PJSC Ufa Engine Production Association (UMPO) launched the largest melting and casting installation for blade casting 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 installation is intended for the production of blanks during the production of engine parts for the promising MS-21 civil aircraft. The new equipment makes it possible to melt from 20 to 150 kg of a special alloy, which makes it possible to fill a large number of blades in just one cycle.
The new ROM 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 resource-efficient technology for the manufacture of hollow cast turbine blades. It will be used in the production of not only aircraft gas turbine engines, but also oil and gas pumping stations,” said Pavel Alinkin, curator of the promising program, deputy head of the technical development and re-equipment department.
At the beginning of November 2015, this project won a subsidy in a competition of the Ministry of Education of the Russian Federation under Resolution No. 218 of the Government of the Russian Federation. The grant will help UMPO reduce the time it takes to introduce innovations into pilot and mass production.
The association has extensive experience in cooperation with Russian universities under Resolution 218. Currently, the company is working on two more technologies: for the production of thin-walled large-sized titanium castings (with MISiS and USATU) and parts made of heat-resistant aluminum (with USATU and other universities). Two projects - also with MISiS and UGATU - have been successfully completed, their results are currently being put into production. This is the technology for manufacturing the turbine support of the VK-2500 helicopter engine and the production of unicycles and blisks using linear friction welding.
For the first time in Russia, it was possible to cast (the method is called investment casting) innovative blades from titanium aluminide alloy, which are twice as light as their nickel-based counterparts. The technology for manufacturing new blades has already been put into production at the Ufa Engine Production Association (UMPO PJSC). It is expected that titanium intermetallic blades will be used in the new Russian PD-14 engine for the Russian short-medium-haul passenger aircraft MS-21. By reducing the weight of the aircraft, the new development will allow it to carry more passengers with less fuel consumption.
“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: the blades can be produced using domestic equipment and from domestic materials,” said the head of the research group, head of the department of “Technology of Foundry Processes and Artistic Processing of Materials” at NUST MISIS, Professor Vladimir Belov in an interview. The transition to 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 of 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 into the atmosphere.
The invention relates to foundry production. The blade of a gas turbine engine is made by investment casting. The blade contains a feather 4, at the end of which there is a heel 5, made in the form of a single part with the feather. The heel contains a platform 5a, in which a first basin 12 with radial surfaces 13 and a bottom 14 is made. The basin 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 the metal to be poured 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 salary f-ly, 4 ill.
Drawings for RF patent 2477196
The present invention relates to a cast metal blade and a method for producing 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, thanks to the presence of radial blades mounted 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 the direction of gas flow in that engine.
The movable turbine blade contains a leg with which it is attached to the rotor disk, a platform forming an element of the internal 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 distant from the leg, the blade ends with an element transverse to the main (main) axis of the airfoil, called a heel, which forms an element of the outer wall delimiting the gas-air path.
On the outer surface of the heel there are one or more radial plates or scallops, which together with the stator wall located opposite them form a labyrinth gasket that ensures tightness against gases; For this purpose, as a rule, the mentioned 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 part. The blade is made by a casting method called “lost wax casting” and is well known to specialists. In this method:
First, a model of a scapula is made from wax;
The model is immersed in a refractory ceramic slip, which, after firing, forms a shell;
The wax is melted and removed, resulting in a "shell mold" of refractory material, the internal volume of which determines the shape of the blade;
Molten metal is poured into a shell mold, and several shell molds are combined into a block to simultaneously pour 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, metal build-ups of relatively large thickness are formed on the metal blade being cast in the mold, which must be machined after the blade is formed. As a rule, metal is poured at the level of the heel of the blade. The diameter of the filling channel and, consequently, the subsequently formed build-up is significant, and the filling occurs near the plates of the labyrinth gasket, which have a small thickness; As a result, if only one pouring point is provided, poor distribution of the metal in the shell mold occurs and porosity problems arise in the blade, particularly at the level of the blade blades.
This problem can be solved by providing two pouring inlets, and the diameter of the pouring channels is correspondingly reduced. Thus, instead of one large-diameter pouring channel, two smaller-diameter pouring channels are obtained, spaced apart from each other, which provides better metal distribution and avoids porosity problems.
However, it is desirable to solve 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 part with a feather, with which it is connected at the level of the interface zone, while the heel contains a platform on which the at least one sealing plate, and a first bath is made in the platform, characterized in that a second bath is made in the first bath at the level of the interface zone between the feather and the heel.
The presence of one bath in another bath at the level of the interface between the blade and the heel avoids too much thickening in this zone and during the molding of the blade by casting ensures better distribution of the liquid metal in the mold. Improved distribution of liquid metal in the mold allows the use of casting molding with a single point of metal pouring. The advantage of manufacturing a blade with one pouring point is the exceptional simplicity of the shell mold and, if necessary, the block of shell molds; the cost of manufacturing blades decreases, while their quality increases.
In addition, the amount of material at the heel level is optimized, which reduces the weight and cost of the paddle.
In addition, we optimize mechanical stress onto the heel and/or feather and are better absorbed by the paddle as 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 bath be made along the main axis of the blade opposite the interface between the heel and the feather.
It is advisable that the blade feather be formed by a solid wall and contain curved surfaces in the mating zone, the second bath contains curved radial surfaces and the bottom surface, and that the curved radial surfaces of the second bath are located essentially parallel to the curved surfaces of the feather in the mating zone, which ensures essentially constant thickness of the blade in the interface zone.
The invention also relates to a turbine comprising at least one blade according to the present invention.
The invention also relates to a gas turbine engine comprising at least one turbine in accordance with the present invention.
The invention also relates to 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, in the first bath at the level of the interface zone between the feather and the heel, the second bath is performed,
A wax spatula is immersed in a refractory slip,
The shell mold is made of fire-resistant material,
Molten metal is poured into the shell mold through a single pouring inlet,
The shell mold is broken and a blade is obtained.
The present invention will be more evident from the following description of a preferred embodiment of a blade in accordance with the present invention and a method for its manufacture 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 - isometric front view outside heels of the scapula.
Fig. 3 is a sectional view of the blade along plane III-III of FIG. 1.
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 generally along a major axis A, which is substantially radial with respect to the axis B of the gas turbine engine comprising the blade 1. B in this case we're talking about about the turbine blade of a turbojet engine. The 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 mating zone 15. Leg 2 is designed to be installed in the rotor socket for mounting on this rotor. 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 limiting the gas-air path of its inside; said wall is formed by all platforms 3 of blades 1 of the turbine stage under consideration, which are adjacent to each other. The feather 4 is mainly located along the main axis A of the blade 1 and has an aerodynamic shape corresponding to its purpose, as is known to specialists. The heel 5 contains a platform 5a, which is made at the outer end of the feather 4 essentially transverse to the main axis A of the blade 1.
As shown in FIG. 2 and 4, the heel platform 5 includes 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 engine). 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 a section 8", 9" respectively, which is substantially transverse or at least angled 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, to dampen the vibrations to which they are subjected during operation, the blades are mounted on the disk with a substantially torsional stress around their main axis A. The heels 5 are designed in such a way that the blades are subjected to a torsional stress when supported by 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 blade or more than two blades. Each plate 10, 11 is made transverse to the axis B of the gas turbine engine, starting from the outer surface of the heel pad 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 designed at a radial angle with respect to the axis B of the gas turbine engine. Indeed, in a 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 inlet to outlet, while its inner surface forms the outer boundary of the gas-air path.
In the platform 5a of the heels 5, a first bath 12 is formed (due to the configuration of the casting 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 substantially radially and in this case are curved on the inside, forming an interface 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 of the platform 5a of the heel 5, following their shape when viewed from above (along the main axis A of the blade 1). Some areas of the heel 5 may not contain such radial surfaces 13, in which case the surface of the bottom 14 of the bath 12 extends directly to the side edge (see edge 9a in FIG. 2) (it should be noted that in FIG. 4 these areas are not in same place).
A tray 12 of this type has already been used in known blades. Its function is to relieve the heel 5 while maintaining it mechanical properties: the thickness of the pad 5a of the heel 5 is significant near the lateral edges 8, 9, the lateral surfaces of which, being in contact with adjacent blades, are subject to strong stresses during the rotation of the blade 1, while the central part of the pad 5a of the heel 5, which is subject 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 called 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 interface between heel 5 and feather 4.
The second bath 16 is a cavity formed by peripheral surfaces 17 forming a rim that connect the surface of the bottom 14 of the first bath 12 with the surface 18 forming the bottom of the second bath 16 (and located on the inner side with respect to the surface of the bottom 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 interface between the surface of the bottom 14 of the first bath 14 and the surface of the bottom 18 of the second bath 16. These curved radial surfaces 17 are substantially parallel to the surfaces of the pen 4, following their shape when viewed from above (along the main axis A of the blade 1) (see Fig. 4).
The second bath 16 is performed during casting molding (in other words, the configuration of the shell mold that allows the blade 1 to be molded is adapted for molding such a bath 16). The blade is made by casting using lost wax models, as indicated above in the description.
The presence of the second bath 16 avoids excessive thickness in the interface zone 15 between the heel 5 and the feather 4. Due to this, during the pouring of metal into the shell mold, the metal is distributed more evenly, which avoids the formation of porosity, even if the metal is poured only at one pouring point.
Thus, the blade 1 can be produced by an investment casting method with a single liquid metal filling inlet for each shell mold, and this 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. Mechanical processing of such a part is simpler.
In addition, the mass and therefore the cost of the blade 1 is reduced by the presence of a second tray 16, while the stresses on the heel 5, as well as the stresses on the blade 4, are better distributed and therefore better absorbed by the blade 1.
In this case, the pen 4 is made in the form of a solid wall, that is, without cooling using a jacket or cavity made in the thickness of its wall. It is preferable that the peripheral surfaces 17 and the bottom surface 18 of the second bath 16 are designed in such a way that the thickness of the blade 1 is substantially constant in the interface area 15 between the heel 5 and the blade 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 tray 16 are made substantially parallel to the curved surfaces 15a, 15b of the feather 4, opposite which they are located. In the illustrated embodiment, the radius of the curved radial surfaces 17 of the second tub 16 is not identical to the radius of the opposing curved surfaces 15a, 15b of the feather 4, but nevertheless these surfaces are essentially parallel.
The part of the second bath 16, located in FIG. 3 on the left is characterized by a continuous curved shape without any flat portion between the curved radial surface 13 of the first bath 12, the bottom 14 of the first bath 12 and the curved radial surface 17 of the second bath 16. However, on the part of the second bath 16 located in FIG. 3 on the right, each of these areas is clearly visible. The execution of different sections between them in the zone 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 used for any blade made by casting and containing a feather, at the end of which a heel is made in the form of a single part with a feather.
CLAIM
1. A 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 part 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 a first bath is made in the platform, characterized in that a second bath is made in the first bath at the level of the interface zone between the feather and the heel.
2. The blade according to claim 1, in which the first bath is limited by radial surfaces and the bottom, and the second bath is made in the bottom of the first bath.
3. The blade according to claim 1, in which the second bath 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 feather is formed by a solid wall and contains curved surfaces in the mating zone, and the second bath contains curved radial surfaces and a bottom surface, while the curved radial surfaces of the second bath are located essentially parallel to the curved surfaces of the feather in the interface zone, which ensures a substantially constant thickness of the blade in the interface zone.
5. A turbine containing at least one blade according to claim 1.
6. Gas turbine engine containing at least one turbine according to claim 5.
1The paper discusses methods for manufacturing high-pressure compressor blades for gas turbine engines. The first method is to process the blade airfoil profile by milling on coordinate machines with numerical control followed by manual modification. The second method is electrochemical processing, which eliminates mechanical and manual processing of the blade feathers. The problems of manufacturing compressor blades using the milling method have been studied. Current problems are presented, the solution of which will improve accuracy, quality and eliminate manual grinding and polishing work. The advantages of electrochemical processing are given. The costs and labor intensity for production preparation, the costs and labor intensity for the manufacture of blades are presented and analyzed. The work also presents the results of measurements of compressor blades. The best results in terms of accuracy and stability of the feather profile geometry were obtained as a result of electrochemical processing.
electrochemical processing
milling
gas turbine engine
1. Galiev V.E., Fatkullina D.Z. Promising technological process for manufacturing precision compressor blades [Text] / V.E. Galiev, D.Z. Fatkullina // Bulletin of UGATU. – 2014. – No. 3. – P. 9–105.
2. Nekhorosheev M.V. Using volumetric and plane modeling of a two-electrode electrochemical cell in the ANSYS program [Text] / M.V. Nekhorosheev, N.D. Pronichev, G.V. Smirnov // Bulletin of Samara University. Aerospace engineering, technology and mechanical engineering. – 2012. – No. 3–3. – pp. 98–102.
3. Lunev A.N. Optimization of milling parameters GTE blades on CNC machines [Text] / A.N. Lunev, L.T. Moiseeva, M.V. Solomina // News of the Higher educational institutions. Aviation technology. – 2007. – No. 2. – P. 52–55.
4. Nekhorosheev M.V. Automation of the design of technology for electrochemical processing of the blades of gas turbine engines based on computer modeling of the shaping process [Text] / M.V. Nekhorosheev., N.D. Pronichev., G.V. Smirnov // News of the Samara Scientific Center of the Russian Academy of Sciences. – 2013. – T. 15, No. 4–6. – pp. 897–900.
5. Pavlinich S.P. Prospects for the use of pulsed electrochemical processing in the production of gas turbine engine parts [Text] / S.P. Pavlinich // Bulletin of UGATU. – 2008. – No. 2. – P. 105–115.
6. Production of gas turbine engines [Text]: reference manual / A.M. Abramov, I.L. Zelikov, M.F. Idzon et al. - M.: Publishing house "MECHANICAL ENGINEERING", 1996. - 472 p.
7. Development of a strategy for creating innovative technological processes [Text]: Textbook / N.D. Pronichev, A.P. Shulepov, L.A. Chempinsky, A.V. Meshcheryakov. – Samara: Samara State Aerospace University, 2011. – 166 p.
8. Technology of production of aviation gas turbine engines [Text]: Textbook for universities / Yu.S. Eliseev, A.G. Boytsov, V.V. Krymov, L.A. Khvorostukhin. – M.: Mashinostroenie, 2003. – 512 p.
9. Tolkachev A.V. Increasing the productivity of vibration polishing of gas turbine engine compressor blades with abrasive granules: diss... cand. those. Sci. – Rybinsk, 2015. – 136 p.
10. Turanov A.V. To the method of calculating the modes of milling the surfaces of gas turbine engine blades on CNC machines [Text]/A.V. Turanov, L.T. Moiseeva, A.N. Lunev // News of higher educational institutions. Aviation technology. – 2005. – No. 2. – P. 60–64.
Compressor blades are critical and massive parts of a gas turbine engine. The service life and final cost of the engine will depend on the correctly chosen blade manufacturing technology.
Ensuring a given blade service life largely depends on a number of technological factors. The condition of the surface layer of the blades, the presence of traces of previous processing (surface roughness), which are stress concentrators, have a significant impact on the long-term and fatigue strength of the blades during operation.
Therefore, the manufacture of blades, even in small-scale production, requires the use of modern technological processes, high-performance equipment and automation of the manufacturing and control process.
One of the widely used technologies for manufacturing compressor blades of a gas turbine engine is milling on coordinate machines with subsequent manual refinement, in particular finishing operations. However, this technology has a number of disadvantages:
Low accuracy and performance;
The need to use manual operations;
Highly qualified worker in final manual operations for finishing the profile of the blades;
Harmful conditions for workers when performing manual grinding and polishing work;
High cost and rapid wear of cutting tools;
100% control required.
Current tasks in the manufacture of gas turbine engine compressor blades are:
Automation of finishing operations for processing the pen profile. Elimination of manual operations will improve the quality and stability of the technological process for manufacturing gas turbine engine blades;
The use of physical and chemical processing methods will eliminate the use of expensive cutting tools and increase processing productivity;
Automation of inspection of gas turbine engine blades.
One of the most effective and promising areas for blade manufacturing is electrochemical processing. The advantages of electrochemical processing are:
Reduced production time for blades and the ability to efficient processing difficult-to-process materials;
The surface quality after electrochemical treatment requires minimal post-finishing;
High tool life;
In addition, it is noted that blades after ECM have increased gas-dynamic stability, a reduced spread of natural vibration frequencies, and increased fatigue strength due to a decrease in residual stresses.
It is known that foreign manufacturers of gas turbine engines (such as General Electric Company, MTU Aero Engines GmbH, Volvo Aero Corporation, etc.) successfully use ECM both as an operation for preliminary shaping of the inter-blade channel of monowheels using non-profiled electrodes, and for dimensional processing of the blade airfoil with profiled ones. electrodes and instruments.
Work has begun in this area and significant progress has been achieved at the NIID (Moscow), Kazan (KAI, KSTU), Samara (SAI) and Ufa (Research Institute of Petrology and Technology ECHO at UGATU) schools of electrochemical processing, etc.
For analysis, two methods were chosen for manufacturing high-pressure compressor blades of a gas turbine engine.
First way. Manufacturing of blades on coordinate milling machines, Fig. 1. A milled parallelepiped, manufactured with an accuracy of 0.1 mm, is used as the initial workpiece. The dovetail lock is formed on a horizontal broaching machine. Next, complex milling of all elements of the flow part of the blade is carried out on coordinate numerically controlled machines with an allowance for finishing. In the complex milling process, the workpiece is supported by a dovetail shank. The final stage of blade manufacturing is manual processing or endless belt processing.
Second way. Manufacturing of blades on electrochemical machines, Fig. 2. A polished parallelepiped, manufactured with an accuracy of 0.02 mm, is used as the initial workpiece. In the process of electrochemical processing, tract surfaces are formed with an allowance for finishing. Next, the dovetail shank is formed on a horizontal broaching machine. The final operation is carried out on a vibration grinding machine.
Let's analyze both methods of manufacturing compressor blades. The most complete picture can be obtained by comparing the costs and labor intensity of production preparation, the costs and labor intensity of manufacturing the part, as well as the accuracy and stability of blade manufacturing. For analysis, two batches of blades were manufactured using the above methods.
Rice. 1. Main stages of manufacturing compressor blades
Rice. 2. Main stages of manufacturing compressor blades
Table 1
Basic costs for production preparation
|
Planned labor intensity n.h. |
Cost 1 piece. rub. |
Incl. material costs |
|||
|
manufacturing |
regrinding |
manufacturing |
regrinding |
||
|
Milling |
|||||
|
Milling cutter No. 1 |
|||||
|
Milling cutter No. 2 |
|||||
|
Milling cutter No. 3 |
|||||
|
Milling cutter No. 4 |
|||||
|
Milling cutter No. 5 |
|||||
|
Milling cutter No. 6 |
|||||
|
Milling cutter No. 7 |
|||||
|
Device |
|||||
|
Electrochemical processing |
|||||
|
Electrode No. 1 |
|||||
|
Electrode No. 2 |
|||||
|
Device |
|||||
Rice. 3. Cost of manufacturing technological equipment
Rice. 4. Labor intensity of manufacturing technological equipment
In the process of designing a technological process, significant factors are time and costs for production preparation (Table 1). In table 1 included the main costs for the manufacture of equipment for milling (first method) and electrochemical processing (second method) of cutting tools and tool electrodes. When considering the table. 1 it becomes obvious that the costs of materials and labor intensity for production preparation for electrochemical processing are higher than for milling.
The total labor intensity and cost of manufacturing technological equipment are presented in Fig. 3 and 4.
The complexity and cost of the main operations for manufacturing blades are presented in Table. 2. High requirements for the accuracy of manufacturing a workpiece for electrochemical processing lead to the use of additional operation"surface grinding". The time spent on processing a complex of compressor blade surfaces using the electrochemical method is lower than when milling. Also from table. 2 shows that the milling technology requires the use of manual finishing work, which increases the cost of the finished product.
The total labor intensity and cost of manufacturing one blade are presented in Fig. 4 and 5.
table 2
Labor intensity and cost of the main operations of blade manufacturing
|
Labor intensity, n.h. |
Cost, rub. |
||||
|
Milling |
Milling |
||||
|
Milling |
93 rub. 90.3 kop. |
93 rub. 90.30 kopecks |
|||
|
Grinding |
26 rub. 27.50 kopecks |
||||
|
Pulling the lock |
7 rub. 43.10 kop. |
7 rub. 43.10 kop. |
|||
|
Treatment of tract surfaces |
100 rub. 00 kop. |
70 rub. 00 kop. |
|||
|
Manual operation |
40 rub. 30.20 kopecks |
||||
|
Vibratory grinding |
5 rub. 40 kopecks |
||||
Rice. 5. Total complexity of manufacturing one part
Rice. 6. Total cost of manufacturing one part
In Fig. Figure 7 shows a comparative analysis of the costs of manufacturing one part. When calculating costs, we took into account the costs of manufacturing technological equipment with their subsequent regrinding and repair. As can be seen from the figure, increasing the part production program reduces the cost of one part. However, significant costs are incurred for blades manufactured using the milling technology. This phenomenon is explained by rapid wear of the cutting tool.
The virtual absence of wear of the electrodes during electrochemical processing reduces the cost of manufacturing blades.
Accuracy of blade manufacturing and stability of technological processes Fig. 1 and 2 are summarized in Fig. 8.
Measurements of finished blades were carried out on a control measuring machine. Measurements were carried out along the entrance and exit edges in four sections. It follows from the figure that the greatest accuracy and repeatability of obtaining the geometric dimensions of the blade edges is achieved by the method of electrochemical processing. A significant increase in the stability and accuracy of blade manufacturing using electrochemical processing is due to the elimination of manual operations.
Taken together, considering the data obtained, the following conclusions can be drawn.
The use of more complex equipment in the process of electrochemical processing significantly increases the costs and time for production preparation. Thus, milling is a more flexible and quickly adaptable processing method. The costs and labor intensity for preparing the production of milling processing are lower than those of electrochemical processing (Fig. 1 and 2).
The cost of manufacturing blades using milling technology is higher than using electrochemical processing. The increase in cost is due to the fact that manual operations are required after the milling operation.
Rice. 7. Comparative graph of costs for manufacturing one part depending on the number of blades produced
Rice. 8. Precision edge manufacturing
The costs of manufacturing blades using the milling technology are higher than those using electrochemical processing (Fig. 7). A significant cost is the purchase of expensive cutting tools.
The accuracy and stability of electrochemical processing is much higher.
Bibliographic link
Valiev A.I. COMPARATIVE ANALYSIS OF THE MANUFACTURE OF GAS TURBINE ENGINE COMPRESSOR BLADES // Basic Research. – 2017. – No. 5. – P. 36-41;URL: http://fundamental-research.ru/ru/article/view?id=41503 (access date: 03/28/2019). We bring to your attention magazines published by the publishing house "Academy of Natural Sciences"
Probably everyone knows that no matter how hard the Chinese try, they cannot copy modern jet engines. All. they copied what they could and got their own SUSHKA, but the engine still has to be bought in the Russian Federation. I just read an article on ViMe: http://www.warandpeace.ru/ru/news/view/74298/ “China still cannot copy a modern jet engine.” Moreover, I understand that there are ultra-modern technologies, developments, mathematics, etc., etc., etc.... But in order to understand in more detail what is actually going on here, I recommend reading the following article.
ENGINES AND MATERIALS
The power of any heat engine is determined by the temperature of the working fluid - in the case of a jet engine, this is the temperature of the gas flowing from the combustion chambers. The higher the gas temperature, the more powerful the engine, the greater its thrust, the higher the efficiency and the better the weight characteristics. A gas turbine engine contains an air compressor. It is driven into rotation by a gas turbine sitting on the same shaft. The compressor compresses atmospheric air up to 6-7 atmospheres and directs it into the combustion chambers, where fuel - kerosene - is injected. The flow of hot gas flowing from the chambers - products of combustion of kerosene - rotates the turbine and, flying out through the nozzle, creates jet thrust and propels the aircraft. The high temperatures arising in the combustion chambers required the creation of new technologies and the use of new materials for the construction of one of the most critical engine elements - the stator and rotor blades of the gas turbine. They must withstand enormous temperatures for many hours, without losing mechanical strength, at which many steels and alloys already melt. First of all, this applies to turbine blades - they perceive a 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 capable of withstanding high mechanical loads for a long time did not exist in our country.
Soon after the end of the Great Patriotic War, a special laboratory at VIAM began work on creating alloys for the manufacture of turbine blades. It was headed by Sergei Timofeevich Kishkin.
TO ENGLAND FOR METAL
The first domestic design of a turbojet engine was created in Leningrad by aircraft engine designer Arkhip Mikhailovich Lyulka even before the war. At the end of the 1930s, he was repressed, but, probably anticipating his arrest, he managed to bury the engine drawings in the courtyard 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 first production combat aircraft was the twin-engine Messerschmitt Me-262 Then Stalin summoned L.P. Beria, who oversaw new military developments, and demanded to find those who are working on jet engines in our country. A.M. Lyulka was quickly released and given him premises in Moscow on Galushkina Street for the first design bureau jet engines. Arkhip Mikhailovich found and dug up his drawings, but the engine according to his design did not work out right away. Then they simply took a turbojet engine bought from the British and repeated it one by one. But it came down to materials that were not available in the Soviet Union, but were available in England, and their composition, of course, was classified, but they still managed to decipher it.
Having arrived in England to get acquainted with the production of engines, S. T. Kishkin appeared everywhere wearing boots with thick microporous soles. And, having visited the plant where turbine blades were processed on a tour, he, near the machine, as if by accident, stepped on 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 subjected to thorough analysis in Moscow. The results of the analysis of English metal and extensive in-house 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 fundamentals of the theory of their structure and production.
It was found that the main carrier of the heat resistance of such alloys are submicroscopic particles of the intermetallic phase based on the Ni3Al compound. Blades made of the first heat-resistant nickel alloys could operate 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 rod to a shape vaguely reminiscent of the finished product, and then carefully and painstakingly machined. But here an unexpected difficulty arose: in order to increase the operating temperature of the material, alloying elements were added to it - tungsten, molybdenum, niobium. But they made the alloy so hard that it became impossible to stamp it - it could not be molded using hot deformation methods.
Then Kishkin suggested casting the blades. The engine designers were indignant: firstly, after casting, the blade would still have to be processed on machines, and most importantly, how can a cast blade be installed in the engine? The metal of stamped blades is very dense, its strength is high, but 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 aviation turbojet engines began to be produced with cast turbine blades.
The first blades were solid and lasted a long time high temperature could not. It was necessary to create a cooling system for them. To do this, they decided to make longitudinal channels in the blades to supply cooling air from the compressor. This idea was not so hot: the more air from the compressor is used for cooling, the less of it will go into the combustion chambers. But there was nowhere to go - the turbine resource 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 this design was ineffective: the air flows through the channel too quickly, the area of the cooled surface is small, and the heat is not removed sufficiently. 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 the channels of a more complex shape. At some point, aircraft engine specialists were seized by a tempting idea - to create an entirely ceramic blade: ceramics can withstand very high temperatures and do 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 TO MAKE A CAST BLADE
The technology for manufacturing turbine blades is called lost-wax casting. First, a wax model of the future blade is made, casting it in a mold, into which quartz cylinders are first placed in place of future cooling channels (later they began to use other materials). The model is covered with liquid ceramic mass. After it dries, the wax is melted with hot water, and ceramic mass burn. The result is a mold that can withstand the temperature of the molten metal from 1450 to 1500 o C, depending on the grade of the alloy. Metal is poured into the mold, which hardens in the form of a finished blade, but with quartz rods instead of channels inside. The rods are removed by dissolving in hydrofluoric acid. This operation is carried out in a hermetically sealed room by a worker in a spacesuit with an air supply hose. The technology is inconvenient, dangerous and harmful.
To eliminate this operation, VIAM began making rods from aluminum oxide with the addition of 10-15% silicon oxide, which dissolves in alkali. The material of the blades does not react with alkali, and the remaining aluminum oxide is removed with a strong stream of water.
IN Everyday life We are accustomed to consider cast products to be very rough and rough. But we managed to select such ceramic compositions that the shape of them is completely smooth and casting requires almost no mechanical processing. This greatly simplifies the work: the blades have a very complex shape and are not easy to process.
New materials required new technologies. No matter how convenient the addition of silicon oxide to the rod material was, 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 during the pouring process.
To ensure that the aluminum oxide mold retains 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 should not change when pouring: the walls of the blades are very thin, and the dimensions must exactly correspond to the calculated ones. That's why permissible value mold shrinkage should not exceed 1%.
WHY WE REFUSED STAMPED BLADES
As already mentioned, after stamping the blade had to be machined. In this case, 90% of the metal went into chips. The task was set: to create such a precision casting technology that would immediately produce a given blade profile, and the finished product would only need to be polished and a heat-protective coating applied to it. No less important is the structure that forms in the body of the blade and performs the task of cooling it.
Thus, it is very important to make a blade that cools efficiently without reducing the working gas temperature and has high long-term strength. This problem was solved by arranging the channels in the body of the blade and the outlet holes from it so that a thin air film appeared around the blade. In this case, they kill two birds with one stone: the hot gases do not come into contact with the material of the blade, and therefore do not heat it up and do not cool themselves.
There is some analogy here with the thermal protection of a space rocket. When a rocket enters the dense layers of the atmosphere at high speed, the so-called sacrificial coating covering the warhead 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 a turbine blade is based on the same principle, only air is used instead of a sacrificial coating. True, the blades also need to be protected from erosion and corrosion.
The procedure for making a blade is as follows. First, a nickel alloy is created with specified parameters for mechanical strength and heat resistance, for which alloying additives are introduced into nickel: 6% aluminum, 6-10% tungsten, tantalum, rhenium and a little ruthenium. They allow you to achieve the maximum high temperature performance for cast nickel-based alloys (it is tempting to increase this further by using more rhenium, but it is insanely expensive). The use of niobium silicide is considered a promising direction, but this is a matter of the distant future.
But the alloy is poured into the mold at a temperature of 1450 o C and cools along with it. The cooling metal crystallizes, forming individual equiaxed, that is, approximately the same size in all directions, grains. The grains themselves can be large or small. They do not adhere reliably, and the working blades were destroyed along the grain boundaries and shattered into pieces. Not a single blade could last longer than 50 hours. Then we proposed introducing a modifier into the material of the casting mold - crystals of cobalt aluminate. They serve as centers, crystallization nuclei, accelerating the process of grain formation. The grains are uniform and small. New blades began to work for 500 hours. This technology, which was developed by E. N. Kablov, still works, and works well. And we at VIAM produce tons of cobalt aluminate and supply it to factories.
The power of the jet engines grew, the temperature and pressure of the gas jet increased. And it became clear that the multi-grain structure of the metal of the blade would not be able to work under the new conditions. Other ideas were needed. They were found and brought to the stage technological development and became known as directional crystallization. This means that the metal, when solidifying, 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 remember the old parable about a broom that cannot be broken, although all its individual twigs break without difficulty.
HOW DIRECTED CRYSTALLIZATION IS PRODUCED
To ensure that the crystals that form the paddle grow properly, the mold containing the molten metal is slowly removed from the heating zone. In this case, the mold with liquid metal stands on a massive copper disk cooled by water. Crystal growth starts from the bottom and goes up at a speed almost equal to the speed at which the mold exits the heater. When creating the technology of directional crystallization, it was necessary to measure and calculate many parameters - the rate of crystallization, the temperature of the heater, the temperature gradient between the heater and the refrigerator, etc. It was necessary to select such a speed of movement of the mold that columnar crystals would grow along the entire length of the blade. If all these conditions are met, 5-7 long columnar crystals grow for every square centimeter of the cross-section of the blade. This technology has enabled the creation of a new generation of aircraft engines. But we went even further.
Having studied the grown columnar crystals using X-ray methods, we realized that the entire blade can be made from one crystal, which will not have intergrain boundaries - the weakest elements of the structure 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 of such a seed is 0-0-1; this means that the crystal grows in the direction of the Z axis, but not in the X-Y direction). The seed was placed in the lower part of the mold and the metal was poured, intensively cooling it from below. The growing single crystal took on the shape of a blade.
American engineers used a water-cooled copper crystallizer for cooling. And after several experiments, we replaced it with a bath of molten tin at a temperature of 600-700 K. This made it possible to more accurately select the required temperature gradient and obtain products High Quality. VIAM built installations with baths for growing single-crystalline blades - very advanced machines with computer control.
In the 1990s, when the USSR collapsed, in the territory East Germany Soviet aircraft remained, mainly MiG fighters. Their engines had blades of our production. The metal of the blades was examined by the Americans, after which quite soon their specialists came to VIAM and asked to show who created it and how. It turned out that they were given the task of making meter-long monocrystalline blades, which they could not solve. We designed a plant 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 already have an almost ready-made 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 energy sector are another interesting problem that VIAM was solving. Aircraft engines that have reached the end of their service life began to be used at gas pipeline compressor stations and in power plants that power oil pipeline pumps. Now it has become urgent to create special engines for these needs that would operate at much lower temperatures and working gas pressure, but for much longer. If the service life of an aircraft engine is about 500 hours, then the turbines on the oil and gas pipeline should operate for 20-50 thousand hours. One of the first to start working on them was the Samara design bureau under the leadership of Nikolai Dmitrievich Kuznetsov.
HEAT-RESISTANT ALLOYS
The monocrystalline blade does not grow solid - inside it has a complex-shaped cavity for cooling. 
Together with CIAM, we have developed a cavity configuration that provides a cooling efficiency coefficient (the ratio of the temperatures of the blade metal and the working gas) of 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-heat-resistant alloys containing up to fifteen elements of the periodic table, including rhenium and ruthenium, and heat-protective coatings, in which include nickel, chromium, aluminum and yttrium, and in the future - ceramic made of zirconium oxide stabilized with yttrium oxide.
The first generation alloys contained small amounts of carbon in the form of titanium or tantalum carbides. Carbides are located along the crystal boundaries 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. We have few rhenium reserves in our country; there are deposits in Kazakhstan, but after the collapse Soviet Union it was completely bought up by the Americans; There 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 alloys, the technology for their production, and the method of casting the finished product. A huge amount of work and knowledge of all VIAM employees has been put into all the blades.
Candidate of Technical Sciences I. DEMONIS, Deputy General Director of VIAM
The “turbine” topic is as complex as it is vast. Therefore, of course, there is no need to talk about its full disclosure. Let’s deal, as always, with “general acquaintance” and “individual interesting points”...
Moreover, the history of the aviation turbine is very short compared to the history of the turbine in general. This means that we cannot do without some kind of theoretical and historical excursion, the content of which for the most part does not relate to aviation, but is the basis for a story about the use of a gas turbine in aircraft engines.
About the hum and roar...
Let's start somewhat unconventionally and remember about "". This is a fairly common phrase, usually used by inexperienced authors in the media when describing the operation of powerful aircraft. Here you can add “roar, whistle” and other loud definitions for the same “aircraft turbines”.
Quite familiar words for many. However, people who understand are well aware that in fact all these “sound” epithets most often characterize the operation of jet engines as a whole or its parts, which have very little to do with turbines as such (except, of course, for the mutual influence during their joint operation in the general turbojet engine cycle).
Moreover, in a turbojet engine (these are the object of rave reviews), as a direct reaction engine that creates thrust by using the reaction of a gas jet, the turbine is just part of it and is rather indirectly related to the “rumbling roar”.
And on those engines where it, as a unit, plays, in some way, a dominant role (these are indirect reaction engines, and it’s not for nothing that they are called gas turbine), the sound is no longer so impressive, or it is created by completely different parts of the aircraft’s power plant, for example, a propeller.
That is, neither hum nor rumble, as such, to aircraft turbine don't really apply. However, despite such sound ineffectiveness, it is a complex and very important unit of a modern turbojet engine (GTE), often determining its main operational characteristics. By definition, no gas turbine engine can do without a turbine.
Therefore, the conversation, of course, is not about impressive sounds and incorrect use of definitions of the Russian language, but about an interesting unit and its relationship to aviation, although this is far from the only area of its application. As a technical device, the turbine appeared long before the very concept of an “aircraft” (or airplane) and even more so a gas turbine engine for it.
History + a little theory...
And even for a very long time. Ever since the invention of mechanisms that convert the energy of natural forces into useful action. The simplest in this regard and therefore one of the first to appear were the so-called rotary engines.
This definition itself, of course, appeared only in our days. However, its meaning precisely determines the simplicity of the engine. Natural energy directly, without any intermediate devices, is converted into mechanical power rotational movement of the main power element such an engine is a shaft.
Turbine– a typical representative of a rotary engine. Looking ahead, we can say that, for example, in a piston engine internal combustion(ICE) the main element is the piston. It performs a reciprocating motion, and to obtain rotation of the output shaft, you need to have an additional crank mechanism, which naturally complicates and makes the design heavier. The turbine is much more profitable in this regard.
For a rotary internal combustion engine, like a heat engine, which, by the way, is a turbojet engine, the name “rotary” is usually used.
Water mill turbine wheel
One of the best known and most ancient applications of turbines are large mechanical mills, used by man since time immemorial for various economic needs (not just for grinding grain). They are treated as water, so wind mechanisms.
For a long period of ancient history (the first mentions from about the 2nd century BC) and the history of the Middle Ages, these were virtually the only mechanisms used by man for practical purposes. The possibility of their use, despite all the primitiveness of the technical circumstances, lay in the simplicity of transformation of the energy of the working fluid used (water, air).
A windmill is an example of a turbine wheel.
In these essentially true rotary engines, the energy of water or air flow is converted into shaft power and, ultimately, useful work. This happens when the flow interacts with the working surfaces, which are water wheel blades or windmill wings. Both of them, in fact, are prototypes of modern blades blade machines, which are the turbines used today (and compressors, by the way, too).
Another type of turbine is known, first documented (apparently invented) by the ancient Greek scientist, mechanic, mathematician and naturalist Heron of Alexandria ( Heron ho Alexandreus,1 1st century AD) in his treatise “Pneumatics”. The invention he described was called aeolipile , which translated from Greek means “ball of Aeolus” (god of the wind, Αἴολος – Aeolus (Greek), pila - ball (lat.)).
Aeolipile of Heron.
In it, the ball was equipped with two oppositely directed nozzle tubes. Steam came out of the nozzles, entering the ball through pipes from the boiler located below and thereby causing the ball to rotate. The action is clear from the figure below. It was a so-called reverse turbine, rotating in the direction opposite to the steam outlet. Turbines This type has a special name - reactive (more details below).
It is interesting that Heron himself hardly imagined what was the working fluid in his machine. In that era, steam was identified with air, even the name testifies to this, because Aeolus commands the wind, that is, the air.
Aeolipile was, in general, a full-fledged heat engine that converted the energy of burned fuel into mechanical energy rotation on the shaft. Perhaps it was one of the first heat engines in history. True, its usefulness was still “not complete”, since useful work did not make the invention.
Aeolipile, among other mechanisms known at that time, was part of the so-called “automata theater”, which was very popular in subsequent centuries, and was in fact just an interesting toy with an unclear future.
From the moment of its creation and, in general, from the era when people in their first mechanisms used only “obviously manifesting themselves” forces of nature (the force of wind or the force of gravity of falling water) to the beginning of the confident use of thermal energy of fuel in newly created heat engines, more than one hundred years have passed years.
The first such units were steam engines. Real working examples were invented and built in England only towards the end of the 17th century and were used to pump water from coal mines. Later, steam engines with a piston mechanism appeared.
Subsequently, as technical knowledge developed, piston internal combustion engines of various designs, more advanced mechanisms with higher efficiency, “came onto the scene.” They already used gas (combustion products) as a working fluid and did not require bulky steam boilers to heat it.
Turbines as the main components of heat engines, also followed a similar path in their development. And although there are separate mentions of some specimens in history, noteworthy and, moreover, documented units, including patented ones, appeared only in the second half of the 19th century.
It all started with a couple...
It was with the use of this working fluid that almost all the basic principles of the design of a turbine (later also a gas turbine) as an important part of a heat engine were developed.
Jet turbine patented by Laval.
The developments of a talented Swedish engineer and inventor are quite characteristic in this regard. Gustave de Laval(Karl Gustaf Patrick de Laval). His research at that time was related to the idea of developing a new milk separator with increased drive speed, which could significantly increase productivity.
Obtain a higher rotation frequency (rpm) by using the then traditional (indeed, the only existing) piston steam engine It was not possible due to the high inertia of the most important element - the piston. Realizing this, Laval decided to try to stop using the piston.
They say that the idea itself came to him while observing the work sandblasting machines. In 1883 he received his first patent (English patent no. 1622) in this field. The patented device was called " Turbine powered by steam and water».
It was an S-shaped tube, at the ends of which tapering nozzles were made. The tube was mounted on a hollow shaft, through which steam was supplied to the nozzles. Fundamentally, all this was no different from the aeolipile of Heron of Alexandria.
The manufactured device worked quite reliably with high speeds for the technology of that time - 42,000 rpm. The rotation speed reached 200 m/s. But with such good parameters turbine had extremely low efficiency. And attempts to increase it with the existing level of technology led to nothing. Why did this happen?
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A little theory... A little more detail about the features....
The mentioned efficiency (for modern aircraft turbines this is the so-called power or effective efficiency) characterizes the efficiency of using the energy expended (available) to drive the turbine shaft. That is, what part of this energy was spent usefully on rotating the shaft, and what part " went down the drain».
It just flew out. For the type of turbine described, called jet, this expression is just right. Such a device receives rotational movement on the shaft under the action of the reaction force of the escaping gas stream (or in this case steam).
A turbine, as a dynamic expansion machine, unlike volumetric machines (piston machines), requires for its operation not only compression and heating of the working fluid (gas, steam), but also its acceleration. Here, expansion (increase in specific volume) and pressure drop occur due to acceleration, in particular in the nozzle. In a piston engine this occurs due to an increase in the volume of the cylinder chamber.
As a result, the large potential energy of the working fluid, which was formed as a result of the supply of thermal energy of burnt fuel to it, turns into kinetic energy (minus various losses, of course). And kinetic (in a jet turbine) through reaction forces - in mechanical work on the shaft.
And the efficiency tells us how completely the kinetic energy transforms into mechanical energy in a given situation. The higher it is, the less kinetic energy the flow leaving the nozzle into environment. This remaining energy is called " losses with output speed", and it is directly proportional to the square of the speed of the outgoing flow (everyone probably remembers mС 2/2).
Operating principle of a jet turbine.
Here we are talking about the so-called absolute speed C. After all, the outgoing flow, or more precisely, each of its particles, participates in a complex movement: rectilinear plus rotational. Thus, the absolute speed C (relative to the fixed coordinate system) is equal to the sum of the turbine rotation speed U and the relative flow speed W (speed relative to the nozzle). The sum is of course vector, shown in the figure.
Segner wheel.
Minimal losses (and maximum efficiency) correspond minimum speed C, ideally, it should be equal to zero. And this is only possible if W and U are equal (as can be seen from the figure). The peripheral speed (U) in this case is called optimal.
Such equality would be easy to achieve on hydraulic turbines (such as Segner wheels), since the speed of liquid outflow from the nozzles for them (similar to the speed W) is relatively small.
But this same speed W for gas or steam is much greater due to the large difference in the densities of liquid and gas. So, at a relatively low pressure of only 5 atm. a hydraulic turbine can produce an exhaust velocity of only 31 m/s, and a steam turbine - 455 m/s. That is, it turns out that even at fairly low pressures (only 5 atm.), the Laval jet turbine should, for reasons of ensuring high efficiency, have a peripheral speed above 450 m/s.
For the then level of technological development, this was simply impossible. It couldn't be done reliable design with these parameters. It also made no sense to reduce the optimal peripheral speed by reducing the relative speed (W), since this can only be done by reducing the temperature and pressure, and therefore the overall efficiency.
Active Laval turbine...
The Laval jet turbine did not lend itself to further improvement. Despite the attempts made, things have reached a dead end. Then the engineer took a different path. In 1889, he patented a different type of turbine, which was later called active. Abroad (in English) it is now called impulse turbine, that is, pulsed.
The device claimed in the patent consisted of one or more fixed nozzles supplying steam to bucket-shaped blades mounted on the rim of a movable turbine wheel (or disk).
Active single-stage steam turbine patented by Laval.
The working process in such a turbine is as follows. The steam accelerates in the nozzles with an increase in kinetic energy and a drop in pressure and falls on the working blades, on their concave part. As a result of the impact on the blades of the impeller, it begins to rotate. Or we can also say that rotation occurs due to the impulse action of the jet. Hence the English name impulseturbine.
Moreover, in the interscapular canals, which have an almost constant cross section, the flow does not change its speed (W) and pressure, but changes direction, that is, it turns at large angles (up to 180°). That is, at the exit from the nozzle and at the entrance to the interblade channel: absolute speed C 1, relative W 1, peripheral speed U.
At the output, respectively, C 2, W 2, and the same U. In this case, W 1 = W 2, C 2< С 1 – из-за того, что часть кинетической энергии входящего потока превращается в механическую на валу турбины (импульсное воздействие) и абсолютная скорость падает.
This process is shown in principle in a simplified figure. Also, to simplify the explanation of the process, it is assumed here that the vectors of absolute and peripheral velocities are almost parallel, the flow changes direction in the impeller by 180°.
Steam (gas) flow in the active turbine stage.
If we consider speeds in absolute values, we see that W 1 = C 1 – U, and C 2 = W 2 – U. Thus, based on the above, for the optimal mode, when the efficiency takes maximum values, and losses from the output speed tend to the minimum (that is, C 2 = 0), we have C 1 = 2U or U = C 1 /2.
We find that for an active turbine optimal peripheral speed half the speed of exhaust from the nozzle, that is, such a turbine is half as loaded as a jet turbine and the task of obtaining a higher efficiency is easier.
Therefore, in the future, Laval continued to develop this type of turbine. However, despite the reduction in the required peripheral speed, it still remained quite large, which entailed equally large centrifugal and vibration loads.
Operating principle of an active turbine.
The consequence of this was structural and strength problems, as well as problems of eliminating imbalances, which were often solved with great difficulty. In addition, there were other unresolved and unsolvable factors under the conditions of that time, which ultimately reduced the efficiency of this turbine.
These included, for example, imperfection of the aerodynamics of the blades, causing increased hydraulic losses, as well as the pulsating effect of individual jets of steam. In fact, only a few or even one blade could be active blades that perceive the action of these jets (or jets) at a time. The rest moved idly, creating additional resistance (in a steam atmosphere).
This one has turbines there was no way to increase power by increasing temperature and steam pressure, since this would lead to an increase in peripheral speed, which was absolutely unacceptable due to the same design problems.
In addition, an increase in power (with an increase in peripheral speed) was also inappropriate for another reason. The consumers of turbine energy were low-speed devices compared to it (electric generators were planned for this). Therefore, Laval had to develop special gearboxes for the kinematic connection of the turbine shaft with the consumer shaft.
The ratio of the masses and dimensions of the active Laval turbine and its gearbox.
Due to the large difference in the speed of these shafts, the gearboxes were extremely bulky and were often significantly larger in size and weight than the turbine itself. An increase in its power would entail an even greater increase in the size of such devices.
Eventually Laval active turbine was a relatively low-power unit (working units up to 350 hp), moreover, expensive (due to a large set of improvements), and complete with a gearbox, it was also quite bulky. All this made it uncompetitive and excluded mass application.
An interesting fact is that constructive principle Laval's active turbine was not actually invented by him. Even 250 years before the appearance of his research, a book by the Italian engineer and architect Giovanni Branca entitled “Le Machine” (“Machines”) was published in Rome in 1629.
Among other mechanisms, it contained a description of the “steam wheel”, which contained all the main components built by Laval: a steam boiler, a steam supply tube (nozzle), an active turbine impeller and even a gearbox. Thus, long before Laval, all these elements were already known, and his merit was that he made them all really work together and was extremely involved in complex issues improving the mechanism as a whole.
Steam active turbine by Giovanni Branca.
Interestingly, one of the most known features his turbine was the design of the nozzle (it was separately mentioned in the same patent), which supplies steam to the rotor blades. Here the nozzle, from an ordinary tapering one, as it was in a jet turbine, became contracting-expanding. Subsequently, this type of nozzle began to be called Laval nozzles. They allow the gas (steam) flow to be accelerated to supersonic speed with fairly low losses. About them .
Thus, the main problem that Laval struggled with when developing his turbines, and which he was never able to overcome, was the high peripheral speed. However, a fairly effective solution to this problem has already been proposed and even, oddly enough, by Laval himself.
Multi-stage….
In the same year (1889), when the above-described active turbine was patented, the engineer developed an active turbine with two parallel rows of rotor blades mounted on one impeller (disk). It was the so-called two-stage turbine.
Steam was supplied to the working blades, just as in the single-stage one, through a nozzle. Between the two rows of working blades, a row of fixed blades was installed, which redirected the flow emerging from the blades of the first stage to the working blades of the second.
If we use the simplified principle proposed above for determining the peripheral speed for a single-stage jet turbine (Laval), it turns out that for a two-stage turbine the rotation speed is no longer two, but four times less than the exhaust speed from the nozzle.
The principle of the Curtis wheel and changing parameters in it.
This is the most effective solution to the problem of low optimal peripheral speed, which was proposed but not used by Laval and which is actively used in modern turbines, both steam and gas. Multi-stage…
It means that the large available energy of the entire turbine can in some way be divided into parts according to the number of stages, and each such part is activated in a separate stage. The lower this energy, the lower the speed of the working fluid (steam, gas) entering the working blades and, therefore, the lower the optimal peripheral speed.
That is, by changing the number of turbine stages, you can change the rotation speed of its shaft and, accordingly, change the load on it. In addition, multi-stage operation makes it possible to operate the turbine with large energy differences, that is, to increase its power while maintaining high efficiency indicators.
Laval did not patent his two-stage turbine, although a prototype was made, so it bears the name of the American engineer Charles Curtis (Curtis wheel (or disk), who in 1896 received a patent for a similar device.
However, much earlier, in 1884, the English engineer Charles Algernon Parsons developed and patented the first real multi-stage steam turbine. There were many statements by various scientists and engineers about the usefulness of dividing available energy into stages before him, but he was the first to translate the idea into hardware.
Multistage active-reaction Parsons turbine (disassembled).
At the same time, his turbine had a feature that brought it closer to modern devices. In it, steam expanded and accelerated not only in nozzles formed by fixed blades, but also partially in channels formed by specially profiled working blades.
This type of turbine is usually called a jet turbine, although the name is quite arbitrary. In fact, it occupies an intermediate position between the purely reactive Heron-Laval turbine and the purely active Laval-Branca turbine. Due to their design, working blades combine active and reactive principle s in the overall process. Therefore, it would be more correct to call such a turbine active-reactive, which is often done.
Diagram of a multistage Parsons turbine.
Parsons worked on various types of multistage turbines. Among his designs were not only the above-described axial ones (the working fluid moves along the axis of rotation), but also radial ones (the steam moves in the radial direction). His three-stage purely active turbine “Heron” is quite well known, in which the so-called Heron wheels are used (the essence is the same as that of the aeolipile).
Jet turbine "Heron".
Subsequently, from the early 1900s, steam turbine construction rapidly gained momentum and Parsons was at its forefront. Its multi-stage turbines were equipped with naval vessels, first experimental ones (ship "Turbinia", 1896, displacement 44 tons, speed 60 km/h - unprecedented for that time), then military ones (example - battleship "Dreadnought", 18000 tons, speed 40 km/ h, turbine power 24,700 hp) and passenger (example - the same type "Mauritania" and "Lusitania", 40,000 tons, speed 48 km/h, turbo power 70,000 hp). At the same time, stationary turbine construction began, for example, by installing turbines as drives in power plants (Edison Company in Chicago).
About gas turbines...
However, let's return to our main topic - aviation and note one fairly obvious thing: such clearly visible success in the operation of steam turbines could have only structural and fundamental significance for aviation, which was rapidly progressing in its development at exactly the same time.
Application steam turbine for obvious reasons, it was extremely doubtful as a power plant on aircraft. Aviation turbine could only be a fundamentally similar, but much more profitable gas turbine. However, not everything was so simple...
According to Lev Gumilyovsky, author of the popular book “Engine Creators” in the 60s, one day, in 1902, during the period of the beginning of the rapid development of steam turbine construction, Charles Parsons, in fact one of the main ideologists of this business at that time, was asked, in general, , a humorous question: “ Is it possible to “parsonize” a gas engine?"(implying a turbine).
The answer was expressed in absolutely decisive form: “ I think that a gas turbine will never be created. No two ways about it." The engineer failed to become a prophet, but he undoubtedly had reasons to say so.
The use of a gas turbine, especially if we mean its use in aviation instead of a steam turbine, was of course tempting, because its positive aspects are obvious. With all its power capabilities, it does not require huge, bulky devices for generating steam - boilers, or equally large devices and systems for its cooling - condensers, cooling towers, cooling ponds, etc.
The heater for a gas turbine engine is a small, compact one, located inside the engine and burning fuel directly in the air flow. And he simply doesn’t have a refrigerator. Or rather, it exists, but it exists as if virtually, because the exhaust gas is discharged into the atmosphere, which is the refrigerator. That is, everything necessary for a heat engine is available, but at the same time everything is compact and simple.
True, a steam turbine plant can also do without a “real refrigerator” (without a condenser) and release steam directly into the atmosphere, but then you can forget about efficiency. An example of this is a steam locomotive - the real efficiency is about 6%, 90% of its energy flies out into the chimney.
But with such tangible advantages, there are also significant disadvantages, which, in general, became the basis for Parsons’ categorical answer.
Compression of the working fluid for subsequent implementation of the work cycle, incl. and in the turbine...
In the operating cycle of a steam turbine plant (Rankine cycle), the work of water compression is small and the requirements for the pump performing this function and its efficiency are therefore also small. In the gas turbine engine cycle, where air is compressed, this work, on the contrary, is very impressive, and most of the available energy of the turbine is spent on it.
This reduces the proportion of useful work that the turbine can be designed for. Therefore, the requirements for an air compression unit in terms of its efficiency and economy are very high. Compressors in modern aviation gas turbine engines (mainly axial), as well as in stationary units, along with turbines, are complex and expensive devices. About them .
Temperature…
This is the main problem for gas turbines, including aviation ones. The fact is that if in a steam turbine installation the temperature of the working fluid after the expansion process is close to the temperature of the cooling water, then in a gas turbine it reaches several hundred degrees.
This means that a large amount of energy is released into the atmosphere (like into a refrigerator), which, of course, negatively affects the efficiency of the entire operating cycle, which is characterized by thermal efficiency: η t = Q 1 – Q 2 / Q 1 . Here Q 2 is the same energy released into the atmosphere. Q 1 – energy supplied to the process from the heater (in the combustion chamber).
In order to increase this efficiency, it is necessary to increase Q 1, which is equivalent to increasing the temperature in front of the turbine (that is, in the combustion chamber). But the fact of the matter is that it is not always possible to raise this temperature. Its maximum value is limited by the turbine itself and the main condition here is strength. The turbine operates under very difficult conditions, when high temperatures are combined with high centrifugal loads.
It is this factor that has always limited the power and thrust capabilities of gas turbine engines (largely dependent on temperature) and has often become the reason for the complexity and rise in cost of turbines. This situation has continued in our time.
And at the time of Parsons, neither the metallurgical industry nor aerodynamic science could yet provide a solution to the problems of creating an efficient and economical compressor and high-temperature turbine. There was neither an appropriate theory nor the necessary heat-resistant and heat-resistant materials.
And yet there were attempts...
Nevertheless, as usually happens, there were people who were not afraid (or maybe did not understand :-)) possible difficulties. Attempts to create a gas turbine did not stop.
Moreover, it is interesting that Parsons himself, at the dawn of his “turbine” activity, in his first patent for a multi-stage turbine, noted the possibility of its operation, in addition to steam, also on fuel combustion products. It was also considered there possible variant gas turbine engine running on liquid fuel with a compressor, combustion chamber and turbine.
Smoke spit.
Examples of the use of gas turbines without any theory behind it have been known for a long time. Apparently, even Heron used the principle of an air jet turbine in the “theater of automata”. The so-called “smoke skewers” are quite widely known.
And in the already mentioned book by the Italian (engineer, architect, Giovanni Branca, Le Machine) Giovanni Branca there is a drawing “ Fire wheel" In it, a turbine wheel rotates with combustion products from a fire (or hearth). It is interesting that Branca himself did not build most of his cars, but only expressed ideas for their creation.
"Wheel of Fire" by Giovanni Branca.
In all these “smoke and fire wheels” there was no air (gas) compression stage, and there was no compressor as such. The conversion of potential energy, that is, the supplied thermal energy of fuel combustion, into kinetic energy (acceleration) for rotation of the gas turbine occurred only due to the action of gravity when the warm masses rose upward. That is, the phenomenon of convection was used.
Of course, such “units” are for real machines, for example, for driving Vehicle could not be used. However, in 1791, the Englishman John Barber patented a “horseless transportation machine”, one of the most important components of which was a gas turbine. This was the first ever officially registered patent for a gas turbine.
John Barber engine with gas turbine.
The machine used gas obtained from wood, coal or oil, heated in special gas generators (retorts), which, after cooling, entered a piston compressor, where it was compressed along with air. Next, the mixture was fed into the combustion chamber, and after that the combustion products were rotated turbine. Water was used to cool the combustion chambers, and the resulting steam was also sent to the turbine.
The level of development of technology at that time did not allow the idea to be brought to life. The current model of the Barber machine with a gas turbine was built only in 1972 by Kraftwerk-Union AG for the Hannover Industrial Fair.
Throughout the 19th century, development of the gas turbine concept progressed extremely slowly for the reasons described above. There were few examples worthy of attention. The compressor and high temperature remained an insurmountable stumbling block. There have been attempts to use a fan to compress air, as well as the use of water and air to cool structural elements.
Engine F. Stolze. 1 - axial compressor, 2 - axial turbine, 3 - heat exchanger.
There is a well-known example of a gas turbine engine by the German engineer Franz Stolze, patented in 1872 and very similar in design to modern gas turbine engines. In it, a multi-stage axial compressor and a multi-stage axial turbine were located on the same shaft.
The air after passing through the regenerative heat exchanger was divided into two parts. One entered the combustion chamber, the second was mixed with the combustion products before entering the turbine, reducing their temperature. This is the so-called secondary air, and its use is a technique widely used in modern gas turbine engines.
The Stolze engine was tested in 1900-1904, but turned out to be extremely inefficient due to the low quality of the compressor and the low temperature in front of the turbine.
For most of the first half of the 20th century, the gas turbine was never able to actively compete with the steam turbine or become part of the gas turbine engine, which could adequately replace the piston internal combustion engine. Its use on engines was mainly auxiliary. For example, as charging units in piston engines, including aviation ones.
But from the beginning of the 40s the situation began to change rapidly. Finally, new heat-resistant alloys were created, which made it possible to radically increase the temperature of the gas in front of the turbine (up to 800˚C and higher), and quite economical ones with high efficiency appeared.
This not only made it possible to build efficient gas turbine engines, but also, thanks to the combination of their power with relative lightness and compactness, to use them on aircraft. The era of jet aviation and aircraft gas turbine engines began.
Turbines in aviation gas turbine engines…
So... The main area of application of turbines in aviation is gas turbine engines. The turbine here does the hard work - it rotates the compressor. Moreover, in a gas turbine engine, as in any heat engine, the work of expansion is greater than the work of compression.
And the turbine is precisely an expansion machine, and it spends only part of the available energy of the gas flow on the compressor. The remaining part (sometimes called free energy) can be used for useful purposes depending on the type and design of the engine.
Scheme of TVAD Makila 1a1 with a free turbine.
Turboshaft engine AMAKILA 1A1.
For indirect reaction engines, such as (helicopter gas turbine engines), it is spent on rotating the propeller. In this case, the turbine is most often divided into two parts. The first one is compressor turbine. The second, driving the screw, is the so-called free turbine. It rotates independently and is connected to the compressor turbine only gasdynamically.
In direct reaction engines (jet engines or jet engines), the turbine is used only to drive the compressor. The remaining free energy, which rotates the free turbine in the TVAD, is activated in the nozzle, turning into kinetic energy to produce jet thrust.
In the middle between these extremes are located. In them, part of the free energy is spent to drive the propeller, and some part forms jet thrust in the output device (nozzle). True, its share in the total engine thrust is small.
Diagram of a single-shaft turboprop engine DART RDa6. Turbine on a common engine shaft.
Rolls-Royce DART RDa6 turboprop single-shaft engine.
By design, turboprop engines can be single-shaft, in which the free turbine is not separated structurally and, being one unit, drives both the compressor and the propeller at once. An example of a Rolls-Royce DART RDa6 theater, as well as our famous AI-20 theater.
There may also be a turboprop engine with a separate free turbine that drives the propeller and is not mechanically connected to the rest of the engine components (gas-dynamic coupling). An example is the PW127 engine of various modifications (airplanes), or the Pratt & Whitney Canada PT6A turboprop engine.
Scheme of the Pratt & Whitney Canada PT6A with a free turbine.
Engine Pratt & Whitney Canada PT6A.
Scheme of a PW127 turboprop engine with a free turbine.
Of course, in all types of gas turbine engines, the payload also includes units that ensure the operation of the engine and aircraft systems. These are usually pumps, fuel and hydraulic generators, electric generators, etc. All these devices are most often driven from the turbocharger shaft.
About types of turbines.
There are actually quite a few types. Just for example, some names: axial, radial, diagonal, radial-axial, rotary-blade, etc. In aviation, only the first two are used, and radial is quite rare. Both of these turbines were named in accordance with the nature of the gas flow in them.
Radial.
In radial it flows along a radius. Moreover, in the radial aircraft turbine a centripetal flow direction is used, which provides higher efficiency (in non-aviation practice there is also a centrifugal direction).
The radial turbine stage consists of an impeller and fixed blades that form the flow at its inlet. The blades are profiled so that the inter-blade channels have a tapering configuration, that is, they are nozzles. All these blades, together with the body elements on which they are mounted, are called nozzle apparatus.
Diagram of a radial centripetal turbine (with explanations).
The impeller is an impeller with specially profiled blades. The impeller spins up when gas passes through the narrowing channels between the blades and acts on the blades.
Impeller of a radial centripetal turbine.
Radial turbines They are quite simple, their impellers have a small number of blades. The possible circumferential speeds of a radial turbine at the same stresses in the impeller are greater than those of an axial turbine, so it can generate larger amounts of energy (heat drops).
However, these turbines have a small flow area and do not provide sufficient gas flow with the same dimensions compared to axial turbines. In other words, they have too large relative diametrical dimensions, which complicates their arrangement in a single engine.
In addition, it is difficult to create multi-stage radial turbines due to large hydraulic losses, which limits the degree of gas expansion in them. It is also difficult to cool such turbines, which reduces the possible maximum gas temperatures.
Therefore, the use of radial turbines in aviation is limited. They are mainly used in low-power units with low gas consumption, most often in auxiliary mechanisms and systems or in the engines of model aircraft and small unmanned aircraft.
The first Heinkel He 178 jet aircraft.
Heinkel HeS3 turbojet engine with radial turbine.
One of the few examples of the use of a radial turbine as a component of a propulsion aircraft jet engine is the engine of the first real jet aircraft, the Heinkel He 178 turbojet Heinkel HeS 3. The photo clearly shows the stage elements of such a turbine. The parameters of this engine were fully consistent with the possibility of its use.
Axial aircraft turbine.
This is the only type of turbine currently used in mid-flight aircraft gas turbine engines. The main source of mechanical work on the shaft obtained from such a turbine in the engine is the impellers, or more precisely, the impeller blades (RL), installed on these impellers and interacting with the energetically charged gas flow (compressed and heated).
The crowns of the stationary blades installed in front of the workers organize the correct direction of the flow and participate in the conversion of the potential energy of the gas into kinetic energy, that is, they accelerate it in the process of expansion with a drop in pressure.
These blades, complete with the housing elements on which they are mounted, are called nozzle apparatus(SA). The nozzle apparatus complete with working blades is turbine stage.
The essence of the process... Generalization of what has been said...
In the process of the above-mentioned interaction with the working blades, the kinetic energy of the flow is converted into mechanical energy, which rotates the engine shaft. Such a transformation in an axial turbine can occur in two ways:
An example of a single-stage active turbine. The change in parameters along the path is shown.
1. Without changing the pressure, and therefore the magnitude of the relative flow velocity (only its direction changes noticeably - the rotation of the flow) in the turbine stage; 2. With a drop in pressure, an increase in the relative flow velocity and some change in its direction in the stage.
Turbines operating using the first method are called active. The gas flow actively (pulses) affects the blades due to a change in its direction as it flows around them. With the second method - jet turbines. Here, in addition to the impulse effect, the flow also affects the rotor blades indirectly (to put it simply), using reactive force, which increases the power of the turbine. Additional reactive action is achieved through special profiling of the rotor blades.
The concepts of activity and reactivity in general, for all turbines (not only aviation ones) were mentioned above. However, modern aviation gas turbine engines use only axial jet turbines.
Changing parameters in the stage of an axial gas turbine.
Since the force effect on the radar is double, such axial turbines are also called active-reactive, which is perhaps more correct. This type of turbine is more aerodynamically advantageous.
The fixed blades of the nozzle apparatus included in the stage of such a turbine have a large curvature, due to which the cross-section of the inter-blade channel decreases from inlet to outlet, that is, the cross-section f 1 is less than the cross-section f 0 . The result is a profile of a tapering jet nozzle.
The working blades that follow them also have greater curvature. In addition, with respect to the oncoming flow (vector W 1), they are located so as to avoid its breakdown and ensure correct flow around the blade. At certain radii, the RL also forms tapering interscapular channels.
Stage work aviation turbine.
The gas approaches the nozzle apparatus with a direction of movement close to axial and a speed C 0 (subsonic). Pressure in the flow P 0, temperature T 0. Passing the interscapular channel, the flow accelerates to speed C 1 with a turn to angle α 1 = 20°-30°. In this case, the pressure and temperature drop to values P 1 and T 1, respectively. Part of the potential energy of the flow is converted into kinetic energy.
Picture of the gas flow movement in the axial turbine stage.
Since the working blades move with a peripheral speed U, the flow enters the interblade channel of the RL with a relative speed W 1, which is determined by the difference between C 1 and U (vectorally). Passing through the channel, the flow interacts with the blades, creating aerodynamic forces P on them, the circumferential component of which P u causes the turbine to rotate.
Due to the narrowing of the channel between the blades, the flow accelerates to speed W 2 (reactive principle), while it also rotates (active principle). The absolute flow velocity C 1 decreases to C 2 - the kinetic energy of the flow is converted into mechanical energy on the turbine shaft. Pressure and temperature drop to values P 2 and T 2, respectively.
The absolute flow velocity as it passes through the stage increases slightly from C 0 to the axial projection of the velocity C 2 . In modern turbines this projection has a value of 200 - 360 m/s for a stage.
The step is profiled so that the angle α 2 is close to 90°. The difference is usually 5-10°. This is done to ensure that the value of C 2 is minimal. This is especially important for the last stage of the turbine (at the first or middle stages, deviations from right angle up to 25°). The reason for this is loss with output speed, which just depend on the magnitude of the speed C 2.
These are the same losses that at one time did not give Laval the opportunity to increase the efficiency of his first turbine. If the engine is jet, then the remaining energy can be used in the nozzle. But, for example, for a helicopter engine that does not use jet thrust, it is important that the flow speed behind the last stage of the turbine is as low as possible.
Thus, in the stage of an active-reactive turbine, the expansion of gas (a decrease in pressure and temperature), the conversion and activation of energy (heat difference) occurs not only in the SA, but also in the impeller. The distribution of these functions between the RC and the SA is characterized by a parameter of the engine theory called degree of reactivity ρ.
It is equal to the ratio of the heat drop in the impeller to the heat drop in the entire stage. If ρ = 0, then the stage (or the entire turbine) is active. If ρ > 0, then the stage is reactive or, more precisely, for our case, active-reactive. Since the profiling of the working blades varies along the radius, this parameter (as well as some others) is calculated according to the average radius (section B-B in the figure of changes in parameters in a stage).
Configuration of the working blade of an active-reaction turbine.
Change in pressure along the length of the radar blade of an active-reactive turbine.
For modern gas turbine engines, the degree of reactivity of turbines is in the range of 0.3-0.4. This means that only 30-40% of the total heat drop of the stage (or turbine) is actuated in the impeller. 60-70% is activated in the nozzle apparatus.
Something about losses.
As already mentioned, any turbine (or its stage) converts the flow energy supplied to it into mechanical work. However, in a real unit this process may have different efficiencies. Part of the available energy is necessarily wasted, that is, it turns into losses, which must be taken into account and measures taken to minimize them in order to increase the efficiency of the turbine, that is, increase its efficiency.
Losses consist of hydraulic and loss with output speed. Hydraulic losses include profile and end losses. Profile is, in fact, friction losses, since the gas, having a certain viscosity, interacts with the surfaces of the turbine.
Typically, such losses in the impeller are about 2-3%, and in the nozzle apparatus - 3-4%. Measures to reduce losses consist of “improving” the flow part by calculation and experiment, as well as correct calculation of velocity triangles for the flow in the turbine stage, or more precisely, choosing the most advantageous peripheral speed U at a given speed C 1 . These actions are usually characterized by the U/C 1 parameter. The peripheral speed at the middle radius in a turbojet engine is 270 – 370 m/s.
The hydraulic perfection of the flow path of a turbine stage takes into account such a parameter as adiabatic efficiency. Sometimes it is also called bladed, because it takes into account friction losses in the stage blades (SA and RL). There is another efficiency factor for a turbine, which characterizes it specifically as a unit for generating power, that is, the degree to which available energy is used to create work on the shaft.
This is the so-called power (or effective) efficiency. It is equal to the ratio of the work on the shaft to the available heat drop. This efficiency takes into account losses with output speed. They usually amount to about 10-12% for turbojet engines (in modern turbojet engines C 0 = 100-180 m/s, C 1 = 500-600 m/s, C 2 = 200-360 m/s).
For modern gas turbine engines, the adiabatic efficiency value is about 0.9 - 0.92 for uncooled turbines. If the turbine is cooled, then this efficiency may be lower by 3-4%. Power efficiency is usually 0.78 - 0.83. It is less than adiabatic by the amount of losses with the output speed.
As for the end losses, these are the so-called “ flow losses" The flow part cannot be completely isolated from the rest of the engine due to the presence of rotating components in combination with stationary ones (cases + rotor). Therefore, gas from areas of high pressure tends to flow into areas of low pressure. In particular, for example, from the area in front of the working blade to the area behind it through the radial gap between the blade airfoil and the turbine housing.
Such a gas does not participate in the process of converting flow energy into mechanical energy, because it does not interact with the blades in this regard, that is, end losses occur (or radial clearance losses). They amount to about 2-3% and negatively affect both adiabatic and power efficiency, reduce the efficiency of gas turbine engines, and quite noticeably.
It is known, for example, that an increase in the radial clearance from 1 mm to 5 mm in a turbine with a diameter of 1 m can lead to an increase specific consumption fuel in the engine by more than 10%.
It is clear that it is impossible to completely get rid of the radial clearance, but they are trying to minimize it. It's quite difficult because aircraft turbine– the unit is heavily loaded. Accurate accounting of all factors influencing the size of the gap is quite difficult.
Engine operating modes often change, which means the amount of deformation of the working blades, the disks on which they are mounted, and the turbine housings changes as a result of changes in temperature, pressure and centrifugal forces.
Labyrinth seal.
Here it is necessary to take into account the amount of residual deformation during long-term operation of the engine. Plus, the evolutions performed by the aircraft affect the deformation of the rotor, which also changes the size of the gaps.
Usually the gap is assessed after stopping the warm engine. In this case, the thin outer casing cools faster than the massive disks and shaft and, decreasing in diameter, touches the blades. Sometimes the radial clearance value is simply selected within 1.5-3% of the length of the blade blade.
The principle of honeycomb compaction.
In order to avoid damage to the blades if they touch the turbine body, special inserts made of a material softer than the material of the blades are often placed in it (for example, metal ceramics). In addition, non-contact seals are used. Usually these are labyrinthine or honeycomb labyrinth seals.
In this case, the working blades are banded at the ends of the feather and seals or wedges (for honeycombs) are already placed on the bandage shelves. In honeycomb seals, due to the thin walls of the honeycomb, the contact area is very small (10 times smaller than a conventional labyrinth), so the unit is assembled without a gap. After running-in, the gap is approximately 0.2 mm.
Application of honeycomb seal. Comparison of losses when using honeycomb (1) and a smooth ring (2).
Similar methods of sealing gaps are used to reduce gas leakage from the flow part (for example, into the inter-disk space).
SOURZ…
These are the so-called passive methods radial clearance control. In addition, many gas turbine engines developed (and being developed) since the late 80s are equipped with so-called “ active radial clearance control systems» (SAURZ - active method). These are automatic systems, and the essence of their work is to control the thermal inertia of the housing (stator) of an aircraft turbine.
The rotor and stator (outer casing) of the turbine differ from each other in material and “massiveness”. Therefore, during transient conditions they expand differently. For example, when an engine switches from a reduced operating mode to an increased one, a high-temperature, thin-walled casing warms up faster (than a massive rotor with disks) and expands, increasing the radial clearance between itself and the blades. Plus to this changes in pressure in the duct and the evolution of the aircraft.
To avoid this, an automatic system (usually a FADEC type main regulator) organizes the supply of cooling air to the turbine housing at required quantities. The heating of the housing is thus stabilized within the required limits, which means that the magnitude of its linear expansion and, accordingly, the magnitude of the radial clearances change.
All this allows you to save fuel, which is very important for modern civil aviation. SAURZ systems are most effectively used in turbines low pressure on turbojet engines such as GE90, Trent 900, and some others.
Much less frequently, but quite effectively, to synchronize the heating rates of the rotor and stator, forced airflow of the turbine disks (and not the housing) is used. Such systems are used on CF6-80 and PW4000 engines.
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Axial clearances in the turbine are also regulated. For example, between the outlet edges of the SA and the inlet RLs, there is usually a gap within 0.1-0.4 from the RL chord at the average radius of the blades. The smaller this gap, the less loss flow energy behind the SA (for friction and alignment of the velocity field behind the SA). But at the same time, the vibration of the radar increases due to the alternating impact of the SA from the areas behind the bodies of the blades into the interscapular areas.
A little general about the design...
Axial aviation turbines modern gas turbine engines may have different design shape of the flow part.
Dav = (Din + Dn) /2
1. Shape with constant body diameter (Dн). Here the internal and average diameters along the tract decrease.
Constant outer diameter.
This design fits well into the dimensions of the engine (and the aircraft fuselage). It has a good distribution of work across stages, especially for twin-shaft turbojet engines.
However, in this scheme the so-called bell angle is large, which is fraught with separation of the flow from the internal walls of the housing and, consequently, hydraulic losses.
Constant inner diameter.
When designing, try not to allow the socket angle to exceed 20°.
2. Mold with constant inner diameter (Dв).
The average diameter and body diameter increase along the tract. This scheme does not fit well with the dimensions of the engine. In a turbojet engine, due to the “divergence” of the flow from the internal casing, it is necessary to turn it further onto the SA, which entails hydraulic losses.
Constant average diameter.
The scheme is more suitable for use in turbofan engines.
3. Shape with constant average diameter (Davg). The diameter of the body increases, the internal diameter decreases.
The scheme has the disadvantages of the previous two. But at the same time, the calculation of such a turbine is quite simple.
Modern aircraft turbines are most often multi-stage. The main reason for this (as mentioned above) is the large available energy of the turbine as a whole. To ensure an optimal combination of peripheral speed U and speed C 1 (U/C 1 - optimal), and therefore high overall efficiency and good economy, it is necessary to distribute all available energy across stages.
An example of a three-stage turbojet turbine.
At the same time, however, she herself turbine structurally becomes more complicated and heavier. Due to the small temperature difference at each stage (it is distributed across all stages), a larger number of the first stages are exposed to high temperatures and often require additional cooling.
Four-stage axial turbine turbine.
Depending on the engine type, the number of stages may vary. For turbojet engines, usually up to three, for dual-circuit engines up to 5-8 stages. Typically, if the engine is multi-shaft, then the turbine has several (according to the number of shafts) cascades, each of which drives its own unit and can itself be multi-stage (depending on the bypass ratio).
Twin-shaft axial aircraft turbine.
For example, in the three-shaft Rolls-Royce Trent 900 engine, the turbine has three stages: a single stage to drive the high-pressure compressor, a single stage to drive the intermediate compressor and a five-stage to drive the fan. The joint operation of cascades and the determination of the required number of stages in cascades is described separately in the “engine theory”.
Herself aircraft turbine, simply put, is a structure consisting of a rotor, stator and various auxiliary structural elements. The stator consists of an outer casing, housings nozzle devices and rotor bearing housings. The rotor is usually a disk structure in which the disks are connected to the rotor and to each other using various additional elements and fastening methods.
An example of a single-stage turbojet turbine. 1 - shaft, 2 - SA blades, 3 - impeller disk, 4 - working blades.
On each disk, as the basis of the impeller, there are working blades. When designing blades, they try to make them with a smaller chord due to the smaller width of the rim of the disk on which they are installed, which reduces its mass. But at the same time, in order to maintain turbine parameters, it is necessary to increase the length of the airfoil, which may entail bandaging the blades to increase strength.
Possible types of locks for fastening the working blades in the turbine disk.
The blade is attached to the disk using lock connection. Such a connection is one of the most loaded structural elements in a gas turbine engine. All loads perceived by the blade are transferred to the disk through the lock and reach very large values, especially since, due to the difference in materials, the disk and blades have different linear expansion coefficients, and besides, due to the unevenness of the temperature field, they heat up differently.
In order to assess the possibility of reducing the load in the locking connection and thereby increasing the reliability and service life of the turbine, research papers, among which experiments on bimetallic blades or the use of blisk impellers in turbines.
When using bimetallic blades, the loads in the locks of their fastening on the disk are reduced due to the manufacture of the locking part of the blade from a material similar to material disk (or similar in parameters). The blade blade is made from another metal, after which they are joined using special technologies (a bimetal is obtained).
Blisks, that is, impellers in which the blades are made integral with the disk, generally eliminate the presence of a locking connection, and therefore unnecessary stress in the material of the impeller. Components of this type are already used in compressors of modern turbofan engines. However, for them the issue of repair is significantly complicated and the possibilities for high-temperature use and cooling in aircraft turbine.
An example of fastening rotor blades to a disk using herringbone locks.
The most common method of attaching blades to heavily loaded turbine disks is the so-called herringbone. If the loads are moderate, then other types of locks that are simpler in design, for example, cylindrical or T-shaped, can be used.
Control…
Since the working conditions aviation turbine extremely heavy, and the issue of reliability, as the most important component of an aircraft, is of paramount priority, then the problem of monitoring the condition of structural elements comes first in ground operation. This is especially true for monitoring the internal cavities of the turbine, where the most loaded elements are located.
Inspection of these cavities is of course impossible without the use of modern equipment. remote visual inspection. For aircraft gas turbine engines, various types of endoscopes (borescopes) serve in this capacity. Modern devices This type is quite advanced and has great capabilities.
Inspection of the gas-air tract of a turbojet engine using a Vucam XO endoscope.
A striking example is the portable measuring video endoscope Vucam XO from the German company ViZaar AG. Possessing small in size and weight (less than 1.5 kg), this device is nevertheless very functional and has impressive capabilities for both inspection and processing of received information.
Vucam XO is completely mobile. Its entire set is located in a small plastic case. Video probe with big amount easily replaceable optical adapters have full 360° articulation, with a diameter of 6.0 mm and can have different lengths (2.2 m; 3.3 m; 6.6 m).
Borescopic inspection of a helicopter engine using a Vucam XO endoscope.
Borescopic inspections using such endoscopes are provided for in the regulations for all modern aircraft engines. In turbines, the flow part is usually inspected. The endoscope probe penetrates the internal cavities aviation turbine through special control ports.
Borescopic inspection ports on the CFM56 turbojet turbine housing.
They are holes in the turbine housing, closed with sealed plugs (usually threaded, sometimes spring-loaded). Depending on the capabilities of the endoscope (probe length), it may be necessary to turn the motor shaft. The blades (SA and RL) of the first stage of the turbine can be inspected through the windows on the combustion chamber housing, and those of the last stage - through the engine nozzle.
Which will raise the temperature...
One of the general directions for the development of gas turbine engines of all schemes is to increase the gas temperature in front of the turbine. This makes it possible to significantly increase thrust without increasing air consumption, which can lead to a decrease in the frontal area of the engine and an increase in specific frontal thrust.
In modern engines, the gas temperature (after the flame) at the exit from the combustion chamber can reach 1650°C (with a tendency to increase), so for normal operation Turbines with such high thermal loads require the adoption of special, often safety measures.
First (and most downtime of this situation)- usage heat-resistant and heat-resistant materials, both metal alloys and (in the future) special composite and ceramic materials, which are used for the manufacture of the most loaded parts of the turbine - nozzles and working blades, as well as disks. The most loaded of them are, perhaps, the working blades.
Metal alloys are mainly nickel-based alloys (melting point - 1455 ° C) with various alloying additives. Up to 16 different alloying elements are added to modern heat-resistant and heat-resistant alloys to obtain maximum high-temperature characteristics.
Chemical exotic...
These include, for example, chromium, manganese, cobalt, tungsten, aluminum, titanium, tantalum, bismuth and even rhenium or, instead, ruthenium and others. Particularly promising in this regard is rhenium (Re – rhenium, used in Russia), which is now used instead of carbides, but it is extremely expensive and its reserves are small. The use of niobium silicide is also considered promising.
In addition, the surface of the blade is often coated with a special coating applied using a special technology. heat-protective layer(anti-thermal coating - thermal-barrier coating or fuel assembly) , significantly reducing the amount of heat flow into the body of the blade (thermal barrier functions) and protecting it from gas corrosion (heat-resistant functions).
An example of a thermal protective coating. The nature of the temperature change across the blade cross section is shown.
The figure (microphoto) shows the heat-protective layer on the high-pressure turbine blade of a modern turbofan engine. Here TGO (Thermally Grown Oxide) is a thermally growing oxide; Substrate – the main material of the blade; Bond coat is a transition layer. The composition of fuel assemblies now includes nickel, chromium, aluminum, yttrium, etc. Experimental work is also being carried out on the use ceramic coatings based on zirconium oxide stabilized by zirconium oxide (developed by VIAM).
For example…
Heat-resistant nickel alloys from Special Metals Corporation - USA, containing at least 50% nickel and 20% chromium, as well as titanium, aluminum and many other components added in small quantities, are quite widely known in the engine industry, starting from the post-war period and currently. .
Depending on their profile purpose (RL, SA, turbine disks, flow parts, nozzles, compressors, etc., as well as non-aviation applications), their composition and properties, they are combined into groups, each of which includes various alloy options.
Rolls-Royce Nene engine turbine blades made from Nimonic 80A alloy.
Some of these groups are: Nimonic, Inconel, Incoloy, Udimet/Udimar, Monel and others. For example, Nimonic 90 alloy, developed back in 1945 and used for the manufacture of elements aircraft turbines(mainly blades), nozzles and parts of aircraft, has the composition: nickel - 54% minimum, chromium - 18-21%, cobalt - 15-21%, titanium - 2-3%, aluminum - 1-2%, manganese – 1%, zirconium -0.15% and other alloying elements (in small quantities). This alloy is still produced today.
In Russia (USSR), the development of this type of alloys and other important materials for gas turbine engines was and is being successfully carried out by VIAM (All-Russian Research Institute of Aviation Materials). In the post-war period, the institute developed deformable alloys (type EI437B), and since the early 60s it has created a whole series of high-quality cast alloys (more on this below).
However, almost all are heat-resistant metal materials withstand temperatures up to approximately ≈ 1050°C without cooling.
That's why:
The second, widely used measure, this is an application various cooling systems blades and others structural elements aircraft turbines. It is still impossible to do without cooling in modern gas turbine engines, despite the use of new high-temperature heat-resistant alloys and special methods for manufacturing elements.
Among cooling systems, there are two areas: systems open And closed. Closed-loop systems can use forced circulation liquid coolant in the blade system - a radiator or use the principle of the “thermosyphon effect”.
In the latter method, the movement of the coolant occurs under the influence of gravitational forces, when warmer layers displace colder ones. The coolant here can be, for example, sodium or an alloy of sodium and potassium.
However, closed systems, due to the large number of difficult-to-solve problems, are not used in aviation practice and are at the stage of experimental research.
Approximate cooling diagram of a multi-stage turbojet turbine. The seals between the CA and the rotor are shown. A - a grid of profiles for swirling air for the purpose of pre-cooling it.
But they are in wide practical use open cooling systems. The refrigerant here is air, usually supplied at different pressures due to different compressor stages into the turbine blades. Depending on the maximum gas temperature at which it is advisable to use these systems, they can be divided into three types: convective, convective-film(or barrier) and porous.
During convective cooling, air is supplied inside the blade through special channels and, washing the most heated areas inside it, goes out into the flow in areas with lower pressure. In this case, various schemes for organizing air flow in the blades can be used, depending on the shape of the channels for it: longitudinal, transverse or loop-shaped (mixed or complicated).
Types of cooling: 1 - convective with a deflector, 2 - convective film, 3 - porous. Blade 4 - heat-protective coating.
The simplest scheme is with longitudinal channels along the feather. Here, the air outlet is usually organized in the upper part of the blade through the bandage shelf. In such a scheme, there is a rather large unevenness of temperature along the blade feather - up to 150-250˚, which adversely affects the strength properties of the blade. The circuit is used on engines with gas temperatures up to ≈ 1130ºС.
Another way convective cooling(1) implies the presence of a special deflector inside the feather (a thin-walled shell is inserted inside the feather), which facilitates the supply of cooling air first to the most heated areas. The deflector forms a kind of nozzle that blows air into the front of the blade. This results in jet cooling of the most heated part. Next, the air, washing the remaining surfaces, exits through narrow longitudinal holes in the feather.
Turbine blade of the CFM56 engine.
In such a scheme, temperature unevenness is much lower, in addition, the deflector itself, which is inserted into the blade under tension along several centering transverse belts, thanks to its elasticity, serves as a damper and dampens vibrations of the blades. This scheme is used at a maximum gas temperature of ≈ 1230°C.
The so-called half-loop design makes it possible to achieve a relatively uniform temperature field in the blade. This is achieved by experimentally selecting the location of various ribs and pins that direct air flows inside the blade body. This scheme allows a maximum gas temperature of up to 1330°C.
The nozzle blades are convectively cooled in the same way as the working blades. They are usually made double-cavity with additional ribs and pins to intensify the cooling process. Air at a higher pressure is supplied to the front cavity at the leading edge than to the rear (due to different stages of the compressor) and is released into various zones of the tract in order to maintain the minimum required pressure difference to ensure the required air speed in the cooling channels.
Examples of possible methods for cooling rotor blades. 1 - convective, 2 - convective-film, 3 convective-film with complicated loop channels in the blade.
Convective film cooling (2) is used at even higher gas temperatures – up to 1380°C. With this method, part of the cooling air is released through special holes in the blade onto its outer surface, thereby creating a kind of barrier film, which protects the blade from contact with the hot gas flow. This method is used for both working and nozzle blades.
The third method is porous cooling (3). In this case, the power rod of the blade with longitudinal channels is covered with a special porous material, which allows for a uniform and dosed release of coolant onto the entire surface of the blade washed by the gas flow.
This is still a promising method, which is not used in the mass practice of using gas turbine engines due to difficulties with the selection of porous material and the high probability of fairly rapid clogging of the pores. However, if these problems are solved, the presumably possible gas temperature with this type of cooling can reach 1650°C.
The turbine disks and CA casings are also cooled by air due to the various stages of the compressor as it passes through the internal cavities of the engine, washing the cooled parts and then releasing them into the flow part.
Due to the fairly high degree of pressure increase in the compressors of modern engines, the cooling air itself can have a fairly high temperature. Therefore, to increase the cooling efficiency, measures are taken to first reduce this temperature.
To do this, before supplying the turbine to the blades and disks, air can be passed through special profile grids, similar to the SA turbine, where the air is twisted in the direction of rotation of the impeller, expanding and cooling at the same time. The cooling amount can be 90-160°.
For the same cooling, air-to-air radiators cooled by secondary circuit air can be used. On the AL-31F engine, such a radiator reduces the temperature to 220° in flight and 150° on the ground.
For cooling needs aviation turbine A fairly large amount of air is taken from the compressor. On various engines - up to 15-20%. This significantly increases losses, which are taken into account in the thermogasdynamic calculation of the engine. Some engines are equipped with systems that reduce the supply of air for cooling (or even shut it off) at reduced engine operating conditions, which has a positive effect on efficiency.
Cooling diagram of the 1st stage of the NK-56 turbofan turbine. Honeycomb seals and a cooling shut-off tape at reduced engine operating conditions are also shown.
When assessing the efficiency of a cooling system, additional hydraulic losses on the blades due to changes in their shape when cooling air is released are usually taken into account. The efficiency of a real cooled turbine is approximately 3-4% lower than that of an uncooled one.
Something about making blades...
On first generation jet engines, turbine blades were mainly made stamping method followed by long-term processing. However, in the 50s, VIAM specialists convincingly proved that it was casting and not wrought alloys that offered the prospect of increasing the level of heat resistance of blades. Gradually, a transition to this new direction was made (including in the West).
Currently, the production uses precision non-waste casting technology, which makes it possible to produce blades with specially profiled internal cavities that are used to operate the cooling system (the so-called technology lost wax casting).
This is, in fact, the only way to obtain cooled blades now. It also improved over time. At the first stages of casting technology, blades with different sizes were produced grains of crystallization, which did not adhere securely to each other, which significantly reduced the strength and service life of the product.
Subsequently, using special modifiers, they began to produce cast cooled blades with homogeneous, equiaxed, fine structural grains. For this purpose, in the 60s, VIAM developed the first serial domestic heat-resistant alloys for casting ZhS6, ZhS6K, ZhS6U, VZHL12U.
Their operating temperature was 200° higher than that of the then common deformable (stamping) alloy EI437A/B (KhN77TYu/YUR). Blades made from these materials worked for at least 500 hours without visually visible signs of destruction. This type of manufacturing technology is still used today. Nevertheless, the grain boundaries remain the weak point of the blade structure, and it is along them that its destruction begins.
Therefore, with the increase in load characteristics of modern aircraft turbines(pressure, temperature, centrifugal loads) there was a need to develop new technologies for manufacturing blades, because the multi-grain structure in many respects no longer satisfied the severe operating conditions.
Examples of the structure of heat-resistant material of working blades. 1 - equiaxed grain size, 2 - directional crystallization, 3 - single crystal.
This is how “ directional crystallization method" With this method, in the solidifying casting of a blade, not individual equiaxed grains of metal are formed, but long columnar crystals elongated strictly along the axis of the blade. This kind of structure significantly increases the blade's fracture resistance. It is similar to a broom, which is very difficult to break, although each of the twigs that make it up breaks without problems.
This technology was subsequently refined to an even more advanced " monocrystal casting method", when one blade is practically one whole crystal. This type of blade is now also installed in modern aircraft turbines. For their manufacture, special alloys are used, including so-called rhenium-containing alloys.
In the 70s and 80s, VIAM developed alloys for casting turbine blades with directional solidification: ZhS26, ZhS30, ZhS32, ZhS36, ZhS40, VKLS-20, VKLS-20R; and in the 90s - long-life corrosion-resistant alloys: ZhSKS1 and ZhSKS2.
Further, working in this direction, from the beginning of 2000 to the present, VIAM has created high-rhenium heat-resistant alloys of the third generation: VZhM1 (9.3% Re), VZhM2 (12% Re), ZhS55 (9% Re) and VZhM5 (4% Re ). To further improve the characteristics, experimental studies have been carried out over the past 10 years, which resulted in rhenium-ruthenium-containing alloys of the fourth - VZhM4 and fifth generations VZhM6.
As assistants...
As mentioned earlier, only reactive (or active-reactive) turbines are used in gas turbine engines. However, in conclusion it is worth remembering that among those used aircraft turbines There are also active ones. They mainly perform secondary tasks and do not take part in the operation of the propulsion engines.
And yet their role is often very important. In this case we are talking about air starters used to launch. There are various types of starter devices used to spin up the rotors of gas turbine engines. The air starter occupies perhaps the most prominent place among them.
Air starter of a turbofan engine.
This unit, in fact, despite the importance of its functions, is fundamentally quite simple. The main unit here is a one- or two-stage active turbine, which rotates the engine rotor (in a turbofan engine, usually a low-pressure rotor) through a gearbox and drive box.
Location of the air starter and its working line on the turbofan engine,
The turbine itself is spun by a flow of air coming from a ground source, either an on-board APU, or from another already running aircraft engine. At a certain point in the starting cycle, the starter is automatically turned off.
In units of this kind, depending on the required output parameters, they can also be used. radial turbines. They can also be used in air conditioning systems in aircraft cabins as an element of a turbo-refrigerator, in which the effect of expansion and reduction of air temperature on the turbine is used to cool the air entering the cabins.
In addition, both active axial and radial turbines are used in turbocharging systems of piston aircraft engines. This practice began even before the turbine was converted into the most important node GTD continues to this day.
An example of the use of radial and axial turbines in auxiliary devices.
Similar systems using turbochargers are used in cars and in general various systems compressed air supply.
Thus, the aircraft turbine also serves people well in an auxiliary sense.
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Well, that's probably all for today. In fact, there is still a lot that can be written about here, both in terms of additional information and in terms of more full description what has already been said. The topic is very broad. However, one cannot embrace the immensity :-). For general information, perhaps, it is enough. Thank you for reading to the end.
Until next time...
Finally, there are pictures that “do not fit” into the text.
An example of a single-stage turbojet turbine.
Model of Heron's aeolipile in the Kaluga Museum of Cosmonautics.
Articulation of the video probe of the Vucam XO endoscope.
Screen of the Vucam XO multifunctional endoscope.
Endoscope Vucam XO.
An example of a thermal protective coating on the CA blades of a GP7200 engine.
Honeycomb plates used for seals.
Possible options for labyrinth seal elements.
Labyrinth honeycomb seal.
