WO2014054825A1 - Film shape structure of cooling hole for cooling gas turbine blades - Google Patents

Film shape structure of cooling hole for cooling gas turbine blades Download PDF

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Publication number
WO2014054825A1
WO2014054825A1 PCT/KR2012/008093 KR2012008093W WO2014054825A1 WO 2014054825 A1 WO2014054825 A1 WO 2014054825A1 KR 2012008093 W KR2012008093 W KR 2012008093W WO 2014054825 A1 WO2014054825 A1 WO 2014054825A1
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Prior art keywords
cooling
hole
gas turbine
film
present
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PCT/KR2012/008093
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French (fr)
Korean (ko)
Inventor
김광용
이기돈
Original Assignee
Kim Kwang-Yong
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Priority to PCT/KR2012/008093 priority Critical patent/WO2014054825A1/en
Publication of WO2014054825A1 publication Critical patent/WO2014054825A1/en

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    • 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
    • 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
    • F05D2250/00—Geometry
    • F05D2250/30—Arrangement of components
    • F05D2250/32—Arrangement of components according to their shape
    • F05D2250/324—Arrangement of components according to their shape divergent
    • 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
    • F05D2260/202—Heat transfer, e.g. cooling by film cooling

Definitions

  • the present invention relates to the shape structure of the membrane cooling hole for cooling the gas turbine blade, and more particularly, to specify the shape of the membrane cooling hole used for the cooling of the gas turbine blade so that the cooling fluid is wider and more effective on the cooling surface.
  • the present invention relates to a shape structure of a membrane cooling hole for cooling of a gas turbine blade, which enables to distribute the gas turbine blade, thereby achieving a significantly improved cooling efficiency compared to the structure of a conventional hole.
  • gas turbine engines are designed to operate at 1,500 ⁇ 1700 ⁇ C, and the trend is to increase the turbine inlet temperature by 20 ⁇ C annually to increase thermal efficiency. to be.
  • the film-cooling method sprays the cooling fluid through holes formed at an angle with the blade surface.
  • this method is most commonly used because of the very effective cooling performance.
  • U.S. Patent Application Publication No. 2008/0031738 discloses a configuration of a bell-shaped film cooling hole.
  • the main technical configuration is an airfoil (22) in fluid communication with a turbine blade cooling circuit, as shown in FIG.
  • the shape of the conventional fan-shaped cooling hole 70 shown in FIG. 2 is also characterized in that the cooling hole includes a diffusion portion so that the cooling fluid passing through the cooling hole 70 can be diffused.
  • the shape of the conventional cooling hole 50 shown in FIG. 1 can be expected to some extent the diffusion effect in the longitudinal direction, but the diffusion effect in the width direction is inadequate and the overall diffusion effect of the cooling fluid is reduced,
  • Figure 2 The conventional fan-shaped cooling hole 70-shaped structure shown in Fig. 2 also has a poor diffusion effect in the width direction and the longitudinal direction, resulting in poor cooling performance, and a large number of cooling holes 70 are required as a whole. There was a problem that was much needed.
  • the present invention has been made to solve the above problems, the object of the present invention is to obtain a significantly improved cooling efficiency compared to the structure of the existing hole by specifying the shape of the film cooling hole used for the cooling of the gas turbine blades.
  • the present invention provides a structure of a film cooling hole for cooling a gas turbine blade.
  • the present invention can extend the life of the gas turbine blades due to the improved cooling performance, the use of less cooling fluid and reducing the number of holes can increase the efficiency of the engine blade cooling of the gas turbine blades
  • Another object is to provide a shape structure.
  • the cylindrical portion having a constant cross-sectional area, the expansion portion connected to the upper end of the cylindrical portion and the central portion of the hole outlet which is gradually widened so as to protrude in the flow direction It characterized in that it comprises a protrusion formed to protrude into the expansion pipe.
  • the film cooling hole is formed to have a slope of 30 ° in the gas turbine blade, when the diameter of the cylindrical portion, D, the length of the cylindrical portion and the expansion portion is characterized in that each of the 3D.
  • the expansion portion is characterized in that the tube is expanded so as to have an inclination of 25 ° in both side directions relative to the cylinder.
  • the protrusion may protrude in a direction in which the bottom portion extends by 2.5 ° in the hole exit direction based on the cylindrical portion.
  • the expansion portion and the protrusions protruding from the expansion portion have an excellent effect of obtaining the cooling efficiency of the gas turbine blades significantly improved compared to the structure of the existing hole.
  • the improved cooling performance can extend the life of the gas turbine blades, increase the efficiency of the engine by using less cooling fluid and reduce the number of holes, as well as the structural stability of the overall gas turbine. It further has an effect to improve.
  • 1 and 2 is a view showing the shape of a conventional film cooling hole.
  • Figure 3 is a perspective view showing a film cooling hole for cooling the gas turbine blade according to the present invention.
  • FIG. 4 is a plan view and a side cross-sectional view of a membrane cooling hole for cooling the gas turbine blade according to the present invention.
  • FIG 5 is a graph showing the film cooling efficiency averaged in the lateral direction on the cooling surface of the conventional fan-shaped hole shown in Figure 2 and the hole shape of the present invention.
  • FIG. 6 is a graph showing the film cooling efficiency averaged at the cooling surface of the conventional fan-shaped hole shown in FIG. 2 and the hole shape of the present invention.
  • 7 (a) and 7 (b) show the film cooling efficiency distribution on the cooling surface of the conventional fan-shaped hole shown in FIG. 2 and the hole shape of the present invention.
  • 8 (a) and 8 (b) are diagrams showing a conventional fan-shaped hole shown in FIG. 2 and a velocity vector at the hole-shaped outlet of the present invention
  • 9 (a) and 9 (b) show a streamline distribution of the conventional fan-shaped hole shown in FIG. 2 and a cooling fluid passing through the hole of the present invention.
  • FIG 3 is a perspective view showing a membrane cooling hole for cooling the gas turbine blade according to the present invention
  • Figure 4 is a plan view and a side cross-sectional view of the membrane cooling hole for cooling the gas turbine blade according to the present invention
  • Figure 5 Fig. 6 is a graph showing the film cooling efficiency averaged laterally from the cooling surface of the conventional fan-shaped hole and the hole shape of the present invention shown in Fig. 6 is shown in the conventional fan-shaped hole shown in Fig. 2 and the hole shape of the present invention
  • 7 is a graph showing the film cooling efficiency averaged at the cooling surface
  • FIGS. 7A and 7B show the distribution of film cooling efficiency at the cooling surface of the conventional fan-shaped hole shown in FIG. 2 and the hole shape of the present invention.
  • FIGS. 8 (a) and 8 (b) show the conventional fan-shaped hole shown in FIG. 2 and the velocity vector at the hole-shaped outlet of the present invention
  • FIGS. ) Is a conventional fan-shaped hole shown in Figure 2 and the present invention
  • the present invention is to specify the shape of the membrane cooling holes used for the cooling of the gas turbine blades so that the cooling fluid can be distributed more widely and effectively on the cooling surface to obtain a significantly improved cooling efficiency compared to the structure of the existing holes
  • the configuration is composed of a cylindrical portion 110, expansion tube 120 and the protrusion 130 as shown in FIG.
  • the cylindrical portion 110 serves as an inlet portion in which the cooling fluid flows, and as shown in FIG. 4, when the width of the gas turbine blade 80 is 3D, the diameter is cylindrical. It is made of a phase, and is formed to have a length of 3D to have a slope of 30 ° in the direction of travel of the cooling fluid.
  • the cooling fluid supplied from the lower portion of the blade 80 flows into the cylindrical portion 110 having a narrow diameter, and the flow velocity thereof is increased.
  • the expansion part 120 is connected to the upper portion of the cylindrical portion 110 so that its cross-sectional area is gradually widened so that the cooling fluid flowing from the cylindrical portion 110 can be spread widely when discharged through the outlet portion of the range of cooling It is to serve to expand the, it is expanded to have an inclination of 25 ° in both side directions when the cylindrical portion 110 relative to the reference, the overall length is configured to be 3D.
  • the protrusion 130 is formed on the bottom surface of the expansion pipe 120 to protrude in the advancing direction of the cooling fluid, and serves to distribute the cooling fluid evenly on the surface of the blade 80.
  • the protrusion 130 is formed to protrude in the direction in which the bottom portion is extended by 2.5 ° in the exit direction of the hole 100 with respect to the cylindrical part 110, and thus the expansion part 120 is formed.
  • ANSYS CFX-11.0 a commercial computational fluid dynamics code, was used, and a hexahedral lattice and shear stress transport turbulence model was used.
  • the working fluid is air (ideal gas, air).
  • boundary conditions heat insulation and adhesion conditions are applied to the walls, and flow conditions are applied to the inlet of the cooling fluid supply passage.
  • High pressure condition 100,400 Pa
  • static pressure condition 93,500 Pa
  • the Mach number of hot gas was 0.3
  • the temperature of hot gas and cooling fluid was 540K and 310K, respectively.
  • the cooling performance on the film cooling surface was evaluated by calculating the film cooling efficiency averaged in the lateral direction and the area according to the change of the injection rate (M). ) And injection rate M are respectively defined as follows.
  • FIG. 5 is a film cooling efficiency averaged in the lateral direction on the cooling surface of the conventional fan-shaped hole 70 and the film cooling hole 100 of the present invention shown in FIG.
  • the numerical analysis results of the fan-shaped hole 70 shows good agreement with the experimental results, and the film cooling hole 100 according to the present invention has a spray rate of 0.5 compared with the fan-shaped hole 70.
  • Film cooling efficiency Increased slightly but the injection rate was 2.5 Can be seen.
  • FIG. 6 shows the film cooling efficiency averaged at the cooling surface of the conventional fan-shaped hole 70 shown in FIG. 2 and the shape of the film cooling hole 100 of the present invention.
  • the film cooling hole 100 according to the present invention has a high film cooling efficiency (compared with the fan-shaped hole 70). ), Especially as the injection rate increases, the film cooling efficiency (higher than the conventional fan-shaped hole 70) is significantly higher ( Can be seen.
  • FIG. 7 (a) and 7 (b) show the film cooling efficiency at the cooling surface of the conventional fan-shaped hole 70 shown in FIG. 2 and the shape of the hole 100 of the present invention when the injection rates are 1.5 and 2.5.
  • the film cooling hole 100 according to the present invention shows the distribution of cooling fluid evenly and widely on the surface, and shows an extremely effective cooling efficiency. It can be seen that as the injection rate increases.
  • 8 (a) and 8 (b) and 9 (a) and 9 (b) are velocity vectors at the exit shape of the conventional fan shape hole 70 shown in FIG. 2 and the hole 100 of the present invention, and It shows a streamline distribution of the cooling fluid, the conventional fan-shaped hole 70 shows a low film cooling efficiency in the central portion of the cooling surface, the cooling fluid is distributed in a narrow region, the film cooling hole 100 according to the present invention In the case of not only shows high cooling efficiency in the central part, but also shows that the cooling fluid is evenly distributed over the entire surface.
  • the structure of the existing hole by the expansion portion 120 and the protrusions 130 protruding from the expansion portion 120 Not only can the cooling efficiency of the gas turbine blade 80 be significantly improved compared to that of the gas turbine blade 80, and thus, the cooling performance of the gas turbine blade can be extended due to the improved cooling performance, the use of less cooling fluid, and the entire membrane cooling hole (100).
  • By reducing the number of installation) can increase the efficiency of the engine and at the same time have various advantages, such as to improve the structural stability of the overall gas turbine.
  • the present invention relates to the shape structure of the membrane cooling hole for cooling the gas turbine blade, and more particularly, to specify the shape of the membrane cooling hole used for the cooling of the gas turbine blade so that the cooling fluid is wider and more effective on the cooling surface.
  • the present invention relates to a shape structure of a membrane cooling hole for cooling of a gas turbine blade, which enables to distribute the gas turbine blade, thereby achieving a significantly improved cooling efficiency compared to the structure of a conventional hole.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Abstract

The present invention relates to a film shape structure of a cooling hole for cooling gas turbine blades, wherein the shape of a film cooling hole which is used for the purpose of cooling gas turbine blades is specifically set so as to more widely and effectively distribute cooling fluid throughout a cooling surface such that remarkably improved cooling efficiency can be obtained when compared with existing hole structures. The present invention relates to the film shape structure of a film cooling hole for cooling gas turbine blades, comprising: a cylindrical part having a predetermined cross-section area, an expansion part connected to the upper end of the cylindrical part and having a cross-section area which gradually increases, and a protrusion part protruding in the center portion of a hole outlet so as to protrude to the expansion part.

Description

가스터빈 블레이드의 냉각을 위한 막냉각 홀의 형상 구조Shape structure of membrane cooling hole for cooling gas turbine blade
본 발명은 가스터빈 블레이드의 냉각을 위한 막냉각 홀의 형상 구조에 관한 것으로, 보다 상세하게는 가스터빈 블레이드의 냉각을 위한 목적으로 사용되는 막냉각 홀의 형상을 특정하여 냉각 유체가 냉각 표면에 보다 넓고 효과적으로 분포할 수 있도록 함으로써 기존의 홀의 구조에 비해 월등히 향상된 냉각 효율을 얻을 수 있도록 하는 가스터빈 블레이드의 냉각을 위한 막냉각 홀의 형상 구조에 관한 것이다.The present invention relates to the shape structure of the membrane cooling hole for cooling the gas turbine blade, and more particularly, to specify the shape of the membrane cooling hole used for the cooling of the gas turbine blade so that the cooling fluid is wider and more effective on the cooling surface. The present invention relates to a shape structure of a membrane cooling hole for cooling of a gas turbine blade, which enables to distribute the gas turbine blade, thereby achieving a significantly improved cooling efficiency compared to the structure of a conventional hole.
일반적으로, 가스터빈 엔진의 효율 및 성능을 높이기 위해 최근 가스터빈 엔진은 1,500~1700ㅇC에서 작동되도록 설계되고 있으며, 열효율을 더욱 높이기 위해 터빈 입구온도를 연평균 20ㅇC씩 꾸준히 상승시켜 설계하는 추세이다.In general, in order to increase the efficiency and performance of gas turbine engines, gas turbine engines are designed to operate at 1,500 ~ 1700 ㅇ C, and the trend is to increase the turbine inlet temperature by 20 ◦ C annually to increase thermal efficiency. to be.
따라서 높은 입구온도로부터 터빈 블레이드를 보호하기 위해 다양한 냉각기법들이 연구 및 개발되고 있는데, 그 중 막냉각(film-cooling) 방법은 블레이드 표면과 일정한 각도를 이루는 홀(hole)을 통해 냉각유체를 분사하여 블레이드 표면에 막을 형성함으로써 고온의 주유동가스로부터 표면을 보호하는 방법으로, 이 방법은 매우 효과적인 냉각성능으로 인해 가장 보편적으로 사용되고 있다.Therefore, various cooling techniques have been researched and developed to protect turbine blades from high inlet temperatures. Among them, the film-cooling method sprays the cooling fluid through holes formed at an angle with the blade surface. By forming a film on the blade surface to protect the surface from the hot main flow gas, this method is most commonly used because of the very effective cooling performance.
막냉각을 위해서는 압축기로부터 추출된 고압의 냉각공기가 사용되므로 과도한 양의 압축공기의 사용은 가스터빈의 효율을 감소시키므로 효과적인 냉각방식의 필요성이 대두 되고 있으며, 홀의 형상은 막냉각 효율에 크게 영향을 끼치기 때문에 막냉각 효율을 높이기 위해 다양한 홀의 형상이 개발되고 있는 실정이다.Since the high-pressure cooling air extracted from the compressor is used for the film cooling, the use of an excessive amount of compressed air reduces the efficiency of the gas turbine, so the necessity of an effective cooling method has emerged. In order to increase the film cooling efficiency, various hole shapes are being developed.
그 중 미국공개특허공보 제2008/0031738호에는 벨 형상의 막냉각 홀의 구성이 개시되어 있는데, 그 주요 기술적 구성은 도 1에 나타낸 바와 같이, 터빈 날개부 냉각 회로를 가진 유체 연통에서 에어포일(22)의 외표면에서 형성된 다수의 냉각홀(50)에 관한 것으로, 냉각홀(50)은 계량부(58)와 에어포일(50)의 외측 표면으로 개구된 확산부(51)를 포함하여 전체적으로 벨형상을 이루도록 구성되어 높은 냉각효율을 갖도록 한 것에 특징이 있다.Among them, U.S. Patent Application Publication No. 2008/0031738 discloses a configuration of a bell-shaped film cooling hole. The main technical configuration is an airfoil (22) in fluid communication with a turbine blade cooling circuit, as shown in FIG. A plurality of cooling holes 50 formed on the outer surface of the), the cooling holes 50, including the metering portion 58 and the diffusion portion 51 opening to the outer surface of the air foil 50 as a whole bell It is characterized by being configured to have a shape to have a high cooling efficiency.
또한, 도 2에 나타낸 종래의 팬 형상의 냉각홀(70)의 형상 또한, 냉각홀이 확산부를 포함하여 냉각홀(70)을 통과하는 냉각유체가 확산될 수 있도록 한 것에 특징이 있다.In addition, the shape of the conventional fan-shaped cooling hole 70 shown in FIG. 2 is also characterized in that the cooling hole includes a diffusion portion so that the cooling fluid passing through the cooling hole 70 can be diffused.
하지만, 상기 도 1에 나타낸 종래의 냉각홀(50) 형상은 길이방향으로의 확산 효과는 어느 정도 기대할 수 있으나, 폭방향으로의 확산 효과가 미비하여 전체적인 냉각유체의 확산효과가 떨어지게 되고, 도 2에 나타낸 종래의 팬 형상의 냉각홀(70) 형상 구조 또한 폭방향 및 길이방향으로의 확산 효과가 떨어져 냉각 성능이 떨어지게 되고, 전체적으로 많은 수의 냉각홀(70)이 필요하게 되어 냉각유체의 사용량이 많이 필요하게 되는 문제점이 있었다. However, the shape of the conventional cooling hole 50 shown in FIG. 1 can be expected to some extent the diffusion effect in the longitudinal direction, but the diffusion effect in the width direction is inadequate and the overall diffusion effect of the cooling fluid is reduced, Figure 2 The conventional fan-shaped cooling hole 70-shaped structure shown in Fig. 2 also has a poor diffusion effect in the width direction and the longitudinal direction, resulting in poor cooling performance, and a large number of cooling holes 70 are required as a whole. There was a problem that was much needed.
본 발명은 상기와 같은 문제점들을 해결하기 위하여 안출된 것으로, 본 발명의 목적은 가스터빈 블레이드의 냉각을 위한 목적으로 사용되는 막냉각 홀의 형상을 특정함으로써 기존의 홀의 구조에 비해 월등히 향상된 냉각 효율을 얻을 수 있는 가스터빈 블레이드의 냉각을 위한 막냉각 홀의 형상 구조를 제공함에 있다.The present invention has been made to solve the above problems, the object of the present invention is to obtain a significantly improved cooling efficiency compared to the structure of the existing hole by specifying the shape of the film cooling hole used for the cooling of the gas turbine blades. The present invention provides a structure of a film cooling hole for cooling a gas turbine blade.
또한, 본 발명은 향상된 냉각성능으로 인해 가스터빈 블레이드의 수명을 연장시킬 수 있고, 적은 냉각유체의 사용 및 홀의 개수를 줄임으로써 엔진의 효율을 상승시킬 수 있는 가스터빈 블레이드의 냉각을 위한 막냉각 홀의 형상 구조를 제공함에 다른 목적이 있다.In addition, the present invention can extend the life of the gas turbine blades due to the improved cooling performance, the use of less cooling fluid and reducing the number of holes can increase the efficiency of the engine blade cooling of the gas turbine blades Another object is to provide a shape structure.
상기한 바와 같은 목적들을 달성하기 위한 본 발명은,The present invention for achieving the above objects,
가스터빈 블레이드의 냉각을 위한 막냉각 홀의 형상 구조에 있어서, 일정한 단면적을 갖는 원통부와, 상기 원통부의 상단에 연결되어 그 단면적이 점점 넓어지는 확관부 및 홀 출구의 중앙부분이 유동방향으로 돌출되도록 하여 확관부에 돌출 형성되는 돌출부를 포함하여 구성된 것을 특징으로 한다.In the shape structure of the membrane cooling hole for cooling the gas turbine blade, the cylindrical portion having a constant cross-sectional area, the expansion portion connected to the upper end of the cylindrical portion and the central portion of the hole outlet which is gradually widened so as to protrude in the flow direction It characterized in that it comprises a protrusion formed to protrude into the expansion pipe.
이때, 상기 막냉각 홀은 가스터빈 블레이드에 30°의 기울기를 갖도록 형성되고, 원통부의 직경을 D라고 했을 때, 원통부와 확관부의 길이는 각각 3D인 것을 특징으로 한다.At this time, the film cooling hole is formed to have a slope of 30 ° in the gas turbine blade, when the diameter of the cylindrical portion, D, the length of the cylindrical portion and the expansion portion is characterized in that each of the 3D.
또한, 상기 확관부는 원통부를 기준으로 하여 양 측면방향으로 25°의 기울기를 갖도록 확관되는 것을 특징으로 한다.In addition, the expansion portion is characterized in that the tube is expanded so as to have an inclination of 25 ° in both side directions relative to the cylinder.
그리고, 상기 돌출부는 원통부를 기준으로 하여 홀 출구 방향으로 저면부분이 2.5°만큼 확관되는 방향으로 돌출 형성된 것을 특징으로 한다.The protrusion may protrude in a direction in which the bottom portion extends by 2.5 ° in the hole exit direction based on the cylindrical portion.
본 발명에 따르면, 확관부 및 확관부에 돌출 형성되는 돌출부에 의해 기존의 홀의 구조에 비해 월등히 향상된 가스터빈 블레이드의 냉각 효율을 얻을 수 있는 뛰어난 효과를 갖는다.According to the present invention, the expansion portion and the protrusions protruding from the expansion portion have an excellent effect of obtaining the cooling efficiency of the gas turbine blades significantly improved compared to the structure of the existing hole.
또한, 본 발명에 따르면 향상된 냉각성능으로 인해 가스터빈 블레이드의 수명을 연장시킬 수 있고, 적은 냉각유체의 사용 및 홀의 개수를 줄임으로써 엔진의 효율을 상승시킬 수 있을 뿐만 아니라, 전체적인 가스터빈의 구조적 안정성을 향상시킬 수 있는 효과를 추가로 갖는다.In addition, according to the present invention, the improved cooling performance can extend the life of the gas turbine blades, increase the efficiency of the engine by using less cooling fluid and reduce the number of holes, as well as the structural stability of the overall gas turbine. It further has an effect to improve.
도 1 및 도 2는 종래의 막냉각 홀의 형상을 나타낸 도면.1 and 2 is a view showing the shape of a conventional film cooling hole.
도 3은 본 발명에 따른 가스터빈 블레이드의 냉각을 위한 막냉각 홀을 나타낸 사시도.Figure 3 is a perspective view showing a film cooling hole for cooling the gas turbine blade according to the present invention.
도 4는 본 발명에 따른 가스터빈 블레이드의 냉각을 위한 막냉각 홀의 평면도 및 측단면도.4 is a plan view and a side cross-sectional view of a membrane cooling hole for cooling the gas turbine blade according to the present invention.
도 5는 도 2에 나타낸 종래의 팬 형상 홀과 본 발명의 홀 형상에 대한 냉각표면에서 측면방향으로 평균한 막냉각 효율을 나타낸 그래프.5 is a graph showing the film cooling efficiency averaged in the lateral direction on the cooling surface of the conventional fan-shaped hole shown in Figure 2 and the hole shape of the present invention.
도 6은 도 2에 나타낸 종래의 팬 형상 홀과 본 발명의 홀 형상에 대한 냉각표면에서 평균한 막냉각 효율을 나타낸 그래프.6 is a graph showing the film cooling efficiency averaged at the cooling surface of the conventional fan-shaped hole shown in FIG. 2 and the hole shape of the present invention.
도 7의 (a),(b)는 도 2에 나타낸 종래의 팬 형상 홀과 본 발명의 홀 형상에 대한 냉각표면에서의 막냉각 효율 분포를 나타낸 도면.7 (a) and 7 (b) show the film cooling efficiency distribution on the cooling surface of the conventional fan-shaped hole shown in FIG. 2 and the hole shape of the present invention.
도 8의 (a),(b)는 도 2에 나타낸 종래의 팬 형상 홀과 본 발명의 홀 형상 출구에서의 속도벡터를 나타낸 도면.8 (a) and 8 (b) are diagrams showing a conventional fan-shaped hole shown in FIG. 2 and a velocity vector at the hole-shaped outlet of the present invention;
도 9의 (a),(b)는 도 2에 나타낸 종래의 팬 형상 홀과 본 발명의 홀을 통과하는 냉각유체의 유선 분포를 나타낸 도면.9 (a) and 9 (b) show a streamline distribution of the conventional fan-shaped hole shown in FIG. 2 and a cooling fluid passing through the hole of the present invention.
이하, 첨부된 도면을 참고로 하여 본 발명에 따른 가스터빈 블레이드의 냉각을 위한 막냉각 홀의 형상 구조의 바람직한 실시예에 대하여 상세히 설명하기로 한다.Hereinafter, with reference to the accompanying drawings will be described in detail a preferred embodiment of the shape structure of the film cooling hole for cooling the gas turbine blade according to the present invention.
도 3은 본 발명에 따른 가스터빈 블레이드의 냉각을 위한 막냉각 홀을 나타낸 사시도이고, 도 4는 본 발명에 따른 가스터빈 블레이드의 냉각을 위한 막냉각 홀의 평면도 및 측단면도이며, 도 5는 도 2에 나타낸 종래의 팬 형상 홀과 본 발명의 홀 형상에 대한 냉각표면에서 측면방향으로 평균한 막냉각 효율을 나타낸 그래프이고, 도 6은 도 2에 나타낸 종래의 팬 형상 홀과 본 발명의 홀 형상에 대한 냉각표면에서 평균한 막냉각 효율을 나타낸 그래프이며, 도 7의 (a),(b)는 도 2에 나타낸 종래의 팬 형상 홀과 본 발명의 홀 형상에 대한 냉각표면에서의 막냉각 효율 분포를 나타낸 도면이고, 도 8의 (a),(b)는 도 2에 나타낸 종래의 팬 형상 홀과 본 발명의 홀 형상 출구에서의 속도벡터를 나타낸 도면이며, 도 9의 (a),(b)는 도 2에 나타낸 종래의 팬 형상 홀과 본 발명의 홀을 통과하는 냉각유체의 유선 분포를 나타낸 도면이다.3 is a perspective view showing a membrane cooling hole for cooling the gas turbine blade according to the present invention, Figure 4 is a plan view and a side cross-sectional view of the membrane cooling hole for cooling the gas turbine blade according to the present invention, Figure 5 Fig. 6 is a graph showing the film cooling efficiency averaged laterally from the cooling surface of the conventional fan-shaped hole and the hole shape of the present invention shown in Fig. 6 is shown in the conventional fan-shaped hole shown in Fig. 2 and the hole shape of the present invention. 7 is a graph showing the film cooling efficiency averaged at the cooling surface, and FIGS. 7A and 7B show the distribution of film cooling efficiency at the cooling surface of the conventional fan-shaped hole shown in FIG. 2 and the hole shape of the present invention. 8 (a) and 8 (b) show the conventional fan-shaped hole shown in FIG. 2 and the velocity vector at the hole-shaped outlet of the present invention, and FIGS. ) Is a conventional fan-shaped hole shown in Figure 2 and the present invention A diagram showing a streamline distribution of a cooling fluid passing through a hole.
본 발명은 가스터빈 블레이드의 냉각을 위한 목적으로 사용되는 막냉각 홀의 형상을 특정하여 냉각 유체가 냉각 표면에 보다 넓고 효과적으로 분포할 수 있도록 함으로써 기존의 홀의 구조에 비해 월등히 향상된 냉각 효율을 얻을 수 있도록 하는 가스터빈 블레이드의 냉각을 위한 막냉각 홀(100)의 형상 구조에 관한 것으로, 그 구성은 도 3에 나타낸 바와 같이, 크게 원통부(110), 확관부(120) 및 돌출부(130)로 이루어진다.The present invention is to specify the shape of the membrane cooling holes used for the cooling of the gas turbine blades so that the cooling fluid can be distributed more widely and effectively on the cooling surface to obtain a significantly improved cooling efficiency compared to the structure of the existing holes It relates to the shape structure of the film cooling hole 100 for cooling the gas turbine blade, the configuration is composed of a cylindrical portion 110, expansion tube 120 and the protrusion 130 as shown in FIG.
보다 상세히 설명하면, 상기 원통부(110)는 냉각유체가 유입되는 입구부 역할을 하는 것으로, 도 4에 나타낸 바와 같이, 가스터빈 블레이드(80)의 폭을 3D라 하였을 때, 직경이 D인 원통형상으로 이루어지고, 냉각유체의 진행방향으로 30°의 기울기를 갖도록 하여 3D의 길이를 갖도록 형성된다.In more detail, the cylindrical portion 110 serves as an inlet portion in which the cooling fluid flows, and as shown in FIG. 4, when the width of the gas turbine blade 80 is 3D, the diameter is cylindrical. It is made of a phase, and is formed to have a length of 3D to have a slope of 30 ° in the direction of travel of the cooling fluid.
따라서, 블레이드(80)의 하부로부터 공급되는 냉각 유체는 직경이 좁은 원통부(110) 내로 유입되어 그 유속이 빨라지게 된다.Therefore, the cooling fluid supplied from the lower portion of the blade 80 flows into the cylindrical portion 110 having a narrow diameter, and the flow velocity thereof is increased.
다음, 상기 확관부(120)는 원통부(110)의 상부에 그 단면적이 점점 넓어지도록 연결되어 원통부(110)로부터 유입되는 냉각유체가 출구부를 통해 배출될 때 넓게 퍼질 수 있도록 하여 냉각의 범위를 확장시킬 수 있도록 하는 역할을 하는 것으로, 원통부(110)를 기준으로 하였을 때 양 측면 방향으로 25°의 기울기를 갖도록 확관되고, 그 전체적인 길이는 3D가 되도록 구성되어 있다.Next, the expansion part 120 is connected to the upper portion of the cylindrical portion 110 so that its cross-sectional area is gradually widened so that the cooling fluid flowing from the cylindrical portion 110 can be spread widely when discharged through the outlet portion of the range of cooling It is to serve to expand the, it is expanded to have an inclination of 25 ° in both side directions when the cylindrical portion 110 relative to the reference, the overall length is configured to be 3D.
다음, 상기 돌출부(130)는 냉각유체의 진행방향으로 돌출되도록 하여 확관부(120)의 저면에 형성된 것으로, 냉각유체를 블레이드(80) 표면에 고르게 분포시키는 역할을 하게 된다.Next, the protrusion 130 is formed on the bottom surface of the expansion pipe 120 to protrude in the advancing direction of the cooling fluid, and serves to distribute the cooling fluid evenly on the surface of the blade 80.
즉, 상기 돌출부(130)는 도 4에 나타낸 바와 같이, 원통부(110)를 기준으로 하여 홀(100)의 출구 방향으로 저면 부분이 2.5°만큼 확관되는 방향으로 돌출 형성되어 확관부(120)를 통과하는 냉각유체가 길이방향으로도 넓게 분포되도록 함으로써 보다 효율적인 블레이드(80)의 냉각이 이루어질 수 있도록 하는 것이다.That is, as shown in FIG. 4, the protrusion 130 is formed to protrude in the direction in which the bottom portion is extended by 2.5 ° in the exit direction of the hole 100 with respect to the cylindrical part 110, and thus the expansion part 120 is formed. By allowing the cooling fluid to pass through a wide distribution in the longitudinal direction it is possible to achieve a more efficient cooling of the blade (80).
상기와 같은 형상적 특징을 갖는 막냉각 홀(100)의 냉각성능을 평가하기 위해서, 유동장과 막냉각에 대한 수치해석을 수행하였다.In order to evaluate the cooling performance of the film cooling hole 100 having the above-described features, numerical analysis of the flow field and the film cooling was performed.
삼차원 Reynolds-averaged Navier-Stokes equation을 풀기 위해 상용전산유체역학 코드인 ANSYS CFX-11.0을 사용하였으며, 육면체 격자 및 shear stress transport 난류모델을 사용하였다. To solve the three-dimensional Reynolds-averaged Navier-Stokes equation, ANSYS CFX-11.0, a commercial computational fluid dynamics code, was used, and a hexahedral lattice and shear stress transport turbulence model was used.
작동유체는 공기(이상기체(ideal gas), air)이며, 경계조건으로는 벽면에 단열조건과 점착조건을 적용하였고, 냉각유체 공급유로의 입구에는 유량조건을 부여하였다. 주유로의 입구에는 전압력 조건(100,400 Pa)을, 출구에는 정압조건(93,500 Pa)을
Figure PCTKR2012008093-appb-I000001
주었으며, 고온가스의 마하수는 0.3을, 고온가스와 냉각유체의 온도는 각각 540K, 310K을 적용하였다.The working fluid is air (ideal gas, air). As boundary conditions, heat insulation and adhesion conditions are applied to the walls, and flow conditions are applied to the inlet of the cooling fluid supply passage. High pressure condition (100,400 Pa) at the inlet of the main passage and static pressure condition (93,500 Pa) at the outlet
Figure PCTKR2012008093-appb-I000001
The Mach number of hot gas was 0.3, and the temperature of hot gas and cooling fluid was 540K and 310K, respectively.
분사율(M)의 변화에 따른 측면방향 및 면적 평균한 막냉각 효율을 계산하여 막냉각 표면에서의 냉각성능을 평가하였으며, 막냉각 효율(
Figure PCTKR2012008093-appb-I000002
) 및 분사율(M)은 각각 다음과 같이 정의된다.The cooling performance on the film cooling surface was evaluated by calculating the film cooling efficiency averaged in the lateral direction and the area according to the change of the injection rate (M).
Figure PCTKR2012008093-appb-I000002
) And injection rate M are respectively defined as follows.
Figure PCTKR2012008093-appb-I000003
(1)
Figure PCTKR2012008093-appb-I000003
(One)
Figure PCTKR2012008093-appb-I000004
(2)
Figure PCTKR2012008093-appb-I000004
(2)
여기서
Figure PCTKR2012008093-appb-I000005
는 단열벽면온도를 의미하며,
Figure PCTKR2012008093-appb-I000006
와
Figure PCTKR2012008093-appb-I000007
는 각각 주유동과 냉각유체의 분사온도를 나타낸다. 또한
Figure PCTKR2012008093-appb-I000008
와
Figure PCTKR2012008093-appb-I000009
는 각각 냉각 유체가 분사되는 지점에서 측정한 냉각유체의 밀도와 속도를 나타내며,
Figure PCTKR2012008093-appb-I000010
와
Figure PCTKR2012008093-appb-I000011
는 고온가스의 입구부근에서 측정한 밀도와 속도를 의미한다. 그리고, D는 본 발명에 따른 막냉각 홀(100)의 원통부(100) 직경이고,
Figure PCTKR2012008093-appb-I000012
와
Figure PCTKR2012008093-appb-I000013
는 각각 냉각 유체가 분사되는 지점으로부터의
Figure PCTKR2012008093-appb-I000014
방향, 즉 분사방향 및
Figure PCTKR2012008093-appb-I000015
방향, 즉 측면방향으로부터의 거리를 나타낸다.here
Figure PCTKR2012008093-appb-I000005
Means the insulation wall temperature,
Figure PCTKR2012008093-appb-I000006
Wow
Figure PCTKR2012008093-appb-I000007
Are the injection temperatures of the main and cooling fluids, respectively. Also
Figure PCTKR2012008093-appb-I000008
Wow
Figure PCTKR2012008093-appb-I000009
Represents the density and velocity of the cooling fluid, respectively, measured at the point where the cooling fluid is injected,
Figure PCTKR2012008093-appb-I000010
Wow
Figure PCTKR2012008093-appb-I000011
Is the density and velocity measured near the inlet of hot gas. And, D is the diameter of the cylindrical portion 100 of the film cooling hole 100 according to the present invention,
Figure PCTKR2012008093-appb-I000012
Wow
Figure PCTKR2012008093-appb-I000013
From the point where the cooling fluid is injected
Figure PCTKR2012008093-appb-I000014
Direction, that is, injection direction and
Figure PCTKR2012008093-appb-I000015
Direction, that is, distance from the lateral direction.
도 5는 도 2에 나타낸 종래의 팬 형상 홀(70)과 본 발명의 막냉각 홀(100) 형상에 대한 냉각표면에서 측면방향으로 평균한 막냉각 효율(
Figure PCTKR2012008093-appb-I000016
)을 나타낸 것으로, 팬 형상 홀(70)의 수치해석결과는 실험결과와 좋은 일치성을 보여주며, 본 발명에 따른 막냉각 홀(100)은 팬 형상 홀(70)과 비교하여 분사율이 0.5일 때는 막냉각 효율(
Figure PCTKR2012008093-appb-I000017
)이 미세하게 증가하였으나 분사율이 2.5일 때는 월등하게 높은 막냉각 효율(
Figure PCTKR2012008093-appb-I000018
)을 보임을 확인할 수 있다.5 is a film cooling efficiency averaged in the lateral direction on the cooling surface of the conventional fan-shaped hole 70 and the film cooling hole 100 of the present invention shown in FIG.
Figure PCTKR2012008093-appb-I000016
The numerical analysis results of the fan-shaped hole 70 shows good agreement with the experimental results, and the film cooling hole 100 according to the present invention has a spray rate of 0.5 compared with the fan-shaped hole 70. Film cooling efficiency
Figure PCTKR2012008093-appb-I000017
) Increased slightly but the injection rate was 2.5
Figure PCTKR2012008093-appb-I000018
Can be seen.
도 6은 도 2에 나타낸 종래의 팬 형상 홀(70)과 본 발명의 막냉각 홀(100) 형상에 대한 냉각표면에서 평균한 막냉각 효율(
Figure PCTKR2012008093-appb-I000019
)을 나타낸 것으로, 본 발명에 따른 막냉각 홀(100)은 팬 형상 홀(70)과 비교하였을 때 높은 막냉각 효율(
Figure PCTKR2012008093-appb-I000020
)을 보이는 것을 확인할 수 있고, 특히 분사율이 증가함에 따라 기존에 사용되어지는 팬 형상 홀(70)과 비교하여 월등히 높은 막냉각 효율(
Figure PCTKR2012008093-appb-I000021
)을 보임을 알 수 있다.FIG. 6 shows the film cooling efficiency averaged at the cooling surface of the conventional fan-shaped hole 70 shown in FIG. 2 and the shape of the film cooling hole 100 of the present invention.
Figure PCTKR2012008093-appb-I000019
), The film cooling hole 100 according to the present invention has a high film cooling efficiency (compared with the fan-shaped hole 70).
Figure PCTKR2012008093-appb-I000020
), Especially as the injection rate increases, the film cooling efficiency (higher than the conventional fan-shaped hole 70) is significantly higher (
Figure PCTKR2012008093-appb-I000021
Can be seen.
도 7의 (a),(b)는 분사율이 1.5와 2.5일 때, 도 2에 나타낸 종래의 팬 형상 홀(70)과 본 발명의 홀(100) 형상에 대한 냉각표면에서의 막냉각 효율 분포를 나타낸 것으로, 종래의 팬 형상 홀(70)과 비교하였을 때, 본 발명에 따른 막냉각 홀(100)의 경우 냉각유체가 표면에 고르며 넓게 분포하며 월등히 효과적인 냉각효율을 보여주며, 이러한 효과는 분사율이 커짐에 따라 증가함을 확인할 수 있다.7 (a) and 7 (b) show the film cooling efficiency at the cooling surface of the conventional fan-shaped hole 70 shown in FIG. 2 and the shape of the hole 100 of the present invention when the injection rates are 1.5 and 2.5. Compared with the conventional fan-shaped hole 70, the film cooling hole 100 according to the present invention shows the distribution of cooling fluid evenly and widely on the surface, and shows an extremely effective cooling efficiency. It can be seen that as the injection rate increases.
도 8의 (a),(b)와, 도 9의 (a),(b)는 도 2에 나타낸 종래의 팬 형상 홀(70)과 본 발명의 홀(100) 형상 출구에서의 속도벡터 및 냉각유체의 유선 분포를 나타낸 것으로, 종래의 팬 형상 홀(70)은 냉각 면의 중앙부분에서 낮은 막냉각 효율을 보이고, 냉각유체가 좁은 영역에 분포하지만, 본 발명에 따른 막냉각 홀(100)의 경우 중앙부분에서도 높은 냉각효율을 보일 뿐만 아니라, 냉각유체가 표면 전체에 걸쳐 고르게 분포함을 확인할 수 있다.8 (a) and 8 (b) and 9 (a) and 9 (b) are velocity vectors at the exit shape of the conventional fan shape hole 70 shown in FIG. 2 and the hole 100 of the present invention, and It shows a streamline distribution of the cooling fluid, the conventional fan-shaped hole 70 shows a low film cooling efficiency in the central portion of the cooling surface, the cooling fluid is distributed in a narrow region, the film cooling hole 100 according to the present invention In the case of not only shows high cooling efficiency in the central part, but also shows that the cooling fluid is evenly distributed over the entire surface.
따라서, 본 발명에 따른 가스터빈 블레이드의 냉각을 위한 막냉각 홀(100)의 형상 구조에 의하면, 확관부(120) 및 확관부(120)에 돌출 형성되는 돌출부(130)에 의해 기존의 홀의 구조에 비해 월등히 향상된 가스터빈 블레이드(80)의 냉각 효율을 얻을 수 있을 뿐만 아니라, 그에 따라 향상된 냉각성능으로 인해 가스터빈 블레이드의 수명을 연장시킬 수 있고, 적은 냉각유체의 사용 및 전체 막냉각 홀(100)의 설치 개수를 줄임으로써 엔진의 효율을 상승시킬 수 있음과 동시에 전체적인 가스터빈의 구조적 안정성을 향상시킬 수 있는 등의 다양한 장점을 갖는 것이다.Therefore, according to the shape structure of the film cooling hole 100 for cooling the gas turbine blade according to the present invention, the structure of the existing hole by the expansion portion 120 and the protrusions 130 protruding from the expansion portion 120 Not only can the cooling efficiency of the gas turbine blade 80 be significantly improved compared to that of the gas turbine blade 80, and thus, the cooling performance of the gas turbine blade can be extended due to the improved cooling performance, the use of less cooling fluid, and the entire membrane cooling hole (100). By reducing the number of installation) can increase the efficiency of the engine and at the same time have various advantages, such as to improve the structural stability of the overall gas turbine.
전술한 실시예들은 본 발명의 가장 바람직한 예에 대하여 설명한 것이지만, 상기 실시예에만 한정되는 것은 아니며, 본 발명의 기술적 사상을 벗어나지 않는 범위 내에서 다양한 변형이 가능하다는 것은 당업자에게 있어서 명백한 것이다.Although the above embodiments have been described with respect to the most preferred examples of the present invention, it is not limited to the above embodiments, and it will be apparent to those skilled in the art that various modifications are possible without departing from the technical spirit of the present invention.
본 발명은 가스터빈 블레이드의 냉각을 위한 막냉각 홀의 형상 구조에 관한 것으로, 보다 상세하게는 가스터빈 블레이드의 냉각을 위한 목적으로 사용되는 막냉각 홀의 형상을 특정하여 냉각 유체가 냉각 표면에 보다 넓고 효과적으로 분포할 수 있도록 함으로써 기존의 홀의 구조에 비해 월등히 향상된 냉각 효율을 얻을 수 있도록 하는 가스터빈 블레이드의 냉각을 위한 막냉각 홀의 형상 구조에 관한 것이다.The present invention relates to the shape structure of the membrane cooling hole for cooling the gas turbine blade, and more particularly, to specify the shape of the membrane cooling hole used for the cooling of the gas turbine blade so that the cooling fluid is wider and more effective on the cooling surface. The present invention relates to a shape structure of a membrane cooling hole for cooling of a gas turbine blade, which enables to distribute the gas turbine blade, thereby achieving a significantly improved cooling efficiency compared to the structure of a conventional hole.

Claims (4)

  1. 가스터빈 블레이드의 냉각을 위한 막냉각 홀의 형상 구조에 있어서,In the shape structure of the film cooling hole for cooling the gas turbine blade,
    일정한 단면적을 갖는 원통부와,A cylindrical section having a constant cross-sectional area,
    상기 원통부의 상단에 연결되어 그 단면적이 점점 넓어지는 확관부 및An expansion part connected to the upper end of the cylindrical part, the cross section being gradually wider;
    홀 출구의 중앙부분이 유동방향으로 돌출되도록 하여 확관부에 돌출 형성되는 돌출부를 포함하여 구성된 것을 특징으로 하는 가스터빈 블레이드의 냉각을 위한 막냉각 홀의 형상 구조.Shaped structure of the film cooling hole for cooling the gas turbine blades, characterized in that the center portion of the hole exit protruding in the flow direction to comprise a projection protruding to the expansion pipe.
  2. 제 1항에 있어서,The method of claim 1,
    상기 막냉각 홀은 가스터빈 블레이드에 30°의 기울기를 갖도록 형성되고, 원통부의 직경을 D라고 했을 때, 원통부와 확관부의 길이는 각각 3D인 것을 특징으로 하는 가스터빈 블레이드의 냉각을 위한 막냉각 홀의 형상 구조.The film cooling hole is formed to have a slope of 30 ° to the gas turbine blade, when the diameter of the cylindrical portion is D, the length of the cylindrical portion and the expansion portion is 3D film for cooling the gas turbine blade, characterized in that each Shape structure of the cooling hole.
  3. 제 1항에 있어서,The method of claim 1,
    상기 확관부는 원통부를 기준으로 하여 양 측면방향으로 25°의 기울기를 갖도록 확관되는 것을 특징으로 하는 가스터빈 블레이드의 냉각을 위한 막냉각 홀의 형상 구조.The expansion pipe is a shape structure of the film cooling hole for cooling the gas turbine blades, characterized in that the pipe is expanded so as to have a slope of 25 ° in both side directions relative to the cylinder.
  4. 제 1항에 있어서,The method of claim 1,
    상기 돌출부는 원통부를 기준으로 하여 홀 출구 방향으로 저면부분이 2.5°만큼 확관되는 방향으로 돌출 형성된 것을 특징으로 하는 가스터빈 블레이드의 냉각을 위한 막냉각 홀의 형상 구조.The protrusion is formed in the shape of the film cooling hole for cooling the gas turbine blades, characterized in that the bottom portion protrudes in a direction extending by 2.5 ° in the hole exit direction based on the cylindrical portion.
PCT/KR2012/008093 2012-10-05 2012-10-05 Film shape structure of cooling hole for cooling gas turbine blades WO2014054825A1 (en)

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CN114151140A (en) * 2021-11-25 2022-03-08 哈尔滨工程大学 Air film cooling structure applied to turbine stationary blade

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JPH1089005A (en) * 1996-09-18 1998-04-07 Toshiba Corp High temperature member cooling device
JP2006009785A (en) * 2004-06-23 2006-01-12 General Electric Co <Ge> Chevron film cooling type wall
JP2006307842A (en) * 2005-03-30 2006-11-09 Mitsubishi Heavy Ind Ltd High temperature member for gas turbine
JP2008008288A (en) * 2006-06-29 2008-01-17 United Technol Corp <Utc> Gas turbine engine, part and method of optimizing its cooling port shape
JP2008248733A (en) * 2007-03-29 2008-10-16 Mitsubishi Heavy Ind Ltd High temperature member for gas turbine
JP2011064207A (en) * 2005-03-30 2011-03-31 Mitsubishi Heavy Ind Ltd High temperature member for gas turbine
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JPH1089005A (en) * 1996-09-18 1998-04-07 Toshiba Corp High temperature member cooling device
JP2006009785A (en) * 2004-06-23 2006-01-12 General Electric Co <Ge> Chevron film cooling type wall
JP2006307842A (en) * 2005-03-30 2006-11-09 Mitsubishi Heavy Ind Ltd High temperature member for gas turbine
JP2011064207A (en) * 2005-03-30 2011-03-31 Mitsubishi Heavy Ind Ltd High temperature member for gas turbine
JP2008008288A (en) * 2006-06-29 2008-01-17 United Technol Corp <Utc> Gas turbine engine, part and method of optimizing its cooling port shape
JP2008248733A (en) * 2007-03-29 2008-10-16 Mitsubishi Heavy Ind Ltd High temperature member for gas turbine
KR20120115006A (en) * 2011-04-08 2012-10-17 인하대학교 산학협력단 The shape of the film cooling hole for cooling gas turbine blades

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* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN114151140A (en) * 2021-11-25 2022-03-08 哈尔滨工程大学 Air film cooling structure applied to turbine stationary blade

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