US20140216055A1 - Gas turbine - Google Patents

Gas turbine Download PDF

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Publication number
US20140216055A1
US20140216055A1 US14/343,369 US201214343369A US2014216055A1 US 20140216055 A1 US20140216055 A1 US 20140216055A1 US 201214343369 A US201214343369 A US 201214343369A US 2014216055 A1 US2014216055 A1 US 2014216055A1
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US
United States
Prior art keywords
transition piece
circumferential direction
combustor
turbine
inclination
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.)
Abandoned
Application number
US14/343,369
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English (en)
Inventor
Yasuro Sakamoto
Keisuke Matsuyama
Keizo Tsukagoshi
Masanori Yuri
Hiroaki KISHIDA
Shunsuke Torii
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 Power Ltd
Original Assignee
Mitsubishi Heavy Industries 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 filed Critical Mitsubishi Heavy Industries Ltd
Assigned to MITSUBISHI HEAVY INDUSTRIES, LTD. reassignment MITSUBISHI HEAVY INDUSTRIES, LTD. ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: KISHIDA, HIROAKI, MATSUYAMA, KEISUKE, SAKAMOTO, YASURO, TORII, SHUNSUKE, TSUKAGOSHI, KEIZO, YURI, MASANORI
Publication of US20140216055A1 publication Critical patent/US20140216055A1/en
Assigned to MITSUBISHI HITACHI POWER SYSTEMS, LTD. reassignment MITSUBISHI HITACHI POWER SYSTEMS, LTD. ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: MITSUBISHI HEAVY INDUSTRIES, LTD.
Abandoned legal-status Critical Current

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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/002—Wall structures
    • 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
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23D—BURNERS
    • F23D11/00—Burners using a direct spraying action of liquid droplets or vaporised liquid into the combustion space
    • F23D11/10—Burners using a direct spraying action of liquid droplets or vaporised liquid into the combustion space the spraying being induced by a gaseous medium, e.g. water vapour
    • F23D11/106—Burners using a direct spraying action of liquid droplets or vaporised liquid into the combustion space the spraying being induced by a gaseous medium, e.g. water vapour medium and fuel meeting at the burner outlet
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23D—BURNERS
    • F23D11/00—Burners using a direct spraying action of liquid droplets or vaporised liquid into the combustion space
    • F23D11/36—Details, e.g. burner cooling means, noise reduction means
    • F23D11/38—Nozzles; Cleaning devices therefor
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
    • F23R3/16—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration with devices inside the flame tube or the combustion chamber to influence the air or gas flow
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/28—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
    • F23R3/283—Attaching or cooling of fuel injecting means including supports for fuel injectors, stems, or lances
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/28—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
    • F23R3/36—Supply of different fuels
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/42—Continuous combustion chambers using liquid or gaseous fuel characterised by the arrangement or form of the flame tubes or combustion chambers
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/42—Continuous combustion chambers using liquid or gaseous fuel characterised by the arrangement or form of the flame tubes or combustion chambers
    • F23R3/46—Combustion chambers comprising an annular arrangement of several essentially tubular flame tubes within a common annular casing or within individual casings
    • 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
    • F05D2250/00—Geometry
    • F05D2250/30—Arrangement of components
    • F05D2250/32—Arrangement of components according to their shape
    • F05D2250/324—Arrangement of components according to their shape divergent
    • 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
    • F05D2250/00—Geometry
    • F05D2250/50—Inlet or outlet
    • F05D2250/52—Outlet
    • 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
    • F05D2250/00—Geometry
    • F05D2250/70—Shape
    • F05D2250/71—Shape curved
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23C—METHODS OR APPARATUS FOR COMBUSTION USING FLUID FUEL OR SOLID FUEL SUSPENDED IN  A CARRIER GAS OR AIR 
    • F23C2201/00—Staged combustion
    • F23C2201/20—Burner staging
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23C—METHODS OR APPARATUS FOR COMBUSTION USING FLUID FUEL OR SOLID FUEL SUSPENDED IN  A CARRIER GAS OR AIR 
    • F23C2900/00—Special features of, or arrangements for combustion apparatus using fluid fuels or solid fuels suspended in air; Combustion processes therefor
    • F23C2900/06043—Burner staging, i.e. radially stratified flame core burners
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23D—BURNERS
    • F23D2214/00—Cooling
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R2900/00—Special features of, or arrangements for continuous combustion chambers; Combustion processes therefor
    • F23R2900/00014—Reducing thermo-acoustic vibrations by passive means, e.g. by Helmholtz resonators
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R2900/00—Special features of, or arrangements for continuous combustion chambers; Combustion processes therefor
    • F23R2900/03341—Sequential combustion chambers or burners

Definitions

  • the present invention relates to a gas turbine including a plurality of combustors configured to mix and combust fuel with compressed air to generate combustion gas, and a turbine having a rotor that is rotated by the combustion gas from the plurality of combustors, and particularly, to a transition piece of a combustor.
  • Gas turbines include a compressor that brings in the outside air to generate compressed air, a plurality of combustors that mix and combust fuel with the compressed air to generate combustion gas, and a turbine having a rotor that is rotated by the combustion gas from the plurality of combustors.
  • the plurality of combustors are annularly arranged with a rotor as a center.
  • Each combustor has a transition piece through which combustion gas is flowed to a gas inlet of the turbine.
  • the combustion gas flows out of the transition piece of the combustor, the combustion gas enters a combustion gas flow channel of the turbine from the gas inlet of the turbine.
  • a Karman's vortex street may be formed in the flow of the combustion gas immediately after flowing out of the transition piece, an unsteady pressure fluctuation that has this Karman's vortex street as a vibration source may resonate at an acoustic eigenvalue, and a large pressure fluctuation may occur, which leads to an operating load.
  • a large pressure fluctuation is suppressed by limiting a dimension in an axis direction between a downstream end of a transition piece and an upstream end of a first stage turbine vane, a dimension in a circumferential direction between the upstream end of the first stage turbine vane and the center between transition pieces that are adjacent to each other in the circumferential direction centered on a rotor, or the like to specific ranges.
  • the technique described in the above PTL 1 can reliably suppress a large pressure fluctuation at the downstream portion of the transition piece. However, it is desired to further suppress the pressure fluctuation at the downstream portion of the transition piece and to further enhance the gas turbine efficiency.
  • an object of the invention is to provide a gas turbine that can further suppress pressure fluctuation at a downstream portion of a transition piece of a combustor and can further enhance gas turbine efficiency, so as to meet such requests.
  • a gas turbine includes a plurality of combustors configured to mix and combust fuel with compressed air to generate combustion gas; and a turbine having a rotor that is rotated by the combustion gas from the plurality of combustors.
  • the plurality of combustors are annularly arranged with the rotor as a center and have a transition piece through which combustion gas is flowed to a gas inlet of the turbine.
  • An inner surface of at least one lateral wall of a pair of lateral walls that constitute a downstream portion of the transition piece of the combustor and face each other in a circumferential direction of the rotor forms an inclination surface that inclines down to a downstream end of the transition piece in a direction approaching the transition piece of another adjacent combustor that gradually draws closer as it goes to the downstream side of the transition piece in an axis direction.
  • a Karman's vortex street may be formed on the downstream side of a downstream end surface of the transition piece.
  • the inner surfaces of the lateral walls on the downstream side of the transition piece form inclination surfaces down to the downstream end of the transition piece. Therefore, flows along the inner surfaces of the lateral walls of the transition piece of the combustor that are adjacent to each other join each other at an angle on the downstream side of the downstream end surface of the transition piece. Therefore, formation of the Karman's vortex street on the downstream side of the downstream end surface of the transition piece can be suppressed, and the pressure fluctuation of the downstream portion of the transition piece can be suppressed.
  • the turbine may include a plurality of first stage turbine vanes arranged annularly along the gas inlet with the rotor as a center, and each of the first stage turbine vanes may be formed such that a chord direction in which a chord extends inclines with respect to the circumferential direction.
  • a side to which a downstream end of the first stage turbine vane is located with respect to an upstream end of the first stage turbine vane in the circumferential direction is defined as a blade inclination side
  • at least one lateral wall of the transition piece may be a lateral wall on the blade inclination side out of the pair of lateral walls of the transition piece facing each other in the circumferential direction.
  • the chord is a line segment that connects the upstream end and downstream end of a turbine vane.
  • both of the inner surfaces of the pair of lateral walls of the transition piece facing each other in the circumferential direction may form the inclination surfaces.
  • a ratio A/B of a dimension A in the axis direction from an upstream end of the inclination surface to a downstream end thereof, to a dimension B in the circumferential direction from the upstream end of the inclination surface to the downstream end thereof may be 1 to 8.
  • the inclination surface may include a curved surface, which swells toward the axis of the transition piece and toward the downstream side, in at least a portion thereof.
  • the ratio of the number of the combustors to the number of the first stage turbine vanes may be an odd number that is equal to or more than 2:3, and a ratio S/P of a dimension S in the circumferential direction from an intermediate point between the transition piece of the combustor and the transition piece of the other combustor to the upstream end of the first stage turbine vane nearest to the intermediate point in the circumferential direction, to a pitch dimension P of the plurality of first stage turbine vanes, may be equal to or less than 0.05, between 0.2 to 0.55, or between 0.7 to 1.0.
  • any first stage turbine vanes are present relatively closely in the circumferential direction, even at the respective downstream ends of the transition piece connected from the respective inner surfaces of the pair of lateral walls of the transition piece of any combustor. Therefore, pressure fluctuation on the downstream side of the respective transition pieces can be suppressed due to the presence of the first stage turbine vanes.
  • a ratio L/P of a dimension L in the axis direction from the downstream end of the transition piece to the upstream end of the first stage turbine vane, to a pitch dimension P of the plurality of first stage turbine vanes may be equal to or less than 0.2.
  • the first stage turbine vane is present relatively closely in the axis direction of the transition piece on the downstream end of the transition piece. Therefore, pressure fluctuation on the downstream side of the respective transition piece can be suppressed due to the presence of the first stage turbine vane.
  • FIG. 1 is an overall side view in which main portions of a gas turbine in an embodiment related to the invention are cut out.
  • FIG. 2 is a cross-sectional view of the circumference of a combustor of a gas turbine in one embodiment related to the invention.
  • FIG. 3 is a perspective view of a transition piece in the embodiment related to the invention.
  • FIG. 4 is a cross-sectional view of a downstream side of the transition piece in the embodiment related to the invention.
  • FIG. 5 is a graph showing the relationship between the inclination rate of an inclination surface and pressure fluctuation range in the embodiment related to the invention.
  • FIG. 6 is an explanatory view showing the positional relationship between the transition piece and first stage turbine vanes in the embodiment related to the invention.
  • FIG. 7 is an explanatory view showing pressure fluctuations on the downstream side of the transition piece in the embodiment related to the invention
  • FIG. 7( a ) shows a case where a circumferential ratio is 10%
  • FIG. 7( b ) shows a case where the circumferential ratio is 22.5%
  • FIG. 7( c ) shows a case where the circumferential ratio is 35%
  • FIG. 7( d ) shows a case where the circumferential ratio is 47.5%.
  • FIG. 8 is a graph showing the relationship between the pressure fluctuation range and the circumferential ratio on the downstream side of the transition piece of the embodiment related to the invention.
  • FIG. 9 is a cross-sectional view on the downstream side of the transition piece for showing modification examples of an inclination surface of the embodiment related to the invention
  • FIG. 9( a ) shows a first modification example of the inclination surface
  • FIG. 9( b ) shows a second modification example of the inclination surface.
  • FIG. 10 is an explanatory view showing the positional relationship between an inclination surface of a transition piece and first stage turbine vanes in a modification example of one embodiment related to the invention.
  • the gas turbine of the present embodiment includes a compressor 1 that compresses the outside air to generate compressed air, a plurality of combustors 10 configured to mix and combust the fuel from a fuel supply source with the compressed air to generate combustion gas, and a turbine 2 that is driven by the combustion gas.
  • the turbine 2 includes a casing 3 , and a turbine rotor 5 that rotates within the casing 3 .
  • the turbine rotor 5 has a rotor body 6 configured such that a plurality of rotor disks are stacked, and a plurality of turbine blades 7 that extend in a radial direction from the rotor disks for each rotor disk. That is, the turbine rotor 5 has a multi-stage turbine blade configuration.
  • a generator (not shown) that generates electricity by the rotation of the turbine rotor 5 is connected to the turbine rotor 5 .
  • a plurality of turbine vanes 4 that extend in a direction approaching the rotor body 6 from an inner peripheral surface of the casing are respectively fixed to the casing 3 on the upstream side of the turbine blade 7 of each stage.
  • the plurality of combustors 10 are fixed to the casing 3 at equal intervals from each other in the circumferential direction with a rotation axis Ar of the turbine rotor 5 as a center.
  • the combustor 10 includes a transition piece 20 through which high-temperature and high-pressure combustion gas G is flowed from a gas inlet 9 of the turbine 2 into a gas flow passage 8 of the turbine 2 , and a fuel supply device 11 that supplies fuel and compressed air Air into the transition piece 20 .
  • the turbine blades 7 and the turbine vanes 4 of the turbine 2 are arranged in the gas flow passage 8 .
  • the fuel supply device 11 includes a pilot burner 12 that supplies pilot fuel X into the transition piece 20 to form a diffusion flame in the transition piece 20 , and a plurality of main nozzles 13 that premix main fuel Y and the compressed air Air to supply premixed gas into the transition piece 20 and forms a premixed flame in the transition piece 20 .
  • the transition piece 20 has a trunk 21 that forms a tubular shape and has the combustion gas G flowing on an inner peripheral side thereof, and an outlet flange 31 that is provided at a downstream end portion of the trunk 21 and spreads in a direction away from the axis Ac of the transition piece 20 .
  • the cross-sectional shape of the trunk 21 on the downstream side has an oblong shape, and the trunk 21 has, at a downstream portion, a pair of lateral walls 22 facing each other in a circumferential direction C centered on the rotation axis Ar of the turbine rotor 5 , and a pair of lateral walls 23 facing each other in a radial direction centered on the rotation axis Ar.
  • the outlet flange 31 provided at the downstream end portion of the trunk 21 has a flange body portion 32 that spreads in the direction away from the downstream end of the trunk 21 with respect to the axis Ac of the transition piece 20 , and a facing portion 33 that extends toward the upstream side from an outer edge of the flange body portion 32 .
  • a downstream end surface of the flange body portion 32 forms a downstream end surface 20 ea of the transition piece 20 .
  • a seal member 35 that seals a space between the transition pieces of the adjacent combustors 10 is provided between this facing portion 33 and the facing portion 33 of the transition piece 20 of an adjacent combustor 10 in the circumferential direction C.
  • the portion of the trunk 21 on the downstream side that is, the lateral walls 22 and 23 of the downstream portion of the trunk 21 , and the flange body portion 32 are formed from an integrally molded article.
  • Respective inner surfaces 24 of the pair of lateral walls 22 facing each other in the circumferential direction C form inclination surfaces 25 that incline down to a downstream end 20 e of the transition piece 20 in the direction approaching the transition piece 20 of another adjacent combustor 10 that gradually draws closer as it goes to the downstream side in the direction of the axis Ac of the transition piece 20 . That is, the downstream ends of the inclination surfaces 25 are downstream ends 20 e of the transition piece 20 .
  • a Karman's vortex street may be formed on the downstream side of the downstream end surface 20 ea of the flange body portion 32 .
  • the angle formed by the downstream end surface 20 ea of the flange body portion 32 and the inner surfaces 24 of the lateral walls 22 is smaller than that in a case where the inner surfaces 24 do not form the inclination surfaces 25 .
  • formation of the Karman's vortex street on the downstream side of the downstream end surface 20 ea of the flange body portion 32 can be suppressed, and pressure fluctuation of the downstream portion of the transition piece 20 can be suppressed.
  • the preferable range Ra of the inclination rate A/B of the inclination surface 25 is 1 to 8, and the more preferable range Rb is 2 to 6.
  • this simulation was performed with the ratio of the number Nc of the combustors 10 and the number Ns of the first stage turbine vanes 4 a being 2:3 and with the inclination rate A/B of the inclination surface 25 of the transition piece 20 being 2.75.
  • this simulation was performed with the ratio L/P (hereinafter referred to as an axis-direction ratio L/P) of a dimension L in the direction of the axis Ac from the downstream end 20 e of the transition piece 20 to the upstream end 4 s of the first stage turbine vane 4 a to the pitch dimension P being 12%.
  • the pressure fluctuation range ⁇ P becomes large.
  • the pressure fluctuation range ⁇ P is smaller than that in a case where the inclination surface 25 is not formed, or the like.
  • the pressure fluctuation range ⁇ P becomes drastically small as shown in FIG. 8 if the circumferential ratio S/P reaches 20%, and an unsteady pressure fluctuation is hardly seen where the circumferential ratio S/P reaches 22.5%.
  • FIG. 7( b ) where the circumferential ratio S/P is 22.5%, an unsteady pressure fluctuation is hardly seen on the downstream side between the transition piece 20 of the specific combustor 10 a and the transition piece 20 of the other combustor 10 b that is adjacent to the specific combustor on one side in the circumferential direction C and even on the downstream side between the transition piece 20 of the specific combustor 10 a and the transition piece 20 of the other combustor 10 c that is adjacent to the specific combustor on the other side in the circumferential direction C.
  • the pressure fluctuation range ⁇ P becomes drastically small where the circumferential ratio S/P reaches 70%, and an unsteady pressure fluctuation is hardly seen where the circumferential ratio S/P becomes 72.5%. An unsteady pressure fluctuation is hardly seen thereafter until the circumferential ratio S/P reaches 100%.
  • the circumferential ratios S/P are 0 to 5%, 20 to 55%, and 70 to 100%, an unsteady pressure fluctuation is hardly seen, and the pressure fluctuation range ⁇ P of the downstream portion of the transition piece 20 becomes extremely small. That is, it can be understood that the preferable ranges Rc of the circumferential ratio S/P are 0 to 5%, 20 to 55%, and 70 to 100%.
  • the pressure fluctuation range ⁇ P becomes small as the axis-direction ratio L/P is made equal to or less than 20%.
  • the absolute value of the pressure fluctuation range ⁇ P when the pressure fluctuation range ⁇ P becomes large becomes smaller at any value of the circumferential ratio S/P or the axis-direction ratio L/P.
  • the number Ns of the first stage turbine vanes 4 a becomes larger, an unsteady pressure fluctuation is hardly seen at any value of the circumferential ratio S/P or the axis-direction ratio L/P.
  • the upstream end 4 s of the first stage turbine vane 4 a can be arranged directly below the position between the transition piece 20 of each combustor 10 and the other transition piece 20 that is adjacent to the combustor of each combustor. Therefore, an unsteady pressure fluctuation can be almost eliminated by arranging the first stage turbine vanes 4 a in this way.
  • the whole surface from the upstream end 25 s of the inclination surface to the downstream end 20 e thereof is a planar surface.
  • the inclination surface 25 does not need to be a planar surface as a whole, and may also include curved surfaces in at least portions thereof.
  • the curved surfaces are curved surfaces 26 a, 26 b, and 26 c that swell smoothly toward the axis Ac of the transition piece 20 and toward the downstream side.
  • the curved surface 26 a as shown in FIG. 9 a , in a boundary region 26 a between the inclination surface 25 and the downstream end surface 20 ea of the flange body portion 32 , that is, a downstream end of the curved surface 26 a coincides with the downstream end 20 e of the inclination surface 25 .
  • an upstream end of the curved surface 26 b coincides with the upstream end 25 s of the inclination surface 25 .
  • the whole inclination surface 25 is the curved surface 26 c.
  • the curved surfaces are adopted in at least a portion of the inclination surface 25 in this way, a place where the flow direction of the combustion gas G changes drastically is eliminated. Therefore, formation of the Karman's vortex street on the downstream side of the downstream end surface 20 ea of the transition piece 20 can be further suppressed, and the pressure fluctuation of the downstream portion of the transition piece 20 can be more effectively suppressed. For this reason, when the curved surfaces are adopted in at least a portion of the inclination surface 25 , the preferable range Ra of the inclination rate A/B of the inclination surface 25 becomes wider than the range of 1 to 8 that is described above.
  • the respective inner surfaces 24 of the pair of lateral walls 22 facing each other in the circumferential direction C form the inclination surfaces 25 .
  • the pressure fluctuation of the downstream portion of the transition piece 20 can be suppressed.
  • the side to which the downstream end 4 e of a first stage turbine vane 4 a is located with respect to the upstream end 4 s of the first stage turbine vane 4 a in the circumferential direction C is defined as a blade inclination side Ca
  • UPSTREAM END OF TURBINE VANE
  • UPSTREAM END OF INCLINATION SURFACE

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Pre-Mixing And Non-Premixing Gas Burner (AREA)
US14/343,369 2011-09-16 2012-09-12 Gas turbine Abandoned US20140216055A1 (en)

Applications Claiming Priority (3)

Application Number Priority Date Filing Date Title
JP2011203016A JP5848074B2 (ja) 2011-09-16 2011-09-16 ガスタービン、尾筒及び燃焼器
JP2011-203016 2011-09-16
PCT/JP2012/073298 WO2013039095A1 (fr) 2011-09-16 2012-09-12 Turbine à gaz

Publications (1)

Publication Number Publication Date
US20140216055A1 true US20140216055A1 (en) 2014-08-07

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US14/343,369 Abandoned US20140216055A1 (en) 2011-09-16 2012-09-12 Gas turbine

Country Status (5)

Country Link
US (1) US20140216055A1 (fr)
EP (1) EP2752622A4 (fr)
JP (1) JP5848074B2 (fr)
CN (1) CN103782103B (fr)
WO (1) WO2013039095A1 (fr)

Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20170030219A1 (en) * 2015-07-28 2017-02-02 Ansaldo Energia Switzerland AG First stage turbine vane arrangement
US20180209282A1 (en) * 2014-08-19 2018-07-26 Mitsubishi Hitachi Power Systems, Ltd. Gas turbine
US11280203B2 (en) * 2017-08-03 2022-03-22 Mitsubishi Power, Ltd. Gas turbine including first-stage stator vanes

Families Citing this family (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20130236301A1 (en) * 2012-03-09 2013-09-12 General Electric Company Apparatus And System For Directing Hot Gas
JP7348784B2 (ja) * 2019-09-13 2023-09-21 三菱重工業株式会社 出口シール、出口シールセット、及びガスタービン
CN114234233B (zh) * 2021-11-30 2023-04-07 中国航发湖南动力机械研究所 一种蒸发管及燃烧室

Citations (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2743579A (en) * 1950-11-02 1956-05-01 Gen Motors Corp Gas turbine engine with turbine nozzle cooled by combustion chamber jacket air
JPS62121835A (ja) * 1985-11-21 1987-06-03 Agency Of Ind Science & Technol 高温空冷ガスタ−ビン
US5761898A (en) * 1994-12-20 1998-06-09 General Electric Co. Transition piece external frame support
US20020141864A1 (en) * 2001-03-29 2002-10-03 Burdgick Steven Sebastian Methods and apparatus for preferential placement of turbine nozzles and shrouds based on inlet conditions
US6840048B2 (en) * 2002-09-26 2005-01-11 General Electric Company Dynamically uncoupled can combustor
US20080202124A1 (en) * 2007-02-27 2008-08-28 Siemens Power Generation, Inc. Transition support system for combustion transition ducts for turbine engines
US20100037617A1 (en) * 2008-08-12 2010-02-18 Richard Charron Transition with a linear flow path with exhaust mouths for use in a gas turbine engine
US20100313567A1 (en) * 2008-02-20 2010-12-16 Mitsubishi Heavy Industries, Ltd. Gas turbine
US9395085B2 (en) * 2009-12-07 2016-07-19 Mitsubishi Hitachi Power Systems, Ltd. Communicating structure between adjacent combustors and turbine portion and gas turbine

Family Cites Families (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3018624A (en) * 1954-03-02 1962-01-30 Bristol Siddeley Engines Ltd Flame tubes for use in combustion systems of gas turbine engines
US3609968A (en) * 1970-04-29 1971-10-05 Westinghouse Electric Corp Self-adjusting seal structure
JPS58208519A (ja) * 1982-05-31 1983-12-05 Hitachi Ltd 燃焼器トランジシヨンピ−ス取付構造
JP2001289003A (ja) * 2000-04-04 2001-10-19 Mitsubishi Heavy Ind Ltd ガスタービンの冷却構造
JP4220947B2 (ja) * 2004-08-13 2009-02-04 三菱重工業株式会社 燃焼器尾筒とタービン入口との連通構造
EP1903184B1 (fr) * 2006-09-21 2019-05-01 Siemens Energy, Inc. Sous-système de turbine à combustion avec conduit de transition tordu
US8091365B2 (en) * 2008-08-12 2012-01-10 Siemens Energy, Inc. Canted outlet for transition in a gas turbine engine
US9822649B2 (en) * 2008-11-12 2017-11-21 General Electric Company Integrated combustor and stage 1 nozzle in a gas turbine and method

Patent Citations (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2743579A (en) * 1950-11-02 1956-05-01 Gen Motors Corp Gas turbine engine with turbine nozzle cooled by combustion chamber jacket air
JPS62121835A (ja) * 1985-11-21 1987-06-03 Agency Of Ind Science & Technol 高温空冷ガスタ−ビン
US5761898A (en) * 1994-12-20 1998-06-09 General Electric Co. Transition piece external frame support
US20020141864A1 (en) * 2001-03-29 2002-10-03 Burdgick Steven Sebastian Methods and apparatus for preferential placement of turbine nozzles and shrouds based on inlet conditions
US6840048B2 (en) * 2002-09-26 2005-01-11 General Electric Company Dynamically uncoupled can combustor
US20080202124A1 (en) * 2007-02-27 2008-08-28 Siemens Power Generation, Inc. Transition support system for combustion transition ducts for turbine engines
US20100313567A1 (en) * 2008-02-20 2010-12-16 Mitsubishi Heavy Industries, Ltd. Gas turbine
US20100037617A1 (en) * 2008-08-12 2010-02-18 Richard Charron Transition with a linear flow path with exhaust mouths for use in a gas turbine engine
US9395085B2 (en) * 2009-12-07 2016-07-19 Mitsubishi Hitachi Power Systems, Ltd. Communicating structure between adjacent combustors and turbine portion and gas turbine

Cited By (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20180209282A1 (en) * 2014-08-19 2018-07-26 Mitsubishi Hitachi Power Systems, Ltd. Gas turbine
US11118465B2 (en) * 2014-08-19 2021-09-14 Mitsubishi Power, Ltd. Gas turbine combustor transition piece including inclined surface at downstream end portions for reducing pressure fluctuations
US20170030219A1 (en) * 2015-07-28 2017-02-02 Ansaldo Energia Switzerland AG First stage turbine vane arrangement
US10233777B2 (en) * 2015-07-28 2019-03-19 Ansaldo Energia Switzerland AG First stage turbine vane arrangement
US11280203B2 (en) * 2017-08-03 2022-03-22 Mitsubishi Power, Ltd. Gas turbine including first-stage stator vanes

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