JPH05312002A - Gas turbine blade - Google Patents
Gas turbine bladeInfo
- Publication number
- JPH05312002A JPH05312002A JP4117461A JP11746192A JPH05312002A JP H05312002 A JPH05312002 A JP H05312002A JP 4117461 A JP4117461 A JP 4117461A JP 11746192 A JP11746192 A JP 11746192A JP H05312002 A JPH05312002 A JP H05312002A
- Authority
- JP
- Japan
- Prior art keywords
- cooling
- gas turbine
- blade
- gas
- turbine blade
- 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.)
- Granted
Links
Landscapes
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
Description
【０００１】[0001]
【産業上の利用分野】本発明は、火力発電などに適用さ
れるガスタービンにおけるガスタービン翼に関する。BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a gas turbine blade in a gas turbine applied to thermal power generation and the like.
【０００２】[0002]
【従来の技術】図３および図４は火力発電に使用されて
いる従来のガスタービンの中空動翼の説明図である。図
において、本動翼には冷却空気が翼根１の底部から流入
し、矢印の方向に流れて本動翼を内部から冷却するよう
になっており、前縁側の入口２ａから流入した冷却空気
は冷却通路を通り堰状のフィン３により冷却効果が増大
されて対流冷却を行った後、チップシンニング（tip th
inning）４が設けられている翼頂部の出口５から流出し
て主ガス流れに合流する。また、後縁側の入口２ｂから
流入した冷却空気はフィン３が設けられている冷却通路
を矢印の方向に流れて対流冷却を行った後、円柱状のピ
ンフィン６を介して翼後縁を冷却し、出口Ａから流出し
て主ガス流れに合流する。2. Description of the Related Art FIGS. 3 and 4 are explanatory views of a hollow rotor blade of a conventional gas turbine used for thermal power generation. In the figure, the cooling air flows into the main blade from the bottom of the blade root 1 and flows in the direction of the arrow to cool the main blade from the inside. Passes through the cooling passage and the cooling effect is increased by the weir-shaped fins 3 to perform convective cooling, and then the tip thinning (tip th
(inning) 4 is discharged from an outlet 5 at the top of the blade, and joins the main gas flow. Further, the cooling air that has flowed in from the inlet 2b on the trailing edge side flows in the direction of the arrow in the cooling passage provided with the fins 3 to perform convection cooling, and then cools the blade trailing edge via the cylindrical pin fins 6. , Exits from outlet A and joins the main gas stream.
【０００３】フィン３は冷却空気の流れに対してほぼ直
角に精密鋳造によって冷却通路と一体に形成されてい
る。冷却空気の流れはフィン３によって二次元的に湾曲
し、その後流側にはほぼ二次元の渦７が発生する。これ
により、冷却通路の壁面における層流境界層が消滅して
乱流が発生し、熱伝達が冷却空気の層流伝熱から乱流伝
熱に移行して行われ、冷却通路の壁面から冷却空気への
熱伝達率が向上して動翼メタルの冷却効果を促進させて
いる。The fins 3 are integrally formed with the cooling passage by precision casting at a right angle to the flow of the cooling air. The flow of the cooling air is two-dimensionally curved by the fins 3, and then a substantially two-dimensional vortex 7 is generated on the flow side. As a result, the laminar boundary layer on the wall surface of the cooling passage disappears to generate turbulent flow, and heat transfer is performed from laminar heat transfer of cooling air to turbulent heat transfer. The heat transfer coefficient of the blade is improved to accelerate the cooling effect of the blade metal.
【０００４】[0004]
【発明が解決しようとする課題】上記のような従来のガ
スタービンの中空動翼において、最近のガスタービンは
高温化に伴って冷却空気量が逐次増加してきており、冷
却空気量の増加はガスタービンにおける熱効率を低下さ
せる原因となり、またフィン３が冷却通路の面積を減少
させるために空気速度が増大して冷却空気における圧力
損失の原因となるなど、これらはともにガスタービンに
おける高温化に対して障害となっている。In the hollow rotor blade of the conventional gas turbine as described above, the amount of cooling air in the recent gas turbine has been gradually increasing with the increase in temperature. These both cause the thermal efficiency in the turbine to decrease, and the fins 3 reduce the area of the cooling passage to increase the air velocity and cause pressure loss in the cooling air. It is an obstacle.
【０００５】[0005]
【課題を解決するための手段】本発明に係るガスタービ
ン翼は上記課題の解決を目的にしており、内部に設けら
れた冷却通路に冷却気体を流して翼メタルの冷却を行う
ガスタービン翼において、上記冷却通路の壁面に突設さ
れ上記冷却気体の上流に向けて低く傾斜するとともに下
流側の端部が立面をなす渦発生体を備えた構成を特徴と
する。SUMMARY OF THE INVENTION A gas turbine blade according to the present invention is intended to solve the above-mentioned problems, and a gas turbine blade for cooling blade metal by flowing a cooling gas through a cooling passage provided inside thereof. The vortex generator is provided so as to project from the wall surface of the cooling passage, is inclined downward toward the upstream side of the cooling gas, and has a vortex generator whose end portion on the downstream side is an upright surface.
【０００６】また、本発明に係るガスタービン翼におい
ては、内部に設けられた冷却通路に冷却気体を流して翼
メタルの冷却を行うガスタービン翼において、く字状に
上記冷却通路の両側壁から折れ曲がって突設され上記冷
却気体の上流に向けて張り出したピンフィンを備えた構
成を特徴とする。Further, in the gas turbine blade according to the present invention, in the gas turbine blade for cooling the blade metal by flowing the cooling gas into the cooling passage provided inside, the gas turbine blade is shaped like a dogleg from the both side walls of the cooling passage. It is characterized in that it is provided with pin fins that are bent and projected so as to project toward the upstream side of the cooling gas.
【０００７】[0007]
【作用】即ち、本発明に係るガスタービン翼において
は、内部に設けられた冷却通路に冷却気体を流して翼メ
タルの冷却を行うガスタービン翼における冷却通路の壁
面に冷却気体の上流に向けて低く傾斜するとともに下流
側の端部が立面をなす渦発生体が突設されており、冷却
通路の壁面に沿って渦が冷却気体の流れ方向および流れ
と直角方向に発生することにより乱流伝熱が三次元的な
渦運動により行われ、冷却通路の壁面から冷却気体への
熱伝達率が向上する。また、渦発生体は従来の堰状のフ
ィンに比べて小形で摩擦係数が小さい形状をしており、
冷却気体に対する流れ抵抗が減少する。That is, in the gas turbine blade according to the present invention, the cooling gas is caused to flow through the cooling passage provided inside to cool the blade metal in the wall surface of the cooling passage in the gas turbine blade toward the upstream side of the cooling gas. A vortex generator, which is inclined low and whose downstream end is an upright surface, is provided so as to project.Vortices are generated along the wall surface of the cooling passage in the flow direction of the cooling gas and in the direction perpendicular to the flow direction, thereby causing turbulent flow. The heat is generated by the three-dimensional vortex motion, and the heat transfer coefficient from the wall surface of the cooling passage to the cooling gas is improved. In addition, the vortex generator is smaller and has a smaller friction coefficient than the conventional dam-like fins.
The flow resistance to the cooling gas is reduced.
【０００８】また、本発明に係るガスタービン翼におい
ては、内部に設けられた冷却通路に冷却気体を流して翼
メタルの冷却を行うガスタービン翼における冷却通路の
両側壁からピンフィンがく字状に折れ曲がって突設され
冷却気体の上流に向けて張り出しており、ピンフィンの
先端から下流側に冷却通路の両側壁に向かって乱流境界
層の剥離による渦が発生し、冷却通路の壁面から冷却気
体への熱伝達率が向上する。また、ピンフィンの形状は
従来の円柱状のピンフィンに比べて摩擦係数が小さく、
冷却気体に対する流れ抵抗が減少する。Further, in the gas turbine blade according to the present invention, the pin fins are bent in a V shape from both side walls of the cooling passage in the gas turbine blade for cooling the blade metal by flowing the cooling gas through the cooling passage provided inside. The turbulent boundary layer is separated from the tip of the pin fins toward the upstream side of the cooling gas toward the both side walls of the cooling passage, and a vortex is generated from the wall of the cooling passage to the cooling gas. The heat transfer coefficient of is improved. Also, the shape of the pin fin has a smaller friction coefficient than the conventional cylindrical pin fin,
The flow resistance to the cooling gas is reduced.
【０００９】[0009]
【実施例】図１は本発明の一実施例に係るガスタービン
の中空動翼の説明図、図２はその応用例に係るガスター
ビンの中空動翼の説明図である。図において、本実施例
に係るガスタービンの中空動翼は火力発電に使用される
ガスタービンの中空動翼で、本動翼には図３における従
来のガスタービンの中空動翼と同様に冷却空気が翼根１
の底部から流入し、矢印の方向に流れて本動翼を内部か
ら冷却するようになっており、前縁側の入口２ａから流
入した冷却空気は冷却通路を通り対流冷却を行った後、
チップシンニング（tip thinning）４が設けられている
翼頂部の出口５から流出して主ガス流れに合流する。ま
た、後縁側の入口２ｂから流入した冷却空気は冷却通路
を矢印の方向に流れて対流冷却を行った後、ピンフィン
６を介して翼後縁を冷却し、出口Ａから流出して主ガス
流れに合流する。1 is an explanatory view of a hollow rotor blade of a gas turbine according to an embodiment of the present invention, and FIG. 2 is an explanatory view of a hollow rotor blade of a gas turbine according to an application example thereof. In the figure, the hollow rotor blade of the gas turbine according to the present embodiment is a hollow rotor blade of a gas turbine used for thermal power generation, and the main rotor blade has the same cooling air as the hollow rotor blade of the conventional gas turbine in FIG. Wing root 1
Of the main blade is cooled from the inside by flowing in the direction of the arrow to cool the main blade from the inside. After cooling air flowing from the inlet 2a on the leading edge side through convection cooling through the cooling passage,
It flows out of an outlet 5 at the tip of the blade where a tip thinning 4 is provided and joins the main gas flow. Further, the cooling air flowing in from the inlet 2b on the trailing edge side flows through the cooling passage in the direction of the arrow for convection cooling, then cools the blade trailing edge via the pin fins 6, and flows out from the outlet A to flow the main gas. To join.
【００１０】最近のガスタービンは高温化に伴って冷却
空気量が逐次増加してきており、この冷却空気の増加は
ガスタービンにおける熱効率を低下させる原因となり、
また冷却空気の圧力損失の原因となるため、本動翼にお
いては図１に示すように冷却通路の壁面に多数の渦発生
体８が壁面と一体に形成されている。この渦発生体８は
タービュレンスプロモータ（turbulence promoter ）と
称ばれ、上面が三角形状で流れに対して先端が鋭角をな
して低く冷却空気の流れの方向に高くなるように傾斜
し、後端には壁面に対して垂直な面が形成されており、
冷却空気の流れはこの渦発生体８により壁面に沿って冷
却空気の流れ方向及び流れの直角方向に湾曲され、渦発
生体８の後縁に２種類の渦を発生するようになってい
る。符号９ａは冷却空気の流れ方向に発生する渦、符号
９ｂは冷却空気の流れの直角方向に発生する渦である。
なお、渦発生体８の先端には或る程度の丸みが設けられ
ていてもよい。In recent gas turbines, the amount of cooling air has been gradually increasing as the temperature rises, and this increase in cooling air causes a decrease in thermal efficiency in the gas turbine.
Further, since this causes a pressure loss of the cooling air, in this rotor blade, as shown in FIG. 1, a large number of vortex generators 8 are formed integrally with the wall surface of the cooling passage. This vortex generator 8 is called a turbulence promoter, and has a triangular upper surface and a slanted tip that forms an acute angle with respect to the flow and becomes higher in the direction of the cooling air flow, and at the rear end. Has a surface perpendicular to the wall,
The flow of the cooling air is curved by the vortex generator 8 along the wall surface in the flow direction of the cooling air and the direction perpendicular to the flow, and two types of vortices are generated at the trailing edge of the vortex generator 8. Reference numeral 9a is a vortex generated in the flow direction of the cooling air, and reference numeral 9b is a vortex generated in the direction perpendicular to the flow of the cooling air.
The tip of the vortex generator 8 may be rounded to some extent.
【００１１】このように、渦発生体８により従来はフィ
ンにより冷却通路の流れ方向にのみ発生している渦が流
れの直角方向にも発生するようになり、従来は二次元的
な渦運動に支配されている乱流伝熱が三次元の渦運動に
より行われるように移行し、冷却通路の壁面から冷却空
気への熱伝達率が向上して少ない冷却空気量で動翼メタ
ルの温度が低減される。また、渦発生体８は従来のフィ
ンに比べて小形で冷却空気の通路面積が大きくなるとと
もに摩擦係数の小さい形状をしているので、冷却空気の
流れに対する抵抗が減少して圧力損失が低下する。これ
らにより、ガスタービンの高温化が行われるとともに、
ガスタービンにおける熱効率が向上する。なお、図２に
示すように冷却通路の壁面に従来のフィンと同様のフィ
ン３を渦発生体８とともに壁面と一体に併設してもよ
く、冷却空気と壁面との間の熱伝達率をさらに向上させ
ることができる。As described above, the vortex generator 8 causes vortices, which are conventionally generated only in the flow direction of the cooling passage by the fins, to be generated also in the direction perpendicular to the flow, which is conventionally a two-dimensional vortex motion. The dominant turbulent heat transfer is transferred as if it were performed by three-dimensional vortex motion, the heat transfer coefficient from the wall of the cooling passage to the cooling air is improved, and the temperature of the blade metal is reduced with a small amount of cooling air. It Further, since the vortex generator 8 is smaller than the conventional fin and has a large passage area for the cooling air and a small friction coefficient, the resistance against the flow of the cooling air is reduced and the pressure loss is reduced. .. With these, the temperature of the gas turbine is raised and
The thermal efficiency in the gas turbine is improved. As shown in FIG. 2, fins 3 similar to the conventional fins may be provided together with the vortex generator 8 on the wall surface of the cooling passage so that the heat transfer coefficient between the cooling air and the wall surface is further increased. Can be improved.
【００１２】図３は本発明の他の実施例に係るガスター
ビンの中空動翼の説明図である。図において、本実施例
に係るガスタービンの中空動翼は上記の実施例に係るガ
スタービンの中空動翼と用途、構造ともに略同一である
が、図に示すようにピンフィン６が冷却通路の両側壁７
ａ，７ｂから対称的にく字状に折れ曲がって突出し、冷
却空気の流れと対峙する方向に迎え角αで前傾して張り
出すように設けられており、この迎え角αは４５〜４０
°となっている。FIG. 3 is an explanatory view of a hollow rotor blade of a gas turbine according to another embodiment of the present invention. In the figure, the hollow blade of the gas turbine according to the present embodiment has substantially the same application and structure as the hollow blade of the gas turbine according to the above-mentioned embodiment, but as shown in the drawing, the pin fins 6 are provided on both sides of the cooling passage. Wall 7
It is provided so as to be bent symmetrically in a V shape from the a and 7b, and to project forward by inclining at an attack angle α in a direction facing the flow of the cooling air. The attack angle α is 45 to 40.
It has become °.
【００１３】このようにピンフィン６を傾斜させてその
先端を冷却空気の流れ内におくことにより、両側壁７
ａ，７ｂから冷却空気への熱伝達率が上昇する。実験結
果によると、ピンフィン６を迎え角４５〜４０°で前傾
させることにより熱伝達率は従来の中空動翼においてピ
ンフィンを直立させた場合に比べて１.７〜１.３倍とな
る。従って、同じ冷却空気量を流した場合に冷却効果は
略１.５倍となるが、このように熱伝達率が上昇するの
はピンフィン６の下流側にピンフィン６の先端から下流
の両側壁７ａ，７ｂに向けて略対称的に乱流境界層の剥
離による渦が発生して冷却通路の壁面から熱を奪って伝
達するためで、これにより冷却空気量を増やしてガスタ
ービンの熱効率を低下させることなく、少ない冷却空気
量で動翼メタルの温度を低下させることができる。ま
た、ピンフィン６は従来の円柱状のピンフィンに比べて
摩擦係数の小さい形状をしているので、冷却空気の流れ
に対する抵抗が減少して冷却空気の圧力損失が低下す
る。これらにより、ガスタービンの高温化と熱効率の向
上を図ることができる。By thus inclining the pin fins 6 and placing their tips in the flow of cooling air, the side walls 7
The heat transfer coefficient from a, 7b to the cooling air increases. According to the experimental results, when the pin fin 6 is tilted forward at an attack angle of 45 to 40 °, the heat transfer coefficient becomes 1.7 to 1.3 times that in the conventional hollow moving blade in which the pin fin is upright. Therefore, when the same amount of cooling air is flowed, the cooling effect is approximately 1.5 times, but the heat transfer coefficient rises in this way on the downstream side of the pin fins 6 on both side walls 7a downstream from the tips of the pin fins 6a. , 7b to generate vortices due to separation of the turbulent boundary layer in a substantially symmetrical manner, and heat is taken from the wall surface of the cooling passage to be transferred. This increases the amount of cooling air and reduces the thermal efficiency of the gas turbine. It is possible to lower the temperature of the blade metal with a small amount of cooling air. Further, since the pin fin 6 has a shape having a smaller friction coefficient than that of the conventional cylindrical pin fin, the resistance against the flow of the cooling air is reduced and the pressure loss of the cooling air is reduced. As a result, the temperature of the gas turbine can be increased and the thermal efficiency can be improved.
【００１４】[0014]
【発明の効果】本発明に係るガスタービン翼は前記のよ
うに構成されており、冷却通路の壁面から冷却気体への
熱伝達率が向上するので、少ない冷却気体量で翼メタル
の温度が低減される。また、冷却気体に対する流れ抵抗
が減少するので、冷却気体の圧力損失が低下する。これ
らにより、ガスタービンにおける高温化が促進される。The gas turbine blade according to the present invention is constructed as described above, and since the heat transfer coefficient from the wall surface of the cooling passage to the cooling gas is improved, the temperature of the blade metal is reduced with a small amount of cooling gas. To be done. Moreover, since the flow resistance to the cooling gas is reduced, the pressure loss of the cooling gas is reduced. These accelerate the temperature rise in the gas turbine.
【図１】図１（ａ）は本発明の一実施例に係るガスター
ビンの中空動翼における冷却通路の斜視図、同図（ｂ）
は断面図である。FIG. 1 (a) is a perspective view of a cooling passage in a hollow rotor blade of a gas turbine according to an embodiment of the present invention, and FIG. 1 (b).
Is a sectional view.
【図２】図２（ａ）はその応用例に係るガスタービンの
中空動翼における冷却通路の斜視図、同図（ｂ）は断面
図である。FIG. 2 (a) is a perspective view of a cooling passage in a hollow rotor blade of a gas turbine according to the application example, and FIG. 2 (b) is a sectional view.
【図３】図３は本発明の他の実施例に係るガスタービン
の中空動翼における冷却通路の断面図である。FIG. 3 is a sectional view of a cooling passage in a hollow rotor blade of a gas turbine according to another embodiment of the present invention.
【図４】図４（ａ）は従来のガスタービンの中空動翼の
一部破断斜視図、同図（ｂ）は同図（ａ）におけるｂ−
ｂ矢視図である。FIG. 4 (a) is a partially cutaway perspective view of a hollow blade of a conventional gas turbine, and FIG. 4 (b) is a sectional view taken along line b- in FIG. 4 (a).
FIG.
【図５】図５（ａ）はその冷却通路の斜視図、同図
（ｂ），（ｃ）は断面図である。5A is a perspective view of the cooling passage, and FIGS. 5B and 5C are cross-sectional views.
１ 翼根 ２ａ 入口 ２ｂ 入口 ３ フィン ４ チップシンニング ５ 出口 ６ ピンフィン ７ａ 側壁 ７ｂ 側壁 ８ 渦発生体 ９ａ 渦 ９ｂ 渦 Ａ 出口 1 Blade Root 2a Inlet 2b Inlet 3 Fin 4 Tip Thinning 5 Outlet 6 Pin Fin 7a Sidewall 7b Sidewall 8 Vortex Generator 9a Vortex 9b Vortex A Outlet
Claims (2)
流して翼メタルの冷却を行うガスタービン翼において、
上記冷却通路の壁面に突設され上記冷却気体の上流に向
けて低く傾斜するとともに下流側の端部が立面をなす渦
発生体を備えたことを特徴とするガスタービン翼。1. A gas turbine blade for cooling a blade metal by flowing a cooling gas through a cooling passage provided inside,
A gas turbine blade provided with a vortex generator which is provided on a wall surface of the cooling passage so as to be inclined downward toward the upstream side of the cooling gas and whose downstream end has an upright surface.
流して翼メタルの冷却を行うガスタービン翼において、
く字状に上記冷却通路の両側壁から折れ曲がって突設さ
れ上記冷却気体の上流に向けて張り出したピンフィンを
備えたことを特徴とするガスタービン翼。2. A gas turbine blade for cooling a blade metal by flowing a cooling gas through a cooling passage provided inside,
A gas turbine blade, comprising pin fins which are bent and protrude from both side walls of the cooling passage and project toward the upstream side of the cooling gas.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP4117461A JP3040590B2 (en) | 1992-05-11 | 1992-05-11 | Gas turbine blades |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP4117461A JP3040590B2 (en) | 1992-05-11 | 1992-05-11 | Gas turbine blades |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| JPH05312002A true JPH05312002A (en) | 1993-11-22 |
| JP3040590B2 JP3040590B2 (en) | 2000-05-15 |
Family
ID=14712255
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| JP4117461A Expired - Lifetime JP3040590B2 (en) | 1992-05-11 | 1992-05-11 | Gas turbine blades |
Country Status (1)
| Country | Link |
|---|---|
| JP (1) | JP3040590B2 (en) |
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|---|---|---|---|---|
| JP2007211670A (en) * | 2006-02-09 | 2007-08-23 | Hitachi Ltd | Member having cooling passage inside, and cooling method for member having cooling passage inside |
| JP2011140951A (en) * | 2010-01-06 | 2011-07-21 | General Electric Co <Ge> | Heat transfer enhancement in internal cavity of turbine engine airfoil part |
| JP2012002228A (en) * | 2011-08-30 | 2012-01-05 | Hitachi Ltd | Member including cooling passage therein |
| WO2015077017A1 (en) * | 2013-11-25 | 2015-05-28 | United Technologies Corporation | Gas turbine engine component cooling passage turbulator |
| CN106988790A (en) * | 2017-06-08 | 2017-07-28 | 哈尔滨工业大学 | To turning the cooling structure in whirlpool at the top of a kind of high-temperature turbine movable vane |
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| US9995148B2 (en) | 2012-10-04 | 2018-06-12 | General Electric Company | Method and apparatus for cooling gas turbine and rotor blades |
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1992
- 1992-05-11 JP JP4117461A patent/JP3040590B2/en not_active Expired - Lifetime
Cited By (10)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JP2007211670A (en) * | 2006-02-09 | 2007-08-23 | Hitachi Ltd | Member having cooling passage inside, and cooling method for member having cooling passage inside |
| US8292578B2 (en) | 2006-02-09 | 2012-10-23 | Hitachi, Ltd. | Material having internal cooling passage and method for cooling material having internal cooling passage |
| JP2011140951A (en) * | 2010-01-06 | 2011-07-21 | General Electric Co <Ge> | Heat transfer enhancement in internal cavity of turbine engine airfoil part |
| JP2012002228A (en) * | 2011-08-30 | 2012-01-05 | Hitachi Ltd | Member including cooling passage therein |
| US9995148B2 (en) | 2012-10-04 | 2018-06-12 | General Electric Company | Method and apparatus for cooling gas turbine and rotor blades |
| US9850762B2 (en) | 2013-03-13 | 2017-12-26 | General Electric Company | Dust mitigation for turbine blade tip turns |
| WO2015077017A1 (en) * | 2013-11-25 | 2015-05-28 | United Technologies Corporation | Gas turbine engine component cooling passage turbulator |
| EP3090145A4 (en) * | 2013-11-25 | 2017-09-13 | United Technologies Corporation | Gas turbine engine component cooling passage turbulator |
| US10364683B2 (en) | 2013-11-25 | 2019-07-30 | United Technologies Corporation | Gas turbine engine component cooling passage turbulator |
| CN106988790A (en) * | 2017-06-08 | 2017-07-28 | 哈尔滨工业大学 | To turning the cooling structure in whirlpool at the top of a kind of high-temperature turbine movable vane |
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| Publication number | Publication date |
|---|---|
| JP3040590B2 (en) | 2000-05-15 |
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