Development of group technologies for manufacturing the main parts of gas turbine engines

The utility model relates to the field of engine building and can be used in the blades of a gas turbine engine (GTE) for aviation, ship and ground (as part of a power plant) application. The utility model solves the problem of increasing the bending fatigue strength of a blade by reducing tensile stresses in its lock in order to avoid premature failure of the blade. Additional task is the possibility of applying the proposed solution to the cooled blades of gas turbine engines. The problem is solved by the fact that the GTE turbine blade contains a Christmas tree lock, on which a voltage concentrator is made in the form of a hole. New in the proposed utility model is that the hole is located along the axis of the GTE blade. The blade may contain a channel that communicates with the hole, forming a single stress concentrator. This design of the herringbone lock of the GTE turbine blade increases the bending fatigue strength of the blade by reducing the tensile stresses in its lock, which makes it possible to avoid premature failure of the blade.

The utility model relates to engine building and can be used in the blades of a gas turbine engine (GTE) for aviation, ship and ground (as part of a power plant) application.

Known for the design of the turbine blades of the gas turbine engine, containing a Christmas tree lock (Skubachevsky G.S. Aircraft gas turbine engines. Design and calculation of parts. - M.: Mashinostroenie, 1981, p. 89, Fig. 3.27).

The disadvantage of a blade with such a lock is that it does not provide for the implementation of the stress concentrator. The absence of a concentrator leads to the destruction of not only the blades, but also the disk when the load is suddenly removed.

Also known is the design of the GTE blade, containing a Christmas tree lock and at least one stress concentrator in the form of a hole in the lock located across the axis of the blade (Patent GB 1468470 dated 03/30/1977).

The disadvantage of this design is that the Christmas tree lock during operation is subject to tensile stresses, the increase of which leads to insufficient bending fatigue strength. The result is premature failure of the GTE blade. Also, this design cannot be used in cooled blades, as there is a leakage of cooling air.

The technical objective of the utility model is to increase the bending fatigue strength of the blade by reducing the tensile stresses in its lock in order to avoid premature failure of the blade.

An additional technical challenge is the possibility of applying the proposed solution to cooled GTE blades.

The problem is solved by the fact that the GTE turbine blade contains a Christmas tree lock, on which a voltage concentrator is made in the form of a hole.

New in the proposed utility model is that the hole is located along the axis of the GTE blade.

In addition, the blade may contain a channel that communicates with the hole, forming a single stress concentrator.

The proposed drawing shows a longitudinal section of a gas turbine turbine blade.

The gas turbine engine blade includes a Christmas tree lock 1. The Christmas tree lock 1 contains a stress concentrator in the form of a hole 2 made along the axis 3 of the blade.

The GTE turbine blade is equipped with a channel 4 for cooling, which is connected with the hole 2.

During operation of the GTE turbine wheel, in the event of a failure due to a sudden removal of the load, the disk rotation speed increases under the influence of increasing centrifugal forces. In turn, centrifugal forces increase the compressive and bending stresses in the Christmas tree lock 1 and in the disk (not shown in the drawing), while the tensile stresses are reduced due to the presence of a stress concentrator in the form of a hole 2 made on the Christmas tree lock 1 along the axis of the blade. This leads to an increase in bending fatigue strength in the blade lock, which avoids premature failure of the blade.

The turbine blade of the gas turbine engine operates as a cooled blade when the air passes through the channel 4 for cooling, which is connected with the hole 2 for cooling the fir-tree lock 1 of the blade.

This design of the GTE turbine blade makes it possible to increase the bending fatigue strength of the blade due to the reduction of tensile stresses in its lock in order to avoid premature destruction of the blade; it can be applied to cooled GTE blades.

Utility model formula

1. A turbine blade of a gas turbine engine containing a Christmas tree lock, on which at least one stress concentrator is made in the form of a hole, characterized in that the hole is made along the axis of the blade.

2. The turbine blade of a gas turbine engine according to claim 1, characterized in that the blade contains at least one channel for cooling, which is in communication with the hole.

Candidate of Technical Sciences I. DEMONIS, Deputy General Director of VIAM.

Jet aviation, which began to be created from the 1940s, required the development of a new type of engine. Those who received the most wide application gas turbine jet engines have revolutionized aviation technology.

Science and life // Illustrations

shoulder blades gas turbine jet engines operate in very difficult conditions: they are surrounded by a stream of hot gases from the combustion chambers.

The cooling air supplied from the side of the turbine axis to the channels of the blade exits from its end face.

Bookmark rods that are placed in a mold for casting a gas turbine blade. After the workpiece is cooled, the rods are dissolved and channels for passing cooling air remain in the finished blade.

Air escaping from holes in the side of the blades creates a thin film of air that insulates the blade from hot gases (left). The channels leading to the holes have a rather complex geometry (right).

The metal of the cast blade solidifies in the form of crystals different sizes coupled not securely enough (left). After the introduction of the modifier into the metal, the crystals became small and uniform, the strength of the product increased (right).

So produce directional crystallization of the material of the blade.

By improving the technology of directional crystallization, it was possible to grow a blade in the form of a single single crystal.

A cooling cavity is created in single-crystal blades complex shape. The latest developments in its configuration made it possible to increase the cooling efficiency of the blades by one and a half times.

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 engine, the greater its thrust, the higher the efficiency and the better the weight characteristics. The gas turbine engine has 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 (entered service with the German troops in 1942. - Note. ed.). 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 to him in Moscow on Galushkin Street a room for the first design department jet engines. Arkhip Mikhailovich found and dug out his drawings, but the engine according to his project did not work out right away. Then they simply took a turbojet engine bought from the British and repeated it one by one. But the matter rested on materials that were absent 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 operating temperature 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 long lasting high temperature could not. 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 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 indoors worker in a spacesuit 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. Our laboratory was engaged in the manufacture of cores, and I myself began to study casting technology, materials for ceramic molds, alloys and protective coatings finished products and now lead this line of research.

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. Therefore, the allowable shrinkage of the mold 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.

There is some analogy with thermal protection here. 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. But for more on this, see page 54.

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). Promising direction the use of niobium silicide is considered, 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 in the direction of the axis Z the crystal grows, and in the direction X-Y- No). 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. By the way, the first publication about this technology appeared in the journal "Science and Life" back in 1971, in No. 1.

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 exhausted their resources began to be used at gas pipeline compressor stations and in power plants that feed oil pipeline pumps (see "Science and Life" No. ). 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 Kaznetsov.

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 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, and the casting technique finished product. Huge work and knowledge of all employees of VIAM has been invested in all the blades.

See in a room on the same topic

The invention relates to foundry production. The blade of a gas turbine engine is made by investment casting. The shoulder blade contains a feather 4, at the end of which there is a heel 5, made in the form of a single piece with a feather. The heel contains a platform 5a, in which the first bath 12 is made with radial surfaces 13 and a bottom 14. The bath 12 reduces the thickness of the heel. In the first bath, at the level of the interface zone 15 between the feather and the heel, a second bath 16 is made, which allows pouring metal into the shell mold at only one point. Due to the uniform distribution of the metal, the formation of porosity in the shovel is prevented. 3 n. and 3 z.p. f-ly, 4 ill.

Drawings to the RF patent 2477196

The present invention relates to a cast metal blade and a method for making the same.

A gas turbine engine, such as a turbojet engine, includes a fan, one or more compressor stages, a combustion chamber, one or more turbine stages, and a nozzle. The gases are driven by the rotors of the fan, compressor and turbine, due to the presence of radial blades fixed on the periphery of the rotors.

The concepts of inboard, outboard, radial, forward or aft position or location should be considered in relation to the main axis of the gas turbine engine and to the direction of gas flow in this engine.

The movable turbine blade contains a leg, with which it is attached to the rotor disk, a platform forming an element of the inner wall that limits the gas-air path, and a feather, which is located mainly along the radial axis and is blown by gases. Depending on the engine and turbine stage, at its end remote from the stem, the blade ends with an element transverse to the main (main) axis of the feather, called the heel, which forms the element outer wall, limiting the gas-air path.

On the outer surface of the heel, one or more radial plates or scallops are made, which together with the opposite stator wall form a labyrinth gasket that provides tightness with respect to gases; for this, as a rule, the said stator wall is made in the form of a ring of abradable material, against which the plates rub. The plates contain a front side and a back side located transversely to the gas flow.

The blade can be monoblock, that is, the leg, platform, feather and heel are made in the form of a single piece. The blade is made by a casting process called "lost wax casting" and is well known to those skilled in the art. In this way:

Previously, a model of the scapula is made from wax;

The model is immersed in a refractory ceramic slip, which forms a shell after firing;

The wax is melted and removed, which makes it possible to obtain a "shell shape" of refractory material, the internal volume of which determines the shape of the blade;

Molten metal is poured into the shell mold, while several shell molds are combined into a block for simultaneous pouring of the metal;

The shell mold is broken, which makes it possible to obtain a metal spatula.

At the points where the metal is poured into the mold, relatively thick metal outgrowths are formed on the metal blade cast in the mold, which must be machined after the blade has been molded. As a rule, metal is poured at the level of the heel of the blade. The diameter of the pouring channel and, therefore, the subsequently formed build-up is significant, and the pouring takes place near the plates of the labyrinth gasket, which have a small thickness; as a result, if only one pouring point is provided, the metal is poorly distributed in the shell mold and there are problems with the porosity of the blade, in particular at the level of its blades.

This problem can be solved by providing two pouring inlets, while the diameter of the pouring channels is correspondingly reduced. Thus, instead of one pour channel large diameter get two filling channels of smaller diameter, remote friend from each other, which ensures better distribution of the metal and avoids porosity problems.

However, it is desirable to address these porosity problems by maintaining only one pour point.

In this regard, the object of the invention is a gas turbine engine blade, made by casting, containing a feather, at the end of which there is a heel, made in the form of a single piece with the feather, with which it is connected at the level of the interface zone, while the heel contains a platform on which, according to at least one sealing plate, and the first bath is made in the platform, characterized in that the second bath is made in the first bath at the level of the interface between the feather and the heel.

The presence of one bath in another bath at the level of the interface zone between the airfoil and the heel makes it possible to avoid too much thickening in this zone and during the molding of the blade by casting provides a better distribution of the liquid metal in the mold. The improved distribution of the liquid metal in the mold allows the casting method to be used with a single metal pour point. The advantage of manufacturing a blade with a single pour point is the exceptional simplicity of the shell mold and, if necessary, the block of shell molds; the cost of manufacturing the blades is reduced, while their quality is improved.

In addition, the amount of material at the heel level is optimized, which reduces the weight and cost of the blade.

In addition, the mechanical stresses on the heel and/or the feather are optimized and are better absorbed by the blade as a better mass distribution is achieved.

Preferably, the first bath is limited by the radial surfaces and the bottom, and the second bath is formed in the bottom of the first bath.

It is also preferable that the second tray is made along the main axis of the blade opposite the interface zone between the heel and the feather.

It is advisable that the blade airfoil be formed by a solid wall and contain curved surfaces in the mating zone, the second bath contains curved radial surfaces and a bottom surface, and that the curved radial surfaces of the second bath be located essentially parallel to the curved surfaces of the airfoil in the mating zone, which provides essentially constant blade thickness in the interface zone.

The object of the invention is also a turbine containing at least one blade in accordance with the present invention.

The object of the invention is also a gas turbine engine containing at least one turbine in accordance with the present invention.

The subject of the invention is also a method for manufacturing a gas turbine engine blade, comprising the following steps:

A wax model of the blade is made, containing a feather, at the end of which a heel is made, forming a single part with the feather, with which it is connected at the level of the interface zone, while the heel contains a platform on which at least one sealing plate is made, while in the first bath is performed on the platform, the second bath is performed in the first bath at the level of the conjugation zone between the feather and the heel,

A spatula made of wax is immersed in a refractory slip,

The shell mold is made of refractory material,

Molten metal is poured into the shell mold through a single pouring inlet,

The shell form is broken and a spatula is obtained.

The present invention will be more apparent from the following description of a preferred embodiment of a blade according to the present invention and a method of making the same with reference to the accompanying drawings.

Fig. 1 is a schematic side view of a turbine blade in accordance with the present invention.

Fig. 2 is an isometric view from the front of the outer side of the heel of the scapula.

Fig. 3 is a sectional view of the blade along plane III-III of FIG. one.

Fig. 4 is an isometric side view of the outer side of the heel of the scapula.

As shown in FIG. 1, the blade 1 according to the present invention is formed essentially along the main axis A, which is essentially radial with respect to the axis B of the gas turbine engine containing the blade 1. In this case, we are talking about the turbine blade of a turbojet engine. The blade 1 contains a leg 2, located with inner sides s, platform 3, feather 4 and heel 5, which is located on the outside. The heel 5 mates with the feather 4 in the interface area 15 . Leg 2 is designed to be installed in the rotor socket for mounting on this rotor. The platform 3 is made between the leg 2 and the feather 4 and contains a surface located transversely with respect to the axis A of the blade 1, forming a wall element that limits the gas-air path of its inside; said wall is formed by all platforms 3 of the blades 1 of the turbine stage in question, which are adjacent to each other. Feather 4 is generally located along the main axis A of the blade 1 and has an aerodynamic shape corresponding to its purpose, as is known to those skilled in the art. The heel 5 contains a platform 5a, which is made at the outer end of the airfoil 4 essentially transversely to the main axis A of the blade 1.

As shown in FIG. 2 and 4, the heel platform 5 comprises a leading edge 6 and a trailing edge 7 directed transversely with respect to the gas flow (the flow is generally parallel to the axis B of the turbojet). These two transverse edges, anterior 6 and posterior 7, are connected by two lateral edges 8, 9, which have Z profile: each side edge 8, 9 contains two longitudinal sections (8a, 8b, 9a, 9b, respectively) connected to each other by a section 8", 9", respectively, which is essentially transverse or is made at least at an angle with respect to direction of the gas flow. It is along the side edges 8, 9 that the heel 5 comes into contact with the heels of two adjacent blades on the rotor. In particular, in order to dampen the vibrations to which they are subjected during operation, the blades are mounted on a disk with essentially torsional stress around their main axis A. The heels 5 are designed in such a way that the blades are subjected to torsional stress when supported on adjacent blades along transverse sections 8" , 9" side edges 8, 9.

Starting from the outer surface of the platform 5a of the heel 5, radial plates 10, 11 or scallops 10, 11 are made, in this case in the amount of two; it is also possible to provide only one plate or more than two plates. Each plate 10, 11 is made transversely to the axis B of the gas turbine engine, starting from the outer surface of the platform of the heel 5, between two opposite longitudinal sections (8a, 8b, 9a, 9b) of the side edges 8, 9 of the heel 5.

The platform 5a of the heel 5 is generally formed at a radial angle with respect to the axis B of the gas turbine engine. Indeed, in the turbine, the cross section of the gas-air path increases from inlet to outlet in order to ensure the expansion of gases; thus, the platform 5a of the heel 5 moves away from the axis B of the gas turbine engine from the inlet to the outlet, while its inner surface forms the outer boundary of the gas-air path.

In the platform 5a of the heel 5, a first bath 12 is formed (due to the configuration of the mold). This first bath 12 is a cavity formed by peripheral surfaces 13 forming a rim, which are made starting from the outer surface of the platform 5a and are connected to the surface 14, forming the bottom 14 of the bath 12. The peripheral surfaces 13 are arranged essentially radially and in this case are curved on the inside, forming a mate between the outer surface of the platform 5a and the surface of the bottom 14 of the bath 12. These curved radial surfaces 15 are generally parallel to the side edges 8, 9 and the transverse edges 6, 7 platforms 5a of the heel 5, following their shape when viewed from above (along the main axis A of the blade 1). Some zones of the heel 5 may not contain such radial surfaces 13, in which case the surface of the bottom 14 of the bath 12 goes directly to the side edge (see edge 9a in Fig. 2) (it should be noted that in Fig. 4 these zones are not in the same place).

A bath 12 of this type has already been used in known spatulas. Its function is to lighten the heel 5 while maintaining its mechanical properties: the thickness of the platform 5a of the heel 5 is significant near the side edges 8, 9, side surfaces which, being in contact with adjacent blades, are subjected to strong stresses during the rotation of the blade 1, while the central part of the platform 5a of the heel 5, which is subjected to less stress, is made with a recess forming the first bath 12.

In addition, the heel contains a tray 16 in the first tray 12, hereinafter referred to as the second tray 16. The second tray 16 is made at the level of the interface zone 15 between the heel 5 and the feather 4. In particular, the second tray is made along the main axis A of the blade 1 opposite the zone 15 pairing between heel 5 and feather 4.

The second bath 16 is a cavity formed by peripheral surfaces 17, forming a side, which connect the surface of the bottom 14 of the first bath 12 with the surface 18, which forms the bottom of the second bath 16 (and located on the inner side with respect to the bottom surface 14 of the first bath 12). The peripheral surfaces 17 are arranged substantially radially, in this case being curved on the outer and inner sides, forming a mate between the bottom surface 14 of the first tub 14 and the bottom surface 18 of the second tub 16. These curved radial surfaces 17 are essentially parallel to the surfaces of the feather 4, following their shape when viewed from above (along the main axis A of the blade 1) (see Fig. 4).

The second tub 16 is made during injection molding (in other words, the configuration of the shell mold allowing the blade 1 to be molded is adapted for molding such a tub 16). The blade is made by casting on lost wax models, as described above in the description.

The presence of the second bath 16 avoids excessive thickness in the zone 15 of the interface between the heel 5 and the feather 4. Due to this, during the pouring of the metal into the shell mold, the metal is distributed more evenly, which makes it possible to avoid the formation of porosity, even if the metal is poured only at one pouring point.

Thus, the blade 1 can be made by an investment casting method with a single liquid metal pouring inlet for each shell mold, and such a method is simpler and cheaper. If the forms are combined into blocks, the method is even simpler. In addition, by pouring into the shell mold through a single pouring inlet, the manufactured blade contains only one residual build-up, which is removed by machining. The machining of such a part is simpler.

In addition, the weight and, consequently, the cost of the blade 1 is reduced due to the presence of the second tray 16, while the stresses on the heel 5, as well as the stresses on the feather 4, are better distributed and, therefore, better perceived by the blade 1.

In this case, the pen 4 is made in the form of a solid wall, that is, without cooling with the help of a jacket or a cavity made in the thickness of its wall. Preferably, the peripheral surfaces 17 and the bottom surface 18 of the second tub 16 are designed in such a way that the thickness of the paddle 1 is substantially constant in the interface 15 between the heel 5 and the feather 4. This hallmark clearly visible in Fig. 3. In particular, if we designate 15a, 15b the curved surfaces of the feather 4 at the level of the interface zone 15 between the feather 4 and the heel 5, then in FIG. 3 it can be seen that the curved radial surfaces 17 of the second bath 16 are substantially parallel to the curved surfaces 15a, 15b of the feather 4, against which they are located. In the illustrated embodiment, the radius of the curved radial surfaces 17 of the second bath 16 is not identical to the radius of the opposite curved surfaces 15a, 15b of the feather 4, but nevertheless these surfaces are substantially parallel.

Part of the second bath 16, located in FIG. 3 on the left, is characterized by a continuous curved shape without any flat area between the curved radial surface 13 of the first tray 12, the bottom 14 of the first tray 12 and the curved radial surface 17 of the second tray 16. However, on the part of the second tray 16, located in FIG. 3 on the right, each of these areas is clearly visible. Execution in between different sites in the area under consideration (in section) depends on the position of the surfaces of the heel 5 in relation to the surfaces of the feather 4.

The invention is described for a movable turbine blade. At the same time, in fact, it can be applied to any blade made by casting and containing a feather, at the end of which a heel is made in the form of a single piece with a feather.

CLAIM

1. The blade of a gas turbine engine, made by casting, containing a feather, at the end of which there is a heel, made in the form of a single piece with a feather, with which it is connected at the level of the interface zone, while the heel contains a platform on which at least one a sealing plate, and the first bath is made in the platform, characterized in that the second bath is made in the first bath at the level of the interface zone between the feather and the heel.

2. A spatula according to claim 1, wherein the first bath is defined by radial surfaces and a bottom, and the second bath is formed in the bottom of the first bath.

3. The blade according to claim 1, in which the second tray is made along the main axis (A) of the blade opposite the interface zone between the heel and the feather.

4. The blade according to claim 3, in which the pen is formed by a solid wall and contains curved surfaces in the interface zone, and the second tray contains curved radial surfaces and a bottom surface, while the curved radial surfaces of the second tray are located essentially parallel to the curved surfaces of the pen in interface zone, which provides a substantially constant blade thickness in the interface zone.

5. Turbine containing at least one blade according to claim 1.

6. Gas turbine engine containing at least one turbine according to claim 5.

PJSC Ufa Motor-Building Production Association (UMPO) launched the largest blade casting melting and pouring plant in Europe at the advanced blade casting site. The dimensions of the equipment are 9 meters wide, 12 meters long and 8.5 meters high. The unit is designed for the manufacture of blanks during the production of engine parts for the promising civil aircraft MS-21. The new equipment makes it possible to melt from 20 to 150 kg of a special alloy, which makes it possible to pour a large number blades in just one cycle.

The new PZU will be actively involved in the implementation of a joint project between UMPO and the Moscow Institute of Steel and Alloys (NUST MISiS) to develop and implement a resource-efficient technology for manufacturing hollow cast turbine blades. It will be used in the production of not only aviation gas turbine engines, but also oil and gas pumping stations,” said the curator of the promising program, deputy head of the department for technical development and re-equipment Pavel Alinkin.

At the beginning of November 2015, this project won a subsidy in the competition of the Ministry of Education of the Russian Federation under Decree No. 218 of the Government of the Russian Federation. The grant will help UMPO reduce the time it takes to introduce innovation into pilot and mass production.

The Association has a rich experience of cooperation with Russian universities under the 218th Decree. Currently, the enterprise is working on two more technologies: for the production of thin-walled large-sized titanium castings (with MISiS and USATU) and heat-resistant aluminum parts (with USATU and other universities). Two projects - also with MISiS and USATU - have been successfully completed, their results are currently being introduced into production. This is a manufacturing technology for the turbine support of the VK-2500 helicopter engine and the production of monowheels and blisks by linear friction welding.

For the first time in Russia, it was possible to cast (a method called investment casting) from an alloy of titanium aluminide innovative blades that are twice as light as their nickel-based counterparts. The technology for manufacturing new blades has already been put into production at the Ufa Motor-Building Production Association (PJSC UMPO). It is expected that titanium intermetallic blades will be used in the new Russian PD-14 engine for Russian short-medium haul passenger aircraft MS-21. By reducing the weight of the aircraft, the new development will allow you to carry more passengers with less fuel.

“Today, the manufacture of products from titanium aluminide is in great demand in civil aviation. Our development is not inferior to world analogues from Europe and the USA. It is very important that this is a completely domestic development: blades can be produced on domestic equipment and from domestic materials,” said in an interview the head of the research group, head of the Department of Technology of Foundry Processes and Artistic Processing of Materials, NUST MISIS, Professor Vladimir Belov. The transition to a new technology will significantly reduce the weight of the engine, as a result, it will be possible to transport more passengers or cargo over long distances. Besides, new technology manufacturing blades will significantly reduce the effective centrifugal stress in the compressor and turbines of aircraft engines, reduce the inertia of turbines and compressors, and thereby reduce fuel consumption and greenhouse gas emissions.

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;

equipping the technological process of manufacturing 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. AT 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 for various stages mechanical processing.

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.