WO2023106125A1 - タービン翼及びガスタービン - Google Patents
タービン翼及びガスタービン Download PDFInfo
- Publication number
- WO2023106125A1 WO2023106125A1 PCT/JP2022/043511 JP2022043511W WO2023106125A1 WO 2023106125 A1 WO2023106125 A1 WO 2023106125A1 JP 2022043511 W JP2022043511 W JP 2022043511W WO 2023106125 A1 WO2023106125 A1 WO 2023106125A1
- Authority
- WO
- WIPO (PCT)
- Prior art keywords
- downstream
- blade
- cooling
- passage
- region
- Prior art date
Links
- 238000001816 cooling Methods 0.000 claims abstract description 199
- 238000011144 upstream manufacturing Methods 0.000 claims abstract description 156
- 239000012809 cooling fluid Substances 0.000 claims abstract description 89
- 239000012720 thermal barrier coating Substances 0.000 claims abstract description 38
- 238000005192 partition Methods 0.000 claims abstract description 14
- 239000000567 combustion gas Substances 0.000 claims description 29
- 230000003746 surface roughness Effects 0.000 claims description 12
- 238000000638 solvent extraction Methods 0.000 abstract description 3
- 239000011295 pitch Substances 0.000 description 36
- 239000007789 gas Substances 0.000 description 22
- 230000007423 decrease Effects 0.000 description 5
- 230000014509 gene expression Effects 0.000 description 5
- 238000009434 installation Methods 0.000 description 5
- 238000010586 diagram Methods 0.000 description 4
- 239000000919 ceramic Substances 0.000 description 3
- 239000000446 fuel Substances 0.000 description 3
- 239000002184 metal Substances 0.000 description 3
- 239000011248 coating agent Substances 0.000 description 2
- 238000000576 coating method Methods 0.000 description 2
- 239000007921 spray Substances 0.000 description 2
- 239000002826 coolant Substances 0.000 description 1
- 230000000694 effects Effects 0.000 description 1
- 239000012530 fluid Substances 0.000 description 1
- 239000000463 material Substances 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 230000004048 modification Effects 0.000 description 1
- 230000003647 oxidation Effects 0.000 description 1
- 238000007254 oxidation reaction Methods 0.000 description 1
- 238000004781 supercooling Methods 0.000 description 1
- 238000007751 thermal spraying Methods 0.000 description 1
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
-
- 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
-
- 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
- F01D5/28—Selecting particular materials; Particular measures relating thereto; Measures against erosion or corrosion
-
- 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
- F01D5/28—Selecting particular materials; Particular measures relating thereto; Measures against erosion or corrosion
- F01D5/288—Protective coatings for blades
-
- 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
-
- 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
-
- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T50/00—Aeronautics or air transport
- Y02T50/60—Efficient propulsion technologies, e.g. for aircraft
Definitions
- a turbine blade of a gas turbine or the like is cooled by flowing a cooling fluid through a serpentine flow path (serpentine flow path) formed inside the turbine blade to cool the turbine blade exposed to a high-temperature gas flow or the like.
- Such turbine blades are provided with cooling holes, turbulators, or the like for effectively cooling the turbine blades with a cooling fluid, or are provided with thermal barrier coatings or the like for suppressing heat input to the turbine blades.
- Patent Document 1 discloses a gas turbine blade in which a serpentine flow path is provided inside the blade body.
- the trailing edge of the turbine blade is provided with a plurality of cooling holes for cooling the trailing edge of the turbine blade using cooling fluid flowing through the serpentine flow path.
- a plurality of fins are provided on the inner wall surface of the cooling passage that forms the serpentine flow path of the turbine blade to promote turbulence in the flow of the cooling fluid in the cooling passage.
- At least one embodiment of the present invention aims to provide a turbine blade and a gas turbine capable of suppressing overcooling or insufficient cooling of the turbine blade due to installation of a bypass section in a serpentine flow path. do.
- a turbine blade comprises: a wing body; a plurality of cooling passages each extending in the blade height direction inside the blade body and connected to each other via folded portions; a bypass portion provided in a partition wall portion that partitions a pair of adjacent cooling passages among the plurality of cooling passages and that allows the pair of cooling passages to communicate with each other; with A turbine blade, wherein the pair of cooling passages includes an upstream passage and a downstream passage positioned downstream of the upstream passage with respect to the flow of cooling fluid, The turbine blade, a plurality of cooling holes formed in the blade body so as to be arranged along the blade height direction, communicating with the downstream passage and opening to the surface of the blade body; A plurality of turbulators provided on the inner wall surface of the downstream passage and arranged along the blade height direction, or A thermal barrier coating covering the surface of the wing body, an upstream region positioned upstream with respect to the flow of the cooling fluid in the downstream passage relative to a position corresponding to the bypass portion in the blade height
- a gas turbine includes: a turbine blade as described above; a combustor for generating combustion gas flowing through a combustion gas passage in which the turbine blades are provided.
- a turbine blade and a gas turbine capable of suppressing overcooling or insufficient cooling of the turbine blade due to installation of a bypass section in a serpentine flow path are provided.
- FIG. 1 is a schematic diagram of a gas turbine according to one embodiment
- FIG. 1 is a schematic cross-sectional view along the blade height direction of a turbine blade according to one embodiment
- FIG. FIG. 3 is a diagram showing an AA cross section of the turbine blade of FIG. 2
- 1 is a schematic cross-sectional view showing an enlarged trailing edge side portion of a turbine blade according to one embodiment
- FIG. 1 is a schematic cross-sectional view showing an enlarged leading edge side portion of a turbine blade according to one embodiment
- FIG. 1 is a schematic cross-sectional view showing an enlarged trailing edge side portion of a turbine blade according to one embodiment
- FIG. 1 is a schematic cross-sectional view showing an enlarged trailing edge side portion of a turbine blade according to one embodiment
- FIG. 1 is a schematic cross-sectional view showing an enlarged trailing edge side portion of a turbine blade according to one embodiment
- FIG. 1 is a schematic cross-sectional view showing an enlarged trailing edge side portion of a turbine
- FIG. 1 is a schematic cross-sectional view showing an enlarged trailing edge side portion of a turbine blade according to one embodiment
- FIG. 1 is a schematic cross-sectional view showing an enlarged trailing edge side portion of a turbine blade according to one embodiment
- FIG. 1 is a schematic cross-sectional view showing an enlarged leading edge side portion of a turbine blade according to one embodiment
- FIG. 1 is a schematic cross-sectional view showing an enlarged trailing edge side portion of a turbine blade according to one embodiment
- FIG. 1 is a schematic cross-sectional view showing an enlarged trailing edge side portion of a turbine blade according to one embodiment
- FIG. 1 is a schematic diagram of a gas turbine as an example of a turbine to which turbine blades according to some embodiments are applied.
- a gas turbine 1 includes a compressor 2 for generating compressed air, a combustor 4 for generating combustion gas using the compressed air and fuel, and a combustion gas that is rotationally driven.
- a turbine 6 configured to:
- the compressor 2 includes a plurality of stator vanes 16 fixed on the side of the compressor casing 10 and a plurality of stator vanes 16 implanted in the rotor 8 so as to be alternately arranged with respect to the stator vanes 16 .
- a rotor blade 18 Air taken in from an air intake port 12 is sent to the compressor 2, and this air passes through a plurality of stationary blades 16 and a plurality of moving blades 18 and is compressed to produce a high temperature and high pressure. of compressed air.
- the combustor 4 is configured to be supplied with fuel and compressed air generated by the compressor 2 , and the fuel is combusted in the combustor 4 to generate combustion gas, which is the working fluid of the turbine 6 .
- the gas turbine 1 has a plurality of combustors 4 arranged in a casing 20 along a circumferential direction around a rotor 8 .
- the turbine 6 has a combustion gas passage 28 formed by the turbine casing 22 and includes a plurality of stator vanes 24 and rotor blades 26 provided in the combustion gas passage 28 .
- the stationary blades 24 are fixed on the turbine casing 22 side, and a plurality of stationary blades 24 arranged along the circumferential direction of the rotor 8 form a row of stationary blades.
- the rotor blades 26 are implanted in the rotor 8, and a plurality of rotor blades 26 arranged along the circumferential direction of the rotor 8 form a rotor blade cascade.
- the row of stationary blades and row of moving blades are alternately arranged in the axial direction of the rotor 8 .
- the combustion gas from the combustor 4 that has flowed into the combustion gas passage 28 passes through the plurality of stationary blades 24 and the plurality of moving blades 26, thereby driving the rotor 8 to rotate.
- a generator may be connected to the rotor 8 and driven by the turbine 6 to generate electric power. Combustion gas after driving the turbine 6 is discharged to the outside through an exhaust chamber 30 .
- At least one of the rotor blades 26 or stator blades 24 of the turbine 6 are turbine blades 40, described below.
- description will be made mainly with reference to the drawing of the moving blade 26 as the turbine blade 40, but basically the same explanation can be applied to the stationary blade 24 as the turbine blade 40 as well.
- FIG. 2 is a schematic cross-sectional view along the blade height direction of a turbine blade 40 (rotating blade 26) according to one embodiment
- FIG. 3 shows an AA cross section of the turbine blade 40 of FIG. It is a diagram.
- the arrows in the figure indicate the direction of flow of the cooling fluid. 2, illustration of the thermal barrier coating 86 (see FIG. 3) is omitted.
- the turbine blade 40 (rotating blade 26) according to one embodiment is provided integrally with a blade body 42, a platform 80, and a rotor 8 (see FIG. 1). and a blade root portion 82 embedded in.
- the blade body 42 is provided so as to extend along the radial direction of the rotor 8 (hereinafter sometimes simply referred to as “radial direction”), and has a base end 50 fixed to the platform 80 and a blade height and a distal end 48 positioned opposite (radially outward) to the proximal end 50 in the longitudinal direction (radial direction of the rotor 8).
- the wing body 42 includes a top plate 49 that forms a distal portion of the wing body 42 that includes a tip 48 .
- the airfoil 42 of the blade 26 also has a leading edge 44 and a trailing edge 46 extending from a proximal end 50 to a distal end 48, and the surfaces of the airfoil 42 between the proximal end 50 and the tip 48, respectively. It includes a pressure surface (ventral surface) 56 and a suction surface (backward surface) 58 extending along the wing height direction.
- cooling passages 60 As shown in FIGS. 2 and 3, a plurality of cooling passages 60a to 60f (hereinafter collectively referred to as cooling passages 60) extending along the blade height direction are provided inside the blade body 42. ing. In the illustrated embodiment, the cooling passages 60a-60f are arranged in this order from the leading edge 44 side to the trailing edge 46 side.
- a plurality of cooling passages 60 are connected to each other via folded portions 59 positioned on the distal end 48 side or the proximal end 50 side to form serpentine flow passages (serpentine flow passages) 61 (61A, 61B).
- a pair of adjacent cooling passages 60 among the plurality of cooling passages 60 are partitioned by a partition wall portion 32 extending along the blade height direction. The direction of flow of the cooling fluid flowing through the serpentine flow passages 61 (61A, 61B) is reversed in the blade height direction at the folding portion 59. As shown in FIG.
- cooling passages 60a to 60c In the exemplary embodiment shown in FIGS. 2 and 3, of the cooling passages 60a to 60c, the adjacent cooling passages 60a and 60b, and the adjacent cooling passages 60b and 60c are arranged with the folded portions 59 therebetween. connected. These cooling passages 60a to 60c form a serpentine passage 61A on the leading edge side.
- cooling passages 60d to 60f the cooling passages 60d and 60e adjacent to each other, and the cooling passages 60e and 60f, which are adjacent to each other, each have a folded portion 59. connected through These cooling passages 60d to 60f form a serpentine passage 61B on the trailing edge side.
- a cooling fluid (for example, air) for cooling the turbine blades 40 is supplied to the serpentine flow passages 61 (61A, 61B).
- internal flow passages 84A and 84B are provided inside the blade root portion 82, and among the plurality of cooling passages 60 forming the serpentine flow passages 61A and 61B, the most upstream cooling passages 60c and 60d are provided. are communicated with the internal flow paths 84A and 84B, respectively. Cooling fluid from the outside is supplied to the serpentine channels 61A and 61B through internal channels 84A and 84B, respectively.
- the cooling fluid introduced into the serpentine flow passages 61A, 61B sequentially flows downstream through the plurality of cooling passages 60 forming the serpentine flow passages 61A, 61B.
- the cooling fluid flowing through the most downstream cooling passages 60a and 60f in the flow direction of the cooling fluid flows through outlet openings 64A and 64B provided on the tip 48 side of the blade body 42. It is adapted to flow out into the combustion gas passage 28 outside the turbine blades 40 .
- the turbine blade 40 has a plurality of cooling holes that are formed in the blade body 42 so as to be arranged along the blade height direction, communicate with the cooling passages 60 and open to the surface of the blade body 42 . Prepare.
- the turbine blades 40 are formed in the blade body 42 so as to be arranged along the blade height direction, and the cooling passages 60a (downstream A plurality of cooling holes 72 (72a to 72c) communicating with the side passages 66) and opening to the front edge surface of the blade 42 are provided.
- a part of the cooling fluid flowing through the cooling passages 60 a passes through the cooling holes 72 and flows out from the openings of the leading edge portions of the blade bodies 42 to the combustion gas passages 28 outside the turbine blades 40 .
- the cooling fluid flowing out in this way flows along the outer surface of the turbine blade 40, the cooling fluid forms a film boundary layer (film-like coolant flow) on the outer surface.
- This film boundary layer suppresses heat transfer from the combustion gas flowing through the combustion gas passage 28 to the blade body 42 (heat input to the turbine blade 40).
- the cooling fluid passing through the cooling holes 72 convectively cools the leading edge portion including the leading edge 44 of the airfoil 42 .
- the turbine blades 40 are formed in the blade body 42 so as to be arranged along the blade height direction, and are positioned closest to the trailing edge.
- a plurality of cooling holes 70 are provided that communicate with a cooling passage 60f (a downstream passage 66 to be described later) and open to the surface of the trailing edge portion of the wing body 42 .
- a portion of the cooling fluid flowing through the cooling passage 60f passes through the cooling holes 70 and flows out from the opening at the trailing edge of the blade body 42 to the combustion gas passage 28 outside the turbine blade 40.
- the passage of the cooling fluid through the cooling holes 70 in this manner convectively cools the trailing edge, including the trailing edge 46 of the airfoil 42 .
- the turbine blade 40 includes a plurality of turbulators 34 provided on the inner wall surface 63 of the cooling passage 60 and arranged along the blade height direction.
- the inner wall surface 63 of each of the plurality of cooling passages 60 is provided with a plurality of rib-like turbulators 34 protruding from the inner wall surface 63 .
- the turbine blade 40 includes a thermal barrier coating 86 (see FIG. 3) that covers the surfaces (the pressure surface 56 and the suction surface 58) of the blade body 42.
- the thermal barrier coating 86 provided on the surface of the turbine blade 40 suppresses heat transfer from the combustion gas flowing through the combustion gas passage 28 to the blade body 42 (heat input to the turbine blade 40).
- the thermal barrier coating 86 may include a ceramic layer formed of ceramic, and a bond layer provided between the ceramic layer and the blade 42 and formed of an oxidation-resistant metal or the like. Thermal barrier coating 86 may be deposited by thermal spraying.
- the turbine blade 40 may include any one or more of the leading edge cooling holes 72 , the trailing edge cooling holes 70 , the turbulators 34 , or the thermal barrier coating 86 . .
- the turbine blade 40 includes an upstream side passage 65 and a downstream side passage 65 which are a pair of adjacent cooling passages 60 among a plurality of cooling passages 60 forming a serpentine flow passage 61 .
- a bypass portion 36 is provided in the partition wall portion 32 that partitions the passage 66 and allows the upstream passage 65 and the downstream passage 66 to communicate with each other.
- the downstream passage 66 is the cooling passage 60 positioned next to the upstream passage 65 among the plurality of cooling passages 60, and the flow of the cooling fluid flowing through the plurality of cooling passages 60 (serpentine passage 61) is , the cooling passage 60 located downstream of the upstream passage 65 .
- the bypass portion 36 may be a hole or slit provided in the partition wall portion 32 .
- the turbine blade 40 includes a plurality of cooling passages 60a to 60c forming a serpentine passage 61A on the leading edge side.
- 60a downstream passage 66
- the cooling passage 60 b upstream passage 65
- the cooling passage 60 a downstream passage 66
- the turbine blade 40 includes adjacent cooling passages 60e (upstream passages 65) among the plurality of cooling passages 60d to 60f forming the serpentine passage 61B on the trailing edge side.
- a bypass portion 36 provided in the partition wall portion 32 partitioning the cooling passage 60f (downstream passage 66) is provided.
- the cooling passage 60 e (upstream passage 65 ) and the cooling passage 60 f (downstream passage 66 ) communicate with each other via the bypass portion 36 .
- an upstream region R1 located upstream with respect to the flow of cooling fluid in the downstream passage 66 (cooling passage 60a or 60f) from the position corresponding to the bypass portion 36 in the blade height direction; With respect to the flow of the cooling fluid in the downstream passage 66 (cooling passage 60a or 60f), the plurality of cooling holes 70 or 72, the plurality of turbulators, and the downstream region R2 located downstream of the upstream region R1.
- 34 or the thermal barrier coating 86 have different values of parameters or average values of the parameters. Examples of parameters will be described later.
- the upstream region R1 located upstream of the position HB of the bypass portion 36 in the blade height direction and the upstream region R1 located downstream of the position HB The value of any of the above parameters is different between the downstream region R2 and the downstream region R2 (that is, the boundary between the upstream region R1 and the downstream region R2 is the position HB of the bypass portion 36 in the blade height direction) .
- the boundary between the upstream region R1 and the downstream region R2 may be displaced from the position HB of the bypass portion 36 in the blade height direction, for example, as in an embodiment shown in FIG. 11 to be described later.
- the temperature of the cooling fluid supplied to the serpentine flow path 61 increases toward the downstream side, and the cooling capacity of the cooling fluid decreases.
- the upstream passage 65 and the downstream passage 66 that are adjacent to each other are communicated through the bypass portion 36. Therefore, the bypass portion 36 in the downstream passage 66
- the downstream region (downstream region R2) can be supplied with relatively low-temperature cooling fluid before the temperature rises. Therefore, it is possible to effectively cool the blade body 42 in the vicinity of the downstream region R2.
- the plurality of cooling holes 70 are provided in the upstream region (upstream region R1) and the downstream region (downstream region R2) of the bypass portion 36 in the downstream passage 66. or 72, the parameters characterizing the plurality of turbulators 34 or thermal barrier coatings 86 are different. Therefore, between the upstream region R1 and the downstream region R2 in the downstream passage 66, the amount of heat removed from the turbine blades 40 by the cooling fluid or the amount of gas flowing through the combustion gas passage 28 in which the turbine blades 40 are arranged (combustion gas, etc.) to the turbine blades 40 can be differentiated.
- FIG. 4 and 6 to 9 are schematic cross-sectional views each showing an enlarged trailing edge side portion of the turbine blade 40 according to one embodiment.
- 5 and 10 are schematic cross-sectional views showing enlarged leading edge side portions of the turbine blade 40 according to one embodiment.
- the above parameter may be a pitch Ph in the blade height direction between a pair of cooling holes 70 or 72 adjacent in the blade height direction among the plurality of cooling holes 70 or 72.
- the pitch Ph2 of the adjacent pair of cooling holes 70 or 72 in the downstream region R2 is greater than the pitch Ph1 in the upstream region R1, for example as shown in FIG. 4 or FIG.
- the plurality of cooling holes 70 or 72 arranged along the blade height direction may have the same diameter.
- the pitch of a pair of adjacent cooling holes in the blade height direction is the distance in the blade height direction between the centers of a pair of adjacent cooling holes.
- the parameter mentioned above may be the diameter D of the plurality of cooling holes 70 or 72 .
- the diameter D2 in the downstream region R2 of the plurality of cooling holes 70 or 72 is smaller than the diameter D1 in the upstream region R1, for example as shown in FIG.
- the pitch in the blade height direction of the plurality of cooling holes 70 or 72 arranged along the blade height direction may be the same.
- the diameter of the cooling hole may be the equivalent diameter (hydraulic diameter) of the cooling hole.
- the parameter mentioned above may be the opening density of the plurality of cooling holes 70 or 72 .
- the density of apertures in downstream region R2 is less than the density of apertures in upstream region R1.
- the ratio Ph/D of the pitch Ph of the cooling holes 70 in the blade height direction to the diameter D of the cooling holes may be adopted.
- a ratio S/Ph between the wetted edge length S at the open end to (that is, the perimeter of the open end on the surface of the blade body 42) and the pitch Ph of the cooling holes in the blade height direction may be adopted, or The number of cooling holes per unit length of the blade body 42 in the blade height direction may be employed.
- the pitch Ph in the blade height direction of the plurality of cooling holes 70 or 72 communicating with the downstream passage 66 and opening on the surface of the blade body 42 is greater in the downstream region R2 than in the upstream region R1. big.
- the diameter D of the plurality of cooling holes 70 or 72 described above is smaller in the downstream region R2 than in the upstream region R1.
- the opening density of the plurality of cooling holes 70 or 72 described above is lower in the downstream region R2 than in the upstream region R1.
- the amount of heat removed from the turbine blades 40 by the cooling fluid in the downstream region R2 is reduced compared to the case where the pitch Ph, diameter D, or aperture density is the same in the downstream region R2 and the upstream region R1. be able to.
- the amount of heat removed from the turbine blades 40 by the cooling fluid in the upstream region R1 is increased compared to the case where the above-described pitch Ph, diameter D, or aperture density is the same in the downstream region R2 and the upstream region R1. be able to. Therefore, overcooling in the downstream region R2 or insufficient cooling in the upstream region R1 of the blade body 42 near the downstream passage 66 can be suppressed.
- the above parameter may be the surface roughness of the inner wall surface 71 (see FIG. 4) of the plurality of cooling holes 70 or 72.
- the surface roughness may be an arithmetic mean roughness Ra.
- the surface roughness in the downstream region R2 of the plurality of cooling holes 70 or 72 is less than the surface roughness in the upstream region R1.
- the surface roughness of the inner wall surfaces 71 of the plurality of cooling holes 70 or 72 that communicate with the downstream passage 66 and open on the surface of the blade body 42 is greater in the downstream region R2 than in the upstream region R1. small. Therefore, compared to the case where the surface roughness of the inner wall surface 71 of the plurality of cooling holes 70 or 72 is the same in the downstream region R2 and the upstream region R1, the cooling fluid from the turbine blade 40 in the downstream region R2 The amount of heat to be removed can be reduced.
- the cooling fluid from the turbine blade 40 in the upstream region R1 A large amount of heat can be removed. Therefore, overcooling in the downstream region R2 or insufficient cooling in the upstream region R1 of the blade body 42 near the downstream passage 66 can be suppressed.
- multiple turbulators 34 are provided such that the heat transfer coefficient between the cooling fluid and the inner wall surface 63 of the downstream passage 66 is less in the downstream region R2 than in the upstream region R1.
- the above parameter may be the pitch PT of the plurality of turbulators 34 in the wing height direction.
- the pitch PT2 in the downstream region R2 of the plurality of turbulators 34 is greater than the pitch PT1 in the upstream region R1, for example as shown in FIG.
- the pitch PT of the turbulators 34 in the blade height direction is the distance in the blade height direction between the centers of a pair of turbulators 34 adjacent in the blade height direction (see FIG. 8).
- the heat transfer coefficient between the cooling fluid and the turbine blades 40 tends to increase as the pitch PT of the turbulators 34 decreases.
- the above parameter may be the height e of the turbulators 34 relative to the inner wall surface 63 of the downstream passage 66 .
- height e2 in downstream region R2 of turbulator 34 is less than height e1 in upstream region R1, for example as shown in FIG.
- the parameters described above are the angle ⁇ (However, ⁇ is 0 degrees or more and 90 degrees or less. Hereinafter, it may also be referred to as the inclination angle ⁇ of the turbulator 34). That is, in some embodiments, the tilt angle ⁇ of the turbulators 34 differs between the downstream region R2 and the upstream region R1.
- the inner wall surface 63 of the downstream passage 66 is provided with a plurality of turbulators 34 arranged along the blade height direction, and the heat transfer coefficient between the inner wall surface 63 and the cooling fluid is
- the downstream region R2 is smaller than the upstream region R1.
- the pitch PT2 in the downstream region R2 of the multiple turbulators 34 is greater than the pitch PT1 in the upstream region R1.
- the height e2 in the downstream region R2 of the turbulator 34 is smaller than the height e1 in the upstream region R1.
- the inclination angle ⁇ 2 in the downstream region R2 of the turbulator 34 is different from the inclination angle ⁇ 1 in the upstream region R1.
- the pitch PT of the turbulators 34, the height e of the turbulators 34, or the angle ⁇ of the turbulators 34 are the same in the downstream region and the upstream region, can reduce the amount of heat removed from the turbine blades 40 by the cooling fluid at .
- the amount of heat removed from the turbine blades 40 by the cooling fluid in the upstream region can be increased compared to the case where the heat transfer coefficient is the same in the downstream region R2 and the upstream region R1. Therefore, overcooling in the downstream region R2 or insufficient cooling in the upstream region R1 of the blade body 42 near the downstream passage 66 can be suppressed.
- the parameter mentioned above may be the thickness T of the thermal barrier coating 86 .
- the thickness T2 of the thermal barrier coating 86 in the downstream region R2 is less than the thickness T1 of the thermal barrier coating 86 in the upstream region R1, for example as shown in FIG.
- the thickness of the thermal barrier coating 86 may be different as described above only in a part of the surface of the turbine blade 40 .
- the thermal barrier coating in the downstream region R2 The thickness T2 of 86 may be smaller than the thickness T1 of the thermal barrier coating 86 in the upstream region R1.
- the thickness T of the thermal barrier coating 86 can be adjusted, for example, by changing the moving speed of the thermal spray gun when the thermal barrier coating 86 is applied. That is, the faster the thermal spray gun moves, the smaller the thickness of the thermal barrier coating to be formed.
- the thickness of the thermal barrier coating 86 covering the surface of the wing body 42 is smaller in the downstream region R2 than in the upstream region R1. Therefore, compared to the case where the thickness T of the thermal barrier coating 86 is the same in the downstream region R2 and the upstream region R1, the metal temperature of the blade body 42 can be made uniform in the blade height direction. Therefore, overcooling in the downstream region R2 or insufficient cooling in the upstream region R1 of the blade body 42 near the downstream passage 66 can be suppressed.
- FIG. 11 is a schematic cross-sectional view showing an enlarged trailing edge side portion of the turbine blade 40 according to one embodiment.
- the bypass section 36 includes a cooling passage 60a or 60f (downstream passage 66) located closest to the leading edge side or the trailing edge of the plurality of cooling passages 60, and the cooling passage 60a or 60f. It is provided in the partition wall portion 32 that separates the adjacent cooling passage 60b or 60e (upstream passage 65).
- the bypass section 36 includes a cooling passage 60f (downstream passage 66) positioned closest to the trailing edge and a cooling passage 60e (upstream passage) next to the cooling passage 60f.
- the cooling passage 60a or 60f positioned closest to the leading edge or trailing edge usually has an outlet opening 64A or 64B at its downstream end (see FIG. 2).
- the direction of the cooling fluid (arrow Fa in FIG. 11) flowing from the upstream passage 65 into the downstream passage 66 via the bypass portion 36 is greatly bent downstream, and the cooling fluid flows into the downstream passage 66. , it is supplied to a position on the downstream side of the bypass section 36 .
- the position where the parameter value indicating the characteristics of the plurality of cooling holes 70 or 72, the plurality of turbulators 34, or the thermal barrier coating 86 is changed is provided downstream of the bypass section 36. . Therefore, even if the direction of the cooling fluid flowing into the downstream passage 66 is greatly bent downstream as described above, the position HB' on the downstream side of the bypass portion 36 (the position corresponding to the bypass portion 36) Between the upstream region (upstream region R1) and the downstream region (downstream region R2), the amount of heat removed from the turbine blades 40 by the cooling fluid, or the combustion gas passage 28 in which the turbine blades 40 are arranged.
- a turbine blade (40) according to at least one embodiment of the present invention, a wing body (42); a plurality of cooling passages (60) each extending along the blade height direction inside the blade body and connected to each other via folded portions (59); a bypass portion (36) provided in a partition wall portion (32) partitioning a pair of adjacent cooling passages among the plurality of cooling passages and allowing the pair of cooling passages to communicate with each other; with
- the pair of cooling passages is a turbine blade including an upstream passage (65) and a downstream passage (66) located downstream of the upstream passage with respect to cooling fluid flow,
- the turbine blade a plurality of cooling holes (70 or 72) formed in the blade body so as to be arranged along the blade height direction, communicating with the downstream passage and opening to the surface of the blade body;
- a plurality of turbulators (34) provided on the inner wall surface (63) of the downstream passage and arranged along the blade height direction, or
- a thermal barrier coating (86) covering the surface of the wing body.
- upstream region (R1) positioned upstream with respect to the flow of the cooling fluid in the downstream passage relative to the position corresponding to the bypass portion in the blade height direction; and the flow of the cooling fluid in the downstream passage.
- the values of the parameters that characterize the plurality of cooling holes, the plurality of turbulators, or the thermal barrier coating differ between the downstream region (R2) located downstream of the upstream region with respect to .
- the upstream passage and the downstream passage that are adjacent to each other are communicated via the bypass portion, the region downstream of the bypass portion in the downstream passage (downstream region) , can supply a relatively cold cooling fluid before the temperature rises. Therefore, the blade body in the vicinity of the downstream region can be effectively cooled.
- the upstream region (upstream region) and the downstream region (downstream region) of the bypass section in the downstream passage have a plurality of cooling holes, a plurality of turbulators, or heat shields. The values of the parameters that characterize the coating are different.
- the turbine blades can be effectively cooled by providing the bypass section, and overcooling or insufficient cooling of the turbine blades due to the installation of the bypass section can be suppressed.
- a plurality of cooling holes (70 or 72) formed in the blade body so as to be arranged along the blade height direction, communicating with the downstream passage and opening to the surface of the blade body; An opening density of the plurality of cooling holes is smaller in the downstream region than in the upstream region.
- the opening density of the plurality of cooling holes communicating with the downstream passage and opening on the surface of the blade body is lower in the downstream region than in the upstream region. Therefore, the amount of heat removed from the turbine blades by the cooling fluid in the downstream area can be reduced compared to the case where the above-described opening density is the same in the downstream area and the upstream area. Alternatively, the amount of heat removed from the turbine blades by the cooling fluid in the upstream region can be increased compared to the case where the above-described aperture density is the same in the downstream region and the upstream region. Therefore, overcooling in the downstream region or insufficient cooling in the upstream region can be suppressed for the blade body near the downstream passage.
- a plurality of cooling holes formed in the blade body so as to be arranged along the blade height direction, communicating with the downstream passage and opening to the surface of the blade body;
- a pitch (Ph) in the blade height direction of a pair of cooling holes adjacent in the blade height direction is larger in the downstream region than in the upstream region.
- the pitch in the blade height direction of the plurality of cooling holes communicating with the downstream passage and opening onto the surface of the blade body is larger in the downstream region than in the upstream region. Therefore, the amount of heat removed from the turbine blades by the cooling fluid in the downstream region can be reduced compared to the case where the pitch of the plurality of cooling holes is the same in the downstream region and the upstream region. Alternatively, the amount of heat removed from the turbine blades by the cooling fluid in the upstream region can be increased compared to the case where the pitch of the plurality of cooling holes is the same in the downstream region and the upstream region. Therefore, overcooling in the downstream region or insufficient cooling in the upstream region can be suppressed for the blade body near the downstream passage.
- a plurality of cooling holes formed in the blade body so as to be arranged along the blade height direction, communicating with the downstream passage and opening to the surface of the blade body;
- a diameter (D) of the plurality of cooling holes is smaller in the downstream region than in the upstream region.
- the diameters of the plurality of cooling holes communicating with the downstream passage and opening onto the surface of the blade body are smaller in the downstream region than in the upstream region. Therefore, the amount of heat removed from the turbine blades by the cooling fluid in the downstream region can be reduced compared to the case where the diameters of the plurality of cooling holes are the same in the downstream region and the upstream region.
- the amount of heat removed from the turbine blades by the cooling fluid in the upstream region can be increased compared to the case where the diameters of the plurality of cooling holes are the same in the downstream region and the upstream region. Therefore, overcooling in the downstream region or insufficient cooling in the upstream region can be suppressed for the blade body near the downstream passage.
- a plurality of cooling holes formed in the blade body so as to be arranged along the blade height direction, communicating with the downstream passage and opening to the surface of the blade body; Surface roughness of inner wall surfaces of the plurality of cooling holes is smaller in the downstream region than in the upstream region.
- the surface roughness of the inner wall surfaces of the plurality of cooling holes communicating with the downstream passage and opening onto the surface of the blade body is smaller in the downstream region than in the upstream region. Therefore, the amount of heat removed from the turbine blades by the cooling fluid in the downstream region can be reduced compared to the case where the surface roughness of the inner wall surfaces of the plurality of cooling holes is the same in the downstream region and the upstream region. . Alternatively, the amount of heat removed from the turbine blades by the cooling fluid in the upstream region can be increased compared to the case where the surface roughness of the inner wall surfaces of the plurality of cooling holes is the same in the downstream region and the upstream region. . Therefore, overcooling in the downstream region or insufficient cooling in the upstream region can be suppressed for the blade body near the downstream passage.
- the inner wall surface of the downstream passage is provided with a plurality of turbulators arranged along the blade height direction, and the heat transfer coefficient between the inner wall surface and the cooling fluid is smaller in the downstream region than in the upstream region. Therefore, the amount of heat removed from the turbine blades by the cooling fluid in the downstream region can be reduced compared to the case where the heat transfer coefficient is the same in the downstream region and the upstream region. Alternatively, the amount of heat removed from the turbine blades by the cooling fluid in the upstream region can be increased compared to the case where the heat transfer coefficient is the same in the downstream region and the upstream region. Therefore, overcooling in the downstream region or insufficient cooling in the upstream region can be suppressed for the blade body near the downstream passage.
- a pitch (PT) of the plurality of turbulators in the blade height direction is larger in the downstream region than in the upstream region.
- the inner wall surface of the downstream passage is provided with a plurality of turbulators arranged along the blade height direction, and the pitch of the plurality of turbulators in the blade height direction is equal to the downstream side larger in area than in the upstream area. Therefore, the amount of heat removed from the turbine blades by the cooling fluid in the downstream region can be reduced compared to the case where the pitches of the plurality of turbulators are the same in the downstream region and the upstream region. Alternatively, the amount of heat removed from the turbine blades by the cooling fluid in the upstream region can be increased compared to when the pitches of the plurality of turbulators are the same in the downstream region and the upstream region. Therefore, overcooling in the downstream region or insufficient cooling in the upstream region can be suppressed for the blade body near the downstream passage.
- a plurality of turbulators arranged along the blade height direction are provided on the inner wall surface of the downstream passage, and the height of the turbulators relative to the inner wall surface is equal to the downstream smaller in the side region than in the upstream region. Therefore, the amount of heat removed from the turbine blades by the cooling fluid in the downstream region can be reduced compared to the case where the turbulators have the same height in the downstream region and the upstream region. Alternatively, the amount of heat removed from the turbine blades by the cooling fluid in the upstream region can be increased compared to the case where the turbulators have the same height in the downstream region and the upstream region. Therefore, overcooling in the downstream region or insufficient cooling in the upstream region can be suppressed for the blade body near the downstream passage.
- a plurality of turbulators arranged along the blade height direction are provided on the inner wall surface of the downstream passage, and the flow direction of the cooling fluid in the downstream passage and the extension of the turbulators are controlled.
- the angle with respect to the existing direction is different between the downstream area and the upstream area. Therefore, the amount of heat removed from the turbine blades by the cooling fluid in the downstream region can be reduced compared to the case where the above angle is the same in the downstream region and the upstream region.
- the amount of heat removed from the turbine blades by the cooling fluid in the upstream region can be increased compared to the case where the above angle is the same in the downstream region and the upstream region. Therefore, overcooling in the downstream region or insufficient cooling in the upstream region can be suppressed for the blade body near the downstream passage.
- thermo barrier coating (86) covering the surface of the wing body A thermal barrier coating (86) covering the surface of the wing body, The thickness (T) of the thermal barrier coating is smaller in the downstream region than in the upstream region.
- the thickness of the thermal barrier coating that covers the surface of the blade body is smaller in the downstream area than in the upstream area. Therefore, compared to the case where the thickness of the thermal barrier coating is the same in the downstream region and the upstream region, the metal temperature of the blade body can be made uniform in the blade height direction. Therefore, overcooling in the downstream region or insufficient cooling in the upstream region can be suppressed for the blade body near the downstream passage.
- the downstream passage is a cooling passage (60a) positioned closest to the leading edge or a cooling passage (60f) positioned closest to the trailing edge in the chord direction of the blade body among the plurality of cooling passages;
- the position corresponding to the bypass portion in the blade height direction is located downstream of the bypass portion with respect to the flow of the cooling fluid in the downstream passage.
- the direction of the cooling fluid flowing from the upstream side passage to the downstream side passage via the bypass portion is greatly bent to the downstream side. is supplied to a position downstream of the bypass section in the downstream passage.
- the position where the value of the parameter indicating the characteristics of the plurality of cooling holes, the plurality of turbulators, or the thermal barrier coating is changed is provided downstream of the bypass section.
- the region ( between the upstream region) and the downstream region (downstream region), the amount of heat removed from the turbine blades by the cooling fluid, or the amount of heat input to the turbine blades from the gas flowing through the gas passage in which the turbine blades 40 are arranged. can make a difference. Therefore, overcooling in the downstream region or insufficient cooling in the upstream region can be effectively suppressed for the blade body in the vicinity of the downstream passage.
- a gas turbine (1) according to at least one embodiment of the present invention, a turbine blade (40) according to any one of (1) to (11) above; a combustor (4) for producing combustion gas flowing through a combustion gas flow path in which said turbine blades are provided.
- the region downstream of the bypass portion in the downstream passage can supply a relatively cold cooling fluid before the temperature rises. Therefore, the blade body in the vicinity of the downstream region can be effectively cooled.
- the upstream region (upstream region) and the downstream region (downstream region) of the bypass section in the downstream passage have a plurality of cooling holes, a plurality of turbulators, or heat shields. The values of the parameters that characterize the coating are different.
- the turbine blades can be effectively cooled by providing the bypass section, and overcooling or insufficient cooling of the turbine blades due to the installation of the bypass section can be suppressed.
- expressions such as “in a certain direction”, “along a certain direction”, “parallel”, “perpendicular”, “central”, “concentric” or “coaxial”, etc. express relative or absolute arrangements. represents not only such arrangement strictly, but also the state of being relatively displaced with a tolerance or an angle or distance to the extent that the same function can be obtained.
- expressions such as “identical”, “equal”, and “homogeneous”, which express that things are in the same state not only express the state of being strictly equal, but also have tolerances or differences to the extent that the same function can be obtained. It shall also represent the existing state.
- expressions representing shapes such as a quadrilateral shape and a cylindrical shape not only represent shapes such as a quadrilateral shape and a cylindrical shape in a geometrically strict sense, but also within the range in which the same effect can be obtained. , a shape including an uneven portion, a chamfered portion, and the like.
- the expressions “comprising”, “including”, or “having” one component are not exclusive expressions excluding the presence of other components.
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Materials Engineering (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
Description
本願は、2021年12月7日に日本国特許庁に出願された特願2021-198469号に基づき優先権を主張し、その内容をここに援用する。
翼体と、
前記翼体の内部においてそれぞれ翼高さ方向に沿って延在し、折り返し部を介して互いに接続される複数の冷却通路と、
前記複数の冷却通路のうち隣り合う一対の冷却通路を仕切る隔壁部に設けられ、前記一対の冷却通路を互いに連通させるバイパス部と、
を備え、
前記一対の冷却通路は、上流側通路と、冷却流体の流れに関して前記上流側通路の下流側に位置する下流側通路と、を含む
タービン翼であって、
前記タービン翼は、
前記翼高さ方向に沿って配列するように前記翼体に形成され、前記下流側通路に連通するとともに前記翼体の表面に開口する複数の冷却孔、
前記下流側通路の内壁面に設けられ、前記翼高さ方向に沿って配列された複数のタービュレータ、又は、
前記翼体の表面を覆う遮熱コーティング
を備え、
前記翼高さ方向において、前記バイパス部に対応する位置よりも前記下流側通路における前記冷却流体の流れに関して上流側に位置する上流側領域と、前記下流側通路における前記冷却流体の流れに関して前記上流側領域よりも下流側に位置する下流側領域との間で、前記複数の冷却孔、前記複数のタービュレータ又は前記遮熱コーティングの特徴を示すパラメータの値が異なる。
上述のタービン翼と、
前記タービン翼が設けられる燃焼ガス通路を流れる燃焼ガスを生成するための燃焼器と、を備える。
図1は、幾つかの実施形態に係るタービン翼が適用されるタービンの一例としてのガスタービンの概略図である。
図2は、一実施形態に係るタービン翼40(動翼26)の翼高さ方向に沿った概略的な断面図であり、図3は、図2のタービン翼40のA-A断面を示す図である。図中の矢印は、冷却流体の流れの向きを示す。なお、図2において、遮熱コーティング86(図3参照)の図示が省略されている。
また、上述の構成を有するタービン翼40では、下流側通路66におけるバイパス部36の上流側の領域(上流側領域R1)と下流側の領域(下流側領域R2)とで、複数の冷却孔70又は72、複数のタービュレータ34又は遮熱コーティング86の特徴を示すパラメータが異なる。よって、下流側通路66における上流側領域R1と下流側領域R2との間で、冷却流体によるタービン翼40からの除熱量、又は、タービン翼40が配置される燃焼ガス通路28を流れるガス(燃焼ガス等)からタービン翼40への入熱量に差をつけることができる。したがって、下流側通路66近傍の翼体42について、下流側領域R2における過冷却、又は、上流側領域R1における冷却不足を抑制することができる。
すなわち、上述の構成を有するタービン翼40よれば、バイパス部36を設けることでタービン翼40を効果的に冷却可能であるとともに、バイパス部36の設置に伴うタービン翼40の過冷却又は冷却不足を抑制することができる。
翼体(42)と、
前記翼体の内部においてそれぞれ翼高さ方向に沿って延在し、折り返し部(59)を介して互いに接続される複数の冷却通路(60)と、
前記複数の冷却通路のうち隣り合う一対の冷却通路を仕切る隔壁部(32)に設けられ、前記一対の冷却通路を互いに連通させるバイパス部(36)と、
を備え、
前記一対の冷却通路は、上流側通路(65)と、冷却流体の流れに関して前記上流側通路の下流側に位置する下流側通路(66)と、を含む
タービン翼であって、
前記タービン翼は、
前記翼高さ方向に沿って配列するように前記翼体に形成され、前記下流側通路に連通するとともに前記翼体の表面に開口する複数の冷却孔(70又は72)、
前記下流側通路の内壁面(63)に設けられ、前記翼高さ方向に沿って配列された複数のタービュレータ(34)、又は、
前記翼体の表面を覆う遮熱コーティング(86)
を備え、
前記翼高さ方向において、前記バイパス部に対応する位置よりも前記下流側通路における前記冷却流体の流れに関して上流側に位置する上流側領域(R1)と、前記下流側通路における前記冷却流体の流れに関して前記上流側領域よりも下流側に位置する下流側領域(R2)との間で、前記複数の冷却孔、前記複数のタービュレータ又は前記遮熱コーティングの特徴を示すパラメータの値が異なる。
また、上記(1)の構成では、下流側通路におけるバイパス部の上流側の領域(上流側領域)と下流側の領域(下流側領域)とで、複数の冷却孔、複数のタービュレータ又は遮熱コーティングの特徴を示すパラメータの値が異なる。よって、下流側通路における上流側領域と下流側領域との間で、冷却流体によるタービン翼からの除熱量、又は、タービン翼が配置されるガス通路を流れるガス(燃焼ガス等)からタービン翼への入熱量に差をつけることができる。したがって、下流側通路近傍の翼体について、下流側領域における過冷却、又は、上流側領域における冷却不足を抑制することができる。
すなわち、上記(1)の構成によれば、バイパス部を設けることでタービン翼を効果的に冷却可能であるとともに、バイパス部の設置に伴うタービン翼の過冷却又は冷却不足を抑制することができる。
前記タービン翼は、
前記翼高さ方向に沿って配列するように前記翼体に形成され、前記下流側通路に連通するとともに前記翼体の表面に開口する複数の冷却孔(70又は72)を備え、
前記複数の冷却孔の開口密度は、前記下流側領域において、前記上流側領域よりも小さい。
前記タービン翼は、
前記翼高さ方向に沿って配列するように前記翼体に形成され、前記下流側通路に連通するとともに前記翼体の表面に開口する複数の冷却孔を備え、
前記翼高さ方向にて隣り合う一対の冷却孔の前記翼高さ方向におけるピッチ(Ph)は、前記下流側領域において、前記上流側領域よりも大きい。
前記タービン翼は、
前記翼高さ方向に沿って配列するように前記翼体に形成され、前記下流側通路に連通するとともに前記翼体の表面に開口する複数の冷却孔を備え、
前記複数の冷却孔の直径(D)は、前記下流側領域において、前記上流側領域よりも小さい。
前記タービン翼は、
前記翼高さ方向に沿って配列するように前記翼体に形成され、前記下流側通路に連通するとともに前記翼体の表面に開口する複数の冷却孔を備え、
前記複数の冷却孔の内壁面の表面粗さは、前記下流側領域において、前記上流側領域よりも小さい。
前記タービン翼は、
前記下流側通路の内壁面(63)に設けられ、前記翼高さ方向に沿って配列された複数のタービュレータ(34)を備え、
前記冷却流体と下流側通路の前記内壁面との間の熱伝達係数が、前記下流側領域において前記上流側領域よりも小さくなるように、前記複数のタービュレータが設けられている。
前記タービン翼は、
前記下流側通路の内壁面に設けられ、前記翼高さ方向に沿って配列された複数のタービュレータを備え、
前記複数のタービュレータの前記翼高さ方向におけるピッチ(PT)は、前記下流側領域において、前記上流側領域よりも大きい。
前記タービン翼は、
前記下流側通路の内壁面に設けられ、前記翼高さ方向に沿って配列された複数のタービュレータを備え、
前記下流側通路の前記内壁面を基準とする前記タービュレータの高さ(e)は、前記下流側領域において、前記上流側領域よりも小さい。
前記タービン翼は、
前記下流側通路の内壁面に設けられ、前記翼高さ方向に沿って配列された複数のタービュレータを備え、
前記下流側通路における前記冷却流体の流れの方向と、前記タービュレータの延在方向との間の角度(θ)は、前記下流側領域と、前記上流側領域とで異なる。
前記タービン翼は、
前記翼体の表面を覆う遮熱コーティング(86)を備え、
前記遮熱コーティングの厚さ(T)は、前記下流側領域において、前記上流側領域よりも小さい。
前記下流側通路は、前記複数の冷却通路のうち、前記翼体のコード方向において最も前縁側に位置する冷却通路(60a)又は最も後縁側に位置する冷却通路(60f)であり、
前記翼高さ方向において前記バイパス部に対応する前記位置は、前記下流側通路における前記冷却流体の流れに関して前記バイパス部よりも下流側に位置する。
上記(1)乃至(11)の何れか一項に記載のタービン翼(40)と、
前記タービン翼が設けられる燃焼ガス流路を流れる燃焼ガスを生成するための燃焼器(4)と、を備える。
また、上記(12)の構成では、下流側通路におけるバイパス部の上流側の領域(上流側領域)と下流側の領域(下流側領域)とで、複数の冷却孔、複数のタービュレータ又は遮熱コーティングの特徴を示すパラメータの値が異なる。よって、下流側通路における上流側領域と下流側領域との間で、冷却流体によるタービン翼からの除熱量、又は、タービン翼が配置されるガス通路を流れるガス(燃焼ガス等)からタービン翼への入熱量に差をつけることができる。したがって、下流側通路近傍の翼体について、下流側領域における過冷却、又は、上流側領域における冷却不足を抑制することができる。
すなわち、上記(12)の構成によれば、バイパス部を設けることでタービン翼を効果的に冷却可能であるとともに、バイパス部の設置に伴うタービン翼の過冷却又は冷却不足を抑制することができる。
例えば、「同一」、「等しい」及び「均質」等の物事が等しい状態であることを表す表現は、厳密に等しい状態を表すのみならず、公差、若しくは、同じ機能が得られる程度の差が存在している状態も表すものとする。
また、本明細書において、四角形状や円筒形状等の形状を表す表現は、幾何学的に厳密な意味での四角形状や円筒形状等の形状を表すのみならず、同じ効果が得られる範囲で、凹凸部や面取り部等を含む形状も表すものとする。
また、本明細書において、一の構成要素を「備える」、「含む」、又は、「有する」という表現は、他の構成要素の存在を除外する排他的な表現ではない。
2 圧縮機
4 燃焼器
6 タービン
8 ロータ
10 圧縮機車室
12 空気取入口
16 静翼
18 動翼
20 ケーシング
22 タービン車室
24 静翼
26 動翼
28 燃焼ガス通路
30 排気室
32 隔壁部
34 タービュレータ
36 バイパス部
40 タービン翼
42 翼体
44 前縁
46 後縁
47 後縁部
48 先端
49 天板
50 基端
56 圧力面
58 負圧面
59 折り返し部
60,60a~60f 冷却通路
61,61A,61B サーペンタイン流路
63 内壁面
64,64A,64B 出口開口
65 上流側通路
66 下流側通路
70 冷却孔
71 内壁面
72 冷却孔
80 プラットフォーム
82 翼根部
84A 内部流路
84B 内部流路
86 遮熱コーティング
PT タービュレータのピッチ
Ph 冷却孔のピッチ
R1 上流側領域
R2 下流側領域
T 厚さ
e 高さ
θ 角度(傾き角)
Claims (12)
- 翼体と、
前記翼体の内部においてそれぞれ翼高さ方向に沿って延在し、折り返し部を介して互いに接続される複数の冷却通路と、
前記複数の冷却通路のうち隣り合う一対の冷却通路を仕切る隔壁部に設けられ、前記一対の冷却通路を互いに連通させるバイパス部と、
を備え、
前記一対の冷却通路は、上流側通路と、冷却流体の流れに関して前記上流側通路の下流側に位置する下流側通路と、を含む
タービン翼であって、
前記タービン翼は、
前記翼高さ方向に沿って配列するように前記翼体に形成され、前記下流側通路に連通するとともに前記翼体の表面に開口する複数の冷却孔、
前記下流側通路の内壁面に設けられ、前記翼高さ方向に沿って配列された複数のタービュレータ、又は、
前記翼体の表面を覆う遮熱コーティング
を備え、
前記翼高さ方向において、前記バイパス部に対応する位置よりも前記下流側通路における前記冷却流体の流れに関して上流側に位置する上流側領域と、前記下流側通路における前記冷却流体の流れに関して前記上流側領域よりも下流側に位置する下流側領域との間で、前記複数の冷却孔、前記複数のタービュレータ又は前記遮熱コーティングの特徴を示すパラメータの値が異なる
タービン翼。 - 前記翼高さ方向に沿って配列するように前記翼体に形成され、前記下流側通路に連通するとともに前記翼体の表面に開口する複数の冷却孔を備え、
前記複数の冷却孔の開口密度は、前記下流側領域において、前記上流側領域よりも小さい
請求項1に記載のタービン翼。 - 前記翼高さ方向に沿って配列するように前記翼体に形成され、前記下流側通路に連通するとともに前記翼体の表面に開口する複数の冷却孔を備え、
前記翼高さ方向にて隣り合う一対の冷却孔の前記翼高さ方向におけるピッチは、前記下流側領域において、前記上流側領域よりも大きい
請求項1又は2に記載のタービン翼。 - 前記翼高さ方向に沿って配列するように前記翼体に形成され、前記下流側通路に連通するとともに前記翼体の表面に開口する複数の冷却孔を備え、
前記複数の冷却孔の直径は、前記下流側領域において、前記上流側領域よりも小さい
請求項1又は2に記載のタービン翼。 - 前記翼高さ方向に沿って配列するように前記翼体に形成され、前記下流側通路に連通するとともに前記翼体の表面に開口する複数の冷却孔を備え、
前記複数の冷却孔の内壁面の表面粗さは、前記下流側領域において、前記上流側領域よりも小さい
請求項1又は2に記載のタービン翼。 - 前記下流側通路の内壁面に設けられ、前記翼高さ方向に沿って配列された複数のタービュレータを備え、
前記冷却流体と下流側通路の前記内壁面との間の熱伝達係数が、前記下流側領域において前記上流側領域よりも小さくなるように、前記複数のタービュレータが設けられている
請求項1又は2に記載のタービン翼。 - 前記下流側通路の内壁面に設けられ、前記翼高さ方向に沿って配列された複数のタービュレータを備え、
前記複数のタービュレータの前記翼高さ方向におけるピッチは、前記下流側領域において、前記上流側領域よりも大きい
請求項1又は2に記載のタービン翼。 - 前記下流側通路の内壁面に設けられ、前記翼高さ方向に沿って配列された複数のタービュレータを備え、
前記下流側通路の前記内壁面を基準とする前記タービュレータの高さは、前記下流側領域において、前記上流側領域よりも小さい
請求項1又は2に記載のタービン翼。 - 前記下流側通路の内壁面に設けられ、前記翼高さ方向に沿って配列された複数のタービュレータを備え、
前記下流側通路における前記冷却流体の流れの方向と、前記タービュレータの延在方向との間の角度は、前記下流側領域と、前記上流側領域とで異なる
請求項1又は2に記載のタービン翼。 - 前記翼体の表面を覆う遮熱コーティングを備え、
前記遮熱コーティングの厚さは、前記下流側領域において、前記上流側領域よりも小さい
請求項1又は2に記載のタービン翼。 - 前記下流側通路は、前記複数の冷却通路のうち、前記翼体のコード方向において最も前縁側又は最も後縁側に位置する冷却通路であり、
前記翼高さ方向において前記バイパス部に対応する前記位置は、前記下流側通路における前記冷却流体の流れに関して前記バイパス部よりも下流側に位置する
請求項1又は2に記載のタービン翼。 - 請求項1又は2に記載のタービン翼と、
前記タービン翼が設けられる燃焼ガス通路を流れる燃焼ガスを生成するための燃焼器と、を備えるガスタービン。
Priority Applications (4)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
KR1020247006561A KR20240031436A (ko) | 2021-12-07 | 2022-11-25 | 터빈 날개 및 가스 터빈 |
DE112022004264.8T DE112022004264T5 (de) | 2021-12-07 | 2022-11-25 | Turbinenschaufel und gasturbine |
CN202280059704.9A CN117897549A (zh) | 2021-12-07 | 2022-11-25 | 涡轮叶片及燃气涡轮 |
JP2023566232A JPWO2023106125A1 (ja) | 2021-12-07 | 2022-11-25 |
Applications Claiming Priority (2)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
JP2021-198469 | 2021-12-07 | ||
JP2021198469 | 2021-12-07 |
Publications (1)
Publication Number | Publication Date |
---|---|
WO2023106125A1 true WO2023106125A1 (ja) | 2023-06-15 |
Family
ID=86730196
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
PCT/JP2022/043511 WO2023106125A1 (ja) | 2021-12-07 | 2022-11-25 | タービン翼及びガスタービン |
Country Status (5)
Country | Link |
---|---|
JP (1) | JPWO2023106125A1 (ja) |
KR (1) | KR20240031436A (ja) |
CN (1) | CN117897549A (ja) |
DE (1) | DE112022004264T5 (ja) |
WO (1) | WO2023106125A1 (ja) |
Citations (3)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
JPS58117303A (ja) * | 1981-12-28 | 1983-07-12 | ユナイテツド・テクノロジ−ズ・コ−ポレイシヨン | 冷却可能なエ−ロフオイル |
JP2010007463A (ja) * | 2008-06-24 | 2010-01-14 | Hitachi Ltd | ガスタービン翼 |
US20130343872A1 (en) * | 2011-02-17 | 2013-12-26 | Rolls-Royce Plc | Cooled component for the turbine of a gas turbine engine |
Family Cites Families (2)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
WO1996012874A1 (en) | 1994-10-24 | 1996-05-02 | Westinghouse Electric Corporation | Gas turbine blade with enhanced cooling |
US10260354B2 (en) | 2016-02-12 | 2019-04-16 | General Electric Company | Airfoil trailing edge cooling |
-
2022
- 2022-11-25 JP JP2023566232A patent/JPWO2023106125A1/ja active Pending
- 2022-11-25 DE DE112022004264.8T patent/DE112022004264T5/de active Pending
- 2022-11-25 CN CN202280059704.9A patent/CN117897549A/zh active Pending
- 2022-11-25 KR KR1020247006561A patent/KR20240031436A/ko unknown
- 2022-11-25 WO PCT/JP2022/043511 patent/WO2023106125A1/ja active Application Filing
Patent Citations (3)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
JPS58117303A (ja) * | 1981-12-28 | 1983-07-12 | ユナイテツド・テクノロジ−ズ・コ−ポレイシヨン | 冷却可能なエ−ロフオイル |
JP2010007463A (ja) * | 2008-06-24 | 2010-01-14 | Hitachi Ltd | ガスタービン翼 |
US20130343872A1 (en) * | 2011-02-17 | 2013-12-26 | Rolls-Royce Plc | Cooled component for the turbine of a gas turbine engine |
Also Published As
Publication number | Publication date |
---|---|
CN117897549A (zh) | 2024-04-16 |
KR20240031436A (ko) | 2024-03-07 |
JPWO2023106125A1 (ja) | 2023-06-15 |
DE112022004264T5 (de) | 2024-06-20 |
Similar Documents
Publication | Publication Date | Title |
---|---|---|
EP2716866B1 (en) | Gas turbine engine components with lateral and forward sweep film cooling holes | |
US20130315710A1 (en) | Gas turbine engine components with cooling hole trenches | |
US9650900B2 (en) | Gas turbine engine components with film cooling holes having cylindrical to multi-lobe configurations | |
US11199098B2 (en) | Flared central cavity aft of airfoil leading edge | |
JP6345319B1 (ja) | タービン翼及びガスタービン | |
JP3213107U (ja) | 翼形部のための衝突システム | |
US11624285B2 (en) | Airfoil and gas turbine having same | |
KR102467118B1 (ko) | 터빈 날개 및 가스 터빈 | |
JP2019183805A5 (ja) | ||
WO2023106125A1 (ja) | タービン翼及びガスタービン | |
JPS59231102A (ja) | ガスタ−ビンの翼 | |
JP6996947B2 (ja) | タービン翼及びガスタービン | |
JP7224928B2 (ja) | タービン動翼及びガスタービン | |
JP2021071085A (ja) | タービン翼及びこれを備えたガスタービン | |
JP2019085973A5 (ja) | ||
US12000304B2 (en) | Turbine blade and gas turbine | |
US11879357B2 (en) | Turbine blade for a gas turbine engine | |
KR20240055099A (ko) | 터빈 정익 | |
JPWO2023095721A5 (ja) |
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: 22904056 Country of ref document: EP Kind code of ref document: A1 |
|
ENP | Entry into the national phase |
Ref document number: 20247006561 Country of ref document: KR Kind code of ref document: A |
|
WWE | Wipo information: entry into national phase |
Ref document number: 1020247006561 Country of ref document: KR |
|
WWE | Wipo information: entry into national phase |
Ref document number: 202280059704.9 Country of ref document: CN |
|
ENP | Entry into the national phase |
Ref document number: 2023566232 Country of ref document: JP Kind code of ref document: A |