JPH0660562B2 - Gas turbine air cooling blade - Google Patents
Gas turbine air cooling bladeInfo
- Publication number
- JPH0660562B2 JPH0660562B2 JP58091557A JP9155783A JPH0660562B2 JP H0660562 B2 JPH0660562 B2 JP H0660562B2 JP 58091557 A JP58091557 A JP 58091557A JP 9155783 A JP9155783 A JP 9155783A JP H0660562 B2 JPH0660562 B2 JP H0660562B2
- Authority
- JP
- Japan
- Prior art keywords
- blade
- film cooling
- cooling
- air
- gas turbine
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Expired - 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
- F01D5/186—Film cooling
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Description
【発明の詳細な説明】 ［発明の技術分野］ 本発明は、ガスタービン空冷翼に関する。Description: TECHNICAL FIELD OF THE INVENTION The present invention relates to a gas turbine air cooling blade.
［発明の技術的背景とその問題点］ ガスタービンにおいては、燃焼ガスによって駆動される
タービンが、バーナに燃焼用空気を供給するための圧縮
機を駆動する。かかるタービンは、比較的高温にて運転
されるが、その熱効率を向上させるため、そのタービン
入口温度を高温化する等の手段がとられる。[Technical Background of the Invention and Problems Thereof] In a gas turbine, a turbine driven by combustion gas drives a compressor for supplying combustion air to a burner. Although such a turbine is operated at a relatively high temperature, measures such as increasing the turbine inlet temperature are taken in order to improve its thermal efficiency.
しかし、特にタービン入口温度が1000゜Ｃを越すように
なると、タービン翼に使用される耐熱金属の使用限界温
度を上まわるため、冷却用空気を流すための中空構造を
持つ空冷翼が使用される。However, especially when the turbine inlet temperature exceeds 1000 ° C, the upper limit temperature of the refractory metal used for the turbine blade is exceeded, so that an air cooling blade having a hollow structure for flowing cooling air is used. .
このような中空構造の空冷翼は、一般に複雑な内部流路
を有している。第１図は従来使用されているガスタービ
ン空冷翼のキャンバ線における断面を模式的に示す図、
第２図は第１図のＡ−Ａ断面図である。図示されるよう
に、翼根部１から翼高さ方向流路２内に導入された冷却
用空気（図中矢印で示す）は、翼有効部３の前縁部４お
よび後縁部５に翼高さ方向にそって各々形成されたフィ
ルム冷却孔列６における複数のフィルム冷却孔６a を通
して翼外に吹出し、そのほとんどが翼外表面を冷却す
る。翼有効部３の中間部７（外表面）もこれによって冷
却される。Such an air cooling blade having a hollow structure generally has a complicated internal flow path. FIG. 1 is a diagram schematically showing a cross section of a conventionally used gas turbine air-cooling blade in a camber line,
FIG. 2 is a sectional view taken along the line AA of FIG. As shown in the figure, the cooling air (indicated by the arrow in the figure) introduced from the blade root portion 1 into the blade height direction flow path 2 is transferred to the leading edge portion 4 and the trailing edge portion 5 of the blade effective portion 3. It blows out of the blade through a plurality of film cooling holes 6a in the film cooling hole row 6 formed along the height direction, and most of them cool the outer surface of the blade. The intermediate portion 7 (outer surface) of the blade effective portion 3 is also cooled by this.
このような構造の空冷翼では、フィルム冷却孔６a を通
して冷却用空気が翼外に吹き出して、翼高さ方向流路２
内を冷却用空気が流通することにより、翼全体がその内
部から対流冷却される。さらに、、タービン入口温度が
より高い場合には、翼有効部３の背側８および腹側９
に、その高さ方向にそってフィルム冷却孔列１０を各々
形成し、各列１０における複数のフィルム冷却孔１０a
に基づくフィルム冷却を併用することによって、空冷翼
の冷却能を一段と向上させている。In the air-cooling blade having such a structure, the cooling air is blown out of the blade through the film cooling hole 6a, and the blade height direction flow path 2
By circulating the cooling air inside, the entire blade is convectively cooled from the inside. Further, when the turbine inlet temperature is higher, the back side 8 and the ventral side 9 of the blade effective portion 3 are
And a plurality of film cooling hole rows 10 are formed along the height direction, and a plurality of film cooling holes 10a in each row 10 are formed.
The cooling capacity of the air cooling blade is further improved by using the film cooling based on the above.
しかしながら、このように翼高さ方向に冷却用空気の流
路を持つ従来の冷却翼においては、以下のような問題が
あった。すなわち、翼根部１から翼高さ方向流路２内に
導入された冷却用空気は、翼高さ方向流路２内を高さ方
向にそって流動する間に、翼根部１に近い方のフィルム
冷却孔６a 、１０a から徐々に翼外に排出される。その
ため、翼高さ方向流路２内の冷却用空気量は、翼先端部
１１付近に至ると非常に少なくなる。その結果、対流冷
却の結果を定める、翼内部の熱伝達率αＣは、第３図の
点線１２に示すように、翼先端付近では、ほとんど０に
近くなってしまう（ただし第３図中、横軸は、第１図に
示すように翼有効部３の下端部１３から翼先端部１１に
向って、翼高さ方向にそってとった座標h であり、翼有
効部３における翼高さ方向流路２の長さをＨとしてい
る）。したがって、翼先端部１１付近が翼の中で最も温
度の高い部分となる傾向があり、そのため、この部分の
温度が空冷翼に使用される耐熱金属の使用限界温度を越
えてしまうおそれがあるので、タービン入口温度を上げ
ることが困難であった。However, the conventional cooling blade having the cooling air flow path in the blade height direction has the following problems. That is, the cooling air introduced from the blade root portion 1 into the blade height direction flow passage 2 flows through the blade height direction flow passage 2 along the height direction, and The film cooling holes 6a and 10a are gradually discharged to the outside of the blade. Therefore, the amount of cooling air in the blade height direction flow passage 2 becomes extremely small when it reaches the vicinity of the blade tip 11. As a result, the heat transfer coefficient α C inside the blade, which determines the result of convection cooling, becomes almost 0 near the blade tip as shown by the dotted line 12 in FIG. 3 (however, in FIG. 3, The horizontal axis is the coordinate h taken along the blade height direction from the lower end portion 13 of the blade effective portion 3 toward the blade tip portion 11 as shown in FIG. The length of the directional channel 2 is H). Therefore, the vicinity of the blade tip portion 11 tends to be the highest temperature portion of the blade, and therefore the temperature of this portion may exceed the operating temperature limit of the refractory metal used for the air cooling blade. It was difficult to raise the turbine inlet temperature.
［発明の目的］ 本発明は、翼高さ方向流路を有するガスタービン空冷翼
における以上のような問題を解消すべくなされたもの
で、翼先端部付近の冷却能を向上させたカスタービン空
冷翼を提供することを目的としている。[Object of the Invention] The present invention has been made to solve the above problems in a gas turbine air-cooling blade having a blade height direction flow path, and a gas turbine air-cooling system with improved cooling performance near the blade tip portion. Intended to provide wings.
［発明の概要］ 本発明は、翼有効部に、翼表面に貫通するようにして形
成した複数のフィルム冷却孔からなる複数のフィルム冷
却孔列を有し、前記フィルム冷却孔列における前記複数
のフィルム冷却孔が、翼有効部の高さ方向にそってその
全体にわたってほぼ均等に配されているガスタービン空
冷翼において、前記翼有効部の先端部に、補助フィルム
冷却孔からなる少なくとも１つの補助フィルム冷却孔列
を前記フィルム冷却孔列よりも密度を高くして形成した
ことを特徴とするガスタービン空冷翼である。SUMMARY OF THE INVENTION The present invention has, in the blade effective portion, a plurality of film cooling hole rows formed of a plurality of film cooling holes formed so as to penetrate the blade surface, and the plurality of film cooling hole rows include the plurality of film cooling hole rows. In a gas turbine air-cooling blade in which film cooling holes are arranged substantially evenly along the height direction of the blade effective portion, at least one auxiliary film cooling hole having an auxiliary film cooling hole is provided at the tip of the blade effective portion. A gas turbine air-cooling blade is characterized in that a film cooling hole array is formed with a higher density than the film cooling hole array.
［発明の実施例］ 第４図は本発明にかかるガスタービン空冷翼のキャンバ
線における断面を模式的に示す図、第５図は第４図のＢ
−Ｂ断面図である。図中第１図および第２図と同一部分
は同一符号で示す。図示されるように、空冷翼は、高さ
方向に長い中空の３つの翼高さ方向流路２を有してい
る。翼有効部３の前縁部４および後縁部５には、複数の
フィルム冷却孔６a からなるフィルム冷却孔列６が１列
ずつ、背側８および腹側９には、複数のフィルム冷却孔
１０a からなるフィルム冷却孔列１０が各々１列ずつ形
成されている。各列における複数のフィルム冷却孔６a
、１０a は、翼有効部３の高さ方向にそってその全体
にわたって配置されている。[Embodiment of the Invention] FIG. 4 is a diagram schematically showing a cross section of a gas turbine air-cooling blade according to the present invention taken along a camber line, and FIG.
It is a -B sectional view. In the figure, the same parts as those in FIGS. 1 and 2 are designated by the same reference numerals. As shown in the figure, the air cooling blade has three hollow blade height direction flow passages 2 that are long in the height direction. The front edge portion 4 and the rear edge portion 5 of the blade effective portion 3 are each provided with a row of film cooling holes 6 composed of a plurality of film cooling holes 6a, and a back side 8 and a vent side 9 are provided with a plurality of film cooling holes. One row 10 of film cooling holes 10a is formed for each row. Multiple film cooling holes 6a in each row
10a are arranged along the height direction of the blade effective portion 3 over the entire area.
翼有効部３の先端部には、各々複数の補助フィルム冷却
孔１４a からなる４列の補助フィルム冷却孔列１４が形
成されている。各列の複数の補助フィルム冷却孔１４a
は、翼有効部３の高さ方向にそって配されている。ま
た、補助フィルム冷却孔列１４は、翼有効部３の背側８
であって後縁部５寄りに１列、そして腹側９であって前
縁部４寄り、中間部および後縁部５寄りに各々１列形成
されている。Four rows of auxiliary film cooling holes 14 each including a plurality of auxiliary film cooling holes 14a are formed at the tip of the blade effective portion 3. A plurality of auxiliary film cooling holes 14a in each row
Are arranged along the height direction of the blade effective portion 3. Further, the auxiliary film cooling hole array 14 is provided on the back side 8 of the blade effective portion 3.
One row is formed near the rear edge portion 5, one row is formed on the ventral side 9 near the front edge portion 4, and one row is formed near the middle portion and the rear edge portion 5.
したがって、以上のような構成によって翼根部１から各
翼高さ方向流路２内に導入された冷却用空気は、翼有効
部３の先端部付近で翼外に比較的多く吹き出される。す
なわち、翼有効部３の先端部には、フィルム冷却孔６a
、１０a の他に、補助フィルム冷却孔１４a があるの
で、そこからも冷却用空気が翼外に吹き出されるからで
ある。その結果、翼内部の熱伝達率αＣは、第３図の実
線１５に示すように、翼先端部付近でも低下しない。し
たがってその高さ方向に比較的均一に空冷翼全体が冷却
される。なお、翼金属は、補助フィルム冷却孔１４a 内
の冷却用空気の対流冷却によっても冷却される。Therefore, the cooling air introduced from the blade root portion 1 into each blade height direction flow path 2 by the above configuration is blown out relatively large outside the blade near the tip of the blade effective portion 3. That is, the film cooling hole 6a is provided at the tip of the blade effective portion 3.
This is because the auxiliary film cooling hole 14a is provided in addition to 10a, and the cooling air is also blown out of the blade from there. As a result, the heat transfer coefficient α C inside the blade does not decrease even near the tip of the blade, as shown by the solid line 15 in FIG. Therefore, the entire air cooling blade is cooled relatively uniformly in the height direction. The blade metal is also cooled by convection cooling of the cooling air in the auxiliary film cooling hole 14a.
さらに、第４図、第５図に示す構造の空冷翼において、
相対的に前縁部４に近い側の翼有効部３の先端部に、複
数の補助フィルム冷却孔を付加することによってフィル
ム冷却が翼先端部に限って強化される。そのため、空冷
翼の冷却能が一層均一化、強化される。Furthermore, in the air cooling blades having the structures shown in FIGS. 4 and 5,
By adding a plurality of auxiliary film cooling holes to the tip portion of the blade effective portion 3 on the side relatively close to the leading edge portion 4, the film cooling is enhanced only in the blade tip portion. Therefore, the cooling ability of the air cooling blade is further uniformed and enhanced.
［発明の効果］ 以上説明したように、本発明によれば、翼先端部の冷却
能を向上させることができ、その結果、一段とタービン
入口温度を上げてガスタービンの熱効率を向上させるこ
とができる。[Effects of the Invention] As described above, according to the present invention, it is possible to improve the cooling performance of the blade tip portion, and as a result, it is possible to further increase the turbine inlet temperature and improve the thermal efficiency of the gas turbine. .
第１図は従来使用されているガスタービン空冷翼のキャ
ンバ線における断面を模式的に示す図、第２図は第１図
のＡ−Ａ断面図、第３図は翼高さ方向流路内部の熱伝達
率αＣの翼高さ方向における分布を示す図、第４図は本
発明にかかるガスタービン空冷翼のキャンバ線における
断面を模式的に示す図、第５図は第４図のＢ−Ｂ断面図
である。 ３……翼有効部、６、１０……フィルム冷却孔列、６a
、１０a ……フィルム冷却孔、１４……補助フィルム
冷却孔列、１４a ……補助フィルム冷却孔FIG. 1 is a diagram schematically showing a cross section of a conventionally used gas turbine air-cooling blade taken along a camber line, FIG. 2 is a sectional view taken along line AA of FIG. 1, and FIG. Showing the distribution of the heat transfer coefficient α C in the blade height direction, FIG. 4 is a view schematically showing the cross section of the gas turbine air cooling blade according to the present invention at the camber line, and FIG. 5 is B of FIG. It is a -B sectional view. 3 ... Blade effective part, 6, 10 ... Film cooling hole array, 6a
10a ... Film cooling holes, 14 ... Auxiliary film cooling hole array, 14a ... Auxiliary film cooling holes
Claims (1)
形成した複数のフィルム冷却孔からなる複数のフィルム
冷却孔列を有し、前記フィルム冷却孔列における前記複
数のフィルム冷却孔が、翼有効部の高さ方向にそってそ
の全体にわたってほぼ均等に配されているガスタービン
空冷翼において、前記翼有効部の先端部に、補助フィル
ム冷却孔からなる少なくとも１つの補助フィルム冷却孔
列を前記フィルム冷却孔列よりも密度を高くして形成し
たことを特徴とするガスタービン空冷翼。1. A blade effective portion has a plurality of film cooling hole rows formed of a plurality of film cooling holes formed so as to penetrate the blade surface, wherein the plurality of film cooling holes in the film cooling hole row are In a gas turbine air-cooling blade arranged substantially evenly along the height direction of the blade effective portion, at least one auxiliary film cooling hole row consisting of auxiliary film cooling holes is provided at the tip of the blade effective portion. Is formed with a higher density than the row of film cooling holes.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP58091557A JPH0660562B2 (en) | 1983-05-26 | 1983-05-26 | Gas turbine air cooling blade |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP58091557A JPH0660562B2 (en) | 1983-05-26 | 1983-05-26 | Gas turbine air cooling blade |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| JPS59218302A JPS59218302A (en) | 1984-12-08 |
| JPH0660562B2 true JPH0660562B2 (en) | 1994-08-10 |
Family
ID=14029806
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| JP58091557A Expired - Lifetime JPH0660562B2 (en) | 1983-05-26 | 1983-05-26 | Gas turbine air cooling blade |
Country Status (1)
| Country | Link |
|---|---|
| JP (1) | JPH0660562B2 (en) |
Families Citing this family (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JP2615823B2 (en) * | 1988-04-28 | 1997-06-04 | 松下電器産業株式会社 | Method for manufacturing blade type blade of impeller |
| DE102006042647A1 (en) | 2006-09-12 | 2008-03-27 | Mtu Aero Engines Gmbh | Turbine of a gas turbine |
Citations (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US2851216A (en) | 1954-01-13 | 1958-09-09 | Schwarzkopf Dev Co | Device adapted for respiration cooling and process of making same |
-
1983
- 1983-05-26 JP JP58091557A patent/JPH0660562B2/en not_active Expired - Lifetime
Patent Citations (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US2851216A (en) | 1954-01-13 | 1958-09-09 | Schwarzkopf Dev Co | Device adapted for respiration cooling and process of making same |
Also Published As
| Publication number | Publication date |
|---|---|
| JPS59218302A (en) | 1984-12-08 |
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