JPH08338202A - Rotor blade of gas turbine - Google Patents
Rotor blade of gas turbineInfo
- Publication number
- JPH08338202A JPH08338202A JP14294995A JP14294995A JPH08338202A JP H08338202 A JPH08338202 A JP H08338202A JP 14294995 A JP14294995 A JP 14294995A JP 14294995 A JP14294995 A JP 14294995A JP H08338202 A JPH08338202 A JP H08338202A
- Authority
- JP
- Japan
- Prior art keywords
- blade
- cooling
- heat transfer
- gas turbine
- leading edge
- 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
Links
Landscapes
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
Description
【０００１】[0001]
【産業上の利用分野】本発明は、ガスタービンにおける
タービン動翼の改良に係り、特にその冷却構造に関する
ものである。BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to an improvement of turbine moving blades in a gas turbine, and more particularly to a cooling structure thereof.
【０００２】[0002]
【従来の技術】ガスタービンは圧縮機により圧縮された
高圧力の空気を酸化剤として燃料を燃焼させ、発生した
高温高圧ガスによりタービンを駆動し、たとえば電力等
のエネルギーに変換するものである。消費された燃料に
たいして得られる電力エネルギーは出来るだけ多い方が
望ましく、すなわちガスタービンの性能向上が期待され
ており、ガスタービンの性能向上を図る手段の一つとし
て作動ガスの高温高圧化が進められている。一方、ガス
タービン作動ガスの高温化を図り、高温排気ガスを利用
した蒸気タービンシステムとのコンバイドプラントによ
って、ガスタービンと蒸気タービンとを含めた総合エネ
ルギー変換効率向上方法も提案されている。ガスタービ
ンの作動ガス温度はタービン翼材が主にガス温度に起因
する熱応力に耐え得る能力によって制限される。作動ガ
ス温度の高温化にさいし、タービン翼の耐用温度を満足
させるため、タービン翼の中空部に冷却空気を供給し、
翼を冷却する方法が良く採られている。具体的には、タ
ービン動翼の場合、図示説明を省略するが翼の内部に１
つあるいはそれ以上の流路を形成させ、冷却空気を通過
させることによってタービン動翼を内部から冷却し、さ
らにタービン動翼の表面、先端あるいは後縁に設けられ
た冷却孔から翼外に出るようにする。かかる冷却空気は
一般に圧縮機から抽気した空気の一部を利用するので、
冷却空気の多量の消費はガスタービン効率の低下をきた
すため、より少ない空気により効率良く冷却することが
重要である。2. Description of the Related Art A gas turbine burns fuel by using high-pressure air compressed by a compressor as an oxidant, and drives the turbine by the generated high-temperature and high-pressure gas to convert it into energy such as electric power. It is desirable to obtain as much power energy as possible for the consumed fuel, that is, it is expected that the performance of the gas turbine will be improved, and as a means for improving the performance of the gas turbine, the high temperature and high pressure of the working gas is being promoted. ing. On the other hand, a method for improving the overall energy conversion efficiency including a gas turbine and a steam turbine has also been proposed in which a working temperature of the gas turbine working gas is increased and a combined plant with a steam turbine system that uses high-temperature exhaust gas is used. The working gas temperature of a gas turbine is limited by the ability of the turbine blade material to withstand thermal stress primarily due to gas temperature. In order to satisfy the service temperature of the turbine blade when the working gas temperature rises, cooling air is supplied to the hollow portion of the turbine blade,
The method of cooling the wings is often adopted. Specifically, in the case of a turbine rotor blade, although not shown in the drawings, it is
One or more flow passages are formed to cool the turbine blade from the inside by passing cooling air, and then to the outside of the blade from the cooling holes provided on the surface, the tip or the trailing edge of the turbine blade. To Since such cooling air generally uses a part of the air extracted from the compressor,
Since a large consumption of cooling air causes a decrease in gas turbine efficiency, it is important to cool efficiently with less air.
【０００３】より高温のガスタービンを実現する為に
は、翼内部の伝熱性能を改善し、供給する冷却空気量に
対して冷却効果をさらに良くすることが肝要であり、冷
却面に対していろいろな伝熱促進対策が施されている。
伝熱促進対策の方法には、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などにより
改善されることが良く知られており、翼内部の冷却面に
多数のリブを設ける方法がある。このような伝熱促進構
造を施した例は、例えば特開昭60-101202号、特開平5-1
0101号公報などに提案されている。特開昭60-101202号
公報に示された伝熱促進構造は、動翼内を複数の冷却流
路を形成しかかる冷却面に冷却空気流に斜めに置かれた
リブを配置しさらに伝熱促進スリットを施した構造であ
る。かかる伝熱促進リブ構造では、スリット後流の空気
流の乱れにより更に高い冷却伝熱性能が得られ、さらに
かかるスリットによってリブ周囲にごみが留まることを
防止し伝熱性能が低下することを防ぐことが出来るとさ
れている。また特開平5-10101号公報に示された伝熱促
進構造は、冷却面中央側のリブ先端を冷却空気流の上流
方向に向け、かつ左右のリブを冷却空気流れ方向に交互
に配置した特徴があり、高い伝熱性能が得られるとされ
ている。In order to realize a higher temperature gas turbine, it is important to improve the heat transfer performance inside the blade and to further improve the cooling effect with respect to the amount of cooling air supplied, and to the cooling surface. Various heat transfer promotion measures are taken.
It is well known that the method of promoting heat transfer is improved by making the air flow on the surface of the heat transfer surface turbulent or by breaking the boundary layer. There is a method of providing the rib. Examples of such a heat transfer promoting structure are disclosed in, for example, JP-A-60-101202 and JP-A-5-151.
It is proposed in Japanese Patent No. 0101. The heat transfer promoting structure disclosed in Japanese Patent Laid-Open No. 60-101202 has a plurality of cooling channels formed in the moving blade, and ribs placed obliquely to the cooling air flow are arranged on the cooling surface to further heat transfer. It is a structure with a promotion slit. With such a heat transfer promoting rib structure, higher cooling heat transfer performance can be obtained due to the turbulence of the air flow after the slit, and further, it is possible to prevent dust from staying around the ribs and prevent the heat transfer performance from being deteriorated by the slit. It is supposed to be possible. Further, the heat transfer promoting structure shown in Japanese Patent Laid-Open No. 5-10101 is characterized in that the rib tips on the center side of the cooling surface face the upstream direction of the cooling air flow, and the left and right ribs are alternately arranged in the cooling air flow direction. It is said that high heat transfer performance can be obtained.
【０００４】[0004]
【発明が解決しようとする課題】前記したごとくタービ
ン翼の冷却空気には圧縮機からの抽気空気を使用するた
め、冷却空気消費量の増加はガスタービンとしての熱効
率を低下させる。したがってガスタービンの冷却には少
ない空気量で効率良く冷却することが肝要であるが、前
記従来のガスタービン動翼の冷却構造ではさらなる作動
ガス温度の高温化に対して、特に熱負荷の高い翼前縁部
に対して冷却空気量を増加させて対処する必要があり、
ガスタービン熱効率の改善効果が小さい欠点が有った。As described above, since the extracted air from the compressor is used as the cooling air for the turbine blades, the increase in the cooling air consumption reduces the thermal efficiency of the gas turbine. Therefore, it is important to cool the gas turbine efficiently with a small amount of air. However, in the conventional cooling structure for a gas turbine moving blade described above, even if the working gas temperature is further increased, the blade having a particularly high heat load is required. It is necessary to increase the cooling air amount for the leading edge,
There was a drawback that the effect of improving the gas turbine thermal efficiency was small.
【０００５】すなわち本発明の目的は、特に翼前縁部を
対象に冷却伝熱性能のさらに良い伝熱促進構造を提供
し、ガスタービン動翼を少量の冷却空気で効果的に冷却
することを可能にし、ひいては熱効率の高い高温ガスタ
ービンを実現可能にすることにある。That is, an object of the present invention is to provide a structure for promoting heat transfer having a better cooling heat transfer performance, particularly for the leading edge of the blade, and to effectively cool the gas turbine moving blade with a small amount of cooling air. It is possible to realize a high temperature gas turbine with high thermal efficiency.
【０００６】[0006]
【課題を解決するための手段】本発明は、翼外壁に囲わ
れた内部空間を翼前縁から翼後縁にかけて翼長手方向に
設けられた複数の仕切壁によって分割されて形成された
複数の冷却流路を有するタービン動翼であって、タービ
ン動翼の前縁冷却流路の前縁冷却面に対向する仕切壁面
に複数の突起を流路に沿って設けることにより、所期の
目的を達成するようにしたものである。According to the present invention, a plurality of partition walls formed by partitioning an inner space surrounded by a blade outer wall by a plurality of partition walls provided in the blade longitudinal direction from a blade leading edge to a blade trailing edge are formed. A turbine rotor blade having a cooling flow passage, wherein a plurality of projections are provided along the flow passage on the partition wall surface facing the front edge cooling surface of the front edge cooling flow passage of the turbine rotor blade to achieve the intended purpose. It was something that was achieved.
【０００７】そして突起は流路方向と交叉する方向に設
けた板片で構成するのがよい。またタービン動翼は、各
冷却流路を形成する外壁内面にそれぞれ外壁内面から突
出する複数の乱流促進リブを冷却流路に沿って有するも
のがよい。It is preferable that the protrusion is formed by a plate piece provided in a direction intersecting with the flow path direction. Further, it is preferable that the turbine rotor blade has a plurality of turbulent flow promoting ribs protruding from the inner surface of the outer wall on the inner surface of the outer wall forming the respective cooling flow paths, along the cooling flow path.
【０００８】[0008]
【作用】すなわちこのように仕切壁に突起を形成する
と、冷却空気流の乱流が促進されること、さらに冷却空
気流が前縁の外壁内面である冷却面に押し付つけられる
ことにより、高い冷却熱伝達率を得ることが出来る。こ
の効果は実験により確認した。When the projection is formed on the partition wall in this way, the turbulent flow of the cooling air flow is promoted, and the cooling air flow is pressed against the cooling surface, which is the inner surface of the outer wall of the leading edge. The cooling heat transfer coefficient can be obtained. This effect was confirmed by experiments.
【０００９】[0009]
【実施例】以下、本発明の一実施例を図１〜図３に基づ
き説明する。図１は本発明を実施したガスタービン動翼
１の縦断面構造を示し、図２は図１におけるＡ−Ａ断面
図、図３は図１における”Ｂ”部拡大図である。図１お
よび図２において、２は軸部、３は軸部２から延びる翼
部、４および５はそれぞれ軸部２から翼部３にわたって
設けた内部流路である。各内部流路４，５は、翼部３に
おいて翼部３の長手方向に延びる複数の仕切壁6a、6b、
6c、6d……により複数の冷却流路7a、7b、7c、7d……に
仕切られ、先端曲部8a、8b…、下端曲部9a、9b…により
折返し流路を形成する。本実施例の場合、第１の内部流
路４は冷却流路7a、先端曲部8a、流路7b、下端曲部9a、
流路7cおよび翼先端壁10に設けた吹き出し孔11により構
成される。同様に第２の内部流路５は冷却流路7d、先端
曲部8b、流路7e、下端曲部9b、流路7fおよび翼後縁12に
設けた吹き出し部13により構成される。タービン動翼１
にはそれを設置したロータ軸(図示省略)などから冷却空
気が内部流路４，５それぞれの空気流入口14に供給さ
れ、内部流路４および５を通過する過程で翼１を内部か
ら冷却する。翼１を冷却した空気流15は、翼先端壁10に
設けた吹き出し孔11および翼後縁12の吹き出し部13よ
り、作動ガス主流中に吹き出される。DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS An embodiment of the present invention will be described below with reference to FIGS. FIG. 1 is a gas turbine rotor blade embodying the present invention.
1 shows a longitudinal section structure, FIG. 2 is a sectional view taken along A-A in FIG. 1, FIG. 3 is a "B" portion enlarged view of FIG. 1 and 2, 2 is a shaft portion, 3 is a blade portion extending from the shaft portion 2, and 4 and 5 are internal flow passages provided from the shaft portion 2 to the blade portion 3, respectively. Each of the internal flow paths 4, 5 has a plurality of partition walls 6a, 6b extending in the longitudinal direction of the wing portion 3 in the wing portion 3,
.. are divided into a plurality of cooling channels 7a, 7b, 7c, 7d .. In the case of the present embodiment, the first internal flow path 4 includes a cooling flow path 7a, a tip curved portion 8a, a flow channel 7b, a lower curved portion 9a,
It is composed of a flow path 7c and a blowout hole 11 provided in the blade tip wall 10. Similarly, the second internal flow path 5 is composed of a cooling flow path 7d, a tip curved portion 8b, a flow channel 7e, a lower end curved portion 9b, a flow channel 7f, and a blowing portion 13 provided at the blade trailing edge 12. Turbine blade 1
In this process, cooling air is supplied to the air inlets 14 of the internal passages 4 and 5 from a rotor shaft (not shown) on which the blades are installed, and the blade 1 is cooled from the inside while passing through the internal passages 4 and 5. To do. The airflow 15 that has cooled the blade 1 is blown into the mainstream of the working gas from the blowing hole 11 provided in the blade tip wall 10 and the blowing portion 13 of the blade trailing edge 12.
【００１０】冷却流路7a、7b、7c、7d……の冷却面、す
なわち翼１の外壁内面には、乱流促進リブ20が一体構造
で設けられており、その乱流促進リブ20の形状はタービ
ン動翼の冷却目標と設計思想により効果的に冷却するた
めに種々の形状が考えられ、本発明をより有効に発揮す
るには現在最も乱流促進に効果のあるＶ型スタッガード
配置構造がより良い。A turbulent flow promoting rib 20 is integrally formed on the cooling surface of the cooling flow paths 7a, 7b, 7c, 7d, that is, on the inner surface of the outer wall of the blade 1 , and the shape of the turbulent flow promoting rib 20 is provided. Various shapes are conceivable for effective cooling depending on the cooling target and design concept of the turbine blade, and the V-type staggered arrangement structure that is most effective for promoting turbulence at present is the most effective for exerting the present invention. Is better.
【００１１】前縁の半円弧状の冷却流路7aには、冷却面
に乱流促進リブ20を形成し、前縁冷却面に対向する仕切
壁6aには、本発明による突起21を一体構造で設置する。
本発明による伝熱促進構造、その作用および効果を図３
と図４により詳しく説明する。図３は、図１の”Ｂ”
部、すなわち前縁冷却流路7aの拡大図を示す。図３にお
いて20a、20b、20c……は前縁冷却流路7aの冷却面19に
一体構造で設けられた複数個の乱流促進リブであり、21
a、21b、21c……は前縁冷却流路 7aを形成する仕切壁6a
に一体構造で設けられた複数個の突起である。冷却空気
流15は、仕切壁6aの突起21a、21b、21c……の後流に発
生する大きな乱流渦22a、22b、22c……により加速さ
れ、かつ前縁冷却流路7aの冷却面19の方向に押し付けら
れる。しかして前縁冷却流路7aの冷却面19には、乱流促
進リブ20a、20b、20c……の後流に従来以上の強力な乱
流渦23a、23b、23c……が発生し、高い乱流促進効果、
すなわち高い冷却熱伝達率を得る伝熱促進効果がある。A turbulent flow promoting rib 20 is formed on the cooling surface in the semicircular cooling passage 7a at the front edge, and a projection 21 according to the present invention is integrally formed on the partition wall 6a facing the front edge cooling surface. Install in.
FIG. 3 shows the structure for promoting heat transfer according to the present invention, and its operation and effect.
Will be described in more detail with reference to FIG. FIG. 3 shows “B” of FIG.
An enlarged view of a portion, that is, the leading edge cooling flow path 7a is shown. In FIG. 3, reference numerals 20a, 20b, 20c ... Indicate a plurality of turbulent flow promoting ribs integrally provided on the cooling surface 19 of the leading edge cooling passage 7a.
a, 21b, 21c ... are partition walls 6a forming the leading edge cooling passage 7a
Is a plurality of protrusions provided as an integral structure. The cooling air flow 15 is accelerated by the large turbulent vortices 22a, 22b, 22c, ... Generated in the wake of the projections 21a, 21b, 21c ... Of the partition wall 6a, and the cooling surface 19 of the leading edge cooling flow path 7a. Is pressed in the direction of. On the cooling surface 19 of the leading edge cooling passage 7a, however, turbulent vortices 23a, 23b, 23c, which are stronger than before, are generated behind the turbulent flow promoting ribs 20a, 20b, 20c. Turbulence promoting effect,
That is, there is a heat transfer promotion effect for obtaining a high cooling heat transfer coefficient.
【００１２】かかる伝熱促進効果をモデル伝熱試験によ
り確認した。表１に実験供試モデルの形状および実験条
件を示す。実験モデルは、流路幅Ｗ(図２参照)：10mm，
流路高さＨ：10mm，頂点半径ｒ：４mmのタービン動翼の
前縁冷却流路を模擬した釣鐘型の半円弧流路を形成して
冷却空気を供給し、一方の半円弧面に乱流促進リブを設
けた加熱伝熱面とし、相対する底面を断熱層として伝熱
面の熱伝達率を測定した。伝熱モデル供試体形状は、比
較の基準として平滑面流路形状、前記従来構造第２
の例、すなわち特開平5-10101号に記載されている乱流
促進リブ構造、本発明構造、すなわち上記乱流促進リ
ブに加えて仕切壁に突起を設けた構造について実施し
た。The effect of promoting heat transfer was confirmed by a model heat transfer test. Table 1 shows the shape of the experimental test model and the experimental conditions. The experimental model is the flow path width W (see FIG. 2): 10 mm,
Cooling air is supplied by forming a bell-shaped semicircular arc flow path simulating the leading edge cooling flow path of a turbine rotor blade with a flow path height H: 10 mm and apex radius r: 4 mm, and one half arc surface is disturbed. The heat transfer coefficient of the heat transfer surface was measured using the heat transfer surface provided with the flow promoting ribs and the opposing bottom surface as the heat insulating layer. The heat transfer model specimen shape is the smooth surface flow path shape, the conventional structure No. 2 as a reference for comparison.
Example, that is, the turbulent flow promoting rib structure described in JP-A-5-10101, and the structure of the present invention, that is, a structure in which projections are provided on the partition wall in addition to the turbulent flow promoting rib.
【００１３】[0013]
【表１】 [Table 1]
【００１４】表１から明らかなように試験体，の伝
熱促進リブは、形状的にほぼ等価(リブ高さ、幅、ピッ
チ(ピッチ/リブ高さ=10)が同じなので)である。実験は
伝熱面側を加熱し、流路側に低温空気を供給して実施し
た。図４に、それぞれの伝熱特性実験結果を比較して示
した。図４では、冷却空気の流れ状況を示す無次元値レ
イノルズ数を横軸とし、熱の流れ状況を示す無次元値平
均ヌセルト数と伝熱促進リブを施していない平滑伝熱面
（試験体）の平均ヌセルト数との比を縦軸として比較
した。本図において同じレイノルズ数(同じ冷却条件)で
縦軸の値が大きいほど冷却性能が良いことを示す。図に
示されるように、従来構造供試体に比較して本発明構
造の伝熱性能は高いことが明らかである。ガスタービン
の定格運転時の冷却空気供給条件にほぼ近いレイノルズ
数5×105では、本発明構造（）の方が従来構造（）
に比較して約20％伝熱性能が高く、本発明構造の効果が
明らかとなった。As is apparent from Table 1, the heat transfer promoting ribs of the test body are substantially equivalent in shape (because the rib height, width and pitch (pitch / rib height = 10) are the same). The experiment was conducted by heating the heat transfer surface side and supplying low temperature air to the flow path side. FIG. 4 shows the results of the heat transfer characteristic experiments for comparison. In FIG. 4, the dimensionless Reynolds number showing the flow of cooling air is taken as the horizontal axis, and the dimensionless average Nusselt number showing the flow of heat and the smooth heat transfer surface without the heat transfer promotion rib (test body) The ratio was compared with the average Nusselt number of the vertical axis. In the figure, the larger the value on the vertical axis under the same Reynolds number (same cooling condition), the better the cooling performance. As shown in the figure, it is clear that the heat transfer performance of the structure of the present invention is higher than that of the conventional structure specimen. When the Reynolds number is 5 × 10 5 , which is close to the cooling air supply condition during the rated operation of the gas turbine, the structure of the present invention () has the conventional structure ().
The heat transfer performance is about 20% higher than that of the above, and the effect of the structure of the present invention is clarified.
【００１５】なお、本発明の説明では、２本の内部流路
を有したリターンフロー型構造を例に示したが、本発明
の適用に冷却流路数および流路構成（折り返し、冷却媒
体の出口位置、乱流促進リブの有無など）の限定を与え
るものではない。また、実施例では冷却媒体を空気とし
て説明したが、蒸気、水、窒素など他の冷却媒体でも良
いことは当然のことである。In the description of the present invention, the return flow type structure having two internal flow paths has been taken as an example, but the number of cooling flow paths and the flow path configuration (folding, cooling medium It does not limit the exit position, the presence of turbulence promoting ribs, etc.). In addition, although the cooling medium is described as air in the embodiments, it goes without saying that other cooling medium such as steam, water, or nitrogen may be used.
【００１６】なお、本発明構造を採用したガスタービン
動翼は、現状の精密鋳造方法により製作可能である。The gas turbine rotor blade adopting the structure of the present invention can be manufactured by the current precision casting method.
【００１７】[0017]
【発明の効果】以上説明のごとく本発明の伝熱促進構造
は、翼前縁部の冷却流路を形成する仕切壁面に設けた突
起により冷却媒体の乱流がさらに促進されること、さら
に冷却媒体の流れが前縁冷却面に押し付けられること、
これらの相乗効果により高い伝熱促進効果すなわち高い
冷却熱伝達率を得ることが出来る。したがって本発明
は、少ない冷却媒体量で効率良く冷却することが可能な
ガスタービン動翼を提供でき、ひいては高温高効率ガス
タービンに寄与出来る。As described above, in the heat transfer promoting structure of the present invention, the turbulent flow of the cooling medium is further promoted by the projections provided on the partition wall surface forming the cooling flow passage at the blade leading edge portion, and further cooling is performed. The flow of media is pressed against the leading edge cooling surface,
Due to these synergistic effects, a high heat transfer promotion effect, that is, a high cooling heat transfer coefficient can be obtained. Therefore, the present invention can provide a gas turbine blade capable of efficiently cooling with a small amount of cooling medium, and can contribute to a high-temperature and high-efficiency gas turbine.
【図１】本発明の一実施例のガスタービン動翼の縦断面
構造図である。FIG. 1 is a vertical cross-sectional structural diagram of a gas turbine rotor blade in one embodiment of the present invention.
【図２】図１のＡ-Ａ断面図である。FIG. 2 is a sectional view taken along line AA of FIG.
【図３】図１の”Ｂ”部拡大図である。FIG. 3 is an enlarged view of a “B” part in FIG. 1.
【図４】本発明のモデルによる伝熱特性実験結果を示す
グラフである。FIG. 4 is a graph showing experimental results of heat transfer characteristics according to the model of the present invention.
１ ガスタービン動翼 ２ 軸部 ３ 翼部 ４、５ 内部流路 ６ａ、６ｂ、６ｃ、６ｄ 仕切壁 ７ａ、７ｂ、７ｃ、７ｄ 冷却流路 ８ａ、８ｂ 先端曲部 ９ａ、９ｂ 下端曲部 １０ 翼先端壁 １１ 翼先端吹き出し孔 １２ 翼後縁 １３ 翼後縁吹き出し部 １４ 空気流入口 １５ 冷却空気流 １９ 前縁冷却面 ２０、２０ａ、２０ｂ、２０ｃ 乱流促進リブ ２１、２１ａ、２１ｂ、２１ｃ 突起 ２２ａ、２２ｂ、２２ｃ 乱流渦 ２３ａ、２３ｂ、２３ｃ 乱流渦 1 Gas Turbine Blade 2 Shaft Part 3 Blade Part 4, 5 Internal Flow Paths 6a, 6b, 6c, 6d Partition Walls 7a, 7b, 7c, 7d Cooling Flow Paths 8a, 8b Tip Curved Section 9a, 9b Bottom Curved Section 10 Blade Tip wall 11 Blade tip blowing hole 12 Blade trailing edge 13 Blade trailing edge blowing portion 14 Air inlet 15 Cooling air flow 19 Leading edge cooling surface 20, 20a, 20b, 20c Turbulent flow promoting rib 21, 21a, 21b, 21c Protrusion 22a , 22b, 22c Turbulent vortex 23a, 23b, 23c Turbulent vortex
フロントページの続き (72)発明者 和田 克夫 茨城県日立市幸町三丁目１番１号 株式会 社日立製作所日立工場内 (72)発明者 金沢 幸夫 宮城県仙台市青葉区一番町三丁目７番１号 東北電力株式会社内 (72)発明者 生越 昭男 宮城県仙台市青葉区一番町三丁目７番１号 東北電力株式会社内Front page continuation (72) Inventor Katsuo Wada 3-1-1, Sachimachi, Hitachi City, Ibaraki Hitachi Ltd. Hitachi factory (72) Inventor Yukio Kanazawa 3-chome, Ichibancho, Aoba-ku, Sendai City, Miyagi Prefecture No. 1 Tohoku Electric Power Co., Inc. (72) Inventor Akio Ikukoshi 3-7-1, Ichibancho, Aoba-ku, Sendai City, Miyagi Prefecture Tohoku Electric Power Co., Inc.
Claims (3)
翼後縁にかけて翼長手方向に設けられた複数の仕切壁に
よって分割されて形成された複数の冷却流路を有するタ
ービン動翼において、翼前縁の外壁内面に対向する仕切
壁に複数の突起を冷却流路方向に設けたことを特徴とす
るタービン動翼。1. A turbine rotor blade having a plurality of cooling channels formed by dividing an inner space surrounded by an outer wall of a blade by a plurality of partition walls provided in a blade longitudinal direction from a blade leading edge to a blade trailing edge. In the turbine blade, a plurality of projections are provided in a cooling flow path direction on a partition wall facing the inner surface of the outer wall of the blade leading edge.
けた板片からなることを特徴とする請求項１記載のター
ビン動翼。2. The turbine rotor blade according to claim 1, wherein the projection is formed of a plate piece provided in a direction intersecting with a flow passage direction.
れぞれ該外壁内面から突出する複数の乱流促進リブを冷
却流路に沿って有することを特徴とする請求項１または
２記載のタービン動翼。3. The turbine according to claim 1, further comprising a plurality of turbulent flow promoting ribs protruding from the inner surface of the outer wall on the inner surface of the outer wall forming the cooling passages, along the cooling passage. Moving blade.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP14294995A JPH08338202A (en) | 1995-06-09 | 1995-06-09 | Rotor blade of gas turbine |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP14294995A JPH08338202A (en) | 1995-06-09 | 1995-06-09 | Rotor blade of gas turbine |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| JPH08338202A true JPH08338202A (en) | 1996-12-24 |
Family
ID=15327402
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| JP14294995A Pending JPH08338202A (en) | 1995-06-09 | 1995-06-09 | Rotor blade of gas turbine |
Country Status (1)
| Country | Link |
|---|---|
| JP (1) | JPH08338202A (en) |
Cited By (10)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JP2008002465A (en) * | 2006-06-22 | 2008-01-10 | United Technol Corp <Utc> | Turbine engine component |
| US9850762B2 (en) | 2013-03-13 | 2017-12-26 | General Electric Company | Dust mitigation for turbine blade tip turns |
| US9957816B2 (en) | 2014-05-29 | 2018-05-01 | General Electric Company | Angled impingement insert |
| US9995148B2 (en) | 2012-10-04 | 2018-06-12 | General Electric Company | Method and apparatus for cooling gas turbine and rotor blades |
| US10233775B2 (en) | 2014-10-31 | 2019-03-19 | General Electric Company | Engine component for a gas turbine engine |
| US10280785B2 (en) | 2014-10-31 | 2019-05-07 | General Electric Company | Shroud assembly for a turbine engine |
| US10364684B2 (en) | 2014-05-29 | 2019-07-30 | General Electric Company | Fastback vorticor pin |
| US10422235B2 (en) | 2014-05-29 | 2019-09-24 | General Electric Company | Angled impingement inserts with cooling features |
| US10563514B2 (en) | 2014-05-29 | 2020-02-18 | General Electric Company | Fastback turbulator |
| US10690055B2 (en) | 2014-05-29 | 2020-06-23 | General Electric Company | Engine components with impingement cooling features |
-
1995
- 1995-06-09 JP JP14294995A patent/JPH08338202A/en active Pending
Cited By (10)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JP2008002465A (en) * | 2006-06-22 | 2008-01-10 | United Technol Corp <Utc> | Turbine engine component |
| US9995148B2 (en) | 2012-10-04 | 2018-06-12 | General Electric Company | Method and apparatus for cooling gas turbine and rotor blades |
| US9850762B2 (en) | 2013-03-13 | 2017-12-26 | General Electric Company | Dust mitigation for turbine blade tip turns |
| US9957816B2 (en) | 2014-05-29 | 2018-05-01 | General Electric Company | Angled impingement insert |
| US10364684B2 (en) | 2014-05-29 | 2019-07-30 | General Electric Company | Fastback vorticor pin |
| US10422235B2 (en) | 2014-05-29 | 2019-09-24 | General Electric Company | Angled impingement inserts with cooling features |
| US10563514B2 (en) | 2014-05-29 | 2020-02-18 | General Electric Company | Fastback turbulator |
| US10690055B2 (en) | 2014-05-29 | 2020-06-23 | General Electric Company | Engine components with impingement cooling features |
| US10233775B2 (en) | 2014-10-31 | 2019-03-19 | General Electric Company | Engine component for a gas turbine engine |
| US10280785B2 (en) | 2014-10-31 | 2019-05-07 | General Electric Company | Shroud assembly for a turbine engine |
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