WO2023234134A1 - 静翼、及びこれを備えているガスタービン - Google Patents

静翼、及びこれを備えているガスタービン Download PDF

Info

Publication number
WO2023234134A1
WO2023234134A1 PCT/JP2023/019272 JP2023019272W WO2023234134A1 WO 2023234134 A1 WO2023234134 A1 WO 2023234134A1 JP 2023019272 W JP2023019272 W JP 2023019272W WO 2023234134 A1 WO2023234134 A1 WO 2023234134A1
Authority
WO
WIPO (PCT)
Prior art keywords
space
wall surface
rear end
blade
end space
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Ceased
Application number
PCT/JP2023/019272
Other languages
English (en)
French (fr)
Japanese (ja)
Inventor
正人 片岡
猛 梅原
渓也 石山
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Mitsubishi Heavy Industries Ltd
Mitsubishi Power Ltd
Original Assignee
Mitsubishi Heavy Industries Ltd
Mitsubishi Power Ltd
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Mitsubishi Heavy Industries Ltd, Mitsubishi Power Ltd filed Critical Mitsubishi Heavy Industries Ltd
Priority to DE112023001579.1T priority Critical patent/DE112023001579T5/de
Priority to CN202380039277.2A priority patent/CN119173682A/zh
Priority to JP2024524775A priority patent/JP7802165B2/ja
Priority to KR1020247038978A priority patent/KR20250002625A/ko
Publication of WO2023234134A1 publication Critical patent/WO2023234134A1/ja
Anticipated expiration legal-status Critical
Ceased legal-status Critical Current

Links

Images

Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/187—Convection cooling
    • F01D5/188—Convection cooling with an insert in the blade cavity to guide the cooling fluid, e.g. forming a separation wall
    • F01D5/189—Convection cooling with an insert in the blade cavity to guide the cooling fluid, e.g. forming a separation wall the insert having a tubular cross-section, e.g. airfoil shape
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D9/00—Stators
    • F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • F01D9/023—Transition ducts between combustor cans and first stage of the turbine in gas-turbine engines; their cooling or sealings
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/182—Transpiration cooling
    • F01D5/183—Blade walls being porous
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/187—Convection cooling
    • F01D5/188—Convection cooling with an insert in the blade cavity to guide the cooling fluid, e.g. forming a separation wall
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D9/00—Stators
    • F01D9/06—Fluid supply conduits to nozzles or the like
    • F01D9/065—Fluid supply or removal conduits traversing the working fluid flow, e.g. for lubrication-, cooling-, or sealing fluids
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/12—Cooling of plants
    • F02C7/16—Cooling of plants characterised by cooling medium
    • F02C7/18—Cooling of plants characterised by cooling medium the medium being gaseous, e.g. air
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2220/00—Application
    • F05D2220/30—Application in turbines
    • F05D2220/32—Application in turbines in gas turbines
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/201—Heat transfer, e.g. cooling by impingement of a fluid
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/221—Improvement of heat transfer
    • F05D2260/2214—Improvement of heat transfer by increasing the heat transfer surface
    • F05D2260/22141—Improvement of heat transfer by increasing the heat transfer surface using fins or ribs

Definitions

  • the present invention relates to a stator blade and a gas turbine equipped with the same.
  • This application claims priority based on Japanese Patent Application No. 2022-089017 filed in Japan on May 31, 2022, the contents of which are incorporated herein.
  • a gas turbine includes a compressor that compresses air to generate compressed air, a combustor that burns fuel in the compressed air to generate combustion gas, and a turbine that is driven by the combustion gas.
  • the turbine includes a turbine rotor that rotates about an axis, a turbine casing that covers the rotor, and a plurality of rows of stationary blades.
  • the turbine rotor has a rotor shaft centered on the axis, and a plurality of rotor blade rows attached to the rotor shaft.
  • the plurality of rotor blade rows are lined up in the axial direction in which the axis extends.
  • Each rotor blade row has a plurality of rotor blades arranged in a circumferential direction with respect to the axis.
  • the plurality of stator blade rows are arranged in the axial direction and attached to the inner peripheral side of the turbine casing.
  • Each of the plurality of stator blade rows is arranged upstream of the axis of any one of the plurality of rotor blade rows.
  • Each stator blade row has a plurality of stator blades arranged in a circumferential direction with respect to the axis.
  • a stationary blade includes a blade body extending in the radial direction with respect to an axis and forming an airfoil shape, an inner shroud provided on the radially inner side of the blade body, and an outer shroud provided on the radial outside of the blade body.
  • the blade body of the stator vane is arranged in a combustion gas flow path through which combustion gas passes.
  • the inner shroud defines a radially inner edge of the combustion gas flow path.
  • the outer shroud defines a radially outer edge of the combustion gas flow path.
  • stator blades are generally cooled with air or the like.
  • a plurality of cooling air spaces through which cooling air passes are formed in a blade body of a stator blade described in Patent Document 1 below. All of the plurality of cooling air spaces extend in the blade height direction, which is the radial direction with respect to the axis.
  • a plurality of trailing end cooling passages extending from the trailing end space closest to the trailing edge to the trailing edge of the wing body are formed.
  • a plurality of pins and a plurality of ribs that partition the inside of the rear end space in the blade height direction are arranged in this rear end space. The plurality of pins protrude from the pressure side inner surface or the negative pressure side inner surface defining the rear end space.
  • the plurality of ribs also protrude from the positive pressure side inner surface or the negative pressure side inner surface defining the rear end space. Furthermore, a cylindrical insert is inserted into an adjacent space of the plurality of cooling air spaces that is in contact with an edge on the leading edge side of the rear end space. This cylindrical insert is formed with a through hole through which cooling air can be ejected from the inner circumferential side toward the outer circumferential side.
  • a plurality of cooling air spaces through which cooling air passes are also formed in the vane body of the stationary blade described in Patent Document 2 below. All of the plurality of cooling air spaces extend in the blade height direction, which is the radial direction with respect to the axis.
  • a plurality of pins are arranged in the rear end space closest to the rear edge. The plurality of pins are joined to a positive pressure side inner surface and a negative pressure side inner surface that define a rear end space.
  • a cylindrical insert is inserted into an adjacent space of the plurality of cooling air spaces that is in contact with an edge on the leading edge side of the rear end space. This cylindrical insert is formed with a through hole through which cooling air can be ejected from the inner circumferential side toward the outer circumferential side.
  • Patent Document 1 and Patent Document 2 by providing a plurality of pins in the rear end space, heat exchange between the wall surface that defines the rear end space and the cooling air in the rear end space is efficiently performed. be able to. Further, as in the technique described in Patent Document 2, when the positive pressure side inner surface and the negative pressure side inner surface that define the rear end space are connected using a plurality of pins in the rear end space, the rear end space is included in the stationary blade. The rigidity of the part can be increased.
  • the pressure within the cooling air space of the blade body is higher than the pressure outside the blade body. Therefore, during operation of the gas turbine, a force is applied to the blade body that causes the blade surface of the blade body to expand outward due to the pressure difference between the inside and outside of the blade body.
  • the phenomenon in which the wing surface undergoes creep deformation in a bulging manner due to this force is called bulging.
  • Bulging phenomenon is likely to occur in parts of the vane body of the stator blade that have low rigidity. When the bulging phenomenon occurs, not only the aerodynamic performance of the stator vane decreases, but also the durability of the stator vane decreases.
  • an object of the present disclosure is to provide a technique that can improve the bulging strength of a stator blade and suppress a decrease in aerodynamic performance and durability of the stator blade.
  • a stationary blade included in a gas turbine includes a blade body having an airfoil-shaped cross section and extending in a blade height direction having a directional component perpendicular to the cross section.
  • the wing body includes a leading edge and a trailing edge that extend in the blade height direction, a pressure surface and a suction surface that extend in the blade height direction and connect the leading edge and the trailing edge, and the leading edge and the rear edge. It has a rear end space and an adjacent space located between the edge and between the positive pressure surface and the negative pressure surface, and a plurality of rear end cooling passages penetrating from the rear end space to the rear edge.
  • the adjacent space is located closer to the front edge than the rear end space.
  • An edge of the rear end space on the front edge side is open and communicates with the rear end space.
  • Both the adjacent space and the rear end space are defined by an inner wall surface.
  • the inner wall surface of the adjacent space and the inner wall surface of the rear end space each include a positive pressure side inner wall surface extending along the positive pressure surface, and a positive pressure side inner wall surface extending along the negative pressure surface and extending from the positive pressure side inner wall surface to the negative pressure side inner wall surface. and a negative pressure side inner wall surface that is remote from the pressure surface side.
  • the positive pressure side inner wall surface of the adjacent space and the positive pressure side inner wall surface of the rear end space are connected.
  • the negative pressure side inner wall surface of the adjacent space and the negative pressure side inner wall surface of the rear end space are connected.
  • a plurality of partition ribs and a plurality of pins are arranged in line in the blade height direction and partition the inside of the rear end space in the blade height direction.
  • Each of the plurality of pins is joined to the positive pressure side inner wall surface of the rear end space and the negative pressure side inner wall surface of the rear end space.
  • At least one of the plurality of partition ribs forms an extended partition rib that extends into the adjacent space and is joined to the inner wall surface defining the adjacent space.
  • the positive pressure side inner wall surface and the negative pressure side inner wall surface that define the rear end space are joined by a plurality of pins. Therefore, the stiffness around the trailing end space in the wing body is higher than the stiffness around the adjacent space in the wing body. In other words, the stiffness around the adjacent space in the wing body is lower than the stiffness around the trailing end space in the wing body.
  • the bulging phenomenon is likely to occur in parts of the wing body that have low rigidity. When this bulging phenomenon occurs, not only the aerodynamic performance of the stator blade deteriorates, but also the durability of the stator blade decreases.
  • high stress is generated near the boundary between a portion of the wing body with low rigidity and a portion of the wing body with high rigidity due to deformation due to the bulging phenomenon.
  • high stress is generated near the boundary between the adjacent space and the rear end space due to deformation due to the bulging phenomenon.
  • At least one of the plurality of partition ribs arranged in the rear end space is extended into the adjacent space to increase the rigidity near the boundary between the adjacent space and the rear end space.
  • the bulging strength of the stator vane is improved, and deterioration in the aerodynamic performance and durability of the stator vane due to the occurrence of the bulging phenomenon can be suppressed.
  • a gas turbine according to one aspect of the invention for achieving the above object includes: The stationary blade according to the one aspect, a rotor that rotates about an axis, and a casing that covers the outer peripheral side of the rotor.
  • the stationary blade is fixed to an inner circumferential surface of the casing.
  • FIG. 1 is a schematic cross-sectional view of a gas turbine in an embodiment according to the present disclosure.
  • FIG. 2 is a perspective view of a stationary blade in an embodiment according to the present disclosure.
  • FIG. 3 is a sectional view taken along the line III-III in FIG. 2 (a vertical sectional view of the main part of the stationary blade).
  • 4 is a cross-sectional view (cross-sectional view of the wing body) taken along the line IV-IV in FIG. 3.
  • FIG. FIG. 5 is an enlarged view of the V section in FIG. 4 (a cross-sectional view of the main part of the wing body).
  • FIG. 7 is a vertical cross-sectional view of a main part of a stator blade in a modification of an embodiment according to the present disclosure.
  • FIG. 7 is a cross-sectional view of a main part of a wing body in another modified example of an embodiment according to the present disclosure.
  • stator vane of the present disclosure and a gas turbine equipped with the stator vane will be described in detail with reference to the drawings.
  • the gas turbine 10 of the present embodiment includes a compressor 20 that compresses air A, and a combustion system that combusts fuel F in the air A compressed by the compressor 20 to generate combustion gas G. 30, and a turbine 40 driven by combustion gas G.
  • the compressor 20 includes a compressor rotor 21 that rotates around an axis Ar, a compressor casing 25 that covers the compressor rotor 21, and a plurality of stator blade rows 26.
  • the turbine 40 includes a turbine rotor 41 that rotates around an axis Ar, a turbine casing 45 that covers the turbine rotor 41, and a plurality of stator blade rows 46.
  • the direction in which the axis Ar extends is referred to as an axial direction Da
  • the circumferential direction around the axis Ar is simply referred to as a circumferential direction Dc
  • the direction perpendicular to the axis Ar is referred to as a radial direction Dr.
  • one side of the axial direction Da is defined as the upstream side of the axis Dau, and the opposite side thereof is defined as the downstream side of the axis Dad.
  • the side approaching the axis Ar in the radial direction Dr is defined as the radially inner side Dri, and the opposite side thereof is defined as the radially outer side Dro.
  • the compressor 20 is arranged on the axial upstream side Dau with respect to the turbine 40.
  • Gas turbine 10 further includes an intermediate casing 16 .
  • This intermediate casing 16 is arranged between the compressor casing 25 and the turbine casing 45 in the axial direction Da.
  • Compressor casing 25, intermediate casing 16, and turbine casing 45 are connected to each other to form gas turbine casing 15.
  • the compressor rotor 21 has a rotor shaft 22 that extends in the axial direction Da centering on the axis Ar, and a plurality of rotor blade rows 23 attached to the rotor shaft 22.
  • the plural rotor blade rows 23 are arranged in the axial direction Da.
  • Each row of rotor blades 23 is composed of a plurality of rotor blades arranged in the circumferential direction Dc.
  • Any one of the plurality of stator blade rows 26 is disposed on the downstream side Dad of each of the plurality of rotor blade rows 23 on the axis line.
  • Each stator blade row 26 is provided inside the compressor casing 25.
  • Each stator blade row 26 is composed of a plurality of stator blades arranged in the circumferential direction Dc.
  • the turbine rotor 41 has a rotor shaft 42 that extends in the axial direction Da centering on the axis Ar, and a plurality of rotor blade rows 43 that are attached to the rotor shaft 42.
  • the plurality of rotor blade rows 43 are arranged in the axial direction Da.
  • Each row of rotor blades 43 is composed of a plurality of rotor blades arranged in the circumferential direction Dc.
  • Any one of the plurality of stator blade rows 46 is arranged on the axial upstream side Dau of each of the plurality of rotor blade rows.
  • Each stationary blade row 46 is provided inside the turbine casing 45.
  • Each stator blade row 46 is composed of a plurality of stator blades arranged in the circumferential direction Dc.
  • the combustor 30 is attached to the intermediate casing 16.
  • the compressor 20 compresses air A to generate compressed air.
  • This compressed air flows into the combustor 30.
  • Fuel F is supplied to the combustor 30.
  • fuel F is combusted in compressed air to generate high-temperature, high-pressure combustion gas G.
  • This combustion gas G is sent from the combustor 30 to an annular combustion gas passage 49 within the turbine casing 45 .
  • the combustion gas G rotates the turbine rotor 41 while flowing in the combustion gas flow path 49 toward the downstream side Dad of the axis.
  • This rotation of the turbine rotor 41 causes the rotor of the generator GEN connected to the gas turbine rotor 11 to rotate.
  • the generator GEN generates electricity.
  • stator blades that constitute the first-stage stator blade row 46 in the turbine 40 and modifications thereof will be described.
  • FIG. 2 is a perspective view of the stationary blade in this embodiment.
  • FIG. 3 is a cross-sectional view taken along the line III--III in FIG. 2 (a vertical cross-sectional view of the main part of the stationary blade).
  • FIG. 4 is a cross-sectional view taken along the line IV-IV in FIG. 3 (a cross-sectional view of the wing body).
  • FIG. 5 is an enlarged view of the V section in FIG. 4 (a cross-sectional view of the main part of the wing body).
  • the stationary blade 50 of this embodiment includes a blade body 51, an inner shroud 80i, an outer shroud 80o, an inner impingement plate 88i, an outer impingement plate 88o, and a first insert 68A. , a second insert 68B, and a third insert 68C.
  • the blade body 51 has an airfoil-shaped cross section and extends in the blade height direction Dh having a directional component perpendicular to this cross section.
  • the outer shroud 80o is provided at one end of the blade body 51 in the blade height direction Dh.
  • the inner shroud 80i is provided at the other end of the blade body 51 in the blade height direction Dh.
  • the wing body 51, the inner shroud 80i, and the outer shroud 80o are integrally formed by casting or the like.
  • the blade height direction Dh becomes the radial direction Dr.
  • the first blade height side Dh1 which is one side in the blade height direction Dh
  • the second blade height side Dh2 which is the other side in the blade height direction Dh
  • the outer shroud 80o is provided on the radially outer side Dro of the blade body 51
  • the inner shroud 80i is provided on the radially inner side Dri of the blade body 51.
  • the blade height direction Dh may be referred to as the radial direction Dr
  • the first blade height side Dh1 may be referred to as the radially outer Dro
  • the blade height second side Dh2 may be referred to as the radially inner Dri.
  • the wing surface which is the outer surface of the wing body 51, includes a leading edge 52, a trailing edge 53, a suction surface 54 that is a convex surface, and a pressure surface 55 that is a concave surface. and has.
  • the negative pressure surface 54 and the positive pressure surface 55 are surfaces that connect the leading edge 52 and the trailing edge 53.
  • the leading edge 52, the trailing edge 53, the suction surface 54, and the pressure surface 55 all extend in the radial direction Dr, which is the blade height direction Dh.
  • Dr the leading edge 52 is located on the upstream side Dau of the axis with respect to the trailing edge 53.
  • the negative pressure surface 54 faces one side in the circumferential direction Dc
  • the positive pressure surface 55 faces the other side in the circumferential direction Dc.
  • This blade body 51 is arranged within the combustion gas flow path 49 of the turbine 40 described using FIG. 1.
  • This blade body 51 has a plurality of cooling air spaces 60 extending in the blade height direction Dh (radial direction Dr) within the blade body 51.
  • the plurality of cooling air spaces 60 are lined up along the camber line CL of the wing body 51 from the axial upstream side Dau to the axial downstream Dad.
  • the cooling air space 60 closest to the axial upstream side Dau is the first space 60A
  • the cooling air space 60 adjacent to the axial downstream side Dad of this first space 60A is the second space 60B.
  • the cooling air space 60 adjacent to Dad on the downstream side of the axis line of this second space 60B is called a third space 60C
  • the cooling air space 60 adjacent to Dad on the downstream side of the axis line of this third space 60C is called a fourth space 60D.
  • the first space 60A and the second space 60B are partitioned by a first partition wall 57 that extends in a direction substantially perpendicular to the camber line CL.
  • the second space 60B and the third space 60C are partitioned by a second partition wall 58 that extends in a direction substantially perpendicular to the camber line CL.
  • a first insert 68A is arranged in the first space 60A
  • a second insert 68B is arranged in the second space 60B
  • a third insert 68C is arranged in the third space 60C.
  • the inner shroud 80i defines a radially inner edge of the annular combustion gas flow path 49.
  • the outer shroud 80o also defines a radially outer edge of the annular combustion gas flow path 49.
  • the inner shroud 80i includes a shroud main body 81i, a peripheral wall 84i, and a retainer 86i.
  • the shroud main body 81i is a square plate-shaped member that extends in a direction including a direction component perpendicular to the radial direction Dr, which is the blade height direction Dh.
  • This shroud body 81i has a gas path surface 82i and an anti-gas path surface 83i.
  • the gas path surface 82i is a surface that faces the radially outer Dro, which is the first blade height side Dh1, and is in contact with the combustion gas G.
  • the anti-gas path surface 83i is a surface facing radially inward Dri, which is the second blade height side Dh2. This anti-gas path surface 83i is in a back-to-back relationship with the gas path surface 82i.
  • the peripheral wall 84i is a wall that protrudes from the shroud main body 81i to the radially inner side Dri along the outer peripheral edge of the shroud main body 81i.
  • a cavity 85i is formed in the inner shroud 80i by the shroud main body 81i and the peripheral wall 84i, and is recessed toward the radially outer side Dro.
  • the retainer 86i is formed on the radially inner side Dri of the peripheral wall 84i. This retainer 86i is connected to the radially outer end Dro of an inner cover (not shown) that is fixed to the gas turbine casing 15, and is used to support the radially inner Dri portion of the stationary blade 50 on the inner cover. take on the role of
  • the outer shroud 80o basically has the same configuration as the inner shroud 80i. Therefore, like the inner shroud 80i, the outer shroud 80o also includes a shroud main body 81o and a peripheral wall 84o. However, the outer shroud 80o does not have a portion corresponding to the retainer 86i of the inner shroud 80i. Like the shroud main body 81i of the inner shroud 80i, the shroud main body 81o of the outer shroud 80o is also a square plate-shaped member, and has a gas path surface 82o and an anti-gas path surface 83o.
  • the gas path surface 82o of the outer shroud 80o faces the radially inner Dri, which is the second blade height side Dh2, and is a surface in contact with the combustion gas G.
  • the anti-gas path surface 83o of the outer shroud 80o is a surface facing the radially outer Dro, which is the first blade height side Dh1.
  • the peripheral wall 84o is a wall that protrudes radially outward from the shroud body 81o along the outer peripheral edge of the shroud body 81o.
  • a cavity 85o is formed in the outer shroud 80o by a shroud main body 81o and a peripheral wall 84o, which is recessed toward the radially inner side Dri.
  • a portion of the peripheral wall 84o serves to attach the stationary blade 50 to the inner peripheral side of the turbine casing 45.
  • the inner impingement plate 88i partitions the cavity 85i of the inner shroud 80i into two spaces in the blade height direction Dh.
  • a plurality of impingement holes 89 penetrating in the blade height direction Dh are formed in this inner impingement plate 88i.
  • the cooling air Ac that has flowed into the cavity 85i of the inner shroud 80i is ejected from the plurality of impingement holes 89 of the inner impingement plate 88i to the radially outer side Dro, collides with the anti-gas path surface 83i of the inner shroud 80i, and performs impingement cooling. do.
  • the cooling air Ac that has impingement-cooled the opposite gas path surface 83i is ejected into the combustion gas flow path 49 through a passage (not shown), for example, to film-cool the gas path surface 82i and the like.
  • the outer impingement plate 88o partitions the cavity 85o of the outer shroud 80o into two spaces in the blade height direction Dh.
  • a plurality of impingement holes 89 penetrating in the blade height direction Dh are formed in this outer impingement plate 88o.
  • the cooling air Ac that has flowed into the cavity 85o of the outer shroud 80o is ejected from the plurality of impingement holes 89 of the outer impingement plate 88o to the radially inner side Dri, collides with the anti-gas path surface 83o of the outer shroud 80o, and performs impingement cooling there. .
  • the cooling air Ac that has impingement-cooled the opposite gas path surface 83o is ejected into the combustion gas flow path 49 through a passage (not shown), for example, to film-cool the gas path surface 82o and the like.
  • the third space 60C is closed at the radially inner end Dri, which is the second blade height side Dh2, and closed at the radially outer Dro end, which is the first blade height side Dh1. It's open.
  • the radially inner Dri end of the blade body 51 forms a part of the anti-gas path surface 83i of the inner shroud 80i
  • the radially outer Dro end of the blade body 51 forms a part of the anti-gas path surface 83o of the outer shroud 80o. form part of Therefore, the opening 61C of the third space 60C opens at the anti-gas path surface 83o of the outer shroud 80o.
  • a plurality of third wing surface cooling passages 67C are formed in the wing body 51, penetrating from the inner wall surface 62C defining the third space 60C to the wing surface of the wing body 51.
  • the first space 60A and the second space 60B are closed at the end of the radially inner Dri, which is the second blade height side Dh2, and the blade height second side Dh2 is closed.
  • the end of the radially outer Dro, which is one side Dh1 is open.
  • the opening of the first space 60A and the opening of the second space 60B are opened at the anti-gas path surface 83o of the outer shroud 80o, similar to the opening 61C of the third space 60C.
  • a plurality of first wing surface cooling passages 67A are formed in the wing body 51, penetrating from the inner wall surface 62A defining the first space 60A to the wing surface of the wing body 51.
  • a plurality of second wing surface cooling passages 67B are formed in the wing body 51, penetrating from the inner wall surface 62B defining the second space 60B to the wing surface of the wing body 51.
  • the third insert 68C disposed in the third space 60C has a cylindrical shape and extends in the radial direction Dr.
  • the radially inner Dri end of the third insert 68C is closed, and the radially outer Dro end of the third insert 68C is open.
  • This third insert 68C is located almost entirely within the third space 60C, but a radially outer Dro portion of the third insert 68C protrudes from the third space 60C to the radially outer Dro, and is located within the third space 60C.
  • the end of the radially outer Dro of the insert 68C is located radially outer Dro than the outer impingement plate 88o.
  • cooling air Ac can flow into the third insert 68C from the opening of the third insert 68C.
  • a plurality of impingement holes 69 penetrating from the inner circumferential side to the outer circumferential side are formed in the cylindrical third insert 68C. Therefore, the cooling air Ac that has flowed into the third insert 68C is ejected from the plurality of impingement holes 69, collides with the inner wall surface 62C that defines the third space 60C, and impingement-cools the space.
  • a part of the cooling air Ac that has impingement-cooled the inner wall surface 62C that defines the third space 60C is injected into the combustion gas flow path 49 through the plurality of third blade surface cooling passages 67C described above, and coats the blade surface with a film. Cooling.
  • the first insert 68A disposed in the first space 60A and the second insert 68B disposed in the second space 60B have substantially the same configuration as the third insert 68C described using FIGS. 3 and 4. It is. That is, the first insert 68A and the second insert 68B, like the third insert 68C, have a cylindrical shape and extend in the radial direction Dr. The radially inner Dri ends of the first insert 68A and the second insert 68B are closed, and the radially outer Dro ends of the first insert 68A and the second insert 68B are open.
  • the first insert 68A is located almost entirely within the first space 60A, but the radially outer Dro portion of the first insert 68A protrudes from the third space 60C to the radially outer Dro, and is located within the first insert 60A.
  • the end of the radially outer Dro of 68A is located further radially outwardly than the outer impingement plate 88o. Therefore, cooling air Ac can flow into the first insert 68A from the opening of the first insert 68A.
  • the cylindrical first insert 68A is also formed with a plurality of impingement holes 69 penetrating from the inner circumferential side to the outer circumferential side.
  • the cooling air Ac that has flowed into the first insert 68A is ejected from the plurality of impingement holes 69, collides with the inner wall surface 62A defining the first space 60A, and impingement-cools the space.
  • the cooling air Ac that has impingement-cooled the inner wall surface 62A that defines the first space 60A is ejected into the combustion gas flow path 49 via the plurality of first blade surface cooling passages 67A described above to film-cool the blade surface.
  • the second insert 68B is almost entirely located within the second space 60B, but the radially outer Dro portion of the second insert 68B protrudes from the second space 60B to the radially outer Dro, and is located within the second space 60B.
  • the end of the radially outer Dro of the second insert 68B is located further radially outward than the outer impingement plate 88o. Therefore, cooling air Ac can flow into the second insert 68B from the opening of the second insert 68B.
  • the cylindrical second insert 68B is also formed with a plurality of impingement holes 69 penetrating from the inner circumferential side to the outer circumferential side. Therefore, the cooling air Ac that has flowed into the second insert 68B is ejected from the plurality of impingement holes 69, collides with the inner wall surface 62B defining the second space 60B, and impingement-cools the space.
  • the cooling air Ac that impingement-cooled the inner wall surface 62B that defines the second space 60B is ejected into the combustion gas flow path 49 through the plurality of second blade surface cooling passages 67B described above to film-cool the blade surface.
  • the fourth space 60D is located furthest Dad on the downstream side of the axis among the plurality of cooling air spaces 60. Therefore, hereinafter, this fourth space 60D will be referred to as a rear end space 60D. Further, the third space 60C is adjacent to the axial upstream side Dau of this rear end space 60D. Therefore, hereinafter, this third space 60C will be referred to as an adjacent space 60C.
  • the rear end space 60D is closed at the radially inner end Dri, which is the second blade height side Dh2, and at the radially outer Dro end, which is the first blade height side Dh1. Further, an edge (edge on the front edge 52 side) of the rear end space 60D on the upstream side of the axis Dau is open and communicates with the adjacent space 60C.
  • the inner wall surface 62D defining the rear end space 60D has a first blade height inner wall surface 64D, a second blade height inner wall surface 65D, a positive pressure side inner wall surface 63Dp, and a negative pressure side inner wall surface 63Dn.
  • the first blade height side inner wall surface 64D faces the second blade height side Dh2 and defines the edge of the first blade height side Dh1 of the rear end space 60D.
  • the second blade height side inner wall surface 65D faces the first blade height side Dh1 and defines the edge of the second blade height side Dh2 of the rear end space 60D.
  • the positive pressure side inner wall surface 63Dp faces the negative pressure surface 54, spreads along the positive pressure surface 55, and defines the edge of the rear end space 60D on the positive pressure surface 55 side.
  • An edge of the positive pressure side inner wall surface 63Dp on the first blade height side Dh1 is connected to the first blade height side inner wall surface 64D.
  • the edge of the second blade height side Dh2 of the positive pressure side inner wall surface 63Dp is connected to the second blade height side inner wall surface 65D.
  • the negative pressure side inner wall surface 63Dn faces the positive pressure surface 55 side, spreads along the negative pressure surface 54, and is separated from the positive pressure side inner wall surface 63Dp toward the negative pressure surface 54 side, and extends toward the negative pressure surface 54 side of the rear end space 60D.
  • the edge of this negative pressure side inner wall surface 63Dn on the first blade height side Dh1 is connected to the first blade height side inner wall surface 64D.
  • the edge of the second blade height side Dh2 of this negative pressure side inner wall surface 63Dn is connected to the second blade height side inner wall surface 65D.
  • the edge of the negative pressure side inner wall surface 63Dn on the downstream side of the axis Dad is connected to the edge of the positive pressure side inner wall surface 63Dp on the downstream side of the axis Dad.
  • a plurality of trailing end cooling passages 67D are formed in the blade body 51, penetrating from the trailing end space 60D to the combustion gas flow path.
  • the rear end cooling passage 67D penetrates from the boundary between the edge of the downstream side Dad of the negative pressure side inner wall surface 63Dn and the edge of the downstream side Dad of the axis line of the positive pressure side inner wall surface 63Dp to the rear edge 53 of the blade body 51.
  • the plurality of trailing end cooling passages 67D are lined up in the blade height direction Dh.
  • a plurality of partition ribs 71 that are arranged in the blade height direction Dh and partition the inside of the rear end space 60D in the blade height direction Dh, and a plurality of cylindrical or polygonal columnar pins 72 are arranged. There is.
  • the plurality of partition ribs 71 extend in a direction substantially perpendicular to the blade height direction Dh, and are joined to the pressure side inner wall surface 63Dp and the negative pressure side inner wall surface 63Dn.
  • the plurality of pins 72 extend in a direction substantially perpendicular to the camber line CL, and are joined to the positive pressure side inner wall surface 63Dp and the negative pressure side inner wall surface 63Dn.
  • At least one of the plurality of partition ribs 71 forms an extended partition rib 71a extending into the adjacent space 60C.
  • the extended partitioning rib 71a is arranged at a position of 30 to 70% of the dimension in the blade height direction Dh of the rear end space 60D from the inner wall surface 65D on the second blade height side that defines the rear end space 60D. ing. That is, the extended partitioning rib 71a is arranged near the center of the blade body 51 in the blade height direction Dh.
  • the inner wall surface 62C defining the adjacent space 60C which is the third space 60C, has a blade height second side inner wall surface 65C, a positive pressure side inner wall surface 63Cp, a negative pressure side inner wall surface 63Cn, and the axis downstream side Dad of the second partition wall 58. and a second partition wall surface 66C.
  • the second blade height side inner wall surface 65C faces the first blade height side Dh1 and defines the edge of the second blade height side Dh2 of the adjacent space 60C.
  • the edge of the axis downstream side Dad of the second blade height side inner wall surface 65C is connected to the axis line upstream edge Dau of the second blade height side inner wall surface 65D that defines the rear end space 60D.
  • the positive pressure side inner wall surface 63Cp faces the negative pressure surface 54 side, spreads along the positive pressure surface 55, and defines the edge of the rear end space 60D on the positive pressure surface 55 side.
  • the edge of the second blade height side Dh2 of the positive pressure side inner wall surface 63Cp is connected to the second blade height side inner wall surface 65C.
  • An edge of the positive pressure side inner wall surface 63Cp on the upstream side Dau of the axis is connected to the second partition wall surface 66C.
  • An edge on the downstream side Dad of the positive pressure side inner wall surface 63Cp is connected to an edge Dau on the upstream side of the axis line of the positive pressure side inner wall surface 63Dp that defines the rear end space 60D.
  • the negative pressure side inner wall surface 63Cn faces the positive pressure surface 55 side, spreads along the negative pressure surface 54 surface, and is separated from the positive pressure side inner wall surface 63Cp toward the negative pressure surface 54 side, and extends toward the negative pressure surface 54 side of the adjacent space 60C. Define the edges.
  • the edge of the second blade height side Dh2 of this negative pressure side inner wall surface 63Cn is connected to the second blade height side inner wall surface 65C.
  • An edge of the axial upstream side Dau of this negative pressure side inner wall surface 63Cn is connected to the second partition wall surface 66C.
  • An edge on the downstream side Dad of the negative pressure side inner wall surface 63Cn is connected to an edge Dau on the upstream side of the axis line of the negative pressure side inner wall surface 63Dn that defines the rear end space 60D.
  • the portions existing in the adjacent space 60C are joined to the positive pressure side inner wall surface 63Dp and the negative pressure side inner wall surface 63Cn of the inner wall surface 62C that defines the adjacent space 60C.
  • the extension dimension L in which the extension partition rib 71a extends from the edge of the axis upstream side Dau (the edge on the front edge 52 side) of the rear end space 60D in the adjacent space 60C is the rear end space. It is smaller than twice the boundary width W, which is the distance between the positive pressure side inner wall surface 63Dp and the negative pressure side inner wall surface 63Dn at the edge on the side of the leading edge 52 at 60D.
  • the edge of the rear end space 60D on the side of the front edge 52 is the boundary between the adjacent space 60C and the rear end space 60D, and is the opening 61D on the side of the front edge 52 in the rear end space 60D. Further, in this embodiment, the edge on the front edge 52 side in the rear end space 60D is the position of the edge on the front edge 52 side of the pin 72 closest to the front edge 52 among the plurality of pins 72.
  • the cooling air Ac that has flowed into the rear end space 60D convectively cools the plurality of partition ribs 71, the plurality of pins 72, and the inner wall surface 62D that defines the rear end space 60D while flowing within the rear end space 60D. .
  • This cooling air Ac flows into the plurality of rear end cooling passages 67D.
  • the cooling air Ac that has flowed into the rear end cooling passage 67D convectively cools the surface defining the rear end space 60D while flowing within the rear end space 60D.
  • This cooling air Ac is ejected from the trailing edge 53 of the blade body 51 into the combustion gas flow path 49 .
  • the positive pressure side inner wall surface 63Dp and the negative pressure side inner wall surface 63Dn that define the rear end space 60D are joined by the plurality of pins 72 and the partition ribs 71.
  • the third insert 68C is inserted into the adjacent space 60C adjacent to the rear end space 60D, the positive pressure side inner wall surface 63Cp and the negative pressure side inner wall surface 63Cn that define the adjacent space 60C are separated by pins, etc. is not joined. Therefore, the rigidity around the adjacent space 60C in the wing body 51 is lower than the rigidity around the rear end space 60D in the wing body 51.
  • the parts of the blade body 51 with low rigidity are caused by the blade surface expanding due to the pressure difference between the inside and outside of the blade body 51 during operation of the gas turbine. Bulging phenomenon caused by creep deformation is likely to occur. When the bulging phenomenon occurs, this portion of the blade surface with low rigidity deforms as shown by the two-dot broken line in FIG. When this bulging phenomenon occurs, not only the aerodynamic performance of the stator blade 50 deteriorates, but also the durability of the stator blade 50 deteriorates.
  • high stress is generated near the boundary between a portion of the blade body 51 with low rigidity and a portion of the blade body 51 with high rigidity due to deformation due to the bulging phenomenon.
  • high stress is generated near the boundary between the adjacent space 60C and the rear end space 60D due to deformation due to the bulging phenomenon.
  • the amount of deformation due to the bulging phenomenon increases near the center of the blade body 51 in the blade height direction Dh.
  • the plurality of partitioning ribs 71 arranged near the center of the wing body 51 in the blade height direction Dh are arranged adjacent to each other. It extends into the space 60C, increasing the rigidity of the center portion of the wing body 51 in the blade height direction Dh, which is near the boundary between the adjacent space 60C and the rear end space 60D. As a result, in this embodiment, the occurrence of the bulging phenomenon is suppressed, and deformation near the center of the blade body 51 in the blade height direction Dh, which is near the boundary between the adjacent space 60C and the rear end space 60D, is suppressed. There is.
  • the bulging strength of the stator vane 50 is improved, and a decrease in the aerodynamic performance and durability of the stator vane 50 due to the occurrence of the bulging phenomenon can be suppressed.
  • the extension dimension L in which the extension partition rib 71a extends from the edge on the front edge 52 side of the rear end space 60D into the adjacent space 60C is the length on the front edge 52 side in the rear end space 60D. is smaller than twice the boundary width W, which is the distance between the positive pressure side inner wall surface 63Dp of the rear end space 60D and the negative pressure side inner wall surface 63Dn of the rear end space 60D at the edge of the rear end space 60D.
  • the extended dimension L of this extended partitioning rib 71a may be larger than twice the boundary width W.
  • the bulging strength of the wing body 51 increases compared to the case where the extension dimension L of the extension partition rib 71a is twice the boundary width W, but the amount of increase in the bulging strength is greater than the increase in the extension dimension L. small. Furthermore, since the stationary blade 50 is formed by casting, if the extension dimension L of the extension partition rib 71a is increased, casting becomes difficult accordingly. Therefore, in this embodiment, the extension dimension L of the extension partition rib 71a is made smaller than twice the boundary width W.
  • the inner wall surface 65D on the second blade height side defining the rear end space 60D is connected to the rear end space 60D.
  • a plurality of partition ribs 71 arranged at positions of 30 to 70% of the dimension in the blade height direction Dh form extended partition ribs 71a.
  • the portion of the extended partitioning rib 71a that exists in the adjacent space 60C is located in the right side of the inner wall surface 62C that defines the adjacent space 60C. It is joined to the pressure side inner wall surface 63Cp and the negative pressure side inner wall surface 63Cn.
  • the portion of the extended partitioning rib 71b that exists within the adjacent space 60C is the negative pressure side inner wall surface of the inner wall surface 62C that defines the adjacent space 60C. 63Cn, and not to the positive pressure side inner wall surface 63Cp.
  • the static pressure that the positive pressure side inner wall surface 63Cp that defines the adjacent space 60C receives from the adjacent space 60C and the static pressure that the negative pressure side inner wall surface 63Cn that defines the adjacent space 60C receives are the same value.
  • the flow velocity of the combustion gas G flowing along the suction surface 54 outside the blade body 51 is higher than the flow velocity of the combustion gas G flowing along the pressure surface 55 outside the blade body 51.
  • the static pressure that 54 receives is lower than the static pressure that pressure surface 55 receives from outside of wing body 51 .
  • the pressure difference between the pressure in the adjacent space 60C and the pressure in the area outside the blade body 51 along the negative pressure surface 54 is equal to the pressure in the adjacent space 60C and the pressure in the area outside the blade body 51 along the positive pressure surface 55. It becomes larger than the pressure difference between Therefore, even if a bulging phenomenon occurs around the adjacent space 60C, the amount of deformation due to the bulging phenomenon is larger on the negative pressure surface 54 side than on the positive pressure surface 55 side.
  • the part of the extended partitioning rib 71b that exists in the adjacent space 60C is joined only to the negative pressure side inner wall surface 63Cn of the inner wall surface 62C that defines the adjacent space 60C, and The rigidity is increased only on the negative pressure surface 54 side near the boundary between the rear end space 60D and the rear end space 60D.
  • the space occupied by the extended partitioning rib 71b in the adjacent space 60C is smaller than in the embodiment described above, so that the insertability of the third insert 68C into the adjacent space 60C and the insertion of the third insert in the adjacent space 60C are reduced.
  • the degree of freedom in layout of the insert 68C can be increased.
  • this modification is a modification of the stationary blade 50 in the embodiment. However, this modification may be applied to the stationary blade 50a in the first modification.
  • stator vanes 50, 50a, and 50b in the above embodiment and each modification each have four spaces from a first space 60A to a fourth space 60D.
  • the vane may have more space.
  • inserts are arranged in all spaces except for the rear end space 60D, which is the fourth space 60D.
  • inserts may not be arranged in all spaces except the rear end space 60D.
  • stator blades 50, 50a, and 50b in the above embodiment and each modification are all stator blades that constitute the first stage stator blade row 46.
  • the stator vane may be a stator vane that constitutes the stator blade row 46 on the downstream side of the axis Dad than the first stage stator vane row 46.
  • the stationary blade in the first aspect is
  • the stationary blades 50, 50a, and 50b included in the gas turbine 10 include a blade body 51 having an airfoil-shaped cross section and extending in a blade height direction Dh having a directional component perpendicular to the cross section.
  • the blade body 51 has a leading edge 52 and a trailing edge 53 extending in the blade height direction Dh, and a pressure surface 55 and a suction surface extending in the blade height direction Dh and connecting the leading edge 52 and the trailing edge 53.
  • a rear end space 60D and an adjacent space 60C located between the front edge 52 and the rear edge 53 and between the positive pressure surface 55 and the negative pressure surface 54, and from the rear end space 60D to the rear It has a plurality of rear end cooling passages 67D extending through the edge 53.
  • the adjacent space 60C is located closer to the front edge 52 than the rear end space 60D.
  • An edge of the rear end space 60D on the front edge 52 side is open and communicates with the rear end space 60D.
  • Both the adjacent space 60C and the rear end space 60D are defined by inner wall surfaces 62C and 62D.
  • the inner wall surface 62C of the adjacent space 60C and the inner wall surface 62D of the rear end space 60D both have positive pressure side inner wall surfaces 63Cp, 63Dp extending along the positive pressure surface 55 and extending along the negative pressure surface 54.
  • negative pressure side inner wall surfaces 63Cn and 63Dn are spaced apart from the positive pressure side inner wall surfaces 63Cp and 63Dp toward the negative pressure surface 54 side.
  • the positive pressure side inner wall surface 63Cp of the adjacent space 60C and the positive pressure side inner wall surface 63Dp of the rear end space 60D are connected.
  • the negative pressure side inner wall surface 63Cn of the adjacent space 60C and the negative pressure side inner wall surface 63Dn of the rear end space 60D are connected.
  • a plurality of partition ribs 71 and a plurality of pins 72 are arranged in line in the blade height direction Dh and partition the inside of the rear end space 60D in the blade height direction Dh. There is.
  • the plurality of pins 72 are all joined to the positive pressure side inner wall surface 63Dp of the rear end space 60D and the negative pressure side inner wall surface 63Dn of the rear end space 60D.
  • At least one partitioning rib 71 among the plurality of partitioning ribs 71 has extended partitioning ribs 71a and 71b that extend into the adjacent space 60C and are joined to the inner wall surface 62C that defines the adjacent space 60C. I will do it.
  • the positive pressure side inner wall surface 63Dp and the negative pressure side inner wall surface 63Dn that define the rear end space 60D are joined by a plurality of pins 72. Therefore, the rigidity around the rear end space 60D in the wing body 51 is higher than the rigidity around the adjacent space 60C in the wing body 51. In other words, the rigidity around the adjacent space 60C in the wing body 51 is lower than the rigidity around the rear end space 60D in the wing body 51.
  • a bulging phenomenon is likely to occur in a portion of the wing body 51 that has low rigidity. When this bulging phenomenon occurs, not only the aerodynamic performance of the stator blade 50 deteriorates, but also the durability of the stator blade 50 deteriorates.
  • high stress is generated near the boundary between a portion of the blade body 51 with low rigidity and a portion of the blade body 51 with high rigidity due to deformation due to the bulging phenomenon.
  • high stress is generated near the boundary between the adjacent space 60C and the rear end space 60D due to deformation due to the bulging phenomenon.
  • At least one partition rib 71 among the plurality of partition ribs 71 arranged in the rear end space 60D is extended into the adjacent space 60C, and the boundary between the adjacent space 60C and the rear end space 60D is Increased rigidity in the vicinity.
  • the occurrence of the bulging phenomenon is suppressed, and deformation near the boundary between the adjacent space 60C and the rear end space 60D is suppressed.
  • the bulging strength of the stator vanes 50, 50a, 50b is improved, and a decrease in the aerodynamic performance and durability of the stator vanes 50, 50a, 50b due to the occurrence of the bulging phenomenon can be suppressed.
  • the stationary blade in the second embodiment is In the stationary blade 50 described in the first aspect, the extended partitioning rib 71a extends from one end of the rear end space 60D in the blade height direction Dh to the blade height direction Dh of the rear end space 60D. It is placed at a position of 30% to 70% of the size of .
  • the amount of deformation due to the bulging phenomenon increases around the adjacent space 60C in the blade body 51 and near the center of the blade body 51 in the blade height direction Dh. Therefore, in this aspect, a position of 30 to 70% of the dimension of the rear end space 60D in the blade height direction Dh from one end of the rear end space 60D in the blade height direction Dh among the plurality of partition ribs 71 is selected.
  • the partition rib 71 disposed at the center of the blade body 51 in the blade height direction Dh is an extended partition rib 71a.
  • the stationary blade in the third aspect is In the stationary blade 50a in the first aspect or the second aspect, all of the plurality of partition ribs 71 are the extended partition ribs 71a, 71b.
  • the stationary blade in the fourth aspect is in the stationary blades 50, 50a, 50b in any one of the first to third aspects, the extended partitioning ribs 71a, 71b extend from the edge of the rear end space 60D on the front edge 52 side.
  • the extension dimension L extending within the adjacent space 60C is a boundary width W that is the distance between the positive pressure side inner wall surface 63Dp and the negative pressure side inner wall surface 63Dn at the edge on the front edge 52 side of the rear end space 60D. less than twice.
  • the extension dimension L of the extension partition ribs 71a and 71b may be larger than twice the boundary width W.
  • the bulging strength of the wing body 51 increases compared to the case where the extension dimension L of the extension partition ribs 71a and 71b is made twice the boundary width W, but the bulging strength increases more than the increase in the extension dimension L. Quantity is small.
  • the stator vanes 50, 50a, 50b are formed by casting, if the extension dimension L of the extension partition ribs 71a, 71b is increased, casting becomes difficult accordingly. Therefore, in this embodiment, the extension dimension L of the extended partitioning ribs 71a, 71b is made smaller than twice the boundary width W by weighing the increase in bulging strength and the difficulty of casting.
  • the stationary blade in the fifth aspect is In the stationary blade 50b in any one of the first to fourth aspects, the extended partitioning rib 71b is joined to the negative pressure side inner wall surface 63Dn of the adjacent space 60C, and the It is not joined to the positive pressure side inner wall surface 63Dp.
  • the static pressure that the positive pressure side inner wall surface 63Cp that defines the adjacent space 60C receives from the adjacent space 60C and the static pressure that the negative pressure side inner wall surface 63Cn that defines the adjacent space 60C receives are the same value.
  • the flow velocity of the combustion gas G flowing along the suction surface 54 outside the blade body 51 is higher than the flow velocity of the combustion gas G flowing along the pressure surface 55 outside the blade body 51.
  • the static pressure that 54 receives is lower than the static pressure that pressure surface 55 receives from outside of wing body 51 .
  • the pressure difference between the pressure in the adjacent space 60C and the pressure in the area outside the blade body 51 along the negative pressure surface 54 is equal to the pressure in the adjacent space 60C and the pressure in the area outside the blade body 51 along the positive pressure surface 55. It becomes larger than the pressure difference between Therefore, even if a bulging phenomenon occurs around the adjacent space 60C, the amount of deformation due to the bulging phenomenon is larger on the negative pressure surface 54 side than on the positive pressure surface 55 side.
  • the portion of the extended partitioning rib 71b that exists in the adjacent space 60C is joined only to the negative pressure side inner wall surface 63Cn of the inner wall surface 62C that defines the adjacent space 60C, and The rigidity is increased only on the negative pressure surface 54 side near the boundary with the rear end space 60D.
  • the stationary blade in the sixth aspect is The stationary blades 50, 50a, 50b in any one of the first to fifth aspects further include a cylindrical insert 68C disposed in the adjacent space 60C.
  • a plurality of impingement holes 69 are formed in the insert 68C, penetrating from the inner circumferential side to the outer circumferential side.
  • the cooling air Ac that has flowed into the insert 68C is ejected from the impingement hole 69, collides with the inner wall surface 62C that defines the adjacent space 60C, and impingement-cools the space. Therefore, in this aspect, the blade surface along the inner wall surface 62C that defines the adjacent space 60C can be efficiently cooled.
  • the gas turbine in the above embodiment can be understood as follows, for example.
  • the gas turbine in the seventh aspect is: Stator blades 50, 50a in any one of the first to sixth aspects. 50b, a rotor 41 that rotates around an axis Ar, and a casing 45 that covers the outer peripheral side of the rotor 41.
  • the stationary blades 50, 50a. 50b is fixed to the inner peripheral surface of the casing 45.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Physics & Mathematics (AREA)
  • Fluid Mechanics (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
PCT/JP2023/019272 2022-05-31 2023-05-24 静翼、及びこれを備えているガスタービン Ceased WO2023234134A1 (ja)

Priority Applications (4)

Application Number Priority Date Filing Date Title
DE112023001579.1T DE112023001579T5 (de) 2022-05-31 2023-05-24 Statorschaufel und gasturbine, die dieselbe umfasst
CN202380039277.2A CN119173682A (zh) 2022-05-31 2023-05-24 静叶片及具备该静叶片的燃气涡轮
JP2024524775A JP7802165B2 (ja) 2022-05-31 2023-05-24 静翼、及びこれを備えているガスタービン
KR1020247038978A KR20250002625A (ko) 2022-05-31 2023-05-24 정익, 및 이를 구비하고 있는 가스 터빈

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
JP2022-089017 2022-05-31
JP2022089017 2022-05-31

Publications (1)

Publication Number Publication Date
WO2023234134A1 true WO2023234134A1 (ja) 2023-12-07

Family

ID=89024891

Family Applications (1)

Application Number Title Priority Date Filing Date
PCT/JP2023/019272 Ceased WO2023234134A1 (ja) 2022-05-31 2023-05-24 静翼、及びこれを備えているガスタービン

Country Status (5)

Country Link
JP (1) JP7802165B2 (https=)
KR (1) KR20250002625A (https=)
CN (1) CN119173682A (https=)
DE (1) DE112023001579T5 (https=)
WO (1) WO2023234134A1 (https=)

Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JPS60111002A (ja) * 1983-10-28 1985-06-17 ウエスチングハウス　エレクトリック　コ−ポレ−ション 燃焼タービン固定子ベーン
JP2002266604A (ja) * 2001-01-05 2002-09-18 General Electric Co <Ge> 切断型リブを有するタービンノズル及び方法
US20070258814A1 (en) * 2006-05-02 2007-11-08 Siemens Power Generation, Inc. Turbine airfoil with integral chordal support ribs
US20220145799A1 (en) * 2020-11-12 2022-05-12 Solar Turbines Incorporated Fin for internal cooling of vane wall

Family Cites Families (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JPS6353131U (https=) 1986-09-24 1988-04-09
JP3316415B2 (ja) 1997-05-01 2002-08-19 三菱重工業株式会社 ガスタービン冷却静翼
JP7618429B2 (ja) 2020-12-03 2025-01-21 キヤノン株式会社 情報処理装置、情報処理方法およびプログラム

Patent Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JPS60111002A (ja) * 1983-10-28 1985-06-17 ウエスチングハウス　エレクトリック　コ−ポレ−ション 燃焼タービン固定子ベーン
JP2002266604A (ja) * 2001-01-05 2002-09-18 General Electric Co <Ge> 切断型リブを有するタービンノズル及び方法
US20070258814A1 (en) * 2006-05-02 2007-11-08 Siemens Power Generation, Inc. Turbine airfoil with integral chordal support ribs
US20220145799A1 (en) * 2020-11-12 2022-05-12 Solar Turbines Incorporated Fin for internal cooling of vane wall

Also Published As

Publication number Publication date
JPWO2023234134A1 (https=) 2023-12-07
KR20250002625A (ko) 2025-01-07
CN119173682A (zh) 2024-12-20
DE112023001579T5 (de) 2025-01-09
JP7802165B2 (ja) 2026-01-19

Similar Documents

Publication Publication Date Title
JP3782637B2 (ja) ガスタービン冷却静翼
US8182203B2 (en) Turbine blade and gas turbine
JP6418667B2 (ja) 翼、及びこれを備えているガスタービン
EP4283094B1 (en) Turbine component with stress relieving cooling circuit
EP3184742B1 (en) Turbine airfoil with trailing edge cooling circuit
US20200063574A1 (en) Flared central cavity aft of airfoil leading edge
US10119406B2 (en) Blade with stress-reducing bulbous projection at turn opening of coolant passages
JP5881369B2 (ja) タービン動翼及びこれを備えたガスタービン
US10605090B2 (en) Intermediate central passage spanning outer walls aft of airfoil leading edge passage
EP3184743A1 (en) Turbine airfoil with trailing edge cooling circuit
JP4823872B2 (ja) ターボ機械の移動ブレードのための中央冷却回路
CN104884741B (zh) 用于涡轮机的叶片
WO2021186796A1 (ja) 静翼及びガスタービン
CN111936724B (zh) 涡轮动叶以及燃气轮机
US12215605B2 (en) Stator blade and gas turbine comprising same
WO2023234134A1 (ja) 静翼、及びこれを備えているガスタービン
JP2017044093A (ja) タービン動翼、及び、ガスタービン
WO2024252886A1 (ja) タービン動翼及びガスタービン
JP7297132B1 (ja) タービン静翼及びガスタービン
JP7778043B2 (ja) 静翼、及びこれを備えているガスタービン
US12385403B2 (en) Turbine stator vane
JP7436725B1 (ja) 動翼、及びこれを備えているガスタービン
US9234428B2 (en) Turbine bucket internal core profile
JP4423660B2 (ja) 分割構造タービン翼
WO2026042628A1 (ja) タービン静翼及びガスタービン

Legal Events

Date Code Title Description
121 Ep: the epo has been informed by wipo that ep was designated in this application

Ref document number: 23815893

Country of ref document: EP

Kind code of ref document: A1

WWE Wipo information: entry into national phase

Ref document number: 2024524775

Country of ref document: JP

ENP Entry into the national phase

Ref document number: 20247038978

Country of ref document: KR

Kind code of ref document: A

WWE Wipo information: entry into national phase

Ref document number: 112023001579

Country of ref document: DE

Ref document number: 1020247038978

Country of ref document: KR

122 Ep: pct application non-entry in european phase

Ref document number: 23815893

Country of ref document: EP

Kind code of ref document: A1