WO1999036675A1 - Stationary blade of gas turbine - Google Patents

Stationary blade of gas turbine Download PDF

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Publication number
WO1999036675A1
WO1999036675A1 PCT/JP1998/000184 JP9800184W WO9936675A1 WO 1999036675 A1 WO1999036675 A1 WO 1999036675A1 JP 9800184 W JP9800184 W JP 9800184W WO 9936675 A1 WO9936675 A1 WO 9936675A1
Authority
WO
WIPO (PCT)
Prior art keywords
cooling
air
gas turbine
passage
steam
Prior art date
Application number
PCT/JP1998/000184
Other languages
French (fr)
Japanese (ja)
Inventor
Kazuo Uematsu
Kiyoshi Suenaga
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
Priority to JP8190717A priority Critical patent/JPH1037704A/en
Application filed by Mitsubishi Heavy Industries, Ltd. filed Critical Mitsubishi Heavy Industries, Ltd.
Priority to DE19880989T priority patent/DE19880989C2/en
Priority to CA002263576A priority patent/CA2263576C/en
Priority to US09/230,751 priority patent/US6315518B1/en
Priority to PCT/JP1998/000184 priority patent/WO1999036675A1/en
Publication of WO1999036675A1 publication Critical patent/WO1999036675A1/en

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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/186—Film cooling
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/02—Preventing or minimising internal leakage of working-fluid, e.g. between stages by non-contact sealings, e.g. of labyrinth type
    • F01D11/04—Preventing or minimising internal leakage of working-fluid, e.g. between stages by non-contact sealings, e.g. of labyrinth type using sealing fluid, e.g. steam
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/187—Convection cooling
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • 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
    • F05D2240/00—Components
    • F05D2240/80—Platforms for stationary or moving blades
    • F05D2240/81—Cooled platforms
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/201—Heat transfer, e.g. cooling by impingement of a fluid
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/205—Cooling fluid recirculation, i.e. after cooling one or more components is the cooling fluid recovered and used elsewhere for other purposes
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/221—Improvement of heat transfer
    • F05D2260/2212—Improvement of heat transfer by creating turbulence
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/221—Improvement of heat transfer
    • F05D2260/2214—Improvement of heat transfer by increasing the heat transfer surface
    • F05D2260/22141—Improvement of heat transfer by increasing the heat transfer surface using fins or ribs
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/232—Heat transfer, e.g. cooling characterized by the cooling medium
    • F05D2260/2322—Heat transfer, e.g. cooling characterized by the cooling medium steam

Definitions

  • the present invention relates to a gas turbine stator vane capable of effectively performing cooling with a simple configuration.
  • the steam passage shall be closed to the outside and have a supply port and a recovery port.
  • cooling air is introduced into the inner shroud 35; cooling air is introduced from the inlet 40; Cooling air flows through 39 to cool the inner shroud 35, and the cooled air flows through the film cooling holes 38 provided in the inner shroud 35 into the main gas flow F to cool the film.
  • the current gas turbine steam-cooled vanes are in the above-mentioned situation and certainly have an advantage in terms of efficiency.However, closed long bent passages are complicated in structure and difficult to manufacture. In addition to this, if the wall thickness is increased to ease the pressure resistance, the whole blade becomes rigid and disadvantageous to thermal stress.
  • the present invention has been provided to solve the above problems. DISCLOSURE OF THE INVENTION
  • the stationary blade of the gas turbine of the present invention takes the following measures.
  • FIG. 1 is a cross-sectional view of an embodiment of the present invention
  • FIG. 2 is an A-A arrow diagram of FIG. 1
  • FIG. 3 is a cross-sectional view of a steam-cooled gas turbine vane of FIG.
  • FIG. 4 is a sectional view taken along the line BB in FIG. BEST MODE FOR CARRYING OUT THE INVENTION
  • the cooling steam flows from the cooling steam inlet 13 on the outer shroud 2 side into the inward cooling passage 8 provided in the wing 6, similarly to the conventional one shown in FIGS. After passing through the passage 8, it turns on the inner shaft 1 side, passes through the outward cooling passage 9 provided in the wing 6, and is collected at the cooling steam outlet 14.
  • the cooling steam is not supplied into the outer shroud, but the cooling steam is supplied to the inward cooling of the wing 6. It is supplied and discharged from the outward cooling passage 9 to the cooling steam outlet 14.
  • the inner shroud 1 and outer shroud 2 of the gas turbine vane are provided with cooling air inlet holes 3 for air cooling.On the other side of the cooling air inlet holes 3,-are many small-diameter passages.
  • a cooling air outlet hole 5 is provided through an accommodation plate 4 having holes.
  • the rear end 7 of the wing 6 which has a thin profile in terms of air performance corresponds to a pair of inward cooling passages 8 and outward cooling passages 9 for cooling the rear end 7 with air.
  • a large number of holes 11 are provided on the tail side of the air passage 1 (3), and as shown by the arrow, the cooling air flowing through the air passage 1 G flows into the main gas flow F discharged from the hole 1.
  • a hole 12 is also formed in the air shroud 1 side of the air passage 10 in the blade height direction, and the hole 12 is a part of the cooling air flowing through the air passage 10.
  • portions other than the rear end portion 7 of the wing portion 6 are connected to the inward cooling passage 8 and the outside cooling passage 8.
  • the inner shroud 1 and the outer shroud 2 are cooled by the cooling steam flowing through the cooling passage 9.
  • the inner shroud 1 and the outer shroud 2 flow through the cooling air inlet hole 3 and flow out of the cooling air outlet hole 5.
  • Cooling which can simplify the structure and locally uncooled
  • the cooling air supplied into the inner shroud 1 and the outer shroud 2 flows through the cooling air outlet hole 5 from the cooling air outlet hole 5 to the blades.
  • the film is discharged to the surface side of the part 6 to cool the surface of the wing part 6 by film, so that the cooling air can be used effectively. Since the air is flowed through the air passage 10, it is lined up, which is advantageous in terms of strength as compared with the case where high-pressure steam is used.
  • the cross-sectional area of the air passage 10 is increased, and the wing S
  • the thickness of the end 7 can be reduced, and effective cooling can be performed, and Binfin cooling can be performed.
  • the cooling air that has cooled the rear end 7 of the wing 6 is discharged from a number of holes 11 provided on the tail side of the rear end 7 of the wing 6 and merges with the main gas flow F. A part of the cooling air is discharged inward from the hole 12 on the inner shroud side of the air passage 10 to become the inner seal air of the combustion gas passage, and the inner seal air can be easily secured.
  • the vane of the gas turbine according to the present invention the complexity of the structure, which has been a problem of the conventional steam-cooled vane, has been reduced, and a part of the vane has been reduced. This can be solved by using air cooling, and the sealer inside the combustion gas passage can be easily secured. In addition, it is advantageous in terms of strength, and it is not unreasonable for ripening stress. Therefore, the reliability of the gas turbine can be further improved.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Abstract

Stationary blade of a gas turbine capable of effective cooling in a state, in which a low-pressure cooling air eases a pressure withstanding strength and with a simple construction. An inner shroud (1) and outer shroud (3) are cooled by a cooling air, which passes through an impingement plate, and a trailing end of a vane portion (6) having a thin configuration is cooled by a cooling air, which flows in an air passage (10) and a part of which is discharged, as an inside seal air for a combustion gas passage, from a hole (12) in a side of the inner shroud (1).

Description

明 細 書 ガスタービンの静翼 擴分野 本発明は、 簡単な構成で冷却を効果的に行なうことができるようにしたガスタ —ビンの静翼に関する。  Description Field of the Invention The present invention relates to a gas turbine stator vane capable of effectively performing cooling with a simple configuration.
ガスタービンのタービン翼の高温部では従来圧縮機の吐出空気又は抽気空気で 冷却を行っていたが、 ガスタービンの効率向上の手段の一つとして空気に代って 蒸気を使用してガスタービンの静翼の冷却を行うことが考えられている。 ガスタ一ビンの静翼の蒸気冷却においては、 蒸気はガスタービンと組合せてコ ンバインドサイクルを構成する蒸気タービンの抽気が使用されていて通常蒸気圧 カは髙い。 また、 蒸気のガスタービン中への洩れは、 蒸気側サイクル上極めて少 くしなければならない。 従って、 ガスタービンの静翼の蒸気冷却においては、In the high-temperature part of the turbine blades of gas turbines, cooling was conventionally performed with the discharge air or bleed air from the compressor. It is considered to cool the vane. In the steam cooling of the stationary vanes of the gas turbine, the steam is extracted from the steam turbine that forms a combined cycle in combination with the gas turbine, and the steam pressure is usually high. In addition, the leakage of steam into the gas turbine must be extremely small in the steam cycle. Therefore, in steam cooling of the gas turbine stationary blade,
( 1 ) 蒸気圧に耐える耐圧強度があること。 (1) There must be pressure resistance to withstand vapor pressure.
( 2 ) 蒸気通路は外部に対して閉じていて供給口と回収口を有すること。  (2) The steam passage shall be closed to the outside and have a supply port and a recovery port.
.( 3 ) 熱応力が低いこと。  (3) Low thermal stress.
( 4 ) 製作が容易であること。 などが要求される。 徒来のガスタービンの静翼においては、 図 3及び図 4に矢印で示すように、 冷 却蒸気は、 外側シュラウド 3 1の冷却蒸気入口 3 2からガスタービンの静翼内に 入り、 ィンビンジメント板 3 9を通過した上翼部 3 3に設けられた内向き冷却通 路 3 4を通って内側シュラウド 3 5でターンし、 翼部 3 3に設けられた外向き冷 却通路 3 6を通って冷却蒸気出口 3 7で回収される。 また、 以上に加えて、 本特許出願人が特許出願した特願平 8— 7 4 9号におけ るように、 内側シュラウド 3 5のみ空気冷却したものも提案されている。 即ち、 特願平 8— 7 4 9号においては、 図 3に示されるように、 以上に加えて、 内側シ ュラウド 3 5内に冷却空^;入口 4 0より冷却空気を導入し、 ィンビンジメント板 3 9を通って冷却空気を流して内側シュラウド 3 5を冷却し、 冷却後の空気を内 側シュラウド 3 5に設けられたフィルム冷却孔 3 8より主ガス流れ F中に流して フィルム冷却を行うようにしている。 現状のガスタ一ビンの蒸気冷却静翼は、 前述のような状況にあり確かに効率上 はこれが有利であるが、 閉じられた長い折れ曲った通路は、 構造が複雑で製作が 困難であり、 またこれに加えて耐圧強度を楽にするために肉厚にすると翼全体が 剛になり熱応力に対して不利になるなどの問題がある。 本発明は、 以上の問題点を解消するために提供されたものである。 発明の開示 本発明のガスタービンの静翼は、 次の手段を講じた。 (4) Being easy to manufacture. Is required. In the conventional gas turbine vanes, as shown by arrows in Figs. 3 and 4, the cooling steam enters the gas turbine vanes from the cooling steam inlet 32 of the outer shroud 31 and the intake plate After passing through 3 9, it turns on the inner shroud 35 through the inward cooling passage 34 provided in the upper wing 33, and passes through the outward cooling passage 36 provided in the wing 33. Collected at the cooling steam outlet 37. Further, in addition to the above, as disclosed in Japanese Patent Application No. Hei 8-7449 filed by the present applicant, a method in which only the inner shroud 35 is air-cooled has been proposed. That is, in Japanese Patent Application No. Hei 8-749, as shown in FIG. 3, in addition to the above, cooling air is introduced into the inner shroud 35; cooling air is introduced from the inlet 40; Cooling air flows through 39 to cool the inner shroud 35, and the cooled air flows through the film cooling holes 38 provided in the inner shroud 35 into the main gas flow F to cool the film. Like that. The current gas turbine steam-cooled vanes are in the above-mentioned situation and certainly have an advantage in terms of efficiency.However, closed long bent passages are complicated in structure and difficult to manufacture. In addition to this, if the wall thickness is increased to ease the pressure resistance, the whole blade becomes rigid and disadvantageous to thermal stress. The present invention has been provided to solve the above problems. DISCLOSURE OF THE INVENTION The stationary blade of the gas turbine of the present invention takes the following measures.
( 1 ) 内側シュラウドと外側シュラウドを空気冷却することを特徴とする。 (1) The inner shroud and the outer shroud are air-cooled.
( 2 ) 薄 (、形状を有する翼部の後端部を空気冷却することを特徵とする。 (2) Thin (The feature is to air-cool the rear end of the wing having the shape.
( 3 ) 前記 （2 ) の本発明において、 翼部の後端部を冷却した空気の一部を、 内側シュラウド側より燃焼ガス通路の内側シールェ了として排出することを特徵 とする。 前記本発明 （1 ) では、 ガスタービンの静翼の内側シュラウドと外側シュラウ ドを空気冷却することによって構成が単純化され、 かつ、 局所的に無冷却部が発 生して焼付けが生ずることを防止することができる。 前記本発明 （2 ) では、 空気性能上薄い形状となるガスタービンの静翼の後凝 部を空気冷却することによって、 圧力の髙 、蒸気で冷却する場合と比较して強度 上有利となると共に、 ビンフィン冷却が可能になる。 前記本癸明 （3 ) では、 前記本発明 （2 ) においてガスタービンの静翼の後端 部を冷却した空気の一部を内側シュラウド側より排出することによって、 構成を 複雑にすることなく燃焼ガス通路の内側シールエアを容易に確保することができ る。 図面の簡単な説明 図 1は本発明の実施の一形態の断面図、図 2は図 1の A— A矢 図、 図 3 は の蒸気冷却方式のガスタービンの静翼の断面図、 図 4は図 3の B— B矢視 断面図である。 究明を実施するための最良の形態 本発明の実施の一形態を、 図 1及び図 2によって説明する。 本実施の形態では 、 図 3及び ΪΙ 4に示す従来のものと同様に、 冷却蒸気は、 外側シュラウド 2側の 冷却蒸気入口 1 3から翼部 6に設けられた内向き冷却通路 8に入り、 同通路 8を 通って内側シユラゥド 1側でターンし、 翼部 6に設けられた外向き冷却通路 9を 通って冷却蒸気出口 1 4で回収される。 ただ、 前記図 4及び図 5に示すものにお けるように、 外側シュラウド内に前記冷却蒸気が供給されるのではなく、 冷却蒸 気は、 翼部 6の前記内向き冷却.通路 8に直接供給され、 かつ、 前記外向き冷却通 路 9から冷却蒸気出口 1 4に排出されるようになっている。 ガスタービンの静翼の内側シュラウド 1及び外側シュラウド 2には、 空気冷却 を行うため冷却空気入口孔 3が設けられており、 同冷却空気入口孔 3の反対側に- は、 多数の小径の通孔を有するィ ンビンジメ ント板 4を介して冷却空気出口孔 5 が設けられている。 また、 空気性能上薄い形拔となる翼部 6の後端部 7には、 同後端部 7を空気冷 却するため、 一組の内向き冷却通路 8と外向き冷却通路 9に相当する部分にビン フィンを設けた一つの空気通路 i 0が設けられており、 同空気通路 1 0内を冷却 空気が、 矢印に示すように、 外側シュラウド 2側から内側シユラゥド 1側へ流れ るようになっている。 前記空気通路 1 (3の翼尾側には多数の孔 1 1が設けられていて、 矢印に示すよ うに、 空気通路 1 G内を流れる冷却空気が孔 1より排出される主ガス流れ Fに合 流する。 また、 空気通路 1 0の内側シュラウド 1側にも翼高さ方向に孔 1 2が設 けられており、 同孔 1 2は空気通路 1 0内を流れた冷却空気の一部を燃焼ガス通 路の内側シールエアとして排出する供給口としている。 . 以上のように構成された本実施の形態では、 翼部 6の後端部 7以外の部分は、 内向き冷却通路 8と外向き冷却通路 9内を流れる冷却'蒸気によつて冷却される。 また、 内側シュラウド 1と外側シュラウド 2は冷却空気入口孔 3より流入し冷 却空気出ロ孔 5から徘出される冷却 気によつて冷却されるようになっていて、 構造を単純化することができ、 かつ、 局所的に無冷却の部分が発生して焼付けを 生ずることを防止することができる。 更に、 内側シュラウド 1内と外側シュラウ ド 2内に供給された冷却空気は、 インビンジメント扳 4を通って冷却空気出口孔 5から翼部 6の表面側に排出されて翼部 6の表面をフィルム冷却し、 冷却空気を 効果的に活用することができる。 また更に、 空気性能上薄い形状になる翼部 6の後端部 7は空気通路 1 Ό内を流 れる空気によって袷却されるので、 圧力の高い蒸気を用いる場合に比较して強度 上有利となり、 空気通路 1 0の断面積を大きぐし、 かつ、 翼部 Sの後端部 7の肉 厚を薄くすることが可能で効果的な冷却を行なうことができ、 かつ、 ビンフィ ン 冷却が可能となる。 前記翼部 6の後端部 7を冷却した冷却空気は翼部 6の後端部 7の翼尾側に設け られた多数の孔 1 1から排出されて主ガス流れ Fに合流すると共に、 前記冷却空 気の一部は空気通路 1 0の内側シュラウド側の孔 1 2より内向きに排出されて燃 焼ガス通路の内側シールエアとなり、 同内側シールエアを容易に確保することが できる。 産 ¾:の利用可 fgft 本発明に係るガスタービンの静翼によれば、 従来の蒸気冷却静翼の問題点であ つた構造の複雑さを製作の因難さを、 静翼の一部を空気冷却としたことによって 解消することができるとともに、 燃焼ガス通路内側のシールェァを容易に確保す ることができる。 また、 強度的に有利となって熟応力に.対しても無理がない。 従 つて、 ガスタービンの信頼性を一段と向上させることができる。 (3) In the present invention of (2), a part of the air cooled at the rear end portion of the wing portion is discharged from the inner shroud side as an inner seal of the combustion gas passage. According to the present invention (1), the structure is simplified by air cooling the inner shroud and the outer shroud of the stationary blade of the gas turbine, and the non-cooling portion is locally generated to cause burning. Can be prevented. According to the present invention (2), by cooling the trailing portion of the stationary blade of the gas turbine, which has a thin shape in terms of air performance, with air, the strength is more advantageous than when cooling with pressure and steam, and , Bin fin cooling becomes possible. According to the present invention (3), a part of the air that has cooled the rear end portion of the stationary blade of the gas turbine in the present invention (2) is discharged from the inner shroud side, thereby burning without complicating the structure. The seal air inside the gas passage can be easily secured. BRIEF DESCRIPTION OF THE DRAWINGS FIG. 1 is a cross-sectional view of an embodiment of the present invention, FIG. 2 is an A-A arrow diagram of FIG. 1, FIG. 3 is a cross-sectional view of a steam-cooled gas turbine vane of FIG. FIG. 4 is a sectional view taken along the line BB in FIG. BEST MODE FOR CARRYING OUT THE INVENTION One embodiment of the present invention will be described with reference to FIGS. In the present embodiment, the cooling steam flows from the cooling steam inlet 13 on the outer shroud 2 side into the inward cooling passage 8 provided in the wing 6, similarly to the conventional one shown in FIGS. After passing through the passage 8, it turns on the inner shaft 1 side, passes through the outward cooling passage 9 provided in the wing 6, and is collected at the cooling steam outlet 14. However, as shown in FIGS. 4 and 5, the cooling steam is not supplied into the outer shroud, but the cooling steam is supplied to the inward cooling of the wing 6. It is supplied and discharged from the outward cooling passage 9 to the cooling steam outlet 14. The inner shroud 1 and outer shroud 2 of the gas turbine vane are provided with cooling air inlet holes 3 for air cooling.On the other side of the cooling air inlet holes 3,-are many small-diameter passages. A cooling air outlet hole 5 is provided through an accommodation plate 4 having holes. In addition, the rear end 7 of the wing 6 which has a thin profile in terms of air performance corresponds to a pair of inward cooling passages 8 and outward cooling passages 9 for cooling the rear end 7 with air. There is one air passage i0 provided with bin fins in the part, so that cooling air flows in the air passage 10 from the outer shroud 2 side to the inner shroud 1 side as shown by the arrow. Has become. A large number of holes 11 are provided on the tail side of the air passage 1 (3), and as shown by the arrow, the cooling air flowing through the air passage 1 G flows into the main gas flow F discharged from the hole 1. A hole 12 is also formed in the air shroud 1 side of the air passage 10 in the blade height direction, and the hole 12 is a part of the cooling air flowing through the air passage 10. In the present embodiment configured as described above, portions other than the rear end portion 7 of the wing portion 6 are connected to the inward cooling passage 8 and the outside cooling passage 8. The inner shroud 1 and the outer shroud 2 are cooled by the cooling steam flowing through the cooling passage 9. The inner shroud 1 and the outer shroud 2 flow through the cooling air inlet hole 3 and flow out of the cooling air outlet hole 5. Cooling, which can simplify the structure and locally uncooled In addition, the cooling air supplied into the inner shroud 1 and the outer shroud 2 flows through the cooling air outlet hole 5 from the cooling air outlet hole 5 to the blades. The film is discharged to the surface side of the part 6 to cool the surface of the wing part 6 by film, so that the cooling air can be used effectively. Since the air is flowed through the air passage 10, it is lined up, which is advantageous in terms of strength as compared with the case where high-pressure steam is used.The cross-sectional area of the air passage 10 is increased, and the wing S The thickness of the end 7 can be reduced, and effective cooling can be performed, and Binfin cooling can be performed. The cooling air that has cooled the rear end 7 of the wing 6 is discharged from a number of holes 11 provided on the tail side of the rear end 7 of the wing 6 and merges with the main gas flow F. A part of the cooling air is discharged inward from the hole 12 on the inner shroud side of the air passage 10 to become the inner seal air of the combustion gas passage, and the inner seal air can be easily secured. According to the vane of the gas turbine according to the present invention, the complexity of the structure, which has been a problem of the conventional steam-cooled vane, has been reduced, and a part of the vane has been reduced. This can be solved by using air cooling, and the sealer inside the combustion gas passage can be easily secured. In addition, it is advantageous in terms of strength, and it is not unreasonable for ripening stress. Therefore, the reliability of the gas turbine can be further improved.

Claims

請 求 の 範 囲 The scope of the claims
1 . 内側シュラウドと外側シュラウドを空気冷却することを特徴 とするガスタービンの静翼。 1. A gas turbine vane characterized by air cooling of the inner shroud and outer shroud.
： 2. 薄い形拔を有する翼部の後端部を空気冷却することを特徴と するガスタービンの静翼。  : 2. A gas turbine vane characterized by air-cooling the rear end of the wing with a thin profile.
.  .
3. 前記翼部の後端部を冷却した空気の一部を、 内側シュラウド 側より燃焼ガス通路の内側シールエアとして排出することを特徵とする請求項 2 に記載のガスタ一ビンの静翼。 3. The gas vane vane according to claim 2, wherein a part of the air that has cooled the rear end of the vane is discharged from the inside shroud side as the inside seal air of the combustion gas passage.
PCT/JP1998/000184 1996-07-19 1998-01-20 Stationary blade of gas turbine WO1999036675A1 (en)

Priority Applications (5)

Application Number Priority Date Filing Date Title
JP8190717A JPH1037704A (en) 1996-07-19 1996-07-19 Stator blade of gas turbine
DE19880989T DE19880989C2 (en) 1998-01-20 1998-01-20 Stationary blade of a gas turbine
CA002263576A CA2263576C (en) 1998-01-20 1998-01-20 Stationary blade of gas turbine
US09/230,751 US6315518B1 (en) 1998-01-20 1998-01-20 Stationary blade of gas turbine
PCT/JP1998/000184 WO1999036675A1 (en) 1996-07-19 1998-01-20 Stationary blade of gas turbine

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
JP8190717A JPH1037704A (en) 1996-07-19 1996-07-19 Stator blade of gas turbine
PCT/JP1998/000184 WO1999036675A1 (en) 1996-07-19 1998-01-20 Stationary blade of gas turbine

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US6647624B2 (en) 2001-07-05 2003-11-18 Alstom (Switzerland) Ltd Method of fitting an impingement plate
EP1526251A1 (en) * 2003-10-22 2005-04-27 General Electric Company Turbine nozzle cooling configuration
US7052233B2 (en) 2001-07-13 2006-05-30 Alstom Switzerland Ltd Base material with cooling air hole
DE10217484B4 (en) 2001-11-02 2018-05-17 Ansaldo Energia Ip Uk Limited Guide vane of a thermal turbomachine
CN112112688A (en) * 2019-06-21 2020-12-22 斗山重工业建设有限公司 Turbine stator blade, turbine including the same, and gas turbine
CN113266436A (en) * 2021-05-14 2021-08-17 西安交通大学 Channel structure for cooling inside of gas turbine stationary blade and gas turbine stationary blade

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JPH1037704A (en) * 1996-07-19 1998-02-10 Mitsubishi Heavy Ind Ltd Stator blade of gas turbine
DE19880989C2 (en) 1998-01-20 2002-01-24 Mitsubishi Heavy Ind Ltd Stationary blade of a gas turbine
JP3494879B2 (en) * 1998-03-25 2004-02-09 株式会社日立製作所 Gas turbine and gas turbine vane
US6406254B1 (en) * 1999-05-10 2002-06-18 General Electric Company Cooling circuit for steam and air-cooled turbine nozzle stage
US6517312B1 (en) * 2000-03-23 2003-02-11 General Electric Company Turbine stator vane segment having internal cooling circuits
US6887039B2 (en) * 2002-07-10 2005-05-03 Mitsubishi Heavy Industries, Ltd. Stationary blade in gas turbine and gas turbine comprising the same
JP2009013837A (en) * 2007-07-03 2009-01-22 Hitachi Ltd Gas turbine facility

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JPH0681675A (en) * 1992-09-03 1994-03-22 Hitachi Ltd Gas turbine and stage device therefor
JPH08165902A (en) * 1994-10-12 1996-06-25 Hitachi Ltd Ceramic stator blade
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JPH09264103A (en) * 1996-03-28 1997-10-07 Mitsubishi Heavy Ind Ltd Gas turbine stationary blade
JPH1037704A (en) * 1996-07-19 1998-02-10 Mitsubishi Heavy Ind Ltd Stator blade of gas turbine

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JPH02241902A (en) * 1989-03-13 1990-09-26 Toshiba Corp Cooling blade of turbine and combined generating plant utilizing gas turbine equipped with this blade
JPH0681675A (en) * 1992-09-03 1994-03-22 Hitachi Ltd Gas turbine and stage device therefor
JPH08165902A (en) * 1994-10-12 1996-06-25 Hitachi Ltd Ceramic stator blade
JPH09189203A (en) * 1996-01-08 1997-07-22 Mitsubishi Heavy Ind Ltd Gas turbine stator
JPH09264103A (en) * 1996-03-28 1997-10-07 Mitsubishi Heavy Ind Ltd Gas turbine stationary blade
JPH1037704A (en) * 1996-07-19 1998-02-10 Mitsubishi Heavy Ind Ltd Stator blade of gas turbine

Cited By (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US6647624B2 (en) 2001-07-05 2003-11-18 Alstom (Switzerland) Ltd Method of fitting an impingement plate
US7052233B2 (en) 2001-07-13 2006-05-30 Alstom Switzerland Ltd Base material with cooling air hole
DE10217484B4 (en) 2001-11-02 2018-05-17 Ansaldo Energia Ip Uk Limited Guide vane of a thermal turbomachine
EP1526251A1 (en) * 2003-10-22 2005-04-27 General Electric Company Turbine nozzle cooling configuration
US6929445B2 (en) 2003-10-22 2005-08-16 General Electric Company Split flow turbine nozzle
CN112112688A (en) * 2019-06-21 2020-12-22 斗山重工业建设有限公司 Turbine stator blade, turbine including the same, and gas turbine
CN112112688B (en) * 2019-06-21 2023-02-17 斗山重工业建设有限公司 Turbine stator blade, turbine including the same, and gas turbine
CN113266436A (en) * 2021-05-14 2021-08-17 西安交通大学 Channel structure for cooling inside of gas turbine stationary blade and gas turbine stationary blade

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