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

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

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

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

By principle of air compression WRDs are divided into:

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

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

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

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

By principle of jet thrust WRDs are divided into:

• direct reaction engines;
• indirect reaction engines.

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

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

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

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

and many other combinations of engines of more complex schemes.

Piston engines (PD)

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

piston engine (English) piston engine ) -

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

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

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

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

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

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

Gas turbine engines (GTE)

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

Single-shaft and multi-shaft engines

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

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

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

Turbojet engine (TRD)

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

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

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

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

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

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

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

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

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

Turbojet engine with afterburner (TRDF)

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

Double-circuit turbojet engine (TRDD)

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

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

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

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

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

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

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

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

All turbofan engines can be divided into 2 groups:

• with mixing flows behind the turbine;
• without mixing.

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

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

Military turbofan EJ200 (m=0.4)

Bypass turbojet engine with afterburner (TRDDF)

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

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

TRDDF with deflectable thrust vector (OVT).

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

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

High Bypass Turbofan / Turbofan Engine

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

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

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

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

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

Due to the fact that the fan of such engines, as a rule, has a large diameter, and the degree of air pressure increase in the fan is not high, the nozzle of the external circuit of such engines is 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 significant air resistance in flight.

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

Turbopropfan engine (TVVD)

Turbopropfan engine (English) turbo propfan engine ) -

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

0

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 "turbine" theme is as complex as it is extensive. Therefore, of course, it is not necessary to talk about its full disclosure. Let's deal, as always, with "general acquaintance" and "separate interesting moments" ...

At the same time, the history of the aviation turbine is very short compared to the history of the turbine in general. This means that one cannot do without some theoretical and historical digression, the content of which for the most part does not apply to aviation, but is the basis for a story about the use of a gas turbine in aircraft engines.

About the hum and rumble...

Let's start somewhat unconventionally and remember about "". This is a fairly common phrase used by usually inexperienced authors in the media when describing the operation of powerful aircraft. Here you can also add "roar, whistle" and other loud definitions for all the same "aircraft turbines".

Pretty familiar words for many. However, people who understand are well aware that in fact all these “sound” epithets most often characterize the operation of jet engines as a whole or its parts, which have very little relation to turbines as such (with the exception, of course, of mutual influence during their joint work). in the general cycle of the turbojet engine).

Moreover, in a turbojet engine (just such are the object of rave reviews), as a direct reaction engine that creates thrust by using the reaction of a gas jet, the turbine is just a part of it and is rather indirectly related to the “roaring roar”.

And on those engines where it, like a node, plays, in some way, a dominant role (these are indirect reaction engines, and they are called gas turbine), there is no longer such an impressive sound, or it is created by completely different parts of the power plant of the aircraft, for example, a propeller.

That is, neither the rumble nor the roar, as such, to aviation turbine don't really apply. However, despite such sound ineffectiveness, it is a complex and very important unit of a modern turbojet engine (GTE), often determining its main operational characteristics. Not a single gas turbine engine, simply by definition, can do without a turbine.

Therefore, the conversation, of course, is not about impressive sounds and incorrect use of the definitions of the Russian language, but about an interesting unit and its relation to aviation, although this is far from the only area of ​​\u200b\u200bits application. As a technical device, the turbine appeared long before the very concept of an “aircraft” (or airplane) arose, and even more so a gas turbine engine for it.

History + some theory ...

And even for a very long time. Ever since the invention of mechanisms that convert the energy of the forces of nature into useful action. The simplest in this regard and therefore one of the first to appear were the so-called rotary engines.

This definition itself, of course, appeared only in our days. However, its meaning just determines the simplicity of the engine. Natural energy directly, without any intermediate devices, is converted into the mechanical power of the rotational movement of the main power element of such an engine - the shaft.

Turbine- a typical representative of a rotary engine. Looking ahead, we can say that, for example, in a piston internal combustion engine (ICE), the main element is the piston. It performs a reciprocating motion, and in order to obtain rotation of the output shaft, it is necessary to have an additional crank mechanism, which naturally complicates and makes the structure heavier. Turbine in this regard is much more profitable.

For a rotary type internal combustion engine, as a heat engine, which, by the way, is a turbojet engine, the name “rotary” is usually used.

Turbine wheel of a water mill

One of the most famous and most ancient uses of the turbine are large mechanical mills used by man since time immemorial for various household needs (not just for grinding grain). They are treated as water, and windmills mechanisms.

Throughout a long period of ancient history (the first mention is from about the 2nd century BC) and the history of the Middle Ages, these were in fact the only mechanisms used by man for practical purposes. The possibility of their application, despite the primitiveness of technical circumstances, consisted in the simplicity of transforming the energy of the used working fluid (water, air).

A windmill is an example of a turbine wheel.

In these, in fact, real rotary engines, the energy of the water or air flow is converted into shaft power and, ultimately, useful work. This happens when the flow interacts with the working surfaces, which are water wheel blades or windmill wings. Both, in fact, are the prototype of the blades of modern blade machines, which are currently used turbines (and compressors, by the way, too).

Another type of turbine is known, first documented (apparently invented) by the ancient Greek scientist, mechanic, mathematician and naturalist Heron of Alexandria ( Heron ho Alexandreus,1 -th century AD) in his treatise Pneumatics. The invention he described was called aeolipil , which in Greek means "ball of Eol" (god of the wind, Αἴολος - Eol (Greek), pila- ball (lat.)).

Aeolipil Heron.

In it, the ball was equipped with two oppositely directed tubes-nozzles. Steam came out of the nozzles, which entered the ball through pipes from a boiler located below and thereby forced the ball to rotate. The action is clear from the figure. It was a so-called inverted turbine, rotating in the direction opposite to the steam outlet. Turbines of this type have a special name - reactive (more details - below).

It is interesting that Heron himself hardly imagined what was the working body in his car. In that era, steam was identified with air, even the name testifies to this, because Eol commands the wind, that is, air.

Eolipil was, in general, a full-fledged heat engine that converted the energy of the burned fuel into mechanical energy of rotation on the shaft. Perhaps it was one of the first heat engines in history. True, its usefulness was still “not complete”, since the invention did not perform useful work.

Eolipil, among other mechanisms known at that time, was part of the so-called “automaton theater”, which was very popular in subsequent centuries, and was actually just an interesting toy with an incomprehensible future.

From the moment of its creation and in general from the era when people in their first mechanisms used only “clearly manifesting themselves” forces of nature (wind force or gravity of falling water) until the start of confident use of the thermal energy of fuel in newly created heat engines, more than one hundred passed years.

The first such units were steam engines. Real working examples were invented and built in England only towards the end of the 17th century and were used to pump water from coal mines. Later, steam engines with a piston mechanism appeared.

In the future, with the development of technical knowledge, piston internal combustion engines of various designs, more advanced and more efficient mechanisms, “entered the stage”. They already used gas (combustion products) as a working fluid and did not require bulky steam boilers to heat it.

Turbines as the main components of thermal engines, also went through a similar path in their development. And although there are separate mentions of some instances in history, but deserving attention and, moreover, documented, including patented, units appeared only in the second half of the 19th century.

It all started with a couple...

It was with the use of this working fluid that almost all the basic principles of the turbine design (later gas turbine) were worked out as an important part of the heat engine.

Jet turbine patented by Laval.

Quite characteristic in this regard were the developments of a talented Swedish engineer and inventor Gustave de Laval(Karl Gustaf Patrik de Laval). His research at that time was connected with the idea of ​​developing a new milk separator with increased drive speed, which made it possible to significantly increase productivity.

It was not possible to obtain a higher rotational speed (revolutions) by using the already traditional (however, the only existing) reciprocating steam engine due to the large inertia of the most important element - the piston. Realizing this, Laval decided to try to abandon the use of the piston.

It is said that the idea itself came to him while observing the work of sandblasters. In 1883 he received his first patent (English Patent No. 1622) in this area. The patented device was called " Turbine powered by steam and water».

It was an S-shaped tube, at the ends of which tapering nozzles were made. The tube was mounted on a hollow shaft through which steam was supplied to the nozzles. In principle, all this did not differ in any way from the eolipil of Heron of Alexandria.

The manufactured device worked quite reliably with high revolutions for the technology of that time - 42,000 rpm. The rotation speed reached 200 m/s. But with such good parameters turbine had extremely low efficiency. And attempts to increase it with the existing state of the art did not lead to anything. Why did it happen?

——————-

A little theory ... A little more about the features ....

The mentioned efficiency factor (for modern aircraft turbines, this is the so-called power or effective efficiency factor) characterizes the efficiency of using the energy expended (available) to drive the turbine shaft. That is, what part of this energy was spent usefully on the rotation of the shaft, and what " went down the pipe».

It just took off. For the type of turbine described, called reactive, this expression is just right. Such a device receives a rotational movement on the shaft under the action of the reaction force of the outgoing gas jet (or in this case, steam).

A turbine, as a dynamic expansion machine, unlike volumetric machines (reciprocating machines), requires for its operation not only compression and heating of the working fluid (gas, steam), but also its acceleration. Here, expansion (increase in specific volume) and pressure drop occur due to acceleration, in particular in the nozzle. In a piston engine, this is due to an increase in the volume of the cylinder chamber.

As a result, that large potential energy of the working fluid, which was formed as a result of the supply of thermal energy of the burnt fuel to it, turns into kinetic energy (minus various losses, of course). And kinetic (in a jet turbine) through reaction forces - into mechanical work on the shaft.

And that's about how fully the kinetic energy goes into mechanical in this situation and tells us the efficiency. The higher it is, the less kinetic energy the flow leaving the nozzle into the environment has. This remaining energy is called " loss with output speed”, and it is directly proportional to the square of the speed of the outgoing stream (everyone probably remembers mС 2 /2).

The principle of operation of a jet turbine.

Here we are talking about the so-called absolute speed C. After all, the outgoing flow, more precisely, each of its particles, participates in a complex movement: rectilinear plus rotational. Thus, the absolute speed C (relative to a fixed coordinate system) is equal to the sum of the turbine rotation speed U and the relative flow speed W (speed relative to the nozzle). The sum is of course vector, shown in the figure.

Segner wheel.

Minimum losses (and maximum efficiency) correspond to the minimum speed C, ideally, it should be equal to zero. And this is possible only if W and U are equal (it can be seen from the figure). The peripheral speed (U) in this case is called optimal.

It would be easy to ensure such equality on hydraulic turbines (such as segner wheel), since the rate of fluid outflow from the nozzles for them (similar to the velocity W) is relatively low.

But the same velocity W for gas or vapor is much greater due to the large difference in the densities of liquid and gas. So, at a relatively low pressure of only 5 atm. a hydraulic turbine can give an exhaust velocity of only 31 m/s, and a steam turbine 455 m/s. That is, it turns out that even at sufficiently low pressures (only 5 atm.), Laval's jet turbine should have, for reasons of high efficiency, a peripheral speed above 450 m / s.

For the then level of development of technology, this was simply impossible. It was impossible to make a reliable design with such parameters. To reduce the optimal circumferential speed by reducing the relative (W) also did not make sense, since this can only be done by reducing the temperature and pressure, and hence the overall efficiency.

Laval active turbine...

Laval's jet turbine did not succumb to further improvement. Despite the attempts made, things came to a standstill. Then the engineer took a different path. In 1889, he patented a different type of turbine, which later received the name active. Abroad (in English) it now bears the name impulse turbine, that is, impulsive.

The device claimed in the patent consisted of one or more fixed nozzles supplying steam to bucket-shaped blades mounted on the rim of a movable working turbine wheel (or disk).

Active single-stage steam turbine patented by Laval.

The working process in such a turbine is as follows. The steam accelerates in the nozzles with an increase in kinetic energy and a drop in pressure and falls on the rotor blades, on their concave part. As a result of the impact on the blades of the impeller, it begins to rotate. Or else you can say that the rotation occurs due to the impulsive action of the jet. Hence the English name impulseturbine.

At the same time, in the interblade channels, which have a practically constant cross section, the flow does not change its speed (W) and pressure, but changes direction, that is, it turns at large angles (up to 180°). That is, we have at the exit from the nozzle and at the entrance to the interblade channel: absolute speed C 1 , relative W 1 , circumferential speed U.

At the output, respectively, C 2, W 2, and the same U. In this case, W 1 \u003d W 2, C 2< С 1 – из-за того, что часть кинетической энергии входящего потока превращается в механическую на валу турбины (импульсное воздействие) и абсолютная скорость падает.

In principle, this process is shown in a simplified figure. Also, to simplify the explanation of the process, it is assumed here that the absolute and circumferential velocity vectors are practically parallel, the flow changes direction in the impeller by 180°.

The flow of steam (gas) in the stage of an active turbine.

If we consider the speeds in absolute terms, then it can be seen that W 1 \u003d C 1 - U, and C 2 \u003d W 2 - U. Thus, based on the foregoing, for the optimal mode, when the efficiency takes maximum values, and losses from the output speed tend to a minimum (that is, C 2 =0) we have C 1 =2U or U=C 1 /2.

We get that for an active turbine optimum circumferential speed half the speed of the outflow from the nozzle, that is, such a turbine is half as loaded as a jet turbine, and the task of obtaining a higher efficiency is facilitated.

Therefore, in the future, Laval continued to develop just this type of turbine. However, despite the reduction in the required circumferential speed, it still remained large enough, which entailed equally large centrifugal and vibration loads.

The principle of operation of an active turbine.

This resulted in structural and strength problems, as well as problems of eliminating imbalances, which were often solved with great difficulty. In addition, there were other unresolved and unsolvable factors in the conditions of that time, which ultimately reduced the efficiency of this turbine.

These included, for example, the imperfection of the aerodynamics of the blades, causing increased hydraulic losses, as well as the pulsating effect of individual steam jets. In fact, only a few or even one blade could be active blades perceiving the action of these jets (or jets) at the same time. The rest at the same time moved idly, creating additional resistance (in a vapor atmosphere).

Such turbines there was no way to increase power due to an increase in temperature and steam pressure, as this would lead to an increase in peripheral speed, which was absolutely unacceptable due to all the same design problems.

In addition, the increase in power (with an increase in peripheral speed) was inappropriate for another reason. The energy consumers of the turbine were low-speed devices compared to it (electric generators were planned for this). Therefore, Laval had to develop special gearboxes for the kinematic connection of the turbine shaft with the consumer shaft.

The ratio of the masses and dimensions of the active Laval turbine and the gearbox to it.

Due to the large difference in the speed of these shafts, the gearboxes were extremely bulky and often significantly exceeded the turbine itself in size and weight. An increase in its power would entail an even greater increase in the size of such devices.

Eventually Laval active turbine It was a relatively low-power unit (working specimens up to 350 hp), moreover, expensive (due to a large set of improvements), and complete with a gearbox, it was also quite bulky. All this made it uncompetitive and excluded mass application.

A curious fact is that the constructive principle of Laval's active turbine was actually not invented by him. Even 250 years before the appearance of his research in Rome in 1629, a book by the Italian engineer and architect Giovanni Branca was published under the title "Le Machine" ("Machines").

In it, among other mechanisms, a description of the “steam wheel” was placed, containing all the main components built by Laval: a steam boiler, a steam supply tube (nozzle), an active turbine impeller, and even a gearbox. Thus, long before Laval, all these elements were already known, and his merit lay in the fact that he made them all really work together and dealt with extremely complex issues of improving the mechanism as a whole.

Steam active turbine Giovanni Branca.

Interestingly, one of the most famous features of his turbine was the design of the nozzle (it was mentioned separately in the same patent), which supplies steam to the rotor blades. Here, the nozzle from an ordinary tapering one, as it was in a jet turbine, became narrowing-expanding. Subsequently, this type of nozzle came to be called Laval nozzles. They make it possible to accelerate the flow of gas (steam) to supersonic speed with sufficiently small losses. About them .

Thus, the main problem that Laval struggled with when developing his turbines, and which he could not cope with, was high peripheral speed. However, a fairly effective solution to this problem has already been proposed and even, oddly enough, by Laval himself.

Multi-stage….

In the same year (1889), when the above-described active turbine was patented, an engineer developed an active turbine with two parallel rows of rotor blades mounted on one impeller (disk). This was the so-called two-stage turbine.

Steam was supplied to the working blades, as in the single-stage one, through the nozzle. Between the two rows of rotor blades, a row of fixed blades was installed, which redirected the flow leaving the blades of the first stage to the rotor blades of the second.

If we use the simplified principle proposed above for determining the circumferential velocity for a single-stage jet turbine (Laval), then it turns out that for a two-stage turbine, the rotation speed is less than the speed of the outflow from the nozzle not by two, but by four times.

The principle of the Curtis wheel and changing the parameters in it.

This is the most effective solution to the problem of low optimum circumferential speed, which was proposed but not used by Laval and which is actively used in modern turbines, both steam and gas. Multistage…

It means that the large available energy for the entire turbine can be divided in some way into parts according to the number of stages, and each such part is worked out in a separate stage. The lower this energy, the lower the speed of the working fluid (steam, gas) entering the rotor blades and, consequently, the lower the optimal circumferential speed.

That is, by changing the number of turbine stages, you can change the frequency of rotation of its shaft and, accordingly, change the load on it. In addition, multi-stage allows you to work on the turbine large differences in energy, that is, to increase its power, and at the same time maintain high efficiency rates.

Laval did not patent his two-stage turbine, although a prototype was made, so it bears the name of the American engineer C. Curtis (wheel (or disk) Curtis), who in 1896 received a patent for a similar device.

However, much earlier, in 1884, the English engineer Charles Algernon Parsons developed and patented the first real multistage steam turbine. There were many statements by various scientists and engineers about the usefulness of dividing the available energy into steps before him, but he was the first to translate the idea into "iron".

Parsons multi-stage active-jet turbine (disassembled).

At the same time, his turbine had a feature that brought it closer to modern devices. In it, steam expanded and accelerated not only in nozzles formed by fixed blades, but also partially in channels formed by specially shaped rotor blades.

It is customary to call this type of turbine a reactive one, although the name is rather arbitrary. In fact, it occupies an intermediate position between the purely reactive Heron-Laval turbine and the purely active Laval-Branca. The rotor blades, due to their design, combine active and reactive principles in the overall process. Therefore, it would be more correct to call such a turbine active-reactive which is often done.

Diagram of a multistage Parsons turbine.

Parsons worked on various types of multistage turbines. Among his designs were not only the above-described axial (the working fluid moves along the axis of rotation), but also radial (steam moves in the radial direction). Quite well known is his three-stage purely active turbine "Heron", in which the so-called Heron's wheels are used (the essence is the same as that of the aeolipil).

Jet turbine "Heron".

Later, from the early 1900s, steam turbine building rapidly gained momentum and Parsons was at the forefront of it. Its multi-stage turbines were equipped with sea vessels, first experimental (ship "Turbinia", 1896, displacement 44 tons, speed 60 km / h - unprecedented for that time), then military ones (for example, the battleship "Dreadnought", 18000 tons, speed 40 km / h). h, turbine power 24,700 hp) and passenger (example - the same type "Mauritania" and "Lusitania", 40,000 tons, speed 48 km / h, turbine power 70,000 hp). At the same time, stationary turbine construction began, for example, by installing turbines as drives in power plants (Edison Company in Chicago).

About gas turbines...

However, let's return to our main topic - aviation and note one fairly obvious thing: such a clearly marked success in the operation of steam turbines could have only constructive and fundamental significance for aviation, which was rapidly progressing in its development just at the same time.

The use of a steam turbine as a power plant in aircraft, for obvious reasons, was extremely doubtful. Aviation turbine only a fundamentally similar, but much more profitable gas turbine could become. However, it wasn't all that easy...

According to Lev Gumilevsky, the author of the popular book “The Engine Makers” in the 60s, once, in 1902, at the beginning of the rapid development of steam turbine building, Charles Parsons, in fact one of the then main ideologists of this business, was asked, in general, joking question: Is it possible to "parsonize" a gas engine?”(implied turbine).

The answer was expressed in an absolutely decisive form: “ I think that a gas turbine will never be created. No two ways about it." The engineer failed to become a prophet, but he certainly had reason to say so.

The use of a gas turbine, especially if we mean its use in aviation instead of steam, of course, was tempting, because its positive aspects are obvious. With all its power capabilities, it does not need huge, bulky devices for creating steam - boilers and also no less large devices and systems for its cooling - condensers, cooling towers, cooling ponds, etc. to work.

The heater for a gas turbine engine is a small, compact one, located inside the engine and burning fuel directly in the air stream. He doesn't even have a refrigerator. Or rather, it exists, but exists as if virtually, because the exhaust gas is discharged into the atmosphere, which is the refrigerator. That is, there is everything you need for a heat engine, but at the same time everything is compact and simple.

True, a steam turbine plant can also do without a “real refrigerator” (without a condenser) and release steam directly into the atmosphere, but then you can forget about efficiency. An example of this is a steam locomotive - the real efficiency is about 6%, 90% of its energy flies into the pipe.

But with such tangible pluses, there are also significant drawbacks, which, in general, became the basis for Parsons' categorical answer.

Compression of the working fluid for the subsequent implementation of the working cycle, incl. and in the turbine...

In the operating cycle of a steam turbine plant (Rankine cycle), the work of compressing water is small and the demands on the pump that performs this function and its efficiency are therefore also small. In the GTE cycle, where air is compressed, this work, on the contrary, is very impressive, and most of the available energy of the turbine is spent on it.

This reduces the amount of useful work that the turbine can be used for. Therefore, the requirements for the air compression unit in terms of its efficiency and economy are very high. Compressors in modern aircraft gas turbine engines (mainly axial), as well as in stationary units, along with turbines, are complex and expensive devices. About them .

Temperature…

This is the main problem for gas turbines, including aviation ones. The fact is that if in a steam turbine plant the temperature of the working fluid after the expansion process is close to the temperature of the cooling water, then in a gas turbine it reaches a value of several hundred degrees.

This means that a large amount of energy is emitted into the atmosphere (like a refrigerator), which, of course, adversely affects the efficiency of the entire operating cycle, which is characterized by thermal efficiency: η t \u003d Q 1 - Q 2 / Q 1. Here Q 2 is the same energy discharged into the atmosphere. Q 1 - energy supplied to the process from the heater (in the combustion chamber).

In order to increase this efficiency, it is necessary to increase Q 1, which is equivalent to increasing the temperature in front of the turbine (that is, in the combustion chamber). But the fact of the matter is that it is far from always possible to raise this temperature. Its maximum value is limited by the turbine itself and strength becomes the main condition here. The turbine operates under very difficult conditions, when high temperatures are combined with high centrifugal loads.

It is this factor that has always limited the power and thrust capabilities of gas turbine engines (largely dependent on temperature) and often became the reason for the complexity and cost of turbines. This situation has continued in our time.

And in Parsons' time, neither the metallurgical industry nor the science of aerodynamics could yet provide a solution to the problems of creating an efficient and economical compressor and high-temperature turbine. There was neither an appropriate theory nor the necessary heat-resistant and heat-resistant materials.

And yet there have been attempts...

Nevertheless, as it usually happens, there were people who are not afraid (or maybe not understanding :-)) of possible difficulties. Attempts to create a gas turbine did not stop.

Moreover, it is interesting that Parsons himself, at the dawn of his “turbine” activity, in his first patent for a multistage turbine, noted the possibility of its operation, in addition to steam, also on fuel combustion products. A possible variant of a gas turbine engine operating on liquid fuel with a compressor, a combustion chamber and a turbine was also considered there.

Smoke spit.

Examples of the use of gas turbines without subsuming any theory have been known for a long time. Apparently, even Heron in the "theater of automata" used the principle of an air jet turbine. The so-called "smoke skewers" are widely known.

And in the already mentioned book by the Italian (engineer, architect, Giovanni Branca, Le Machine) Giovanni Branca there is a drawing “ fire wheel". In it, the turbine wheel is rotated by the products of combustion from the fire (or hearth). Interestingly, Branca himself did not build most of his machines, but only expressed ideas for their creation.

The Fire Wheel by Giovanni Branca.

In all these "smoke and fire wheels" there was no air (gas) compression stage, and there was no compressor as such. The transformation of potential energy, that is, the supplied thermal energy of fuel combustion, into kinetic (acceleration) for the rotation of a gas turbine occurred only due to the action of gravity when the warm masses rose up. That is, the phenomenon of convection was used.

Of course, such "units" for real machines, for example, could not be used to drive vehicles. However, in 1791, the Englishman John Barber patented a “horseless transport machine”, one of the most important components of which was a gas turbine. It was the first officially registered gas turbine patent in history.

John Barber gas turbine engine.

The machine used gas obtained from wood, coal or oil, heated in special gas generators (retorts), which, after cooling, entered the reciprocating compressor, where it was compressed together with air. Next, the mixture was fed into the combustion chamber, and after that the combustion products were rotated turbine. Water was used to cool the combustion chambers, and the resulting steam was also sent to the turbine.

The level of development of the then technologies did not allow to bring the idea to life. The working model of the Barber machine with a gas turbine was built only in 1972 by Kraftwerk-Union AG for the Hanover Industrial Exhibition.

Throughout the 19th century, the development of the gas turbine concept was extremely slow for the reasons described above. There were few samples worthy of attention. The compressor and heat remained an insurmountable stumbling block. There have been attempts to use a fan to compress air, as well as the use of water and air to cool structural elements.

Engine F. Stolze. 1 - axial compressor, 2 - axial turbine, 3 - heat exchanger.

An example of a gas turbine engine by the German engineer Franz Stolze, patented in 1872 and very similar in design to modern gas turbine engines, is known. In it, a multi-stage axial compressor and a multi-stage axial turbine were located on the same shaft.

The air after passing through the regenerative heat exchanger was divided into two parts. One entered the combustion chamber, the second was mixed with the combustion products before they entered the turbine, reducing their temperature. This so-called secondary air, and its use is a technique widely used in modern gas turbine engines.

The Stolze engine was tested in 1900-1904, but it turned out to be extremely inefficient due to the low quality of the compressor and the low temperature in front of the turbine.

For most of the first half of the 20th century, the gas turbine was not able to actively compete with the steam turbine or become part of the gas turbine engine, which could adequately replace the reciprocating internal combustion engine. Its use on engines was mainly auxiliary. For example, as pressurization units in piston engines, including aviation ones.

But from the beginning of the 1940s, the situation began to change rapidly. Finally, new heat-resistant alloys were created, which made it possible to radically raise the temperature of the gas in front of the turbine (up to 800 ° C and higher), and quite economical ones with high efficiency appeared.

This not only made it possible to build efficient gas turbine engines, but also, due to the combination of their power with relative lightness and compactness, to use them on aircraft. The era of jet aircraft and aircraft gas turbine engines began.

Turbines in aircraft gas turbine engines ...

So ... The main area of ​​application of turbines in aviation is gas turbine engines. The turbine here does the hard work - it rotates the compressor. At the same time, in a gas turbine engine, as in any heat engine, the work of expansion is greater than the work of compression.

And the turbine is just an expansion machine, and it consumes only a part of the available energy of the gas flow for the compressor. The remainder (sometimes referred to as free energy) can be used for useful purposes depending on the type and design of the engine.

Scheme TVAD Makila 1a1 with a free turbine.

Turboshaft engine AMAKILA 1A1.

For indirect reaction engines, such as (helicopter GTE), it is spent on the rotation of the propeller. In this case, the turbine is most often divided into two parts. The first one is compressor turbine. The second one, which drives the screw, is the so-called free turbine. It rotates independently and is only gas-dynamically connected to the compressor turbine.

In direct reaction engines (jet engines or VREs), the turbine is only used to drive the compressor. The remaining free energy, which rotates a free turbine in the TVAD, is used up in the nozzle, turning into kinetic energy to obtain jet thrust.

In the middle between these extremes are located. Some of their free energy is used to drive the propeller, and some of it forms jet thrust in the output device (nozzle). True, its share in the total thrust of the engine is small.

Scheme of a single-shaft theater DART RDa6. Turbine on a common shaft of the engine.

Turboprop single-shaft engine Rolls-Royce DART RDa6.

By design, HPTs can be single-shaft, in which the free turbine is not structurally allocated and, being one unit, drives both the compressor and the propeller at once. An example of a Rolls-Royce DART RDa6 TVD, as well as our well-known AI-20 TVD.

There may also be a TVD with a separate free turbine that drives the propeller and is not mechanically connected to the rest of the engine components (gas-dynamic connection). An example is the PW127 engine of various modifications (aircraft), or the Pratt & Whitney Canada PT6A theater.

Scheme of the Pratt & Whitney Canada PT6A theater with a free turbine.

Pratt & Whitney Canada PT6A engine.

Scheme of a PW127 TVD with a free turbine.

Of course, in all types of gas turbine engines, the payload also includes units that ensure the operation of the engine and aircraft systems. These are usually pumps, fuel and hydro-, electric generators, etc. All these devices are most often driven from the turbocharger shaft.

On the types of turbines.

There are actually quite a few types. Just for example, some names: axial, radial, diagonal, radial-axial, rotary-blade, etc. In aviation, only the first two are used, and radial is quite rare. Both of these turbines were named in accordance with the nature of the movement of the gas flow in them.

Radial.

In radial it flows along the radius. Moreover, in the radial aviation turbine centripetal flow direction is used, which provides higher efficiency (in non-aviation practice, there is also centrifugal).

The stage of a radial turbine consists of an impeller and fixed blades that form the flow at its inlet. The blades are profiled so that the interblade channels have a tapering configuration, that is, they are nozzles. All these blades, together with the body elements on which they are mounted, are called nozzle apparatus.

Scheme of a radial centripetal turbine (with explanations).

The impeller is an impeller with specially profiled blades. The spinning of the impeller occurs when the gas passes through the narrowing channels between the blades and acts on the blades.

The impeller of a radial centripetal turbine.

Radial turbines are quite simple, their impellers have a small number of blades. The possible circumferential speeds of a radial turbine at the same stresses in the impeller are greater than those of an axial turbine, therefore, larger amounts of energy (heat drops) can be generated on it.

However, these turbines have a small flow area and do not provide sufficient gas flow for the same size compared to axial turbines. In other words, they have too large relative diametrical dimensions, which complicates their arrangement in a single engine.

In addition, it is difficult to create multi-stage radial turbines due to large hydraulic losses, which limits the degree of gas expansion in them. It is also difficult to cool such turbines, which reduces the possible maximum gas temperatures.

Therefore, the use of radial turbines in aviation is limited. They are mainly used in low-power units with low gas consumption, most often in auxiliary mechanisms and systems or in engines of model aircraft and small unmanned aircraft.

The first Heinkel He 178 jet aircraft.

TRD Heinkel HeS3 with a radial turbine.

One of the few examples of the use of a radial turbine as a main air jet engine is the engine of the first real jet aircraft, the Heinkel He 178 turbojet Heinkel HeS 3. The photo clearly shows the elements of the stage of such a turbine. The parameters of this engine were quite consistent with the possibility of its use.

Axial aviation turbine.

This is the only type of turbine currently used in sustainer aviation gas turbine engines. The main source of mechanical work on the shaft obtained from such a turbine in the engine are impellers or, more precisely, rotor blades (RL) mounted on these wheels and interacting with an energetically charged gas flow (compressed and heated).

The rims of fixed blades installed in front of the workers organize the correct direction of the flow and participate in the conversion of the potential energy of the gas into kinetic energy, that is, they accelerate it in the process of expansion with a drop in pressure.

These blades, complete with the body elements on which they are mounted, are called nozzle apparatus(SA). The nozzle apparatus complete with working blades is turbine stage.

The essence of the process ... Generalization of what has been said ...

In the process of the above interaction with the rotor blades, the kinetic energy of the flow is converted into mechanical energy that rotates the engine shaft. Such a transformation in an axial turbine can occur in two ways:

An example of a single-stage active turbine. The change of parameters along the path is shown.

1. Without changing the pressure, and hence the magnitude of the relative flow rate (only its direction changes noticeably - the turn of the flow) in the turbine stage; 2. With a drop in pressure, an increase in the relative flow velocity and some change in its direction in the stage.

Turbines operating according to the first method are called active. The gas flow actively (impulsively) acts on the blades due to a change in its direction as it flows around them. In the second way - jet turbines. Here, in addition to the impulse action, the flow affects the rotor blades also indirectly (to put it simply), with the help of a reactive force, which increases the power of the turbine. Additional reactive action is achieved due to the special profiling of the rotor blades.

The concepts of activity and reactivity in general, for all turbines (not only aviation ones) were mentioned above. However, modern aircraft gas turbine engines use only axial jet turbines.

Change of parameters in the stage of an axial gas turbine.

Since the force effect on the radar is double, such axial turbines are also called active-reactive which is perhaps more correct. This type of turbine is more advantageous in terms of aerodynamics.

The stationary blades of the nozzle apparatus included in the stage of such a turbine have a large curvature, due to which the cross section of the interblade channel decreases from inlet to outlet, that is, the section f 1 is less than the section f 0 . It turns out the profile of a tapering jet nozzle.

The working blades following them also have a large curvature. In addition, with respect to the oncoming flow (vector W 1), they are located in such a way as to avoid its stall and ensure the correct flow around the blade. At certain radii, the RL also form narrowing interscapular channels.

Step work aviation turbine.

The gas approaches the nozzle apparatus with a direction of movement close to axial and a speed of C 0 (subsonic). Pressure in the flow Р 0 , temperature Т 0 . Passing the interblade channel, the flow accelerates to speed C 1 with a turn to an angle α 1 = 20°-30°. In this case, the pressure and temperature fall to the values ​​of P 1 and T 1, respectively. Part of the potential energy of the flow is converted into kinetic energy.

Pattern of gas flow movement in the stage of an axial turbine.

Since the working blades move with peripheral speed U, the flow enters the interblade channel of the RL already at a relative speed W 1, which is determined by the difference between C 1 and U (vector). Passing through the channel, the flow interacts with the blades, creating aerodynamic forces P on them, the circumferential component of which P u makes the turbine rotate.

Due to the narrowing of the channel between the blades, the flow accelerates to the speed W 2 (reactive principle), while it also turns (active principle). The absolute flow rate C 1 decreases to C 2 - the kinetic energy of the flow is converted into mechanical energy on the turbine shaft. The pressure and temperature drop to P 2 and T 2 , respectively.

The absolute flow rate during the passage of the stage slightly increases from C 0 to the axial projection of the velocity C 2 . In modern turbines, this projection has a value of 200-360 m/s for a stage.

The step is profiled so that the angle α 2 is close to 90°. The difference is usually 5-10°. This is done so that the value of C 2 is minimal. This is especially important for the last stage of the turbine (on the first or middle stages, a deviation from a right angle of up to 25 ° is allowed). The reason for that is loss with output speed, which just depend on the magnitude of the velocity C 2 .

These are the same losses that at one time did not give Laval the opportunity to increase the efficiency of his first turbine. If the engine is reactive, then the remaining energy can be generated in the nozzle. But, for example, for a helicopter engine that does not use jet propulsion, it is important that the flow velocity behind the last stage of the turbine is as low as possible.

Thus, in the stage of an active-jet turbine, gas expansion (pressure and temperature reduction), energy conversion and operation (heat drop) occur not only in the SA, but also in the impeller. The distribution of these functions between the RC and SA characterizes the parameter of the theory of engines, called degree of reactivity ρ.

It is equal to the ratio of the heat drop in the impeller to the heat drop in the entire stage. If ρ = 0, then the stage (or the entire turbine) is active. If ρ > 0, then the stage is reactive or, more precisely, for our case, active-reactive. Since the profile of the rotor blades varies along the radius, this parameter (as well as some others) is calculated according to the average radius (section В-В in the figure of changing parameters in the stage).

The configuration of the pen of the working blade of an active-jet turbine.

Change in pressure along the length of the radar pen of an active-jet turbine.

For modern gas turbine engines, the degree of reactivity of turbines is in the range of 0.3-0.4. This means that only 30-40% of the total heat drop of the stage (or turbine) is exhausted in the impeller. 60-70% is worked out in the nozzle apparatus.

Something about losses.

As already mentioned, any turbine (or its stage) converts the flow energy supplied to it into mechanical work. However, in a real unit, this process may have different efficiency. Part of the available energy is necessarily wasted, that is, it turns into losses, which must be taken into account and measures must be taken to minimize them in order to increase the efficiency of the turbine, that is, increase its efficiency.

Losses are made up of hydraulic and loss with output speed. Hydraulic losses include profile and end losses. Profile is, in fact, friction losses, since the gas, having a certain viscosity, interacts with the surfaces of the turbine.

Typically, such losses in the impeller are about 2-3%, and in the nozzle apparatus - 3-4%. Measures to reduce losses are to "ennoble" the flow path by calculation and experiment, as well as the correct calculation of the velocity triangles for the flow in the turbine stage, more precisely, the choice of the most advantageous circumferential velocity U at a given velocity C 1 . These actions are usually characterized by the parameter U/C 1 . The circumferential speed at the average radius in the turbojet engine is 270 - 370 m/s.

The hydraulic perfection of the flow part of the turbine stage takes into account such a parameter as adiabatic efficiency. Sometimes it is also called bladed, because it takes into account friction losses in the stage blades (SA and RL). There is another efficiency factor for the turbine, which characterizes it precisely as a unit for generating power, that is, the degree of use of available energy to create work on the shaft.

This so-called power (or effective) efficiency. It is equal to the ratio of work on the shaft to the available heat drop. This efficiency takes into account losses with the output speed. They usually make up about 10-12% for turbojet engines (in modern turbojet engines C 0 = 100-180 m/s, C 1 = 500-600 m/s, C 2 = 200-360 m/s).

For turbines of modern gas turbine engines, the value of the adiabatic efficiency is about 0.9 - 0.92 for uncooled turbines. If the turbine is cooled, then this efficiency can be lower by 3-4%. Power efficiency is usually 0.78 - 0.83. It is less than adiabatic by the amount of losses with the output speed.

As for the end losses, these are the so-called " leakage losses". The flow part cannot be completely isolated from the rest of the engine due to the presence of rotating assemblies in combination with fixed ones (casings + rotor). Therefore, gas from areas of high pressure tends to flow into areas of low pressure. In particular, for example, from the area in front of the working blade to the area behind it through the radial gap between the blade airfoil and the turbine casing.

Such a gas does not participate in the process of converting the flow energy into mechanical energy, because it does not interact with the blades in this regard, that is, there are end losses (or radial clearance loss). They make up about 2-3% and negatively affect both the adiabatic and power efficiency, reduce the efficiency of the gas turbine engine, and quite noticeably.

It is known, for example, that an increase in the radial clearance from 1 mm to 5 mm in a turbine with a diameter of 1 m can lead to an increase in the specific fuel consumption in the engine by more than 10%.

It is clear that it is impossible to completely get rid of the radial clearance, but they try to minimize it. It's hard enough because aviation turbine- the unit is heavily loaded. Accurate consideration of all factors affecting the size of the gap is quite difficult.

The engine operating modes often change, which means that the deformations of the rotor blades, the disks on which they are fixed, and the turbine housings change as a result of changes in temperature, pressure and centrifugal forces.

labyrinth seal.

Here it is necessary to take into account the value of residual deformation during long-term operation of the engine. Plus, the evolutions performed by the aircraft affect the deformation of the rotor, which also changes the size of the gaps.

The clearance is usually assessed after the warm engine is stopped. In this case, the thin outer casing cools faster than the massive discs and shaft and, decreasing in diameter, touches the blades. Sometimes the value of the radial clearance is simply chosen in the range of 1.5-3% of the length of the blade airfoil.

The principle of honeycomb sealing.

In order to avoid damage to the blades, if they touch the turbine housing, special inserts are often placed in it from a material that is softer than the material of the blades (for example, cermet). In addition, non-contact seals are used. These are usually labyrinthine or honeycomb labyrinth seals.

In this case, the working blades are shrouded at the ends of the airfoil, and seals or wedges (for honeycombs) are already placed on the shroud shelves. In honeycomb seals, due to the thin walls of the honeycomb, the contact area is very small (10 times smaller than a conventional labyrinth), so the assembly of the assembly is carried out without a gap. After running in, the gap is about 0.2 mm.

Application of honeycomb seal. Comparison of losses when using honeycombs (1) and a smooth ring (2).

Similar gap sealing methods are used to reduce gas leakage from the flow path (for example, into the interdisk space).

SAURZ…

These are the so-called passive methods radial clearance control. In addition, on many gas turbine engines developed (and being developed) since the late 80s, the so-called " systems for active regulation of radial clearances» (SAURZ - active method). These are automatic systems, and the essence of their work is to control the thermal inertia of the housing (stator) of an aircraft turbine.

The rotor and stator (outer casing) of the turbine differ from each other in material and in “massiveness”. Therefore, in transient regimes, they expand in different ways. For example, when the engine is switched from a reduced operating mode to an increased one, a high-temperature, thin-walled housing warms up and expands faster (than a massive rotor with disks), increasing the radial clearance between itself and the blades. Plus, pressure changes in the tract and the evolution of the aircraft.

To avoid this, an automatic system (usually the main regulator of the FADEC type) organizes the supply of cooling air to the turbine housing in the required quantities. The heating of the housing is thus stabilized within the required limits, which means that the value of its linear expansion and, accordingly, the value of the radial clearances change.

All this allows saving fuel, which is very important for modern civil aviation. SAURZ systems are most effectively used in low-pressure turbines on turbojet engines of the GE90, Trent 900, and some other types.

Much less often, but quite effectively, forced blowing of the turbine disks (rather than the housing) is used to synchronize the rates of heating of the rotor and stator. Such systems are used on CF6-80 and PW4000 engines.

———————-

In the turbine, axial clearances are also regulated. For example, between the output edges of the SA and the input RL, there is usually a gap within 0.1-0.4 of the RL chord at the average radius of the blades. The smaller this gap, the lower the flow energy loss behind the SA (for friction and equalization of the velocity field behind the SA). But at the same time, the vibration of the RL increases due to the alternate hit from the areas behind the bodies of the SA blades to the interblade areas.

A little about the design...

Axial aviation turbines modern gas turbine engines in a constructive plan can have different flow path shape.

Dav = (Din+Dn) /2

1. Form with a constant body diameter (Dn). Here, the inner and average diameters along the path are reduced.

Constant outside diameter.

Such a scheme fits well into the dimensions of the engine (and the aircraft fuselage). It has a good distribution of work in stages, especially for twin-shaft turbojet engines.

However, in this scheme, the so-called bell angle is large, which is fraught with flow separation from the inner walls of the housing and, consequently, hydraulic losses.

Constant inside diameter.

When designing, they try not to allow the angle of the socket to be more than 20 °.

2. Form with a constant inner diameter (Dv).

The average diameter and body diameter increase along the path. Such a scheme does not fit well into the dimensions of the engine. In a turbojet engine, due to the "run-up" of the flow from the inner casing, it is necessary to turn it on the SA, which entails hydraulic losses.

Constant average diameter.

The scheme is more appropriate for use in turbofan engines.

3. Form with a constant average diameter (Dav). The diameter of the body increases, the inner diameter decreases.

The scheme has the disadvantages of the previous two. But at the same time, the calculation of such a turbine is quite simple.

Modern aircraft turbines are most often multistage. The main reason for this (as mentioned above) is the large available energy of the turbine as a whole. To ensure an optimal combination of circumferential speed U and speed C 1 (U / C 1 - optimal), and therefore high overall efficiency and good economy, it is necessary to distribute all available energy in steps.

An example of a three-stage turbojet turbine.

At the same time, however, she turbine structurally more complex and heavier. Due to the small temperature difference in each stage (spread across all stages), more of the first stages are exposed to high temperatures and often require additional cooling.

Four-stage axial turbine TVD.

Depending on the type of engine, the number of stages may be different. For turbojet engines, usually up to three, for bypass engines up to 5-8 steps. Usually, if the engine is multi-shaft, then the turbine has several (according to the number of shafts) cascades, each of which drives its own unit and can itself be multi-stage (depending on the degree of bypass).

Twin-shaft axial aircraft turbine.

For example, in the Rolls-Royce Trent 900 three-shaft engine, the turbine has three stages: one stage for driving the high pressure compressor, one stage for driving the intermediate compressor, and five stages for driving the fan. The joint operation of cascades and the determination of the required number of stages in cascades is described separately in the "engine theory".

Herself aviation turbine, to put it simply, is a structure consisting of a rotor, a stator and various auxiliary structural elements. The stator consists of an outer housing, housings nozzle devices and rotor bearing housings. The rotor is usually a disk structure in which the disks are connected to the rotor and to each other using various additional elements and fastening methods.

An example of a single-stage turbojet turbine. 1 - shaft, 2 - SA blades, 3 - impeller disk, 4 - rotor blades.

On each disk, as the basis of the impeller, there are working blades. When designing the blades, they try to perform with a smaller chord due to the smaller width of the disk rim on which they are installed, which reduces its mass. But at the same time, in order to maintain the parameters of the turbine, it is necessary to increase the length of the feather, which may entail shrouding the blades to increase strength.

Possible types of locks for fastening the working blades in the turbine disk.

The blade is attached to the disk with lock connection. Such a connection is one of the most loaded structural elements in a gas turbine engine. All loads perceived by the blade are transferred to the disk through the lock and reach very large values, especially since, due to the difference in materials, the disk and blades have different coefficients of linear expansion, and besides, due to the unevenness of the temperature field, they heat up differently.

In order to assess the possibility of reducing the load in the interlock and thereby increasing the reliability and service life of the turbine, research work is being carried out, among which experiments on bimetallic blades or application in blisk impeller turbines.

When using bimetallic blades, the loads in the locks of their fastening on the disk are reduced due to the manufacture of the locking part of the blade from a material similar to the material of the disk (or close in parameters). The blade feather is made of another metal, after which they are connected using special technologies (a bimetal is obtained).

Blisks, that is, impellers in which the blades are made in one piece with the disk, generally exclude the presence of a lock connection, and hence unnecessary stresses in the material of the impeller. Units of this type are already used in modern turbofan compressors. However, for them, the issue of repair is much more complicated and the possibilities of high-temperature use and cooling in aviation turbine.

An example of fastening the working blades in the disk using herringbone locks.

The most common way of fastening blades in heavily loaded turbine disks is the so-called herringbone. If the loads are moderate, then other types of locks that are structurally simpler, for example, cylindrical or T-shaped, can be used.

Control…

Since the working conditions aviation turbine extremely heavy, and the issue of reliability, as the most important unit of the aircraft, is of paramount priority, then the problem of monitoring the state of structural elements is in the first place in ground operation. In particular, this concerns the control of the internal cavities of the turbine, where the most loaded elements are located.

Inspection of these cavities is of course impossible without the use of modern equipment. remote visual control. For aircraft gas turbine engines, various types of endoscopes (borescopes) act in this capacity. Modern devices of this type are quite perfect and have great capabilities.

Inspection of the gas-air duct of the turbojet engine using the Vucam XO endoscope.

A vivid example is the portable measuring video endoscope Vucam XO of the German company ViZaar AG. Despite its small size and weight (less than 1.5 kg), this device is nevertheless very functional and has impressive capabilities for both inspection and processing of the information received.

Vucam XO is completely mobile. The whole set is housed in a small plastic case. The video probe with a large number of easily replaceable optical adapters has a full 360° articulation, 6.0 mm in diameter and can have different lengths (2.2m; 3.3m; 6.6m).

Borescopic inspection of a helicopter engine using a Vucam XO endoscope.

Borescopic checks using such endoscopes are provided for in the regulations for all modern aircraft engines. In turbines, the flow path is usually inspected. Endoscope probe penetrates into internal cavities aviation turbine through special control ports.

Borescopic control ports on the CFM56 turbojet turbine housing.

They are holes in the turbine housing, closed with sealed plugs (usually threaded, sometimes spring-loaded). Depending on the capabilities of the endoscope (probe length), it may be necessary to rotate the motor shaft. The blades (SA and RL) of the first stage of the turbine can be viewed through windows on the combustion chamber housing, and the blades of the last stage through the engine nozzle.

That will raise the temperature ...

One of the general directions for the development of gas turbine engines of all schemes is to increase the gas temperature in front of the turbine. This allows a significant increase in thrust without increasing air consumption, which can lead to a decrease in the frontal area of ​​the engine and an increase in the specific frontal thrust.

In modern engines, the gas temperature (after the torch) at the exit from the combustion chamber can reach 1650°C (with a tendency to increase), therefore, for normal operation of the turbine at such high thermal loads, it is necessary to take special, often protective measures.

First (and most simple of this situation)- usage heat-resistant and heat-resistant materials, both metal alloys and (in the future) special composite and ceramic materials, which are used to manufacture the most loaded turbine parts - nozzle and rotor blades, as well as disks. The most loaded of them are, perhaps, the working blades.

Metal alloys are mainly nickel-based alloys (melting point - 1455 ° C) with various alloying additives. Up to 16 types of various alloying elements are added to modern heat-resistant and heat-resistant alloys to obtain maximum high-temperature characteristics.

Chemical exotic...

Among them, for example, chromium, manganese, cobalt, tungsten, aluminum, titanium, tantalum, bismuth and even rhenium or instead of ruthenium and others. Particularly promising in this regard is rhenium (Re - rhenium, used in Russia), which is now used instead of carbides, but it is extremely expensive and its reserves are small. The use of niobium silicide is also considered promising.

In addition, the surface of the blade is often coated with a special coating applied using a special technology. heat-shielding layer(anti-thermal coating - thermal-barrier coating or TVS) , which significantly reduces the amount of heat flow into the body of the blade (thermal barrier functions) and protects it from gas corrosion (heat-resistant functions).

An example of a thermal protective coating. The nature of the temperature change over the blade cross section is shown.

The figure (microphoto) shows a heat-shielding layer on a high-pressure turbine blade of a modern turbofan engine. Here TGO (Thermally Grown Oxide) is a thermally growing oxide; Substrate - the main material of the blade; Bond coat - transition layer. The composition of fuel assemblies now includes nickel, chromium, aluminum, yttrium, etc. Experimental work is also being carried out on the use of ceramic coatings based on zirconium oxide stabilized by zirconium oxide (development by VIAM).

For example…

Quite widely known in engine building, starting from the post-war period and currently used are heat-resistant nickel alloys from Special Metals Corporation - USA, containing at least 50% nickel and 20% chromium, as well as titanium, aluminum and many other components added in small quantities. .

Depending on the profile purpose (RL, SA, turbine disks, elements of the flow path, nozzles, compressors, etc., as well as non-aeronautical applications), their composition and properties, they are combined into groups, each of which includes different types of alloys.

Rolls-Royce Nene turbine blades made from Nimonic 80A alloy.

Some of these groups are Nimonic, Inconel, Incoloy, Udimet/Udimar, Monel and others. For example, Nimonic 90 alloy, developed back in 1945 and used to make elements aircraft turbines(mainly blades), nozzles and parts of aircraft, has a composition: nickel - 54% minimum, chromium - 18-21%, cobalt - 15-21%, titanium - 2-3%, aluminum - 1-2%, manganese - 1%, zirconium -0.15% and other alloying elements (in small quantities). This alloy is produced to this day.

In Russia (USSR), VIAM (All-Russian Research Institute of Aviation Materials) has been and is successfully developing this type of alloys and other important materials for gas turbine engines. In the post-war period, the institute developed deformable alloys (EI437B type), since the beginning of the 60s it has created a whole series of high-quality cast alloys (more on this below).

However, almost all heat-resistant metallic materials can withstand temperatures up to about ≈ 1050°C without cooling.

That's why:

The second widely used measure this application various cooling systems blades and other structural elements aircraft turbines. It is still impossible to do without cooling in modern gas turbine engines, despite the use of new high-temperature heat-resistant alloys and special methods for manufacturing elements.

Among the cooling systems, there are two areas: systems open And closed. Closed systems can use forced circulation of the heat transfer fluid in the blade-radiator system, or use the "thermosiphon effect" principle.

In the latter method, the movement of the coolant occurs under the action of gravitational forces, when warmer layers displace colder ones. Here, for example, sodium or an alloy of sodium and potassium can be used as a heat carrier.

However, closed systems are not used in aviation practice due to the large number of problems that are difficult to solve and are at the stage of experimental research.

Approximate cooling scheme for a multistage turbojet turbine. The seals between the SA and the rotor are shown. A - a lattice of profiles for swirling air in order to pre-cool it.

But in wide practical application are open cooling systems. The refrigerant here is air, which is usually supplied at different pressures due to the different stages of the compressor inside the turbine blades. Depending on the maximum gas temperature at which it is advisable to use these systems, they can be divided into three types: convective, convective-film(or barrage) and porous.

With convective cooling, air is supplied inside the blade through special channels and, washing the most heated areas inside it, goes out into the stream in areas with lower pressure. In this case, various schemes for organizing the flow of air in the blades can be used, depending on the shape of the channels for it: longitudinal, transverse or loop-shaped (mixed or complicated).

Types of cooling: 1 - convective with a deflector, 2 - convective-film, 3 - porous. Blade 4 - heat-shielding coating.

The simplest scheme with longitudinal channels along the feather. Here, the air outlet is usually organized in the upper part of the blade through the shroud shelf. In such a scheme, there is a rather large temperature non-uniformity along the blade airfoil - up to 150-250˚, which adversely affects the strength properties of the blade. The scheme is used on engines with gas temperatures up to ≈ 1130ºС.

Another way convective cooling(1) implies the presence of a special deflector inside the feather (a thin-walled shell is inserted inside the feather), which contributes to the supply of cooling air first to the most heated areas. The deflector forms a kind of nozzle that blows air into the front of the blade. It turns out jet cooling of the most heated part. Further, the air, washing the rest of the surface, exits through the longitudinal narrow holes in the pen.

Turbine blade of the CFM56 engine.

In such a scheme, the temperature unevenness is much lower, in addition, the deflector itself, which is inserted into the blade under tension along several centering transverse belts, due to its elasticity, serves as a damper and dampens the vibrations of the blades. This scheme is used at a maximum gas temperature of ≈ 1230°C.

The so-called half-loop scheme makes it possible to achieve a relatively uniform temperature field in the blade. This is achieved by experimental selection of the location of various ribs and pins that direct air flows inside the body of the blade. This circuit allows a maximum gas temperature of up to 1330°C.

Nozzle blades are convectively cooled similarly to workers. They are usually made double-cavity with additional ribs and pins to intensify the cooling process. Higher pressure air is supplied to the front cavity at the leading edge than to the rear one (due to different compressor stages) and is released into different zones of the tract in order to maintain the minimum necessary pressure difference to ensure the required air velocity in the cooling channels.

Examples of possible methods for cooling rotor blades. 1 - convective, 2 - convective-film, 3 - convective-film with complicated loop channels in the blade.

Convective-film cooling (2) is used at an even higher gas temperature - up to 1380°C. With this method, part of the cooling air through special holes in the blade is released onto its outer surface, thereby creating a kind of barrier film, which protects the blade from contact with the hot gas stream. This method is used for both working and nozzle blades.

The third way is porous cooling (3). In this case, the power rod of the blade with longitudinal channels is covered with a special porous material, which makes it possible to carry out a uniform and dosed release of the coolant to the entire surface of the blade, washed by the gas flow.

This is still a promising method, which is not used in the mass practice of using gas turbine engines because of the difficulties with the selection of porous material and the high probability of fairly rapid clogging of pores. However, if these problems are solved, the supposedly possible gas temperature with this type of cooling can reach 1650°C.

Turbine disks and CA housings are also cooled by air due to different stages of the compressor as it passes through the internal cavities of the engine with washing of the cooled parts and subsequent release into the flow path.

Due to the rather high pressure ratio in the compressors of modern engines, the cooling air itself can have a rather high temperature. Therefore, to improve the cooling efficiency, measures are taken to reduce this temperature in advance.

To do this, the air, before being fed into the turbine on the blades and disks, can be passed through special profile gratings, similar to the SA turbine, where the air is twisted in the direction of rotation of the impeller, expanding and cooling at the same time. The amount of cooling can be 90-160°.

For the same cooling, air-to-air radiators cooled by secondary air can be used. On the AL-31F engine, such a radiator reduces the temperature to 220° in flight and 150° on the ground.

for cooling needs aviation turbine a sufficiently large amount of air is taken from the compressor. On various engines - up to 15-20%. This significantly increases the losses that are taken into account in the thermogasdynamic calculation of the engine. Some engines have systems that reduce the air supply for cooling (or close it altogether) at low engine operating conditions, which has a positive effect on efficiency.

Cooling scheme of the 1st stage of the turbofan engine NK-56. Also shown are honeycomb seals and a cooling cut-off tape at reduced engine operating modes.

When evaluating the efficiency of the cooling system, additional hydraulic losses on the blades due to a change in their shape during the release of cooling air are usually taken into account. The efficiency of a real cooled turbine is about 3-4% lower than that of an uncooled one.

Something about blade making...

On jet engines of the first generation, turbine blades were mainly manufactured stamping method followed by lengthy processing. However, in the 1950s, VIAM specialists convincingly proved that it was cast alloys and not wrought alloys that opened the prospect of increasing the level of heat resistance of blades. Gradually, a transition was made to this new direction (including in the West).

At present, the technology of precision waste-free casting is used in production, which makes it possible to produce blades with specially profiled internal cavities that are used for the operation of the cooling system (the so-called technology investment casting).

This is, in fact, the only way now to obtain cooled blades. It also improved over time. At the first stages, using injection molding technology, blades with different sizes were produced. grains of crystallization, which unreliably interlocked with each other, which significantly reduced the strength and service life of the product.

Later, with the use of special modifiers, they began to produce cast cooled blades with uniform, equiaxed, fine structural grains. To this end, in the 1960s, VIAM developed the first serial domestic heat-resistant alloys for casting ZhS6, ZhS6K, ZhS6U, VZhL12U.

Their operating temperature was 200° higher than that of the deformable (forging) alloy EI437A/B (KhN77TYu/YuR), which was then common. Blades made from these materials have worked for at least 500 hours without visually visible signs of failure. This type of manufacturing technology is still used today. Nevertheless, grain boundaries remain a weak point of the blade structure, and it is along them that its destruction begins.

Therefore, with the growth of the load characteristics of the work of modern aircraft turbines(pressure, temperature, centrifugal loads), it became necessary to develop new technologies for the manufacture of blades, because the multi-grain structure no longer satisfies the heavy operating conditions in many respects.

Examples of the structure of the heat-resistant material of rotor blades. 1 - equiaxed grain size, 2 - directional crystallization, 3 - single crystal.

Thus appeared " directional crystallization method". With this method, not individual equiaxed metal grains are formed in the hardening casting of the blade, but long columnar crystals, elongated strictly along the axis of the blade. This kind of structure significantly increases the fracture resistance of the blade. It is like a broom, which is very difficult to break, although each of its constituent twigs breaks without problems.

This technology was subsequently developed into an even more advanced " single crystal casting method”, when one blade is practically one whole crystal. This type of blade is now also installed in modern aviation turbines. For their manufacture, special alloys are used, including the so-called rhenium-containing alloys.

In the 70s and 80s, VIAM developed alloys for casting turbine blades with directional crystallization: ZhS26, ZhS30, ZhS32, ZhS36, ZhS40, VKLS-20, VKLS-20R; and in the 90s - corrosion-resistant alloys with a long service life: ZhSKS1 and ZhSKS2.

Further, working in this direction, VIAM from the beginning of 2000 to the present has created high-rhenium heat-resistant alloys of the third generation: VZhM1 (9.3% Re), VZhM2 (12% Re), ZhS55 (9% Re) and VZhM5 (4% ​​Re ). To further improve the characteristics over the past 10 years, experimental studies have been carried out, which resulted in rhenium-ruthenium-containing alloys of the fourth - VZhM4 and fifth generations VZhM6.

As assistants...

As mentioned earlier, only reactive (or active-reactive) turbines are used in gas turbine engines. However, in conclusion, it is worth remembering that among the used aircraft turbines there are also active ones. They mainly perform secondary tasks and do not take part in the operation of main engines.

And yet their role is often very important. In this case, it's about air starters used to run . There are various types of starter devices used to spin up the rotors of gas turbine engines. The air starter occupies perhaps the most prominent place among them.

Air starter turbofan.

This unit, in fact, despite the importance of functions, is fundamentally quite simple. The main unit here is a one- or two-stage active turbine, which rotates the engine rotor through a gearbox and a drive box (usually a low-pressure rotor in a turbofan engine).

The location of the air starter and its working line on the turbofan engine,

The turbine itself is spun by a stream of air coming from a ground source, or an onboard APU, or from another, already running aircraft engine. At a certain point in the start cycle, the starter will automatically disengage.

In such units, depending on the required output parameters, one can also use radial turbines. They can also be used in air conditioning systems in aircraft cabins as an element of a turbo-cooler, in which the effect of expansion and decrease in air temperature on the turbine is used to cool the air entering the cabins.

In addition, both active axial and radial turbines are used in turbocharging systems of reciprocating aircraft engines. This practice began even before the turbine became the most important GTE unit and continues to this day.

An example of the use of radial and axial turbines in auxiliary devices.

Similar systems using turbochargers are used in automobiles and in general in various compressed air supply systems.

Thus, the aviation turbine serves people well in an auxiliary sense.

———————————

Well, that's probably all for today. In fact, there is still a lot to be written about both in terms of additional information and in terms of a more complete description of what has already been said. The topic is very broad. However, it is impossible to grasp the immensity :-). For a general acquaintance, perhaps, it is enough. Thank you for reading to the end.

Until we meet again…

At the end of the picture, "out of place" in the text.

An example of a single-stage turbojet turbine.

Heron's aeolipil model in the Kaluga Museum of Cosmonautics.

Articulation of the Vucam XO endoscope video probe.

The screen of the Vucam XO multifunctional endoscope.

Endoscope Vucam XO.

An example of a thermal protective coating on the CA blades of a GP7200 engine.

Honeycomb plates used for seals.

Possible variants of labyrinth seal elements.

Labyrinth honeycomb seal.

Send your good work in the knowledge base is simple. Use the form below

Students, graduate students, young scientists who use the knowledge base in their studies and work will be very grateful to you.

Posted on http://www.allbest.ru/

Ministry of Education and Science of the Russian Federation

Federal Agency for Education

Samara State Aerospace University

named after Academician S.P. Queen

Department of the Theory of Aircraft Engines

Course work

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

Axial turbine designaviationengineJT9 D20

Samara 2008

Exercise

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

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

Table 1. Initial data for turbine design

high pressure turbine
Parameter	Numerical value	Dimension
T*TND = T*T
R*TND = R*T

Essay

Coursework on thermogasdynamic design of the JT9D20 axial turbine.

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

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

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

Conventions

D - diameter, m;

Relative bushing diameter;

h - blade height, m;

F - cross-sectional area, m 2;

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

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

i - specific enthalpy, kJ/kg;

k is the isentropic index;

l - length, m;

M - Mach number;

n - speed, 1/min;

Р - pressure, kPa;

Reduced speed;

s - flow velocity, m/s;

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

R - gas constant, kJ/kggrad;

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

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

S - axial width of the crown, m;

T - temperature, K;

Assigned resource, h;

V - flight speed, m/s;

z - number of steps;

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

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

Mass flow change factor;

U - circumferential speed, m/s;

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

Gap size, m;

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

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

Indices

a - axial component;

c - air section at the compressor inlet

vent - fan

vzl - takeoff;

w - bushing section;

d - gases section at the outlet of the turbine

k - compressor section at the outlet of the compressor

kr - critical

ks - combustion chamber

n - cross section of the undisturbed flow

on - guide apparatus;

cool - cooling;

n - flight parameter, peripheral diameter;

pr - given parameters;

ps - retaining stage

s - isentropic parameters;

c - second section at the exit of the nozzle

cp - average parameter;

st - step parameter;

t - turbine fuel section at the turbine inlet

h - hourly

* - braking parameters.

Abbreviations

HP - high pressure;

LP - low pressure;

VNA - input guide vane;

GDF - gas dynamic functions

GTE - gas turbine engine

Efficiency - efficiency factor;

ON - guide vane;

RK - impeller;

SA - turbine nozzle apparatus;

SAU - standard atmospheric conditions

Turbofan engine - turbojet bypass engine.

Introduction

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

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

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

2. Gas-dynamic calculation of the HP turbine

2.1 Distribution of heat drop by steps

2.2 Calculation of the step by average diameter

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

2.4 Calculation of flow parameters at different radii

Conclusion

List of sources used

Introduction

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

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

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

pressure

1.1 Calculation geometric and regime parameters HP turbines

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

Figure 1. - Geometric model of an axial turbine

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

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

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

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

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

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

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

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

5. Average diameter at the HP turbine outlet

6. Peripheral diameter at the outlet of the valve:

7. Sleeve diameter at the outlet of the valve:

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

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

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

10. Sleeve diameter at the HP turbine inlet:

11. HP turbine rotor speed:

1.2 Construction of the meridional section of the flowparts

HP turbines

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

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

accept kSA = 0.06

2. First stage impeller ring width:

accept kRK = 0.045

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

4. Second stage impeller ring width:

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

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

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

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

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

2 . G azodynamic calculation of the turbine VD

2.1 Distributionheat drop reduction by steps

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

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

.

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

Accept:

kJ/kg

Then: kJ/kg

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

m

; ; .

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

; ;

; ; .

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

Accept:

.

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

; .

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

.

7. Determine the total pressure at the stage inlet:

,

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

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

; .

10. Static pressure downstream

.

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

; .

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

.

2.2 Step calculation according to average at diameter at

Flow parameters behind the nozzle

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

.

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

;

3. Speed ​​coefficient CA is accepted:

.

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

; .

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

.

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

;

Where.

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

.

8. Effective angle at the outlet of the nozzle array

.

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

Accept: ;

;

.

10. Blade profile chord SA

.

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

12. Optimal SA lattice spacing in the first approximation

.

13. Optimum number of SA blades

.

We accept.

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

.

15. The size of the throat of the SA channel

.

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

; .

17. Static pressure in the gap between SA and RK

.

18. Actual gas velocity at the outlet of the SA

.

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

;

; .

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

.

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

;

.

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

.

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

.

24. Relative flow velocity at the inlet to the AC

.

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

;

; .

26. Reduced flow velocity in relative motion

.

27. Total pressure in relative air movement

.

Flow parameters at the outlet of the RC

28. Thermodynamic flow parameters

;

;.

29. Isentropic flow velocity in relative motion

.

30. Reduced isentropic flow velocity in relative motion:

.

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

31. Reduced flow velocity in relative motion

Let's accept:

,

Then:

; .

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

.

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

Let's calculate:

;

.

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

.

35. Effective angle at the outlet of the DC

.

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

Let's calculate:;

.

37. Blade profile chord RK

.

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

.

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

.

40. Optimum number of blades RK

.

We accept.

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

.

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

.

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

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

; .

45. Density of gas at the outlet of the RC

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

;

.

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

48. Absolute gas velocity behind the RK

.

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

50. Total enthalpy of gas behind the RC

.

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

disk and in the radial clearance

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

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

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

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

We accept the flow coefficient of the labyrinth seal:

The gap area is determined from the expression:

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

; .

Peripheral pressure

Seal pressure ratio

We accept the number of scallops:

Leakage loss

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

,

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

54. Total energy loss due to leakage and disk friction

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

;

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

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

58. Actual effective operation of a stage

59. Actual efficiency steps

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

which is 0.78%.

2.4 Calculation of parameters flow at different radii

turbine pressure blade wheel

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

Determination of parameters for the spigot section of the blade

1. Relative bushing diameter

2. Flow exit angle in absolute motion

3. Speed ​​ratio

4. Absolute flow rate at the outlet of the SA

5. Circumferential component of absolute speed

6. Axial component of absolute velocity

7. Isentropic velocity of gas outflow from SA

8. Thermodynamic parameters at the outlet of the SA

; ;

;

; .

9. Static pressure

.

10. Gas density

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

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

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

.

14. Relative speed at the hub

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

,

,

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

17. Reduced relative speed at the entrance to the RC

Parameters in peripheral section

18. Relates. peripheral section diameter

19. Angle of flow exit from SA in absolute motion

20. Speed ​​ratio

21. Absolute speed at the exit from the SA

22. Circumferential and axial components of absolute speed

23. Isentropic velocity of gas outflow from SA

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

;

, ; .

25. Static pressure

26. Gas Density

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

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

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

.

30. Relative flow velocity at the periphery

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

,

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

.

33. Reduced relative velocity at the entrance to the RC

Calculation of flow parameters at the outlet of the RC

34. Relative bushing diameter

35. Flow angle in absolute motion

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

37. Static pressure at the outlet of the valve

38. Thermodynamic parameters in RK

,

39. Isentropic flow velocity at the outlet of the RC

40. Reduced isentropic speed

41. Flow velocity behind the RK in relative motion.

, Where

speed factor.

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

;

43. Gas density behind the working crown

44. Flow exit angle in relative motion

45. Circumferential and axial components of the relative flow velocity

46. ​​Absolute speed at the output of the working crown

47. Circumferential component of absolute speed

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

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

;

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

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

51. Relative diameter of the peripheral section

52. Flow angle in absolute motion

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

54. Static pressure at the outlet of the valve

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

;

56. Isentropic flow velocity at the outlet of the RC

57. Reduced isentropic speed

58. Flow velocity behind RK in relative motion

Speed ​​ratio;

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

;

60. Gas density behind the working crown

61. Flow outlet angle in relative motion

62. Circumferential and axial components of the relative flow velocity

63. Absolute exit speed from RK

64. Circumferential component of absolute speed

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

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

;

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

3. Profiling of the impeller blade

Table 2. - Initial data for profiling of RV blades

Initial parameter and calculation formula	Dimension	Control sections
D (according to the drawing of the flow part of the stage)

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

Value			Average diameter	Periphery

Conclusion

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

List of sources used

1. Thermogasdynamic design of axial turbines for aircraft gas turbine engines using p-i-T functions: Proc. allowance / N.T. Tikhonov, N.F. Musatkin, V.N. Matveev, V.S. Kuzmichev; Samar. state aerospace un-t. - Samara, 2000. - 92. p.

2. Mamaev B.I., Musatkin N.F., Aronov B.M. Gas-dynamic design of axial turbines for aircraft gas turbine engines: Textbook. - Kuibyshev: KuAI, 1984 - 70 p.

3. Design calculation of the main parameters of aircraft GTE turbochargers: Proc. allowance / V.S. Kuzmichev, A.A. Trofimov; KuAI. - Kuibyshev, 1990. - 72 p.

4. Thermogasdynamic calculation of gas turbine power plants. / Dorofeev V.M., Maslov V.G., Pervyshin N.V., Svatenko S.A., Fishbein B.D. - M., "Engineering", 1973 - 144 p.

Hosted on Allbest.ru

Similar Documents

Calculation of flow parameters and construction of gratings of compressor and turbine stage profiles. Profiling of the combustion chamber, the jet nozzle of the designed engine and the profile gratings of the impeller of the high-pressure turbine. Construction of blade profiles.

term paper, added 02/27/2012


Profiling the blades of the first stage of a high-pressure turbine. Calculation and construction of lattice profiles of a subsonic axial compressor. Profiling of grids of impeller profiles along the radius. Calculation and construction of lattice profiles of the RK turbine on a PC.

term paper, added 02/04/2012


Determination of the main geometric dimensions of the meridional section of the turbine stage. Calculation of flow parameters in the stage nozzle at the average diameter. Establishment of flow parameters along the radius of the flow path when profiling blades.

term paper, added 11/14/2017


Designing a centrifugal compressor in a transport gas turbine engine: calculation of outlet flow parameters, geometrical parameters of the impeller outlet section, profiling of the meridional outlet, estimation of the maximum blade load.

term paper, added 04/05/2010


Thermogasdynamic calculation of the engine, selection and justification of parameters. Coordination of compressor and turbine parameters. Gas-dynamic calculation of the turbine and profiling of the turbine blades of the first stage of the turbine on a computer. Calculation of the turbine blade lock for strength.

thesis, added 03/12/2012


Calculation and profiling of engine design elements: working blades of the first stage of an axial compressor, turbine. Method for calculating velocity triangles. The procedure for determining the parameters of the combustion chamber, the geometric parameters of the flow path.

term paper, added 02/22/2012


Calculation and profiling of the working blade of the compressor stage, high-pressure gas turbine, annular combustion chamber and outlet device. Determination of the components of triangles of velocities and geometrical parameters of lattices of profiles on three radii.

term paper, added 02/17/2012


Thermogasdynamic calculation of the engine. Coordination of compressor and turbine operation. Gas-dynamic calculation of an axial turbine on a computer. Profiling of high-pressure turbine blades. Description of the engine design, calculation for the strength of the turbine disc.

thesis, added 01/22/2012


Choice and substantiation of the power and rotational speed of the gas turbine drive: thermogasdynamic calculation of the engine, pressure in the compressor, matching the parameters of the compressor and turbine. Calculation and profiling of lattice profiles of the turbine impeller.

term paper, added 12/26/2011


Profiling the blades of the first stage of a high-pressure compressor. Computer calculation of a turbine blade. Combustion chamber design. Gas-dynamic calculation of the nozzle. Formation of initial data. Computer profiling of the ejector nozzle.