The technology for manufacturing new generation blades has been put into production at umpo. From single-crystal uncooled blades to turbine blades with penetrating (transpiration) cooling, manufactured using additive technologies (review on the technology of

The blades of gas turbine engines (GTE) are the most massive parts in the production of these power plants.

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

Perfection technological processes manufacturing blades of gas turbine engines (GTE) should primarily solve the problem of increasing economic indicators process, namely: increasing the utilization rate of the material; reducing the complexity of manufacturing; reduction of the technological cycle for the manufacture of parts and the cost of technological preparation of production.

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

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

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

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

On the basis of uniform signs of similarity and difference of parts, the following groups with characteristic features can be formed: turbine rotor blades; nozzle blades; compressor blades; rings; disks; shafts; deflectors; supports, etc. Thus, a group of parts is given - GTE compressor blades, which should be manufactured within the framework of one standard technology.

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

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

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

However, existing industries already include big number technologies created in the previous period, which should also be combined within the group technology, while retaining their existing classification system for parts, technological processes, tooling, etc.

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

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

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. An additional task 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. 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 spruce 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 spruce 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.

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. IN 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. As construction materials for compressor blades of modern gas turbine engines, titanium alloys are used (VT22, VT3-1, VT6, VT8, VT33), heat-resistant steels (EN961 Sh, EP517Sh), 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 fatigue 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 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 airfoil, called the heel, which forms an element of the outer wall that limits the gas-air path.

On the outer surface of the heel, one or more radial plates or scallops are made, forming, together with the stator wall opposite, 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 to produce 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 molded metal blade, which must be machined after the molding of the blade. As a rule, metal is poured at the level of the heel of the blade. The diameter of the pouring channel and, consequently, 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 casting point is provided, there is a poor distribution of the metal 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 two smaller diameter casting channels are obtained, spaced apart, which provides 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 avoids 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 process for 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 - front isometric view outer side blade heels.

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 a major 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 shoulder blade 1 contains a leg 2 located on the inside, a platform 3, a feather 4 and a 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, front 6 and rear 7, are connected by two side edges 8, 9, which have a Z-shaped profile: each side edge 8, 9 contains two longitudinal sections (8a, 8b, 9a, 9b respectively) connected to each other section 8", 9", respectively, which is essentially transverse or made at least at an angle with respect to the 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 disc with substantially 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 keeping it mechanical properties: the thickness of the platform 5a of the heel 5 is significant near the side edges 8, 9, the side surfaces of which, in contact with the 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 bath 16 in the first bath 12, hereinafter referred to as the second bath 16. The second bath 16 is made at the level of the interface zone 15 between the heel 5 and the feather 4. In particular, the second bath 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 bottom surface 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 conjugation 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. The execution between them of different sections 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 mating 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.

The relevance of the work

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

Figure 1 - GTE compressor blade model (10 frames, 20 cycles)

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

Purpose and tasks of the work

Objective- increasing the durability and quality of the GTE compressor blades by improving the structural and technological support for the manufacturing processes of the GTE compressor blades.

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

Scientific novelty of the work

The scientific novelty of the work lies in the development of recommendations for improving the structural and technological support for the manufacturing processes of GTE compressor blades and creating an optimal structure for the technological process of processing GTE compressor blades. Also, this work provides a solution to the problem of durability and wear resistance of the GTE compressor blades.

Main part

Compressor blades of a gas turbine engine

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

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

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

Formation of the surface layer at the finishing operations for the manufacture of blades

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

Vibroabrasive processing of blades

Blades are characteristic mass parts of aircraft gas turbine engines, they operate under conditions of high static, dynamic and thermal loads and largely determine the life and reliability of the engine as a whole.
For their manufacture, high-strength titanium alloys, stainless steels, nickel-based heat-resistant alloys, as well as composite materials are used.
The complexity of manufacturing blades in most designs of gas turbine engines is 30-40% of the total complexity of the engine. This feature, along with the operating conditions of the blade in the engine, requires the use of advanced methods for obtaining blanks in the production, modern technologies processing, especially at finishing operations, mechanization and automation of technological processes.
In the operation of aircraft gas turbine engines, out of all failures due to the reasons for the strength failure of parts, blades account for about 60%. The vast majority of blade failures are of a fatigue nature. This is often facilitated by damage to the blades caused by solid particles entering the engine tract (stones when taxiing on the ground, birds in flight, etc.). This causes the need to have a sufficiently high margin of cyclic strength of the blades, as well as to take special technological and design measures to increase their survivability in the event of damage (dents).
Depending on the operating conditions in the engine, the level of alternating stresses in the blades is usually in the range of 40-160 MPa, and taking into account the necessary safety margin, their endurance limit is usually required in the range of 300-500 MPa. The fatigue resistance of a blade depends on the material, the design of the blade, and the technology of its manufacture, but in any case, the state of the surface layer greatly affects the value of the endurance limit. The main factors affecting the quality of the surface layer are:
- residual stresses - their sign, magnitude, depth, nature of distribution over the section of the part, etc.;
- surface microrelief - the size and nature of microroughnesses, the presence of scratches;
- structure of the surface layer.
The urgency of the task of increasing the fatigue resistance of the blades has led to the development and implementation of special processing methods and the introduction in the industry of a number of special methods for processing their surface.
The place of vibroabrasive processing in the technological process of mechanical processing of blades, as a rule, is the finishing process performed at the final stage of processing. Depending on the material of the blade, the type of previous processing, and the initial value of surface microroughness and some other factors, processing modes are assigned - the frequency and magnitude of the oscillation amplitude, the characteristics of the working bodies (abrasive breakage, molded vibrating bodies, ceramic, glass or metal balls, wooden cubes, etc. .), mass ratios, etc. This makes it possible to achieve the desired result in a fairly wide range of initial surface states. So, for compressor blades of small and medium dimensions made of steel and titanium alloys, the final shaping operation is cold rolling followed by rounding the edges with an abrasive wheel. In this case, the surface roughness is Ra = 1.6 and higher, therefore, “soft” vibration treatment modes are used to level out microroughnesses on the surface and create compressive stresses in the surface layer. In this case, bulk processing is used (without fixing parts) in toroidal vibrators. In some cases, the processing technology provides for abrasive grinding in the final operations, followed by polishing the surface of the blade airfoil. Such blades are subjected to more intense vibroabrasive treatment to remove microroughness and provide residual compressive stresses in the surface layer.
It is much more difficult to implement effective vibration processing of large blades of turbomachines. A large mass of such parts, taking into account the weight of the container and working environment make it problematic to create a vibratory machine with an acceptable frequency and amplitude of oscillations in two or three coordinates due to a sharp increase in the required drive power and dynamic overloads of machine elements. Moreover, these details are worst quality the original surface, which reduces the productivity of processing.
At the Motor Sich enterprise, the method of longitudinal single-coordinate vibration treatment in a closed container (POVO) is used.
In traditional domestic and foreign vibroabrasive machines, loose filler is driven from oscillatory movements the bottom of the container, which is always at the bottom. In this case, the filler is returned back free fall. The effectiveness of this method is not high enough.
The process of vibroabrasive machining of parts is significantly activated and intensified inside a closed container with two bottoms located opposite each other, if the bulk filler actively oscillates between them, receiving kinetic energy from each bottom. The intensity of impacts of the filler with the workpiece increases significantly. The side walls of the container are inclined (conical), which creates an additional compression of the filler during its movement, which increases the forces of dynamic action between the abrasive filler and the walls of the container, inside which the machined parts of the gas turbine engine are located in a fixed or free state.
When vibrating by this method with abrasive granules and hardened steel balls, metal removal from the surface and surface microdeformation of parts are more intense than in traditional vibrators, which increases the magnitude and depth of surface compressive stresses and increases the fatigue resistance of parts.
Figure 3 shows the curves of changes in the surface roughness of blades made of steel 14Kh17N2Sh on the duration of treatment on a vibratory unit with a U-shaped container.

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

Achieving the roughness Ra=1.5 µm by the POOH method, as follows from Fig. 3, occurs in about 30 minutes, and by conventional vibroabrasive processing - 1.5 hours.
The study of vibroabrasive processing of turbine and compressor blades shows the advantages of this process compared to manual polishing and polishing. The results of the study showed that the endurance limit of the blades subjected to vibrogrinding and vibropolishing is 410 MPa and meets the requirements of TS. The magnitude and nature of the residual stresses of the investigated blades are more favorable than on the blades with manual polishing and glossing.

Conclusion

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

Literature

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JSC "Naro-Fominsk Machine-Building Plant" GTE compressor blades
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Doctor of technical sciences Yury Eliseev, General Director of FSPC MMPP "Salyut", Advanced technologies for the production of GTE blades

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