JPS5918204A - Blade of gas turbine - Google Patents
Blade of gas turbineInfo
- Publication number
- JPS5918204A JPS5918204A JP12561682A JP12561682A JPS5918204A JP S5918204 A JPS5918204 A JP S5918204A JP 12561682 A JP12561682 A JP 12561682A JP 12561682 A JP12561682 A JP 12561682A JP S5918204 A JPS5918204 A JP S5918204A
- Authority
- JP
- Japan
- Prior art keywords
- blade
- main body
- flow channel
- blade main
- passage
- 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
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
- 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
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
Description
【発明の詳細な説明】
〔発明の技術分野〕
本発明は、ガスタービンの翼に係り、特に、冷却構造を
改良した翼に関する。DETAILED DESCRIPTION OF THE INVENTION [Technical Field of the Invention] The present invention relates to a gas turbine blade, and particularly to a blade with an improved cooling structure.
一般的にガスタービンは往復機関に比較して小型軽量で
犬馬力が得られるなどの多くの利点、を有している。こ
のようなガスタービンハ、 鴻常、１つの軸に圧縮機と
７４ワータービンとを連−シ、圧縮機で圧縮された高圧
空気で燃焼器内自圧力を高め、この状態で燃焼器内に燃
料を噴軸して燃焼させ、この燃焼によって生じた高温。In general, gas turbines have many advantages over reciprocating engines, such as being smaller, lighter, and able to provide a higher horsepower. This kind of gas turbine has a compressor and a 74-hour turbine connected to one shaft, and the high-pressure air compressed by the compressor increases the internal pressure in the combustor, and in this state, the pressure inside the combustor is increased. Fuel is combusted through a jet shaft, and the high temperature generated by this combustion.
高圧のガスをパワータービンに導いて膨張させることに
よシ回転動力を得るように構成されている。圧縮機は、
通常、案内翼と回転翼とを軸方向に配列した軸流型に構
成され、また、パワータービンも動翼と静翼とを軸方向
に交互に配列して構成されている。It is configured to obtain rotational power by guiding high-pressure gas to a power turbine and expanding it. The compressor is
Generally, the turbine is constructed as an axial flow type in which guide vanes and rotor vanes are arranged in the axial direction, and power turbines are also constructed in such a manner that rotor blades and stator vanes 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、出力効
率を高めるには、ノヤワータービンのしている。By the way, in the above-mentioned gas turbine, the Noyawer turbine is used to increase the output efficiency.
上記のように構成されたタービンの翼にあっては確カー
に翼本体１の前縁部Ａおよび後縁部Ｂを冷却することが
できる。In the turbine blade configured as described above, the leading edge A and the trailing edge B of the blade body 1 can be cooled efficiently.
しかしながら、後縁部Ｂを冷却するだめの第２０流路１
２は、内部に多数のビンフィン１３を設置したものとな
っているので、この第２０流路１２の構造が複雑となシ
、この結果、鋳造時の型の製作が困難で製作に長時間を
要する問題があった。また、ビンフィン１３による冷却
効果を増すためには、ビンフィン１３の数を増す必要が
あるので、必然的に翼全体の重量が増加し、特に、動翼
に適用した場合には遠心応力に耐えさせるために各部の
肉厚を増す必要があり、この結果重量がさらに増加する
問題があった。However, the 20th flow path 1 for cooling the trailing edge portion B
2 has a large number of bottle fins 13 installed inside, so the structure of this 20th channel 12 is complicated, and as a result, it is difficult to make a mold for casting and it takes a long time to make. There was a problem. In addition, in order to increase the cooling effect of the bin fins 13, it is necessary to increase the number of bin fins 13, which inevitably increases the weight of the entire blade, and especially when applied to a rotor blade, it is necessary to increase the number of bin fins 13. Therefore, it was necessary to increase the thickness of each part, which resulted in a further increase in weight.
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、特に、翼本体の重量を増すこと
なく、また、製作の困難化を招くことなく翼本体の後縁
部を良好に冷却で□きるガスタービンの翼を提供するこ
とにある。The present invention has been made in view of the above circumstances, and its purpose is to improve the trailing edge of the wing body without increasing the weight of the wing body or complicating manufacturing. The objective is to provide gas turbine blades that can be cooled well.
、〔発明の概要〕
本発明に係るガスタービンの翼は、翼本体の向へ複数設
けたことを特徴としている。, [Summary of the Invention] The gas turbine blade according to the present invention is characterized in that a plurality of blades are provided toward the blade body.
上記構成であると、第２０流路は上記突条の存在によっ
て内表面積が十分に拡大化される。With the above configuration, the inner surface area of the 20th flow path is sufficiently expanded due to the presence of the ridges.
したがって、熱交換面積が十分広くなシ、上記第２０流
路内を通流する冷却流体によって翼本体の後縁部が良好
に冷却されることになる。また、上記突条を翼本体の高
さ方向、つまシ、第２の流路内を通流する冷却流体の通
流方向と＃１ぼ直交する関係に設けているので、上記突
条によって通流する冷却流体を積極的に攪拌させること
ができ、これによって、さらに後縁部を良好に冷却する
ことができる。また、突条を上記関係に設けているので
、いわゆる高さの低い突条であっても良好な攪拌機能を
発揮させることができる。したがって、突条の高さを増
さなくきる。また、第１の流路と第２の流路とを仕切る
仕切壁に設けられる小孔をその軸心線がキャンバ線の接
線に対して傾くように設けると前記突条による攪拌機能
をさらに増進させることができる。Therefore, since the heat exchange area is sufficiently large, the trailing edge portion of the blade body can be cooled well by the cooling fluid flowing through the 20th flow path. In addition, since the above-mentioned protrusion is provided in a relationship that is approximately perpendicular to the height direction of the blade body, the blade, and the flow direction of the cooling fluid flowing through the second flow path, the above-mentioned protrusion allows the passage of water The flowing cooling fluid can be actively stirred, and thereby the trailing edge can be further cooled well. Further, since the protrusions are provided in the above relationship, even the so-called low protrusions can exhibit a good stirring function. Therefore, the height of the protrusion can be removed without increasing it. Furthermore, if the small hole provided in the partition wall that partitions the first flow path and the second flow path is provided so that its axis line is inclined with respect to the tangent line of the camber line, the stirring function by the protrusion can be further enhanced. can be done.
第４図は本発明の一実施例に係る翼をキャンバ線に沿っ
て切断して示す図であり、第２図と同一部分は同一符号
で示しである。したがって、重複する部分の説明は省略
する。FIG. 4 is a view showing a wing according to an embodiment of the present invention cut along a camber line, and the same parts as in FIG. 2 are designated by the same reference numerals. Therefore, the explanation of the overlapping parts will be omitted.
この実施例においては、翼本体１の後縁部Ｂ内にコード
方向に設けられる第２の流路１２の内面で、かつ翼本体
１の腹側および背側に位置する側内面に第５図にも示す
ように、翼本体１の高さ方向に延びる突条２１を各面相
互で位相が異なる関係にコード方向に３本ずつ突設し、
体は仕切壁１６に設けられた小孔２２を介して第２０流
路１２内へ流れ込むことになる。第２０流路１２内へ噴
流となって流れ込んだ冷却流体は小孔２２が前記関係に
設けられているので第２０流路１２の内面に衝突し、さ
らに突条２１によって十分攪拌されながら上記第２０流
路１２内をコード方向に流れ、この間に後縁部Ｂから熱
を奪う。そして最終的に後縁部Ｂの後縁端に設けられた
開口から真性へと流れる。In this embodiment, the inner surface of the second flow path 12 provided in the trailing edge B of the wing body 1 in the chord direction, and the side inner surfaces located on the ventral side and the dorsal side of the wing body 1, as shown in FIG. As shown in FIG. 2, three protrusions 21 extending in the height direction of the wing body 1 are provided protruding in the chord direction with different phases on each surface.
The body flows into the twentieth channel 12 through the small hole 22 provided in the partition wall 16. The cooling fluid flowing into the twentieth flow path 12 as a jet collides with the inner surface of the twentieth flow path 12 because the small holes 22 are provided in the above-mentioned relationship, and is further stirred sufficiently by the ridges 21. 20 flows in the chord direction in the flow path 12, and during this time, heat is taken away from the trailing edge B. Finally, it flows from the opening provided at the trailing edge end of the trailing edge portion B to the essence.
そして、この場合には、第２の流路１２の内面に前記関
係に突条２１を設けたことが有効に作用し、十分に広い
熱交換面に十分に攪拌された冷却流体が接触することに
なυ、後縁部Ｂが良好に冷却されることになる。また、
突条２１を冷却流体の通流方向にほぼ直交するように設
けたことと突条２１と云った単純形状のものを用いてい
ることとが相俟って翼本体の大重量化を招くことなく、
シかも製作の困難化を招くことなく良好な冷却特性を発
揮させることができるので、結局、前述した効果が得ら
れることに重複する部分の説明は省略する。In this case, providing the protrusions 21 in the above-mentioned relationship on the inner surface of the second flow path 12 effectively works to ensure that the sufficiently agitated cooling fluid comes into contact with a sufficiently wide heat exchange surface. In other words, the trailing edge B is well cooled. Also,
Providing the protrusions 21 so as to be substantially orthogonal to the flow direction of the cooling fluid and using a simple-shaped protrusion 21 combine to increase the weight of the blade body. Without,
Since it is possible to exhibit good cooling characteristics without complicating the manufacturing process, the explanation of the parts that overlap with the above-mentioned effects will be omitted.
この実施例においては第１の流路を仕切壁２３によって
流路ＩＩ＠とｌｌｂとに区分し、小孔１５から噴出する
冷却流体の流量と、第２の流路１２内を流れる冷却流体
の流量とをそれぞれ独立して調整できるようにしている
。In this embodiment, the first flow path is divided into flow paths II@ and Ilb by a partition wall 23, and the flow rate of the cooling fluid ejected from the small hole 15 and the flow rate of the cooling fluid flowing inside the second flow path 12 are controlled. The flow rate can be adjusted independently.
このように構成しても前記実施例と同様な効果が得られ
ることは勿論である。Of course, even with this configuration, the same effects as in the above embodiment can be obtained.
なお、各実施例において、翼本体の腹側外面および背側
外面に沿って冷却流体をフィルム状に排出させるフィル
ム冷却用の孔を設けるようにしてもよい。In each embodiment, film cooling holes may be provided along the ventral outer surface and the dorsal outer surface of the wing body to discharge the cooling fluid in the form of a film.
第１図は冷却手段を施した従来のガスタービンの動翼の
斜視図、第２図は第１図におけるＰ−Ｐ線に沿って切断
し矢印方向にみた断面図、がスタービンの翼を第３図に
対応させて示す横−断面図、第６図は本発明の他の実施
例に係る翼を第４図に対応させて示す縦断面図、第７図
は門真を第５図に対応させて示す横断面図である。
１・・・翼本体、２・・・翼根部、１ノ・・・第１０流
路、１２・・・第２の流路、１６・・・仕切壁、２１・
・・突条、２２・・・小孔。
出願人 工業技術院長 石 坂 誠 −第１図
第２図
１１３図
窄４図
第５１１
ｑｐ ６閃
慎７図
１９−Figure 1 is a perspective view of the rotor blades of a conventional gas turbine equipped with a cooling means, and Figure 2 is a sectional view taken along line P-P in Figure 1 and viewed in the direction of the arrow. 3 is a transverse sectional view corresponding to FIG. 3, FIG. 6 is a vertical sectional view showing a blade according to another embodiment of the present invention corresponding to FIG. 4, and FIG. 7 is a Kadoma shown in FIG. FIG. DESCRIPTION OF SYMBOLS 1...Blade body, 2...Blade root, 1...10th channel, 12...2nd channel, 16...Partition wall, 21...
... Protrusion, 22... Small hole. Applicant Makoto Ishizaka, Director of the Agency of Industrial Science and Technology - Figure 1 Figure 2 Figure 2 113 Figure 511 qp 6 Senshin 7 Figure 19-
Claims (2)
流路を設けるとともに上記翼本体の後縁部内圧一端側が
後縁端に開口する関係にコード方向に延びる第２０流路
を設け、上記第１の餌路に導かれた冷却流体を、上記第
１の流路と前記第２の流路との間を仕切る仕切壁に設け
られた複数の小孔を介して上記第２の流路に流入させる
ようにした冷却路を備えたガスタービンの翼において、
前記第２の流路は、前記翼本体の腹側および背側に位置
する内面に翼本体の高さ方向に延びる突条を複数コード
方向に有してなることを特徴とするガスタービンの翼。(1) A first flow path is provided in the blade body along the height direction of the blade body, and a 20th flow path extends in the chord direction such that one end side of the internal pressure of the trailing edge of the blade body opens to the trailing edge end. A passage is provided, and the cooling fluid guided to the first bait passage is passed through a plurality of small holes provided in a partition wall that partitions between the first passage and the second passage. In a gas turbine blade equipped with a cooling passage configured to flow into a second flow passage,
The second flow path has a plurality of protrusions extending in the height direction of the blade body on the inner surfaces located on the ventral side and the dorsal side of the blade body in the cord direction. .
が翼型のキャンバ線の接線に対して傾いて設けられてい
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
ービンの翼。(2) The gas according to claim 1, wherein the small hole provided in the partition wall is provided with an axial center line inclined with respect to a tangent to a camber line of the airfoil. turbine blade.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP12561682A JPS5918204A (en) | 1982-07-21 | 1982-07-21 | Blade of gas turbine |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP12561682A JPS5918204A (en) | 1982-07-21 | 1982-07-21 | Blade of gas turbine |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| JPS5918204A true JPS5918204A (en) | 1984-01-30 |
Family
ID=14914485
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| JP12561682A Pending JPS5918204A (en) | 1982-07-21 | 1982-07-21 | Blade of gas turbine |
Country Status (1)
| Country | Link |
|---|---|
| JP (1) | JPS5918204A (en) |
Cited By (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JPH03141801A (en) * | 1990-09-19 | 1991-06-17 | Hitachi Ltd | Cooling blade of gas turbine |
| EP1035302A3 (en) * | 1999-03-05 | 2002-02-06 | General Electric Company | Multiple impingement airfoil cooling |
| GB2460936A (en) * | 2008-06-18 | 2009-12-23 | Gen Electric | Turbine airfoil cooling |
| JP2018112187A (en) * | 2017-01-10 | 2018-07-19 | ドゥサン ヘヴィー インダストリーズ アンド コンストラクション カンパニー リミテッド | Blade, cut-back of blade or vane, and gas turbine having the same |
-
1982
- 1982-07-21 JP JP12561682A patent/JPS5918204A/en active Pending
Cited By (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JPH03141801A (en) * | 1990-09-19 | 1991-06-17 | Hitachi Ltd | Cooling blade of gas turbine |
| EP1035302A3 (en) * | 1999-03-05 | 2002-02-06 | General Electric Company | Multiple impingement airfoil cooling |
| GB2460936A (en) * | 2008-06-18 | 2009-12-23 | Gen Electric | Turbine airfoil cooling |
| US8210814B2 (en) | 2008-06-18 | 2012-07-03 | General Electric Company | Crossflow turbine airfoil |
| GB2460936B (en) * | 2008-06-18 | 2012-11-28 | Gen Electric | Crossflow turbine airfoil |
| JP2018112187A (en) * | 2017-01-10 | 2018-07-19 | ドゥサン ヘヴィー インダストリーズ アンド コンストラクション カンパニー リミテッド | Blade, cut-back of blade or vane, and gas turbine having the same |
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