Probably everyone knows that no matter how hard the Chinese try, they cannot copy modern jet engines. All. what they could - they copied and got their DRY, but the engine still has to be bought in the Russian Federation. I just read an article on WIM: http://www.warandpeace.ru/ru/news/view/74298/ "China is still unable to copy a modern jet engine." Moreover, I understand that there are ultra-modern technologies, developments, mathematics, and so on, so on, so on ... But in order to understand in more detail what the matter is, I recommend reading the following article.

ENGINES AND MATERIALS

The power of any heat engine determines the temperature of the working fluid - in the case of a jet engine, this is the temperature of the gas flowing from the combustion chambers. The higher the gas temperature, the more powerful the engine, the greater its thrust, the higher the efficiency and the better the weight characteristics. The gas turbine engine has an air compressor. It is driven by a gas turbine that sits on the same shaft with it. The compressor compresses atmospheric air up to 6-7 atmospheres and sends it to the combustion chambers, where fuel - kerosene - is injected. The flow of hot gas flowing out of the chambers - products of combustion of kerosene - rotates the turbine and, flying out through the nozzle, creates jet thrust, propels the aircraft. The high temperatures that occur in the combustion chambers required the creation of new technologies and the use of new materials for the design of one of the most critical elements of the engine - the stator and rotor blades of a gas turbine. They must, for many hours, without losing their mechanical strength, withstand the enormous temperature at which many steels and alloys already melt. First of all, this applies to turbine blades - they perceive the flow of hot gases heated to temperatures above 1600 K. Theoretically, the gas temperature in front of the turbine can reach 2200 K (1927 o C). At the time of the birth of jet aviation - immediately after the war - materials from which it was possible to make blades that could withstand high mechanical loads for a long time did not exist in our country.
Shortly after the end of the Great Patriotic War work on the creation of alloys for the manufacture of turbine blades was started by a special laboratory at VIAM. It was headed by Sergei Timofeevich Kishkin.

TO ENGLAND FOR METAL

Even before the war, the first domestic design of a turbojet engine was created in Leningrad by the designer of aircraft engines, Arkhip Mikhailovich Lyulka. In the late 1930s, he was repressed, but, probably anticipating his arrest, he managed to bury the drawings of the engine in the yard of the institute. During the war, the country's leadership learned that the Germans had already created jet aircraft (the first aircraft with a turbojet engine was the German "Heinkel" He-178, designed in 1939 as a flying laboratory; the twin-engine "Messerschmit" Me-262 became the first serial combat aircraft "Then Stalin called L.P. Beria, who was in charge of new military developments, and demanded to find those who are engaged in jet engines in our country. A.M. Lyulka was quickly released and given him a room in Moscow on Galushkin Street for the first design bureau jet engines. Arkhip Mikhailovich found his drawings and dug them out, but the engine according to his project did not work right away. Then they simply took the turbojet engine bought from the British and repeated it one to one. But the matter came up against materials that were not available in the Soviet Union, but were available in England, and their composition, of course, was classified.And yet it was possible to decipher it.
Arriving in England to get acquainted with the production of engines, S. T. Kishkin appeared everywhere in boots with thick microporous soles. And, having visited with a tour the plant where turbine blades were processed, near the machine, as if by chance, he stepped on the chips that had fallen from the part. A piece of metal crashed into soft rubber, got stuck in it, and then was taken out and already in Moscow subjected to a thorough analysis. The results of the analysis of the English metal and extensive own research carried out at VIAM made it possible to create the first heat-resistant nickel alloys for turbine blades and, most importantly, to develop the foundations of the theory of their structure and production.

It was found that the main carriers of the heat resistance of such alloys are submicroscopic particles of the intermetallic phase based on the Ni 3 Al compound. Blades made of the first heat-resistant nickel alloys could work for a long time if the gas temperature in front of the turbine did not exceed 900-1000 K.

CASTING INSTEAD OF STAMPING

The blades of the first engines were stamped from an alloy cast into a bar to a shape that vaguely resembles a finished product, and then long and carefully machined. But here an unexpected difficulty arose: in order to increase the working temperature of the material, alloying elements were added to it - tungsten, molybdenum, niobium. But they made the alloy so hard that it became impossible to stamp it - it could not be molded by hot deformation methods.
Then Kishkin suggested casting the shoulder blades. Engineers were indignant: firstly, after casting, the blade will still have to be machined, and most importantly, how can a cast blade be put into the engine? The metal of stamped blades is very dense, its strength is high, and cast metal remains looser and obviously less durable than stamped metal. But Kishkin managed to convince the skeptics, and VIAM created special casting heat-resistant alloys and blade casting technology. Tests were carried out, after which almost all aircraft turbojet engines began to be produced with cast turbine blades.
The first blades were solid and could not withstand high temperatures for a long time. It was necessary to create a system for their cooling. To do this, we decided to make longitudinal channels in the blades for supplying cooling air from the compressor. This idea was not so hot: the more air from the compressor goes to cooling, the less it goes into the combustion chambers. But there was nowhere to go - the resource of the turbine must be increased at all costs.

They began to design blades with several through cooling channels located along the axis of the blade. However, it soon became clear that such a design was inefficient: air flows through the channel too quickly, the area of ​​the cooled surface is small, and heat is not sufficiently removed. They tried to change the configuration of the internal cavity of the blade by inserting a deflector there, which deflects and delays the air flow, or to make channels of a more complex shape. At some point, aircraft engine experts came up with a tempting idea - to create an entirely ceramic blade: ceramics withstand very high temperatures, and it does not need to be cooled. Almost fifty years have passed since then, but so far no one in the world has made an engine with ceramic blades, although attempts continue.

HOW THE CAST SHOVEL IS MADE

The technology for manufacturing turbine blades is called investment casting. First, a wax model of the future blade is made, casting it in a mold, in which quartz cylinders are first placed in place of future cooling channels (later they began to use other materials). The model is covered with a liquid ceramic mass. After it dries, the wax is melted out with hot water, and the ceramic mass is fired. It turns out a form that can withstand the temperature of the molten metal from 1450 to 1500 ° C, depending on the grade of the alloy. Metal is poured into the mold, which solidifies in the form of a finished blade, but with quartz rods instead of channels inside. The rods are removed by dissolving in hydrofluoric acid. This operation is carried out in a hermetically sealed room by a worker in a space suit with a hose for air supply. Technology is inconvenient, dangerous and harmful.
To exclude this operation, VIAM began to make aluminum oxide rods with the addition of 10-15% silicon oxide, which dissolves in alkali. The material of the blades does not react with alkali, and the remains of aluminum oxide are removed with a strong jet of water.
AT Everyday life we are accustomed to consider cast products as very rough and rough. But we managed to choose such ceramic compositions that the shape of them is completely smooth and almost no machining is required. This greatly simplifies the work: the blades have a very complex shape, and it is not easy to process them.
New materials demanded new technologies. No matter how convenient the addition of silicon oxide to the material of the rods, it had to be abandoned. The melting point of aluminum oxide Al 2 O 3 is 2050 o C, and silicon oxide SiO 2 is only about 1700 o C, and new heat-resistant alloys destroyed the rods already in the pouring process.
In order for the aluminum oxide mold to retain its strength, it is fired at a temperature higher than the temperature of the liquid metal that is poured into it. In addition, the internal geometry of the mold during pouring should not change: the walls of the blades are very thin, and the dimensions must exactly match the calculated ones. That's why admissible value mold shrinkage should not exceed 1%.

WHY REJECTED STAMPED SHOVEL

As already mentioned, after stamping, the blade had to be machined. At the same time, 90% of the metal went into chips. The task was set: to create such a precision casting technology so that a given blade profile is immediately obtained, and the finished product would only have to be polished and applied with a heat-shielding coating. No less important is the design that is formed in the body of the blade and performs the task of cooling it.
Thus, it is very important to make a blade that is efficiently cooled without lowering the temperature of the working gas and has high long-term strength. This problem was solved by arranging the channels in the body of the blade and the outlets from it so that a thin air film appeared around the blade. At the same time, they kill two birds with one stone: hot gases do not come into contact with the blade material, and therefore do not heat it up and do not cool themselves.
Here there is some analogy with the thermal protection of a space rocket. When a rocket enters the dense layers of the atmosphere at high speed, the so-called sacrificial coating that covers the head begins to evaporate and burn. It takes on the main heat flow, and the products of its combustion form a kind of protective cushion. The design of the turbine blade is based on the same principle, only air is used instead of a sacrificial coating. True, the blades must also be protected from erosion and corrosion.

The procedure for making a blade is as follows. First, a nickel alloy is created with specified parameters for mechanical strength and heat resistance, for which alloying additives are introduced into nickel: 6% aluminum, 6-10% tungsten, tantalum, rhenium and a little ruthenium. They allow for maximum high temperature performance for cast nickel-based alloys (there is a temptation to increase them further by using more rhenium, but it is insanely expensive). A promising direction is the use of niobium silicide, but this is a matter of the distant future.
But here the alloy is poured into a mold at a temperature of 1450 ° C and cools along with it. The cooling metal crystallizes, forming separate equiaxed, that is, approximately the same size in all directions, grains. The grains themselves can be both large and small. They adhere unreliably, and the working blades collapsed along the grain boundaries and shattered to smithereens. Not a single blade could last longer than 50 hours. Then we proposed to introduce a modifier into the casting mold material - cobalt aluminate crystals. They serve as centers, nuclei of crystallization, accelerating the process of grain formation. The grains are uniform and fine. New blades began to work for 500 hours. This technology, which was developed by E. N. Kablov, is still working, and it works well. And we at VIAM produce tons of cobalt aluminate and supply it to factories.
The power of jet engines grew, the temperature and pressure of the gas jet increased. And it became clear that the multi-grain structure of the blade metal would not be able to work under the new conditions. Other ideas were needed. They were found, brought to the stage of technological development and became known as directed crystallization. This means that the metal, when solidified, does not form equiaxed grains, but long columnar crystals elongated strictly along the axis of the blade. A blade with such a structure will resist fracture very well. I immediately recall the old parable about a broom that cannot be broken, although all its twigs individually break without difficulty.

HOW DIRECTIONAL CRYSTALLIZATION IS PERFORMED

In order for the crystals forming the blade to grow properly, the molten metal mold is slowly removed from the heating zone. At the same time, the form with liquid metal stands on a massive copper disk cooled by water. The growth of crystals starts from below and proceeds upwards at a rate practically equal to the rate at which the mold exits the heater. When creating the directional crystallization technology, it was necessary to measure and calculate many parameters - the crystallization rate, the heater temperature, the temperature gradient between the heater and the cooler, etc. It was necessary to choose such a mold movement speed so that columnar crystals would grow over the entire length of the blade. Under all these conditions, 5-7 long columnar crystals grow for each square centimeter of the blade section. This technology has enabled the creation of a new generation of aircraft engines. But we went even further.
Having studied the grown columnar crystals by X-ray diffraction methods, we realized that the entire blade can be made entirely from one crystal, which will not have grain boundaries - the weakest structural elements along which destruction begins. To do this, they made a seed that allowed only one crystal to grow in a given direction (the crystallographic formula for such a seed is 0-0-1; this means that the crystal grows in the Z direction, but not in the X-Y direction). The seed was placed in the lower part of the mold and the metal was poured, intensively cooling it from below. The growing single crystal acquired the shape of a blade.
American engineers used a copper water-cooled crystallizer for cooling. And after several experiments, we replaced it with a bath with molten tin at a temperature of 600-700 K. This made it possible to more accurately select the required temperature gradient and obtain high-quality products. At VIAM, installations with baths for growing single-crystal blades were built - very advanced machines with computer control.
In the 1990s, when the USSR collapsed, Soviet aircraft remained in East Germany, mainly MiG fighters. They had blades of our production in their engines. The metal of the blades was examined by the Americans, after which, quite soon, their specialists arrived at VIAM and asked to show who created it and how. It turned out that they were given the task of making single-crystal blades of a meter length, which they could not solve. We designed an installation for high-gradient casting of large blades for power turbines and tried to offer our technology to Gazprom and RAO "UES of Russia", but they showed no interest. Nevertheless, we have almost ready an industrial installation for casting meter-long blades, and we will try to convince the management of these companies of the need to implement it.

By the way, turbines for the power industry is another interesting task that VIAM solved. Aircraft engines that had run out of service life began to be used at compressor stations of gas pipelines and in power plants that feed oil pipeline pumps. Now it has become an urgent task to create special engines for these needs that would operate at much lower temperatures and pressure of the working gas, but much longer. If the resource of an aircraft engine is about 500 hours, then the turbines on the oil and gas pipeline should work 20-50 thousand hours. One of the first to deal with them was the Samara design bureau under the leadership of Nikolai Dmitrievich Kuznetsov.

HEAT RESISTANT ALLOYS

A single-crystal blade does not grow solid - inside it has a cavity of complex shape for cooling. Together with CIAM, we have developed a cavity configuration that provides a cooling efficiency coefficient (ratio of temperatures of the blade metal and working gas) equal to 0.8, almost one and a half times higher than that of serial products.

These are the blades we offer for new generation engines. Now the gas temperature in front of the turbine barely reaches 1950 K, and in new engines it will reach 2000-2200 K. For them, we have already developed high-temperature alloys containing up to fifteen elements of the periodic table, including rhenium and ruthenium, and heat-shielding coatings, in which include nickel, chromium, aluminum and yttrium, and in the future - ceramic from zirconium oxide stabilized with yttrium oxide.

In the first generation alloys, a small amount of carbon was present in the form of titanium or tantalum carbides. Carbides are located along the boundaries of the crystals and reduce the strength of the alloy. We got rid of carbide and replaced it with rhenium, increasing its concentration from 3% in the first samples to 12% in the last ones. There are few reserves of rhenium in our country; there are deposits in Kazakhstan, but after the collapse of the Soviet Union, it was completely bought up by the Americans; remains the island of Iturup, which is claimed by the Japanese. But we have a lot of ruthenium, and in new alloys we have successfully replaced rhenium with it.
The uniqueness of VIAM lies in the fact that we are able to develop both alloys and the technology for their production, as well as the method of casting the finished product. Huge work and knowledge of all employees of VIAM has been invested in all the blades.

Candidate of Technical Sciences I. DEMONIS, Deputy CEO VIAM

The production of GTE blades occupies a special place in the aircraft engine industry, which is due to a number of factors, the main of which are:

complex geometric shape of the airfoil and blade shank;

high manufacturing precision;

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

mass production of blades;

equipment technological process production of blades with expensive specialized equipment;

overall manufacturing complexity.

Compressor and turbine blades are the most massive 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 feather of the scapula has an extended complex spatial shape

The length of the working part of the pen is from 30-500 mm 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 interlock. In the cross sections, the calculated values ​​of the points that determine the profile of the back and trough of the blade in the coordinate system are given. The values ​​of these coordinates are given in a tabular way. 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 allowable deviation from the given nominal values ​​of each airfoil profile point. In the example, this is 0.5 mm, while the angular error in the twist of the pen should not exceed 20 ’.

The thickness of the pen has small values; at the inlet and outlet of the air flow to the compressor, it varies from 1.45 mm to 2.5 mm for various 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 in this case changes from 0.5 mm to 0.8 mm.

The roughness of the blade airfoil profile must be at least 0.32 µm.

In the middle part of the blade airfoil there are supporting shroud shelves of a complex profile design. These shelves play the role of auxiliary design surfaces of the blades, and hard-alloy coatings of tungsten carbide and titanium carbide are applied to their bearing surfaces. The middle shroud shelves, connecting with 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 lock shelf, which has a complex spatial shape with variable cross-sectional parameters. The lower shelves of the blades create a closed circuit in the compressor wheel and provide smooth air supply to the compressor. Changing the gap between these shelves is carried out within 0.1 ... 0.2 mm. The upper part of the blade airfoil has a shaped surface, the generatrix of which is exactly located relative to the profile of the lock and the leading edge of the airfoil. The clearance 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 shroud blade feather and the lock is subjected to hardening processing methods in order to create compressive stresses on the generatrix surfaces. High requirements are also imposed 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. For a batch of blades, a special sample is also prepared, 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 for further processing determine the output quality and economic indicators of production. The initial blanks of compressor blades 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 original workpiece, which was obtained by hot stamping of ordinary accuracy. When creating such a workpiece, ways have been identified that reduce the complexity of manufacturing and the implementation of the listed indicators, the quality of the compressor blades.

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

Creation of the initial blank by hot stamping with a minimum allowance for the blade feather.

Creation of technological profits 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 blade airfoil profile in order to distribute (optimize) the allowance at various stages of machining.

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

The use of finishing methods of processing by grinding and polishing with the guarantee of quality indicators of surfaces.

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

Route technology for the manufacture of blades. Stamping and all related operations are carried out using conventional precision hot stamping technology. Processing is carried out on crank presses in accordance with technical requirements. Stamping slopes are 7…10°. The transition radii of the stamping surfaces are performed within R=4mm. Tolerances for horizontal and vertical dimensions in accordance with IT-15. Permissible displacement along the parting line of stamps is not more than 2 mm. Feather of the original workpiece is subjected to profiled running. Flash traces 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 effects from dynamic and static stresses, high-temperature gas flow containing abrasive particles, as well as oxidative products of the environment and combustion. fuel. At the same time, it should be noted that, depending on the geographical location of operation and the mode of operation of the engine, the temperature along its path ranges from -50 ... -40 ° C to

700…800 С° in the compressor. Titanium alloys (VT22, VT3-1, VT6, VT8, VT33), heat-resistant steels (EN961 Sh, EP517Sh) are used as structural materials for compressor blades of modern gas turbine engines, and nickel-based cast alloys (ZhS6U, ZhS32) are used for turbine blades. .

The experience of operating and repairing engines for military aircraft shows that the provision of the assigned resource of 500-1500 hours largely depends on the level of damage to the compressor and turbine blades. At the same time, 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 blades of the 4th stage on the basis of 20 * 10 6 cycles is 30% (from 480 MPa for blades without defects, to 340 MPa for repair blades), although the maximum stresses on the repaired blades of the 4th stage, although they decrease, still significantly exceed the stress on blade edges without nicks. The nicks on the compressor rotor blades lead to a significant loss of fatigue strength of the new blades. A significant number of blades are rejected and irretrievably lost, as they have nicks that go beyond the repair tolerance limit. Structures made of titanium with a relatively low weight have high corrosion resistance, good mechanical properties and a beautiful appearance.

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

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

About the hum and rumble...

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

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

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

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

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

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

History + some theory ...

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

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

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

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

Turbine wheel of a water mill

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

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

A windmill is an example of a turbine wheel.

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

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

Aeolipil Heron.

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

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

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

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

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

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

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

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

It all started with a couple...

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

Jet turbine patented by Laval.

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

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

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

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

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

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

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

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

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

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

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

The principle of operation of a jet turbine.

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

Segner wheel.

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

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

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

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

Laval active turbine...

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

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

Active single-stage steam turbine patented by Laval.

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

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

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

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

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

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

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

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

The principle of operation of an active turbine.

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

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

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

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

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

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

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

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

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

Steam active turbine Giovanni Branca.

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

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

Multi-stage….

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

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

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

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

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

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

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

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

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

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

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

It is customary to call this type of turbine a reactive one, although the name is rather arbitrary. In fact, it occupies an intermediate position between the purely reactive Heron-Laval turbine and the purely active Laval-Branca. The rotor blades, due to their design, combine active and reactive principles in common 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 multistage turbines. Among his designs were not only the above-described axial (the working fluid moves along the axis of rotation), but also radial (steam moves in the radial direction). Quite well known is his three-stage purely active turbine "Heron", in which the so-called Heron's wheels are used (the essence is the same as that of the aeolipil).

Jet turbine "Heron".

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

About gas turbines...

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

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

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

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

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

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

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

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

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

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

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

Temperature…

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

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

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

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

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

And yet there have been attempts...

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

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

Smoke spit.

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

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

The Fire Wheel by Giovanni Branca.

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

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

John Barber gas turbine engine.

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

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

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

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

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

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

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

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

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

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

Turbines in aircraft gas turbine engines ...

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

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

Scheme TVAD Makila 1a1 with a free turbine.

Turboshaft engine AMAKILA 1A1.

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

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

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

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

Turboprop single-shaft engine Rolls-Royce DART RDa6.

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

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

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

Pratt & Whitney Canada PT6A engine.

Scheme of a PW127 TVD with a free turbine.

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

On the types of turbines.

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

Radial.

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

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

Scheme of a radial centripetal turbine (with explanations).

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

The impeller of a radial centripetal turbine.

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

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

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

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

The first Heinkel He 178 jet aircraft.

TRD Heinkel HeS3 with a radial turbine.

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

Axial aviation turbine.

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

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

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

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

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

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

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

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

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

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

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

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

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

Step work aviation turbine.

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

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

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

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

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

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

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

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

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

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

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

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

Something about losses.

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

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

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

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

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

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

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

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

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

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

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

labyrinth seal.

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

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

The principle of honeycomb sealing.

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

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

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

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

SAURZ…

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

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

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

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

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

———————-

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

A little about the design...

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

Dav = (Din+Dn) /2

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

Constant outside diameter.

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

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

Constant inside diameter.

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

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

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

Constant average diameter.

The scheme is more appropriate for use in turbofan engines.

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

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

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

An example of a three-stage turbojet turbine.

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

Four-stage axial turbine TVD.

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

Twin-shaft axial aircraft turbine.

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

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

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

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

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

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

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

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

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

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

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

Control…

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

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

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

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

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

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

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

Borescopic control ports on the CFM56 turbojet turbine housing.

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

That will raise the temperature ...

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

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

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

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

Chemical exotic...

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

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

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

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

For example…

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

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

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

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

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

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

That's why:

The second widely used measure this application various cooling systems blades and 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 the cooling systems, there are two areas: systems open and closed. Closed systems can use forced circulation of the heat transfer fluid in the blade-radiator system, or use the "thermosiphon effect" principle.

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

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

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

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

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

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

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

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

Turbine blade of the CFM56 engine.

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

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

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

Examples possible ways blade cooling. 1 - convective, 2 - convective-film, 3 - convective-film with complicated loop channels in the blade.

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

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

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

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

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

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

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

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

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

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

Something about blade making...

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

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

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

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

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

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

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

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

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

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

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

As assistants...

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

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

Air starter turbofan.

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

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

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

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

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

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

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

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

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

Until we meet again…

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

An example of a single-stage turbojet turbine.

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

Articulation of the Vucam XO endoscope video probe.

The screen of the Vucam XO multifunctional endoscope.

Endoscope Vucam XO.

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

Honeycomb plates used for seals.

Possible variants of labyrinth seal elements.

Labyrinth honeycomb seal.

Introduction

The production of gas turbine engine blades occupies a special place in modern mechanical engineering. This is due to the following features of the manufacture of blades.
1. Responsible purpose of the blades in the engine. Blades to a decisive extent determine the reliability and non-failure operation of gas turbine engines. The service life of the engine is determined, as a rule, by the performance of the blades. In this regard, the manufacturing and control technology of the blades should ensure the stability of the quality of their manufacture and exclude the possibility of installing blades with deviations in geometric dimensions, surface quality, metallurgical and other defects in the engine.
2. The complexity of geometric shapes and the requirements for high precision in the manufacture of blades. The feather of the blade is a blade of variable section, limited by surfaces of a complex shape and precisely oriented in space with respect to the lock. The manufacturing accuracy of the pen is within 0.05x0.15 mm. The lock part, with which the blades are attached to the disks, is made with an accuracy of 0.01-0.02 mm.
3. Mass production of blades. A modern engine with an axial compressor has up to 2000 blades. In this regard, even with the production of prototype engines, the manufacture of blades is of a serial nature.
4. The use of expensive and scarce materials for the manufacture of blades. In this regard, the technological process for the production of blades must guarantee a minimum percentage of rejects.
5. Poor machinability of materials used for the manufacture of blades. Turbine blades are made from nickel-based alloys, which have relatively high hardness and high toughness.
The combination of these factors determined the specificity of blade production.
The production of blades is being improved at the present time, mainly in the direction of mechanization and automation. The elimination of manual labor allows not only to reduce labor intensity, but also to improve the quality of blade manufacturing.
Significant progress has been made lately in the field of intensification of the processing modes of heat-resistant and titanium steels and alloys, as well as in the field of manufacturing ceramic blades.

1. Purpose and design of nozzle blades

Guiding and working blades, according to their official purpose, are the main parts of steam and blade engines. Together, they form the flow part of the turbine, in which the thermal energy of the working medium (steam, gas) is converted into the mechanical work of a rotating rotor. The combination of guide and working blades is called the turbine blade apparatus.
The blade apparatus is the most expensive and most critical part of the turbine. The efficiency of a turbine—its efficiency—depends primarily on the quality of the blade apparatus. The complexity of manufacturing the blades of a modern powerful steam turbine reaches 42-45% of the total labor intensity of manufacturing all its parts.
Turbine blades operate under very difficult conditions. They are subjected to strong centrifugal force, bending and pulsating action of the working medium, causing vibrations of the blades, in which resonant vibrations can easily be excited. All this happens in the first stages of the turbine at high temperatures of the working medium, which acts on the blades both chemically and mechanically; in the last stages, the leading edges of the blades are corroded (eroded) by particles of water contained in the wet steam.
These conditions require a particularly careful approach to the design of blades, the choice of materials for them and the organization of their production. Particular care should be taken to fulfill all the dimensions forming the shape of the blades and to comply with the technical requirements established for their manufacture. Deviations from the drawings can cause additional stresses in the blades that are not provided for by the calculations, which, in turn, can lead to a serious failure of the turbines.
The nozzle apparatus of the first stage is washed by gas, the temperature of which, taking into account the unevenness after the combustion chamber, can be 100-120 ° C higher than the average mass in front of the turbine. Therefore, in high-temperature gas turbines, it is cooled very intensively. Under the mass average temperature in front of the turbine, the weighted average stagnation temperature directly in front of the rotor blades should be considered. This allows air to be more freely used to cool the nozzle vanes of the first stage, however, at the same time, small aerodynamic losses in the nozzle apparatus itself and the most uniform flow in temperature, pressure and direction directly in front of the rotor blades of this stage should be ensured.
The nozzle vanes are usually slightly twisted along the radius, and therefore the cooling systems used can be implemented with almost all laws of stage swirling.
The nozzle apparatus of the first stage of the turbine is usually made collapsible with double-support nozzles, since it perceives the greatest pressure drop, but with the necessary freedom of thermal expansion (Fig. 1, a). All new ones have cooled nozzle vanes with air outlet mostly to the trailing edge. This air, mixed with the main gas flow, works in subsequent turbine rims, so its consumption does not cause much damage to the efficiency of the turbine. Hollow cooled nozzle vanes are made by precision casting (according to investment models). The first turbine stage of the GTK-16 TMZ unit has brazed blades.
For nozzle devices of subsequent stages in stationary practice, cantilever-fixed blades are used (Fig. 1, b). At the turbo engine plant, they are combined into packages (segments) of three or four pieces, and between the packages they leave

Solo blade designs

a)

b)

in)

a - double-support air-cooled nozzle blade; b - cantilevered
turbine guide vane; c - adjustable nozzle apparatus with spherical limiting surfaces.

Rice. one

Sections of the profile part of the cooled nozzle blades

a - convective cooling with a deflector; b - convective-film cooling; c - penetrating cooling; d - intra-wall cooling;
1 - deflector; 2 - cast blade; 3 - porous coating; 4 - heat-shielding coating.
Rice. 2

Non-separable nozzle devices are used in the form of welded diaphragms. They require special design measures to ensure thermoelasticity and to avoid leash. Hollow and thin-walled diaphragms without a horizontal split are preferred.
Uncooled nozzle vanes are also desirable to be hollow to reduce thermal stresses in trailing edges during sudden stops. In all cases, it is necessary to minimize the heat removal from the nozzle vanes to the stator parts that fasten them.
Two-shaft and three-shaft nozzles require a strict tolerance for the area of ​​the outlet section of the first stage of each turbine to ensure the calculated distribution of heat drops between them. In working condition, the area of ​​the high and low pressure turbines increases by a different amount.
special attention in the design require adjustable nozzle devices. To reduce the radial clearances at the ends of the blades, the meridional surfaces adjacent to the rotary guide vanes should be made along spheres described by radii from the center located at the intersection of the blade trunnion axis with the turbine axis (Fig. 1, c). Simplification of the design is achieved with a relatively small number of wide blades, however, the axial clearance between the nozzle and working blades changes more strongly when they turn. The required operating range for changing the nozzle area is ±10%.
Among the various designs of cooled nozzle blades, deflector blades are more common than others (Fig. 2, a). The outer load-bearing shell is usually made by precision casting. The plug-in thin-walled deflector makes it possible to organize good convective cooling of the walls and jet cooling from the inside of the leading edge of the blade. The coolant leaves the blade most often through the leading edge, performed by the hollow, or near it. In such blades, the coolant moves across the axis of the blade. In the early designs of cooled nozzle apparatuses of the first stage, a longitudinal flow of the coolant was used without air venting into the edge. Now, due to the small cooling effect, such designs are rarely used and only for the second or third stage.
Advantages of a blade with a plug-in deflector for cooler cross-section:
convergence of the heat transfer coefficients of air and gas, which gives a uniform temperature over the cross section of the blade;
the possibility of implementing differentiated cooling of the blade sections in height and in cross section due to the location and number of holes in the deflector;
the ability to control the depth of cooling of the blade in the process of finishing or increasing the resource;
comparative simplicity of intensifying heat exchange from the air due to various turbulators.
The deflector is a thin-walled shell of two parts, connected by spot or roller welding, sometimes soldering. It is possible to manufacture a deflector by deformation and reaming of a thin-walled tube. Perforation of the deflector in certain places makes it possible to intensify convective heat transfer due to jet cooling. The concentration of jet cooling in one place is called shower cooling.
Nozzle blades with convective-film cooling are used for higher gas temperatures (Tg > 1200 - 1250 ° C) than with purely convective. This consumes more cooling air than without blown cooling film. However, for nozzle vanes of the first stage, this is not critical. The advantage of convective-film cooling of the blades (Fig. 2, b) is the possibility of an additional decrease in the metal temperature by 100 °C or more. Another advantage is the possibility of eliminating local overheating of the blade by creating an additional blowing slot in front of the section with an overestimated temperature. However, the film is quickly washed out and the blowing slots must be repeated. In addition, the effect of the injected film on the boundary layer causes an increase in aerodynamic losses. During film cooling, there is usually an uneven temperature across the blade cross section.
In domestic driven nozzle blades with convective-film cooling at the end of the 80s were not yet widespread, but they appear in the new 90s.
Among the cooling systems for nozzle blades that are being developed, but not put into practice, we will mention blades with penetrating cooling and blades with intra-wall cooling.
Penetrating cooling, in which air passes through small holes (pores) in the blade wall, is designed for very high temperatures, for example Tg = 1600 °C. It is under these conditions that a significant reduction in cooling air consumption can be achieved compared to convective-film cooling. Penetrating cooling is more closely related to the technology of manufacturing blade walls than other cooling methods. As a rule, nozzle vanes with penetrating cooling are sleeve vanes, i.e. a thin shell covers the hard core of the blade (Fig. 2, c). Significant disadvantages are the need for thorough purification of the cooling air and the risk of pores being blocked by dispersed particles contained in the combustion products.
Another promising type of sleeve (shell) blades is blades with intra-wall cooling. Here, the longitudinal flow of the coolant is used (Fig. 2, d).

2. Materials used for the manufacture of blades

The temperature of the metal of the nozzle vanes is determined by the temperature of the working fluid, washing the blades of a given stage, and by the cooling system. The bending stresses arising under the action of a gas flow are 50-80 MPa, and in promising high-temperature powerful ones they reach 130 MPa.
The blades are subjected to static and dynamic effects of the gas flow. In this case, temperature drops such as thermal shocks up to 400 0С are possible, and in promising ones up to 600 -700 0С. For drive turbines, the number of starts per resource reaches 200, for peak ones - 5000. The blades are also exposed to erosive and corrosive effects of the flow of combustion products at a speed of up to 700 m/s. The dust content of the flow with solid particles up to 100 microns in size can reach a concentration of 0.3 mg/m3. Under unfavorable atmospheric conditions, these values ​​can briefly increase up to 250 µm and 2.5 mg/m3, respectively. In the presence of air cleaning devices, the dust content of the air flow should not exceed the established norms.
An analysis of the conditions in which the blades operate and a study of typical accidents of blades determined the following requirements for the material of turbine nozzle blades:
A) high heat resistance, i.e. maintaining high strength at high operating temperature;
B) high plasticity required for uniform distribution of stresses over the entire cross-sectional area of ​​the blade; good resistance to local stresses;
C) high fatigue strength (endurance);
D) high damping factor;
E) the stability of the structure, ensuring the invariability of mechanical properties during the operation of turbines;
E) high resistance to oxidation and scale formation at high temperatures;
G) favorable technological properties, allowing the use of more rational methods of processing the blades (primarily by cutting) and ensuring the exact execution of the profile size and high purity of processing. Blade metal should be well forged, stamped, riveted without cracking, bent well and rolled in a cold state. In the case of welded structures, good weldability is required from the metal of the blades.
H) High resistance to erosion.
Cast or wrought nickel-based alloys are used as the material for the nozzle blades of the first stages. At gas temperatures up to 700 °C, austenitic steels were previously used. For the blades of the last stages at a gas temperature of less than 580 ° C, it is also possible to use alloyed chromium steels. For blades operating at temperatures above 650 to 8000 C, nickel-based heat-resistant metal alloys are used. Among them are ZhS6K, EI929VD, EI893, N70VMYUT, KhN80TBYu, etc.
At gas temperatures of 800°C and above, and in the presence of sulfur in the fuel gas and at 720°C, it is necessary to apply protective coatings on nozzle and rotor blades having a chromium content in the alloy of less than 20%, by chromium aluminizing, chromosiliconizing or chromoaluminosilicizing, etc. The thickness of the protective coating is 30 - 60 microns. Enamel coatings are also used, and for cooled blades, heat-shielding coatings.

3. Type of workpiece

The following types of blanks are used for the manufacture of blades: flat steel, sheet steel, forgings, stampings, hot-rolled profile strips (the so-called light-rolled profile) and investment casting. The most common blanks for blades are light-rolled sections and forgings.
The type of workpiece has a great influence on the subsequent technological process of processing, therefore, when choosing rational workpieces, one should take into account all the specific production conditions and, in particular, the shape of the blades, their number and the timing of orders.
The main method of manufacturing nozzle blades is precision investment casting, mainly from casting alloys LK4, ZhS6, ZhS6-K, etc.
The use of precision investment casting makes it possible to obtain blanks with a minimum feather allowance. Mechanical restoration blanks of such blades is mainly in the processing of the locks of the blades.
Lost wax casting has the following advantages over other methods for producing blanks for nozzle blades;
1) the possibility of obtaining workpieces of complex shape, with a surface finish of 5-b and an accuracy within the 4th class;
2) the possibility of obtaining hollow blades with a wall thickness of up to 0.5 mm.
The disadvantages of this method include:
1) the need to use expensive alloys and auxiliary materials for casting;
2) the duration of the production cycle.
In some engines, the blades of the nozzle apparatus began to be made from sheet heat-resistant material by cold stamping, followed by electric welding of the trailing edge.

4.Basic requirements for blade machining

The good quality of the blades, like all other parts of the turbine, depends on the correct implementation of the design dimensions specified in the drawings and the cleanliness of the surface treatment. Each part of the blade (tail, working part and head) has a different purpose. The tail serves to securely fasten the blades in the turbine housing. The working part is intended for the perception of steam pressure, and the head for fixing the bandage. If at the tail of the blade, in accordance with its official purpose, the degree of accuracy with which all landing dimensions of the tail are made is of great importance, then for the working part, the dimensions of which are not landing, the degree of cleanliness of processing is of great importance. A well-polished surface of the working part helps to reduce steam losses due to friction on the surface of the blade, while at the same time increasing the anti-corrosion resistance of the blade.
All sizes of blades, according to the requirements for their accuracy, can be divided into three groups.
First: the dimensions on which the nature of the connection of the blades with other parts of the turbine depends, i.e. landing details. These include, first of all, the dimensions of the tails and spikes for the attachment of bandage tapes. The diameter of the stud (with a round stud) and the width and thickness of the stud (with a rectangular stud) are performed according to running landings of the 4th class.
Second: dimensions that are not landing, but require increased accuracy. These include the dimensions of the sections of the working parts; dimensions that determine the installation of the blades and the location of the holes for the fastening wire, etc. These dimensions are performed either according to the third and fourth accuracy classes, or according to free non-standard tolerances ranging from 0.1 mm to 0.5 mm, depending on the size of the blade.
Third: free dimensions, which usually include the dimensions of fillets, chamfers and other less critical elements of the blades. The accuracy of free dimensions is either not standardized at all or is limited to tolerances of the 7th accuracy class. However, even in the case when no tolerances are established for free dimensions, they are usually carried out according to the tolerances established for free dimensions by special technological instructions issued by a given enterprise.
The cleanliness of the processing of the seating surfaces is maintained within the 6th class, working profiles and fillets at the working parts - 8-9th class.
The landing dimensions of the tail connections are the most responsible. These dimensions, as well as the finish of processing, must be ensured by the appropriate precision of machining and the quality of the cutting tool. A drawing of a typical nozzle blade is shown in fig. 3.

Drawing of a typical nozzle blade

a)

b)

a - lockless design, b-with a lock.

Rice. 3

The accuracy of manufacturing the main surfaces of the blades is characterized by the following data:
nib profile thickness tolerance ………………… +0.5 - 0,2;
edge thickness tolerance ………………………. ±0.2;
profile curvature..……………………. 0.8mm;
trailing edge curvature……………. 0.8mm;
tolerance for wall thickness of hollow blades ..... ± 0.3 mm;
cleanliness of the surface of the lock ……………………………………………………………………………………………………………………………………………………………………………………………………………………………………………………………………………………………………………………………………

5. Typical machining process

The technological process of processing any new blade can be easily and quickly developed by a technologist in the presence of a classifier and typical technological operations.
The alloys from which the blades are made are poorly processed by cutting (especially with a metal tool). In this regard, the operations for processing these blades are usually performed by grinding.
For blanks of blades of the nozzle apparatus, made by precision casting with a feather allowance for grinding, the main type of machining is grinding of locks.
The blade feathers are usually finished by hand on polishing headstocks. The initial stripping of the pen is carried out abrasive wheels grain size 46-60.The route technological process of machining the blades of the nozzle apparatus (with locks) consists of the following operations:

№ operations	the name of the operation	Equipment
	Workpiece control
	Grinding reference planes	Surface grinding machine MSZ
	Bench cleaning of the trailing edge flush with the main surface
	Lapping of the side planes of the castle from the side of the trough	Lapping machine
	Grinding the planes of the castle	Surface grinding machine MSZ
	Sprue grinding	Surface grinding machine MSZ
	Grinding two planes of the lock from the back	Surface grinding MSZ
	Electroerosive machining of holes in the lock	Special installation
	flushing	Washing machine
	Milling a groove on the sole of the lock	Vertical milling machine
	Locksmith (dulling of sharp edges after machining)
	Washing and blowing	Washing machine
	Ultimate control
	Color flaw detection	Special installation
	Cleaning of defective areas after color flaw detection	polishing head
	Etching
	Control after cleaning of defective places
	Luminescent control
	Cleaning of defects after luminescent control	polishing head
	Washing and wiping	Washing machine

The route technological process of machining the blades of a lockless nozzle apparatus consists of the following operations:

operation number	the name of the operation	Equipment
	Blank - precision casting without allowance for machining on pen
	Grinding the end of the pen	Surface grinding machine MSZ
	Radius milling from input side­ edge	Horizontal milling machine
	Radius milling from input side edge	Horizontal milling machine
	Bench deburring after milling and blunting sharp edges	polishing head
	Washing and blowing	Washing machine
	Ultimate control
	Color flaw detection	Special installation
	Cleaning defects after color flaw detection	polishing head
	Etching
	Control after stripping
	Luminescent control	Special installation
	Deburring after fluorescent inspection	polishing head
	Washing and wiping	Washing machine

Next, the pen is polished with felt circles with a glued abrasive. Polishing is carried out in three passes. The grit size of the abrasive used in this treatment is 60, 180 and 220, respectively.

6. Type of machines

Due to the high labor intensity of manual adjustment of the profile at individual plants, attempts were made to mechanize these operations.
On fig. 4 shows a modernized PSL machine for polishing the back of the blades of the nozzle apparatus. This machine can process several parts at the same time.
MSH-81 and MSH-82 machines of the Moscow plant of grinding machines (Fig. 5) are designed for processing lockless nozzle blades, the back and trough of which have a constant profile in all sections. The pen is processed with a profile circle, which is corrected with a special profile cutter. On fig. 6 shows a special device used on cylindrical grinding machines for grinding the back of the blades of the nozzle apparatus.
The device consists of a mechanism for synchronous rotation of the spindle of the grinding wheel and the spindle of the front beam, a mechanism for dressing the grinding wheel and a mechanism for driving the copier.
Spindle 3 of the headstock receives rotation from the spindle of the grinding head through a system of gears to ensure the synchronism of the rotation of the wheel and the workpiece.
From the spindle of the product, rotation with a gear ratio of 2: 1 is transmitted to a volumetric copier 2, which serves to dress the grinding wheel. Circle 9 is corrected using a special mechanism. On the shaft 10 of the mechanism for dressing the circle, a lever rigidly sits, carrying a profiling tool 8. At the other end of the shaft 10, a roller 11 is mounted, connected to the roller 6, abutting against the volume copier 12. The dressing mechanism moves along the axis of rotation of the grinding wheel. For preliminary grinding of the volume copier, a reference blade 6 is used, against which the disk 7 rests, replacing the grinding wheel.
When the reference blade 6 rotates, the disk 7 receives a horizontal movement, which is transmitted through the lever of the shaft 10 of the dressing mechanism to the grinding wheel mechanism, which grinds the profile of the three-dimensional copier.
After grinding the volumetric copier, instead of the grinding wheel, a roller 11 is installed, the diameter of which is equal to the diameter of the circle. Instead of a disk sector, diamond 8 is installed, which profiles the grinding wheel. After dressing the grinding wheel, the back of the blade installed instead of the reference blade is processed.
The blades of the nozzle apparatus of a number of gas turbine engines are made by the method of precision investment casting with an allowance for grinding on the feather.
In this case, the technological process of processing the blades includes (in addition to the indicated operations) also operations for grinding the blade profile, performed on machines XSh-185V, XSh-186 and on modernized universal grinding machines.
Hollow design nozzle blades are widely used in high-temperature gas turbine engines. Such blades are also made by investment casting, with ceramic or other rods forming an internal cavity.
The locks of the blades of the nozzle apparatus are processed on surface grinders. The processed blade is installed in a special cassette. In this case, the surface of the trough and the edge of the pen serve as bases. The clamp is carried out on the surface of the back. The required arrangement of the planes of the locks is achieved by turning the cassette and installing it with the corresponding surfaces fig. 7.
Processing of the bases of the blades of the nozzle apparatus can be carried out on a semi-automatic surface grinding machine model BS-200. The machine works on a semi-automatic cycle and provides a uniform distribution of the allowance between the back and trough. The machine has an electronic device for uniform distribution of the allowance along the profile of the pen, as well as a device for diamond-free dressing of the circle. Parts are fastened in a special fixture with a quick-release clamp.

7. Fixing workpieces

During processing, the workpiece (part) is oriented accordingly, must be stationary. This is achieved by fixing it in a fixture or on a machine.
In contrast to the basing of the workpiece, when a different number of bonds are imposed on it and it is deprived of three, four, five and six degrees of freedom, in all cases of fastening the workpiece must be deprived of six degrees of freedom.
For this purpose, various clamping devices(mechanical, hydraulic, pneumatic, magnetic, vacuum, etc.) based on the use of friction forces.
The clamping devices in the fixtures must create a constant contact between the bases and the reference points (ensure correct basing) and immobility of the workpiece during its processing (fixing the workpiece).
It should be noted that the smaller the number of bases and reference points used when basing workpieces, the simpler, more productive and cheaper the design of fixtures is. Therefore, when basing workpieces to be processed, it is necessary to strive to use the smallest number of bases with the smallest number of reference points, at which the dimensions and shape of the part specified by the drawing can be ensured.

Polishing the back of the feather blades of the nozzle apparatus
on a modernized PSL machine

General view and working area of ​​the surface grinder
MSh-81 and MSh-82 models

Rice. 5

Grinding the back of the nozzle blade
on a modernized copy-grinding machine

1—stops, 2—copier, 3—spindle, 4—frame for fixing the reference blade, 5—blade, 6—reference blade, 7—disk, 8—diamond, 9—grinding wheel, 10—shafts of the dressing mechanism, 11— roller, 12—copier disk.
Rice. 6

Grinding the planes of the locks of the blades of the nozzle apparatus

Rice. 7

8. Technical control of blades

The blades are checked both in the process of machining and after its completion. Blade control includes:
detection of external and internal material defects; checking the roughness of the treated surfaces in accordance with the requirements of the drawing; checking the dimensions, shape of the pen profiles (back, trough) and locks and their relative position; determination of the mass and frequency of natural oscillations of the blades; selective testing of turbine and compressor rotor blades for fatigue. In hollow cooled LPT working blades, the water flow rate through the internal cavity is checked (blade shedding tests).
The control of external and internal defects in the material of the blades makes it possible to identify cracks and hairlines on the surface, shells, porosity, delamination, foreign inclusions and flakes in the material. For this purpose, etching, color flaw detection, luminescent, magnetic and ultrasonic control methods are used.
The magnetic particle method is based on the attraction of iron powder particles to the magnetic poles that form near a magnetized part in places of discontinuity. The magnetic particle method reveals cracks with an opening width of 0.001 mm or more, a depth of 0.01 mm or more. The relative simplicity and rather high reliability of this method contributed to its widespread implementation.
Color and luminescent inspection methods (capillary flaw detection methods) are used to detect defects that appear on the surface of a part. The color flaw detection method is based on the ability of a special red paint to penetrate deep into surface defects and white paint to absorb red paint from a defect. The method detects cracks wide from 0.01 mm, in depth from 0.05 mm and in length from 0.3 mm.
Luminescent method(LUM-A) is based on the ability of some liquids to glow when irradiated with ultraviolet light. The LUM-A luminescent method reliably detects cracks, pores, friability, oxide films, blockages, etc. emerging on the surface. It detects cracks as small as 0.01 mm wide, as deep as 0.05 mm, and as wide as 0.2 mm. The sensitivity of the LUM-A method is slightly higher than the color flaw detection method. Internal defects in the material of the blades are checked by X-ray and ultrasonic methods.
The X-ray method for detecting defects is based on the attenuation of X-ray radiation by the material of the part, in which the shadow image of the translucent part is recorded on the X-ray film. The advantage of the method is its high sensitivity to the detection of internal pores, shells, foreign inclusions, etc. in the material.
For transillumination of cast turbine blades, mobile cable X-ray machines such as RUP-100-10, RUP-150-10-1, etc. are used.
The ultrasonic testing method using surface waves makes it possible to detect surface cracks and metallurgical defects of the material. This method is usually used to detect cracks in the leading and trailing edges, less often on the surface of the back and trough, which occur during the manufacture and operation of the blade. echoes) from defects.
Control of geometric dimensions, shape of pen and lock profiles and their relative position. Operations of this type of technical control of the blades are the most time-consuming. The devices used in these operations can be divided into two main groups: non-contact - optical-projection and contact - mechanical, optical-mechanical, pneumatic and pneumohydraulic.
The blade feather is checked in the calculated cross sections by non-contact and contact methods. One of the non-contact methods of control is the profile check on projectors used in single production. We have not found any use for them.
On a small scale of production, the profile of the blade feather is sometimes checked with templates. The deviation of the profile of the back and trough from the template is determined visually through the light or using a probe. Pen control by templates is inefficient, subjective and requires a cumbersome template-measuring economy.
In mass production, mechanical instruments with dial indicators adjusted according to the reference blade were used. They are simple and easy to use, but inefficient.
Multidimensional instruments and measuring machines are productive. They can be quickly reconfigured to control other blades using a reference blade. The base for fastening the blade is a lock or center recesses, two of which are on the side surfaces of the lock and one at the end of the feather. Such devices include universal multidimensional optical-mechanical devices of the POMKL type for simultaneous control of the airfoil profile, displacement of the airfoil from the lock axis, twist angle and airfoil thickness in cross sections of the compressor blade.
The main geometric parameters of the turbine and compressor blade locks are usually checked by mechanical instruments with indicator clocks adjusted according to the standard.
The flow of water through the inner cavity of the feather of the cooled LPT blades is checked on a special installation. The spatula is installed in the fixture and spilled with water when overpressure at 4 ± 0.05 kgf / cm2 (0.3 ± 0.005 MPa) and a temperature of 20 ± 5 "C for 20 s. Check the capacity of the internal channel for the entire I set of blades of this stage. Compare the average flow rate with the result of the passage of each blades in the set The difference in water consumption for the working blades in the set (difference) should be no more than 13 ... 15% of the average water flow in the set of blades
The frequencies of natural oscillations of the turbine and compressor blades are checked on electrodynamic vibration stands.
The rotor blades of the turbine and compressor are weighed on a balance of the VTK-500 type with an accuracy of 0.1 g.

9. Real implementation of the technological process at UTMZ

Let's consider a real technological process on the example of the guide vane of the first stage GTN-6U. Workpiece type - investment casting, workpiece material - KhN648MKYUT alloy - USZMI - ZU.
Real execution of the technological process in the factory for guide vanes
6-11 stages of the GT-6-750 turbine are presented in Table. 3.
Table 3

Operation No.	Name and content of the operation	Equipment
	Input control
	Milling and centering. Trim ends and center on 2 sides.	center. milling MP-71
	Horizontal milling. Mill the planes of the tail from the side of the inner and outer profile in the centers.	Horizontal milling 6M82G
	Grinding. Grind the plane of the tail from the side of the outer profile in the centers.	surface grinding 3B-722
	Grinding. Grind the plane of the tail from the side of the inner profile	surface grinding 3B-722
	Horizontal milling. Mill the plane of the tail at an angle from the side of the gas outlet in advance in 2 passes.	Horizontal milling 6M83G
	Vertical milling. Milling the plane of the tail at an angle from the side of the gas outlet is clean.	Vertical milling 6M13P
	Horizontal milling. Mill the plane of the tail from the input side at an angle beforehand.	Horizontal milling 6M82G
	Vertical milling. Mill the plane of the tail from the entry side at a clean angle	Vertical milling 6M13P
	Turning. Sharpen the threaded shank.	Turning P.U. 16K20F3
	Vertical milling. Mill the inlet and outlet sides along the length of the working part.	Vertical milling FK-300
	Horizontal milling. Milling the fillet on the gas inlet side is clean.	Horizontal milling 6M83G
	Horizontal milling. Mill the fillet on the gas outlet side cleanly.	Horizontal milling 6M83G
	Vertical milling. Mill the fillet of the inner and outer profile at an angle of 1050’ in 11 rows (except for the 11th step) flush with the main profile.	Vertical milling 4FSL-4A
	Vertical milling. Mill the fillet of the inner and outer profile in a straight line for 11 lines flush with the main profile.	Vertical milling 4FSL-4A
	Grinding. Sanding the inside and outer profiles simultaneously in the centers for 400 lines	grinding LSh-1A
	Control. Operation control 16.
	Locksmith. Saw down the radii on the shoulders from the side of the inner and outer profile of the entrance and exit according to the templates; chamfer 1x450
	Grinding. Grind the fillet of the inner and outer profile flush with the main profile; grind the leading edge.	Polishing
	Locksmith. File the exit edge.
	Ultimate control.
	Cut-off. Cut off the base from the end of the working part.	Abrasive cutting
	Grinding. Polish the outer and inner profile, leading edge and fillets.	Polishing DSh-96
	Locksmith. Polish the trailing edge by hand.
	Locksmith. Mark the designation of the blade.
	Control. Check for cracks.
	flushing
	Ultimate control	control plate
	vibration test

10. Suggestions for improving the technological process

The expansion of the serial production of steam and steam engines, caused by the tasks of developing the country's energy and gas industry, contributed to accelerated technical progress in turbine building.
Particularly significant progress in this direction has been achieved in the production of turbine blades. At all stages of the technological process, starting with the preparation of the main base surfaces, special machines and CNC machines are used. The introduction of multi-spindle machines for circular milling of the inner and outer profiles of the working parts of long blades with transverse lines became the most important measure to increase labor productivity and improve quality.
The transfer of the processing of a certain range of blades to machine tools with program control made it possible to combine several operations into one and thereby reduce the blade preparation cycle, free the worker from performing heavy manual work, increase the accuracy of processing in terms of size and roughness by eliminating resets and working in design modes cutting.
Among promising works required scientific substantiation and implementation should be called the following:
- improving the production of stamped blanks in terms of reducing machining allowances;
- mechanization of grinding work on fine-tuning the profiles of the working parts of long blades;
- carrying out research work to determine scientifically substantiated parameters of permissible deviations from the design dimensions of the profile parts, respectively, of the length and width of the working and guide vanes.
Significant technical progress in turbine building will be achieved through the organization of centralized design and manufacture of blades at one specialized plant with a wide typification of blades and, thus, the transfer of their machining in flow and automatically operating lines, preparation for which is practically already being carried out at the present time at the turbine plant. blades (LZTD).
An important factor technical progress this event will bring the process of designing blades closer to their production. GTU-UPI 2002

Please read before asking a question:

The invention relates to the field of mechanical engineering, and in particular to methods for manufacturing blades of aircraft gas turbine engines (GTE) from materials that can be deformed in a cold or hot state. The blade blank is made. An aerodynamic profile is formed in each section of the feather. Form a shank. Carry out finishing operations. The formation of an aerodynamic profile and a shank is carried out by simultaneously twisting the feather and the shank and calibrating them in a die. A flat billet is made with sections, the area and length of which are equal, respectively, to the area of ​​the corresponding sections of the stamped blade and the length of the chords of these sections. As a result, an increase in the metal utilization factor and manufacturing accuracy, an increase in the quality of wide-chord GTE blades and a reduction in time costs are ensured. 2 ill.

The present invention relates to the field of mechanical engineering, and in particular to methods for manufacturing blades of aircraft gas turbine engines (GTE) from materials that can be deformed in a cold or hot state.

In modern designs of aircraft engine fans, large-sized wide-chord blades are widely used, which make it possible to significantly reduce fan noise, increase thrust, and generally increase the efficiency of a gas turbine engine.

Traditional technologies for the production of blades are known, including the manufacture of a blade blank by stamping with a phased twisting of the blade profile and allowances for the blade and lock, followed by removal of allowances by machining, electrophysical and other methods (Krymov V.V., Eliseev Yu.S., Zudin KI Production of blades for gas turbine engines, M., "Engineering / Engineering - Flight", 2002, pp. 66-100, 101-197).

This method becomes extremely time-consuming and metal-consuming in the production of wide-chord blades due to their large dimensions (the length can reach 1.5 m, with a height-to-chord ratio of less than 2) and complex geometric shape.

The complex configuration of the preliminary transitions reduces the manufacturability of the accompanying operations, from the cleaning of stamping defects to the use of specialized lodgements for heating before the next stamping transition.

Reducing the machining allowance for the profile of the pen leads to an increase in the specific forging forces, and at the same time obtaining its final configuration requires an increase in the rigidity of the die set assembly to absorb high shear forces during forging.

Simultaneous final finishing of the profile of the pen in terms of thickness and configuration, despite the known methods of mechanical, chemical and electrochemical milling, is a highly laborious operation.

A known method for manufacturing gas turbine engine blades (RF patent No. 2257277) is a prototype. The essence of the method lies in the fact that at the first stage of designing the technological process, the design drawing of the blade is processed, unwinding and pushing apart the calculated sections of the pen, while “laying” the chords of the unwound sections in one plane. The resulting modified drawing of the blade is the basis for the design of the blank-forging. A stamping blank with an untwisted feather profile is manufactured by volumetric stamping with an allowance for the feather and lock for further cutting. After removing the rough allowance, for example, by milling, the pen profile is twisted in a hot state using special devices. Subsequently, the workpiece made in this way is subjected to all the traditional stages of the technological process of manufacturing the blade.

The disadvantage of this method is that the determination of the power parameters by calculating the process of hot twisting of the blade feather, having an airfoil cross section, is problematic along the length. the analysis of existing mathematical models for determining the force parameters during twisting is limited to the consideration of rods with elementary geometric sections (circle, ellipse, square, rectangle). Therefore, deformations during the twisting of the product inevitably lead to a distortion of the airfoil, which can exceed the tolerance field. The selection of technological modes and geometric parameters of the workpiece requires a large number of laborious and time-consuming experimental work for each type of wide-chord blade size. The process is not stable, depends on many factors and requires special equipment.

To eliminate the above negative aspects, it is proposed to separate the operations: the formation of the delivery thickness of the pen profile and the formation of its contour. Additionally, this allows you to significantly expand the range of equipment for performing the first stage, and all related operations of adjusting and machining of this stage are carried out on a straightened contour that is more technologically advanced in processing.

The present invention attempts to provide new method production of gas turbine engine blades with contoured, one-pass isothermal flashless final stamping (twisting + calibration), which reduces or solves the above problems.

The invention solves the problem of manufacturing wide-chord GTE blades of complex geometric shape using standard equipment.

The technical result of the present invention is to improve the quality of manufacturing wide-chord GTE blades, as well as the stability of the process while reducing costs.

A method for manufacturing blades of a gas turbine engine, including manufacturing a blank of a blade, forming an aerodynamic profile in each section of the blade feather, forming a shank and performing finishing operations, forming an aerodynamic profile in each section of the feather of a blade and forming a shank is carried out by simultaneously twisting the feather and the shank and calibrating them in a die by isothermal forging, at the same time, a flat blank is made, made with sections, the area and length of which are equal, respectively, to the area of ​​the corresponding sections of the stamped blade and the length of the chords of these sections.

The essence of the invention is illustrated by drawings, which show:

figure 1 - wide-chord blade 1, made, for example, of titanium or one of its alloys;

figure 2 - straightened workpiece wide-chord blades.

Proposed according to the invention, the method of manufacturing blades of gas turbine engines is as follows.

1. The production of a slab 4 (figure 2) by extrusion and (or) precision stamping, as well as rolling and (or) upsetting and (or) machining of flat or section products.

2. Preparation of basic elements 3 for subsequent finishing machining of the pen and at the same time laying elements for single-pass stamping or at the stage of precision stamping of the workpiece and (or) additional fur. processing previously obtained workpieces or obtained by welding to the workpiece 4 and additional fur. processing.

3. Preparation of a planned projection of the workpiece for single-pass stamping or at the stage of precision stamping of the workpiece and (or) additional fur. processing of previously obtained blanks (this ensures the equality of the chords of the pen blank 6 and the chords of the finished product 7).

4. Preparation of high-altitude dimensions of the workpiece for single-pass stamping or at the stage of precision stamping of the workpiece and (or) additional fur. processing previously obtained blanks.

5. Application of heat and pressure to the workpiece for isothermal stamping (simultaneous twisting of the airfoil ("feather") 1 and shank ("lock") 2 with simultaneous calibration) and the production of essentially the required finished external configuration and dimensions of the profile of the feather. For high-angle swirling of the airfoil (more than 40°) and calibration of wide-chord fan blades, specially introduced holding elements of die equipment (not shown) are used.

6. Finishing the product to remove excess material from the leading and trailing edges (5) of the isothermally stamped outer configuration to obtain a finished feather profile.

7. Removal of the base (laying) elements 3 of Fig.1.

8. Mechanical processing of the blade shank (“lock”) 2.

An example of a specific implementation. An experimental stamping of a wide-chord GTE blade in a closed die was carried out. Material - titanium alloy brand VT6. The stamping temperature is not more than 850°C. The tool was heated to a temperature not exceeding 850°C. Finished blade dimensions: length - 1200 mm, maximum chord width 620 mm.

The proposed method for the manufacture of wide-chord blades makes it possible to develop an effective technology that can be used to produce a number of blades for gas turbine engines from progressive metals and alloys.

The advantage of the proposed technical solution allows you to expand the technological capabilities of standard equipment, to conduct the process with minimal time. The utilization rate of the metal is significantly increased, the accuracy of manufacturing and the stability of the technological process are increased.

A method for manufacturing blades of a gas turbine engine, including manufacturing a blank of a blade, forming an aerodynamic profile in each section of the blade airfoil, forming a tail and performing finishing operations, characterized in that the formation of an airfoil in each section of the blade airfoil and the formation of the tail are carried out by simultaneously twisting the airfoil and the tail, and their calibration in a stamp by isothermal forging, while making a flat workpiece made with sections, the area and length of which are respectively equal to the area of ​​the corresponding sections of the stamped blade and the length of the chords of these sections.

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