JPH03194101A - Gas turbine cooling moving blade - Google Patents
Gas turbine cooling moving bladeInfo
- Publication number
- JPH03194101A JPH03194101A JP33230389A JP33230389A JPH03194101A JP H03194101 A JPH03194101 A JP H03194101A JP 33230389 A JP33230389 A JP 33230389A JP 33230389 A JP33230389 A JP 33230389A JP H03194101 A JPH03194101 A JP H03194101A
- Authority
- JP
- Japan
- Prior art keywords
- blade
- cooling
- cooling passage
- shroud
- effective part
- 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
- 238000001816 cooling Methods 0.000 title claims abstract description 81
- 238000005192 partition Methods 0.000 claims abstract description 5
- 238000004891 communication Methods 0.000 abstract description 2
- 230000000149 penetrating effect Effects 0.000 abstract 1
- 239000007789 gas Substances 0.000 description 22
- 230000000694 effects Effects 0.000 description 6
- 238000005495 investment casting Methods 0.000 description 5
- 239000000567 combustion gas Substances 0.000 description 4
- 238000003754 machining Methods 0.000 description 3
- 239000000463 material Substances 0.000 description 2
- 238000000034 method Methods 0.000 description 2
- 238000007789 sealing Methods 0.000 description 2
- 239000000919 ceramic Substances 0.000 description 1
- 239000000470 constituent Substances 0.000 description 1
- 239000002826 coolant Substances 0.000 description 1
- 238000005260 corrosion Methods 0.000 description 1
- 230000007797 corrosion Effects 0.000 description 1
- 238000009792 diffusion process Methods 0.000 description 1
- 238000007599 discharging Methods 0.000 description 1
- 238000005516 engineering process Methods 0.000 description 1
- 239000012530 fluid Substances 0.000 description 1
- 230000013011 mating Effects 0.000 description 1
- 239000002184 metal Substances 0.000 description 1
- 230000003647 oxidation Effects 0.000 description 1
- 238000007254 oxidation reaction Methods 0.000 description 1
- 230000002093 peripheral effect Effects 0.000 description 1
- XLYOFNOQVPJJNP-UHFFFAOYSA-N water Substances O XLYOFNOQVPJJNP-UHFFFAOYSA-N 0.000 description 1
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05B—INDEXING SCHEME RELATING TO WIND, SPRING, WEIGHT, INERTIA OR LIKE MOTORS, TO MACHINES OR ENGINES FOR LIQUIDS COVERED BY SUBCLASSES F03B, F03D AND F03G
- F05B2240/00—Components
- F05B2240/80—Platforms for stationary or moving blades
- F05B2240/801—Platforms for stationary or moving blades cooled platforms
Landscapes
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
Description
【発明の詳細な説明】
〔発明の目的〕
（産業上の利用分野）
本発明はガスタービン冷却動翼に係り、特にチップシュ
ラウド内に冷却通路が形成され、チップシュラウドを均
一に冷却できるようにしたガスタービン冷却動翼に関す
る。[Detailed Description of the Invention] [Object of the Invention] (Industrial Application Field) The present invention relates to a gas turbine cooling rotor blade, and in particular, a cooling passage is formed in a tip shroud so that the tip shroud can be cooled uniformly. The present invention relates to a gas turbine cooling rotor blade.
（従来の技術）
一般にこの種のガスタービンでは、作動流体である高温
燃焼ガスを流通させるためのガス通路がケーシング内に
形成されている。このガス通路内には静翼と動翼とが交
互に配置されており、静翼はケーシングに固着されガス
通路内で移動しないよう保持される一方、動翼はその根
元がタービンロータに放射状に植設され、タービンロー
タと一体的に回転できるようになっている。また、この
動翼の翼端と対向するケーシングの内面にはシュラウド
セグメントが固着されている。(Prior Art) Generally, in this type of gas turbine, a gas passage is formed in a casing for flowing high-temperature combustion gas, which is a working fluid. Stator blades and rotor blades are arranged alternately within this gas passage.The stator blades are fixed to the casing and are held so as not to move within the gas passage, while the roots of the rotor blades are radially connected to the turbine rotor. It is installed so that it can rotate integrally with the turbine rotor. Further, a shroud segment is fixed to the inner surface of the casing facing the blade tip of the rotor blade.
ところで、ガスタービンの出力を増大させたり、出力効
率を向上させるためにはガスタービン入口温度を上昇さ
せればよいが、この場合、上記ガス通路内への入熱量も
増大し、ガス通路を構成する各部材の温度も相当に高温
度となる。このため、材料強度の低下や高温酸化等によ
る材料腐食を生じるおそれがある。By the way, in order to increase the output of the gas turbine or improve the output efficiency, it is sufficient to increase the gas turbine inlet temperature, but in this case, the amount of heat input into the gas passage also increases, causing the gas passage to become The temperature of each member becomes considerably high. Therefore, there is a risk that material strength may decrease or material corrosion may occur due to high-temperature oxidation.
そこで、これを防止するために上記構成部材を冷却する
種々の方法がとられている。たとえば、冷却媒体として
は空気、蒸気、水等が利用されるが、ガスタービンの圧
縮機から抽気した低温の抽気空気を利用した冷却方式が
一般的である。Therefore, in order to prevent this, various methods have been taken to cool the above-mentioned constituent members. For example, air, steam, water, etc. are used as the cooling medium, but a cooling method that uses low-temperature bleed air extracted from a compressor of a gas turbine is common.
一方、この種のガスタービンにおいては、動翼の空力性
能の向上と共振現象の防止のために動翼の翼端にチップ
シュラウドが形成されている。しかし、このチップシュ
ラウドは上記目的を達成できる反面、回転時の遠心力に
より大きな引張応力を受ける。この種のチップシュラウ
ドは冷却手段を有していないか、冷却手段として第１３
図及び第１４図に示したように放出孔をチップシュラウ
ドの表面に配置したものもある。すなわち、第１３図に
おいて放出孔３１．３１はチップシュラウド３２を貫通
するように設けられており、植込部３３からシャンク３
４を介して翼有効部３５内に形成されている冷却通路（
図示せず）を流れる冷却空気の一部を放出し、チップシ
ュラウドをフィルム冷却できるようになっている。On the other hand, in this type of gas turbine, a tip shroud is formed at the tip of the rotor blade in order to improve the aerodynamic performance of the rotor blade and prevent resonance phenomena. However, although this chip shroud can achieve the above purpose, it is subjected to large tensile stress due to centrifugal force during rotation. This type of chip shroud does not have a cooling means, or has a cooling means as a cooling means.
There is also one in which the discharge holes are arranged on the surface of the tip shroud as shown in FIG. 14 and FIG. That is, in FIG. 13, the discharge hole 31.31 is provided so as to penetrate the tip shroud 32, and the discharge hole 31.31 is provided so as to pass through the tip shroud 32.
A cooling passage (
A portion of the cooling air flowing through the tip shroud (not shown) can be released to provide film cooling of the chip shroud.
（発明が解決しようとする課題）
しかしながら、上述のフィルム冷却方式では冷却空気の
流れが狭い範囲にしか形成されず、また、端部３６．３
６・・・では十分な冷却効果が得られない。このため、
チップシュラウドの表面の各部においてメタル温度が不
均一になり、チップシュラウドに十分な強度が得られな
いという問題が生じる。この問題を回避するために高圧
段の動翼では第１５図及び第１６図に示したようにチッ
プシュラウドを設けない形式の動翼も用いられている。(Problems to be Solved by the Invention) However, in the film cooling method described above, the cooling air flow is formed only in a narrow range, and the end portion 36.
6..., a sufficient cooling effect cannot be obtained. For this reason,
The metal temperature becomes non-uniform in various parts of the surface of the chip shroud, causing a problem that sufficient strength cannot be obtained in the chip shroud. In order to avoid this problem, rotor blades of a type without a tip shroud are also used in high-pressure stage rotor blades, as shown in FIGS. 15 and 16.
しかし、この動翼では翼端と上記シュラウドセグメント
との間に形成されるクリアランスに翼有効部の腹側３５
ａから背側３５ｂにかけて図中矢印で示したようなチッ
プリークが発生し、動翼の翼端での損失が大きくなる。However, in this rotor blade, the clearance formed between the blade tip and the shroud segment includes the ventral side 35 of the blade effective part.
A tip leak as shown by the arrow in the figure occurs from a to the dorsal side 35b, and the loss at the tip of the rotor blade increases.
このため、チップシュラウドを有する動翼に比し、ガス
タービン性能が大幅に低下するという問題がある。For this reason, there is a problem in that the gas turbine performance is significantly reduced compared to rotor blades having a tip shroud.
そこで、本発明の目的は、上述した従来の技術が有する
問題点を解消し、チップシュラウドにおいても均一な冷
却効果を有し、高圧段のガスタービン性能を向上できる
ようなガスタービンの冷却動翼を提供することにある。SUMMARY OF THE INVENTION Therefore, an object of the present invention is to provide a cooling rotor blade for a gas turbine that solves the problems of the conventional technology described above, has a uniform cooling effect even in the tip shroud, and can improve the performance of the gas turbine in the high pressure stage. Our goal is to provide the following.
（課題を解決するための手段）
上記目的を達成するために、本発明は、タービン円板に
埋設される植込部にシャンクと翼有効部とを連接すると
ともに、この翼有効部の翼端にチップシュラウドを取着
し、上記シャンク内と翼有効部内とを貫通するようにし
て冷却通路を形成したガスタービン冷却動翼において、
上記チップシュラウドの内部に中空部を形成し、この中
空部に隔壁を配置して冷却通路とし、この冷却通路と上
記翼有効部内の冷却通路とを連結孔により連通させると
ともに、上記チップシュラウドの表面に放出孔を設け、
翼有効部内からチップシュラウド内に冷却空気を供給さ
せるようにしたことを特徴とするものである。(Means for Solving the Problems) In order to achieve the above object, the present invention connects a shank and an effective blade part to an embedded part embedded in a turbine disk, and also provides a blade tip of the effective blade part. In a gas turbine cooling rotor blade, a tip shroud is attached to the blade, and a cooling passage is formed so as to pass through the inside of the shank and the inside of the blade effective part,
A hollow part is formed inside the tip shroud, a partition wall is arranged in the hollow part to form a cooling passage, and this cooling passage and a cooling passage in the effective blade part are communicated through a connecting hole, and the surface of the tip shroud is A discharge hole is provided in the
This is characterized in that cooling air is supplied into the tip shroud from within the blade effective section.
（作 用）
本発明によれば、チップシュラウドの内部に中空部を形
成し、この中空部に隔壁を配置して冷却通路とし、この
冷却通路と上記翼有効部内の冷却通路とを連結孔により
連通させるとともに、上記チップシュラウドの表面に放
出孔を設け、翼有効部内からチップシュラウド内に冷却
空気を供給させるようにしたので、チップシュラウド内
を流れる冷却空気によりチップシュラウドを均一に冷却
することができる。(Function) According to the present invention, a hollow part is formed inside the tip shroud, a partition is arranged in this hollow part to form a cooling passage, and this cooling passage and the cooling passage in the above-mentioned blade effective part are connected by a connecting hole. In addition to providing communication, a discharge hole is provided on the surface of the tip shroud to supply cooling air from within the effective part of the blade to the tip shroud, so that the tip shroud can be uniformly cooled by the cooling air flowing inside the tip shroud. can.
（実施例）
以下本発明によるガスタービン冷却動翼の一実施例を第
１図乃至第６図を参照して説明する。本実施例において
は冷却効率の高いリターンフロー型対流冷却方式を採用
している冷却動翼を例に説明するが、他の冷却方式にお
いてもその作用効果は同様である。(Example) An example of a gas turbine cooling rotor blade according to the present invention will be described below with reference to FIGS. 1 to 6. In this embodiment, a cooling rotor blade employing a return flow type convection cooling method with high cooling efficiency will be explained as an example, but the operation and effect are similar for other cooling methods.
第１図および第２図はガスタービン冷却動翼の全体を示
しており、図中ね号１は植込部を示しており、この植込
部１は先端にかけて尖る樹形状をなし、従来例と同様に
タービン円板（図示せず）に埋設できるようになってい
る。この植込部１の上方にはシャンク２と翼有効部３と
が連接されている。上記シャンク２は略四角錘台形をな
し、ガスに対するシール効果を高めるためにプラットフ
ォーム２ａ、、２ａ・・・が形成されている。また、翼
有効部３の翼型は根元３ａと翼端３ｂとかねじるように
形成されており、」二記植込部１、シャンク２及び翼有
効部３とを貫通するように内部に図示しない冷却通路が
形成されている。この冷却通路内の冷却空気は翼有効部
３を対流冷却するとともに、その一部は翼前縁３Ｃをフ
ィルム冷却し、その他は翼後縁の吹出孔等から燃焼ガス
中に放出される。Figures 1 and 2 show the entire gas turbine cooling rotor blade, and number 1 in the figures indicates the implanted part. Similarly, it can be embedded in a turbine disc (not shown). A shank 2 and an effective blade portion 3 are connected above the implanted portion 1. The shank 2 has a substantially quadrangular trapezoidal shape, and platforms 2a, 2a, . . . are formed to enhance the sealing effect against gas. In addition, the airfoil of the blade effective part 3 is formed so that the root 3a and the blade tip 3b are twisted, and the airfoil (not shown) is formed inside so as to penetrate through the implanted part 1, the shank 2, and the blade effective part 3. A cooling passage is formed. The cooling air in this cooling passage convectively cools the blade effective section 3, a part of which cools the blade leading edge 3C as a film, and the rest is released into the combustion gas from the blow-off holes at the blade trailing edge.
さらに、第３図及び第４図に示したように上記翼有効部
３の翼端３ｂにはチップシュラウド４が形成されている
。このチップシュラウド４は平面形状が略Ｚ字形をなし
、上記翼有効部３と一体的に形成されたチップシュラウ
ド下部４Ａとこのチップシュラウド下部４Ａに積重する
ように固着されたチップシュラウド上蓋４Ｂとから構成
されている。上記チップシュラウド下部４Ａは精密鋳造
時に翼有効部３と一体的に形成される一方、チップシュ
ラウド上蓋４Ｂは精密鋳造あるいは機械加工により製造
され、この上面には２列の翼端シール用のフィン５．５
が配設されている。また、上記チップシュラウド上蓋４
Ｂとチップシュラウド下部４Ａとはその合わせ面に機械
加工が施され、拡散接合法により接合されている。Further, as shown in FIGS. 3 and 4, a tip shroud 4 is formed at the blade tip 3b of the blade effective section 3. This chip shroud 4 has a substantially Z-shaped planar shape, and includes a lower chip shroud 4A formed integrally with the effective blade section 3 and a chip shroud upper lid 4B fixed to the lower chip shroud 4A in a stacked manner. It consists of The tip shroud lower part 4A is formed integrally with the blade effective part 3 during precision casting, while the tip shroud upper cover 4B is manufactured by precision casting or machining, and its upper surface has two rows of fins 5 for wing tip sealing. .5
is installed. In addition, the chip shroud upper cover 4
B and the chip shroud lower part 4A are machined on their mating surfaces and are joined by a diffusion bonding method.
一方、第５図及び第６図はチップシュラウド４内に形成
された冷却通路６を示している。この冷却通路６は上記
チップシュラウド下部４Ａの内部に一体的に形成されて
おり、外周壁と複数の隔壁とからなっている。この冷却
通路６は翼体の精密鋳造時に形成されるか、あらかじめ
平面状に鋳造されたチップシュラウド下部４Ａに機械加
工により形成されるようになっている。さらにこの冷却
通路６の底面には上記翼有効部３内の冷却通路３ｄと連
通ずるための２つの連結孔７．７が形成されている。こ
の連結孔７は通常セラミック中子等により翼体の精密鋳
造時に同時に形成される。On the other hand, FIGS. 5 and 6 show cooling passages 6 formed within the chip shroud 4. As shown in FIG. The cooling passage 6 is integrally formed inside the chip shroud lower portion 4A, and consists of an outer peripheral wall and a plurality of partition walls. The cooling passage 6 is formed during precision casting of the blade body, or is formed by machining in the tip shroud lower part 4A, which is previously cast into a flat shape. Further, two connecting holes 7.7 are formed in the bottom surface of the cooling passage 6 to communicate with the cooling passage 3d in the blade effective section 3. This connecting hole 7 is usually formed using a ceramic core or the like at the same time as precision casting of the blade body.
他方、上記チップシュラウド上蓋４Ｂの上面には３つの
小径の放出孔８が穿設されている。この放出孔８は冷却
通路６内に流入した冷却空気を燃焼ガス内に放出する役
割を果たし、上記チップシュラウド上Ｍ４Ｂを精密鋳造
する際に同時に形成されるか、後に機械加工により孔あ
けされるようになっている。On the other hand, three small-diameter discharge holes 8 are bored in the upper surface of the tip shroud upper cover 4B. This discharge hole 8 plays the role of discharging the cooling air that has flowed into the cooling passage 6 into the combustion gas, and is formed at the same time as precision casting of the M4B on the tip shroud, or is bored later by machining. It looks like this.
以上に述べた構成によれば、チップシュラウド４の冷却
通路６内の冷却空気の流れは第５図及び第６図に示した
ようになる。すなわち、翼有効部３内の冷却通路３ｄを
流れる冷却空気の一部が上記連結孔７を通じてチップシ
ュラウド４の冷却通路６内に流入し、内部を対流して放
出孔８から放出される。これによりチップシュラウド４
内を均一に冷却することができる。According to the configuration described above, the flow of cooling air in the cooling passage 6 of the chip shroud 4 is as shown in FIGS. 5 and 6. That is, a portion of the cooling air flowing through the cooling passage 3d in the blade effective portion 3 flows into the cooling passage 6 of the tip shroud 4 through the connection hole 7, convects inside, and is discharged from the discharge hole 8. This allows the chip shroud 4
The interior can be cooled evenly.
第７図及び第８図は他の実施例として伝熱促進体として
ピンフィン９．９・・・を設けた冷却通路６を示してお
り、このビンフィン９は円柱状をなし、冷却通路６内に
一定の間隔をあけて立設されている。また、第９図及び
第１０図は伝熱促進体としてタービュレンスプロモータ
１０．１０・・・を配設した冷却通路６を示している。7 and 8 show a cooling passage 6 provided with pin fins 9, 9, etc. as heat transfer accelerators as another embodiment. They are set up at regular intervals. Moreover, FIGS. 9 and 10 show a cooling passage 6 in which turbulence promoters 10, 10, . . . are disposed as heat transfer promoters.
このタービュレンスプロモータ１０は冷却空気の流れ方
向に直交するような方向に延在する棒状の凸部であり、
冷却通路６の上下面の両面に一定の間隔をあけて突設さ
れている。このように伝熱促進体を用いることにより冷
却通路６内の表面積を大きくすることができ、より冷却
効率を高めることができる。This turbulence promoter 10 is a rod-shaped convex portion extending in a direction perpendicular to the flow direction of cooling air,
They are provided protrudingly on both the upper and lower surfaces of the cooling passage 6 at regular intervals. By using the heat transfer promoter in this way, the surface area within the cooling passage 6 can be increased, and the cooling efficiency can be further improved.
第１１図及び第１２図は他の実施例として放出孔８をチ
ップシュラウド下部４Ａの下面に穿設した冷却通路６を
示している。このように放出孔８を設けることで冷却空
気をガス通路内に放出し、翼有効部をフィルム冷却する
ことができる。このとき、上記放出孔８は放出された冷
却空気が上記燃焼ガスの流れを乱さないように後方に向
かって斜めに形成されている。FIGS. 11 and 12 show another embodiment of a cooling passage 6 in which discharge holes 8 are formed in the lower surface of the lower tip shroud 4A. By providing the discharge holes 8 in this manner, cooling air can be discharged into the gas passage, and the effective portion of the blade can be film-cooled. At this time, the discharge hole 8 is formed obliquely toward the rear so that the discharged cooling air does not disturb the flow of the combustion gas.
以上の説明から明らかなように、本発明によれば、チッ
プシュラウドの内部に冷却通路を形成し、この冷却通路
によりチップシュラウド内に冷却空気を供給させるよう
にしたので、チップシュラウドを均一に冷却することが
でき、この結果、高圧段の動翼にもチップシュラウドを
設けることが可能となり、ガスタービンの性能が向上す
る等の効果が期待できる。As is clear from the above description, according to the present invention, a cooling passage is formed inside the chip shroud, and cooling air is supplied into the chip shroud by this cooling passage, so that the chip shroud is uniformly cooled. As a result, it becomes possible to provide a tip shroud even on the rotor blades of the high pressure stage, and effects such as improving the performance of the gas turbine can be expected.
第１図は本発明によるガスタービン冷却動翼の一実施例
の全体概略を示した平面図、第２図は同正面図、第３図
は本発明によるチップシュラウドの外形を示すために冷
却動翼の一部を拡大した平面図、第４図は同正面図、第
５図及び第６図は本発明によるチップシュラウド冷却通
路を示すために冷却動翼の一部を拡大した断面図、第７
図及び第８図は他の実施例としてビンフィンを設けたチ
ップシュラウド冷却通路を示した断面図、第９図及び第
１０図は他の実施例としてタービュレンスプロモータを
設けたチップシュラウド冷却通路を示した断面図、第１
１図は他の実施例として放出孔をチップシュラウドの下
面に設けたチップシュラウド冷却通路を示した平面図、
第１２図は同正面図、第１３図は従来のチップシュラウ
ドを有する動翼の一例の全体概略を示した平面図、第１
４図は同正面図、第１５図は従来のチップシュラウドを
有しない動翼の一例の全体概略を示した平面図、第１６
図は同正面図である。
３・・・翼有効部、４・・・チップシュラウド、６・・
・冷却通路、７・・・連結孔、８・・・放出孔。FIG. 1 is a plan view showing an overall outline of an embodiment of a gas turbine cooling rotor blade according to the present invention, FIG. 2 is a front view of the same, and FIG. FIG. 4 is an enlarged plan view of a portion of the blade, FIG. 4 is a front view thereof, and FIGS. 7
8 and 8 are cross-sectional views showing a chip shroud cooling passage provided with bin fins as another embodiment, and FIGS. 9 and 10 show a chip shroud cooling passage provided with a turbulence promoter as another embodiment. cross-sectional view, 1st
Figure 1 is a plan view showing a chip shroud cooling passage in which discharge holes are provided on the lower surface of the chip shroud as another embodiment;
FIG. 12 is a front view of the same, FIG. 13 is a plan view schematically showing an example of the rotor blade having a conventional tip shroud, and FIG.
4 is a front view of the same, FIG. 15 is a plan view schematically showing an example of a rotor blade without a conventional tip shroud, and FIG.
The figure is a front view of the same. 3... Wing effective part, 4... Chip shroud, 6...
- Cooling passage, 7... connection hole, 8... discharge hole.
Claims (1)
とを連接するとともに、この翼有効部の翼端にチップシ
ュラウドを取着し、上記シャンク内と翼有効部内とを貫
通するようにして冷却通路を形成したガスタービン冷却
動翼において、上記チップシュラウドの内部に中空部を
形成し、この中空部に隔壁を配置して冷却通路とし、こ
の冷却通路と上記翼有効部内の冷却通路とを連結孔によ
り連通させるとともに、上記チップシュラウドの表面に
放出孔を設け、翼有効部からチップシユラウド内に冷却
空気を供給させるようにしたことを特徴とするガスター
ビン冷却動翼。The shank and the blade effective part are connected to an embedded part embedded in the turbine disc, and a tip shroud is attached to the blade tip of the blade effective part so as to pass through the inside of the shank and the blade effective part. In a gas turbine cooled rotor blade having a cooling passage formed therein, a hollow part is formed inside the tip shroud, a partition is arranged in the hollow part to form a cooling passage, and this cooling passage and a cooling passage in the effective part of the blade are connected. A gas turbine cooling rotor blade, characterized in that the blades are communicated with each other through a connecting hole, and a discharge hole is provided on the surface of the tip shroud, so that cooling air is supplied from an effective part of the blade into the tip shroud.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP33230389A JPH03194101A (en) | 1989-12-21 | 1989-12-21 | Gas turbine cooling moving blade |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP33230389A JPH03194101A (en) | 1989-12-21 | 1989-12-21 | Gas turbine cooling moving blade |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| JPH03194101A true JPH03194101A (en) | 1991-08-23 |
Family
ID=18253453
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| JP33230389A Pending JPH03194101A (en) | 1989-12-21 | 1989-12-21 | Gas turbine cooling moving blade |
Country Status (1)
| Country | Link |
|---|---|
| JP (1) | JPH03194101A (en) |
Cited By (20)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US5320485A (en) * | 1992-06-11 | 1994-06-14 | Societe Nationale D'etude Et De Construction De Moteurs D'aviation (S.N.E.C.M.A.) | Guide vane with a plurality of cooling circuits |
| US5536143A (en) * | 1995-03-31 | 1996-07-16 | General Electric Co. | Closed circuit steam cooled bucket |
| WO1999000584A1 (en) * | 1997-06-26 | 1999-01-07 | Mitsubishi Heavy Industries, Ltd. | Tip shroud for moving blades of gas turbine |
| JP2000291405A (en) * | 1999-04-05 | 2000-10-17 | General Electric Co <Ge> | Cooling circuit for gas turbine bucket and upper shroud |
| US6152695A (en) * | 1998-02-04 | 2000-11-28 | Mitsubishi Heavy Industries, Ltd. | Gas turbine moving blade |
| EP1126136A2 (en) * | 1999-12-28 | 2001-08-22 | ALSTOM (Schweiz) AG | Turbine blade with air cooled tip shroud |
| WO2006029983A1 (en) * | 2004-09-16 | 2006-03-23 | Alstom Technology Ltd | Turbine engine vane with fluid cooled shroud |
| JP2006316750A (en) * | 2005-05-16 | 2006-11-24 | Hitachi Ltd | Gas turbine moving blade, gas turbine using the same, and its power generation plant |
| JP2007327493A (en) * | 2006-06-07 | 2007-12-20 | General Electric Co <Ge> | Serpentine cooling circuit and method for cooling shroud |
| JP2008169845A (en) * | 2007-01-12 | 2008-07-24 | General Electric Co <Ge> | Impingement cooled bucket shroud, turbine rotor incorporating the same, and cooling method |
| JP2009168018A (en) * | 2008-01-10 | 2009-07-30 | General Electric Co <Ge> | Turbine moving blade tip shroud |
| JP2009168017A (en) * | 2008-01-10 | 2009-07-30 | General Electric Co <Ge> | Turbine blade tip shroud |
| JP2009168014A (en) * | 2008-01-10 | 2009-07-30 | General Electric Co <Ge> | Turbine blade tip shroud |
| JP2010031865A (en) * | 2008-07-29 | 2010-02-12 | General Electric Co <Ge> | Rotor blade and method of fabricating the same |
| JP2011001919A (en) * | 2009-06-21 | 2011-01-06 | Toshiba Corp | Turbine moving blade |
| US20130142667A1 (en) * | 2011-10-27 | 2013-06-06 | Mitsubishi Heavy Industries, Ltd. | Turbine blade and gas turbine having the same |
| JP2013117227A (en) * | 2011-12-01 | 2013-06-13 | General Electric Co <Ge> | Cooled turbine blade and method for cooling turbine blade |
| EP2881541A1 (en) * | 2013-12-05 | 2015-06-10 | Rolls-Royce Deutschland Ltd & Co KG | Tip cooling of a turbine rotor blade of a gas turbine |
| JP6025941B1 (en) * | 2015-08-25 | 2016-11-16 | 三菱日立パワーシステムズ株式会社 | Turbine blade and gas turbine |
| US10746029B2 (en) | 2017-02-07 | 2020-08-18 | General Electric Company | Turbomachine rotor blade tip shroud cavity |
-
1989
- 1989-12-21 JP JP33230389A patent/JPH03194101A/en active Pending
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| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US5320485A (en) * | 1992-06-11 | 1994-06-14 | Societe Nationale D'etude Et De Construction De Moteurs D'aviation (S.N.E.C.M.A.) | Guide vane with a plurality of cooling circuits |
| US5536143A (en) * | 1995-03-31 | 1996-07-16 | General Electric Co. | Closed circuit steam cooled bucket |
| WO1999000584A1 (en) * | 1997-06-26 | 1999-01-07 | Mitsubishi Heavy Industries, Ltd. | Tip shroud for moving blades of gas turbine |
| US6152694A (en) * | 1997-06-26 | 2000-11-28 | Mitsubishi Heavy Industries, Ltd. | Tip shroud for moving blades of gas turbine |
| US6152695A (en) * | 1998-02-04 | 2000-11-28 | Mitsubishi Heavy Industries, Ltd. | Gas turbine moving blade |
| JP4514877B2 (en) * | 1999-04-05 | 2010-07-28 | ゼネラル・エレクトリック・カンパニイ | Cooling circuit for gas turbine bucket and upper shroud |
| JP2000291405A (en) * | 1999-04-05 | 2000-10-17 | General Electric Co <Ge> | Cooling circuit for gas turbine bucket and upper shroud |
| EP1126136A2 (en) * | 1999-12-28 | 2001-08-22 | ALSTOM (Schweiz) AG | Turbine blade with air cooled tip shroud |
| EP1126136A3 (en) * | 1999-12-28 | 2004-05-19 | ALSTOM Technology Ltd | Turbine blade with air cooled tip shroud |
| WO2006029983A1 (en) * | 2004-09-16 | 2006-03-23 | Alstom Technology Ltd | Turbine engine vane with fluid cooled shroud |
| US7427188B2 (en) | 2004-09-16 | 2008-09-23 | Alstom Technology Ltd | Turbomachine blade with fluidically cooled shroud |
| AU2005284134B2 (en) * | 2004-09-16 | 2008-10-09 | General Electric Technology Gmbh | Turbine engine vane with fluid cooled shroud |
| JP4628865B2 (en) * | 2005-05-16 | 2011-02-09 | 株式会社日立製作所 | Gas turbine blade, gas turbine using the same, and power plant |
| JP2006316750A (en) * | 2005-05-16 | 2006-11-24 | Hitachi Ltd | Gas turbine moving blade, gas turbine using the same, and its power generation plant |
| JP2007327493A (en) * | 2006-06-07 | 2007-12-20 | General Electric Co <Ge> | Serpentine cooling circuit and method for cooling shroud |
| US7686581B2 (en) * | 2006-06-07 | 2010-03-30 | General Electric Company | Serpentine cooling circuit and method for cooling tip shroud |
| JP2008169845A (en) * | 2007-01-12 | 2008-07-24 | General Electric Co <Ge> | Impingement cooled bucket shroud, turbine rotor incorporating the same, and cooling method |
| JP2009168014A (en) * | 2008-01-10 | 2009-07-30 | General Electric Co <Ge> | Turbine blade tip shroud |
| JP2009168017A (en) * | 2008-01-10 | 2009-07-30 | General Electric Co <Ge> | Turbine blade tip shroud |
| JP2009168018A (en) * | 2008-01-10 | 2009-07-30 | General Electric Co <Ge> | Turbine moving blade tip shroud |
| JP2010031865A (en) * | 2008-07-29 | 2010-02-12 | General Electric Co <Ge> | Rotor blade and method of fabricating the same |
| JP2011001919A (en) * | 2009-06-21 | 2011-01-06 | Toshiba Corp | Turbine moving blade |
| US9371741B2 (en) * | 2011-10-27 | 2016-06-21 | Mitsubishi Heavy Industries, Ltd. | Turbine blade and gas turbine having the same |
| US20130142667A1 (en) * | 2011-10-27 | 2013-06-06 | Mitsubishi Heavy Industries, Ltd. | Turbine blade and gas turbine having the same |
| JP2013117227A (en) * | 2011-12-01 | 2013-06-13 | General Electric Co <Ge> | Cooled turbine blade and method for cooling turbine blade |
| EP2881541A1 (en) * | 2013-12-05 | 2015-06-10 | Rolls-Royce Deutschland Ltd & Co KG | Tip cooling of a turbine rotor blade of a gas turbine |
| JP6025941B1 (en) * | 2015-08-25 | 2016-11-16 | 三菱日立パワーシステムズ株式会社 | Turbine blade and gas turbine |
| WO2017033920A1 (en) * | 2015-08-25 | 2017-03-02 | 三菱日立パワーシステムズ株式会社 | Turbine rotor blade, and gas turbine |
| CN107614835A (en) * | 2015-08-25 | 2018-01-19 | 三菱日立电力***株式会社 | Turbine rotor blade and gas turbine |
| CN107614835B (en) * | 2015-08-25 | 2019-11-12 | 三菱日立电力***株式会社 | Turbine rotor blade and gas turbine |
| US10890073B2 (en) | 2015-08-25 | 2021-01-12 | Mitsubishi Power, Ltd. | Turbine blade and gas turbine |
| DE112016002559B4 (en) | 2015-08-25 | 2021-09-09 | Mitsubishi Power, Ltd. | TURBINE BLADE AND GAS TURBINE |
| US10746029B2 (en) | 2017-02-07 | 2020-08-18 | General Electric Company | Turbomachine rotor blade tip shroud cavity |
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