Production of monowheels and impellers of gas turbine engines. Development and analysis of the technological process for processing nozzle blades TND

The production of gas turbine engine blades occupies a special place in aircraft engine manufacturing, which is determined by a number of factors, the main of which are:

complex geometric shape of the feather and shank of the blades;

high manufacturing precision;

the use of expensive and scarce materials for the manufacture of blades;

mass production of blades;

equipping the blade production process with expensive specialized equipment;

total manufacturing complexity.

Compressor and turbine blades are the most common parts of gas turbine engines. Their number in one engine kit reaches 3000, and the labor intensity of manufacturing is 25...35% of the total labor intensity of the engine.

The blade feather has an extended complex spatial shape

The length of the working part of the pen ranges from 30-500mm with a variable profile in cross sections along the axis. These sections are strictly oriented relative to the base design plane and the profile of the locking part. In the cross sections, the calculated values ​​of the points defining the profile of the back and trough of the blade in the coordinate system are specified. The values ​​of these coordinates are specified in a tabular manner. The cross sections are rotated relative to each other and create a twist of the blade feather.

The accuracy of the blade airfoil profile in the coordinate system is determined by the permissible deviation from the specified nominal values ​​of each point of the airfoil profile. In the example, this is 0.5 mm; the angular error in pen twisting should not exceed 20 '.

The thickness of the feather has small values; at the inlet and outlet of the air flow into the compressor, it varies from 1.45 mm to 2.5 mm for different sections. In this case, the thickness tolerance ranges from 0.2 to 0.1 mm. High demands are also placed on the formation of the transition radius at the inlet and outlet of the blade airfoil. The radius varies from 0.5mm to 0.8mm.

The roughness of the blade blade profile must be no lower than 0.32 µm.

In the middle part of the blade feather there are supporting bandage shelves of a complex profile design. These flanges play the role of auxiliary design surfaces of the blades, and carbide coatings of tungsten carbide and titanium carbide are applied to their supporting surfaces. The middle bandage shelves, connecting to each other, create a single support ring in the first wheel of the compressor rotor.

In the lower part of the blade there is a locking flange, which has a complex spatial shape with variable section parameters. The lower flanges of the blades create a closed loop in the compressor wheel and ensure smooth air supply to the compressor. The gap between these shelves is changed within 0.1...0.2 mm. The upper part of the blade feather has a shaped surface, the generatrix of which is precisely located relative to the profile of the lock and the inlet edge of the blade blade. The gap between the tops of the blades and the housing of the compressor stator wheel depends on the accuracy of this profile.

The working profile of the blade blade of the shroud flanges and the lock is subjected to hardening processing methods in order to create compressive stresses on the forming surfaces. High demands are also placed on the condition of the blade surfaces, on which cracks, burns and other manufacturing defects are not allowed.

The blade material belongs to the second control group, which provides for a thorough quality check of each blade. A special sample is also prepared for a batch of blades, which is subjected to laboratory analysis. The requirements for the quality of compressor blades are very high.

Methods for obtaining initial blanks for such parts and the use of traditional and special methods During further processing, the output quality and economic indicators of production are determined. Initial compressor blade blanks are obtained by stamping. In this case, workpieces of increased accuracy can be obtained, with small allowances for machining. Below we consider the technological process of manufacturing compressor blades, the initial blank of which is obtained by hot stamping of standard accuracy. When creating such a workpiece, ways were identified that reduce the labor intensity of manufacturing and achieving the listed quality indicators of compressor blades.

When developing the technological process, the following tasks were set:

Creation of the initial workpiece by hot stamping with a minimum allowance along the blade blade.

Creation of technological gains for orientation and reliable fastening of the workpiece in the technological system.

Development of technological equipment and application of the method of orienting the initial workpiece in the technological system relative to the profile of the blade airfoil in order to distribute (optimize) the allowance at various stages of machining.

Using a CNC machine to process complex contours in milling operations.

Using finishing methods such as grinding and polishing to guarantee the quality of surfaces.

Creation of a quality control system for the execution of operations at the main stages of production.

Route technology for manufacturing blades. Stamping and all related operations are performed using hot stamping technology of normal precision. Processing is carried out on crank presses in accordance with technical requirements. Stamping slopes are 7...10°. The transition radii of stamping surfaces are made within the limits of R=4mm. Tolerances for horizontal and vertical dimensions in accordance with IT-15. The permissible displacement along the parting line of the dies is no more than 2 mm. The feather of the original blank is subjected to profile rolling. Flash marks along the entire contour of the workpiece should not exceed 1 mm.

Compressor blades are one of the most critical and mass-produced engine products and, having a service life from several hours to several tens of thousands of hours, experience a wide range of impacts from dynamic and static stresses, high-temperature gas flow containing abrasive particles, as well as oxidative products of the environment and combustion fuel. It should be noted that, depending on the geographical location of operation and the operating mode of the engine, the temperature along its path ranges from -50...-40 C° to

700…800 C° in the compressor. Titanium alloys (VT22, VT3-1, VT6, VT8, VT33), heat-resistant steels (EN961 Ш, EP517Ш) are used as structural materials for compressor blades of modern gas turbine engines, and for turbine blades casting nickel-based alloys (ZhS6U, ZhS32) .

Experience in the operation and repair of engines for military aircraft shows that ensuring the assigned resource of 500-1500 hours largely depends on the level of damage to the compressor and turbine rotor blades. Moreover, in most cases it is associated with the appearance of nicks, fatigue and thermal fatigue cracks, pitting and gas corrosion, and erosive wear.

The drop in the endurance limit for stage 4 blades based on 20*10 6 cycles is 30% (from 480 MPa for blades without defects, to 340 MPa for repair blades), the maximum stress on repaired stage 4 blades, although decreasing, still significantly exceeds the stress on the edges of the blades without nicks. Nicks on compressor rotor blades lead to a significant loss of fatigue strength of new blades. A significant number of blades are rejected and irretrievably lost, since they have nicks that exceed the repair tolerance limit. Structures made of titanium with a relatively low weight have high resistance to corrosion, good mechanical properties and a beautiful appearance.

Relevance of the work

The service life and reliability of aircraft engines are mainly determined by the load-bearing capacity of compressor blades (Fig. 1), which are the most critical and highly loaded parts that experience significant alternating and cyclic loads during operation, which act on them at high frequencies. Compressor blades are the most massive, highly loaded and critical part of an aircraft engine.
A feature of compressor blades, which have thin inlet and outlet edges and are made of titanium alloys that are very sensitive to stress concentration, is that they are the first to encounter a foreign body (birds, hail, etc.) that gets into the engine tract.
Risks, nicks, erosion damage and other defects significantly increase the level of local vibration stresses, which sharply reduces the strength characteristics of the blades. Therefore, the creation of a favorable combination of properties of the surface layer during finishing finishing and strengthening operations has a great influence on increasing bearing capacity GTE blades. An urgent task is to assess the influence of surface strain hardening on the impact strength of blades in collisions with foreign objects.

Figure 1 - Model of a gas turbine engine compressor blade (10 frames, 20 cycles)

Currently, in the manufacture of compressor blades, methods of plastic deformation and mechanical processing, as well as complex technologies for finishing operations of the technological process, are widely used.
Vibroabrasive processing (VA) on special installations has found wide application in the production of compressor blades from titanium alloys. The use of chemically active liquids together with abrasives has a positive effect on the effectiveness of vibroabrasive processing.
Ultrasonic treatment of balls (UST) makes it possible to form a favorable combination of characteristics of the surface layer of compressor blades, which have low rigidity, high manufacturing precision, complex configuration and thin edges.
Pneumatic shot blasting (PDB) is characterized by a sliding collision of balls with the surface of the blade blade, preventing them from overworking. It has been established that PDO is accompanied by a decrease in structural heterogeneity and makes the structure, phase distribution and residual compressive stresses more uniform in the surface layer of the blade airfoil. The proposed pneumatic shot blasting method of finishing and hardening treatment effectively neutralizes technological microdefects of the surface layer formed at previous stages of the technological process, is accompanied by a significant increase in the endurance limit, a decrease in durability dissipation and does not require subsequent finishing of thin edges by hand polishing.
One of the promising methods of finishing and strengthening treatment is the method of magnetic abrasive polishing (MAP). A distinctive feature of MAP is the ability to process parts with different configurations and combine finishing and hardening operations in one process.
The problem of erosion of gas turbine engine blades is generally recognized. The intensity and type of erosion of compressor blades depend not only on the conditions of particle collision with the blade surface, but also on the combination of characteristics of the surface layer.
To increase the wear resistance of blades, various types of complex technologies have become increasingly used - the application of plasma coatings in combination with various finishing and strengthening methods.
The development and introduction into mass production of engines is currently accompanied by progressive design and technological solutions, expressed in the appearance of new parts, the use of fundamentally new structural materials, as well as improvements in production, assembly and testing technologies. Progressive technological processes of mechanical processing based on the concept of high-speed cutting are widely used, and methods of finishing, hardening and heat treatment are being improved.
The close relationship between the design and production technology of engines has predetermined a number of current issues related to increasing the load-bearing capacity of complex-profile parts using technological methods.

Purpose and objectives of the work

Goal of the work- increasing the durability and quality of gas turbine engine compressor blades by improving the structural and technological support for the manufacturing processes of gas turbine engine compressor blades.

Main tasks of the work:
1.) Conduct an analysis of the current state of structural and technological support for the manufacturing processes of gas turbine engine compressor blades;
2.) Explore the possibilities of increasing the durability of compressor blades by applying ion-plasma coatings;
3.) Perform experiments to study the properties of wear-resistant ion-plasma coating;
4.) Development of recommendations for improving the structural and technological support for the manufacturing processes of gas turbine engine compressor blades.

Scientific novelty of the work

The scientific novelty of the work lies in the development of recommendations for improving the structural and technological support for the manufacturing processes of gas turbine engine compressor blades and the creation of an optimal structure for the technological process of machining gas turbine engine compressor blades. This work also provides a solution to the problem of durability and wear resistance of gas turbine engine compressor blades.

Main part

Compressor blades of a gas turbine engine

GTE blades operate under high temperatures, reaching over 1200°C for a turbine and over 600°C for a compressor. Repeated changes in the thermal operating conditions of the engine - rapid heating at the time of start-up and rapid cooling when the engine is stopped - causes a cyclic change in thermal stress, characterized as thermal fatigue (Fig. 2). In addition, the profile part of the airfoil and the blade shank, in addition to stretching and bending from centrifugal forces, bending and torque from a high-speed gas flow, experience alternating stresses from vibration loads, the amplitude and frequency of which vary over a wide range.

Figure 2 - Scheme of the movement of gas flows in a gas turbine engine (3 frames)

The reliability of the working blades of a compressor and turbine depends not only on their structural strength, resistance to cyclic and long-term static loads, but also on their manufacturing technology, which directly affects the quality of the surface layer of the blade shank and blade. Structural and technological stress concentrators are formed in the surface layer; it is influenced by work hardening and internal residual stresses from mechanical processing. In addition, the surface layer is exposed to external loads under the main types of stress (bending, tension, torsion) external environment. These negative factors can lead to blade destruction and, consequently, to failure of the gas turbine engine.
The production of gas turbine engine blades occupies a special place in the aircraft engine industry, which is determined by a number of factors, the main of which are:
complex geometric shape feather and shank of blades;
high manufacturing precision;
the use of expensive materials such as alloy steels and titanium alloys;
mass production of blades;
equipment of the technological process is expensive specialized equipment;
high manufacturing complexity.
The production of gas turbine engine blades today is characterized by the following types of machining:
stretching;
milling;
rolling;
polishing;
vibration polishing or vibration grinding;
heat treatment

Formation of the surface layer during the finishing operations of blade manufacturing

During the manufacture of gas turbine engine blades, microroughnesses and scratches are formed on their surfaces, and structural and phase transformations occur in the surface layer. In addition, an increase in the hardness of the metal and the formation of residual stresses are observed in the surface layer.
Under operating conditions, the surface layer takes the greatest loads and is subjected to physical and chemical effects: mechanical, thermal, corrosive, etc.
In most cases, the surface properties of gas turbine engine blades begin to deteriorate due to wear, erosion, corrosion, and the initiation of fatigue cracks, which can lead to failure.
After finishing treatment, the following surface defects are distinguished: scratches, scratches, dents, pores, cracks, burrs, etc.
The physical and mechanical properties of the surface layer created during the manufacture of blades during operation change significantly under the influence of force, temperature and other factors.
The surface of the part has a number of features compared to the core. The atoms that are on the surface have one-way bonds with the metal, therefore they are in an unstable state and have excess energy compared to the atoms located inside.
As a result of diffusion, especially when exposed to elevated temperatures, chemical compounds of the base metal with substances penetrating from outside appear in the surface layer. At elevated temperatures, the diffusion mobility of atoms increases, leading to a redistribution of the concentration of alloying elements. Diffusion in the surface layer has a noticeable effect on the properties of metals. This is especially true for an operation such as grinding, where there is high temperature in the processing area.
The main reasons for the occurrence of macrostresses during machining are inhomogeneity of plastic deformation and local heating of the metal of the surface layer, as well as phase transformations.
The degree and depth of hardening of the surface layer of parts are determined by the mechanical processing modes and are directly related to the increase in the number of dislocations, vacancies and other defects in the metal crystal lattice.
The surface layer of gas turbine engine parts is formed as a result of interrelated phenomena occurring in the deformation zone and adjacent zones: multiple elastic-plastic deformations, changes in the plastic properties of the metal, friction, changes in micro and macrostructure, etc.
During hardening, as a result of deformation of the surface metal and friction work, heat is released, which heats the part. Under intensive processing modes, local areas of the surface layers are heated, when smoothing - up to 600-700 °C, with impact methods - up to 800-1000 °C.
Such heating leads to a decrease in the level of residual compressive stresses at the surface, which can lead to a decrease in the hardening effect. In some cases, compressive stresses transform into tensile stresses.
The main reason for hardening is an increase in the density of dislocations accumulating near shear lines and their subsequent stoppage in front of various kinds of obstacles formed during the deformation process or that existed before it. The crushing of metal volumes contained between the sliding planes into blocks, the rotation of these blocks, the curvature of the sliding planes and the accumulation of crystal lattice destruction products on them contribute to an increase in irregularities along the sliding planes, and, consequently, to hardening.
During mechanical processing of parts, the formation of residual stresses is associated with uneven plastic deformation of surface layers, which occurs during the interaction of force and thermal factors.
Deformation is accompanied by uneven in depth and interconnected processes of shift, reorientation, fragmentation, elongation or shortening of the components of the structure. Depending on the nature of the deformations, an increase in the density of the material of the part is observed.
Under severe hardening conditions, overworking can occur, as a result of which dangerous microcracks appear in the surface layer and the formation of particles of peeling metal is expected. Overhardening is an irreversible process in which heating does not restore the original structure of the metal and its mechanical properties.

Vibroabrasive treatment of blades

Blades are typical mass-produced parts of aircraft gas turbine engines; they operate under conditions of high static, dynamic and thermal loads and largely determine the service life and reliability of the engine as a whole.
High-strength titanium alloys are used for their manufacture, stainless steels, heat-resistant nickel-based alloys, as well as composite materials.
The labor intensity of manufacturing blades in most gas turbine engine designs is 30-40% of the total labor intensity of the engine. This feature, along with the operating conditions of the blade in the engine, requires the use in production of progressive methods for obtaining workpieces, modern processing technologies, especially in finishing operations, mechanization and automation of technological processes.
In the operation of aircraft gas turbine engines, of all failures due to strength failures of parts, blades account for about 60%. The vast majority of blade failures are of a fatigue nature. This is often facilitated by damage to the blades caused by solid particles entering the engine path (stones when taxiing on the ground, birds in flight, etc.). This creates a need to have a sufficiently high margin of cyclic strength of the blades, as well as to take special technological and design measures to increase their survivability in the event of damage (nicks).
Depending on the operating conditions in the engine, the level of alternating stresses in the blades is usually in the range of 40-160 MPa, and taking into account the required safety margin, their endurance limit is usually required in the range of 300-500 MPa. The fatigue resistance of a blade depends on the material, design of the blade, and its manufacturing technology, but in any case, the value of the endurance limit is greatly influenced by the condition of the surface layer. The main factors influencing the quality of the surface layer are:
- residual stresses - their sign, magnitude, depth, nature of distribution over the cross-section of the part, etc.;
- surface microrelief - the size and nature of micro-irregularities, the presence of scratches;
- structure of the surface layer.
The urgency of the tasks of increasing the fatigue resistance of blades has led to the development and implementation of special processing methods and the introduction in the industry of a number of special methods for treating their surface.
The place of vibroabrasive processing in the technological process of mechanical processing of blades is, as a rule, a finishing process performed at the final stage of processing. Depending on the material of the blade, the type of previous processing, and original value surface microirregularities and some other factors, processing modes are assigned - frequency and amplitude of vibrations, characteristics of working media (abrasive chips, molded vibrating bodies, ceramic, glass or metal balls, wooden cubes, etc.), mass ratios, etc. This allows you to achieve the desired result in a fairly wide range of initial surface states. Thus, for compressor blades of small and medium dimensions made of steel and titanium alloys, the final shaping operation is cold rolling followed by rounding of the edges with an abrasive wheel. In this case, the surface roughness is Ra = 1.6 and higher, so “soft” vibration treatment modes are used to level out micro-irregularities on the surface and create compressive stresses in the surface layer. In this case, “bulk” processing is used (without securing parts) in toroidal vibrating machines. In some cases, the processing technology involves abrasive grinding in the final operations followed by polishing the surface of the blade blade. Such blades are subjected to more intense vibration-abrasive treatment to remove micro-roughness and provide residual compressive stresses in the surface layer.
It is much more difficult to implement effective vibration treatment of large turbomachinery blades. The large mass of such parts, taking into account the weight of the container and the working environment, makes it problematic to create a vibrating machine with an acceptable frequency and amplitude of oscillations in two or three coordinates due to a sharp increase in the required drive power and dynamic overloads of machine elements. In addition, such parts have a worse initial surface quality, which reduces processing productivity.
The Motor Sich enterprise uses the method of longitudinal single-axis vibration processing in a closed container (POVO).
In traditional domestic and foreign vibro-abrasive machines, the bulk filler is driven by the oscillatory movements of the bottom of the container, which is always located below. In this case, the filler returns back free fall. The effectiveness of this method is not high enough.
The process of vibroabrasive processing of parts is significantly activated and intensified inside a closed container with two bottoms located opposite each other, if the bulk filler actively oscillates between them, receiving the kinetic energy of the push from each bottom. The intensity of collisions between the filler and the workpiece increases significantly. The side walls of the container are made inclined (conical), this creates additional compression of the filler during its movement, which increases the dynamic forces between the abrasive filler and the walls of the container, inside which the processed parts of the gas turbine engine are located in a fixed or free state.
When vibrating using this method with abrasive granules and hardened steel balls, more intense metal removal from the surface and surface microdeformation of parts occur than in traditional vibrating machines, which increases the magnitude and depth of surface compressive stresses and increases the fatigue resistance of parts.
Figure 3 shows the curves of changes in the surface roughness of blades made of steel 14Х17Н2Ш from the duration of processing on a vibration installation with a U-shaped container.

Figure 3 – Dependence of roughness on vibroabrasive treatment in a U-shaped container (1) and the POVO method (2)

Achieving a roughness of Ra=1.5 µm using the POVO method, as follows from Fig. 3, occurs in approximately 30 minutes, and by conventional vibroabrasive treatment - 1.5 hours.
A study of vibroabrasive machining of turbine and compressor blades shows the advantages of this process compared to manual polishing and glossing. The results of the study showed that the endurance limit of blades subjected to vibration grinding and vibration polishing is 410 MPa and meets the requirements of the technical specifications. The magnitude and nature of residual stresses on the blades under study are more favorable than on blades polished and polished by hand.

Conclusion

Great importance In solving the problem of ensuring the service life and reliability of aviation gas turbine engines, as well as the creation of new generations of engines, there is the development, improvement and creation of new technological processes, methods of processing parts and equipment that increase not only productivity, but also manufacturing quality.
The emergence of modern types and modifications of aircraft engines is continuously accompanied by new design solutions that entail technological difficulties. To overcome them in a timely manner and reduce the gap between the “ideal”, from the point of view of design, and the “real”, from the point of view of manufacturing technology, it is necessary to actively introduce into production progressive methods of mechanical and finishing-hardening processing.

Literature

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2. Driggs I. G., Pancaster O. E. Aviation gas turbines. Per. from English G.G. Mironov. - M., Oborongiz, 1957 - 265 p.
3. Zhiritsky G. S. Aviation gas turbines. -M., Oborongiz, 1950 - 511 p. 4. Doronin Yu.V., Makarov V.F. Reasons for the formation of defects on the profile of the feather of titanium blades during polishing. // Ibid. – 1991. - No. 12. – pp. 17-19
5. Koloshchuk E.M., Shabotenko A.G., Khazanovich S.V. Volumetric vibroabrasive processing of gas turbine engine parts. // Aviation it happened. – 1973. - No. 6. S7 13 -16
6. Boguslaev V.A., Yatsenko V.K., Zhemanyuk P.D., Pukhalskaya G.V., Pavlenko D.V., Ben V.P. Finishing and strengthening processing of gas turbine engine parts - Zaporozhye, ed. JSC MotorSich, 2005 – 559 p.
7. Demin F.I., Pronichev N.D., Shitarev I.L. Manufacturing technology of main parts of gas turbine engines: Textbook. allowance. - M.: Mechanical Engineering. 2002. - 328 pp.; ill.
8. Sulima A.M., Shulov V.A., Yagodkin Yu.D. Surface layer and operational properties machine parts. M.: Mechanical Engineering Yu, 1988.240p.
9. Skubachevsky G.S. Aviation gas turbine engines: A textbook for students of aviation universities. M.: Mechanical Engineering, 1969-544 p.
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11. http://www.nfmz.ru/lopatki.htm
JSC Naro-Fominsk Machine-Building Plant GTE compressor blades
12. http://www.nfmz.ru/lopatki.htm
Doctor of Technical Sciences Yuri Eliseev, General Director of Federal Scientific and Production Center MMPP "Salyut", Advanced technologies for the production of gas turbine engine blades

Important note!
When writing this abstract, the master's thesis has not yet been completed. Final completion: December 2009. The full text of the work and materials on the topic can be obtained from the author or his supervisor after the specified date.

The “turbine” topic is as complex as it is vast. Therefore, of course, there is no need to talk about its full disclosure. Let’s deal, as always, with “general acquaintance” and “individual interesting points”...

Moreover, the history of the aviation turbine is very short compared to the history of the turbine in general. This means that we cannot do without some kind of theoretical and historical excursion, the content of which for the most part does not relate to aviation, but is the basis for a story about the use of a gas turbine in aircraft engines.

About the hum and roar...

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

Quite 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 to do with turbines as such (except, of course, for the mutual influence during their joint operation in the general turbojet engine cycle).

Moreover, in a turbojet engine (these 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 part of it and is rather indirectly related to the “rumbling roar”.

And on those engines where it, as a unit, plays, in some way, a dominant role (these are indirect reaction engines, and it’s not for nothing that they are called gas turbine), the sound is no longer so impressive, or it is created by completely different parts of the aircraft’s power plant, for example, a propeller.

That is, neither hum nor rumble, as such, to aircraft 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. By definition, no gas turbine engine can do without a turbine.

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

History + a little theory...

And even for a very long time. Ever since the invention of mechanisms that convert the energy of natural forces 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 precisely determines the simplicity of the engine. Natural energy is directly, without any intermediate devices, converted into 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 to obtain rotation of the output shaft, you need to have an additional crank mechanism, which naturally complicates and makes the design heavier. The turbine is much more profitable in this regard.

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

Water mill turbine wheel

One of the best known and most ancient applications of turbines are large mechanical mills, used by man since time immemorial for various economic needs (not just for grinding grain). They are treated as water, so wind mechanisms.

For a long period of ancient history (the first mentions from about the 2nd century BC) and the history of the Middle Ages, these were virtually the only mechanisms used by man for practical purposes. The possibility of their use, despite all the primitiveness of the technical circumstances, lay in the simplicity of transformation of the energy of the working fluid used (water, air).

A windmill is an example of a turbine wheel.

In these essentially true rotary engines, the energy of 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 of them, in fact, are prototypes of modern blades blade machines, which are the turbines used today (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 1st century AD) in his treatise “Pneumatics”. The invention he described was called aeolipile , which translated from Greek means “ball of Aeolus” (god of the wind, Αἴολος – Aeolus (Greek), pila - ball (lat.)).

Aeolipile of Heron.

In it, the ball was equipped with two oppositely directed nozzle tubes. Steam came out of the nozzles, entering the ball through pipes from the boiler located below and thereby causing the ball to rotate. The action is clear from the figure below. It was a so-called reverse turbine, rotating to the side, reverse side steam release. Turbines This type has a special name - reactive (more details below).

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

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

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

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

Later, as technical knowledge developed, piston internal combustion engines “came onto the scene” various designs, more advanced and higher efficiency mechanisms. 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 heat engines, also followed a similar path in their development. And although there are separate mentions of some specimens in history, noteworthy and, moreover, documented units, including patented ones, 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 basic principles turbine design (hereinafter also gas turbine) as an important part of a heat engine.

Jet turbine patented by Laval.

The developments of a talented Swedish engineer and inventor are quite characteristic in this regard. Gustave de Laval(Karl Gustaf Patrick de Laval). His research at that time was related to the idea of ​​​​developing a new milk separator with increased drive speed, which could significantly increase productivity.

It was not possible to obtain a high rotational speed (rpm) by using the then traditional (albeit the only existing) piston steam engine due to the high inertia of the most important element - the piston. Realizing this, Laval decided to try to stop using the piston.

They say that the idea itself came to him while observing the work of sandblasting machines. In 1883 he received his first patent (English patent no. 1622) in this field. 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. Fundamentally, all this was no different from the aeolipile of Heron of Alexandria.

The manufactured device worked quite reliably with high speeds 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 level of technology led to nothing. Why did this happen?

——————-

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

The mentioned efficiency (for modern aircraft turbines this is the so-called power or effective efficiency) 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 rotating the shaft, and what part " went down the drain».

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

A turbine, as a dynamic expansion machine, unlike volumetric machines (piston 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 occurs due to an increase in the volume of the cylinder chamber.

As a result, the large potential energy of the working fluid, which was formed as a result of the supply of thermal energy of 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 the efficiency tells us how completely the kinetic energy transforms into mechanical energy in a given situation. The higher it is, the less kinetic energy the flow leaving the nozzle into the environment has. This remaining energy is called " losses with output speed", and it is directly proportional to the square of the speed of the outgoing flow (everyone probably remembers mС 2/2).

Operating principle of a jet turbine.

Here we are talking about the so-called absolute speed C. After all, the outgoing flow, or more precisely, each of its particles, participates in a complex movement: rectilinear plus rotational. Thus, the absolute speed C (relative to the 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 only possible if W and U are equal (as can be seen from the figure). The peripheral speed (U) in this case is called optimal.

Such equality would be easy to achieve on hydraulic turbines (such as Segner wheels), since the speed of liquid outflow from the nozzles for them (similar to the speed W) is relatively small.

But this same speed W for gas or steam 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 produce an exhaust velocity of only 31 m/s, and a steam turbine - 455 m/s. That is, it turns out that even at fairly low pressures (only 5 atm.), the Laval jet turbine should, for reasons of ensuring high efficiency, have a peripheral speed above 450 m/s.

For the then level of technological development, this was simply impossible. It couldn't be done reliable design with these parameters. It also made no sense to reduce the optimal peripheral speed by reducing the relative speed (W), since this can only be done by reducing the temperature and pressure, and therefore the overall efficiency.

Active Laval turbine...

The Laval jet turbine did not lend itself to further improvement. Despite the attempts made, things have reached a dead end. Then the engineer took a different path. In 1889, he patented a different type of turbine, which was later called active. Abroad (in English) it is now called impulse turbine, that is, pulsed.

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 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 working blades, on their concave part. As a result of the impact on the blades of the impeller, it begins to rotate. Or we can also say that rotation occurs due to the impulse action of the jet. Hence English name impulseturbine.

At the same time, in the interscapular channels, which have an almost 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, at the exit from the nozzle and at the entrance to the interblade channel: absolute speed C 1, relative W 1, peripheral speed U.

At the output, respectively, C 2, W 2, and the same U. In this case, W 1 = 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.

This process is shown in principle in a simplified figure. Also, to simplify the explanation of the process, it is assumed here that the vectors of absolute and peripheral velocities are almost parallel, the flow changes direction in the impeller by 180°.

Steam (gas) flow in the active turbine stage.

If we consider speeds in absolute values, we see that W 1 = C 1 – U, and C 2 = W 2 – U. Thus, based on the above, for the optimal mode, when the efficiency takes maximum values, and losses from the output speed tend to the minimum (that is, C 2 = 0), we have C 1 = 2U or U = C 1 /2.

We find that for an active turbine optimal peripheral speed half the speed of exhaust 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 easier.

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

Operating principle of an active turbine.

The consequence of this was 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 under the conditions of that time, which ultimately reduced the efficiency of this turbine.

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

This one has turbines there was no way to increase power by increasing temperature and steam pressure, since this would lead to an increase in peripheral speed, which was absolutely unacceptable due to the same design problems.

In addition, an increase in power (with an increase in peripheral speed) was also inappropriate for another reason. The consumers of turbine energy 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 its gearbox.

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

Eventually Laval active turbine was a relatively low-power unit (working units 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.

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

Among other mechanisms, it contained a description of the “steam wheel”, which contained 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 was that he made them all really work together and dealt with extremely complex issues of improving the mechanism as a whole.

Steam active turbine by Giovanni Branca.

Interestingly, one of the most famous features of his turbine was the design of the nozzle (it was separately mentioned in the same patent) supplying steam to the rotor blades. Here the nozzle, from an ordinary tapering one, as it was in a jet turbine, became contracting-expanding. Subsequently, this type of nozzle began to be called Laval nozzles. They allow the gas (steam) flow to be accelerated to supersonic speed with fairly low losses. About them .

Thus, the main problem that Laval struggled with when developing his turbines, and which he was never able to overcome, was the 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, the engineer developed an active turbine with two parallel rows of rotor blades mounted on one impeller (disk). It was the so-called two-stage turbine.

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

If we use the simplified principle proposed above for determining the peripheral speed for a single-stage jet turbine (Laval), it turns out that for a two-stage turbine the rotation speed is no longer two, but four times less than the exhaust speed from the nozzle.

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

This is the most effective solution to the problem of low optimal peripheral speed, which was proposed but not used by Laval and which is actively used in modern turbines, both steam and gas. Multi-stage…

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

That is, by changing the number of turbine stages, you can change the rotation speed of its shaft and, accordingly, change the load on it. In addition, multi-stage operation allows operation on the turbine big differences energy, that is, increase its power, and at the same time maintain high efficiency indicators.

Laval did not patent his two-stage turbine, although a prototype was made, so it bears the name of the American engineer Charles Curtis (Curtis wheel (or disk), 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 multi-stage steam turbine. There were many statements by various scientists and engineers about the usefulness of dividing available energy into stages before him, but he was the first to translate the idea into hardware.

Multistage active-reaction Parsons 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 profiled working blades.

This type of turbine is usually called a jet turbine, although the name is quite arbitrary. In fact, it occupies an intermediate position between the purely reactive Heron-Laval turbine and the purely active Laval-Branca turbine. Due to their design, working blades combine active and reactive principles in general 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 ones (the working fluid moves along the axis of rotation), but also radial ones (the steam moves in the radial direction). His three-stage purely active turbine “Heron” is quite well known, in which the so-called Heron wheels are used (the essence is the same as that of the aeolipile).

Jet turbine "Heron".

Subsequently, from the early 1900s, steam turbine construction rapidly gained momentum and Parsons was at its forefront. Its multi-stage turbines were equipped with naval vessels, first experimental ones (ship "Turbinia", 1896, displacement 44 tons, speed 60 km/h - unprecedented for that time), then military ones (example - battleship "Dreadnought", 18000 tons, speed 40 km/ h, turbine power 24,700 hp) and passenger (example - the same type "Mauritania" and "Lusitania", 40,000 tons, speed 48 km/h, turbo 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 clearly visible success in the operation of steam turbines could have only structural and fundamental significance for aviation, which was rapidly progressing in its development at exactly the same time.

The use of a steam turbine as a power plant in aircraft was, for obvious reasons, extremely questionable. Aviation turbine could only be a fundamentally similar, but much more profitable gas turbine. However, not everything was so simple...

According to Lev Gumilyovsky, author of the popular book “Engine Creators” in the 60s, one day, in 1902, during the period of the beginning of the rapid development of steam turbine construction, Charles Parsons, in fact one of the main ideologists of this business at that time, was asked, in general, , a humorous question: “ Is it possible to “parsonize” a gas engine?"(implying a turbine).

The answer was expressed in 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 undoubtedly had reasons to say so.

The use of a gas turbine, especially if we mean its use in aviation instead of a steam turbine, was of course tempting, because its positive aspects are obvious. With all its power capabilities, it does not require huge, bulky devices for generating steam - boilers, or equally large devices and systems for its cooling - condensers, cooling towers, cooling ponds, etc.

The heater for a gas turbine engine is a small, compact one, located inside the engine and burning fuel directly in the air flow. And he simply doesn’t have a refrigerator. Or rather, it exists, but it exists as if virtually, because the exhaust gas is discharged into the atmosphere, which is the refrigerator. That is, everything necessary for a heat engine is available, 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 out into the chimney.

But with such tangible advantages there are also significant shortcomings, which, in general, became the basis for Parsons’ categorical answer.

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

In the operating cycle of a steam turbine plant (Rankine cycle), the work of water compression is small and the requirements for the pump performing this function and its efficiency are therefore also small. In the gas turbine engine 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 proportion of useful work that the turbine can be designed for. Therefore, the requirements for an air compression unit in terms of its efficiency and economy are very high. Compressors in modern aviation 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 installation 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 several hundred degrees.

This means that it is released into the atmosphere (like into a refrigerator). a large number of energy, which, of course, negatively affects the efficiency of the entire operating cycle, which is characterized by thermal efficiency: η t = Q 1 – Q 2 / Q 1. Here Q 2 is the same energy released 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 not always possible to raise this temperature. Its maximum value is limited by the turbine itself and the main condition here is strength. 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 has often become the reason for the complexity and rise in cost of turbines. This situation has continued in our time.

And at the time of Parsons, neither the metallurgical industry nor aerodynamic science 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 were attempts...

Nevertheless, as usually happens, there were people who were not afraid (or perhaps did not understand :-)) 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 multi-stage turbine, noted the possibility of its operation, in addition to steam, also on fuel combustion products. A possible version of a gas turbine engine running on liquid fuel with a compressor, combustion chamber and turbine was also considered there.

Smoke spit.

Examples of the use of gas turbines without any theory behind it have been known for a long time. Apparently, even Heron used the principle of an air jet turbine in the “theater of automata”. The so-called “smoke skewers” ​​are quite 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, a turbine wheel rotates with combustion products from a fire (or hearth). It is interesting that Branca himself did not build most of his cars, but only expressed ideas for their creation.

"Wheel of Fire" 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 conversion of potential energy, that is, the supplied thermal energy of fuel combustion, into kinetic energy (acceleration) for rotation of the gas turbine occurred only due to the action of gravity when the warm masses rose upward. That is, the phenomenon of convection was used.

Of course, such “units” are for real machines, for example, for driving Vehicle could not be used. However, in 1791, the Englishman John Barber patented a “horseless transportation machine”, one of the most important components of which was a gas turbine. This was the first ever officially registered patent for a gas turbine.

John Barber engine with gas turbine.

The machine used gas obtained from wood, coal or oil, heated in special gas generators (retorts), which, after cooling, entered a piston compressor, where it was compressed along 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 technology at that time did not allow the idea to be brought to life. The current model of the Barber machine with a gas turbine was built only in 1972 by Kraftwerk-Union AG for the Hannover Industrial Fair.

Throughout the 19th century, development of the gas turbine concept progressed extremely slowly for the reasons described above. There were few examples worthy of attention. The compressor and high temperature 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.

There is a well-known 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. 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 entering the turbine, reducing their temperature. This is the 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 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 never able to actively compete with the steam turbine or become part of the gas turbine engine, which could adequately replace the piston internal combustion engine. Its use on engines was mainly auxiliary. For example, as charging units in piston engines, including aviation ones.

But from the beginning of the 40s the situation began to change rapidly. Finally, new heat-resistant alloys were created, which made it possible to radically increase 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, thanks to the combination of their power with relative lightness and compactness, to use them on aircraft. The era of jet aviation and aircraft gas turbine engines began.

Turbines in aviation gas turbine engines…

So... The main area of ​​application of turbines in aviation is gas turbine engines. The turbine here makes hard work— the compressor rotates. Moreover, in a gas turbine engine, as in any heat engine, the expansion work more work compression.

And the turbine is precisely an expansion machine, and it spends only part of the available energy of the gas flow on the compressor. The remaining part (sometimes called free energy) can be used for useful purposes depending on the type and design of the engine.

Scheme of TVAD Makila 1a1 with a free turbine.

Turboshaft engine AMAKILA 1A1.

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

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

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

Diagram of a single-shaft turboprop engine DART RDa6. Turbine on a common engine shaft.

Rolls-Royce DART RDa6 turboprop single-shaft engine.

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

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

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

Engine Pratt & Whitney Canada PT6A.

Scheme of a PW127 turboprop engine 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 hydraulic generators, electric generators, etc. All these devices are most often driven from the turbocharger shaft.

About 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 gas flow in them.

Radial.

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

The radial turbine stage consists of an impeller and fixed blades that form the flow at its inlet. The blades are profiled so that the inter-blade 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.

Diagram of a radial centripetal turbine (with explanations).

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

Impeller of a radial centripetal turbine.

Radial turbines They 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, so it can generate larger amounts of energy (heat drops).

However, these turbines have a small flow area and do not provide sufficient gas flow with the same dimensions 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 the engines of model aircraft and small unmanned aircraft.

The first Heinkel He 178 jet aircraft.

Heinkel HeS3 turbojet engine with radial turbine.

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

Axial aircraft turbine.

This is the only type of turbine currently used in mid-flight aircraft gas turbine engines. The main source of mechanical work on the shaft obtained from such a turbine in the engine is the impellers, or more precisely, the impeller blades (RL), installed on these impellers and interacting with the energetically charged gas flow (compressed and heated).

The crowns of the stationary 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 housing 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-mentioned interaction with the working blades, the kinetic energy of the flow is converted into mechanical energy, which 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 in parameters along the path is shown.

1. Without changing the pressure, and therefore the magnitude of the relative flow velocity (only its direction changes noticeably - the rotation 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 using the first method are called active. The gas flow actively (pulses) affects the blades due to a change in its direction as it flows around them. With the second method - jet turbines. Here, in addition to the impulse effect, the flow also affects the rotor blades indirectly (to put it simply), using reactive force, which increases the power of the turbine. Additional reactive action is achieved through 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 aviation gas turbine engines use only axial jet turbines.

Changing 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 aerodynamically advantageous.

The fixed 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 inter-blade channel decreases from inlet to outlet, that is, the cross-section f 1 is less than the cross-section f 0 . The result is a profile of a tapering jet nozzle.

The working blades that follow them also have greater curvature. In addition, with respect to the oncoming flow (vector W 1), they are located so as to avoid its breakdown and ensure correct flow around the blade. At certain radii, the RL also forms tapering interscapular channels.

Stage work aviation turbine.

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

Picture of the gas flow movement in the axial turbine stage.

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

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

The absolute flow velocity as it passes through the stage increases slightly 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 to ensure that the value of C 2 is minimal. This is especially important for the last stage of the turbine (at the first or middle stages, a deviation from a right angle of up to 25° is allowed). The reason for this is loss with output speed, which just depend on the magnitude of the speed 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 jet, then the remaining energy can be used in the nozzle. But, for example, for a helicopter engine that does not use jet thrust, it is important that the flow speed behind the last stage of the turbine is as low as possible.

Thus, in the stage of an active-reactive turbine, the expansion of gas (a decrease in pressure and temperature), the conversion and activation of energy (heat difference) occurs not only in the SA, but also in the impeller. The distribution of these functions between the RC and the SA is characterized by a parameter of the engine theory 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 profiling of the working blades varies along the radius, this parameter (as well as some others) is calculated according to the average radius (section B-B in the figure of changes in parameters in a stage).

Configuration of the working blade of an active-reaction turbine.

Change in pressure along the length of the radar blade of an active-reactive 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 actuated in the impeller. 60-70% is activated 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 efficiencies. Part of the available energy is necessarily wasted, that is, it turns into losses, which must be taken into account and measures taken to minimize them in order to increase the efficiency of the turbine, that is, increase its efficiency.

Losses consist 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 consist of “improving” the flow part by calculation and experiment, as well as correct calculation of velocity triangles for the flow in the turbine stage, or more precisely, choosing the most advantageous peripheral speed U at a given speed C 1 . These actions are usually characterized by the U/C 1 parameter. The peripheral speed at the middle radius in a turbojet engine is 270 – 370 m/s.

The hydraulic perfection of the flow path of a 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 a turbine, which characterizes it specifically as a unit for generating power, that is, the degree to which available energy is used to create work on the shaft.

This is the so-called power (or effective) efficiency. It is equal to the ratio of the work on the shaft to the available heat drop. This efficiency takes into account losses with output speed. They usually amount to 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 modern gas turbine engines, the adiabatic efficiency value is about 0.9 - 0.92 for uncooled turbines. If the turbine is cooled, then this efficiency may 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 “ flow losses" The flow part cannot be completely isolated from the rest of the engine due to the presence of rotating components in combination with stationary ones (cases + 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 housing.

Such a gas does not participate in the process of converting flow energy into mechanical energy, because it does not interact with the blades in this regard, that is, end losses occur (or radial clearance losses). They amount to about 2-3% and negatively affect both adiabatic and power efficiency, reduce the efficiency of gas turbine engines, 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 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 are trying to minimize it. It's quite difficult because aircraft turbine– the unit is heavily loaded. Accurate accounting of all factors influencing the size of the gap is quite difficult.

Engine operating modes often change, which means the amount of deformation of the working blades, the disks on which they are mounted, and the turbine housings changes as a result of changes in temperature, pressure and centrifugal forces.

Labyrinth seal.

Here it is necessary to take into account the amount 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.

Usually the gap is assessed after stopping the warm engine. In this case, the thin outer casing cools faster than the massive disks and shaft and, decreasing in diameter, touches the blades. Sometimes the radial clearance value is simply selected within 1.5-3% of the length of the blade blade.

The principle of honeycomb compaction.

In order to avoid damage to the blades if they touch the turbine body, special inserts made of a material softer than the material of the blades are often placed in it (for example, metal ceramics). In addition, non-contact seals are used. Usually these are labyrinthine or honeycomb labyrinth seals.

In this case, the working blades are banded at the ends of the feather and seals or wedges (for honeycombs) are already placed on the bandage 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 unit is assembled without a gap. After running-in, the gap is approximately 0.2 mm.

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

Similar methods of sealing gaps are used to reduce gas leakage from the flow part (for example, into the inter-disk space).

SOURZ…

These are the so-called passive methods radial clearance control. In addition, many gas turbine engines developed (and being developed) since the late 80s are equipped with so-called “ active radial clearance control systems» (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 “massiveness”. Therefore, during transient conditions they expand differently. For example, when an engine switches from a reduced operating mode to an increased one, a high-temperature, thin-walled casing warms up faster (than a massive rotor with disks) and expands, increasing the radial clearance between itself and the blades. Plus to this changes in pressure in the duct and the evolution of the aircraft.

To avoid this, an automatic system (usually a FADEC type main regulator) 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 magnitude of its linear expansion and, accordingly, the magnitude of the radial clearances change.

All this allows you to save 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 others types.

Much less frequently, but quite effectively, to synchronize the heating rates of the rotor and stator, forced airflow of the turbine disks (not the housing) is used. Such systems are used on CF6-80 and PW4000 engines.

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Axial clearances in the turbine are also regulated. For example, between the outlet edges of the SA and the inlet RLs, there is usually a gap within 0.1-0.4 from the RL chord at the average radius of the blades. The smaller this gap, the less energy loss of the flow behind the SA (for friction and equalization of the velocity field behind the SA). But at the same time, the vibration of the radar increases due to the alternating impact of the SA from the areas behind the bodies of the blades into the interscapular areas.

A little general about the design...

Axial aircraft turbines modern gas turbine engines may have different design shape of the flow part.

Dav = (Din + Dn) /2

1. Shape with constant body diameter (Dн). Here the internal and average diameters along the tract decrease.

Constant outer diameter.

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

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

Constant inner diameter.

When designing, try not to allow the socket angle to exceed 20°.

2. Mold with constant inner diameter (Dв).

The average diameter and body diameter increase along the tract. This scheme does not fit well with the dimensions of the engine. In a turbojet engine, due to the “divergence” of the flow from the internal casing, it is necessary to turn it further onto the SA, which entails hydraulic losses.

Constant average diameter.

The scheme is more suitable for use in turbofan engines.

3. Shape with constant average diameter (Davg). The diameter of the body increases, the internal 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 multi-stage. 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 peripheral 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 across stages.

An example of a three-stage turbojet turbine.

At the same time, however, she herself turbine structurally becomes more complicated and heavier. Due to the small temperature difference at each stage (it is distributed across all stages), a larger number of the first stages are exposed to high temperatures and often require additional cooling.

Four-stage axial turbine turbine.

Depending on the engine type, the number of stages may vary. For turbojet engines, usually up to three, for dual-circuit engines up to 5-8 stages. Typically, 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 bypass ratio).

Twin-shaft axial aircraft turbine.

For example, in the three-shaft Rolls-Royce Trent 900 engine, the turbine has three stages: a single stage to drive the high-pressure compressor, a single stage to drive the intermediate compressor and a five-stage to drive 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 aircraft turbine, simply put, is a structure consisting of a rotor, stator and various auxiliary structural elements. The stator consists of an outer casing, 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 methods of fastening.

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

On each disk, as the basis of the impeller, there are working blades. When designing blades, they try to make them with a smaller chord due to the smaller width of the rim of the disk on which they are installed, which reduces its mass. But at the same time, in order to maintain turbine parameters, it is necessary to increase the length of the airfoil, which may entail bandaging 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 using 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 linear expansion coefficients, 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 locking connection and thereby increasing the reliability and service life of the turbine, research work is being carried out, among which experiments on bimetallic blades or the use of blisk impellers in 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 similar in parameters). The blade blade is made from another metal, after which they are joined using special technologies (a bimetal is obtained).

Blisks, that is, impellers in which the blades are made integral with the disk, generally eliminate the presence of a locking connection, and therefore unnecessary stress in the material of the impeller. Components of this type are already used in compressors of modern turbofan engines. However, for them the issue of repair is significantly complicated and the possibilities for high-temperature use and cooling in aircraft turbine.

An example of fastening rotor blades to a disk using herringbone locks.

The most common method of attaching blades to heavily loaded turbine disks is the so-called herringbone. If the loads are moderate, then other types of locks that are simpler in design, 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 component of an aircraft, is of paramount priority, then the problem of monitoring the condition of structural elements comes first in ground operation. This is especially true for monitoring 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 inspection. For aircraft gas turbine engines, various types of endoscopes (borescopes) serve in this capacity. Modern devices of this type are quite advanced and have great capabilities.

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

A striking example is the portable measuring video endoscope Vucam XO from the German company ViZaar AG. Having a 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 received information.

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

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

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

Borescopic inspection 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 turn the motor shaft. The blades (SA and RL) of the first stage of the turbine can be inspected through the windows on the combustion chamber housing, and those of the last stage - through the engine nozzle.

Which 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 makes it possible to significantly increase thrust without increasing air consumption, which can lead to a decrease in the frontal area of ​​the engine and an increase in specific frontal thrust.

In modern engines, the gas temperature (after the flame) 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 safety measures.

First (and most downtime 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 for the manufacture of the most loaded parts of the turbine - nozzles and working 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 different alloying elements are added to modern heat-resistant and heat-resistant alloys to obtain maximum high-temperature characteristics.

Chemical exotic...

These include, for example, chromium, manganese, cobalt, tungsten, aluminum, titanium, tantalum, bismuth and even rhenium or, instead, 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-protective layer(anti-thermal coating - thermal-barrier coating or fuel assembly) , significantly reducing the amount of heat flow into the body of the blade (thermal barrier functions) and protecting it from gas corrosion (heat-resistant functions).

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

The figure (microphoto) shows the heat-protective layer on the 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 is a transition layer. The composition of fuel assemblies now includes nickel, chromium, aluminum, yttrium, etc. experimental work on the use of ceramic coatings based on zirconium oxide stabilized by zirconium oxide (developed by VIAM).

For example…

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, are quite widely known in the engine industry, starting from the post-war period and currently. .

Depending on their profile purpose (RL, SA, turbine disks, flow parts, nozzles, compressors, etc., as well as non-aviation applications), their composition and properties, they are combined into groups, each of which includes various options alloys

Rolls-Royce Nene engine 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 for the manufacture of elements aircraft turbines(mainly blades), nozzles and parts of aircraft, has the 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 still produced today.

In Russia (USSR), the development of this type of alloys and other important materials for gas turbine engines was and is being successfully carried out by VIAM (All-Russian Research Institute of Aviation Materials). In the post-war period, the institute developed deformable alloys (type EI437B), and from the early 60s created a whole series high-quality cast alloys (more on this below).

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

That's why:

The second, widely used measure, this is an application various cooling systems blades and others 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 cooling systems, there are two areas: systems open And closed. Closed systems can use forced circulation coolant liquid in the blade system - a radiator or use the principle of the “thermosyphon effect”.

In the latter method, the movement of the coolant occurs under the influence of gravitational forces, when warmer layers displace colder ones. The coolant here can be, for example, sodium or an alloy of sodium and potassium.

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

Approximate cooling diagram of a multi-stage turbojet turbine. The seals between the CA and the rotor are shown. A - a grid of profiles for swirling air for the purpose of pre-cooling it.

But they are in wide practical use open systems cooling. The refrigerant here is air, usually supplied at different pressures due to different compressor stages into 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 barrier) and porous.

During convective cooling, air is supplied inside the blade through special channels and, washing the most heated areas inside it, goes out into the flow in areas with lower pressure. In this case, various schemes for organizing air flow 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-protective coating.

The simplest scheme is with longitudinal channels along the feather. Here, the air outlet is usually organized in the upper part of the blade through the bandage shelf. In such a scheme, there is a rather large unevenness of temperature along the blade feather - up to 150-250˚, which adversely affects the strength properties of the blade. The circuit 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 facilitates 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. This results in jet cooling of the most heated part. Next, the air, washing the remaining surfaces, exits through narrow longitudinal holes in the feather.

Turbine blade of the CFM56 engine.

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

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

The nozzle blades are convectively cooled in the same way as the working blades. They are usually made double-cavity with additional ribs and pins to intensify the cooling process. Air at a higher pressure is supplied to the front cavity at the leading edge than to the rear (due to different stages of the compressor) and is released into various zones of the tract in order to maintain the minimum required pressure difference to ensure the required air speed 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 even higher gas temperatures – up to 1380°C. With this method, part of the cooling air is released through special holes in the blade onto its outer surface, thereby creating a kind of barrier film, which protects the blade from contact with the hot gas flow. This method is used for both working and nozzle blades.

The third method is porous cooling (3). In this case, the power rod of the blade with longitudinal channels is covered with a special porous material, which allows for a uniform and dosed release of coolant onto the entire surface of the blade washed by the gas flow.

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

The turbine disks and CA casings are also cooled by air due to the various stages of the compressor as it passes through the internal cavities of the engine, washing the cooled parts and then releasing them into the flow part.

Due to the fairly high degree of pressure increase in the compressors of modern engines, the cooling air itself can have a fairly high temperature. Therefore, to increase the cooling efficiency, measures are taken to first reduce this temperature.

To do this, before supplying the turbine to the blades and disks, air can be passed through special profile grids, 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 cooling amount can be 90-160°.

For the same cooling, air-to-air radiators cooled by secondary circuit 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 fairly large amount of air is taken from the compressor. On various engines - up to 15-20%. This significantly increases losses, which are taken into account in the thermogasdynamic calculation of the engine. Some engines are equipped with systems that reduce the supply of air for cooling (or even shut it off) at reduced engine operating conditions, which has a positive effect on efficiency.

Cooling diagram of the 1st stage of the NK-56 turbofan turbine. Honeycomb seals and a cooling shut-off tape at reduced engine operating conditions are also shown.

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

Something about making blades...

On jet engines first generation turbine blades were mainly manufactured stamping method followed by long-term processing. However, in the 50s, VIAM specialists convincingly proved that it was casting and not wrought alloys that offered the prospect of increasing the level of heat resistance of blades. Gradually, a transition to this new direction was made (including in the West).

Currently, the production uses precision non-waste casting technology, which makes it possible to produce blades with specially profiled internal cavities that are used to operate the cooling system (the so-called technology lost wax casting).

This is, in fact, the only way to obtain cooled blades now. It also improved over time. At the first stages of casting technology, blades with different sizes were produced grains of crystallization, which did not adhere securely to each other, which significantly reduced the strength and service life of the product.

Subsequently, using special modifiers, they began to produce cast cooled blades with homogeneous, equiaxed, fine structural grains. For this purpose, in the 60s, 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 then common deformable (stamping) alloy EI437A/B (KhN77TYu/YUR). Blades made from these materials worked for at least 500 hours without visually visible signs of destruction. This type of manufacturing technology is still used today. Nevertheless, grain boundaries remain weak point structures of the scapula, and it is along them that its destruction begins.

Therefore, with the increase in load characteristics of modern aircraft turbines(pressure, temperature, centrifugal loads) there was a need to develop new technologies for manufacturing blades, because the multi-grain structure in many respects no longer satisfied the severe operating conditions.

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

This is how “ directional crystallization method" With this method, in the solidifying casting of a blade, not individual equiaxed grains of metal are formed, but long columnar crystals elongated strictly along the axis of the blade. This kind of structure significantly increases the blade's fracture resistance. It is similar to a broom, which is very difficult to break, although each of the twigs that make it up breaks without problems.

This technology was subsequently refined to an even more advanced " monocrystal casting method", when one blade is practically one whole crystal. This type of blade is now also installed in modern aircraft turbines. For their manufacture, special alloys are used, including 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 - long-life corrosion-resistant alloys: ZhSKS1 and ZhSKS2.

Further, working in this direction, from the beginning of 2000 to the present, VIAM 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, experimental studies have been carried out over the past 10 years, 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 those used aircraft turbines There are also active ones. They mainly perform secondary tasks and do not take part in the operation of the propulsion engines.

And yet their role is often very important. In this case we are talking about air starters used to launch. 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 of a turbofan engine.

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

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

The turbine itself is spun by a flow of air coming from a ground source, either an on-board APU, or from another already running aircraft engine. At a certain point in the starting cycle, the starter is automatically turned off.

In units of this kind, depending on the required output parameters, they can also be used. radial turbines. They can also be used in air conditioning systems in aircraft cabins as an element of a turbo-refrigerator, in which the effect of expansion and reduction of 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 piston aircraft engines. This practice began even before the turbine became the most important component of a gas turbine engine 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 cars and in general various systems compressed air supply.

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

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Well, that's probably all for today. In fact, there is still a lot that can be written about here, 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, one cannot embrace the immensity :-). For general information, perhaps, it is enough. Thank you for reading to the end.

Until next time...

Finally, there are pictures that “do not fit” into the text.

An example of a single-stage turbojet turbine.

Model of Heron's aeolipile in the Kaluga Museum of Cosmonautics.

Articulation of the video probe of the Vucam XO endoscope.

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 options for labyrinth seal elements.

Labyrinth honeycomb seal.

Blades of gas turbine engines (GTE) are the most widely used parts in the production of these power plants.

The total number of blades in the rotor and stator of a gas turbine engine, depending on its design, can reach several thousand pieces with a range of two to three dozen items, while their standard sizes can range from several tens of millimeters to one to one and a half meters. Turbine blades are the most difficult to manufacture and the most critical to operate. The labor intensity of manufacturing these parts in the total labor costs for the production of gas turbine engines is at least 70 - 80%.

Improving technological processes for manufacturing blades of gas turbine engines (GTE) must solve, first of all, the problem of increasing the economic indicators of the process, namely: increasing the material utilization rate; reducing the labor intensity of manufacturing; reduction of the technological cycle for manufacturing parts and costs for technological preparation of production.

The basis for solving this problem is the development of group technologies for manufacturing the main parts of a gas turbine engine, which determine its cost. Such parts primarily include turbine and compressor blades, open and semi-closed impellers. The choice of one or another technology for their production depends on design features details. However, for the same blade design, various technological processes can be used, the choice of the most optimal of which is determined by the economic feasibility of its use within the framework of a particular production program, i.e. in the manufacture of the same part at different stages of production development - from single to serial - are used various technologies, while the transition from one technology to another can be significantly reduced if certain general principles are observed.

These principles must meet the conditions of automated production, where the achievement of the required geometric accuracy and quality of the surface layer is guaranteed by compliance with one or another group technology implemented on multi-purpose machines and the use of special processes.

One of the outstanding Soviet scientists and designers was Mikhail Mil. This unique person worked as the chief designer of helicopter manufacturing. Using his outstanding knowledge, the Mi-1, Mi-2, Mi-4, Mi-6, Mi-8, Mi-10, Mi-12, Mi-24 and others helicopters were created.

Group technology is based on standard parts designs. Classification of the latter into Various types carried out taking into account the similarity of their design features and functional purpose. This allows the use of similar technologies when processing parts of a particular group. The basis for the formation of groups of similar parts is the many different parts used in gas turbine engines (GTE).

Based on uniform features of similarities and differences between parts, the following groups with characteristic features can be formed: turbine blades; nozzle blades; compressor blades; rings; disks; shafts; deflectors; supports, etc. Thus, a group of parts is given - gas turbine engine compressor blades, which must be manufactured within the framework of one standard technology.

The use of group technology as one of the production stages requires its mandatory coding, based on a parts classification system. This system is built on the principle of distribution of parts into groups by the product designer. The geometric similarity of parts plays a decisive role in this case. This similarity determines another commonality - the similarity of processing methods, i.e. the same sequence of operations, cutting methods and, accordingly, the same technological equipment for their manufacture.

The next stage of classification is the use of codes (numbers) of group technology operations. The operation code must imply a specific technological operation that defines one or another stage of the group technology.

For example, operation 005 – production of technological bases for machining from foundry bases; operation 095 – processing of surfaces mating with another part from the technological base, etc. Thus, when compiling a new technology for the manufacture of a part included in a particular group, the operation number (code) is used to integrate this part into the technological capacities involved in this operation.

However, existing production already includes a large number of technologies created in the previous period, which must also be combined within the framework of a group technology, while maintaining their existing system of classification of parts, technological processes, equipment, etc.

In addition, within one group there may be parts with design differences that entail the introduction of additional operations into the technology. These operations do not fundamentally change the group technology; they are carried out within its framework. However, they significantly change the technology of a specific part included in this group. Due to these design differences to perform one or another stage of group technology for a specific part, a different number of technological operations and, accordingly, devices, cutting and measuring tools, etc. can be used.

Thus, the technological system of group technologies is intended, on the one hand, to generalize the experience of the previous stages of enterprise development, and on the other, to create an orderly system of technological preparation of production for the subsequent development of the enterprise.