JP2007255224A - Turbine blade and gas turbine - Google Patents

Abstract

<P>PROBLEM TO BE SOLVED: To provide a turbine blade and a gas turbine, capable of improving the performance and enhancing the durability by effectively reducing thermal stress. <P>SOLUTION: The gas turbine comprises a compressor 11, a combustor 12, a turbine 13 and an exhaust chamber 14. A stationary blade 21 of the turbine 13 is divided in a blade section 31, an outer shroud 32 and an inner shroud 33. The blade section 31 having a cooling passage penetrates through a through hole 32b of the outer shroud 32. The blade section 31 and the outer shroud 32 are connected to each other by fastening bolts 38 at positions where flange parts 31a, 32c thereof are in contact with each other. The blade section 31 also penetrates through a through hole 33b of the inner shroud 33 and is fastened to the inner shroud 33 by fastening bolts 40. <P>COPYRIGHT: (C)2008,JPO&INPIT

Description

  The present invention relates to a gas turbine that supplies fuel to compressed compressed air and burns it, supplies the generated combustion gas to a turbine to obtain rotational power, and a turbine blade used in the gas turbine.

  The gas turbine is composed of a compressor, a combustor, and a turbine, and the air taken in from the air intake port is compressed by the compressor to become high-temperature / high-pressure compressed air. The fuel is supplied and burned, and the high-temperature and high-pressure combustion gas drives the turbine, and the generator connected to the turbine through the rotor is driven. In this case, the turbine is configured by alternately arranging a plurality of stationary blades and moving blades in the vehicle interior, and rotationally drives an output shaft connected to the generator by driving the moving blades with combustion gas. ing. The combustion gas that has driven the turbine is released to the atmosphere after the dynamic pressure is converted to static pressure by the diffuser in the exhaust casing.

  In such a gas turbine, a stationary blade as a turbine blade has a blade portion, an outer shroud, and an inner shroud integrally manufactured by precision casting. In order to increase the output and temperature of the gas turbine and to ensure high reliability, it is desirable that the stationary blades have a single crystal structure using a high performance material, and a cooling structure corresponding to a high heat load is provided. Need to apply. For this purpose, it is conceivable that the blade portion, the outer shroud, and the inner shroud constituting the stationary blade are divided and manufactured by precision casting and then joined.

  Examples of such a stationary blade of a gas turbine include those described in Patent Documents 1 and 2 below. In the turbine nozzle described in Patent Document 1, a hollow nozzle vane, an outer band, and an inner band are manufactured by die casting, and one end portion and an outer band of the nozzle vane, the other end portion of the nozzle vane, and an inner band. Are manufactured by brazing and joining. Moreover, the turbine blade described in Patent Document 2 is made of ceramics for the turbine blade portion, the outer shroud, and the inner shroud, and the turbine blade portion is interposed between the outer shroud and the inner shroud and fastened with a stay bolt. Thus, the turbine blades are clamped by the outer shroud and the inner shroud.

JP 2002-138802 A Japanese Examined Patent Publication No. 04-025404

  However, in the turbine nozzle of Patent Document 1 described above, each end of the nozzle vane is joined to the outer band and the inner band by brazing, and in this case, the temperature difference under the high heat load condition of the gas turbine. And thermal stress due to thermal expansion becomes excessive, and sufficient reliability may not be ensured. That is, when the combustion gas flows through the gas passage, the nozzle vane, the outer band, and the inner band are heated on the front side, and the outer band and the inner band are curved due to the difference in thermal expansion between the front side and the rear side. There is a possibility that cracks and breakage may occur due to excessive stress acting on the joint with the nozzle vane.

  Further, in the turbine blade described in Patent Document 2, a turbine blade portion is interposed between a ceramic outer shroud and an inner shroud and fastened with a stay bolt. However, the use of ceramics as a material leads to an increase in cost, and a sufficient cooling passage cannot be secured in the turbine blade portion, resulting in a decrease in cooling performance.

  The present invention solves the above-described problems, and an object of the present invention is to provide a turbine blade and a gas turbine that are improved in durability and are improved in durability by efficiently relieving thermal stress.

  In order to achieve the above object, a turbine blade of the invention of claim 1 includes a blade portion having a cooling passage therein, an outer shroud connected to each longitudinal end of the blade portion, and an inner shroud. In the above, at least one end portion of the blade portion passes through the outer shroud or the inner shroud and is bolted outside the combustion gas passage.

  In the turbine blade of the invention according to claim 2, a flange portion is formed at a base end portion of the blade portion, the blade portion passes through either the outer shroud or the inner shroud, and the flange portion abuts. It is characterized by being bolted in place.

  In the turbine blade of the invention of claim 3, the blade portion passes through the outer shroud and the inner shroud, and is bolted to either the outer shroud or the inner shroud at a position where the flange portion abuts. The tip portion is bolted to the other side.

  In the turbine blade according to claim 4, one end portion of the blade portion passes through either the outer shroud or the inner shroud and is bolted, and the other end portion of the blade portion is connected to the other by welding. It is characterized by that.

  In the turbine blade of the fifth aspect of the invention, one end portion of the blade portion passes through either the outer shroud or the inner shroud and is bolted, and the other end portion of the blade portion has a seal member on the other. It is characterized by being connected through the network.

  In the turbine blade according to the invention of claim 6, a plurality of the blade portions are provided, and at least one end portion of the plurality of blade portions penetrates the outer shroud or the inner shroud and is bolted outside the combustion gas passage. It is characterized by.

  The turbine blade according to the invention of claim 7 is characterized in that a moving blade shroud adjacent to the outer shroud is integrally provided.

  A gas turbine according to an eighth aspect of the present invention is a gas turbine in which fuel is supplied to a compressed air compressed by a compressor and burned by a combustor, and rotational power is obtained by supplying the generated combustion gas to the turbine. The stationary blade constituting the turbine is characterized in that at least one end portion of a blade portion having a cooling passage passes through the outer shroud or the inner shroud and is bolted outside the combustion gas passage.

  According to the turbine blade of the first aspect of the present invention, since at least one end portion of the blade portion having the cooling passage therein passes through the outer shroud or the inner shroud and is bolted outside the combustion gas passage, the blade portion and the outer shroud And the inner shroud can be divided into a single crystal structure using a high-performance material, and a cooling structure corresponding to a high heat load can be applied to improve performance. In addition, one end of the wing and the outer shroud or inner shroud are bolted outside the combustion gas passage, and thermal stress can be efficiently relieved by sliding at the fastening part, improving durability Can do.

  According to the turbine blade of the second aspect of the present invention, the flange portion is formed at the base end portion of the blade portion, the blade portion passes through either the outer shroud or the inner shroud, and the bolt is positioned at the position where the flange portion abuts. Since it is fastened, when the wing part is assembled to the outer shroud or the inner shroud, positioning and temporary assembly are possible by abutting the flange part, and bolt fastening can be performed properly.

  According to the turbine blade of the third aspect of the present invention, the blade portion passes through the outer shroud and the inner shroud, is bolted to either the outer shroud or the inner shroud at the position where the flange portion abuts, and the tip portion is the other Since the base end portion and the tip end portion of the wing portion are bolted to the outer shroud and the inner shroud, thermal stress can be reliably relieved by sliding at the fastening portion.

  According to the turbine blade of the invention of claim 4, since one end of the blade is passed through either the outer shroud or the inner shroud and bolted, and the other end of the blade is connected to the other by welding. By bolting one end of the wing to the outer shroud or the inner shroud, thermal stress can be effectively relieved by sliding at the fastening portion, and the other end of the wing is connected to the outer or inner shroud. By being connected to each other by welding, leakage of combustion gas can be prevented and performance can be improved.

  According to the turbine blade of the fifth aspect of the invention, one end of the blade is passed through either the outer shroud or the inner shroud and bolted, and the other end of the blade is connected to the other via the seal member. Therefore, by bolting one end of the wing to the outer shroud or the inner shroud, thermal stress can be effectively relieved by sliding at the fastening portion, and the other end of the wing is also connected to the outer shroud. Alternatively, by connecting to the inner shroud via a seal member, it is possible to prevent leakage of combustion gas and improve performance.

  According to the turbine blade of the sixth aspect of the invention, a plurality of blade portions are provided, and at least one end portion of the plurality of blade portions penetrates the outer shroud or the inner shroud and is bolted outside the combustion gas passage. The blade portion can be supported by the outer shroud and the inner shroud, and by reducing the number of connecting portions between the shrouds, the leakage of combustion gas can be prevented and the performance can be improved.

  According to the turbine blade of the seventh aspect of the present invention, since the blade shroud adjacent to the outer shroud is provided integrally, the outer shroud and the blade shroud are integrated, so that the connection portion between the shrouds is eliminated, and the combustion gas The leakage can be prevented and the performance can be improved.

  According to the gas turbine of the eighth aspect of the present invention, the fuel is supplied to the compressed air compressed by the compressor and burned by the combustor, and the generated combustion gas is supplied to the turbine to obtain the rotational power. The stationary vane constituting the turbine is configured such that at least one end portion of the blade portion having the cooling passage passes through the outer shroud or the inner shroud and is bolted outside the combustion gas passage, so that the blade portion, the outer shroud, and the inner shroud are formed. By using a split structure, a high-performance material can be used to form a single crystal structure, and a cooling structure that can handle high heat loads can be applied to increase the output and temperature of the gas turbine. By ensuring high reliability, performance can be improved, and one end of the wing and the outer shroud or inner shroud are bolted outside the combustion gas passage. Cage, it is possible to effectively relax the thermal stress by sliding in the fastening portion, it is possible to improve the durability.

  Exemplary embodiments of a turbine blade and a gas turbine according to the present invention will be described below in detail with reference to the accompanying drawings. In addition, this invention is not limited by this Example.

  FIG. 1 is a cross-sectional view of a stationary blade as a turbine blade according to a first embodiment of the present invention, FIG. 2 is a perspective view of the stationary blade according to the first embodiment, and FIG. 3 is a stationary blade according to the first embodiment. FIG. 4 is a schematic configuration diagram of the gas turbine according to the first embodiment, and FIG. 5 is a schematic diagram illustrating a turbine blade arrangement in the gas turbine according to the first embodiment.

  As shown in FIG. 4, the gas turbine according to the first embodiment includes a compressor 11, a combustor 12, a turbine 13, and an exhaust chamber 14. The turbine 13 is not illustrated via a rotor (turbine shaft) 24 described later. A generator is connected. The compressor 11 has an air intake port 15 for taking in air, a plurality of stationary blades 17 and moving blades 18 are alternately arranged in a compressor casing 16, and a bleed manifold 19 is provided on the outside thereof. ing. The combustor 12 is combustible by supplying fuel to the compressed air compressed by the compressor 11 and igniting it with a burner. In the turbine 13, a plurality of stationary blades 21 and moving blades 22 are alternately arranged in a turbine casing 20. The exhaust chamber 14 has an exhaust diffuser 23 that is continuous with the turbine 13. Further, the rotor 24 is positioned so as to pass through the central portion of the compressor 11, the combustor 12, the turbine 13, and the exhaust chamber 14, and the end portion on the compressor 11 side is rotatably supported by the bearing portion 25. On the other hand, the end portion on the exhaust chamber 14 side is rotatably supported by the bearing portion 26. A plurality of disk plates are fixed to the rotor 24, the rotor blades 18 and 22 are connected, and a drive shaft of a generator (not shown) is connected to an end portion on the compressor 11 side.

  Therefore, the air taken in from the air intake port 15 of the compressor 11 passes through the plurality of stationary blades 21 and the moving blades 22 and is compressed to become high-temperature and high-pressure compressed air. Combustion occurs when a predetermined fuel is supplied to the compressed air. The high-temperature and high-pressure combustion gas that is the working fluid generated in the combustor 12 passes through the plurality of stationary blades 21 and the moving blades 22 constituting the turbine 13 to drive and rotate the rotor 24. While the generator connected to 24 is driven, the exhaust gas is discharged into the atmosphere after the dynamic pressure is converted to static pressure by the exhaust diffuser 23 in the exhaust chamber 14.

  In the turbine 13 described above, as shown in FIG. 5, a plurality of stationary blades 21 and moving blades 22 are alternately arranged in a turbine casing 20 (chamber casing). The stationary blade 21 is configured by connecting an outer shroud 32 and an inner shroud 33 to each end in the longitudinal direction of the blade portion 31, and the turbine blade 20 is assembled in a state where the plurality of stationary blades 21 are assembled in an annular shape. It is supported. The rotor blade 22 is configured to be fixed to an outer peripheral portion of a rotor disk 34 integrally coupled to the rotor 24 (see FIG. 4) via an inner shroud 35, and is outside the turbine casing 20 on the outer peripheral side with a predetermined gap. A shroud 36 is fixed in an annular shape. A passage in which a plurality of stationary blades 21 and moving blades 22 are alternately arranged is a combustion gas passage 37.

  In this embodiment, the stationary blade 21 as a turbine blade is divided into a blade portion 31, an outer shroud 32, and an inner shroud 33, and each end portion of the blade portion 31, the outer shroud 32, and the inner shroud 33 are burned. Bolts are fastened outside the gas passage 37.

  That is, as shown in FIGS. 1 to 3, the wing part 31 has a hollow shape so that a cooling passage for supplying cooling air to the inside can be formed, and a flange part 31a is formed at the base end part. A plurality of fastening holes 31b are formed in the flange portion 31a, and a plurality of fastening holes 31c are formed in the tip portion.

  The outer shroud 32 is formed with a through hole 32b by forming an annular folded portion 32a at the center, and a flange portion 32c is formed in the folded portion 32a, and a plurality of fastening holes 32d are formed in the flange portion 32c. Is formed. Further, the inner shroud 33 has a through hole 33b formed by forming an annular folded portion 33a at the center, and a plurality of fastening holes 33c are formed in the folded portion 32a.

  In order to assemble the wing part 31, the outer shroud 32, and the inner shroud 33, first, the wing part 31 is passed through the through hole 32b of the outer shroud 32, and the flange part 31a is in contact with the flange part 32c. Fastening bolts 38 are inserted into the fastening holes 31b and 32d at the positions where the fastening holes 31b and 32d are fitted, and fastened using the nut 39. Next, the wing portion 31 is passed through the through hole 33 b of the inner shroud 33, and the fastening bolt 40 is inserted into the fastening holes 31 c and 33 c at a position where the fastening holes 31 c and 33 c are fitted, and fastened using the nut 41. Thus, the stationary blade 21 is manufactured.

  Accordingly, the combustion gas flows through the combustion gas passage 37, the blade 31 of the stationary blade 21, the outer shroud 32, and the inner shroud 33 are heated, and the outer shroud 32 and the inner shroud 33 are heated on the front and back sides. Due to the expansion difference, as shown by a two-dot chain line in FIG. However, the connecting portion between the blade portion 31 and the outer shroud 32 and the connecting portion between the blade portion 31 and the inner shroud 33 are located outside the combustion gas passage 37 and fastened by fastening bolts 38 and 40. The stress acting on each connecting portion is absorbed by the slippage at, so that the occurrence of cracks and breakage is prevented.

  In addition, since the blade portion 31 of the stationary blade 21 has a hollow shape and is fastened outside the combustion gas passage 37 with the outer shroud 32 and the inner shroud 33, a cooling passage can be easily formed in the blade portion 31. A partition wall can be inserted, and sufficient cooling performance can be secured for the stationary blade 21.

  Thus, in the stationary blade and gas turbine of Example 1, the compressor 11, the combustor 12, the turbine 13, and the exhaust chamber 14 constitute a gas turbine, and the stationary blade 21 of the turbine 13 is replaced with the blade portion 31. The outer shroud 32 and the inner shroud 33 are divided, and the blade 31 having a cooling passage passes through the through hole 32b of the outer shroud 32 and is connected by the fastening bolt 38 at a position where the flange portions 31a and 32c are in contact with each other. At the same time, the wing portion 31 passes through the through hole 33 b of the inner shroud 33 and is fastened by the fastening bolt 40.

  Therefore, by making the wing portion 31, the outer shroud 32, and the inner shroud 33 into a divided structure, a high-performance material can be used for a single crystal structure, and a cooling structure corresponding to a high heat load can be applied. The performance can be improved. Further, the connecting portion of the blade portion 31, the outer shroud 32, and the inner shroud 33 is fastened by bolts 38, 40 outside the combustion gas passage 37, and the thermal stress can be effectively relieved by sliding at the fastening portion. And durability can be improved.

  In addition, a flange portion 31a is formed at the base end portion of the wing portion 31, a flange portion 32c is formed in the outer shroud 32, the wing portion 31 is passed through the outer shroud 32, and the flange portions 31a and 32c are in contact with each other. The bolt is fastened in position. Therefore, when the wing portion 31 is assembled to the outer shroud 32, the flange portions 31a and 32c can be brought into contact with each other so that the assembly position of both can be positioned, and temporary assembly is possible. The bolt can be fastened.

  FIG. 6 is a cross-sectional view of a stationary blade as a turbine blade according to a second embodiment of the present invention.

  In the second embodiment, as shown in FIG. 6, a stationary blade 51 as a turbine blade is a divided structure of a blade portion 52, an outer shroud 53, and an inner shroud 54, and includes one end portion of the blade portion 52 and the outer shroud 53. Bolts are fastened outside the combustion gas passage 55, and the other end of the blade 52 and the inner shroud 54 are connected by welding.

  That is, the wing portion 52 has a hollow shape so that a cooling passage for supplying cooling air can be formed therein, and a flange portion 52a is formed at the base end portion, and a plurality of fastening holes 52b are formed in the flange portion 52a. Is formed. The outer shroud 53 is formed with a through hole 53b by forming an annular folded portion 53a at the center, and a flange portion 53c is formed in the folded portion 53a, and a plurality of fastening holes 53d are formed in the flange portion 53c. Is formed. Further, the inner shroud 54 has a through hole 54b formed by forming an annular folded portion 54a at the center, and a stepped portion 54c is formed in the folded portion 54a.

  Then, the fastening bolt 56 is inserted into the fastening hole 52b at a position where the fastening holes 52b and 53d are fitted in a state where the wing part 52 is passed through the through hole 53b of the outer shroud 53 and the flange part 52a is in contact with the flange part 53c. , 53d and fastened using a nut 57. Further, the stationary blade 51 is manufactured by passing the blade portion 52 through the through hole 54b of the inner shroud 54 and joining the tip portion with the welding portion 58 in contact with the stepped portion 54c.

  Accordingly, the combustion gas flows through the combustion gas passage 55, and the blade 52 of the stationary blade 51, the outer shroud 53, and the surface side of the inner shroud 54 are heated, and the outer shroud 53 and the inner shroud 54 are heated on the front side and the rear side. It deforms into a curved shape due to the expansion difference. However, since the connecting portion between the base end portion of the blade portion 52 and the outer shroud 53 is positioned outside the combustion gas passage 55 and fastened by the fastening bolt 56, the stress acting on the connecting portion due to the slip here is generated. It will be absorbed and cracking and breakage will be prevented.

  Moreover, since the front-end | tip part of the wing | blade part 52 and the inner shroud 54 are joined by welding, sufficient intensity | strength can be ensured and the leakage of the combustion gas from the combustion gas channel | path 55 is prevented.

  As described above, in the stationary blade and the gas turbine according to the second embodiment, the stationary blade 51 is divided into the blade portion 52, the outer shroud 53, and the inner shroud 54, and the blade portion 52 having the cooling passage is the outer shroud 53. It penetrates through the through-hole 53b and is connected by a fastening bolt 56 at a position where the flanges 52a and 53c are in contact with each other, and the wing 52 penetrates through the through-hole 54b of the inner shroud 54, and the leading end is connected to the step 54c. They are connected by welding 58.

  Therefore, by making the blade 52, the outer shroud 53, and the inner shroud 54 into a divided structure, a single crystal structure can be formed using a high-performance material, and a cooling structure corresponding to a high heat load can be applied. The performance can be improved. Further, the connecting portion between the base end portion of the wing portion 52 and the outer shroud 53 is fastened by a bolt 56 outside the combustion gas passage 55, and the thermal stress can be efficiently relieved by sliding at the fastening portion, Durability can be improved. Further, the tip of the blade portion 52 is joined to the step portion 54c of the inner shroud 54 by welding, so that sufficient strength can be secured and leakage of combustion gas from the combustion gas passage 55 can be prevented. it can.

  7 is a schematic diagram showing a turbine blade arrangement in a gas turbine according to a third embodiment of the present invention. FIG. 8 is a cross-sectional view of a main part of a stationary blade as a turbine blade of the third embodiment. FIGS. -2 is a schematic diagram illustrating a mounting structure of a seal member in the stationary blade of the third embodiment. In addition, the same code | symbol is attached | subjected to the member which has the same function as what was demonstrated in the Example mentioned above, and the overlapping description is abbreviate | omitted.

  In the gas turbine of the third embodiment, as shown in FIG. 7, a plurality of stationary blades 61 and moving blades 22 are alternately arranged in a turbine casing 20 (chamber casing). The stationary blade 61 is configured by connecting an outer shroud 63 and an inner shroud 64 to each end portion in the longitudinal direction of the blade portion 62, and in the turbine casing 20 with the plurality of stationary blades 61 assembled in an annular shape. It is supported. The rotor blade 22 is configured to be fixed to an outer peripheral portion of a rotor disk 34 integrally coupled with a rotor (not shown) via an inner shroud 35, and an outer shroud 36 with a predetermined gap in the turbine casing 20 on the outer peripheral side. Is fixed in an annular shape. A passage in which a plurality of stationary blades 61 and moving blades 22 are alternately arranged is a combustion gas passage 37.

  In this embodiment, the stationary blade 61 as a turbine blade is divided into a blade portion 62, an outer shroud 63, and an inner shroud 64, and one end portion of the blade portion 62 and the outer shroud 63 are arranged outside the combustion gas passage 37. The other end of the wing 62 and the inner shroud 64 are connected via a seal member.

  That is, as shown in FIGS. 7 and 8, the wing portion 62 has a hollow shape so that a cooling passage for supplying cooling air to the inside can be formed, a flange portion 62a is formed at the proximal end portion, and the distal end portion is formed. An annular groove 62b is formed. The outer shroud 63 is formed with an annular folded portion 63a and a flange portion 63b at the center, and a fitting portion 63c is formed on the upstream side in the flow direction of the combustion gas, while an engaging portion 63d is provided on the downstream side. Is formed. In addition, the inner shroud 64 has a through hole 64a formed in the central portion, annular grooves 64b and 64c are formed in the through hole 64a, and a fitting portion 64d is formed on the upstream side in the combustion gas flow direction. On the other hand, a mounting portion 64e is formed on the downstream side.

  Then, the wing portion 62 is passed through the outer shroud 63 and is fastened by using the fastening bolt 65 in a state where the flange portion 62a is in contact with the flange portion 63b. Further, the wing portion 62 is passed through the through hole 64a of the inner shroud 64, and as shown in FIGS. 9-1 and 9-2, half-seal seal members 66a and 66b are attached to the groove portion 62b of the tip portion. As shown in FIG. 8, in a state where the seal members 66a and 66b are accommodated in the groove portion 64b of the inner shroud 64, the locking member 67 is fitted into the groove portion 64c and fastened by the fastening bolt 68, whereby the stationary blade 61 is Manufactured. In this case, a predetermined gap is secured between the groove portion 64 b of the inner shroud 64, the seal members 66 a and 66 b, and the locking member 67.

  Thereafter, as shown in FIG. 8, the engaging portion 63d of the outer shroud 63 in the stationary blade 61 is engaged with the recess 20a of the turbine casing 20, and the fitting portions 63c and 64d of the shrouds 63 and 64 are sealed members. The stationary blades 61 are assembled at predetermined positions by closely contacting with each other via the joints 69 and 70 and fastening the connecting portion 64e of the inner shroud 64 to the connecting portion 20b of the turbine casing 20.

  Accordingly, the combustion gas flows through the combustion gas passage 37, the blade 62 of the stationary blade 61, the outer shroud 63, and the inner shroud 64 are heated, and the outer shroud 63 and the inner shroud 64 are heated on the front and back sides. It deforms into a curved shape due to the expansion difference. However, since the connecting portion between the base end portion of the blade portion 62 and the outer shroud 63 is located outside the combustion gas passage 37 and fastened by the fastening bolt 65, the stress acting on the connecting portion due to the slip here is generated. It will be absorbed and cracking and breakage will be prevented.

  Further, since the tip end portion of the wing portion 62 and the inner shroud 64 are connected with a predetermined gap through the seal members 66a and 66b, the stress acting on the connecting portion between the wing portion 62 and the inner shroud 64 due to this gap. As a result, the occurrence of cracks and breakage is prevented, and the leakage of the combustion gas from the combustion gas passage 55 is prevented by the seal members 66a and 66.

  As described above, in the stationary blade and the gas turbine of the third embodiment, the stationary blade 61 is divided into the blade portion 62, the outer shroud 63, and the inner shroud 64, and the blade portion 62 having the cooling passage is formed into the outer shroud 63. It penetrates and is connected by the fastening bolt 65 at the position where the flange portions 62a and 63b are in contact with each other, and the tip end portion of the wing portion 62 penetrates the through-hole 64a of the inner shroud 64, and passes through the seal members 66a and 66b. It is configured to be connected with a predetermined gap.

  Therefore, by making the blade portion 62, the outer shroud 63, and the inner shroud 64 into a divided structure, a high-performance material can be used for a single crystal structure, and a cooling structure corresponding to a high heat load can be applied. The performance can be improved. Further, the base end portion of the blade portion 62 and the connecting portion of the outer shroud 63 are fastened by the bolt 65 outside the combustion gas passage 37, while the tip portion and the inner shroud 64 are connected with the seal members 66a and 66b and a predetermined gap. In addition, the thermal stress can be efficiently relieved by the sliding and movement of the blade portion 62 at the connecting portion, and the durability can be improved. The tip of the blade portion 62 is connected to the inner shroud 64 via the seal members 66a and 66b, and the leakage of the combustion gas from the combustion gas passage 37 can be prevented.

  FIG. 10 is a cross-sectional view of a stationary blade as a turbine blade according to a fourth embodiment of the present invention.

  In the fourth embodiment, as shown in FIG. 10, a stationary blade 71 as a turbine blade is a divided structure of a blade portion 72, an outer shroud 73, and an inner shroud 74, and includes one end portion of the blade portion 72 and the outer shroud 73. Bolts are fastened outside the combustion gas passage 75, and the other end of the blade 72 and the inner shroud 74 are connected via a labyrinth seal.

  That is, the wing portion 72 has a hollow shape so that a cooling passage for supplying cooling air to the inside can be formed, and a flange portion 72a is formed at the base end portion, and a plurality of fastening holes 72b are formed in the flange portion 72a. Is formed. The outer shroud 73 is formed with a through hole 73b by forming an annular folded portion 73a in the center, and a flange portion 73c is formed in the folded portion 73a, and a plurality of fastening holes 73d are formed in the flange portion 73c. Is formed. Further, the inner shroud 74 has a through hole 74a at the center. And while the convex part 72c is formed in the front-end | tip part of the wing | blade part 72, the recessed part 74b is formed in the through-hole 74a of the inner shroud 74, and this convex part 72c fits into the recessed part 74b with a predetermined clearance gap. Here, a labyrinth seal 76 is formed.

  Then, the fastening bolt 77 is inserted into the fastening hole 72b at a position where the fastening holes 72b and 73d are fitted in a state where the wing part 72 passes through the through-hole 73b of the outer shroud 53 and the flange part 72a contacts the flange part 73c. , 73d and fastened with a nut 78. Further, the labyrinth seal 76 is formed by fitting the blade portion 72 through the through hole 74a of the inner shroud 74 and fitting the convex portion 72c into the concave portion 74b with a predetermined gap, and the stationary blade 71 is manufactured. In this case, the outer shroud 53 and the inner shroud 74 are supported by a casing casing (not shown), so that a gap between the labyrinth seals 76 is secured.

  Accordingly, the combustion gas flows through the combustion gas passage 75, the blade 72 of the stationary blade 71, the outer shroud 73, and the inner shroud 74 are heated, and the outer shroud 73 and the inner shroud 74 are heated on the front and back sides. It deforms into a curved shape due to the expansion difference. However, since the connecting portion between the base end portion of the blade portion 72 and the outer shroud 73 is located outside the combustion gas passage 75 and fastened by the fastening bolt 77, the stress acting on the connecting portion due to the slip here is generated. It will be absorbed and cracking and breakage will be prevented.

  Further, since the tip end portion of the wing portion 72 and the inner shroud 74 are connected via the labyrinth seal 76, the stress acting on the connecting portion is absorbed by the gap of the labyrinth seal 76, and cracks and breakage occur. The labyrinth seal 76 prevents combustion gas from leaking from the combustion gas passage 75.

  As described above, in the stationary blade and the gas turbine according to the fourth embodiment, the stationary blade 71 is divided into the blade portion 72, the outer shroud 73, and the inner shroud 74, and the blade portion 72 having the cooling passage is the outer shroud 73. It penetrates through the through-hole 73b and is connected by a fastening bolt 77 at a position where the flanges 72a and 73c are in contact with each other, and the wing part 72 penetrates the through-hole 74a of the inner shroud 74 and the convex part 72c becomes the concave part 74b. The labyrinth seal 76 is formed by fitting with a predetermined gap.

  Therefore, by making the wing portion 72, the outer shroud 73, and the inner shroud 74 into a divided structure, a high-performance material can be used for a single crystal structure, and a cooling structure corresponding to a high heat load can be applied. The performance can be improved. In addition, the base end portion of the blade portion 72 and the connecting portion of the outer shroud 73 are fastened by a bolt 76 outside the combustion gas passage 75, and the tip portion of the blade portion 72 and the inner shroud 74 are connected via a labyrinth seal 76. In addition, the thermal stress can be efficiently relieved by sliding at the fastening portion, and the durability can be improved. Further, the labyrinth seal 76 can prevent the combustion gas from leaking from the combustion gas passage 55.

  FIG. 11 is a perspective view of a stationary blade as a turbine blade according to the fifth embodiment of the present invention.

  In the fifth embodiment, as shown in FIG. 11, a stationary blade 81 as a turbine blade is a divided structure of two blade portions 82, an outer shroud 83, and an inner shroud 84, and each end portion of the two blade portions 82. The outer shroud 83 and the inner shroud 84 are bolted to the outside of the combustion gas passage 85.

  That is, in this embodiment, two wing portions 82 are supported by a pair of outer shroud 83 and inner shroud 84, and each wing portion 82 penetrates through outer shroud 83 and abuts on each flange portion 82a, 83a. In addition to being fastened by fastening bolts 86 at the positions, each wing 72 passes through the inner shroud 74 and the tip is fastened to the inner shroud 84 by fastening bolts 87. In this case, three or more wing parts 82 may be supported by the outer shroud 83 and the inner shroud 84.

  Accordingly, the combustion gas flows through the combustion gas passage 85, and the surface sides of the two blade portions 82, the outer shroud 83, and the inner shroud 84 are heated, and the outer shroud 83 and the inner shroud 84 are subjected to a difference in thermal expansion between the front surface side and the rear surface side. To deform into a curved shape. However, since the connection part of the base end part and the front-end | tip part of each wing | blade part 82 and each shroud 83 and 84 is located in the outer side of the combustion gas channel | path 85, and is fastened by the fastening bolts 86 and 87, it is slippery here. As a result, stress acting on the connecting portion is absorbed, and cracks and breakage are prevented.

  As described above, in the stationary blade and the gas turbine according to the fifth embodiment, the stationary blade 81 is divided into the two blade portions 82, the outer shroud 83, and the inner shroud 84, and each blade portion 82 having a cooling passage is used. Are fastened by fastening bolts 86 and 87 outside the combustion gas passage 85.

  Therefore, by making the two wings 82, the outer shroud 83, and the inner shroud 84 into a divided structure, a single crystal structure can be formed using a high-performance material, and a cooling structure corresponding to a high heat load is applied. And the performance can be improved. Further, the end portions of the two wing portions 82 and the connecting portions of the outer shroud 83 and the inner shroud 84 are fastened by bolts 86 and 87 outside the combustion gas passage 85, and the thermal stress is made efficient by slipping at the fastening portion. Can be relaxed and the durability can be improved. Further, two wing portions 82 are supported by a pair of outer shroud 83 and inner shroud 84, and by reducing the number of annularly connecting each shroud 83, 84, the connecting portion between each shroud 83, 84 is provided. This reduces the leakage of the combustion gas from the combustion gas passage 55.

  FIG. 12 is a schematic diagram illustrating a turbine blade arrangement in the gas turbine according to the sixth embodiment of the present invention. In addition, the same code | symbol is attached | subjected to the member which has the same function as what was demonstrated in the Example mentioned above, and the overlapping description is abbreviate | omitted.

  In the gas turbine of the sixth embodiment, as shown in FIG. 12, a plurality of stationary blades 91 and moving blades 22 are alternately arranged in a turbine casing 20 (chamber casing). The stationary blade 91 is configured by connecting an outer shroud 93 and an inner shroud 94 to each end portion in the longitudinal direction of the blade portion 92, and a plurality of stationary blades 91 are annularly assembled in the turbine casing 20. It is supported. The rotor blade 22 is configured to be fixed to an outer peripheral portion of a rotor disk 34 integrally coupled with a rotor (not shown) via an inner shroud 35. A passage in which a plurality of stationary blades 91 and moving blades 22 are alternately arranged is a combustion gas passage 37.

  In this embodiment, the stationary blade 91 is divided into a blade portion 92, an outer shroud 93, and an inner shroud 94, and each end portion of the blade portion 92 and the outer shroud 63 and the inner shroud 94 are connected to the combustion gas passage 37. The bolts are fastened outside. The outer shroud 93 is integrally provided with the outer shroud of the adjacent moving blade 22.

  That is, the outer shroud 93 has a plate shape that covers the outer peripheral side of the stationary blade 91 and the moving blade 22, and an annular folded portion 93a is formed at a position corresponding to the stationary blade 91, and a flange portion is formed on the folded portion 93a. 93b is formed. And the wing | blade part 92 penetrates this outer side shroud 9, and it is fastened using the fastening bolt 95 in the position where flange part 92a, 93b contact | abutted. Further, the wing portion 92 penetrates the inner shroud 94, and the tip portions are fastened using fastening bolts 96.

  Further, the fitting portion 93 c formed at the front portion of the outer shroud 93 is in close contact with the concave portion 20 c of the turbine casing 20 via the seal member 97, and the locking portion 93 d formed at the rear portion is a concave portion of the turbine casing 20. The stationary blade 91 is assembled at a predetermined position by engaging with the concave portion 20c of the turbine casing 20 through the seal member 98 while the engagement portion 94a formed in the front portion of the inner shroud 94 is brought into contact with the concave portion 20c. Will be.

  Accordingly, the combustion gas flows through the combustion gas passage 37 and the blades 92 of the stationary blade 91, the outer shroud 93, and the inner shroud 94 are heated, and the outer shroud 93 and the inner shroud 94 are heated on the front and back sides. It deforms into a curved shape due to the expansion difference. However, since the base end portion and the tip end portion of the wing portion 92 and the connecting portions of the shrouds 93 and 94 are located outside the combustion gas passage 37 and are fastened by the fastening bolts 95 and 96, The stress acting on the connecting portion is absorbed, and the occurrence of cracks and breakage is prevented.

  As described above, in the stationary blade and the gas turbine according to the sixth embodiment, the stationary blade 91 has a divided structure of the blade portion 92, the outer shroud 93, and the inner shroud 94, and the outer shroud 93 of the adjacent moving blade 22 is disposed. Each end portion of the blade portion 92 extending to the outside and having a cooling passage is fastened by fastening bolts 95 and 96 outside the combustion gas passage 37.

  Therefore, by making the wing portion 92, the outer shroud 93, and the inner shroud 94 into a divided structure, a high-performance material can be used for a single crystal structure, and a cooling structure corresponding to a high heat load can be applied. The performance can be improved. Further, each end portion of the wing portion 92 and the connecting portion of the outer shroud 93 and the inner shroud 94 are fastened by bolts 95 and 96 outside the combustion gas passage 37, and thermal stress is efficiently caused by slipping at the fastening portion. It can be relaxed and durability can be improved. Further, the outer shroud 93 extends to the outside of the moving blade 22, so that it is not necessary to connect the stationary blade 91 and each shroud located outside the moving blade 22, and combustion is achieved by reducing the connecting portion between the shrouds. Leakage of combustion gas from the gas passage 37 can be prevented.

  In the turbine blade and the gas turbine according to the present invention, the turbine blade has a split structure, and the end of the blade portion and the shroud are bolted to the outside of the combustion gas passage to effectively reduce thermal stress and improve durability. It can be applied to any kind of turbine blades and gas turbines.

It is sectional drawing of the stationary blade as a turbine blade which concerns on Example 1 of this invention. It is the perspective view seen from the upper direction of the stationary blade of Example 1. FIG. It is the perspective view seen from the downward direction of the stationary blade of Example 1. FIG. 1 is a schematic configuration diagram of a gas turbine according to Embodiment 1. FIG. 1 is a schematic diagram illustrating a turbine blade arrangement in a gas turbine according to Embodiment 1. FIG. It is sectional drawing of the stationary blade as a turbine blade which concerns on Example 2 of this invention. It is the schematic showing the turbine blade arrangement | sequence in the gas turbine which concerns on Example 3 of this invention. FIG. 6 is a cross-sectional view of a main part of a stationary blade as a turbine blade of Example 3. 6 is a schematic diagram showing a mounting structure of a seal member in a stationary blade of Example 3. FIG. 6 is a schematic diagram showing a mounting structure of a seal member in a stationary blade of Example 3. FIG. It is sectional drawing of the stationary blade as a turbine blade which concerns on Example 4 of this invention. It is a perspective view of the stationary blade as a turbine blade which concerns on Example 5 of this invention. It is the schematic showing the turbine blade arrangement | sequence in the gas turbine which concerns on Example 6 of this invention.

Explanation of symbols

11 Compressor 12 Combustor 13 Turbine 14 Exhaust chamber 20 Turbine casing 21, 51, 61, 71, 81, 91 Stator blade 22 Rotor blade 24 Rotor 31, 52, 62, 72, 82, 92 Blade section 31 a, 52 a, 63a, 72a, 82a, 92a Flange portion 32, 53, 63, 73, 83, 93 Outer shroud 32c, 53c, 63b, 73c, 83a, 93b Flange portion 33, 54, 64, 74, 84, 94 Inner shroud 37, 55, 75, 85 Combustion gas passage 38, 40, 56, 65, 77, 86, 87, 95, 96 Fastening bolt 58 Welding 66a, 66b Seal member 76 Labyrinth seal

Claims (8)

  In a turbine blade having a blade portion having a cooling passage therein and an outer shroud and an inner shroud connected to respective longitudinal ends of the blade portion, at least one end portion of the blade portion includes the outer shroud or the inner shroud. A turbine blade that is penetrated and bolted outside the combustion gas passage.   2. The turbine blade according to claim 1, wherein a flange portion is formed at a base end portion of the blade portion, the blade portion passes through either the outer shroud or the inner shroud, and the flange portion abuts. A turbine blade that is bolted in position.   The turbine blade according to claim 2, wherein the blade portion passes through the outer shroud and the inner shroud, and is bolted to either the outer shroud or the inner shroud at a position where the flange portion abuts. A turbine blade characterized in that a tip portion is bolted to the other end.   2. The turbine blade according to claim 1, wherein one end portion of the blade portion passes through either the outer shroud or the inner shroud and is bolted, and the other end portion of the blade portion is connected to the other by welding. Turbine blades characterized by that.   2. The turbine blade according to claim 1, wherein one end portion of the blade portion passes through either the outer shroud or the inner shroud and is bolted, and the other end portion of the blade portion has a seal member on the other. Turbine blades connected with each other.   The turbine blade according to any one of claims 1 to 5, wherein a plurality of the blade portions are provided, and at least one end portion of the plurality of blade portions penetrates the outer shroud or the inner shroud, A turbine blade characterized by being bolted on the outside.   The turbine blade according to any one of claims 1 to 6, wherein a blade shroud adjacent to the outer shroud is integrally provided.   In a gas turbine that obtains rotational power by supplying fuel to compressed air compressed by a compressor and burning with a combustor and supplying the generated combustion gas to the turbine, the stationary blades constituting the turbine have cooling passages. A gas turbine characterized in that at least one end portion of a blade portion has an outer shroud or an inner shroud and is bolted to the outside of a combustion gas passage.

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