necessary cooling of the last stage of the low-pressure turbine in all modes of engine operation, while 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.

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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 disc.

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 crown. 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 disc.

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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