JP3426841B2 - Gas turbine blade - Google Patents
Gas turbine bladeInfo
- Publication number
- JP3426841B2 JP3426841B2 JP09220096A JP9220096A JP3426841B2 JP 3426841 B2 JP3426841 B2 JP 3426841B2 JP 09220096 A JP09220096 A JP 09220096A JP 9220096 A JP9220096 A JP 9220096A JP 3426841 B2 JP3426841 B2 JP 3426841B2
- Authority
- JP
- Japan
- Prior art keywords
- cooling
- steam
- passage
- blade
- gas turbine
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Expired - Fee Related
Links
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05B—INDEXING SCHEME RELATING TO WIND, SPRING, WEIGHT, INERTIA OR LIKE MOTORS, TO MACHINES OR ENGINES FOR LIQUIDS COVERED BY SUBCLASSES F03B, F03D AND F03G
- F05B2240/00—Components
- F05B2240/80—Platforms for stationary or moving blades
- F05B2240/801—Platforms for stationary or moving blades cooled platforms
Landscapes
- Turbine Rotor Nozzle Sealing (AREA)
Description
【発明の詳細な説明】
【０００１】
【発明の属する技術分野】本発明は，ガスタービン翼の
冷却技術に属する。
【０００２】
【従来の技術】従来から，コンバインドプラント等に用
いる高温のガスタービンにおいては，高温ガスの熱から
翼を保護するために，翼の内部に通路を設けて比較的低
温の冷却空気を流し，翼の温度をガス温度よりも低く抑
えている。この空気による翼の冷却方式では，翼根部か
ら供給された冷却空気を，翼内部の冷却通路を通過させ
た後，翼の外縁に設けられた穴から翼外部（タービンの
主流ガス中）に放出している。
【０００３】この空気冷却方式に対し，近年，蒸気によ
る翼の冷却方式が考えられている。蒸気による翼冷却方
式では，冷却に供した蒸気を放出せずに回収することと
しており，これによってガスタービンの熱効率向上が期
待できる。また，コンバインドプラントにおいては，回
収した蒸気を蒸気タービンに送ってプラント全体の効率
を改善することも可能となる。
【０００４】
【発明が解決しようとする課題】ガスタービンの動翼に
おいて，高温ガスの影響を直接受けるのは翼形部とプラ
ットホーム部であり，これらの部分には全体にわたって
均一に冷却を施す必要がある。通常，このような冷却を
施すには，内側面を翼形部またはプラットホーム部の外
面に沿わせた形状の蛇行通路（いわゆる，サーペンタイ
ン流路）を冷却通路として内部に設けることが考えられ
る。
【０００５】この場合，翼形部はある程度の厚みを有し
ているため，精密鋳造で作成した場合であっても比較的
容易に上記のような蛇行通路を内部に設けることが可能
である。これに対して，プラットホーム部は薄く広い構
造となっているため，精密鋳造で内部全体に上記のよう
な蛇行通路を設けることは困難であり，不経済でもあ
る。
【０００６】また，コンバインドプラント全体の効率を
上げるためには，ガスタービン翼の冷却過程で生じる圧
力損失を抑えて，できるだけ高い圧力を保ったまま回収
蒸気を蒸気タービンへ供給する必要があるが，プラット
ホーム部の内部全体に設けた蛇行通路に冷却蒸気を通す
と，圧力損失が非常に大きくなってしまい，さほどの効
率向上は見込めなくなってしまう。
【０００７】
【課題を解決するための手段】上記の課題を解決するた
め、本発明は、ガスタービン動翼において、翼根部に蒸
気供給口及び蒸気回収口を設け、該蒸気供給口及び蒸気
回収口と連通した蛇行通路を翼形部の内部に備え、前記
蛇行通路は翼前縁側から翼中央部に向かって蛇行し前記
蒸気回収口に通じる第１通路と翼後縁側から翼中央部に
向かって蛇行し前記蒸気回収口に通じる第２通路とから
形成され、かつシール空気を通してプラットホーム部を
対流冷却する対流冷却通路またはフィルム冷却するフィ
ルム冷却孔を該プラットホーム部に備えたことを特徴と
する。
【０００８】このような構成を採用したことにより，翼
形部は蒸気で冷却され，プラットホーム部は空気で冷却
されるようになる。プラットホーム部を冷却した空気は
タービン主流ガス中に放出されるが，該空気は元々ター
ビン主流ガス中に放出されるシール空気であるため，同
タービン主流ガス中に余分な冷却媒体を放出する必要は
ない。
【０００９】
【発明の実施の形態】図１は，本発明にかかるガスター
ビン動翼の一実施形態を示した縦断面図である。同図に
おいて，翼形部１０１の内部には，内側面を翼形部の外
面に沿わせた形状の蛇行通路（サーペンタイン流路）１
０３が設けてあり，翼根部には蒸気供給口１０４及び蒸
気回収口１０５が設けられている。サーペンタイン流路
１０３は，翼前縁側に位置する流路と翼後縁側に位置す
る流路に分割されている。
【００１０】蒸気供給口１０４から供給された冷却蒸気
は，翼根部内に設けられた二股通路によって２方向に分
割され，一方は翼前縁側のサーペンタイン流路１０３に
供給され，他方は翼後縁側のサーペンタイン流路１０３
に供給される。図１の矢印で示したように，どちらの流
路においても，蒸気供給口１０４から供給された冷却蒸
気は，翼の縁側から翼中央部に向かって蛇行して進むよ
うになっている。そして，翼中央部に送られた蒸気は翼
根部内の通路を通って前記蒸気回収口１０５へ進み，回
収される。
【００１１】図２は，図１のＡ−Ａ断面図である。同図
において，プラットホーム部１０２上には，複数のフィ
ルム冷却孔１０８が開口しており，同フィルム冷却孔１
０８から吹き出す冷却空気によってプラットホーム部１
０２をフィルム冷却する。フィルム冷却孔１０８に向か
って延びている破線は，同フィルム冷却孔１０８へ冷却
空気を供給する冷却空気通路である。
【００１２】図３は，図２のＢ−Ｂ断面図である。同図
において，プラットホーム部１０２には，冷却空気を流
すことによって対流冷却を行う対流冷却通路１０７が設
けられている。図４は，図２のＣ−Ｃ断面図を表してい
る。
【００１３】図５は，本発明の一実施形態にかかる冷却
蒸気及び冷却空気の供給，回収経路を示している。同図
において，冷却蒸気１０６は，タービンロータ１１０を
通って，第１段動翼に供給される。第１段動翼を冷却し
た蒸気は，タービンロータ１１０内を通過し，第２段動
翼を冷却した後に同タービンロータ１１０を通って回収
される。
【００１４】また，プラットホーム部１０２の冷却に
は，圧縮器から抽気されたシール空気を用いている。シ
ール側とタービン主流側との間には圧力差が存在するの
で，同空気はプラットホーム部１０２に設けた対流冷却
通路１０７及びフィルム冷却孔１０８を介してタービン
主流ガス中に流出する。その際，プラットホーム部１０
２の冷却に寄与することとなる。
【００１５】
【発明の効果】本発明によれば、翼形部の内部に蒸気供
給口及び蒸気回収口と連通した蛇行通路を上記のように
翼前縁側から翼中央部に向かって蛇行し前記蒸気回収口
に通じる第１通路と翼後縁側から翼中央部に向かって蛇
行し前記蒸気回収口に通じる第２通路とから形成される
ことにより、蒸気供給口から供給された蒸気にて最初に
高温になりやすい動翼の前縁部や後縁部を冷却すること
ができる。また、シール空気を通してプラットホーム部
を対流冷却する対流冷却通路またはフィルム冷却するフ
ィルム冷却孔をプラットホーム部に備えているので、翼
形部及びプラットホーム部の全体にわたって均一に冷却
することができる。また、翼形部及びプラットホーム部
の冷却性能を低下させることなく、ガスタービンの熱効
率を向上し、また翼の製造コストを抑えることができる
ようになる。また、本発明をコンバインドプラントにお
けるガスタービンに適用すれば、プラント全体の効率ア
ップにつながることとなる。Description: BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a technology for cooling gas turbine blades. 2. Description of the Related Art Conventionally, in a high-temperature gas turbine used in a combined plant or the like, in order to protect the blade from the heat of the high-temperature gas, a passage is provided inside the blade to supply relatively low-temperature cooling air. The temperature of the wing is kept lower than the gas temperature. In this blade cooling method using air, the cooling air supplied from the root of the blade passes through the cooling passage inside the blade, and is then discharged to the outside of the blade (in the mainstream gas of the turbine) from a hole provided at the outer edge of the blade. are doing. In recent years, a method of cooling blades using steam has been considered in contrast to the air cooling method. In the blade cooling method using steam, the steam used for cooling is collected without being released, which is expected to improve the thermal efficiency of the gas turbine. In a combined plant, the recovered steam can be sent to a steam turbine to improve the efficiency of the entire plant. [0004] In the moving blade of a gas turbine, the hot gas directly affects the airfoil portion and the platform portion, and these portions need to be uniformly cooled throughout. There is. Usually, in order to perform such cooling, it is conceivable to provide a meandering passage (a so-called serpentine flow passage) having a shape in which the inner surface is along the outer surface of the airfoil portion or the platform portion as a cooling passage inside. In this case, since the airfoil has a certain thickness, it is possible to relatively easily provide the above-mentioned meandering passage inside even when it is made by precision casting. On the other hand, since the platform has a thin and wide structure, it is difficult and uneconomical to provide such a meandering passage in the entire interior by precision casting. Further, in order to increase the efficiency of the combined plant as a whole, it is necessary to suppress the pressure loss generated in the cooling process of the gas turbine blades and supply the recovered steam to the steam turbine while maintaining the pressure as high as possible. When cooling steam is passed through a meandering passage provided in the entire interior of the platform, the pressure loss becomes very large, and it is not possible to expect much improvement in efficiency. [0007] In order to solve the above-mentioned problems, the present invention provides a gas turbine moving blade, wherein a steam supply port and a steam recovery port are provided at a blade root portion, and the steam supply port and the steam recovery port are provided. includes a serpentine passage through the mouth and with the interior of the airfoil, the
The meandering passage meanders from the leading edge of the wing toward the center of the wing.
The first passage leading to the steam recovery port and from the trailing edge to the center of the blade
From the second passage leading to the steam recovery port
The platform portion is provided with a convection cooling passage formed therein and convectively cooling the platform portion through seal air or a film cooling hole for film cooling. By adopting such a configuration, the airfoil portion is cooled by steam, and the platform portion is cooled by air. The air that has cooled the platform part is released into the mainstream gas of the turbine, but since this air is originally seal air that is released into the mainstream gas of the turbine, it is not necessary to discharge extra cooling medium into the mainstream gas of the turbine. Absent. FIG. 1 is a longitudinal sectional view showing one embodiment of a gas turbine rotor blade according to the present invention. In FIG. 1, a meandering passage (serpentine flow passage) 1 having an inner surface along the outer surface of the airfoil is provided inside the airfoil 101.
A steam supply port 104 and a steam recovery port 105 are provided at the blade root. The serpentine channel 103 is divided into a channel located on the blade leading edge side and a channel located on the blade trailing edge side. The cooling steam supplied from the steam supply port 104 is divided into two directions by a forked passage provided in the blade root, one of which is supplied to the serpentine flow channel 103 on the blade leading edge side, and the other is supplied to the blade trailing edge side. Serpentine channel 103
Supplied to As shown by the arrows in FIG. 1, in both flow paths, the cooling steam supplied from the steam supply port 104 meanders from the edge of the blade toward the center of the blade. Then, the steam sent to the center of the blade travels to the steam recovery port 105 through a passage in the root of the blade and is recovered. FIG. 2 is a sectional view taken along line AA of FIG. In the figure, a plurality of film cooling holes 108 are opened on a platform 102,
08 by the cooling air blown out from the
02 is film cooled. A broken line extending toward the film cooling hole 108 is a cooling air passage for supplying cooling air to the film cooling hole 108. FIG. 3 is a sectional view taken along line BB of FIG. In the figure, a convection cooling passage 107 for performing convection cooling by flowing cooling air is provided in a platform section 102. FIG. 4 is a cross-sectional view taken along the line CC of FIG. FIG. 5 shows a supply and recovery path of cooling steam and cooling air according to an embodiment of the present invention. In the figure, cooling steam 106 is supplied to a first-stage bucket through a turbine rotor 110. The steam that has cooled the first stage rotor blades passes through the turbine rotor 110, cools the second stage rotor blades, and is recovered through the turbine rotor 110. The cooling of the platform 102 is performed by using seal air extracted from the compressor. Since there is a pressure difference between the seal side and the turbine mainstream side, the air flows out into the turbine mainstream gas through the convection cooling passage 107 and the film cooling hole 108 provided in the platform 102. At that time, the platform unit 10
2 contributes to cooling. According to the present invention , a steam supply is provided inside the airfoil.
The meandering passage communicating with the supply port and the steam recovery port as described above
Meandering from the leading edge of the blade toward the center of the blade, the steam recovery port
From the trailing edge side of the wing to the center of the wing
And a second passage leading to the steam recovery port.
By using the steam supplied from the steam supply port,
Cool the leading and trailing edges of the moving blades, which are likely to be hot
Can be. In addition, the platform section through the seal air
Convection cooling passage or film cooling
Because the cooling holes are provided in the platform,
Uniform cooling over the profile and platform
can do. Further, the thermal efficiency of the gas turbine can be improved without lowering the cooling performance of the airfoil portion and the platform portion, and the blade manufacturing cost can be reduced. Further, if the present invention is applied to a gas turbine in a combined plant, the efficiency of the whole plant will be improved.
【図面の簡単な説明】
【図１】本発明の一実施形態にかかるガスタービン動翼
の縦断面図。
【図２】図１におけるＡ−Ａ断面図。
【図３】図２におけるＢ−Ｂ断面図。
【図４】図２におけるＣ−Ｃ断面図。
【図５】本発明の一実施形態にかかる冷却蒸気及び冷却
空気の供給，回収経路を示したガスタービン内部断面
図。
【符号の説明】
１０１ 翼形部
１０２ プラットホーム部
１０３ サーペンタイン流路
１０４ 蒸気導入口
１０５ 蒸気回収口
１０６ 冷却蒸気
１０７ 対流冷却通路
１０８ フィルム冷却孔
１０９ 冷却空気
１１０ タービンロータBRIEF DESCRIPTION OF THE DRAWINGS FIG. 1 is a longitudinal sectional view of a gas turbine blade according to an embodiment of the present invention. FIG. 2 is a sectional view taken along line AA in FIG. FIG. 3 is a sectional view taken along line BB in FIG. 2; FIG. 4 is a sectional view taken along line CC in FIG. 2; FIG. 5 is an internal cross-sectional view of a gas turbine showing supply and recovery paths of cooling steam and cooling air according to an embodiment of the present invention. DESCRIPTION OF SYMBOLS 101 Airfoil 102 Platform 103 Serpentine flow passage 104 Steam introduction port 105 Steam recovery port 106 Cooling steam 107 Convection cooling passage 108 Film cooling hole 109 Cooling air 110 Turbine rotor
───────────────────────────────────────────────────── フロントページの続き (56)参考文献 特開 平７−332004（ＪＰ，Ａ) 特開 平８−42302（ＪＰ，Ａ) 特開 平８−240102（ＪＰ，Ａ) 特開 昭55−117004（ＪＰ，Ａ) 特開 平６−257405（ＪＰ，Ａ) 特開 平７−305602（ＪＰ，Ａ) 特開 昭54−13809（ＪＰ，Ａ) 特開 昭55−5490（ＪＰ，Ａ) 特開 平７−189604（ＪＰ，Ａ) 特許2851575（ＪＰ，Ｂ２) 特許2851578（ＪＰ，Ｂ２) (58)調査した分野(Int.Cl.7，ＤＢ名) F01D 1/00 - 11/10 ──────────────────────────────────────────────────続 き Continuation of the front page (56) References JP-A-7-332004 (JP, A) JP-A 8-42302 (JP, A) JP-A 8-240102 (JP, A) JP-A 55- 117004 (JP, A) JP-A-6-257405 (JP, A) JP-A-7-305602 (JP, A) JP-A-54-13809 (JP, A) JP-A-55-5490 (JP, A) JP-A-7-189604 (JP, A) Patent 2851575 (JP, B2) Patent 2851578 (JP, B2) (58) Fields investigated (Int. Cl. 7 , DB name) F01D 1/00-11/10
Claims (1)
気供給口及び蒸気回収口を設け、該蒸気供給口及び蒸気
回収口と連通した蛇行通路を翼形部の内部に備え、前記
蛇行通路は翼前縁側から翼中央部に向かって蛇行し前記
蒸気回収口に通じる第１通路と翼後縁側から翼中央部に
向かって蛇行し前記蒸気回収口に通じる第２通路とから
形成され、かつシール空気を通してプラットホーム部を
対流冷却する対流冷却通路またはフィルム冷却するフィ
ルム冷却孔を該プラットホーム部に備えたことを特徴と
するガスタービン動翼。(1) In a gas turbine rotor blade, a steam supply port and a steam recovery port are provided at a blade root portion, and a meandering passage communicating with the steam supply port and the steam recovery port is formed in an airfoil portion. provided in the interior, the
The meandering passage meanders from the leading edge of the wing toward the center of the wing.
The first passage leading to the steam recovery port and from the trailing edge to the center of the blade
From the second passage leading to the steam recovery port
A gas turbine rotor blade having a convection cooling passage formed therein and a convection cooling passage for convectively cooling the platform portion through seal air or a film cooling hole for film cooling.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP09220096A JP3426841B2 (en) | 1996-04-15 | 1996-04-15 | Gas turbine blade |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP09220096A JP3426841B2 (en) | 1996-04-15 | 1996-04-15 | Gas turbine blade |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| JPH09280002A JPH09280002A (en) | 1997-10-28 |
| JP3426841B2 true JP3426841B2 (en) | 2003-07-14 |
Family
ID=14047811
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| JP09220096A Expired - Fee Related JP3426841B2 (en) | 1996-04-15 | 1996-04-15 | Gas turbine blade |
Country Status (1)
| Country | Link |
|---|---|
| JP (1) | JP3426841B2 (en) |
Families Citing this family (9)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CA2263576C (en) * | 1998-01-20 | 2003-08-05 | Mitsubishi Heavy Industries, Ltd. | Stationary blade of gas turbine |
| CA2262064C (en) | 1998-02-23 | 2002-09-03 | Mitsubishi Heavy Industries, Ltd. | Gas turbine moving blade platform |
| ATE483098T1 (en) * | 1999-09-24 | 2010-10-15 | Gen Electric | GAS TURBINE BLADE WITH IMPACT-COOLED PLATFORM |
| US6422817B1 (en) * | 2000-01-13 | 2002-07-23 | General Electric Company | Cooling circuit for and method of cooling a gas turbine bucket |
| US6390774B1 (en) | 2000-02-02 | 2002-05-21 | General Electric Company | Gas turbine bucket cooling circuit and related process |
| EP1577497A1 (en) * | 2004-03-01 | 2005-09-21 | ALSTOM Technology Ltd | Internally cooled turbomachine blade |
| US7870742B2 (en) * | 2006-11-10 | 2011-01-18 | General Electric Company | Interstage cooled turbine engine |
| CN102859120B (en) * | 2011-04-14 | 2016-06-01 | 三菱重工业株式会社 | Gas turbine rotor blade and gas turbine |
| CN106661946B (en) * | 2014-09-08 | 2018-05-22 | 西门子能源公司 | Include the cooling turbine guide vane platform of forepart, centre and blade trailing cooling chamber wherein |
-
1996
- 1996-04-15 JP JP09220096A patent/JP3426841B2/en not_active Expired - Fee Related
Also Published As
| Publication number | Publication date |
|---|---|
| JPH09280002A (en) | 1997-10-28 |
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