JPH06102963B2 - Gas turbine air cooling blade - Google Patents
Gas turbine air cooling bladeInfo
- Publication number
- JPH06102963B2 JPH06102963B2 JP58241001A JP24100183A JPH06102963B2 JP H06102963 B2 JPH06102963 B2 JP H06102963B2 JP 58241001 A JP58241001 A JP 58241001A JP 24100183 A JP24100183 A JP 24100183A JP H06102963 B2 JPH06102963 B2 JP H06102963B2
- Authority
- JP
- Japan
- Prior art keywords
- blade
- cooling
- gas turbine
- cooling air
- return flow
- 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 - Lifetime
Links
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
Description
【発明の詳細な説明】 ［発明の属する技術分野］ 本発明はガスタービン空冷翼に係り、特に冷却性能を向
上させた翼に関する。Description: TECHNICAL FIELD The present invention relates to a gas turbine air cooling blade, and more particularly to a blade having improved cooling performance.
［従来技術とその問題点］ 周知のように、ガスタービンは往復機関に比較して小型
軽量で大馬力が得られるなどの多くの利点を有してい
る。[Prior Art and its Problems] As is well known, a gas turbine has many advantages such as small size and light weight and large horsepower as compared with a reciprocating engine.
このようなガスタービン、たとえば等圧燃焼式のものを
例にとると、通常第１図に示すように筒状のケーシング
１内に軸２を回転自在に設け、この軸２の両端部とケー
シンズ１との間にそれぞれ圧縮機３とパワータービン４
とを構成し、圧縮機３で圧縮された高圧空気で燃焼器５
内の圧力を高め、この状態で燃料を噴射させて燃焼さ
せ、この燃焼によって生じた超高圧の高温ガスをパワー
タービン４に導いて膨張させることにより、軸２の回転
動力を得るように構成されている。そして、圧縮機３
は、図の場合では案内羽根６と回転羽根７とを軸方向へ
配列して軸流型とし、また、パワータービン４は軸２に
固定された動翼８とケーシング１に固定された静翼９と
を軸方向へ交互に配列して構成されている。Taking such a gas turbine, for example, an isobaric combustion type as an example, a shaft 2 is normally rotatably provided in a cylindrical casing 1 as shown in FIG. 1, and both ends of the shaft 2 and the casings are 1 and compressor 3 and power turbine 4 respectively
And the high pressure air compressed by the compressor 3 constitutes the combustor 5
The internal pressure is increased, fuel is injected and burned in this state, and the high-pressure gas of ultrahigh pressure generated by this combustion is guided to the power turbine 4 and expanded, whereby rotational power of the shaft 2 is obtained. ing. And the compressor 3
In the case of the figure, the guide vanes 6 and the rotary vanes 7 are arranged in the axial direction to form an axial flow type, and the power turbine 4 has a moving blade 8 fixed to the shaft 2 and a stationary blade fixed to the casing 1. 9 and 9 are alternately arranged in the axial direction.
ところで、上記のようなガスタービンにおいて、効率を
向上させる為にはパワータービン４の入口におけるガス
温度を高めることが最も有効な手段であると云われてい
る。しかし、パワータービン４を構成する金属材料の許
容温度は、一般的に、850℃程度であり、これ以上にガ
ス温度を上げるにはパワータービン４を構成する部材、
特に翼を効率よく冷却する必要がある。By the way, in the gas turbine as described above, increasing the gas temperature at the inlet of the power turbine 4 is said to be the most effective means for improving the efficiency. However, the allowable temperature of the metal material that constitutes the power turbine 4 is generally about 850 ° C., and in order to raise the gas temperature above this, the members that constitute the power turbine 4,
Especially, it is necessary to cool the blades efficiently.
従来用いられている空気冷却方式を採用した代表的な例
を第２図，第３図に示す。ここでは翼根元10から供給さ
れる冷却空気は、ひとつは翼前縁部11へ供給され、対流
冷却並びに翼前縁12に設けられたフィルム孔13から冷却
空気を吹き出し、フィルム冷却がなされる。翼中央部14
に供給される冷却空気は翼後縁から翼前縁方向へ向かっ
てのリターンフロー流路が形成され対流冷却がなされ
る。当然ながら途中、翼の腹側、背側に設けられたフィ
ルム孔15,16からは冷却空気が吹き出され、フィルム冷
却がなされる。翼内壁には熱伝達を促進させる為の突起
であるタービュレンスプロモータ17が流れに対向して設
けられている。翼後縁部18へ供給される冷却空気は内部
流れに対向して設けられたピンフィン19によるピンフィ
ン対流冷却がなされ、翼後縁20から吹き抜ける構造とな
っている。A typical example using the conventionally used air cooling method is shown in FIGS. 2 and 3. Here, one of the cooling air supplied from the blade root 10 is supplied to the blade leading edge portion 11, and the film is cooled by convection cooling and blowing the cooling air from the film hole 13 provided in the blade leading edge 12. Central wing 14
The return air flow path from the trailing edge of the blade toward the leading edge of the blade is formed in the cooling air to be convectively cooled. As a matter of course, cooling air is blown out from the film holes 15 and 16 provided on the ventral side and the dorsal side of the blade on the way to cool the film. A turbulence promoter 17, which is a projection for promoting heat transfer, is provided on the inner wall of the blade so as to face the flow. The cooling air supplied to the blade trailing edge portion 18 is subjected to pin fin convection cooling by a pin fin 19 provided so as to face the internal flow, and is blown out from the blade trailing edge 20.
このような翼においては、特に翼中央部のリターンフロ
ー部で翼内部を通過する冷却空気温度は流れに沿って徐
々に上昇する為に最後のリターン流路部では適切な冷却
効果が得られない等の問題があった。これは第４図，第
５図の翼外周部の熱伝達率分布に示すように、翼の背側
で最後のリターン流路に位置するＳ領域では熱伝達率の
値がかなり高くなっており、従って翼外面からの熱移動
が大きく、特に翼内部冷却が必要とされるにもかかわら
ず、上述の理由から翼内部での対流冷却が悪化し翼金属
温度が局所的に高くなってしまうからである。仮にこの
ような問題を取除く為に翼内部流路を逆方向に流すとす
ると、今度は翼の腹側に位置するＰ領域の翼金属温度が
上述と同理由から局所的に高くなり、翼母材を一様な温
度に保つのが困難になる。又、翼先端キャップにおいて
も、翼内部からの有効な冷却効果が得られない等の問題
もあった。In such a blade, the temperature of the cooling air passing through the inside of the blade in the return flow portion in the central portion of the blade gradually rises along the flow, so an appropriate cooling effect cannot be obtained in the final return flow passage portion. There was a problem such as. As shown in the heat transfer coefficient distributions on the outer circumference of the blade in FIGS. 4 and 5, the value of the heat transfer coefficient is considerably high in the S region located in the last return passage on the back side of the blade. Therefore, although heat transfer from the outer surface of the blade is large and cooling inside the blade is particularly required, the convective cooling inside the blade deteriorates and the blade metal temperature locally rises due to the above reason. Is. If, in order to eliminate such a problem, the flow path inside the blade is made to flow in the opposite direction, then the blade metal temperature in the P region located on the ventral side of the blade becomes locally high for the same reason as described above, and It becomes difficult to maintain the base material at a uniform temperature. Further, the blade tip cap also has a problem that an effective cooling effect from the inside of the blade cannot be obtained.
近年高効率のガスタービン装置の開発が進められてお
り、ますます主流ガス温度が上昇する傾向にあり、冷却
効果の優れたガスタービン冷却翼の出現が強く望まれて
いる。In recent years, highly efficient gas turbine devices have been developed, the mainstream gas temperature tends to rise more and more, and the emergence of gas turbine cooling blades having excellent cooling effect is strongly desired.
［発明の目的］ 本発明は、このような事情に鑑みてなされたもので、そ
の目的とするところは、高温のガスにさらされるガスタ
ービン翼の冷却性能の向上にあり、特に翼の背側、腹側
共翼内で効率の良い対流冷却を行い、翼母材の温度低
減、温度分布の一様化を図ったガスタービン空冷翼を提
供することにある。[Object of the Invention] The present invention has been made in view of such circumstances, and an object of the present invention is to improve the cooling performance of a gas turbine blade exposed to high-temperature gas, and particularly to the back side of the blade. The purpose of the present invention is to provide a gas turbine air-cooling blade in which efficient convection cooling is performed in the ventral co-blade to reduce the temperature of the blade base material and uniform the temperature distribution.
［発明の概要］ 本発明は高温、高圧ガスにさらされるガスタービン空冷
翼において翼内部中央のリターンフロー流路を有する翼
にあっては、その冷却空気供給部を翼の腹側と翼の背側
で独自に設け、翼腹側部では翼後縁方向から翼前縁方向
に向かって冷却空気が流動し、翼背側部では翼前縁方向
から翼後縁方向へ向かって冷却空気が流動するように構
成するとともに、いずれか一方のリターンフロー流路の
終端部を翼先端内部に設けられたセルへ導き、さらに翼
先端キャップの腹側に設けられた小孔から冷却空気を吹
き出すことによって翼母材の温度低減、温度分布を均一
化させたことを特徴とするガスタービン空冷翼を提供す
る。SUMMARY OF THE INVENTION The present invention relates to a gas turbine air-cooling blade that is exposed to high temperature and high pressure gas and has a return flow channel in the center of the blade. The cooling air flows from the trailing edge direction of the blade to the leading edge direction of the blade on the ventral side of the blade, and the cooling air flows from the leading edge direction of the blade to the trailing edge direction of the blade on the back side of the blade. In addition, the end portion of either one of the return flow channels is guided to a cell provided inside the blade tip, and cooling air is blown out from a small hole provided on the ventral side of the blade tip cap. Provided is a gas turbine air-cooling blade, characterized in that the temperature of a blade base material is reduced and the temperature distribution is made uniform.
［発明の効果］ 本発明によれば、リターンフロー冷却方式による翼背
側、腹側の翼母材温度分布の均一化並びに温度低減化さ
らには翼先端部付近の冷却性能の向上、翼先端とケーシ
ングからの主流ガスリークによる翼列性能低下の防止が
施されたガスタービン空冷翼を提供することが可能とな
る。[Advantages of the Invention] According to the present invention, the temperature distribution of the blade base metal on the blade back side and the ventral side is made uniform and the temperature is reduced by the return flow cooling method, and further, the cooling performance near the blade tip portion is improved, and It is possible to provide a gas turbine air-cooling blade in which deterioration of blade cascade performance due to mainstream gas leakage from the casing is prevented.
［発明の実施例］ 本発明の実施例を図面により説明する。Embodiments of the Invention Embodiments of the present invention will be described with reference to the drawings.
第６図は本発明による翼コード方向の断面図を示すもの
であり、翼中央部に位置するリターンフロー部は翼の背
側、腹側で二分され、それぞれ独自に冷却空気が供給、
流通される。そして、翼外周面の熱伝達率が高いとされ
る位置を冷却空気供給口21,22とし、翼の背側では翼の
前縁方向から翼後縁方向へ、翼の腹側ではその逆にして
互いに対向して流れるようになっている。第７図は第６
図における線A-B-C-Dで結ぶ半径方向の翼背側の断面を
示すものであり、翼中央部の冷却空気供給口21へ供給さ
れた冷却空気は翼後縁方向へ向かってリターンフロー流
路を辿り、翼先端内部に設けられたセル23へ導かれる。
又、ここでは翼内側の熱伝達率を促進させるためのター
ビュレンスプロモータの突起24が流れに対向して設けら
れている。さらに翼先端キャップ25には第９図に示すよ
うに翼腹側に沿って複数個のフィルム孔26が設けられて
おり、この部分から冷却空気が吹き出され、翼先端部の
冷却並びに翼先端とケーシング間のエアーシールも合せ
て行われる。第８図は第６図における線A-E-C-Dで結ば
れる半径方向の翼腹側の断面図を示すものであり、翼中
央部の冷却空気供給口22に供給される冷却空気は翼後縁
部から翼前縁方向に向かってフィルム冷却を行いながら
リターンフロー流路を辿る。FIG. 6 is a cross-sectional view of the blade cord direction according to the present invention, in which the return flow portion located in the blade central portion is divided into two on the back side and the ventral side of the blade, and cooling air is supplied independently to each of them.
Be distributed. Then, the positions where the heat transfer coefficient of the outer peripheral surface of the blade is considered to be high are set as the cooling air supply ports 21 and 22. And flow opposite each other. FIG. 7 is the sixth
It shows a cross section of the blade back side in the radial direction connected by a line ABCD in the figure, the cooling air supplied to the cooling air supply port 21 at the blade central portion follows the return flow channel toward the blade trailing edge direction, It is guided to a cell 23 provided inside the blade tip.
Further, here, a protrusion 24 of the turbulence promoter for promoting the heat transfer coefficient inside the blade is provided so as to face the flow. Further, as shown in FIG. 9, the blade tip cap 25 is provided with a plurality of film holes 26 along the blade ventral side. Cooling air is blown from this portion to cool the blade tip and to remove the blade tip. Air sealing between the casings is also performed. FIG. 8 is a cross-sectional view of the ventral side of the blade in the radial direction, which is connected by the line AECD in FIG. 6, and the cooling air supplied to the cooling air supply port 22 in the central portion of the blade is transferred from the trailing edge of the blade to the blade. Follow the return flow channel while cooling the film toward the front edge.
以上のような構造にすることにより、従来困難であった
リターンフロー冷却方式による翼背側、腹側の翼金属温
度分布の均一化並びに温度低減化さらには翼先端部付近
の冷却性能向上、翼先端とケーシング間からの主流ガス
リークによる翼列性能低下の防止が施された優れたガス
タービン空冷翼を提供することが可能となった。By adopting the above structure, the return flow cooling method, which has been difficult to achieve in the past, makes the temperature distribution of the blade metal on the blade back and ventral sides uniform and reduces the temperature, and improves the cooling performance near the blade tip. It has become possible to provide an excellent gas turbine air-cooling blade in which deterioration of blade cascade performance due to mainstream gas leakage between the tip and the casing is prevented.
［発明の他の実施例］ 本発明による冷却翼内部構造は特に翼背側へ供給された
冷却空気の一部が翼先端部の冷却に使用されているが翼
内部流動条件から翼腹側へ供給される冷却空気の一部を
翼先端部の冷却に用いてもいっこうにかまわない。[Other Embodiments of the Invention] In the internal structure of the cooling blade according to the present invention, a part of the cooling air supplied to the blade back side is used for cooling the blade tip portion. It does not matter if a part of the supplied cooling air is used for cooling the blade tip.
又、本発明に係る翼は高温、高圧用のガスタービン装置
において、高圧段動、静翼の冷却を必要とされる翼に広
く適用できる。Further, the blade according to the present invention can be widely applied to a gas turbine device for high temperature and high pressure, which requires high pressure stepping and cooling of the stationary blade.
第１図はガスタービンを一部切欠して示す側面図、第２
図は従来用いられる空冷タービン翼の構造図、第３図は
第２図におけるZ-Z断面図、第４図は翼外周面での熱伝
達率分布を説明する為に示す図、第５図は翼外周面を説
明するために示す図、第６図は本発明におけるガスター
ビン空冷翼の一実施例の断面図、第７図は第６図におけ
るA-B-C-D断面を示す縦断面図、第８図は第６図におけ
るA-E-C-D断面を示す縦断面図、第９図は第７図，第８
図におけるX-X断面を示す横断面図である。 21,22……冷却空気供給口、23……セル、25……キャッ
プFIG. 1 is a side view showing a partially cutaway view of a gas turbine, and FIG.
The figure shows the structure of a conventional air-cooled turbine blade, Fig. 3 shows the ZZ cross section in Fig. 2, Fig. 4 shows the heat transfer coefficient distribution on the outer peripheral surface of the blade, and Fig. 5 shows the blade. FIG. 6 is a view for explaining the outer peripheral surface, FIG. 6 is a cross-sectional view of an embodiment of a gas turbine air-cooling blade according to the present invention, FIG. 7 is a vertical cross-sectional view showing an ABCD cross-section in FIG. 6, and FIG. FIG. 6 is a vertical sectional view showing an AEDC section in FIG. 6, FIG. 9 is FIGS.
It is a cross-sectional view showing a XX cross section in the drawing. 21,22 …… Cooling air supply port, 23 …… Cell, 25 …… Cap
Claims (2)
フローによって冷却されるガスタービン空冷翼におい
て、翼内部のリターンフロー流路が翼の腹側部と背側部
とに独立して設けられ、翼腹側部では翼後縁方向から翼
前縁方向に向かって冷却空気が流動し、翼背側部では翼
前縁方向から翼後縁方向へ向かって冷却空気が流動する
ように構成されていることを特徴とするガスタービン空
冷翼。1. A gas turbine air-cooling blade, to which cooling air is supplied from the blade root and is cooled by a return flow, is provided with return flow passages inside the blade independently on the ventral side and the backside of the blade. , The cooling air flows from the trailing edge direction of the blade toward the leading edge direction of the blade, and the cooling air flows from the leading edge direction of the blade toward the trailing edge direction of the blade at the back side of the blade. The gas turbine air cooling blade is characterized by having
側及び翼背側のどちらか一方のリターンフロー流路の終
端部は、翼先端の内部空間へ導かれ、さらに翼先端部の
翼腹側に面して設けられた複数個の小孔から冷却空気を
吹き出すように構成されていることを特徴とする特許請
求の範囲第１項記載のガスタービン空冷翼。2. An end portion of the return flow channel on either the ventral side or the back side of the blade in the return flow channel inside the blade is guided to the internal space of the blade tip, and further the blade belly of the blade tip portion. The gas turbine air-cooling blade according to claim 1, characterized in that cooling air is blown out from a plurality of small holes provided facing the side.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP58241001A JPH06102963B2 (en) | 1983-12-22 | 1983-12-22 | Gas turbine air cooling blade |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP58241001A JPH06102963B2 (en) | 1983-12-22 | 1983-12-22 | Gas turbine air cooling blade |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| JPS60135606A JPS60135606A (en) | 1985-07-19 |
| JPH06102963B2 true JPH06102963B2 (en) | 1994-12-14 |
Family
ID=17067848
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| JP58241001A Expired - Lifetime JPH06102963B2 (en) | 1983-12-22 | 1983-12-22 | Gas turbine air cooling blade |
Country Status (1)
| Country | Link |
|---|---|
| JP (1) | JPH06102963B2 (en) |
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| DE60237350D1 (en) * | 2002-05-09 | 2010-09-30 | Gen Electric | Turbine blade with triple backward winding cooling channels |
| US6932570B2 (en) | 2002-05-23 | 2005-08-23 | General Electric Company | Methods and apparatus for extending gas turbine engine airfoils useful life |
| EP1630354B1 (en) | 2004-08-25 | 2014-06-18 | Rolls-Royce Plc | Cooled gas turbine aerofoil |
| JP2006105084A (en) * | 2004-10-08 | 2006-04-20 | Mitsubishi Heavy Ind Ltd | Gas turbine moving blade |
| FR2887287B1 (en) | 2005-06-21 | 2007-09-21 | Snecma Moteurs Sa | COOLING CIRCUITS FOR MOBILE TURBINE DRIVE |
| US7744347B2 (en) | 2005-11-08 | 2010-06-29 | United Technologies Corporation | Peripheral microcircuit serpentine cooling for turbine airfoils |
| FR2893974B1 (en) * | 2005-11-28 | 2011-03-18 | Snecma | CENTRAL COOLING CIRCUIT FOR MOBILE TURBINE DRIVE |
| US7296973B2 (en) * | 2005-12-05 | 2007-11-20 | General Electric Company | Parallel serpentine cooled blade |
| US7413403B2 (en) * | 2005-12-22 | 2008-08-19 | United Technologies Corporation | Turbine blade tip cooling |
| US7699583B2 (en) * | 2006-07-21 | 2010-04-20 | United Technologies Corporation | Serpentine microcircuit vortex turbulatons for blade cooling |
| US7513744B2 (en) * | 2006-07-18 | 2009-04-07 | United Technologies Corporation | Microcircuit cooling and tip blowing |
| JP5730418B2 (en) * | 2014-01-06 | 2015-06-10 | 三菱日立パワーシステムズ株式会社 | Wing body and gas turbine |
| US10119405B2 (en) * | 2015-12-21 | 2018-11-06 | General Electric Company | Cooling circuit for a multi-wall blade |
| US9926788B2 (en) * | 2015-12-21 | 2018-03-27 | General Electric Company | Cooling circuit for a multi-wall blade |
| US10053989B2 (en) * | 2015-12-21 | 2018-08-21 | General Electric Company | Cooling circuit for a multi-wall blade |
| US10851996B2 (en) * | 2018-07-06 | 2020-12-01 | Rolls-Royce North American Technologies Inc. | Turbulators for cooling heat shield of a combustor |
Citations (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US4312624A (en) | 1980-11-10 | 1982-01-26 | United Technologies Corporation | Air cooled hollow vane construction |
Family Cites Families (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB1404757A (en) * | 1971-08-25 | 1975-09-03 | Rolls Royce | Gas turbine engine blades |
| US4078930A (en) * | 1975-10-28 | 1978-03-14 | Xerox Corporation | Developer compositions comprising toner and carrier |
-
1983
- 1983-12-22 JP JP58241001A patent/JPH06102963B2/en not_active Expired - Lifetime
Patent Citations (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US4312624A (en) | 1980-11-10 | 1982-01-26 | United Technologies Corporation | Air cooled hollow vane construction |
Also Published As
| Publication number | Publication date |
|---|---|
| JPS60135606A (en) | 1985-07-19 |
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