JPS61205301A - Gas turbine blade - Google Patents
Gas turbine bladeInfo
- Publication number
- JPS61205301A JPS61205301A JP60042678A JP4267885A JPS61205301A JP S61205301 A JPS61205301 A JP S61205301A JP 60042678 A JP60042678 A JP 60042678A JP 4267885 A JP4267885 A JP 4267885A JP S61205301 A JPS61205301 A JP S61205301A
- Authority
- JP
- Japan
- Prior art keywords
- blade
- cooling
- gas turbine
- turbine blade
- cooling fluid
- 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
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】
〔発明の利用分野〕
本発明は内部に冷却構造をもつガスタービン翼に係シ、
特に冷却効果にすぐれた冷却流体の通路を備えたガスタ
ービン翼に関するものでおる。[Detailed Description of the Invention] [Field of Application of the Invention] The present invention relates to a gas turbine blade having an internal cooling structure.
In particular, the present invention relates to a gas turbine blade equipped with a cooling fluid passage having an excellent cooling effect.
カスタービンの翼は高温の状態で作動し、翼に使用する
材料の許容温度範囲内で用いるために翼の冷却か必要を
なっている。その際に要求されることは、少ない冷却流
体の流量で翼の最高温度および平均温度を翼の材料の許
容温度範囲内に保つこと、興の内部の温度差を小さくす
ること％かめる。Cast turbine blades operate at high temperatures and require cooling to stay within the acceptable temperature range of the materials used in the blades. What is required in this case is to maintain the maximum and average temperatures of the blade within the permissible temperature range of the blade material with a small flow rate of cooling fluid, and to reduce the temperature difference inside the blade.
今日ガスタービン翼の冷却方法として採用されている中
で、インピンジメント冷却は冷却流体を絞って被冷却面
に衝突させるため平均的な冷却効果は大きいものの冷却
流路全体の圧力損失も大きく、更に衝突流の中心と周囲
の被冷却面上で冷却効果かかなシネ均一になる欠点があ
る。またフィルム冷却は翼外表面に開口した吹出し孔が
燃焼ガス中の不純物により閉１されることが、１、使用
する燃料によっては冷却効果が著しく損われる欠点かあ
る。それに対して対流冷却は翼内部の冷却流路に沿って
冷却流体を流すことによって翼の冷却をはかるものでフ
ィルム冷却のような翼外部への吹出し孔かないため燃料
の性質によシ冷却効果か低下することなく、また翼の製
造技術の進歩によシ微細な内部冷却流路を製造すること
が可能になったためインビジ、メント冷却よりも小さな
圧力損失で大きな冷却効果を得られることから、多様な
燃料の使用に対応できる対流冷却を基本とした内部冷却
構造を有する翼の開発が進められている。Impingement cooling, which is used today as a cooling method for gas turbine blades, squeezes the cooling fluid and causes it to collide with the surface to be cooled, so although the average cooling effect is large, the pressure loss in the entire cooling flow path is also large. There is a drawback that the cooling effect is rather uniform at the center of the impinging flow and on the surrounding surface to be cooled. In addition, film cooling has the drawbacks that the blow-off holes opened on the outer surface of the blade are blocked by impurities in the combustion gas, and the cooling effect may be significantly impaired depending on the fuel used. On the other hand, convection cooling cools the blade by flowing cooling fluid along the cooling channel inside the blade, and since there are no blow holes to the outside of the blade like in film cooling, the cooling effect may vary depending on the nature of the fuel. Furthermore, advancements in blade manufacturing technology have made it possible to manufacture fine internal cooling channels, so it is possible to obtain a large cooling effect with smaller pressure loss than in-visible cooling. Development of blades with an internal cooling structure based on convection cooling that can handle the use of various types of fuel is underway.
対流冷却によってタービン挑を冷却するモノトして特開
昭５８−１１９９０２号公報があげられる。Japanese Unexamined Patent Publication No. 119902/1983 is an example of cooling a turbine through convection cooling.
これは冷却通路に沿って冷却流体の温度が上昇し、それ
Ｋよって冷却効果が下がることから冷却流路出口の断面
積を入口よりも小さくすることによシ冷却流路出ロ付近
の流速を高め、冷却効果が下がるのを防いでいる。これ
を第９図、第１０図によって説明する。第９図はガスタ
ービン翼の横断面図、第１０図はその縦断面図を示す。This is because the temperature of the cooling fluid increases along the cooling passage, which reduces the cooling effect, so by making the cross-sectional area of the cooling passage outlet smaller than the inlet, the flow velocity near the cooling passage outlet can be reduced. This prevents the cooling effect from decreasing. This will be explained with reference to FIGS. 9 and 10. FIG. 9 shows a cross-sectional view of the gas turbine blade, and FIG. 10 shows a longitudinal cross-sectional view thereof.
第１０図において、冷却流路２２は翼状のけた材２１の
周囲にある薄板状の殻材２５の間で翼外表面付近に配置
されている。冷却流体は冷却流路入口２４から加圧中空
内部２６に送られ、翼根元部２Ｂ付近でけた材２１を貫
通する孔２７から冷却流路２２へ導びかれている。この
ような冷却流路を設けることにより、比較的長さの短い
冷却流路に高速で冷却流体を送る九め、限られた圧力損
失・流量で効果的な対流冷却か可能となシ、更に冷却流
路出口付近の断面積を小さくして冷却流路出口付近の流
速を上げているため、冷却流路に沿って温度上昇する冷
却流体による冷却効果が下がるのを防ぐことはできる。In FIG. 10, the cooling channels 22 are arranged near the outer surface of the blade between thin plate-like shell members 25 around the blade-like spar members 21. In FIG. The cooling fluid is sent from the cooling passage inlet 24 to the pressurized hollow interior 26, and is led to the cooling passage 22 from a hole 27 penetrating the spar 21 near the blade root portion 2B. By providing such a cooling channel, it is possible to send cooling fluid at high speed through a relatively short cooling channel, and effective convection cooling is possible with limited pressure loss and flow rate. Since the cross-sectional area near the exit of the cooling channel is reduced to increase the flow velocity near the exit of the cooling channel, it is possible to prevent the cooling effect of the cooling fluid whose temperature increases along the cooling channel from decreasing.
しかしながら、ガスタービン翼の冷却流路の出口付近の
温度が最も高いわけではなく第４図に示したように真中
央部の温度が高く、翼端部の万が低い状態にある。第４
図においてガス温度の具体的な数値をあげれば真中央部
のＴ、、！は１３９０Ｃ翼の平均温度Ｔ ａ ｖは１２
６０Ｃ，翼端部のＴム。However, the temperature near the outlet of the cooling flow path of the gas turbine blade is not the highest; as shown in FIG. 4, the temperature is high at the center, and the temperature at the blade tip is low. Fourth
In the figure, if we give a specific numerical value for the gas temperature, it is T at the true center...! is 1390C The average temperature of the blade T a v is 12
60C, Tum of the wing tip.
Ｔｃは１０００Ｃでめる。このような翼内部の温度分布
を考慮していなかった従来のガスタービンの冷却翼では
、翼の温度の不均一を解消することはできず、翼の最高
温度を材料の許容温度以下にするためには多量の冷却流
体を必要としなければならなかつ九。更に、翼の最高温
度と最低温度の温度差が大きいと材料の熱ひずみか大き
くなり、材料の強度上の問題があった。Tc is measured at 1000C. Conventional cooling blades for gas turbines do not take this kind of temperature distribution inside the blade into consideration, and it is not possible to eliminate the uneven temperature of the blade. (9) It requires a large amount of cooling fluid. Furthermore, if the temperature difference between the maximum and minimum temperatures of the blade is large, the thermal strain of the material will be large, which poses problems in terms of the strength of the material.
本発明の目的は、ガスタービン翼の最高温度を材料の許
容温度以下にするとともに、少ない流量で真の温度の不
均一を解消することのできる冷却効果のすぐれた内部冷
却構造を備え九ガス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 nine gas turbine equipped with an internal cooling structure that has an excellent cooling effect that can lower the maximum temperature of a gas turbine blade to below the permissible temperature of the material, and eliminate true temperature non-uniformity with a small flow rate. It's about providing wings.
本発明の特徴とするところは、ガスタービン翼の内部に
冷却流体を流す冷却流路と、竺却痺体を翼端部から冷却
流路へ送る導入流路と冷却流路からの冷却流体を真端部
または翼外部へ流出させる排出流路を備え、冷却流路の
側面金属の外殻と隔壁で構成されたガスタービン翼にお
いて、翼の外殻の外表面から内面に至る伝熱通路の熱伝
達抵抗をほは一定にして、かつ翼の中央部の冷却流路断
面積を翼の根元部および翼の先端部よりも小さくしたこ
とにある。The features of the present invention include a cooling channel that allows cooling fluid to flow inside the gas turbine blade, an introduction channel that sends the cooling body from the blade tip to the cooling channel, and a cooling fluid that flows from the cooling channel. In a gas turbine blade that is equipped with a discharge flow path that flows out to the true end or outside of the blade, and is composed of a metal outer shell and a partition wall on the side of the cooling flow path, the heat transfer passage from the outer surface of the outer shell to the inner surface of the blade is The heat transfer resistance is kept fairly constant, and the cross-sectional area of the cooling flow path in the center of the blade is smaller than that in the root and tip of the blade.
以上のような構造にしたことによシ、冷却流路の断面積
が翼の根元部では大きく、翼の中央部で小さくなり、翼
の先端部で再び大きくなる。よって冷却流体の流速およ
び流速とほぼ比例関係をもつ熱伝達率は興の根元部で小
さく、舅の中央部で大きく、翼の中央部で大きくなり、
異の先端部で小さくなる。このようにして翼の中央部付
近の冷却効果を高めることができる。その際に翼内面の
伝熱面積や翼の外殻の厚さを変化させることは上述した
効果を弱めることにつながるため冷却流路断面積の変化
は、隔壁を突出させることによって行っている。With the above structure, the cross-sectional area of the cooling flow path is large at the root of the blade, becomes small at the center of the blade, and becomes large again at the tip of the blade. Therefore, the flow velocity of the cooling fluid and the heat transfer coefficient, which has a nearly proportional relationship with the flow velocity, are small at the root of the wing, large at the center of the wing, and large at the center of the wing.
It becomes smaller at the different tip. In this way, the cooling effect near the center of the blade can be enhanced. At this time, changing the heat transfer area on the inner surface of the blade or the thickness of the outer shell of the blade will weaken the above-mentioned effects, so the cross-sectional area of the cooling flow path is changed by protruding the partition wall.
この構造よシ、少ない冷却流体で翼の中央部の温度を下
げることかでき、翼の温度の均一化がはかれる。With this structure, the temperature in the center of the blade can be lowered with less cooling fluid, and the temperature of the blade is more uniform.
本発明の一実施例を第１図および第２図に示す。 An embodiment of the invention is shown in FIGS. 1 and 2.
これはリターンフロ一対流冷却翼に適用した例である。This is an example of application to a return flow convection cooling blade.
ガスタービン翼ｌの断面内には中心線に沿って４本の翼
高さ方向の冷却流路２〜５が配設され、各流路の側面は
翼の外殻による翼内面６と隔壁７で得成されている。冷
却流路２・３へは翼根元部８内部の導入流路９から冷却
流体１０が供給され、冷却流路２では翼先端部１１へ向
って単純に通過し排出孔１２から排出される（第１の系
統）が、冷却流路３では翼先端部１１まで通過後折返し
て冷却流路４・５を順に迎向きに流れ排出流路１３から
排出される（第２の系統）。冷却流路２〜４の隔壁７に
は突出部１４が設けられ、諷高さ中央（Ｂ−Ｂ〜［面）
の流路断面積をｈ根元部付近（Ｃ−Ｃ断面）や真先端部
付近（Ａ−Ａ断面）よりも小さくするとともに、熱交換
を行う翼内面６の表面積は流路断面積変化の影響を受け
ないようにしている。そのため各冷却流路２〜４内の冷
却流体１０の流速は翼の中央部が翼根元部、先端部付近
よりも犬きくなシ、同様に冷却側熱伝達率も翼の中央部
が高くなる。このように冷却流路の断面積変化による冷
却側の熱伝達率への影響は大きく、翼内面６に設置され
た板・柱状の小突起１５によって渦流を発生させて冷却
効果を高めることと併用することで、翼の温度の均一と
冷却効果を更に向上させることができる。従来用いられ
ていた板・柱状の小突起なる慎熱促進簀累たけを用いて
冷却効果を高めようとすることは冷却流体の流れによど
み等が発生したシする九め冷却能力には限界がアシ、ガ
スタービン翼内の温度差が１０００程度までしか縮める
ことかできなかったか、本発明を用い九場合は、ガスタ
ービン翼内の温度差を８０Ｃ以下に抑えることができた
。Four cooling passages 2 to 5 in the blade height direction are arranged along the center line in the cross section of the gas turbine blade l, and the side surfaces of each passage are formed by the blade inner surface 6 and the partition wall 7 formed by the blade outer shell. It has been achieved by Cooling fluid 10 is supplied to the cooling channels 2 and 3 from the introduction channel 9 inside the blade root section 8, and in the cooling channel 2, it simply passes toward the blade tip section 11 and is discharged from the discharge hole 12 ( After passing through the cooling channel 3 to the blade tip 11, the air is turned back, flows in the forward direction through the cooling channels 4 and 5, and is discharged from the discharge channel 13 (second channel). A protrusion 14 is provided on the partition wall 7 of the cooling channels 2 to 4, and a protrusion 14 is provided in the partition wall 7 of the cooling channels 2 to 4.
The cross-sectional area of the flow passage is made smaller than that near the root (C-C section) and near the true tip (A-A cross-section), and the surface area of the inner surface 6 of the blade that performs heat exchange is affected by the change in the cross-sectional area of the flow passage. I try not to get it. Therefore, the flow velocity of the cooling fluid 10 in each of the cooling channels 2 to 4 is faster at the center of the blade than at the root and near the tip, and similarly, the heat transfer coefficient on the cooling side is higher at the center of the blade. . In this way, the change in the cross-sectional area of the cooling channel has a large effect on the heat transfer coefficient on the cooling side, and it is also used in conjunction with generating vortices by the plate/column-shaped small protrusions 15 installed on the inner surface 6 of the blade to increase the cooling effect. By doing so, it is possible to further improve the uniformity of the temperature of the blade and the cooling effect. Attempting to increase the cooling effect by using the conventional heat-reducing heat-promoting cages, which are small protrusions in the form of plates or columns, would result in stagnation in the flow of the cooling fluid, and the cooling capacity would be limited. However, when the present invention was used, the temperature difference within the gas turbine blade could be reduced to 80C or less.
本発明における第２の実施例を第３図、第４図および第
５図に示す。A second embodiment of the present invention is shown in FIGS. 3, 4, and 5.
これは上記の実施例と同様なリターンフロ一対流冷却翼
であるが、隔壁を突出させるかわシに翼中央壁１６を突
出させることで真性表面付近に配設され九冷却流路２〜
４の断面積を変化させている。上記の実施例と同様の原
理によシ、第４図のように不均一なガス温度分布（Ｔ、
、、：最高温度、Ｔａｖ：平均温度、ＴＡ 、 Ｔｍ
、 Ｔｃ ：各断面位置での温度）による翼高さ方向の
外部熱負荷変化に対し、翼高さ中央の冷却流路断面積ｔ
−挑根元・先端部付近よシ小さくすることで真中央部を
流れる冷却流体の流速を高められ、ガスタービン翼内部
における温度の均一化を達成することかできる。This is a return flow convection cooling blade similar to the above embodiment, but by making the blade center wall 16 protrude in addition to the protruding partition walls, nine cooling channels 2 to 2 are arranged near the intrinsic surface.
The cross-sectional area of 4 is changed. Based on the same principle as the above embodiment, the non-uniform gas temperature distribution (T,
, ,: maximum temperature, Tav: average temperature, TA, Tm
, Tc: temperature at each cross-sectional position), the cooling flow passage cross-sectional area at the center of the blade height t
- By making the area near the root and tip of the blade smaller, the flow velocity of the cooling fluid flowing through the center can be increased, making it possible to achieve uniform temperature inside the gas turbine blade.
第６図および第７図に本発明の第３の実施例を示す。こ
れは翼高さ方向を用いている冷却流路を興外表面付近に
配設し九対流冷却翼（シェル・スパ一対流冷却翼も含ま
れる）に適用した例である。A third embodiment of the present invention is shown in FIGS. 6 and 7. This is an example in which a cooling channel using the blade height direction is arranged near the outer surface of the blade and applied to nine convection cooling blades (including shell and spa one convection cooling blades).
ガスタービン翼１は内部空洞１８をもつ中空構造になっ
ておシ、薄い翼壁１７内に翼高さ方向の微細な冷却流路
２が並べられて翼前手部分を冷却し、内部空洞１８から
後縁部に通ずる翼弦方向の排出流路１３で後縁部が冷却
される。翼高さ方向の冷却流路２は翼高さ中央（Ｂ−Ｂ
断面）の断面積（または高さ）か翼根光・先端部付近（
Ｃ−Ｃ断面・Ａ−Ａ断面）よシ小さくなっている。これ
によシ、この部分での冷却流体の流速そして熱伝達率を
大きくして、上記の説明の第４図のようなガス温度不均
一による翼高さ方向の外部熱負荷変化に対応し、温度の
高い真中央部の冷却効果を高めることかでき翼の温度の
均一化を達成することができる。The gas turbine blade 1 has a hollow structure with an internal cavity 18. Fine cooling channels 2 are lined up in the blade height direction within a thin blade wall 17 to cool the front part of the blade. The trailing edge is cooled by a chordwise exhaust flow path 13 leading from the blade to the trailing edge. The cooling flow path 2 in the blade height direction is located at the center of the blade height (B-B
The cross-sectional area (or height) of the blade root light/tip area (
(C-C cross section / A-A cross section). Accordingly, the flow velocity and heat transfer coefficient of the cooling fluid in this part are increased to cope with external heat load changes in the blade height direction due to non-uniform gas temperature as shown in Fig. 4 of the above explanation. It is possible to improve the cooling effect in the center, where the temperature is high, and to achieve uniform temperature of the blades.
以上のように本発明によれば、カスタービン翼の冷却効
果を向上することができるー、方、翼の温度分布を均一
化することかできた、具体的な数値をあければ翼内部の
最大温度と最小温度の差が従来では１０００程度あった
ものが本発明を用いることによシ温度差を８０Ｃ以下に
することができた。本発明は翼の内部において温度の高
い部分を下げることによって温度の均一化をはかつてい
るため、冷却流体の流量を増やすことなく目的か達成で
き、更に、翼の温夏差が１１、ｊ ＜なったことによシ
、材料の一部に熱負荷が大きくかかるようなことかなく
なシ、材料の熱ひずみか小さくなるとともに材料の長寿
命化につ危がる。As described above, according to the present invention, the cooling effect of the cast turbine blade can be improved, and the temperature distribution of the blade can be made uniform. Conventionally, the difference between the temperature and the minimum temperature was about 1000 degrees, but by using the present invention, the temperature difference could be reduced to 80C or less. Since the present invention aims to equalize the temperature by lowering the high temperature part inside the blade, the objective can be achieved without increasing the flow rate of the cooling fluid, and furthermore, the temperature difference of the blade is reduced to 11, j < As a result, a large thermal load will not be applied to a part of the material, the thermal strain of the material will be reduced, and the life of the material will be extended.
第１図は本発明のガスタービン翼の内部冷却構造の本発
明による第１の実施例における翼の横断面図、第２図は
その翼断面中心線に沿った縦断面図、第３図は本発明の
第２の実施例における興の横断面図、第４図は翼高さ方
向の温度分布図、第５図は第２の実施例の翼の正面図、
第６図は本発明の第３の実施例における翼の横断面図、
第７図はその流路形状を示すため拡大した部分横断面図
、第８図は第３の実施例の翼の正面図、第９図、第１０
図は従来例を示す。
１・・・カスタービン具、２〜５・・・冷却流路、６・
・・翼内面、７・・・隔壁、９・・・導入流路、１０・
・・冷却流体、゛・−−−Ｉ
番Ｉ目
蓼３０
第しい
寮９０FIG. 1 is a cross-sectional view of a blade in a first embodiment of the internal cooling structure for a gas turbine blade according to the present invention, FIG. 2 is a longitudinal cross-sectional view along the center line of the blade cross section, and FIG. A cross-sectional view of the blade in the second embodiment of the present invention, FIG. 4 is a temperature distribution diagram in the blade height direction, and FIG. 5 is a front view of the blade in the second embodiment.
FIG. 6 is a cross-sectional view of a blade in a third embodiment of the present invention;
FIG. 7 is a partial cross-sectional view enlarged to show the shape of the flow path, FIG. 8 is a front view of the blade of the third embodiment, and FIGS. 9 and 10.
The figure shows a conventional example. DESCRIPTION OF SYMBOLS 1... Caster turbine tool, 2-5... Cooling channel, 6...
...Blade inner surface, 7...Partition wall, 9...Introduction channel, 10.
・・Cooling fluid, ゛・---I No. 1 30 No. 1 dormitory 90
Claims (1)
、冷却流体を翼端部から前記冷却流路へ送る導入流路と
、前記冷却流路からの冷却流体を翼端部または翼外部へ
流出させる排出流路を備え、前記冷却流路の側面は外殻
と隔壁で構成されたガスタービン翼において、前記外殻
の外表面から内面に至る伝熱通路の熱伝達抵抗をほぼ一
定にし、翼の根元部および翼の先端部よりも翼の中央部
の前記冷却流路の断面積を小さくしたことを特徴とする
ガスタービン翼。 ２、特許請求の範囲第１項において、翼の中央部の前記
隔壁を冷却通路に突出させることにより前記冷却流路の
断面積を小さくしたことを特徴とするガスタービン翼。 ３、特許請求の範囲第１項において、前記冷却流路の側
面は外殻と隔壁と翼中央壁で構成し、翼の中央部の前記
翼中央壁を冷却通路に突出させることにより前記冷却流
路の断面積を小さくしたことを特徴とするガスタービン
翼。 ４、特許請求の範囲第１項において、前記冷却流路の側
面に板あるいは柱状の突起部を設けたことを特徴とする
ガスタービン翼。 ５、ガスタービン翼の内部に冷却流体を流す第１の冷却
流路と、冷却流体を翼端部から前記第１の冷却流路へ送
る導入流路と、前記第１の冷却流路からの冷却流体を翼
端部または翼外部へ流出させる排出流路を備え、前記第
１の冷却流路を構成する外殻の中に多数の第２の冷却流
路をもつガスタービン翼において、翼の根元部および翼
の先端部よりも翼の中央部の前記第２の冷却流路の断面
積を小さくしたことを特徴とするガスタービン翼。[Scope of Claims] 1. A cooling channel for flowing cooling fluid into the inside of a gas turbine blade, an introduction channel for sending cooling fluid from the blade tip to the cooling channel, and a cooling fluid flowing from the cooling channel. In a gas turbine blade that includes a discharge flow path for discharging to the blade tip or outside of the blade, and the side surface of the cooling flow path is composed of an outer shell and a partition wall, the heat of the heat transfer passage from the outer surface of the outer shell to the inner surface of the outer shell is removed. A gas turbine blade characterized in that the transmission resistance is kept almost constant, and the cross-sectional area of the cooling flow path in the center of the blade is smaller than that in the root and tip of the blade. 2. A gas turbine blade according to claim 1, characterized in that the partition wall at the center of the blade projects into the cooling passage to reduce the cross-sectional area of the cooling passage. 3. In claim 1, the side surface of the cooling flow path is constituted by an outer shell, a partition wall, and a blade center wall, and the cooling flow is controlled by making the blade center wall in the center of the blade protrude into the cooling path. A gas turbine blade characterized by a narrow cross-sectional area. 4. A gas turbine blade according to claim 1, characterized in that a plate or columnar protrusion is provided on a side surface of the cooling flow path. 5. A first cooling channel that allows cooling fluid to flow inside the gas turbine blade, an introduction channel that sends the cooling fluid from the blade tip to the first cooling channel, and a flow path that flows the cooling fluid from the first cooling channel. A gas turbine blade having a discharge passage for discharging cooling fluid to a blade tip or outside the blade, and having a large number of second cooling passages in an outer shell constituting the first cooling passage. A gas turbine blade, characterized in that the cross-sectional area of the second cooling channel in the center of the blade is smaller than that in the root and the tip of the blade.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP60042678A JPS61205301A (en) | 1985-03-06 | 1985-03-06 | Gas turbine blade |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP60042678A JPS61205301A (en) | 1985-03-06 | 1985-03-06 | Gas turbine blade |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| JPS61205301A true JPS61205301A (en) | 1986-09-11 |
Family
ID=12642685
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| JP60042678A Pending JPS61205301A (en) | 1985-03-06 | 1985-03-06 | Gas turbine blade |
Country Status (1)
| Country | Link |
|---|---|
| JP (1) | JPS61205301A (en) |
Cited By (12)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JPH03141801A (en) * | 1990-09-19 | 1991-06-17 | Hitachi Ltd | Cooling blade of gas turbine |
| US5215431A (en) * | 1991-06-25 | 1993-06-01 | Societe Nationale D'etude Et De Construction De Moteurs D'aviation "S.N.E.C.M.A." | Cooled turbine guide vane |
| US5511309A (en) * | 1993-11-24 | 1996-04-30 | United Technologies Corporation | Method of manufacturing a turbine airfoil with enhanced cooling |
| US5669759A (en) * | 1995-02-03 | 1997-09-23 | United Technologies Corporation | Turbine airfoil with enhanced cooling |
| JP2001214703A (en) * | 2000-02-02 | 2001-08-10 | General Electric Co <Ge> | Gas turbine bucket cooling circuit and its cooling method |
| EP1630353A2 (en) | 2004-08-25 | 2006-03-01 | Rolls-Royce Plc | Internally cooled gas turbine aerofoil |
| JP2017529479A (en) * | 2014-07-24 | 2017-10-05 | シーメンス アクチエンゲゼルシヤフトＳｉｅｍｅｎｓ Ａｋｔｉｅｎｇｅｓｅｌｌｓｃｈａｆｔ | Turbine blade cooling system with a flow blocker extending in the blade length direction |
| US10626733B2 (en) | 2017-10-03 | 2020-04-21 | United Technologies Corporation | Airfoil having internal hybrid cooling cavities |
| US10626734B2 (en) | 2017-10-03 | 2020-04-21 | United Technologies Corporation | Airfoil having internal hybrid cooling cavities |
| US10633980B2 (en) | 2017-10-03 | 2020-04-28 | United Technologies Coproration | Airfoil having internal hybrid cooling cavities |
| US10704398B2 (en) | 2017-10-03 | 2020-07-07 | Raytheon Technologies Corporation | Airfoil having internal hybrid cooling cavities |
| CN111868352A (en) * | 2018-04-17 | 2020-10-30 | 三菱动力株式会社 | Turbine blade and gas turbine |
-
1985
- 1985-03-06 JP JP60042678A patent/JPS61205301A/en active Pending
Cited By (15)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JPH03141801A (en) * | 1990-09-19 | 1991-06-17 | Hitachi Ltd | Cooling blade of gas turbine |
| US5215431A (en) * | 1991-06-25 | 1993-06-01 | Societe Nationale D'etude Et De Construction De Moteurs D'aviation "S.N.E.C.M.A." | Cooled turbine guide vane |
| US5511309A (en) * | 1993-11-24 | 1996-04-30 | United Technologies Corporation | Method of manufacturing a turbine airfoil with enhanced cooling |
| US5669759A (en) * | 1995-02-03 | 1997-09-23 | United Technologies Corporation | Turbine airfoil with enhanced cooling |
| JP2001214703A (en) * | 2000-02-02 | 2001-08-10 | General Electric Co <Ge> | Gas turbine bucket cooling circuit and its cooling method |
| EP1630353A3 (en) * | 2004-08-25 | 2009-10-21 | Rolls-Royce Plc | Internally cooled gas turbine aerofoil |
| EP1630353A2 (en) | 2004-08-25 | 2006-03-01 | Rolls-Royce Plc | Internally cooled gas turbine aerofoil |
| US8052389B2 (en) | 2004-08-25 | 2011-11-08 | Rolls-Royce Plc | Internally cooled airfoils with load carrying members |
| JP2017529479A (en) * | 2014-07-24 | 2017-10-05 | シーメンス アクチエンゲゼルシヤフトＳｉｅｍｅｎｓ Ａｋｔｉｅｎｇｅｓｅｌｌｓｃｈａｆｔ | Turbine blade cooling system with a flow blocker extending in the blade length direction |
| US10626733B2 (en) | 2017-10-03 | 2020-04-21 | United Technologies Corporation | Airfoil having internal hybrid cooling cavities |
| US10626734B2 (en) | 2017-10-03 | 2020-04-21 | United Technologies Corporation | Airfoil having internal hybrid cooling cavities |
| US10633980B2 (en) | 2017-10-03 | 2020-04-28 | United Technologies Coproration | Airfoil having internal hybrid cooling cavities |
| US10704398B2 (en) | 2017-10-03 | 2020-07-07 | Raytheon Technologies Corporation | Airfoil having internal hybrid cooling cavities |
| US11649731B2 (en) | 2017-10-03 | 2023-05-16 | Raytheon Technologies Corporation | Airfoil having internal hybrid cooling cavities |
| CN111868352A (en) * | 2018-04-17 | 2020-10-30 | 三菱动力株式会社 | Turbine blade and gas turbine |
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