JP5848074B2 - Gas turbine, tail cylinder and combustor - Google Patents

Gas turbine, tail cylinder and combustor Download PDF

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JP5848074B2
JP5848074B2 JP2011203016A JP2011203016A JP5848074B2 JP 5848074 B2 JP5848074 B2 JP 5848074B2 JP 2011203016 A JP2011203016 A JP 2011203016A JP 2011203016 A JP2011203016 A JP 2011203016A JP 5848074 B2 JP5848074 B2 JP 5848074B2
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transition piece
downstream
turbine
circumferential direction
combustor
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JP2013064535A (en
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康朗 坂元
康朗 坂元
敬介 松山
敬介 松山
敬三 塚越
敬三 塚越
由里　雅則
雅則 由里
岸田　宏明
宏明 岸田
鳥井　俊介
俊介 鳥井
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Mitsubishi Power Ltd
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Mitsubishi Hitachi Power Systems Ltd
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Priority to JP2011203016A priority Critical patent/JP5848074B2/en
Priority to CN201280043745.5A priority patent/CN103782103B/en
Priority to EP20120832651 priority patent/EP2752622A4/en
Priority to US14/343,369 priority patent/US20140216055A1/en
Priority to PCT/JP2012/073298 priority patent/WO2013039095A1/en
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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

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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)

Description

本発明は、燃料を圧縮空気に混合して燃焼させ燃焼ガスを生成する複数の燃焼器と、複数の燃焼器からの燃焼ガスにより回転するロータを有するタービンと、を備えているガスタービンに関し、特に、燃焼器の尾筒に関する。   The present invention relates to a gas turbine comprising: a plurality of combustors that mix and burn 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. In particular, it relates to a combustor tail tube.

ガスタービンは、外気を取り込んで圧縮空気を生成する圧縮機と、燃料を圧縮空気に混合して燃焼させ燃焼ガスを生成する複数の燃焼器と、複数の燃焼器からの燃焼ガスにより回転するロータを有するタービンと、を備えている。複数の燃焼器は、ロータを中心として環状に配置されている。各燃焼器は、タービンのガス入り口に燃焼ガスを送る尾筒を有している。   The gas turbine includes a compressor that takes in outside air to generate compressed air, a plurality of combustors that generate combustion gas by mixing fuel with the compressed air, and a rotor that is rotated by the combustion gas from the plurality of combustors. Having a turbine. The plurality of combustors are annularly arranged around the rotor. Each combustor has a tail tube that delivers combustion gas to the gas inlet of the turbine.

燃焼ガスは、燃焼器の尾筒から流出すると、タービンのガス入り口からタービンの燃焼ガス流路内に入り込む。この際、燃焼ガスは、尾筒から流出した直後に、その流れの中にカルマン渦列が形成され、このカルマン渦列を振動源とする非定常圧力変動が音響固有値に共振し、大きな圧力変動が生じて、運転負荷になる場合がある。   When the combustion gas flows out of the transition piece of the combustor, it enters the combustion gas flow path of the turbine from the gas inlet of the turbine. At this time, immediately after the combustion gas flows out of the transition piece, a Karman vortex street is formed in the flow, and unsteady pressure fluctuations using this Karman vortex street as a vibration source resonate with the acoustic eigenvalue, resulting in large pressure fluctuations. May occur, resulting in an operating load.

そこで、以下の特許文献１に記載の技術では、尾筒の下流端と第一段静翼の上流端との間の軸線方向における寸法や、第一段静翼の上流端とロータを中心とした周方向で隣接する尾筒間の中心との間の周方向の寸法等に関して、特定の範囲に限定することで、大きな圧力変動を抑えている。   Therefore, in the technique described in Patent Document 1 below, the dimension in the axial direction between the downstream end of the transition piece and the upstream end of the first stage stationary blade, the upstream end of the first stage stationary blade and the rotor are the center. A large pressure fluctuation is suppressed by limiting the dimension in the circumferential direction between the center portions adjacent to each other in the circumferential direction to a specific range.

特開２００９−１９７６５０号公報JP 2009-197650 A

上記特許文献１に記載の技術は、確かに、尾筒の下流側部分における大きな圧力変動を抑えることができる。しかしながら、尾筒の下流側部分における圧力変動をより抑え、よりガスタービン効率を高めることが望まれている。   The technique described in Patent Document 1 can certainly suppress a large pressure fluctuation in the downstream portion of the transition piece. However, it is desired to further suppress the pressure fluctuation in the downstream portion of the transition piece and further increase the gas turbine efficiency.

そこで、本発明は、このような要望に応えるべく、燃焼器の尾筒の下流側部分における圧力変動をより抑え、よりガスタービン効率を高めることができるガスタービン、尾筒及び燃焼器を提供することを目的とする。 Therefore, the present invention provides a gas turbine , a tail cylinder, and a combustor that can further suppress the pressure fluctuation in the downstream portion of the tail pipe of the combustor and further improve the gas turbine efficiency in order to meet such a demand. For the purpose.

上記目的を達成するための発明に係るガスタービンは、燃料を圧縮空気に混合して燃焼させ燃焼ガスを生成する複数の燃焼器と、複数の該燃焼器からの燃焼ガスにより回転するロータを有するタービンと、を備え、複数の燃焼器は、前記ロータを中心として環状に配置され、前記タービンのガス入口に燃焼ガスを送る尾筒を有するガスタービンにおいて、前記燃焼器の前記尾筒の下流部で、前記ロータの周方向で互いに対向する一対の側壁のうち、少なくとも一方の側壁の内面は、該尾筒の軸線方向の下流側に向うに連れて、次第に隣接する他の燃焼器の尾筒に近づく向きに、該尾筒の下流端に至るまで傾斜している傾斜面を成しており、前記タービンは、前記ロータを中心として環状に且つ前記ガス入口に沿って配置された複数の第一段静翼を有し、該第一段静翼は、翼弦が伸びる翼弦方向が前記周方向に対して傾斜しており、前記周方向で、前記第一段静翼の上流端に対して該第一段静翼の下流端が存在する側を翼傾斜側とした場合、前記尾筒の前記一方の側壁の傾斜面は、前記尾筒における前記周方向で互いに対向する一対の前記側壁のうちの該翼傾斜側の側壁であり、前記翼傾斜側の側壁の傾斜面の下流端は、前記第一段静翼における上流側を向く面であって該第一段静翼の上流端から下流端にわたって延びる腹側の面に対して、前記軸線方向に対向していることを特徴とする。
また、上記ガスタービンでは、前記一方の側壁のうちの下流側の端部のみに前記傾斜面を有することが好ましい。
さらに、上記ガスタービンでは、前記翼傾斜側の側壁の傾斜面によって下流側に導かれる燃焼ガスが、前記軸線方向に対向する腹側の面によってさらに下流側に導かれることが好ましい。
また、本発明に係る尾筒は、タービンのロータを中心として隣接するように環状に複数配置されて前記タービンのガス入口に燃焼ガスを送る燃焼器の尾筒であって、前記ロータの周方向で互いに対向する一対の側壁のうち、一方の側壁の内面のみが、該尾筒の軸線方向の下流側に向うに連れて、次第に隣接する他の燃焼器の尾筒に近づく向きに、該尾筒の下流端に至るまで傾斜している傾斜面を成していることを特徴とする。
さらに、前記一方の側壁は、前記周方向で互いに対向する一対の前記側壁のうち、前記ロータとともに回転する動翼の回転方向前方側の側壁であることが好ましい。
また、上記尾筒は、周方向において、前記ロータを中心として環状に且つ前記ガス入口に沿って配置されて翼弦が伸びる翼弦方向が前記周方向に対して傾斜する第一段静翼の上流端に対して該第一段静翼の下流端が存在する側を翼傾斜側とした場合、前記一方の側壁は、前記周方向で互いに対向する一対の前記側壁のうちの該翼傾斜側の側壁であることが好ましい。
さらに、本発明に係る燃焼器は、上記の尾筒を備えることを特徴とする。
そして、本発明に係るガスタービンは、上記の燃焼器と、前記第一段静翼を有する前記タービンと、を備えることを特徴とする。 A gas turbine according to an invention for achieving the above object includes a plurality of combustors that generate combustion gas by mixing fuel with compressed air and a rotor that is rotated by the combustion gas from the plurality of combustors. And a plurality of combustors arranged annularly around the rotor, and having a tail tube that sends combustion gas to a gas inlet of the turbine, a downstream portion of the tail tube of the combustor Of the pair of side walls facing each other in the circumferential direction of the rotor, the inner surface of at least one of the side walls gradually moves toward the downstream side in the axial direction of the tail cylinder, and the tail cylinder of the other combustor gradually adjoins. The turbine is inclined to reach the downstream end of the transition piece, and the turbine has a plurality of second rings arranged around the rotor and along the gas inlet. One step still In the first stage stationary blade, the chord direction in which the chord extends is inclined with respect to the circumferential direction, and the first stage stationary blade is inclined with respect to the upstream end of the first stage stationary blade in the circumferential direction. When the side where the downstream end of the first stage stationary blade exists is the blade inclined side, the inclined surface of the one side wall of the tail tube is the one of the pair of side walls facing each other in the circumferential direction of the tail tube. The blade inclined side wall, and the downstream end of the inclined surface of the blade inclined side wall is a surface facing the upstream side of the first stage stationary blade, and is the downstream end from the upstream end of the first stage stationary blade It is opposed to the abdominal surface extending in the axial direction .
Moreover, in the said gas turbine, it is preferable to have the said inclined surface only in the downstream edge part of said one side wall.
Further, in the gas turbine, it is preferable that the combustion gas guided downstream by the inclined surface of the blade inclined side wall is guided further downstream by the ventral surface facing the axial direction.
Further, the transition piece according to the present invention is a transition piece of a combustor that is arranged in a plurality of annular shapes so as to be adjacent to each other around the rotor of the turbine and sends combustion gas to the gas inlet of the turbine, and the circumferential direction of the rotor Of the pair of side walls opposed to each other, only the inner surface of one side wall gradually approaches the tail cylinder of another combustor adjacent to the tail cylinder as it goes downstream in the axial direction of the tail cylinder. An inclined surface that is inclined to reach the downstream end of the cylinder is formed.
Furthermore, it is preferable that the one side wall is a side wall on the front side in the rotational direction of the moving blade rotating together with the rotor, of the pair of side walls facing each other in the circumferential direction.
In addition, the tail tube is a first stage stationary blade that is arranged annularly with respect to the rotor in the circumferential direction and along the gas inlet so that the chord direction in which the chord extends extends with respect to the circumferential direction. When the side where the downstream end of the first stage stationary blade exists with respect to the upstream end is the blade inclined side, the one side wall is the blade inclined side of the pair of side walls facing each other in the circumferential direction. It is preferable that it is a side wall.
Furthermore, the combustor which concerns on this invention is equipped with said tail tube, It is characterized by the above-mentioned.
And the gas turbine which concerns on this invention is provided with said combustor and the said turbine which has said 1st stage stationary blade, It is characterized by the above-mentioned.

尾筒内を下流側に向って流れる燃焼ガスは、尾筒内から流出した後も尾筒の側壁内面に沿った方向に流れようとするため、尾筒の下流端面の下流側にカルマン渦列が形成されることがある。   Since the combustion gas flowing toward the downstream side in the transition piece tries to flow in the direction along the inner surface of the side wall of the transition piece after flowing out of the transition piece, the Karman vortex street is located downstream of the downstream end face of the transition piece. May be formed.

当該ガスタービンでは、尾筒の下流側の側壁内面が尾筒の下流端に至るまで傾斜面を成しているため、互いに隣接し合う燃焼器の尾筒の側壁内面に沿う流れ同士が、尾筒の下流端面の下流側に角度をもって合流するので、尾筒の下流端面の下流側にカルマン渦列が形成されるのを抑制でき、尾筒の下流側部分の圧力変動を抑えることができる。
また、尾筒における周方向で互いに対向する一対の側壁のうちの翼傾斜側の側壁のみの内面が傾斜面である場合、この傾斜面により導かれる燃焼ガスの流れの向きと、第一段静翼により導かれる燃焼ガスの流れの向きとがほぼ同じなり、尾筒から第一段静翼への燃焼ガスの流れがスムーズになる。このため、翼傾斜側の側壁のみの内面が傾斜面であっても、尾筒の下流側部分の圧力変動を効果的に抑えることができる。なお、翼弦とは、静翼の上流端と下流端とを結んだ線分である。 In the gas turbine, since the inner surface of the side wall on the downstream side of the transition piece forms an inclined surface until it reaches the downstream end of the transition piece, the flows along the inner surface of the side wall of the transition piece adjacent to each other are Since it joins at an angle to the downstream side of the downstream end face of the cylinder, it is possible to suppress the formation of a Karman vortex street on the downstream side of the downstream end face of the tail cylinder, and to suppress the pressure fluctuation in the downstream part of the tail cylinder.
Further, when the inner surface of only the side wall on the blade inclined side of the pair of side walls facing each other in the circumferential direction in the tail cylinder is an inclined surface, the flow direction of the combustion gas guided by the inclined surface and the first stage static The direction of the flow of the combustion gas guided by the blade is almost the same, and the flow of the combustion gas from the transition piece to the first stage stationary blade becomes smooth. For this reason, even if the inner surface of only the side wall on the blade inclination side is an inclined surface, the pressure fluctuation in the downstream portion of the tail tube can be effectively suppressed. The chord is a line segment connecting the upstream end and the downstream end of the stationary blade.

また、前記ガスタービンにおいて、前記尾筒における前記周方向で互いに対向する一対の前記側壁の両方の前記内面が前記傾斜面を成していることが好ましい。   In the gas turbine, it is preferable that the inner surfaces of both of the pair of side walls facing each other in the circumferential direction of the tail tube form the inclined surface.

当該ガスタービンでは、尾筒の一対の側壁の内面のそれぞれから連なる尾筒の各下流端面の下流側にカルマン渦列が形成されるのを抑制できる。   In the gas turbine, it is possible to suppress the formation of a Karman vortex street on the downstream side of each downstream end surface of the transition piece that is continuous from the inner surfaces of the pair of side walls of the transition piece.

また、前記ガスタービンにおいて、前記傾斜面の上流端から下流端までの前記周方向の寸法Ｂを基準として、該上流端から該下流端までの前記軸線方向の寸法Ａの割合Ａ／Ｂは、１以上で８以下であることが好ましい。   Further, in the gas turbine, on the basis of the circumferential dimension B from the upstream end to the downstream end of the inclined surface, the ratio A / B of the axial dimension A from the upstream end to the downstream end is: It is preferable that it is 1 or more and 8 or less.

本発明に係るガスタービンは、燃料を圧縮空気に混合して燃焼させ燃焼ガスを生成する複数の燃焼器と、複数の該燃焼器からの燃焼ガスにより回転するロータを有するタービンと、を備え、複数の燃焼器は、前記ロータを中心として環状に配置され、前記タービンのガス入口に燃焼ガスを送る尾筒を有するガスタービンにおいて、前記燃焼器の前記尾筒の下流部で、前記ロータの周方向で互いに対向する一対の側壁のうち、少なくとも一方の側壁の内面は、該尾筒の軸線方向の下流側に向うに連れて、次第に隣接する他の燃焼器の尾筒に近づく向きに、該尾筒の下流端に至るまで傾斜している傾斜面を成しており、前記傾斜面は、前記尾筒の軸線に近づき且つ下流側に向う側に滑らかに膨らんだ曲面を少なくとも一部に含む、ことを特徴とする。A gas turbine according to the present invention includes a plurality of combustors that generate combustion gas by mixing fuel with compressed air, 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 around the rotor, and have a tail tube that sends combustion gas to a gas inlet of the turbine. In the gas turbine, the periphery of the rotor is disposed downstream of the tail tube of the combustor. The inner surface of at least one of the pair of side walls facing each other in the direction is gradually closer to the tail cylinder of the other adjacent combustor as it goes downstream in the axial direction of the tail cylinder. An inclined surface that is inclined to reach the downstream end of the transition piece, and the inclined surface includes at least part of a curved surface that approaches the axis of the transition piece and smoothly swells toward the downstream side, It is characterized by .

本発明に係るガスタービンは、燃料を圧縮空気に混合して燃焼させ燃焼ガスを生成する複数の燃焼器と、複数の該燃焼器からの燃焼ガスにより回転するロータを有するタービンと、を備え、複数の燃焼器は、前記ロータを中心として環状に配置され、前記タービンのガス入口に燃焼ガスを送る尾筒を有するガスタービンにおいて、前記燃焼器の前記尾筒の下流部で、前記ロータの周方向で互いに対向する一対の側壁のうち、少なくとも一方の側壁の内面は、該尾筒の軸線方向の下流側に向うに連れて、次第に隣接する他の燃焼器の尾筒に近づく向きに、該尾筒の下流端に至るまで傾斜している傾斜面を成しており、前記燃焼器の数と前記第一段静翼の数との比が２：３以上の奇数であり、複数の前記第一段静翼のピッチ寸法Ｐを基準にして、前記燃焼器の前記尾筒と前記他の燃焼器の尾筒との中間地点から前記周方向で最も近い第一段静翼の上流端までの該周方向の寸法Ｓの割合Ｓ／Ｐは、０．０５以下、０．２から０．５５の間、及び０．７から１．０の間である、ことを特徴とする。A gas turbine according to the present invention includes a plurality of combustors that generate combustion gas by mixing fuel with compressed air, 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 around the rotor, and have a tail tube that sends combustion gas to a gas inlet of the turbine. In the gas turbine, the periphery of the rotor is disposed downstream of the tail tube of the combustor. The inner surface of at least one of the pair of side walls facing each other in the direction is gradually closer to the tail cylinder of the other adjacent combustor as it goes downstream in the axial direction of the tail cylinder. An inclined surface that is inclined to the downstream end of the transition piece, and a ratio of the number of the combustors to the number of the first stage stationary blades is an odd number of 2: 3 or more, Based on the pitch dimension P of the first stage stationary blade, The ratio S / P of the circumferential dimension S from the intermediate point between the transition piece of the combustor and the transition piece of the other combustion unit to the upstream end of the first stage stationary blade closest in the circumferential direction is 0.05 or less, between 0.2 and 0.55, and between 0.7 and 1.0.

当該ガスタービンでは、いずれの燃焼器の尾筒の一対の側壁の内面のそれぞれから連なる尾筒の各下流端に対しても、いずれかの第１段静翼が周方向において比較的近くに存在することになるため、この第一段静翼の存在により、各尾筒の下流側における圧力変動を抑えることができる。   In the gas turbine, one of the first stage stationary blades exists relatively close in the circumferential direction to each downstream end of the transition piece connected from each of the inner surfaces of the pair of side walls of the transition piece of any combustor. Therefore, the presence of this first stage stationary blade can suppress the pressure fluctuation on the downstream side of each transition piece.

また、前記ガスタービンにおいて、複数の前記第一段静翼のピッチ寸法Ｐを基準にして、前記尾筒の下流端から前記第一段静翼の上流端までの前記軸線方向の寸法Ｌの割合Ｌ／Ｐが０．２以下であることが好ましい。   Further, in the gas turbine, the ratio of the dimension L in the axial direction from the downstream end of the transition piece to the upstream end of the first stage stationary blade, based on the pitch dimension P of the plurality of first stage stationary blades L / P is preferably 0.2 or less.

当該ガスタービンでは、尾筒の下流端に対して、第一段静翼が尾筒の軸線方向において比較的に近くに存在することになるため、この第一段静翼の存在により、各尾筒の下流側における圧力変動を抑えることができる。   In the gas turbine, since the first stage stationary blades are relatively close to the downstream end of the tail cylinder in the axial direction of the tail cylinder, the presence of the first stage stationary blades causes each tail Pressure fluctuation on the downstream side of the cylinder can be suppressed.

本発明では、尾筒の下流端面の下流側にカルマン渦列が形成されるのを抑制でき、尾筒の下流側部分の圧力変動を抑えることができる。このため、本発明によれば、ガスタービン効率を高めることがでる。   In the present invention, it is possible to suppress the formation of a Karman vortex street on the downstream side of the downstream end face of the transition piece, and it is possible to suppress the pressure fluctuation in the downstream portion of the transition piece. For this reason, according to the present invention, the gas turbine efficiency can be increased.

本発明に係る一実施形態におけるガスタービンの要部を切り欠いた全体側面図である。It is the whole side view which notched the principal part of the gas turbine in one Embodiment which concerns on this invention. 本発明に係る一実施形態におけるガスタービンの燃焼器周りの断面図である。It is sectional drawing around the combustor of the gas turbine in one Embodiment which concerns on this invention. 本発明に係る一実施形態における尾筒の斜視図である。It is a perspective view of a transition piece in one embodiment concerning the present invention. 本発明に係る一実施形態における尾筒の下流側の断面図である。It is sectional drawing of the downstream of the transition piece in one Embodiment which concerns on this invention. 本発明に係る一実施形態における傾斜面の傾斜率と圧力変動幅との関係を示すグラフである。It is a graph which shows the relationship between the inclination rate of the inclined surface in one Embodiment which concerns on this invention, and a pressure fluctuation range. 本発明に係る一実施形態における尾筒と第一段静翼との位置関係を示す説明図である。It is explanatory drawing which shows the positional relationship of the transition piece and 1st stage stationary blade in one Embodiment which concerns on this invention. 本発明に係る一実施形態における尾筒の下流側の圧力変動を示す説明図で、同図（ａ）は周方向割合が１０％の場合を示し、同図（ｂ）は周方向割合が２２．５％の場合を示し、同図（ｃ）は周方向割合が３５％の場合を示し、同図（ｄ）は周方向割合が４７．５％の場合を示す。It is explanatory drawing which shows the pressure fluctuation of the downstream of the transition piece in one Embodiment which concerns on this invention, The figure (a) shows the case where the circumferential direction ratio is 10%, The figure (b) shows the circumferential direction ratio. FIG. 5C shows a case where the circumferential ratio is 35%, and FIG. 4D shows a case where the circumferential ratio is 47.5%. 本発明に係る一実施形態の尾筒の下流側の圧力変動幅と周方向割合との関係を示すグラフである。It is a graph which shows the relationship between the pressure fluctuation width of the downstream of the transition piece of one Embodiment which concerns on this invention, and the circumferential direction ratio. 本発明に係る一実施形態の傾斜面の変形例を示すための尾筒の下流側の断面図であり、同図（ａ）は傾斜面の第一変形例を示し、同図（ｂ）は傾斜面の第二変形例を示す。It is sectional drawing of the downstream of the transition piece for showing the modification of the inclined surface of one Embodiment which concerns on this invention, The figure (a) shows the 1st modification of an inclined surface, The figure (b) The 2nd modification of an inclined surface is shown. 本発明に係る一実施形態の変形例における尾筒の傾斜面と第１段静翼との位置関係を示す説明図である。It is explanatory drawing which shows the positional relationship of the inclined surface of a transition piece, and a 1st step | paragraph stationary blade in the modification of one Embodiment which concerns on this invention.

以下、本発明に係るガスタービンの一実施形態について、図面を参照して詳細に説明する。   Hereinafter, an embodiment of a gas turbine according to the present invention will be described in detail with reference to the drawings.

本実施形態のガスタービンは、図１に示すように、外気を圧縮して圧縮空気を生成する圧縮機１と、燃料供給源からの燃料を圧縮空気に混合して燃焼させ燃焼ガスを生成する複数の燃焼器１０と、燃焼ガスにより駆動するタービン２と、を備えている。   As shown in FIG. 1, the gas turbine according to the present embodiment generates a combustion gas by combusting a compressor 1 that compresses outside air to generate compressed air and fuel from a fuel supply source mixed with the compressed air. A plurality of combustors 10 and a turbine 2 driven by combustion gas are provided.

タービン２は、ケーシング３と、このケーシング３内で回転するタービンロータ５とを備えている。タービンロータ５は、複数のロータディスクが積層されて構成されているロータ本体６と、複数のロータディスク毎にそのロータディスクから放射方向に延びる複数の動翼７と、を有している。すなわち、このタービンロータ５は、多段動翼構成である。このタービンロータ５には、例えば、このタービンロータ５の回転で発電する発電機（図示されていない。）と接続されている。また、ケーシング３には、各段の動翼７の各上流側に、その内周面からロータ本体６に近づく方向に延びる複数の静翼４が固定されている。   The turbine 2 includes a casing 3 and a turbine rotor 5 that rotates in the casing 3. The turbine rotor 5 includes a rotor main body 6 configured by laminating a plurality of rotor disks, and a plurality of moving blades 7 extending in a radial direction from the rotor disks for each of the plurality of rotor disks. That is, the turbine rotor 5 has a multistage rotor blade configuration. The turbine rotor 5 is connected to, for example, a generator (not shown) that generates electricity by the rotation of the turbine rotor 5. A plurality of stationary blades 4 extending from the inner peripheral surface of the casing 3 in the direction approaching the rotor body 6 are fixed to the upstream side of the moving blades 7 of each stage.

複数の燃焼器１０は、タービンロータ５の回転軸線Ａｒを中心として、周方向に互いに等間隔でケーシング３に固定されている。   The plurality of combustors 10 are fixed to the casing 3 at equal intervals in the circumferential direction around the rotation axis Ar of the turbine rotor 5.

燃焼器１０は、図２に示すように、高温・高圧の燃焼ガスＧをタービン２のガス入口９からタービン２のガス流路８内に送る尾筒２０と、この尾筒２０内に燃料及び圧縮空気Ａｉｒを供給する燃料供給器１１と、を備えている。タービン２の動翼７及び静翼４は、このガス流路８中に配置されている。燃料供給器１１は、パイロット燃料Ｘを尾筒２０内に供給して、この尾筒２０内に拡散火炎を形成するパイロットバーナ１２と、メイン燃料Ｙ及び圧縮空気Ａｉｒを予混合して、予混合気体として尾筒２０内に供給し、この尾筒２０内に予混合火炎を形成する複数のメインノズル１３と、を備えている。   As shown in FIG. 2, the combustor 10 includes a tail cylinder 20 that sends high-temperature and high-pressure combustion gas G from a gas inlet 9 of the turbine 2 into the gas flow path 8 of the turbine 2, and fuel and fuel in the tail cylinder 20. And a fuel supplier 11 for supplying the compressed air Air. The moving blade 7 and the stationary blade 4 of the turbine 2 are arranged in the gas flow path 8. The fuel supplier 11 supplies pilot fuel X into the tail cylinder 20, premixes the pilot burner 12 that forms a diffusion flame in the tail cylinder 20, the main fuel Y, and the compressed air Air, and performs premixing. A plurality of main nozzles 13 that supply gas as a gas into the transition piece 20 and form a premixed flame in the transition piece 20 are provided.

尾筒２０は、図２及び図３に示すように、筒状を成し、内周側に燃焼ガスＧが流れる胴体２１と、胴体２１の下流端部に設けられ、尾筒２０の軸線Ａｃから遠ざかる向きに広がる出口フランジ３１とを有している。   As shown in FIGS. 2 and 3, the transition piece 20 has a cylindrical shape, and is provided at a body 21 through which combustion gas G flows on the inner peripheral side, and at a downstream end portion of the body 21. And an outlet flange 31 that spreads away from the head.

胴体２１の下流側の断面形状は長方形状を成し、この胴体２１は、その下流側に、タービンロータ５の回転軸線Ａｒを中心とした周方向Ｃで互いに対向する一対の側壁２２と、この回転軸線Ａｒを中心とした放射方向で互いに対向する一対の側壁２３と、を有している。   The cross-sectional shape of the downstream side of the body 21 is rectangular, and the body 21 has a pair of side walls 22 facing each other in the circumferential direction C around the rotation axis Ar of the turbine rotor 5, and the downstream side. A pair of side walls 23 facing each other in the radial direction about the rotation axis Ar.

図４に示すように、胴体２１の下流端部に設けられている出口フランジ３１は、胴体２１の下流端から尾筒２０の軸線Ａｃに対して遠ざかる向きに広がるフランジ本体部３２と、このフランジ本体部３２の外縁から上流側に向って延びる対向部３３と、を有している。このフランジ本体部３２の下流端面は、尾筒２０の下流端面２０ｅａを成している。また、この対向部３３と、周方向Ｃで隣接する燃焼器１０の尾筒２０の対向部３３との間には、隣接する燃焼器１０の尾筒相互間をシールするシール部材３５が設けられている。なお、本実施形態では、胴体２１の下流側の部分、つまり胴体２１の下流側の側壁２２，２３とフランジ本体部３２とは、一体成形品で形成されている。   As shown in FIG. 4, the outlet flange 31 provided at the downstream end of the body 21 includes a flange main body 32 that extends from the downstream end of the body 21 in a direction away from the axis Ac of the tail cylinder 20, and this flange. And an opposing portion 33 extending from the outer edge of the main body portion 32 toward the upstream side. The downstream end surface of the flange main body portion 32 forms the downstream end surface 20ea of the transition piece 20. Further, a seal member 35 is provided between the facing portion 33 and the facing portion 33 of the transition piece 20 of the combustor 10 adjacent in the circumferential direction C to seal between the transition pieces of the adjacent combustors 10. ing. In the present embodiment, the downstream portion of the body 21, that is, the side walls 22 and 23 on the downstream side of the body 21, and the flange main body 32 are formed as an integrally molded product.

周方向Ｃで互いに対向する一対の側壁２２のそれぞれの内面２４は、尾筒２０の軸線Ａｃ方向の下流側に向うに連れて、次第に隣接する他の燃焼器１０の尾筒２０に近づく向きに、尾筒２０の下流端２０ｅに至るまで傾斜している傾斜面２５を成している。すなわち、傾斜面２５の下流端は尾筒２０の下流端２０ｅである。   The inner surfaces 24 of the pair of side walls 22 facing each other in the circumferential direction C gradually approach the tail cylinder 20 of another combustor 10 adjacent to the downstream side in the axis Ac direction of the tail cylinder 20. An inclined surface 25 that is inclined to reach the downstream end 20e of the transition piece 20 is formed. That is, the downstream end of the inclined surface 25 is the downstream end 20 e of the tail cylinder 20.

尾筒２０内を下流側に向って流れる燃焼ガスＧは、尾筒２０内から流出した後も側壁２２の内面２４に沿った方向に流れようとするため、フランジ本体部３２の下流端面２０ｅａの下流側にカルマン渦列が形成されることがある。   The combustion gas G flowing toward the downstream side in the transition piece 20 tends to flow in the direction along the inner surface 24 of the side wall 22 even after flowing out of the transition piece 20, so that the downstream end face 20 ea of the flange main body portion 32 A Karman vortex street may be formed downstream.

本実施形態では、尾筒２０の下流側の側壁２２の内面２４が傾斜面２５を成しているため、側壁２２の内面２４に対してフランジ本体部３２の下流端面２０ｅａの成す角度が小さくなる。よって、本実施形態では、フランジ本体部３２の下流端面２０ｅａの下流側にカルマン渦列が形成されるのを抑制でき、尾筒２０の下流側部分の圧力変動を抑えることができる。   In the present embodiment, since the inner surface 24 of the side wall 22 on the downstream side of the tail cylinder 20 forms an inclined surface 25, the angle formed by the downstream end surface 20 ea of the flange main body portion 32 with respect to the inner surface 24 of the side wall 22 becomes small. . Therefore, in the present embodiment, Karman vortex streets can be prevented from being formed on the downstream side of the downstream end face 20ea of the flange main body 32, and pressure fluctuations in the downstream portion of the tail cylinder 20 can be suppressed.

ここで、この傾斜面２５の傾斜率を変えた際の尾筒２０の下流側部分の圧力変動幅についてシミュレートしたので、このシミュレート結果について説明する。なお、このシミュレートでは、図４に示すように、傾斜面２５の上流端２５ｓから下流端２０ｅ（＝尾筒２０の下流端）までの周方向Ｃの寸法Ｂを基準として、上流端２５ｓから下流端２０ｅまでの軸線Ａｃ方向の寸法Ａの割合Ａ／Ｂを、傾斜面２５の傾斜率と定義している。   Here, since the pressure fluctuation width of the downstream side portion of the transition piece 20 when the inclination rate of the inclined surface 25 is changed is simulated, the simulation result will be described. In this simulation, as shown in FIG. 4, from the upstream end 25 s on the basis of the dimension B in the circumferential direction C from the upstream end 25 s of the inclined surface 25 to the downstream end 20 e (= the downstream end of the transition piece 20). The ratio A / B of the dimension A in the axis Ac direction to the downstream end 20 e is defined as the inclination rate of the inclined surface 25.

このシミュレートの結果、図５に示すように、傾斜率Ａ／Ｂが１以上で８以下であるとき、燃焼器１０の下流側部分の圧力変動幅ΔＰが小さくなることが分かった。これは、傾斜率Ａ／Ｂが１未満のとき、及び８より大きいときには、傾斜が急すぎる又は緩すぎるため、傾斜面２５としての効果を十分に得ることができないためである。さらに、傾斜面２５の傾斜率Ａ／Ｂが２以上で６以下であると、圧力変動幅ΔＰが極めて小さくなることも分かった。なお、尾筒２０内を流れる燃焼ガスＧの流速が高くなるに連れて、圧力変動幅ΔＰも大きくなるが、傾斜面２５の傾斜率Ａ/Ｂと圧力変動幅ΔＰとの関係は、尾筒２０内を流れる燃焼ガスＧの流速が変化しても基本的に同じである。   As a result of this simulation, as shown in FIG. 5, it was found that when the gradient A / B is 1 or more and 8 or less, the pressure fluctuation width ΔP in the downstream portion of the combustor 10 becomes small. This is because when the slope ratio A / B is less than 1 and greater than 8, the slope is too steep or too slow, so that the effect as the slope 25 cannot be sufficiently obtained. Furthermore, it was also found that the pressure fluctuation width ΔP becomes extremely small when the inclination ratio A / B of the inclined surface 25 is 2 or more and 6 or less. Note that, as the flow velocity of the combustion gas G flowing in the transition piece 20 increases, the pressure fluctuation range ΔP also increases, but the relationship between the inclination rate A / B of the inclined surface 25 and the pressure fluctuation range ΔP is as follows. This is basically the same even if the flow velocity of the combustion gas G flowing through 20 changes.

このため、傾斜面２５の傾斜率Ａ/Ｂの好ましい範囲Ｒａは、１以上で８以下であり、より好ましい範囲Ｒｂは、２以上で６以下であると言える。   For this reason, it can be said that the preferable range Ra of the inclination rate A / B of the inclined surface 25 is 1 or more and 8 or less, and the more preferable range Rb is 2 or more and 6 or less.

さらに、ここでは、尾筒２０と第一段静翼４ａとの相対位置を変えた際の燃焼器１０の下流側部分での圧力変動幅についてもシミュレートしたので、このシミュレート結果についても説明する。具体的に、図６に示すように、複数の第一段静翼４ａの周方向Ｃのピッチ寸法Ｐを基準にして、特定の燃焼器１０ａの尾筒２０とこれに周方向Ｃの一方の側に隣接する他の燃焼器１０ｂの尾筒２０との中間地点Ｍから周方向Ｃの一方の側で最も近い第一段静翼４ａの上流端４ｓまでの周方向Ｃの寸法Ｓの割合Ｓ／Ｐ（以下、周方向割合Ｓ／Ｐとする）と圧力変動幅ΔＰとの関係についてシミュレートした。   Furthermore, since the pressure fluctuation width in the downstream portion of the combustor 10 when the relative position between the transition piece 20 and the first stage stationary blade 4a is changed is simulated here, the simulation result is also described. To do. Specifically, as shown in FIG. 6, with reference to the pitch dimension P in the circumferential direction C of the plurality of first stage stationary blades 4a, the transition piece 20 of the specific combustor 10a and one of the circumferential direction C The ratio S of the dimension S in the circumferential direction C from the intermediate point M with the transition piece 20 of the other combustor 10b adjacent to the side to the upstream end 4s of the first stage stationary blade 4a closest to one side in the circumferential direction C The relationship between / P (hereinafter referred to as circumferential ratio S / P) and pressure fluctuation range ΔP was simulated.

なお、このシミュレートは、燃焼器１０の数Ｎｃと第一段静翼４ａの数Ｎｓとの比が２：３で、尾筒２０の傾斜面２５の傾斜率Ａ／Ｂが２．７５で行った。さらに、このシミュレートは、ピッチ寸法Ｐを基準にして、尾筒２０の下流端２０ｅから第一段静翼４ａの上流端４ｓまでの軸線Ａｃ方向の寸法Ｌの割合Ｌ／Ｐ（以下、軸線方向割合Ｌ／Ｐとする）を１２％で行った。   In this simulation, the ratio between the number Nc of the combustors 10 and the number Ns of the first stage stationary blades 4a is 2: 3, and the inclination rate A / B of the inclined surface 25 of the tail cylinder 20 is 2.75. went. Further, this simulation is based on the pitch dimension P, and the ratio L / P of the dimension L in the axis Ac direction from the downstream end 20e of the transition piece 20 to the upstream end 4s of the first stage stationary blade 4a (hereinafter referred to as the axis line). The direction ratio L / P) was 12%.

図８に示すように、軸線方向割合Ｌ／Ｐが１２％の場合、周方向割合Ｓ／Ｐが０％〜５％では非定常圧力変動はほとんどなかった。しかしながら、周方向割合Ｓ／Ｐが５％を越えると、大きな圧力変動幅ΔＰが見られ始め、周方向割合Ｓ／Ｐが１０％になると、圧力変動幅ΔＰがより大きくなった。   As shown in FIG. 8, when the axial ratio L / P was 12%, there was almost no unsteady pressure fluctuation when the circumferential ratio S / P was 0% to 5%. However, when the circumferential direction ratio S / P exceeded 5%, a large pressure fluctuation range ΔP began to be seen, and when the circumferential direction ratio S / P reached 10%, the pressure fluctuation range ΔP became larger.

図７（ａ）に示すように、周方向割合Ｓ／Ｐが１０％の場合、特定の燃焼器１０ａの尾筒２０とこれに周方向Ｃの一方の側に隣接する他の燃焼器１０ｂの尾筒２０との間の下流側では、非定常圧力変動がみられないものの、特定の燃焼器１０ａの尾筒２０とこれに周方向Ｃの他方の側に隣接する他の燃焼器１０ｃの尾筒２０との間の下流側では、カルマン渦列ｖが発生し、比較的大きな非定常圧力変動が生じた。これは、特定の燃焼器１０ａの尾筒２０とこれに周方向Ｃの他方の側に隣接する他の燃焼器１０ｃの尾筒２０との間から、周方向Ｃで最も近い第一段静翼４ａの上流端４ｓまでの周方向Ｃの寸法が大きくなっているためである、と考えられる。なお、図８中の細線は、等静圧線を示している。   As shown in FIG. 7 (a), when the circumferential ratio S / P is 10%, the transition piece 20 of a specific combustor 10a and the other combustor 10b adjacent to one side in the circumferential direction C of the specific combustor 10a. Although no unsteady pressure fluctuation is observed on the downstream side with respect to the transition piece 20, the tail section 20 of the specific combustor 10 a and the tails of the other combustors 10 c adjacent to the other side in the circumferential direction C. A Karman vortex street v was generated on the downstream side between the cylinder 20 and a relatively large unsteady pressure fluctuation occurred. This is because the first stage stationary blade closest in the circumferential direction C from between the transition piece 20 of the specific combustor 10a and the transition piece 20 of the other combustor 10c adjacent to the other side in the circumferential direction C. This is probably because the dimension in the circumferential direction C up to the upstream end 4s of 4a is increased. In addition, the thin line in FIG. 8 has shown the isostatic pressure line.

よって、周方向割合Ｓ／Ｐが１０％になると、圧力変動幅ΔＰが大きくなったと考えられる。但し、このシミュレートでは、傾斜面２５の傾斜率Ａ/Ｂは、より好ましい範囲Ｒｂ（２以上で６以下）内である２．７５であるから、傾斜面２５が形成されていない場合等よりも、圧力変動幅ΔＰより小さい。   Therefore, it is considered that when the circumferential ratio S / P is 10%, the pressure fluctuation range ΔP is increased. However, in this simulation, the inclination rate A / B of the inclined surface 25 is 2.75 which is in a more preferable range Rb (2 or more and 6 or less), and therefore, when the inclined surface 25 is not formed. Is smaller than the pressure fluctuation range ΔP.

圧力変動幅ΔＰは、周方向割合Ｓ／Ｐが２０％になると、図８に示すように、急激に小さくなり、周方向割合Ｓ／Ｐが２２.５％になると、非定常圧力変動はほとんど見られなくなった。図７（ｂ）に示すように、周方向割合Ｓ／Ｐが２２.５％の場合、特定の燃焼器１０ａの尾筒２０とこれに周方向Ｃの一方の側に隣接する他の燃焼器１０ｂの尾筒２０との間の下流側、及び特定の燃焼器１０ａの尾筒２０とこれに周方向Ｃの他方の側に隣接する他の燃焼器１０ｃの尾筒２０との間の下流側でも、非定常圧力変動はほとんど見られなくなった。   As shown in FIG. 8, the pressure fluctuation width ΔP decreases rapidly as the circumferential ratio S / P becomes 20%, and when the circumferential ratio S / P becomes 22.5%, the unsteady pressure fluctuation is almost not. I can no longer see it. As shown in FIG. 7B, when the circumferential ratio S / P is 22.5%, the transition piece 20 of the specific combustor 10a and another combustor adjacent to this on one side in the circumferential direction C. The downstream side between the transition piece 20 of the 10b and the downstream side between the transition piece 20 of the specific combustor 10a and the transition piece 20 of the other combustor 10c adjacent to the other side in the circumferential direction C. However, almost no steady pressure fluctuation was observed.

非定常圧力変動は、図７（ｃ）及び（ｄ）に示すように、周方向割合Ｓ／Ｐが３５％、４７．５％でも、ほとんど見られなかった。しかし、図８に示すように、周方向割合Ｓ／Ｐが５５％を越えると、再び、大きな圧力変動幅ΔＰが見られ始め、周方向割合Ｓ／Ｐが６０％になると、圧力変動幅ΔＰがより大きくなった。これは、特定の燃焼器１０ａの尾筒２０とこれに周方向Ｃの一方の側に隣接する他の燃焼器１０ｂの尾筒２０との間から、周方向Ｃで最も近い第一段静翼４ａの上流端４ｓまでの周方向Ｃの寸法が大きくなっているためである、と考えられる。すなわち、周方向割合Ｓ／Ｐが６０％のときには、周方向割合Ｓ／Ｐが１０％のときと、基本的に同様の理由で同様の現象が生じている、と考えられる。   As shown in FIGS. 7C and 7D, unsteady pressure fluctuation was hardly observed even when the circumferential ratio S / P was 35% or 47.5%. However, as shown in FIG. 8, when the circumferential ratio S / P exceeds 55%, a large pressure fluctuation range ΔP starts again, and when the circumferential ratio S / P reaches 60%, the pressure fluctuation range ΔP. Became larger. This is because the first stage stationary blade closest in the circumferential direction C from between the transition piece 20 of the specific combustor 10a and the transition piece 20 of the other combustor 10b adjacent to the one side in the circumferential direction C. This is probably because the dimension in the circumferential direction C up to the upstream end 4s of 4a is increased. That is, when the circumferential ratio S / P is 60%, it is considered that the same phenomenon occurs for basically the same reason as when the circumferential ratio S / P is 10%.

この圧力変動幅ΔＰは、周方向割合Ｓ／Ｐが７０％になると、急激に小さくなり、周方向割合Ｓ／Ｐが７２．５％になると、非定常圧力変動がほとんど見らなくなった。以降、周方向割合Ｓ／Ｐが１００％になるまで、非定常圧力変動はほとんど見なれない。   This pressure fluctuation range ΔP decreased rapidly when the circumferential ratio S / P was 70%, and almost no unsteady pressure fluctuation was observed when the circumferential ratio S / P was 72.5%. Thereafter, almost no unsteady pressure fluctuation is observed until the circumferential ratio S / P reaches 100%.

以上のように、周方向割合Ｓ／Ｐが０％〜５％、２０％〜５５％、７０％〜１００％では、非定常圧力変動がほとんどみられず、尾筒２０の下流側部分の圧力変動幅ΔＰが非常に小さくなることが分かった。すなわち、周方向割合Ｓ／Ｐで好ましい範囲Ｒｃは、０％〜５％、２０％〜５５％、７０％〜１００％であることが分かった。   As described above, when the circumferential ratio S / P is 0% to 5%, 20% to 55%, and 70% to 100%, there is almost no unsteady pressure fluctuation, and the pressure in the downstream portion of the tail tube 20 It was found that the fluctuation range ΔP is very small. That is, it was found that preferable ranges Rc in the circumferential direction ratio S / P are 0% to 5%, 20% to 55%, and 70% to 100%.

さらに、ここで、軸線方向割合Ｌ／Ｐを変えて、以上と同様に、周方向割合Ｓ／Ｐと圧力変動幅ΔＰとの関係についてシミュレートした。   Furthermore, here, the axial direction ratio L / P was changed, and the relationship between the circumferential direction ratio S / P and the pressure fluctuation range ΔP was simulated in the same manner as described above.

図８に示すように、軸線方向割合Ｌ／Ｐが１８％のとき、２０％のとき、２７％のときでも、周方向割合Ｓ／Ｐの変化に対する圧力変動幅ΔＰの変化は、前述の軸線方向割合Ｌ／Ｐが１２％のときと同様の傾向が見られた。すなわち、前述の軸線方向割合Ｌ／Ｐが１２％のときと同様、周方向割合Ｓ／Ｐが０％〜５％、２０％〜５５％、７０％〜１００％の場合、周方向割合Ｓ／Ｐが５％〜２０％、５５％〜７０％の場合と比べて、燃焼器１０の下流側部分の圧力変動幅ΔＰが相対的に小さくなることが分かった。   As shown in FIG. 8, even when the axial direction ratio L / P is 18%, 20%, or 27%, the change in the pressure fluctuation width ΔP with respect to the change in the circumferential direction ratio S / P The same tendency as when the direction ratio L / P was 12% was observed. That is, when the axial direction ratio L / P is 12%, the circumferential direction ratio S / P is 0% to 5%, 20% to 55%, and 70% to 100%. It was found that the pressure fluctuation width ΔP in the downstream portion of the combustor 10 is relatively small as compared with the cases where P is 5% to 20% and 55% to 70%.

しかしながら、軸線方向割合Ｌ／Ｐが１８％及び２０％のとき、周方向割合Ｓ／Ｐが０％〜５％、２０％〜５５％、７０％〜１００％の場合、前述の軸線方向割合Ｌ／Ｐが１２％のときと同様、非定常圧力変動がほとんど見られず、圧力変動幅ΔＰの絶対値が小さいものの、軸線方向割合Ｌ／Ｐが２７％のときでは、周方向割合Ｓ／Ｐが０％〜５％、２０％〜５５％、７０％〜１００％の場合でも、圧力変動幅ΔＰの絶対値が大きくなることが分かった。   However, when the axial direction ratio L / P is 18% and 20%, when the circumferential direction ratio S / P is 0% to 5%, 20% to 55%, 70% to 100%, the axial direction ratio L described above. As in the case where / P is 12%, there is almost no unsteady pressure fluctuation and the absolute value of the pressure fluctuation width ΔP is small, but when the axial direction ratio L / P is 27%, the circumferential ratio S / P It was found that the absolute value of the pressure fluctuation range ΔP is large even when 0 is 5%, 20% to 55%, and 70% to 100%.

すなわち、軸線方向割合Ｌ／Ｐは２０％以下にすることで、圧力変動幅ΔＰが小さくなることが分かった。   That is, it was found that the pressure fluctuation range ΔP is reduced by setting the axial direction ratio L / P to 20% or less.

また、以上のシミュレーション結果は、燃焼器１０の数Ｎｃと第一段静翼４ａの数Ｎｓとの比が２：３の場合のものであるが、Ｎｃ：Ｎｓ＝２：５、Ｎｃ：Ｎｓ＝２：７、Ｎｃ：Ｎｓ＝２：９以上の場合でも、尾筒２０と第一段静翼４ａとの相対位置と尾筒２０の下流側部分での圧力変動幅ΔＰとの関係は、以上と同様の結果が得られると考えられる。但し、燃焼器１０の数Ｎｃに対して第一段静翼４ａの数Ｎｓが多くなるほど、周方向割合Ｓ／Ｐや軸線方向割合Ｌ／Ｐがいずれの値でも、圧力変動幅ΔＰが大きくなる際の圧力変動幅ΔＰの絶対値は小さくなり、第一段静翼４ａの数Ｎｓがより多くなれば、周方向割合Ｓ／Ｐや軸線方向割合Ｌ／Ｐがいずれの値でも、非定常圧力変動がほとんど見られなくなると考えられる。   The above simulation results are obtained when the ratio of the number Nc of the combustors 10 and the number Ns of the first stage stationary blades 4a is 2: 3. Nc: Ns = 2: 5, Nc: Ns Even when = 2: 7, Nc: Ns = 2: 9 or more, the relationship between the relative position between the transition piece 20 and the first stage stationary blade 4a and the pressure fluctuation width ΔP in the downstream portion of the transition piece 20 is It is considered that the same result as above can be obtained. However, as the number Ns of the first stage stationary blades 4a increases with respect to the number Nc of the combustors 10, the pressure fluctuation width ΔP increases regardless of the circumferential ratio S / P and the axial ratio L / P. When the absolute value of the pressure fluctuation width ΔP at the time becomes smaller and the number Ns of the first stage stationary blades 4a becomes larger, the unsteady pressure can be obtained regardless of the circumferential direction ratio S / P and the axial direction ratio L / P. It is thought that almost no change will be seen.

なお、燃焼器１０の数Ｎｃと第一段静翼４ａの数Ｎｓとの比が１：自然数のときには、各燃焼器１０の尾筒２０と、それに隣接する他の尾筒２０との間の直下に第一段静翼４ａの上流端４ｓを配置することができるため、このように第一段静翼４ａを配置することで、非定常圧力変動をほとんどなくすことができる。   When the ratio between the number Nc of the combustors 10 and the number Ns of the first stage stationary blades 4a is 1: a natural number, the ratio between the transition piece 20 of each combustor 10 and the other transition piece 20 adjacent to it. Since the upstream end 4s of the first stage stationary blade 4a can be disposed immediately below, the unsteady pressure fluctuation can be almost eliminated by arranging the first stage stationary blade 4a in this way.

次に、尾筒２０の傾斜面２５の各種変形例について説明する。   Next, various modifications of the inclined surface 25 of the transition piece 20 will be described.

以上の実施形態の傾斜面２５は、その上流端２５ｓから下流端２０ｅまでの全体が平面であるが、この傾斜面２５は、全体が平面である必要はなく、少なくとも一部に曲面を含んでもよい。   The inclined surface 25 of the above embodiment is entirely flat from the upstream end 25s to the downstream end 20e. However, the inclined surface 25 does not need to be entirely flat, and may include a curved surface at least partially. Good.

この曲面は、具体的に、図９に示すように、尾筒２０の軸線Ａｃに近づき且つ下流側に向う側に滑らかに膨らんだ曲面２６ａ，２６ｂ，２６ｃである。例えば、曲面２６ａは、同図（ａ）に示すように、傾斜面２５とフランジ本体部３２の下流端面２０ｅａとの境界領域に、つまり、この曲面２６ａの下流端が傾斜面２５の下流端２０ｅと一致する。また、曲面２６ｂは、この曲面２６ｂの上流端が傾斜面２５の上流端２５ｓと一致する。また、同図（ｂ）に示すように、傾斜面２５の全体が曲面２６ｃである。   Specifically, as shown in FIG. 9, the curved surfaces are curved surfaces 26 a, 26 b, and 26 c that swell smoothly toward the downstream side toward the axis Ac of the tail cylinder 20. For example, the curved surface 26a is formed in the boundary region between the inclined surface 25 and the downstream end surface 20ea of the flange main body 32, that is, the downstream end of the curved surface 26a is the downstream end 20e of the inclined surface 25, as shown in FIG. Matches. Further, the curved surface 26 b has the upstream end of the curved surface 26 b coincident with the upstream end 25 s of the inclined surface 25. Further, as shown in FIG. 5B, the entire inclined surface 25 is a curved surface 26c.

このように、傾斜面２５の少なくとも一部に曲面を採用すると、燃焼ガスＧの流れの向きが急激に変化する箇所がなくなるため、尾筒２０の下流端面２０ｅａの下流側にカルマン渦列が形成されるのをより抑制でき、尾筒２０の下流側部分の圧力変動をより効果的に抑えることができる。このため、傾斜面２５の少なくとも一部に曲面を採用した場合、傾斜面２５の傾斜率Ａ/Ｂの好ましい範囲Ｒａは、以上で説明した１以上で８以下よりも広がる。   As described above, when a curved surface is employed as at least a part of the inclined surface 25, there is no place where the flow direction of the combustion gas G changes rapidly, and thus a Karman vortex street is formed on the downstream side of the downstream end surface 20ea of the tail cylinder 20. This can be suppressed more effectively, and the pressure fluctuation in the downstream portion of the transition piece 20 can be more effectively suppressed. For this reason, when a curved surface is adopted as at least a part of the inclined surface 25, the preferable range Ra of the inclination ratio A / B of the inclined surface 25 is 1 or more and 8 or less as described above.

また、以上の実施形態では、周方向Ｃで互いに対向する一対の側壁２２のそれぞれの内面２４が傾斜面２５を成しているが、いずれか一方の内面２４のみが傾斜面２５を成しても、尾筒２０の下流側部分の圧力変動を抑えることができる。この場合、図１０に示すように、周方向Ｃにおいて、第一段静翼４ａの上流端４ｓに対して第一段静翼４ａの下流端４ｅが存在する側を翼傾斜側Ｃａとした場合、周方向Ｃで互いに対向する一対の側壁２２（図１０の場合では２２ａ及び２２ｂ）のうちの翼傾斜側Ｃａの側壁２２（図１０の場合では２２ｂ）の内面２４が傾斜面２５を成していることが好ましい。これは、傾斜面２５により導かれる燃焼ガスＧの流れの向きと、第一段静翼４ａの上流端４ｓと下流端４ｅとの繋ぐ線分である翼弦の方向、つまり第一段静翼４ａにより導かれる燃焼ガスＧの流れの向きとがほぼ同じなり、尾筒２０から第一段静翼４ａへの燃焼ガスＧの流れがスムーズになり、尾筒２０の下流側部分の圧力変動を効果的に抑えることができるからである。   In the above embodiment, the inner surfaces 24 of the pair of side walls 22 facing each other in the circumferential direction C form the inclined surface 25, but only one of the inner surfaces 24 forms the inclined surface 25. In addition, the pressure fluctuation in the downstream portion of the transition piece 20 can be suppressed. In this case, as shown in FIG. 10, in the circumferential direction C, the side where the downstream end 4e of the first stage stationary blade 4a exists with respect to the upstream end 4s of the first stage stationary blade 4a is the blade inclined side Ca. Of the pair of side walls 22 (22a and 22b in the case of FIG. 10) facing each other in the circumferential direction C, the inner surface 24 of the side wall 22 (22b in the case of FIG. 10) on the blade inclination side Ca forms an inclined surface 25. It is preferable. This is because the direction of the flow of the combustion gas G guided by the inclined surface 25 and the chord direction which is a line segment connecting the upstream end 4s and the downstream end 4e of the first stage stationary blade 4a, that is, the first stage stationary blade. The flow direction of the combustion gas G guided by 4a is almost the same, the flow of the combustion gas G from the tail cylinder 20 to the first stage stationary blade 4a becomes smooth, and the pressure fluctuation in the downstream portion of the tail cylinder 20 is reduced. This is because it can be effectively suppressed.

なお、以上のように、周方向Ｃで互いに対向する一対の側壁２２のうち、いずれか一方の側壁２２（図１０の場合では２２ａ及び２２ｂ）の内面２４のみが傾斜面２５を成していても、尾筒２０と第一段静翼４ａとの相対位置と尾筒２０の下流側部分での圧力変動幅ΔＰとの関係は、以上のシミュレート結果と基本的に同様である。   As described above, of the pair of side walls 22 facing each other in the circumferential direction C, only the inner surface 24 of one of the side walls 22 (22a and 22b in the case of FIG. 10) forms the inclined surface 25. However, the relationship between the relative position of the transition piece 20 and the first stage stationary blade 4a and the pressure fluctuation width ΔP in the downstream portion of the transition piece 20 is basically the same as the above simulation result.

１：圧縮機、２：タービン、３：ケーシング、４：静翼、４ａ：第１段静翼、４ｓ：（静翼の）上流端、４ｅ：（静翼の）下流端、５：タービンロータ、６：ロータ本体、７：動翼、８：ガス流路、９：ガス入口、１０：燃焼器、２０：尾筒、２０ｅ：（尾筒のｏｒ傾斜面の）下流端、２０ｅａ：（尾筒のｏｒフランジ本体部の）下流端面、２１：胴体、２２，２２ａ，２２ｂ，２３：側壁、２４：内面、２５：傾斜面、２５ｓ：（傾斜面の）上流端、２６ａ，２６ｂ，２６ｃ：曲面   1: compressor, 2: turbine, 3: casing, 4: stationary blade, 4a: first stage stationary blade, 4s: upstream end (of stationary blade), 4e: downstream end (of stationary blade), 5: turbine rotor, 6 : Rotor body, 7: moving blade, 8: gas flow path, 9: gas inlet, 10: combustor, 20: tail tube, 20e: downstream end (of or inclined surface of tail tube), 20ea: (of tail tube) or downstream end surface (of flange main body), 21: body, 22, 22a, 22b, 23: side wall, 24: inner surface, 25: inclined surface, 25s: upstream end (of inclined surface), 26a, 26b, 26c: curved surface

Claims (14)

燃料を圧縮空気に混合して燃焼させ燃焼ガスを生成する複数の燃焼器と、複数の該燃焼器からの燃焼ガスにより回転するロータを有するタービンと、を備え、
複数の燃焼器は、前記ロータを中心として環状に配置され、前記タービンのガス入口に燃焼ガスを送る尾筒を有するガスタービンにおいて、
前記燃焼器の前記尾筒の下流部で、前記ロータの周方向で互いに対向する一対の側壁のうち、少なくとも一方の側壁の内面は、該尾筒の軸線方向の下流側に向うに連れて、次第に隣接する他の燃焼器の尾筒に近づく向きに、該尾筒の下流端に至るまで傾斜している傾斜面を成しており、
前記タービンは、前記ロータを中心として環状に且つ前記ガス入口に沿って配置された複数の第一段静翼を有し、該第一段静翼は、翼弦が伸びる翼弦方向が前記周方向に対して傾斜しており、
前記周方向で、前記第一段静翼の上流端に対して該第一段静翼の下流端が存在する側を翼傾斜側とした場合、前記尾筒の前記一方の側壁の傾斜面は、前記尾筒における前記周方向で互いに対向する一対の前記側壁のうちの該翼傾斜側の側壁であり、
前記翼傾斜側の側壁の傾斜面の下流端は、前記第一段静翼における上流側を向く面であって該第一段静翼の上流端から下流端にわたって延びる腹側の面に対して、前記軸線方向に対向することを特徴とするガスタービン。 A plurality of combustors for mixing combustion with compressed air to generate combustion gas, and a turbine having a rotor rotated by the combustion gas from the plurality of combustors,
A plurality of combustors are annularly arranged around the rotor, and a gas turbine having a tail tube that sends combustion gas to a gas inlet of the turbine,
The inner surface of at least one of the pair of side walls facing each other in the circumferential direction of the rotor at the downstream portion of the transition piece of the combustor is directed toward the downstream side in the axial direction of the transition piece, An inclined surface that is inclined to the downstream end of the transition piece in a direction approaching the transition piece of another adjacent combustor gradually ,
The turbine has a plurality of first stage stationary blades arranged annularly around the rotor and along the gas inlet, and the first stage stationary blade has a chord direction in which the chord extends in the circumferential direction. Inclined with respect to the direction,
When the side where the downstream end of the first stage stationary blade exists with respect to the upstream end of the first stage stationary blade in the circumferential direction is a blade inclined side, the inclined surface of the one side wall of the tail tube is The side wall on the blade inclination side of the pair of side walls facing each other in the circumferential direction in the transition piece,
The downstream end of the inclined surface of the side wall on the blade inclined side is a surface facing the upstream side in the first stage stationary blade, and is a surface on the ventral side extending from the upstream end to the downstream end of the first stage stationary blade. The gas turbine is opposed to the axial direction. 前記一方の側壁のうちの下流側の端部のみに前記傾斜面を有することを特徴とする請求項１に記載のガスタービン。The gas turbine according to claim 1, wherein the inclined surface is provided only at an end portion on the downstream side of the one side wall. 前記翼傾斜側の側壁の傾斜面によって下流側に導かれる燃焼ガスが、前記軸線方向に対向する腹側の面によってさらに下流側に導かれることを特徴とする請求項１又は２に記載のガスタービン。3. The gas according to claim 1, wherein the combustion gas guided downstream by the inclined surface of the blade inclined side wall is guided further downstream by the ventral surface facing the axial direction. Turbine. 燃料を圧縮空気に混合して燃焼させ燃焼ガスを生成する複数の燃焼器と、複数の該燃焼器からの燃焼ガスにより回転するロータを有するタービンと、を備え、A plurality of combustors for mixing combustion with compressed air to generate combustion gas, and a turbine having a rotor rotated by the combustion gas from the plurality of combustors,
複数の燃焼器は、前記ロータを中心として環状に配置され、前記タービンのガス入口に燃焼ガスを送る尾筒を有するガスタービンにおいて、A plurality of combustors are annularly arranged around the rotor, and a gas turbine having a tail tube that sends combustion gas to a gas inlet of the turbine,
前記燃焼器の前記尾筒の下流部で、前記ロータの周方向で互いに対向する一対の側壁のうち、少なくとも一方の側壁の内面は、該尾筒の軸線方向の下流側に向うに連れて、次第に隣接する他の燃焼器の尾筒に近づく向きに、該尾筒の下流端に至るまで傾斜している傾斜面を成しており、The inner surface of at least one of the pair of side walls facing each other in the circumferential direction of the rotor at the downstream portion of the transition piece of the combustor is directed toward the downstream side in the axial direction of the transition piece, An inclined surface that is inclined to the downstream end of the transition piece in a direction approaching the transition piece of another adjacent combustor gradually,
前記傾斜面は、前記尾筒の軸線に近づき且つ下流側に向う側に滑らかに膨らんだ曲面を少なくとも一部に含む、The inclined surface includes, at least in part, a curved surface that approaches the axis of the tail tube and smoothly bulges toward the downstream side.
ことを特徴とするガスタービン。A gas turbine characterized by that. 燃料を圧縮空気に混合して燃焼させ燃焼ガスを生成する複数の燃焼器と、複数の該燃焼器からの燃焼ガスにより回転するロータを有するタービンと、を備え、A plurality of combustors for mixing combustion with compressed air to generate combustion gas, and a turbine having a rotor rotated by the combustion gas from the plurality of combustors,
複数の燃焼器は、前記ロータを中心として環状に配置され、前記タービンのガス入口に燃焼ガスを送る尾筒を有するガスタービンにおいて、  A plurality of combustors are annularly arranged around the rotor, and a gas turbine having a tail tube that sends combustion gas to a gas inlet of the turbine,
前記燃焼器の前記尾筒の下流部で、前記ロータの周方向で互いに対向する一対の側壁のうち、少なくとも一方の側壁の内面は、該尾筒の軸線方向の下流側に向うに連れて、次第に隣接する他の燃焼器の尾筒に近づく向きに、該尾筒の下流端に至るまで傾斜している傾斜面を成しており、  The inner surface of at least one of the pair of side walls facing each other in the circumferential direction of the rotor at the downstream portion of the transition piece of the combustor is directed toward the downstream side in the axial direction of the transition piece, An inclined surface that is inclined to the downstream end of the transition piece in a direction approaching the transition piece of another adjacent combustor gradually,
前記燃焼器の数と前記第一段静翼の数との比が２：３以上の奇数であり、A ratio of the number of the combustors and the number of the first stage stationary blades is an odd number of 2: 3 or more;
複数の前記第一段静翼のピッチ寸法Ｐを基準にして、前記燃焼器の前記尾筒と前記他の燃焼器の尾筒との中間地点から前記周方向で最も近い第一段静翼の上流端までの該周方向の寸法Ｓの割合Ｓ／Ｐは、０．０５以下、０．２から０．５５の間、及び０．７から１．０の間である、  Based on the pitch dimension P of the plurality of first stage stator blades, the first stage stator blades closest in the circumferential direction from an intermediate point between the tail cylinder of the combustor and the tail cylinder of the other combustor The ratio S / P of the circumferential dimension S to the upstream end is 0.05 or less, between 0.2 and 0.55, and between 0.7 and 1.0.
ことを特徴とするガスタービン。A gas turbine characterized by that. 請求項４又は５に記載のガスタービンにおいて、The gas turbine according to claim 4 or 5,
前記タービンは、前記ロータを中心として環状に且つ前記ガス入口に沿って配置された複数の第一段静翼を有し、該第一段静翼は、翼弦が伸びる翼弦方向が前記周方向に対して傾斜しており、  The turbine has a plurality of first stage stationary blades arranged annularly around the rotor and along the gas inlet, and the first stage stationary blade has a chord direction in which the chord extends in the circumferential direction. Inclined with respect to the direction,
前記周方向で、前記第一段静翼の上流端に対して該第一段静翼の下流端が存在する側を翼傾斜側とした場合、前記尾筒の前記少なくとも一方の側壁は、前記尾筒における前記周方向で互いに対向する一対の前記側壁のうちの該翼傾斜側の側壁である、In the circumferential direction, when the side where the downstream end of the first stage stationary blade exists with respect to the upstream end of the first stage stationary blade is a blade inclined side, the at least one side wall of the tail tube is Of the pair of side walls facing each other in the circumferential direction in the tail cylinder, the side wall on the blade inclination side,
ことを特徴とするガスタービン。A gas turbine characterized by that. 請求項１から６のいずれか一項に記載のガスタービンにおいて、
前記尾筒における前記周方向で互いに対向する一対の前記側壁の両方の前記内面が前記傾斜面を成している、
ことを特徴とするガスタービン。 The gas turbine according to any one of claims 1 to 6 ,
The inner surfaces of both of the pair of side walls facing each other in the circumferential direction in the tail tube form the inclined surface.
A gas turbine characterized by that. タービンのロータを中心として隣接するように環状に複数配置されて前記タービンのガス入口に燃焼ガスを送る燃焼器の尾筒であって、A plurality of annularly arranged so as to be adjacent to each other centering on a rotor of a turbine, and a combustor of a combustor that sends combustion gas to a gas inlet of the turbine,
前記ロータの周方向で互いに対向する一対の側壁のうち、一方の側壁の内面のみが、該尾筒の軸線方向の下流側に向うに連れて、次第に隣接する他の燃焼器の尾筒に近づく向きに、該尾筒の下流端に至るまで傾斜している傾斜面を成していることを特徴とする尾筒。Of the pair of side walls opposed to each other in the circumferential direction of the rotor, only the inner surface of one of the side walls gradually approaches the tail cylinder of another combustor that is adjacent to the inner side of the tail cylinder toward the downstream side in the axial direction. A tail tube characterized by having an inclined surface that is inclined in the direction to the downstream end of the tail tube. 前記一方の側壁は、前記周方向で互いに対向する一対の前記側壁のうち、前記ロータとともに回転する動翼の回転方向前方側の側壁であることを特徴とする請求項８に記載の尾筒。The tail pipe according to claim 8, wherein the one side wall is a side wall on a front side in a rotation direction of a moving blade rotating together with the rotor among a pair of the side walls facing each other in the circumferential direction. 周方向において、前記ロータを中心として環状に且つ前記ガス入口に沿って配置されて翼弦が伸びる翼弦方向が前記周方向に対して傾斜する第一段静翼の上流端に対して該第一段静翼の下流端が存在する側を翼傾斜側とした場合、前記一方の側壁は、前記周方向で互いに対向する一対の前記側壁のうちの該翼傾斜側の側壁であることを特徴とする請求項８又は９に記載の尾筒。In the circumferential direction, the first chord is arranged in an annular shape around the rotor and along the gas inlet, and the chord direction in which the chord extends extends relative to the upstream end of the first stage stationary blade inclined with respect to the circumferential direction. When the side where the downstream end of the first stage stationary blade exists is a blade inclined side, the one side wall is a side wall on the blade inclined side of the pair of side walls facing each other in the circumferential direction. The tail tube according to claim 8 or 9. 請求項８から１０のいずれか一項に記載の尾筒を備えることを特徴とする燃焼器。A combustor comprising the transition piece according to any one of claims 8 to 10. 請求項１１に記載の燃焼器と、
前記第一段静翼を有する前記タービンと、を備えることを特徴とするガスタービン。 A combustor according to claim 11;
A gas turbine comprising: the turbine having the first stage stationary blade . 請求項１から７、請求項１２のいずれか一項に記載のガスタービンにおいて、
前記傾斜面の上流端から下流端までの前記周方向の寸法Ｂを基準として、該上流端から該下流端までの前記軸線方向の寸法Ａの割合Ａ／Ｂは、１以上で８以下である、
ことを特徴とするガスタービン。 The gas turbine according to any one of claims 1 to 7 and claim 12 ,
On the basis of the circumferential dimension B from the upstream end to the downstream end of the inclined surface, the ratio A / B of the axial dimension A from the upstream end to the downstream end is 1 or more and 8 or less. ,
A gas turbine characterized by that. 請求項１から７及び請求項１２のいずれか一項に記載のガスタービンにおいて、
複数の前記第一段静翼のピッチ寸法Ｐを基準にして、前記尾筒の下流端から前記第一段静翼の上流端までの前記軸線方向の寸法Ｌの割合Ｌ／Ｐが０．２以下である、
ことを特徴とするガスタービン。 The gas turbine according to any one of claims 1 to 7 and claim 12 ,
The ratio L / P of the dimension L in the axial direction from the downstream end of the transition piece to the upstream end of the first stage stationary blade is 0.2 based on the pitch dimension P of the plurality of first stage stationary blades. Is
A gas turbine characterized by that.
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JP2013064535A (en) 2013-04-11
CN103782103A (en) 2014-05-07
US20140216055A1 (en) 2014-08-07
WO2013039095A1 (en) 2013-03-21
EP2752622A4 (en) 2015-04-29
CN103782103B (en) 2015-12-23

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