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US20140003962A1 - Turbine blade - Google Patents

Turbine blade Download PDF

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Publication number
US20140003962A1
US20140003962A1 US14/004,309 US201214004309A US2014003962A1 US 20140003962 A1 US20140003962 A1 US 20140003962A1 US 201214004309 A US201214004309 A US 201214004309A US 2014003962 A1 US2014003962 A1 US 2014003962A1
Authority
US
United States
Prior art keywords
portions
cooling
blade
rows
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.)
Abandoned
Application number
US14/004,309
Other languages
English (en)
Inventor
Yoji Okita
Chiyuki Nakamata
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
IHI Corp
Original Assignee
IHI Corp
Priority date (The priority date 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 date listed.)
Filing date
Publication date
Application filed by IHI Corp filed Critical IHI Corp
Assigned to IHI CORPORATION reassignment IHI CORPORATION ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: NAKAMATA, CHIYUKI, OKITA, YOJI
Publication of US20140003962A1 publication Critical patent/US20140003962A1/en
Abandoned legal-status Critical Current

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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • 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
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/187—Convection cooling
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D9/00—Stators
    • F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/12—Cooling of plants
    • F02C7/16—Cooling of plants characterised by cooling medium
    • F02C7/18—Cooling of plants characterised by cooling medium the medium being gaseous, e.g. air
    • 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
    • F05D2240/00—Components
    • F05D2240/20—Rotors
    • F05D2240/30—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor
    • F05D2240/304—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor related to the trailing edge of a rotor blade
    • 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
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/221—Improvement of heat transfer
    • F05D2260/2212—Improvement of heat transfer by creating turbulence

Definitions

  • the present invention relates to a turbine blade.
  • Priority is claimed on Japanese Patent Application No. 2011-054253, filed Mar. 11, 2011, the contents of which are incorporated herein by reference.
  • Turbine blades that are provided in a turbine are generally exposed to high-temperature fluid bodies.
  • the turbine blades provide in gas turbines are exposed to high-temperature combustion gas that has been discharged from a combustion chamber, they are exposed to an extremely high-temperature environment.
  • a cooling gas such as cooling air is supplied to the interior of the turbine blades.
  • an area around the rear edge of a turbine blade it is desirable for the thickness of an area around the rear edge of a turbine blade to be as thin as possible from the standpoint of the aerodynamic performance of the turbine. For this reason, in many cases it is difficult to form a flow path in the area around the rear edge of a turbine blade to enable the cooling gas to flow into the interior of the turbine blade. Namely, the area around the rear edge of a turbine blade can be said to be an area that is difficult to cool. Because of this, in, for example, Patent document 1 and Patent document 2, a technique is proposed in which the rear edge area is made thinner by forming notch portions in the ventral side of the rear edge area of a turbine blade, and film cooling is performed on the rear edge area by blowing cooling gas onto the ventral side surface that is exposed by these notch portions.
  • the present invention was conceived in view of the above-described problems, and it is an object thereof to improve the cooling efficiency in an area around the rear edge of a turbine blade by suppressing any deterioration in the film cooling efficiency in the turbine blade, and by improving the heat transfer rate between the blade portion and the cooling gas.
  • the present invention employs the following structure as a means of solving the above-described problems.
  • a first aspect of the present invention is a turbine blade that is provided with a plurality of rows of cooling portions that have notch portions that discharge cooling gas that has been introduced into an internal portion of a blade portion onto a ventral side blade surface, and that are formed in rows that are stacked in a direction between the blade front edge and the blade rear edge, and that includes: turbulence promoting cooling portions that are provided in the row located furthest to the downstream side from among the plurality of rows, and that have turbulence promoting devices in areas exposed by the notch portions; and film cooling portions that are provided in at least one of other rows from among the plurality of rows, and that form a film cooling layer in the areas exposed by the notch portions.
  • a second aspect of the present invention is the turbine blade according to the first or second aspects wherein, in at least one of the plurality of rows, the plurality of cooling portions are placed discretely from each other in the height direction of the blade portion.
  • a third aspect of the present invention is the turbine blade according to any of the first through third aspects wherein, in at least one of the plurality of rows, the cooling portions are provided continuously in an elongated shape in the height direction of the blade portion.
  • a fourth aspect of the present invention is the turbine blade according to any of the first through fourth aspects wherein the turbulence promoting device is formed by a plurality of pin portions that stand upright in the areas exposed by the notch portions.
  • a fifth aspect of the present invention is the turbine blade according to any of the first through fourth aspects wherein the turbulence promoting device is formed by a plurality of step portions that stand upright in the areas exposed by the notch portions and that are stacked in rows in the flow direction of the cooling gas.
  • a sixth aspect of the present invention is the turbine blade according to any of the first through fourth aspects wherein the turbulence promoting device is formed by a plurality of dimple portions that are formed in the areas exposed by the notch portions.
  • a plurality of rows of cooling portions having notch portions are provided in rows that are stacked in a direction from the front edge of a blade to the rear edge of the blade.
  • the row located furthest to the downstream side from among this plurality of rows of cooling portions forms a turbulence promoting cooling portion that is provided with turbulence promoting devices and cools by means of turbulence.
  • All of the other rows form film cooling portions that form a film cooling layer.
  • FIG. 1 is a perspective view of a turbine blade according to an embodiment of the present invention.
  • FIG. 2A is a cross-sectional view of a turbine blade according to an embodiment of the present invention.
  • FIG. 2B is a cross-sectional view of a turbine blade according to an embodiment of the present invention.
  • FIG. 3A is a perspective view showing in enlargement principal portions of a turbine blade according to an embodiment of the present invention.
  • FIG. 3B is a perspective view showing in enlargement principal portions of a turbine blade according to an embodiment of the present invention.
  • FIG. 4A is an enlarged perspective view showing a variant example of a turbine blade according to an embodiment of the present invention.
  • FIG. 4B is an enlarged perspective view showing a variant example of a turbine blade according to an embodiment of the present invention.
  • FIG. 5 is a plan view showing a variant example of a turbine blade according to an embodiment of the present invention.
  • FIG. 1 is a perspective view of a turbine blade 1 of the present embodiment.
  • FIGS. 2A and 2B are cross-sectional views of the turbine blade 1
  • FIG. 2A is a cross-sectional view taken along a line A-A in FIG. 1
  • FIG. 2B is a cross-sectional view taken along a line B-B in FIG. 1 .
  • the turbine blade 1 of the present embodiment is provided with a blade portion 2 , a root portion 3 , and cooling portions 4 .
  • the blade portion 2 is set in a blade shape having a front edge 21 (i.e., a blade front edge), a rear edge 22 (i.e., a blade rear edge), a ventral side blade surface 23 (i.e., a blade belly), and a rear side blade surface 24 .
  • an internal portion 25 i.e., a blade interior portion of the blade portion 2 is provided with a hollow portion into which a cooling gas can be introduced. This blade portion 2 is positioned with the front edge 21 facing towards the upstream side of the fluid body.
  • a flow in which the fluid body flows from the front edge 21 towards the rear edge 22 is formed on the surfaces of the ventral side blade surface 23 and the rear side blade surface 24 of the blade portion 2 .
  • a flow from the front edge 21 towards the rear edge 22 along the surfaces of the ventral side blade surface 23 and the rear side blade surface 24 is referred to as a main flow.
  • the root portion 3 supports the blade portion 2 , and is fixed to a turbine disk. Note that the root portion 3 has an internal flow path (not shown). Cooling gas is supplied to the internal portion 25 of the blade portion 2 through this internal flow path in the root portion 3 .
  • the cooling portions 4 discharge the cooling gas introduced into the internal portion 25 of the blade portion 2 onto the ventral side blade surface 23 side of the rear edge 22 side, and cool the rear edge area including the rear edge 22 .
  • These cooling portions 4 have notch portions 41 that are formed on the ventral side blade surface 23 side, and these notch portions 41 form slots 42 .
  • the cooling portions 4 have through holes 43 that are provided in the wall portion on the front edge 21 side of the slot 42 , and that penetrate into the interior portion 25 of the blade portion 2 . These through holes 43 discharge cooling gas into the notch portions 41 .
  • the cooling portions 4 of the turbine blade 1 of the present embodiment have notch portions 41 that discharge the cooling gas introduced into the internal portion 25 of the blade portion 2 onto the ventral side blade surface 23 side.
  • the turbine blade 1 of the present embodiment is provided with two rows (i.e., a plurality of rows) of the cooling portions 4 that are stacked in a direction from the front edge 21 towards the rear edge 22 .
  • Each of these rows is formed by a plurality of the cooling portions 4 that are placed discretely from each other in the height direction of the blade portion 2 (i.e., in the top-bottom direction of the sheet of paper on which FIG. 1 is shown) and at equal distances from each other.
  • the cooling portions 4 that form the row on the front edge 21 side, and the cooling portions 4 that form the row on the rear edge 22 side are offset from each other in the height direction of the blade portion 2 .
  • the cooling portions 4 are arranged in a zigzag pattern.
  • the cooling portions 4 that form the row on the front edge 21 side make up film cooling portions 4 A, while the cooling portions 4 that form the row on the rear edge 22 side make up film cooling portions 413 .
  • FIG. 3A is an enlarged perspective view of a film cooling portion 4 A.
  • the film forming portion 4 A is provided with a flat area R that is exposed by the notch portion 41 .
  • This film cooling portion 4 A forms a film cooling layer using the cooling gas that is discharged from the through hole 43 , and thereby performs film cooling.
  • the film cooling portion 4 A forms a film cooling layer in the area R that is exposed by the notch portion 41 , and is thereby able to perform film cooling. Namely, the area around the rear edge of the blade portion 2 is able to be cooled by this structure.
  • FIG. 3B is an enlarged perspective view of a film turbulence promoting cooling portion 4 B.
  • the turbulence promoting cooling portion 4 B is provided with a plurality of pin portions 44 (i.e., turbulence promoting devices) that stand upright on the flat area R that is exposed by the notch portions 41 .
  • This turbulence promoting cooling portion 4 B promotes turbulence as a result of the laminar flow expelled from the through hole 43 colliding with the plurality of pin portions 44 . Namely, using this turbulence, it is possible to promote a convective heat transfer between the film air and the blade portion.
  • FIG. 3B is an enlarged perspective view of a film turbulence promoting cooling portion 4 B.
  • the turbulence promoting cooling portion 4 B is provided with a plurality of pin portions 44 (i.e., turbulence promoting devices) that stand upright on the flat area R that is exposed by the notch portions 41 .
  • the plurality of pin portions 44 are arranged in a zigzag pattern on the area R.
  • the height of each pin portion 44 is set such that the pin portions 44 do not protrude above the surface of the ventral side blade surface 23 in order to avoid any collision with the aforementioned main flow. Note that, as is shown in FIG. 3B , in the present embodiment all of the pin portions 44 are set at the same height, however, it is not essential for the height of the pin portions 44 to be standardized.
  • the plurality of rows of cooling portions 4 that have the notch portions 41 are arranged so as to form rows that are stacked from the front edge 21 towards the rear edge 22 .
  • the row from among the plurality of rows of cooling portions 4 that is located on the rear edge 22 side forms the turbulence promoting cooling portions 4 B, while the row on the front edge 21 side forms the film cooling portions 4 A. Because of this, according to the turbine blade 1 of the present embodiment, a convective heat transfer is promoted by the turbulence promoting cooling portions 4 B, and a film cooling layer is formed by the film cooling portions 4 A.
  • the convective heat transfer that is generated between the film cooling gas whose temperature is lower than that of the main flow gas and the blade portion is promoted by the turbulence promoting cooling portion, rear edge portions of a blade portion can be cooled efficiently. Accordingly, according to the present invention, by providing the notch portions that form the film cooling layer separately from the notch portions that have a structure that promotes the transfer of heat, it is possible to suppress any reduction in the film cooling efficiency, and to improve the heat transfer rate between the blade portions and the cooling gas. As a result, the cooling efficiency of the area around the rear edge of a blade portion can be improved.
  • the cooling portions 4 that form the row furthest to the rear edge 22 side form the turbulence promoting cooling portions 413 .
  • the rear side blade surface of the rear edge side of the blade portion is where the temperature of the blade surface is most easily increased, and it is necessary to strengthen the cooling in this region. Namely, by providing notch portions in the wall surface of the side of the blade portion having this region that is cooled (i.e., the blade belly side), it is possible to achieve a heat transfer in this region at a high heat transfer rate with cooling gas at the lowest possible temperature.
  • the turbine blade 1 of the present embodiment as a result of the cooling portions 4 that form the row furthest to the rear edge 22 side forming the turbulence promoting cooling portions 413 , the effects of this turbulence on the downstream side of the flow can be reduced. Namely, it is possible to prevent any deterioration in the film cooling efficiency and to improve the heat transfer rate between the blade portion and the cooling gas.
  • a structure is employed in which turbulence is created by the pin portions 44 . Because of this, turbulence can be created by means of a simple structure.
  • step portions 45 i.e., turbulence promoting devices
  • the cooling gas discharged from the through holes 43 is able to collide with the step portions 45 so that turbulence is promoted by this structure as well. Namely, turbulence can be promoted by means of a simple structure.
  • the plurality of cooling portions 4 that make up a row are placed discretely from each other in the height direction of the blade portion. As a consequence, it is possible to leave a thick area of the blade portion between cooling portions 4 so that the strength of the turbine blade 1 remains unimpaired.
  • cooling portion 4 i.e., the cooling portion 4 A in FIG. 5
  • the cooling portion 4 A it is also possible to provide a cooling portion 4 (i.e., the cooling portion 4 A in FIG. 5 ) that is provided continuously in an elongated shape in the height direction of the blade portion.
  • this type of cooling portion 4 it is possible to cool an even broader range in the height direction of the blade portion.
  • the cooling portion 4 is formed in an elongated shape, then the surface area of the apertures of the through holes also increases.
  • the flow rate of cooling gas that is discharged from a row that is formed by this cooling portion 4 that is provided continuously in an elongated shape will be greater than the flow rate of cooling gas discharged from a row that is formed by cooling portions 4 that are arranged discretely. Because of this, if the cooling portion 4 is formed in an elongated shape, it is preferable for the surface area of the apertures of the through holes in the cooling portion 4 that is provided continuously in an elongated shape to be narrowed down, so that sufficient cooling gas can be discharged from the cooling portions 4 that are arranged discretely.
  • the typical view of a turbine blade in FIG. 1 shows a moving blade.
  • the present structure can also be applied to a stationary blade.
  • two rows of cooling portions 4 are provided so as to form rows that are stacked in the flow direction of the main flow.
  • the present invention is not limited to this and it is also possible to provide the plurality of rows of cooling portions 4 in three or more rows in the flow direction of the main flow.
  • those cooling portions in the row located furthest to the downstream side form the turbulence promoting cooling portions 4 B, while the cooling portions in at least one of the other rows form the film cooling portions 4 A.
  • the cooling portions 4 are arranged in a zigzag pattern.
  • the present invention is not limited to this and it is also possible for the cooling portions 4 to be arranged in a lattice configuration.
  • the through holes 43 are not formed in the turbulence promoting cooling portions 4 B.
  • notch portions that form a film separately from notch portions that have a structure that promotes the transfer of heat, it is possible to suppress any reduction in the film cooling efficiency, and to improve the heat transfer rate between the blade portions and the cooling gas. As a result, the cooling efficiency of the area around the rear edge of a turbine blade can be improved.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
US14/004,309 2011-03-11 2012-03-07 Turbine blade Abandoned US20140003962A1 (en)

Applications Claiming Priority (3)

Application Number Priority Date Filing Date Title
JP2011054253A JP2012189026A (ja) 2011-03-11 2011-03-11 タービン翼
JP2011-054253 2011-03-11
PCT/JP2012/055876 WO2012124578A1 (fr) 2011-03-11 2012-03-07 Aube de turbine

Publications (1)

Publication Number Publication Date
US20140003962A1 true US20140003962A1 (en) 2014-01-02

Family

ID=46830651

Family Applications (1)

Application Number Title Priority Date Filing Date
US14/004,309 Abandoned US20140003962A1 (en) 2011-03-11 2012-03-07 Turbine blade

Country Status (7)

Country Link
US (1) US20140003962A1 (fr)
EP (1) EP2685049B1 (fr)
JP (1) JP2012189026A (fr)
KR (1) KR20130116323A (fr)
CN (1) CN103403299B (fr)
CA (1) CA2829742C (fr)
WO (1) WO2012124578A1 (fr)

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US20160003071A1 (en) * 2014-05-22 2016-01-07 United Technologies Corporation Gas turbine engine stator vane baffle arrangement
WO2021019156A1 (fr) * 2019-07-30 2021-02-04 Safran Aircraft Engines Aube mobile de turbomachine a circuit de refroidissement ayant une double rangee de fentes d'evacuation
US11773727B2 (en) 2020-03-18 2023-10-03 Safran Aircraft Engines Turbine blade comprising three types of orifices for cooling the trailing edge
US11867083B2 (en) 2020-06-22 2024-01-09 Siemens Energy Global GmbH & Co. KG Turbine blade and method for machining same

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CN103912316A (zh) * 2014-04-11 2014-07-09 北京航空航天大学 一种涡轮导叶缝气膜冷却结构
WO2017082907A1 (fr) * 2015-11-12 2017-05-18 Siemens Aktiengesellschaft Surface portante de turbine avec bord de fuite refroidi
KR20180082118A (ko) * 2017-01-10 2018-07-18 두산중공업 주식회사 가스 터빈의 블레이드 또는 베인의 컷백
CN107013255A (zh) * 2017-06-01 2017-08-04 西北工业大学 一种带有连续直肋的涡轮叶片尾缘扰流半劈缝冷却结构
CN107060893A (zh) * 2017-06-01 2017-08-18 西北工业大学 一种带有v型肋的涡轮叶片尾缘扰流半劈缝冷却结构
CN107013254A (zh) * 2017-06-01 2017-08-04 西北工业大学 一种带有球面凸块的涡轮叶片尾缘扰流半劈缝冷却结构
CN107035421A (zh) * 2017-06-01 2017-08-11 西北工业大学 一种带有阵列针肋的涡轮叶片尾缘扰流半劈缝冷却结构
US10844728B2 (en) * 2019-04-17 2020-11-24 General Electric Company Turbine engine airfoil with a trailing edge
FR3102794B1 (fr) * 2019-10-31 2022-09-09 Safran Aircraft Engines Composant de turbomachine comportant des orifices de refroidissement ameliores
FR3107562B1 (fr) * 2020-02-20 2022-06-10 Safran Aube de turbomachine comportant des fentes de refroidissement de son bord de fuite équipées de perturbateurs
KR102456633B1 (ko) 2020-10-23 2022-10-18 두산에너빌리티 주식회사 터빈 블레이드의 트레일링 에지 냉각 구조

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Publication number Priority date Publication date Assignee Title
US20160003071A1 (en) * 2014-05-22 2016-01-07 United Technologies Corporation Gas turbine engine stator vane baffle arrangement
WO2021019156A1 (fr) * 2019-07-30 2021-02-04 Safran Aircraft Engines Aube mobile de turbomachine a circuit de refroidissement ayant une double rangee de fentes d'evacuation
FR3099522A1 (fr) * 2019-07-30 2021-02-05 Safran Aircraft Engines Aube mobile de turbomachine à circuit de refroidissement ayant une double rangée de fentes d’évacuation
CN114207249A (zh) * 2019-07-30 2022-03-18 赛峰航空器发动机 具有带有双排排放槽的冷却回路的涡轮发动机移动叶片
US11913354B2 (en) 2019-07-30 2024-02-27 Safran Aircraft Engines Turbomachine moving blade with cooling circuit having a double row of discharge slots
US11773727B2 (en) 2020-03-18 2023-10-03 Safran Aircraft Engines Turbine blade comprising three types of orifices for cooling the trailing edge
US11867083B2 (en) 2020-06-22 2024-01-09 Siemens Energy Global GmbH & Co. KG Turbine blade and method for machining same

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KR20130116323A (ko) 2013-10-23
CA2829742A1 (fr) 2012-09-20
CN103403299A (zh) 2013-11-20
EP2685049A1 (fr) 2014-01-15
EP2685049A4 (fr) 2014-10-01
JP2012189026A (ja) 2012-10-04
CA2829742C (fr) 2017-01-17
EP2685049B1 (fr) 2018-05-23
WO2012124578A1 (fr) 2012-09-20
CN103403299B (zh) 2016-06-29

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