JP2007292006A - Turbine blade having cooling passage inside thereof - Google Patents

Turbine blade having cooling passage inside thereof Download PDF

Info

Publication number
JP2007292006A
JP2007292006A JP2006122884A JP2006122884A JP2007292006A JP 2007292006 A JP2007292006 A JP 2007292006A JP 2006122884 A JP2006122884 A JP 2006122884A JP 2006122884 A JP2006122884 A JP 2006122884A JP 2007292006 A JP2007292006 A JP 2007292006A
Authority
JP
Japan
Prior art keywords
blade
trailing edge
cooling
edge side
wall surface
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.)
Pending
Application number
JP2006122884A
Other languages
Japanese (ja)
Inventor
Ryo Akiyama
陵 秋山
Yasuhiro Horiuchi
康広 堀内
Shinya Marushima
信也 圓島
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Hitachi Ltd
Original Assignee
Hitachi Ltd
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Hitachi Ltd filed Critical Hitachi Ltd
Priority to JP2006122884A priority Critical patent/JP2007292006A/en
Priority to EP07008241A priority patent/EP1849960A3/en
Publication of JP2007292006A publication Critical patent/JP2007292006A/en
Pending legal-status Critical Current

Links

Images

Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/187—Convection cooling
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/20—Specially-shaped blade tips to seal space between tips and stator
    • 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
    • F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
    • F01D9/041—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector using blades
    • 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

Landscapes

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

Abstract

<P>PROBLEM TO BE SOLVED: To provide a reliable turbine blade without reducing efficiency of a gas turbine by effectively cooling the turbine blade. <P>SOLUTION: In the turbine blade equipped with the cooling passage inside thereof and a plurality of protrusions or connecting parts (for example, a cylindrical pin fin extending from a blade rear end side body side wall surface to a blade rear end side back side wall surface) connecting both of the blade rear end side body side wall surface and the blade rear end side back side wall surface on the blade rear end side body side wall surface and the blade rear end side back side wall surface of the cooling passage, the protrusion or the part connecting the both wall surfaces on the blade rear end side are arranged so that the flow passage resistance is different in accordance with a position on the wall surface of the blade rear end side and effective cooling is performed, and required quantity of a cooling medium is reduced. <P>COPYRIGHT: (C)2008,JPO&INPIT

Description

本発明は、内部に冷却通路を有するタービン翼に係わる。   The present invention relates to a turbine blade having a cooling passage therein.

ガスタービンは、圧縮機で圧縮した空気と燃料とを燃焼器で混合燃焼させて得た高温高圧の燃焼ガスでタービンを駆動させ、回転エネルギーを得る装置である。この回転エネルギーは、発電機を用いて電気エネルギーを得るため、またはポンプなどを駆動するための動力源として使用される。   A gas turbine is a device that obtains rotational energy by driving a turbine with high-temperature and high-pressure combustion gas obtained by mixing and burning air and fuel compressed by a compressor in a combustor. This rotational energy is used as a power source for obtaining electrical energy using a generator or driving a pump or the like.

最近では、ガスタービンと蒸気タービンとを組み合わせたコンバインドサイクルの効率向上に大きな期待がよせられている。効率向上の一手段として、作動媒体の高温高圧力比化があげられる。   Recently, great expectations have been placed on improving the efficiency of a combined cycle that combines a gas turbine and a steam turbine. One means for improving the efficiency is to increase the temperature and pressure ratio of the working medium.

また、ガスタービンの吸気空気に湿分を添加して高効率化を図る高湿分ガスタービン
（ＨＡＴ）発電プラントが注目されている。これは、燃焼用空気に水分を付加して増湿することでエネルギーポテンシャルを上げてガスタービン出力の増加を図り、発電プラントとして熱効率を向上させるものである。 In addition, a high humidity gas turbine (HAT) power plant that increases the efficiency by adding moisture to the intake air of the gas turbine has attracted attention. This increases moisture potential by adding moisture to the combustion air, thereby increasing the energy potential and increasing the output of the gas turbine, thereby improving the thermal efficiency of the power plant.

ガスタービン作動ガスの高温化や、水分を付加した高温ガスにより、タービン高温部の熱負荷は年々上昇する傾向にある。   Due to the high temperature of the gas turbine working gas and the high temperature gas to which moisture is added, the heat load of the turbine high temperature part tends to increase year by year.

タービン高温部の部品の一つであるタービン翼では、内部に中空の冷却通路を設け、この冷却通路内に冷却媒体を流通させることで、翼材の温度を許容温度以下に抑えられるように冷却している。この冷却媒体としては、圧縮機の抽気空気や吐出空気が用いられることが多いが、圧縮機からの空気を用いた場合、冷却空気流量の増加は燃焼用空気量の減少を意味し、ガスタービンの効率低下を招く。したがって、タービン翼の冷却は、より少量の空気で効率良く冷却することが望ましい。   Turbine blades, one of the components in the high temperature section of the turbine, are cooled so that the temperature of the blade material can be kept below the allowable temperature by providing a hollow cooling passage inside and circulating a cooling medium in the cooling passage. is doing. As this cooling medium, the bleed air or discharge air of the compressor is often used. However, when the air from the compressor is used, an increase in the cooling air flow rate means a decrease in the amount of combustion air. This leads to a decrease in efficiency. Therefore, it is desirable to cool the turbine blades efficiently with a smaller amount of air.

タービン翼を効率的に冷却するためには、冷却通路内における伝熱を促進する必要がある。伝熱を促進するためには、伝熱面表面の空気の流れを乱流とすること、あるいは境界層を破壊することが有効であり、翼内部の冷却面に多数の突起や障害物を設ける方法がある。タービン翼後縁側の冷却に着目したものとして、例えば、特開平６−１３７１０２号公報には、翼後縁側の被冷却部にピンフィンを有するガスタービン動翼の翼後縁側先端付近の冷却効果を高めるために、複数の冷却空気通路を備え、これら複数の冷却通路の出口を翼の後縁部に半径方向に並べて配置する技術が開示されている。同公報には、翼後縁側の被冷却部にピンフィンを有するガスタービン動翼の翼後縁側先端付近の冷却効率を高めるために、冷却空気の一部を翼先端側へ向かわせる仕切板を設ける技術も開示されている。また、特開平５−１５６９０１号公報には、翼後縁側の被冷却部にピンフィンを有するガスタービン静翼の後縁側を冷却するために、ピンフィン設置領域の翼先端側と根元側から冷却空気を供給する構成が開示されている。   In order to efficiently cool the turbine blade, it is necessary to promote heat transfer in the cooling passage. In order to promote heat transfer, it is effective to make the air flow on the surface of the heat transfer surface turbulent or to destroy the boundary layer, and provide many protrusions and obstacles on the cooling surface inside the blade. There is a way. For example, Japanese Patent Application Laid-Open No. 6-137102 discloses a cooling effect in the vicinity of the tip of the blade trailing edge of a gas turbine rotor blade having a pin fin in a cooled portion on the blade trailing edge. Therefore, a technique is disclosed in which a plurality of cooling air passages are provided, and outlets of the plurality of cooling passages are arranged side by side in the radial direction at the rear edge of the blade. In this publication, a partition plate for directing a part of the cooling air toward the tip of the blade is provided in order to increase the cooling efficiency in the vicinity of the tip of the blade trailing edge of the gas turbine rotor blade having pin fins on the cooled portion on the blade trailing edge. Technology is also disclosed. Japanese Patent Laid-Open No. 5-156901 discloses cooling air from the blade tip side and the root side in the pin fin installation region in order to cool the trailing edge side of the gas turbine stationary blade having pin fins on the cooled portion on the blade trailing edge side. A supply arrangement is disclosed.

特開平６−１３７１０２号公報JP-A-6-137102 特開平５−１５６９０１号公報JP-A-5-156901

しかし、複数の冷却通路の出口を翼の後縁部に半径方向に並べて配置する構成では、翼後縁部以外の被冷却部の冷却が不十分である。冷却通路の本数を多くすると、冷却空気の必要量が多くなってガスタービンの効率が低下してしまい、本数を少なくすると、流路が太くなり流速が遅いためピンフィン設置部分以外の冷却効果が小さくなってしまう。冷却空気の一部を翼先端部へ向かって変向させる仕切板を設ける構成では、翼後縁側先端部へ供給される冷却空気量を多少増加させられる可能性はあるが、冷却空気の大部分は翼根元側から流出するため、翼先端部まで流通せず翼後縁側先端付近を十分に冷却できない。また、ピンフィン設置領域の翼先端側と根元側から冷却空気を供給する構成では、翼先端側と根元側の中間付近の領域に冷却空気が流通しづらく、この領域を充分に冷却できない。つまり、他の被冷却部の冷却効率を犠牲にすることなく、翼後縁部の被冷却部を全体にわたって充分に冷却できるような翼の冷却通路の構成が望まれる。   However, in the configuration in which the outlets of the plurality of cooling passages are arranged in the radial direction at the trailing edge of the blade, cooling of the cooled parts other than the trailing edge of the blade is insufficient. If the number of cooling passages is increased, the required amount of cooling air increases and the efficiency of the gas turbine decreases.If the number of cooling passages is decreased, the flow path becomes thicker and the flow velocity is slow, so the cooling effect other than the pin fin installation part is small. turn into. In the configuration in which the partition plate that changes a part of the cooling air toward the blade tip part is provided, there is a possibility that the amount of cooling air supplied to the blade trailing edge side tip part may be slightly increased. Since it flows out from the blade root side, it does not flow to the blade tip and cannot cool the vicinity of the tip of the blade trailing edge. Further, in the configuration in which the cooling air is supplied from the blade tip side and the root side in the pin fin installation region, it is difficult for the cooling air to flow through a region near the blade tip side and the root side, and this region cannot be cooled sufficiently. That is, it is desirable to have a blade cooling passage configuration that can sufficiently cool the portion to be cooled at the trailing edge of the blade without sacrificing the cooling efficiency of other portions to be cooled.

本発明の目的は、タービン翼を効率的に冷却することで、ガスタービンの効率を低下させることなく、信頼性の高いタービン翼を提供することにある。   An object of the present invention is to provide a highly reliable turbine blade without efficiently reducing the efficiency of the gas turbine by efficiently cooling the turbine blade.

内部に冷却通路を備え、前記冷却通路の翼後縁側の腹側壁面と翼後縁側の背側壁面に凸部または前記両壁面をつなぐ部分を複数有し、前記凸部または翼後縁側の前記両壁面をつなぐ部分は、翼後縁側の前記壁面上の位置により流路抵抗を異ならしめるように配置する。   A cooling passage is provided therein, and has a plurality of portions connecting the convex portion or the both wall surfaces to the abdominal side wall surface on the blade trailing edge side and the back side wall surface on the blade trailing edge side of the cooling passage, and the convex portion or the blade trailing edge side The portion connecting both wall surfaces is arranged so that the flow path resistance varies depending on the position on the wall surface on the blade trailing edge side.

本発明によると、タービン翼を効率的に冷却することで、ガスタービンの効率を向上でき、かつ信頼性に優れたタービン翼を提供できる。   According to the present invention, by efficiently cooling a turbine blade, the efficiency of the gas turbine can be improved, and a turbine blade excellent in reliability can be provided.

本発明は、ガスタービンの動翼や静翼に適用できる。以下実施例には、冷却媒体として圧縮機抽気または吐出空気を利用した例を示す。
（実施例１）
図１，図２，図３を用い、本発明の実施例を詳細に説明する。図１は、本発明の一実施例であるタービン動翼の縦断面図を示す。タービン動翼１の根元側はシャンク部２，先端側は翼部３であり、シャンク部２の内部から翼部３の内部にかけて、中空の冷却通路４，５が設けられている。冷却通路４，５には圧縮機抽気または吐出空気がそれぞれ冷却媒体供給孔１４，１５から供給される。なお、方向１２は翼後縁側方向を示す。 The present invention can be applied to a moving blade and a stationary blade of a gas turbine. In the following examples, an example in which compressor bleed air or discharged air is used as a cooling medium will be shown.
(Example 1)
The embodiment of the present invention will be described in detail with reference to FIGS. FIG. 1 is a longitudinal sectional view of a turbine rotor blade that is an embodiment of the present invention. The turbine rotor blade 1 has a shank portion 2 on the base side and a blade portion 3 on the tip side, and hollow cooling passages 4 and 5 are provided from the inside of the shank portion 2 to the inside of the blade portion 3. Compressor bleed air or discharge air is supplied to the cooling passages 4 and 5 from the cooling medium supply holes 14 and 15, respectively. The direction 12 indicates the blade trailing edge side direction.

冷却通路４は、翼部３において仕切壁６ａ，６ｂにより冷却通路７ａ，７ｂ，７ｃに仕切られ、先端曲部８ａ，下端曲部９ａとともに折れ曲り通路であるサーペンタイン通路を形成する。冷却通路５は、翼部３において仕切壁６ｃ，６ｄにより冷却通路７ｄ，７ｅ，７ｆに仕切られ、先端曲部８ｂ，下端曲部９ｂとともにサーペンタイン通路を形成する。   The cooling passage 4 is partitioned into cooling passages 7a, 7b, and 7c by partition walls 6a and 6b in the wing portion 3, and forms a serpentine passage that is a bent passage together with the leading end bending portion 8a and the lower end bending portion 9a. The cooling passage 5 is partitioned into cooling passages 7d, 7e, and 7f by partition walls 6c and 6d in the wing portion 3, and forms a serpentine passage together with the leading end bent portion 8b and the lower end bent portion 9b.

冷却通路７ｆより後縁側には、冷却通路５を流れた冷却空気が翼外部に流出するように吹き出し部１３が設けられている。吹き出し部１３には、ピンフィン１６が翼と一体構造で複数設けられている。本実施例ではピンフィン１６の形状として略円筒状のものを用いた場合について説明する。   A blowing portion 13 is provided on the rear edge side of the cooling passage 7f so that the cooling air flowing through the cooling passage 5 flows out of the blades. A plurality of pin fins 16 are provided in the blowing portion 13 in an integral structure with the wing. In this embodiment, the case where a substantially cylindrical shape is used as the shape of the pin fin 16 will be described.

タービン動翼１には、翼を保持するロータディスク（図示省略）等から冷却空気が供給孔１４，１５に供給され、冷却通路４，５を通過する過程で動翼１を内部から冷却する。供給孔１４から流入した冷却空気の大部分は翼先端に設けた吹き出し孔１１から、供給孔
１５から流入した冷却空気の大部分は翼後縁の吹き出し部１３から外部に流出する。 Cooling air is supplied to the turbine blades 1 from a rotor disk (not shown) that holds the blades to the supply holes 14 and 15, and the blades 1 are cooled from the inside in the process of passing through the cooling passages 4 and 5. Most of the cooling air flowing in from the supply hole 14 flows out from the blowing hole 11 provided at the tip of the blade, and most of the cooling air flowing in from the supply hole 15 flows out from the blowing part 13 at the trailing edge of the blade.

本実施例のタービン動翼１は、翼部３を冷却する冷却通路として、翼根元側から翼先端側へ向かう流路と翼先端側から翼根元側へ向かう流路をそれぞれ３本ずつ、計６本有している。このように流路を多く有することで、流路１本あたりの断面積が小さいために、流路が少ない場合と比べて冷却空気の流速が早く、翼部３の冷却効果が大きい。本実施例で示した翼と同様の用途に用いられるタービン動翼では、充分な冷却効果を得るために、翼部３を冷却する翼根元側から翼先端側へ向かう流路と翼先端側から翼根元側へ向かう流路は合計４本以上設けることが望ましい。   The turbine rotor blade 1 according to the present embodiment has three flow paths from the blade root side to the blade tip side and three flow paths from the blade tip side to the blade root side as cooling passages for cooling the blade portion 3. Has six. By having many flow paths in this way, since the cross-sectional area per flow path is small, the flow velocity of the cooling air is faster than when there are few flow paths, and the cooling effect of the blade portion 3 is great. In the turbine blade used for the same application as the blade shown in the present embodiment, in order to obtain a sufficient cooling effect, a flow path from the blade root side to the blade tip side for cooling the blade portion 3 and the blade tip side are used. It is desirable to provide a total of four or more channels toward the blade root side.

本実施例のタービン動翼１はまた、翼部３を冷却する冷却通路を、供給孔１４から冷却空気を供給する冷却通路４，供給孔１５から冷却空気を供給する冷却通路５の合計２系統有している。各系統の供給孔にはそれぞれ充分な流量の冷却空気を供給しなければならないため、冷却通路の系統数を多くすると冷却空気の必要供給量は多くなる。冷却空気には、作動媒体の一部である圧縮機の抽気または突出空気が用いられることが多い。そのため、冷却空気必要量の増加は燃焼器に供給される燃焼空気量の減少につながり、ガスタービンの効率を低下させる。   The turbine rotor blade 1 of the present embodiment also has two cooling passages for cooling the blade portion 3, that is, a cooling passage 4 for supplying cooling air from the supply hole 14 and a cooling passage 5 for supplying cooling air from the supply hole 15. Have. Since a sufficient flow rate of cooling air must be supplied to the supply holes of each system, the required supply amount of cooling air increases when the number of systems of cooling passages is increased. As the cooling air, compressor bleed air or protruding air that is a part of the working medium is often used. Therefore, an increase in the required amount of cooling air leads to a decrease in the amount of combustion air supplied to the combustor, thereby reducing the efficiency of the gas turbine.

本実施例では冷却系統の総数を２本と少数に抑えているため冷却空気の供給量を少なくでき、ガスタービンの効率を向上できる。ガスタービンの効率を考慮すると、本実施例で示した翼と同様の用途に用いられるタービン動翼では、冷却系統の総数は合計２本以内とすることが望ましい。   In this embodiment, since the total number of cooling systems is limited to two, the amount of cooling air supplied can be reduced and the efficiency of the gas turbine can be improved. In consideration of the efficiency of the gas turbine, it is desirable that the total number of cooling systems in the turbine rotor blade used for the same application as the blade shown in the present embodiment is not more than two.

図２は、図１に示した動翼のＡ−Ａに沿う断面図を示す。方向２２は翼後縁側方向を示す。ピンフィン１６は、翼腹側壁２８から翼背側壁２９に伸びる略円柱形状の構成要素である。   FIG. 2 shows a cross-sectional view along AA of the rotor blade shown in FIG. A direction 22 indicates a blade trailing edge side direction. The pin fin 16 is a substantially cylindrical component that extends from the blade abdominal side wall 28 to the blade back side wall 29.

図３は図１に示したタービン動翼１の冷却通路７ｆ，吹き出し部１３の拡大図を示す。流れ１８，１９ａ−１９ｅは冷却空気の流れを示す。冷却通路７ｆでは流れ１８として冷却空気の一部は翼後縁側先端まで導かれる。冷却空気は流れ１８として翼先端まで導かれる過程において、流れ１８の一部は流れ１９ａ−１９ｄとして徐々に吹き出し部１３から外部へ流出する。   FIG. 3 shows an enlarged view of the cooling passage 7f and the blowing portion 13 of the turbine rotor blade 1 shown in FIG. Streams 18, 19a-19e indicate the flow of cooling air. In the cooling passage 7f, a part of the cooling air is guided to the tip of the blade trailing edge as the flow 18. In the process in which the cooling air is guided to the blade tip as the flow 18, a part of the flow 18 gradually flows out from the blowing portion 13 as the flow 19a to 19d.

冷却通路７ｆの領域までピンフィンが設けられている（冷却通路７ｆが設けられていない）タービン翼では、吹き出し部１３から吹き出す冷却空気の大部分が、翼根元側の流れである１９ａ−１９ｃとして流出してしまい、翼先端側まで到達する冷却空気は僅かである。そのため、翼後縁側先端付近の高温部分を充分に冷却することができず、タービン翼の信頼性が低下する。   In a turbine blade in which pin fins are provided up to the region of the cooling passage 7f (no cooling passage 7f is provided), most of the cooling air blown out from the blowing portion 13 flows out as a blade root side flow 19a-19c. As a result, the amount of cooling air reaching the blade tip side is small. For this reason, the high-temperature portion near the tip of the blade trailing edge cannot be sufficiently cooled, and the reliability of the turbine blade decreases.

これに対し、本実施例では、ピンフィンが設けられた吹き出し部１３の翼前縁側に冷却通路７ｆを設けている。このような構成とすることにより、冷却通路７ｆが設けられていないガスタービン翼と比べ、より多くの空気をより確実に翼後縁側先端部まで導くことができる。つまり、冷却空気の流れ１８の下流側および冷却空気の流れ１９ｄ，１９ｅの流量が増すことで、この流れにより冷却される被冷却部の冷却効果が大きくなり、被冷却部のメタル温度が低下してタービン翼の信頼性が向上する。また、冷却空気が届きにくい翼後縁側先端部まで効率良く冷却できるようになるため、冷却空気の供給量を削減することができ、さらにガスタービンの効率も向上する。   In contrast, in this embodiment, the cooling passage 7f is provided on the blade leading edge side of the blowing portion 13 provided with the pin fins. By adopting such a configuration, it is possible to more reliably guide more air to the blade trailing edge side tip than in the gas turbine blade not provided with the cooling passage 7f. That is, the downstream side of the cooling air flow 18 and the flow rates of the cooling air flows 19d and 19e increase, so that the cooling effect of the cooled part cooled by this flow increases, and the metal temperature of the cooled part decreases. This improves the reliability of the turbine blade. In addition, since the cooling air can be efficiently cooled down to the tip portion on the blade trailing edge side where it is difficult for the cooling air to reach, the supply amount of the cooling air can be reduced and the efficiency of the gas turbine is also improved.

また、本実施例では、翼根元部では大きなピンフィン１６ｂを、翼先端部では小さなピンフィン１６ａを設けた。これにより、吹き出し部１３の翼根元側領域である大きなピンフィン１６ｂが設けられた領域から外部に吹き出す際の冷却媒体の流路抵抗を、小さなピンフィン１６ａが設けられた領域から外部に吹き出す際の冷却媒体の流路抵抗よりも大きくすることができる。つまり、流れ１９ａ−１９ｃの流量を少なく、流れ１９ｄ，１９ｅの流量を多くすることができる。このような構成とすることにより、ピンフィン１６の大きさが全て同程度であるタービン翼と比べ、より多くの空気をより確実に翼後縁側先端部まで導くことができるようになる。つまり、冷却通路７ｆを設けたことによる効果と同様、タービン翼信頼性の向上，ガスタービンの効率の向上という２つの効果を得ることができる。   In this embodiment, a large pin fin 16b is provided at the blade root and a small pin fin 16a is provided at the blade tip. As a result, the flow path resistance of the cooling medium when blowing out from the region where the large pin fin 16b, which is the blade root side region of the blowing portion 13, is blown out is cooled when blowing out from the region where the small pin fin 16a is blown out. It can be larger than the flow path resistance of the medium. That is, the flow rates of the streams 19a-19c can be reduced, and the flow rates of the streams 19d and 19e can be increased. By adopting such a configuration, more air can be more reliably guided to the tip of the blade trailing edge side as compared with a turbine blade having the same size of the pin fins 16. That is, the two effects of improving the turbine blade reliability and improving the efficiency of the gas turbine can be obtained in the same manner as the effect of providing the cooling passage 7f.

なお、本実施例で流路抵抗とは、冷却空気が冷却通路７ｆから、ピンフィン１６が設けられた領域である吹き出し部１３を通過して外部に流出する際の、吹き出し部１３における冷却空気の流れにくさを意味する。   In this embodiment, the flow path resistance means that the cooling air flows from the cooling passage 7f to the outside through the blowing portion 13 which is an area where the pin fins 16 are provided and flows out to the outside. Means difficulty in flowing.

また、本実施例でピンフィン１６が大きいとは、ピンフィン１６が設けられている面
（翼腹側壁２８または翼背側壁２９）に略平行な面の断面積が大きいことを意味する。ピンフィン１６が小さいとは、同様にピンフィン１６の断面積が小さいことを意味する。大きなピンフィンと小さなピンフィンの断面積比は、矢印１８の上流側と下流側に必要とされる冷却能力によって、最適に設計されるべきであるが、１.４ 〜４倍程度を選ぶことが一般的である。本実施例のようにピンフィンの断面形状が円形である場合には、円の直径比で１.２〜２倍程度に相当する。 Further, in the present embodiment, the fact that the pin fin 16 is large means that the cross-sectional area of a surface substantially parallel to the surface (the blade belly side wall 28 or the blade back side wall 29) on which the pin fin 16 is provided is large. The small pin fin 16 means that the cross-sectional area of the pin fin 16 is small similarly. The cross-sectional area ratio between the large pin fin and the small pin fin should be optimally designed according to the cooling capacity required on the upstream side and downstream side of the arrow 18, but it is generally selected to be about 1.4 to 4 times. Is. When the cross-sectional shape of the pin fin is circular as in this embodiment, the diameter ratio of the circle corresponds to about 1.2 to 2 times.

本実施例では、翼後縁側先端部の冷却能力を向上させる観点から、翼先端側に設置するピンフィン１６ａを小さなものとしたが、翼根元側等他の部分の冷却能力を向上させたい場合には、適当な部分のピンフィン１６を大きなものとしたり小さなものとしたりすることも考えられる。
（実施例２）
図４，図５を用い、本発明の他の実施例を説明する。図４は、本発明の一実施例であるタービン動翼１の縦断面図を示す。図５は図４に示したタービン動翼１の冷却通路７ｆ，吹き出し部１３の拡大図を示す。 In this embodiment, from the viewpoint of improving the cooling capability of the blade trailing edge side tip, the pin fin 16a installed on the blade tip side is made small, but when the cooling capability of other parts such as the blade root side is desired to be improved. It is also conceivable to make the pin fin 16 of an appropriate part large or small.
(Example 2)
Another embodiment of the present invention will be described with reference to FIGS. FIG. 4 is a longitudinal sectional view of the turbine rotor blade 1 which is an embodiment of the present invention. FIG. 5 shows an enlarged view of the cooling passage 7f and the blowing portion 13 of the turbine rotor blade 1 shown in FIG.

本実施例は、吹き出し部１３における単位面積あたりに設けられたピンフィン１６の本数、すなわちピンフィン１６の設置密度を異ならしめた例を示す。具体的には、吹き出し部１３における流れ１９ｄ，１９ｅの流路に相当する領域(翼先端側)のピンフィン１６ｃの設置密度を、吹き出し部１３における流れ１９ａ−１９ｃの流路に相当する領域（翼根元側）のピンフィン１６ｄの設置密度よりも低くした。   The present embodiment shows an example in which the number of pin fins 16 provided per unit area in the blowing portion 13, that is, the installation density of the pin fins 16 is made different. Specifically, the installation density of the pin fins 16c in the region (blade tip side) corresponding to the flow paths 19d and 19e in the blowing portion 13 is set to the region corresponding to the flow 19a-19c flow passage (blade in the blowing portion 13). It is lower than the installation density of the pin fins 16d on the base side.

このような構成とすることで、吹き出し部１３の翼根元側領域であるピンフィン１６ｃが疎に設けられた領域から外部に吹き出す際の冷却媒体の流路抵抗を、ピンフィン１６ｄが設けられた領域から外部に吹き出す際の冷却媒体の流路抵抗よりも小さくすることができる。つまり、流れ１９ａ−１９ｃの流量を少なく、流れ１９ｄ，１９ｅの流量を多くすることができる。このような構成とすることにより、ピンフィン１６の設置密度が全て同程度であるタービン翼と比べ、より多くの空気をより確実に翼後縁側先端部まで導くことができるようになる。つまり、実施例１のタービン翼で得られる効果と同様、タービン翼信頼性の向上，ガスタービンの効率の向上という２つの効果を得ることができる。   By adopting such a configuration, the flow path resistance of the cooling medium when the pin fins 16c, which are the blade root side regions of the blowing portion 13, are blown out from the region where the pin fins 16d are blown out from the region where the pin fins 16d are provided. It can be made smaller than the flow path resistance of the cooling medium when blowing out to the outside. That is, the flow rates of the streams 19a-19c can be reduced, and the flow rates of the streams 19d and 19e can be increased. By adopting such a configuration, more air can be more reliably guided to the tip of the blade trailing edge side as compared with the turbine blade in which the installation density of the pin fins 16 is almost the same. That is, the two effects of improving the turbine blade reliability and improving the efficiency of the gas turbine can be obtained in the same manner as the effect obtained by the turbine blade of the first embodiment.

本実施例ではピンフィン１６を配置するピッチを変化させている。このピッチ比は、翼先端側と翼根元側で必要とされる冷却能力によって最適に設計されるべきであるが、一例を挙げると、１.１〜２倍程度を選ぶことが一般的である。   In this embodiment, the pitch for arranging the pin fins 16 is changed. This pitch ratio should be optimally designed according to the cooling capacity required on the blade tip side and blade root side. To give an example, it is common to select about 1.1 to 2 times. .

本実施例では、翼後縁側先端部の冷却能力を向上させる観点から、翼先端側に設置するピンフィン１６ｃの設置密度を小さくしたが、翼根元側等他の部分の冷却能力を向上させたい場合には、適当な部分のピンフィン１６の密度を大きくしたり小さくしたりすることも考えられる。
（実施例３）
図６，図７を用い、本発明の他の実施例を説明する。図６は、本発明の一実施例であるタービン動翼の縦断面図を示す。図７は図６に示したタービン動翼１の冷却通路７ｆ，吹き出し部１３の拡大図を示す。 In the present embodiment, the installation density of the pin fins 16c to be installed on the blade tip side is reduced from the viewpoint of improving the cooling capability of the blade trailing edge side tip portion, but it is desired to improve the cooling capacity of other portions such as the blade root side. It is also conceivable to increase or decrease the density of the pin fins 16 at appropriate portions.
Example 3
Another embodiment of the present invention will be described with reference to FIGS. FIG. 6 is a longitudinal sectional view of a turbine rotor blade that is an embodiment of the present invention. FIG. 7 shows an enlarged view of the cooling passage 7f and the blowing portion 13 of the turbine rotor blade 1 shown in FIG.

本実施例は、吹き出し部１３におけるピンフィン１６の設置密度を異ならしめ、かつ、ピンフィンの大きさに差をつけた例を示す。具体的には、吹き出し部１３において、翼先端側には小さなピンフィン１６ｅを、翼根元側には大きなピンフィン１６ｆを設け、小さなピンフィン１６ｅの設置密度を大きなピンフィン１６ｆの設置密度よりも小さくなるように構成した。このような構成とすることで、実施例１，２で述べた通り、タービン翼信頼性の向上，ガスタービンの効率の向上という効果を奏する。本実施例のタービン翼では、ピンフィンの大きさに差をつけた上で設置密度にも差を設けているため、実施例１や実施例２のタービン翼よりも大きな効果を得ることができる。   The present embodiment shows an example in which the installation density of the pin fins 16 in the blowing portion 13 is made different and the size of the pin fins is made different. Specifically, in the blowing portion 13, a small pin fin 16e is provided on the blade tip side and a large pin fin 16f is provided on the blade root side so that the installation density of the small pin fins 16e is smaller than the installation density of the large pin fins 16f. Configured. With such a configuration, as described in the first and second embodiments, the effects of improved turbine blade reliability and improved efficiency of the gas turbine are achieved. In the turbine blade of the present embodiment, a difference is also provided in the installation density after making a difference in the size of the pin fins, so that a larger effect can be obtained than in the turbine blades of the first and second embodiments.

ここまではタービン翼の例として、ガスタービン動翼１を例にとって説明した。しかし本発明の適用対象は動翼に限られたものではない。次に実施例４を用い、本発明をガスタービン静翼に適用した例を示す。
（実施例４）
図８を用いて本発明の実施例４を説明する。図８は、本発明の一実施例であるタービン静翼の縦断面図を示す。 Up to this point, the gas turbine rotor blade 1 has been described as an example of the turbine blade. However, the application target of the present invention is not limited to the moving blade. Next, the example which applied Example 4 to the gas turbine stationary blade using Example 4 is shown.
Example 4
Embodiment 4 of the present invention will be described with reference to FIG. FIG. 8 shows a longitudinal sectional view of a turbine vane that is an embodiment of the present invention.

タービン静翼６１は、翼部６３の内部に冷却通路６７ａ，６７ｂを有する。方向７２は翼後縁側方向、８５は翼根元側方向、方向８６は翼先端側方向を示す。冷却通路６７ａ，６７ｂは仕切壁６６ａで仕切られている。冷却通路６７ｂの翼後縁側の吹き出し部７３には、ピンフィン７６が複数設けられている。流れ６４，６５，７９ａ−７９ｅは冷却空気の流れを示す。   The turbine stationary blade 61 has cooling passages 67 a and 67 b inside the blade portion 63. A direction 72 indicates a blade trailing edge side direction, 85 indicates a blade root side direction, and a direction 86 indicates a blade tip side direction. The cooling passages 67a and 67b are partitioned by a partition wall 66a. A plurality of pin fins 76 are provided in the blowing portion 73 on the blade trailing edge side of the cooling passage 67b. Flows 64, 65, 79a-79e indicate the flow of cooling air.

冷却空気供給孔７４から供給された冷却空気は、冷却通路６７ａを翼先端側方向に流れ、吹き出し孔８３から流出する。冷却空気供給孔７５から供給された冷却空気は、冷却通路６７ｂを翼先端側方向に流れる。この翼先端側方向に向かう流れ６５は流れ７９ａ，
７９ｂ，７９ｃ，７９ｄ，７９ｅに分岐して吹き出し部７３を冷却し、翼外部へ流出する。 The cooling air supplied from the cooling air supply hole 74 flows in the cooling passage 67 a in the direction of the blade tip side and flows out from the blowing hole 83. The cooling air supplied from the cooling air supply hole 75 flows through the cooling passage 67b in the blade tip side direction. This flow 65 toward the blade tip side is a flow 79a,
It branches into 79b, 79c, 79d, 79e, cools the blowing part 73, and flows out of the blade.

本実施例のタービン静翼６１は、翼中央付近領域８７付近でメタル温度が特に高くなる。したがって、タービン静翼６１の後縁側領域である吹き出し部７３においては、領域
８７付近に冷却空気を多く供給することが望まれる。 The turbine stator blade 61 of the present embodiment has a particularly high metal temperature in the vicinity of the blade center vicinity region 87. Therefore, it is desirable to supply a large amount of cooling air in the vicinity of the region 87 in the blowing portion 73 that is the region on the trailing edge side of the turbine stationary blade 61.

本実施例では、吹き出し部７３において、領域８７に小さなピンフィン７６ａを設置し、それ以外の領域には大きなピンフィン７６ｂを設置した。このように構成することで、吹き出し部７３における、領域８７から流れ７９ｃとして流出する冷却空気の流路抵抗を他の領域から流れ７９ａ，７９ｂ，７９ｄ，７９ｅとして流出する冷却空気の流路抵抗より高くでき、充分な量の冷却空気を供給することができる。   In the present embodiment, in the blowing portion 73, the small pin fins 76a are installed in the region 87, and the large pin fins 76b are installed in the other regions. With this configuration, the flow path resistance of the cooling air that flows out from the region 87 as the flow 79c in the blowing unit 73 is greater than the flow path resistance of the cooling air that flows out as flows 79a, 79b, 79d, and 79e from other regions. The cooling air can be increased and a sufficient amount of cooling air can be supplied.

このように、翼後縁側で特にメタル温度が高くなる領域８７の冷却能力を高めることで、タービン静翼６１の冷却効率を向上させることができ、タービン翼信頼性の向上，ガスタービンの効率の向上という２つの効果を得ることができる。   Thus, by increasing the cooling capacity of the region 87 where the metal temperature is particularly high on the blade trailing edge side, the cooling efficiency of the turbine stationary blade 61 can be improved, and the turbine blade reliability can be improved and the efficiency of the gas turbine can be improved. Two effects of improvement can be obtained.

なお、本実施例ではピンフィン７６の大きさを異ならしめることで、吹き出し部７３から流出する冷却空気の流路抵抗を異ならしめる例を示したが、ピンフィン７６の配置密度を異ならしめる等、他の方法で流路抵抗を異ならしめることも考えられる。   In this embodiment, the example in which the flow path resistance of the cooling air flowing out from the blowing portion 73 is made different by changing the size of the pin fins 76 is shown, but other arrangements such as making the arrangement density of the pin fins 76 different are shown. It is also conceivable to make the flow path resistance different by a method.

以上説明した各実施例では、ピンフィン１６の大きさや設置密度に差を持たせることで、翼後縁側先端部等の冷却空気が届きにくい領域や特に冷却が必要とされる領域への冷却空気到達を容易にすることでタービン翼を効果的に冷却する例を説明した。この効果的な冷却は、ピンフィンが設置された吹き出し部に、吹き出し部から流出する冷却空気の流路抵抗が高い領域と低い領域を設けることにより達成される。すなわち、吹き出し部における冷却空気の流出位置によって流路抵抗を異ならしめることができる構成であれば、タービン翼信頼性の向上，ガスタービンの効率の向上という２つの効果を得ることができる。   In each of the embodiments described above, by providing a difference in the size and installation density of the pin fins 16, the cooling air reaches an area where the cooling air is difficult to reach such as the tip of the blade trailing edge side or an area where cooling is particularly necessary. An example in which turbine blades are effectively cooled by facilitating the above has been described. This effective cooling is achieved by providing regions where the flow resistance of cooling air flowing out from the blowing portion is high and low in the blowing portion where the pin fins are installed. That is, if the flow path resistance can be made different depending on the cooling air outflow position in the blow-out portion, two effects of improving the turbine blade reliability and improving the efficiency of the gas turbine can be obtained.

ピンフィン１６の大きさや設置密度に差を持たせるということは、吹き出し部１３の翼腹側壁面または翼背側壁面上における、単位面積あたりのピンフィン設置面積に差を持たせることである。このような構成を備えたタービン翼であれば上記２つの効果を得ることができる。   Giving a difference in the size and installation density of the pin fins 16 means making a difference in the pin fin installation area per unit area on the blade belly side wall surface or blade back side wall surface of the blowing portion 13. The above two effects can be obtained with a turbine blade having such a configuration.

また、各実施例ではピンフィン１６として、翼と一体構造で設けられた略円筒形状のものを用いた例を説明したが、これに限られたものではない。ピンフィン１６と同様の役割を果たす構成として、例えば、翼腹側壁２８から翼背側壁２９にのびる三角柱状のものを用いてもよいし、翼腹側壁２８や翼背側壁２９に凸部を設けてもよい。要は、吹き出し部１３から外部に流出する冷却空気の流路抵抗を増すことができるものを用いればよい。   In each embodiment, the pin fin 16 is described as having an approximately cylindrical shape provided integrally with the wing. However, the present invention is not limited to this. As a configuration that plays the same role as the pin fin 16, for example, a triangular prism shape extending from the blade abdominal side wall 28 to the blade back side wall 29 may be used, or a protrusion is provided on the blade abdominal side wall 28 or the blade back side wall 29. Also good. In short, what can increase the flow path resistance of the cooling air flowing out from the blowing portion 13 to the outside may be used.

本発明の実施例１であるタービン動翼の縦断面図を示す。1 is a longitudinal sectional view of a turbine rotor blade that is Embodiment 1 of the present invention. 図１に示したタービン動翼１のＡ−Ａに沿う断面図を示す。Sectional drawing which follows AA of the turbine rotor blade 1 shown in FIG. 1 is shown. 図１に示したタービン動翼１の冷却通路７ｆ，吹き出し部１３の拡大図を示す。The enlarged view of the cooling channel | path 7f and the blowing part 13 of the turbine rotor blade 1 shown in FIG. 1 is shown. 本発明の実施例２であるタービン動翼の縦断面図を示す。The longitudinal cross-sectional view of the turbine rotor blade which is Example 2 of this invention is shown. 図４に示したタービン動翼１の冷却通路７ｆ，吹き出し部１３の拡大図を示す。The enlarged view of the cooling channel | path 7f and the blowing part 13 of the turbine rotor blade 1 shown in FIG. 4 is shown. 本発明の実施例３であるタービン動翼の縦断面図を示す。The longitudinal cross-sectional view of the turbine rotor blade which is Example 3 of this invention is shown. 図６に示したタービン動翼１の冷却通路７ｆ，吹き出し部１３の拡大図を示す。The enlarged view of the cooling channel | path 7f and the blowing part 13 of the turbine rotor blade 1 shown in FIG. 6 is shown. 本発明の実施例４であるタービン静翼の縦断面図を示す。The longitudinal cross-sectional view of the turbine stationary blade which is Example 4 of this invention is shown.

符号の説明Explanation of symbols

１…動翼、２…シャンク部、３，６３…翼部、４，５，７ａ，７ｂ，７ｃ，７ｄ，７ｅ，７ｆ，６７ａ，６７ｂ，…冷却通路、６ａ，６ｂ，６ｃ，６ｄ，６６…仕切壁、８ａ，８ｂ…先端曲部、９ａ，９ｂ…下端曲部、１１，８３…吹き出し孔、１２，２２…翼後縁側方向、１３，７３…吹き出し部、１４，１５，７４，７５…供給孔、１６，１６ａ，
１６ｂ，１６ｃ，１６ｄ，７６…ピンフィン、１８，１９ａ，１９ｂ，１９ｄ，１９ｅ，６４，６５，７９ａ，７９ｂ，７９ｄ，７９ｅ…流れ、２８…翼腹側壁、２９…翼背側壁、６１…静翼、８５，８６…方向、８７…領域。

DESCRIPTION OF SYMBOLS 1 ... Moving blade, 2 ... Shank part, 3,63 ... Blade part, 4, 5, 7a, 7b, 7c, 7d, 7e, 7f, 67a, 67b, ... Cooling passage, 6a, 6b, 6c, 6d, 66 ... partition wall, 8a, 8b ... tip curved part, 9a, 9b ... bottom end curved part, 11, 83 ... blowing hole, 12, 22 ... wing trailing edge side direction, 13, 73 ... blowing part, 14, 15, 74, 75 ... Supply hole 16, 16a,
16b, 16c, 16d, 76 ... pin fins, 18, 19a, 19b, 19d, 19e, 64, 65, 79a, 79b, 79d, 79e ... flow, 28 ... blade side wall, 29 ... blade back side wall, 61 ... stationary blade , 85, 86... Direction, 87.

Claims (8)

内部に冷却通路を備え、前記冷却通路の翼後縁側の腹側壁面と翼後縁側の背側壁面に凸部または前記両壁面をつなぐ部分を複数有するタービン翼において、
前記凸部または翼後縁側の前記両壁面をつなぐ部分は、翼後縁側の前記壁面上の位置により流路抵抗を異ならしめるように配置されていることを特徴とするタービン翼。 In a turbine blade comprising a cooling passage inside, and having a plurality of portions connecting the convex portion or the both wall surfaces to the abdominal side wall surface on the blade trailing edge side of the cooling passage and the back side wall surface on the blade trailing edge side,
The turbine blade according to claim 1, wherein a portion connecting the two wall surfaces on the convex portion or the blade trailing edge side is arranged so as to make a flow path resistance different depending on a position on the wall surface on the blade trailing edge side. 請求項１に記載のタービン翼において、
前記各壁面における、単位面積あたりの、前記凸部または前記両壁面をつなぐ部分の設置面積が翼後縁側の前記壁面上の位置により異なることで、翼後縁側の前記壁面上の位置により流路抵抗が異なることを特徴とするタービン翼。 The turbine blade according to claim 1,
In each of the wall surfaces, the installation area of the convex portion or the portion connecting the both wall surfaces per unit area varies depending on the position on the wall surface on the blade trailing edge side, and the flow path depends on the position on the wall surface on the blade trailing edge side. Turbine blades characterized by different resistances. 請求項１に記載のタービン翼において、
前記凸部または前記両壁面をつなぐ部分は、翼後縁側の前記両壁面をつなぐ略円筒状部材であり、前記略円筒状部材の設置密度が翼後縁側の前記壁面上の位置により異なることを特徴とするタービン翼。 The turbine blade according to claim 1,
The convex portion or the portion connecting the both wall surfaces is a substantially cylindrical member that connects the both wall surfaces on the blade trailing edge side, and the installation density of the substantially cylindrical member varies depending on the position on the wall surface on the blade trailing edge side. Characteristic turbine blade. 請求項１に記載のタービン翼において、
前記凸部または前記両壁面をつなぐ部分は、翼後縁側の前記両壁面をつなぐ略円筒状部材であり、前記略円筒状部材の、翼後縁側における前記壁面に略平行な面における断面積が、翼後縁側の前記壁面上の位置によって異なることを特徴とするタービン翼。 The turbine blade according to claim 1,
The convex portion or the portion connecting the both wall surfaces is a substantially cylindrical member that connects the both wall surfaces on the blade trailing edge side, and the cross-sectional area of the substantially cylindrical member in a plane substantially parallel to the wall surface on the blade trailing edge side is A turbine blade characterized in that it varies depending on the position on the wall surface on the blade trailing edge side. 請求項１に記載のタービン翼において、
前記流路抵抗は、翼根元側よりも翼先端側の方が低いことを特徴とするタービン翼。 The turbine blade according to claim 1,
The turbine blade according to claim 1, wherein the flow path resistance is lower on a blade tip side than on a blade root side. 内部に冷却通路を備え、前記冷却通路の翼後縁側の腹側壁面と翼後縁側の背側壁面に凸部または前記両壁面をつなぐ部分を複数有するタービン翼において、
前記凸部または前記両壁面をつなぐ部分を複数有する領域の翼前縁側に隣接する、冷却媒体を翼根元方向から翼先端部まで導く冷却通路を備え、前記凸部または前記両壁面をつなぐ部分は、翼後縁側の前記壁面上の位置によって流路抵抗を異ならしめるように配置されていることを特徴とするタービン翼。 In a turbine blade comprising a cooling passage inside, and having a plurality of portions connecting the convex portion or the both wall surfaces to the abdominal side wall surface on the blade trailing edge side of the cooling passage and the back side wall surface on the blade trailing edge side,
A cooling passage for guiding the cooling medium from the blade root direction to the blade tip is adjacent to the blade leading edge side of the region having a plurality of portions connecting the convex portions or the both wall surfaces, and the portion connecting the convex portions or the both wall surfaces is A turbine blade, characterized in that the flow passage resistance is varied depending on the position on the wall surface on the blade trailing edge side. 内部に、入口が翼根元部、出口の少なくとも一つが翼後縁部である冷却通路を備え、前記冷却通路の翼後縁側の腹側壁面と翼後縁側の背側壁面に凸部または前記両壁面をつなぐ部分を複数有するタービン翼において、
翼根元方向から翼先端方向に向かって冷却媒体を流通させる通路と、翼先端方向から翼根元方向に向かって前記冷却媒体を流通させる通路を組み合わせた、翼部を根元側から先端側にわたって冷却するための冷却通路を備え、翼部を根元側から先端側にわたって冷却した前記冷却媒体の一部または全部を翼根元方向から翼先端部まで導く、前記凸部または前記両壁面をつなぐ部分を複数有する領域の翼前縁側に隣接する冷却通路を備え、前記凸部または前記両壁面をつなぐ部分は、流路抵抗を異ならしめるように配置されていることを特徴とするタービン翼。 Inside, a cooling passage having an inlet as a blade root portion and at least one outlet as a blade trailing edge is provided, and a convex portion or both of the protrusions are formed on the abdominal side wall surface on the blade trailing edge side and the back side wall surface on the blade trailing edge side of the cooling passage. In a turbine blade having a plurality of portions connecting wall surfaces,
The wing portion is cooled from the root side to the tip side by combining a passage through which the cooling medium flows from the blade root direction toward the blade tip direction and a passage through which the cooling medium flows from the blade tip direction toward the blade root direction. And a plurality of portions connecting the convex portions or the two wall surfaces for guiding a part or all of the cooling medium, which has cooled the blade portion from the root side to the tip side, from the blade root direction to the blade tip portion. A turbine blade comprising a cooling passage adjacent to a blade leading edge side of a region, wherein a portion connecting the convex portion or the two wall surfaces is arranged so as to have different flow path resistances. 内部に冷却通路を備え、前記冷却通路の翼後縁側の腹側壁面と翼後縁側の背側壁面に凸部または前記両壁面をつなぐ部分を複数有するタービン翼の冷却方法において、
前記凸部または翼後縁側の前記両壁面をつなぐ部分を、翼後縁側の前記壁面上の位置により流路抵抗を異ならしめるように配置し、前記冷却通路に冷却媒体を流通させることを特徴とするタービン翼の冷却方法。

In the cooling method of a turbine blade, comprising a cooling passage inside, and having a plurality of portions connecting the convex portion or the both wall surfaces to the abdominal side wall surface on the blade trailing edge side of the cooling passage and the back side wall surface on the blade trailing edge side,
The convex portion or a portion connecting the wall surfaces on the blade trailing edge side is arranged so that the flow resistance varies depending on the position on the wall surface on the blade trailing edge side, and the cooling medium is circulated in the cooling passage. Method for cooling turbine blades.

JP2006122884A 2006-04-27 2006-04-27 Turbine blade having cooling passage inside thereof Pending JP2007292006A (en)

Priority Applications (2)

Application Number Priority Date Filing Date Title
JP2006122884A JP2007292006A (en) 2006-04-27 2006-04-27 Turbine blade having cooling passage inside thereof
EP07008241A EP1849960A3 (en) 2006-04-27 2007-04-23 Turbine blade having internal cooling passage

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
JP2006122884A JP2007292006A (en) 2006-04-27 2006-04-27 Turbine blade having cooling passage inside thereof

Publications (1)

Publication Number Publication Date
JP2007292006A true JP2007292006A (en) 2007-11-08

Family

ID=38007269

Family Applications (1)

Application Number Title Priority Date Filing Date
JP2006122884A Pending JP2007292006A (en) 2006-04-27 2006-04-27 Turbine blade having cooling passage inside thereof

Country Status (2)

Country Link
EP (1) EP1849960A3 (en)
JP (1) JP2007292006A (en)

Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JP2011515618A (en) * 2008-03-28 2011-05-19 アルストム テクノロジー リミテッド Gas turbine stationary blade and gas turbine equipped with such a stationary blade
US8465255B2 (en) 2008-05-14 2013-06-18 Mitsubishi Heavy Industries, Ltd. Gas turbine blade and gas turbine having the same
JP2014047787A (en) * 2012-08-31 2014-03-17 General Electric Co <Ge> Airfoil and method for manufacturing airfoil

Families Citing this family (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2009109462A1 (en) * 2008-03-07 2009-09-11 Alstom Technology Ltd Vane for a gas turbine
EP2602439A1 (en) * 2011-11-21 2013-06-12 Siemens Aktiengesellschaft Coolable hot gas component for a gas turbine
US20150204237A1 (en) * 2014-01-17 2015-07-23 General Electric Company Turbine blade and method for enhancing life of the turbine blade
EP3099901B1 (en) * 2014-01-30 2019-10-09 United Technologies Corporation Turbine blade with airfoil having a trailing edge cooling pedestal configuration
US10208604B2 (en) 2016-04-27 2019-02-19 United Technologies Corporation Cooling features with three dimensional chevron geometry
US10605096B2 (en) * 2016-05-12 2020-03-31 General Electric Company Flared central cavity aft of airfoil leading edge
KR102114681B1 (en) * 2018-09-21 2020-05-25 두산중공업 주식회사 Turbine blade having pin-fin array

Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4278400A (en) * 1978-09-05 1981-07-14 United Technologies Corporation Coolable rotor blade
JPH05156901A (en) * 1991-12-02 1993-06-22 Hitachi Ltd Gas turbine cooling stationary blade
JPH06137102A (en) * 1992-10-26 1994-05-17 Mitsubishi Heavy Ind Ltd Hollow moving blade of gas turbine
US5538394A (en) * 1993-12-28 1996-07-23 Kabushiki Kaisha Toshiba Cooled turbine blade for a gas turbine
DE19807563A1 (en) * 1997-02-25 1998-09-24 Mitsubishi Heavy Ind Ltd Cooling construction for mounting plate of turbine rotor blade
EP1538305A2 (en) * 2003-11-19 2005-06-08 United Technologies Corporation Airfoil with variable density array of pedestals at the trailing edge

Family Cites Families (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2473621A1 (en) * 1980-01-10 1981-07-17 Snecma DAWN OF TURBINE DISPENSER
US4775296A (en) * 1981-12-28 1988-10-04 United Technologies Corporation Coolable airfoil for a rotary machine
US5772397A (en) * 1996-05-08 1998-06-30 Alliedsignal Inc. Gas turbine airfoil with aft internal cooling
DE19921644B4 (en) * 1999-05-10 2012-01-05 Alstom Coolable blade for a gas turbine
US7014424B2 (en) * 2003-04-08 2006-03-21 United Technologies Corporation Turbine element
US7150601B2 (en) * 2004-12-23 2006-12-19 United Technologies Corporation Turbine airfoil cooling passageway
US7438527B2 (en) * 2005-04-22 2008-10-21 United Technologies Corporation Airfoil trailing edge cooling

Patent Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4278400A (en) * 1978-09-05 1981-07-14 United Technologies Corporation Coolable rotor blade
JPH05156901A (en) * 1991-12-02 1993-06-22 Hitachi Ltd Gas turbine cooling stationary blade
JPH06137102A (en) * 1992-10-26 1994-05-17 Mitsubishi Heavy Ind Ltd Hollow moving blade of gas turbine
US5538394A (en) * 1993-12-28 1996-07-23 Kabushiki Kaisha Toshiba Cooled turbine blade for a gas turbine
DE19807563A1 (en) * 1997-02-25 1998-09-24 Mitsubishi Heavy Ind Ltd Cooling construction for mounting plate of turbine rotor blade
EP1538305A2 (en) * 2003-11-19 2005-06-08 United Technologies Corporation Airfoil with variable density array of pedestals at the trailing edge

Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JP2011515618A (en) * 2008-03-28 2011-05-19 アルストム テクノロジー リミテッド Gas turbine stationary blade and gas turbine equipped with such a stationary blade
US8465255B2 (en) 2008-05-14 2013-06-18 Mitsubishi Heavy Industries, Ltd. Gas turbine blade and gas turbine having the same
JP2014047787A (en) * 2012-08-31 2014-03-17 General Electric Co <Ge> Airfoil and method for manufacturing airfoil

Also Published As

Publication number Publication date
EP1849960A3 (en) 2010-03-10
EP1849960A2 (en) 2007-10-31

Similar Documents

Publication Publication Date Title
JP2007292006A (en) Turbine blade having cooling passage inside thereof
US8790083B1 (en) Turbine airfoil with trailing edge cooling
US7887294B1 (en) Turbine airfoil with continuous curved diffusion film holes
US7527475B1 (en) Turbine blade with a near-wall cooling circuit
US7563072B1 (en) Turbine airfoil with near-wall spiral flow cooling circuit
JP5325664B2 (en) Crossflow turbine airfoil
US7611330B1 (en) Turbine blade with triple pass serpentine flow cooling circuit
US7637720B1 (en) Turbulator for a turbine airfoil cooling passage
US6379118B2 (en) Cooled blade for a gas turbine
US7901183B1 (en) Turbine blade with dual aft flowing triple pass serpentines
US7789626B1 (en) Turbine blade with showerhead film cooling holes
US8556583B2 (en) Blade cooling structure of gas turbine
US20060210390A1 (en) Film cooling for microcircuits
US8052390B1 (en) Turbine airfoil with showerhead cooling
US8459935B1 (en) Turbine vane with endwall cooling
US8585365B1 (en) Turbine blade with triple pass serpentine cooling
JP2006105152A (en) Stepped outlet turbine airfoil part
JP4929097B2 (en) Gas turbine blade
US6988872B2 (en) Turbine moving blade and gas turbine
JP2009281380A (en) Gas turbine airfoil
JP2000038901A (en) Hollow aerofoil
EP1818504B1 (en) Material having internal cooling passage and method for cooling material having internal cooling passage
US20130084191A1 (en) Turbine blade with impingement cavity cooling including pin fins
US8491264B1 (en) Turbine blade with trailing edge cooling
JP6438662B2 (en) Cooling passage of turbine blade of gas turbine engine

Legal Events

Date Code Title Description
A621 Written request for application examination

Free format text: JAPANESE INTERMEDIATE CODE: A621

Effective date: 20080602

A131 Notification of reasons for refusal

Free format text: JAPANESE INTERMEDIATE CODE: A131

Effective date: 20101109

A02 Decision of refusal

Free format text: JAPANESE INTERMEDIATE CODE: A02

Effective date: 20110308