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EP2230383A1 - Aube de turbine avec refroidissement de l'extrémité - Google Patents

Aube de turbine avec refroidissement de l'extrémité Download PDF

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Publication number
EP2230383A1
EP2230383A1 EP09155437A EP09155437A EP2230383A1 EP 2230383 A1 EP2230383 A1 EP 2230383A1 EP 09155437 A EP09155437 A EP 09155437A EP 09155437 A EP09155437 A EP 09155437A EP 2230383 A1 EP2230383 A1 EP 2230383A1
Authority
EP
European Patent Office
Prior art keywords
blade
cooling
cooling holes
holes
trailing edge
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.)
Withdrawn
Application number
EP09155437A
Other languages
German (de)
English (en)
Inventor
Shailendra Naik
Gaurav Pathak
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.)
General Electric Technology GmbH
Original Assignee
Alstom Technology AG
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 Alstom Technology AG filed Critical Alstom Technology AG
Priority to EP09155437A priority Critical patent/EP2230383A1/fr
Priority to RU2011141997/06A priority patent/RU2531712C2/ru
Priority to PCT/EP2010/053286 priority patent/WO2010108809A1/fr
Publication of EP2230383A1 publication Critical patent/EP2230383A1/fr
Priority to US13/234,592 priority patent/US20120070308A1/en
Withdrawn legal-status Critical Current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
    • F01D5/20—Specially-shaped blade tips to seal space between tips and stator
    • 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
    • F05D2240/307—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 tip 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
    • F05D2250/00—Geometry
    • F05D2250/20—Three-dimensional
    • F05D2250/23—Three-dimensional prismatic
    • F05D2250/231—Three-dimensional prismatic cylindrical
    • 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
    • F05D2250/20—Three-dimensional
    • F05D2250/23—Three-dimensional prismatic
    • F05D2250/232—Three-dimensional prismatic conical
    • 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
    • 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 field of gas turbine technology, and relates in particular to a cooled blade for a gas turbine as claimed in the preamble of claim 1.
  • the efficiency of gas turbines depends substantially on the temperature of the hot gas that expands in the turbine when performing work.
  • the components must not only be produced from particularly resistant materials but must also be cooled as effectively as possible during operation.
  • Various methods have been developed for blade cooling in the prior art, and can be used alternatively or together.
  • One method is to pass a cooling medium, generally compressed cooling air, from the gas-turbine compressor, through the interior of the blades in cooling channels, and to allow it to emerge into the hot gas channel through cooling holes arranged in a distributed manner.
  • the cooling channels may in this case pass through the interior of the blade more than once in a serpentine shape (see for example WO-A1-2005/068783 ).
  • the heat transfer between the cooling medium and the walls of the blade can in this case be improved by using suitable elements (turbulators) to produce additional turbulence in the cooling medium flow, or by using impingement cooling.
  • the cooling medium can emerge from the interior of the blade such that a film of cooling medium is formed on the blade surface, and protects the blade (film cooling).
  • the blade tip is furthest away from the blade root, through which the cooling air is supplied. Particular attention must therefore be paid to its cooling. Furthermore, cooling that is as uniform as possible must be achieved in all operating states, and the consumption of cooling medium should be restricted to what is necessary, in order to avoid disadvantageously influencing the efficiency of the machine.
  • DE-A1-199 44 923 discloses a comparatively complex solution for cooling the blade tip.
  • the object of the invention is therefore to provide a cooled blade for a gas turbine which is distinguished in particular by better cooling in the area of the blade tip.
  • the object is achieved by the totality of the features of independent claim 1.
  • the major aspect of the invention is that first cooling holes for convection cooling are provided on the pressure face of the blade, and second cooling holes for film cooling are provided on the suction face of the blade, through the cap of the blade, in the blade tip from the cooling channels, and distributed over the blade width.
  • the combination of convection cooling on the pressure face and film cooling on the suction face of the blade tip results in particularly effective and stable cooling without this having any disadvantageous influence on the efficiency.
  • the first and second cooling holes comprise at least sections in the form of cylindrical bores with a predetermined first diameter.
  • the first cooling holes are in the form of long cylindrical bores which run obliquely upwards and include a first angle of between 25° and 35°, preferably of approximately 30°, with the outer surface of the blade.
  • the first cooling holes open into the environment of the blade with a fan-shaped section of the bore.
  • those of the first cooling holes arranged outside the trailing edge of the blade open into the environment of the blade with a 3D symmetric fan-shaped section of the bore, whereby said 3D symmetric fan-shaped section has a first aperture angle having a range of 10° to 50°, and being preferably about 24°, and a second aperture angle perpendicular to said first aperture angle, said second aperture angle having a range of 5° to 25°, and being preferably about 12°.
  • those of the first cooling holes arranged outside the trailing edge of the blade include a second angle of between 15° and 45°, preferably of approximately 30°, with the outer surface of the blade.
  • those of the first cooling holes arranged at the trailing edge of the blade open into the environment of the blade with a 2D symmetric fan-shaped section of the bore, whereby said 2D symmetric fan-shaped section has a third aperture angle having a range of 10° to 40°, and being preferably about 20°.
  • those of the first cooling holes arranged at the trailing edge of the blade include a third angle of between 5° and 45°, preferably of approximately 30°, with the outer surface of the blade.
  • those of the first cooling holes arranged at the trailing edge of the blade have a bore of a predetermined first length, which is subdivided into said 2D symmetric fan-shaped section and a cylindrical section of a predetermined second length, whereby the ratio of said second length and said first length is in the range of 0.2 to 0.7, and is preferably about 0.5.
  • the first cooling holes are arranged along the pressure face in a row with a predetermined first periodicity, and the ratio between said first periodicity and said first diameter is in the range of 3 to 8, and is preferably about 6.
  • the second cooling holes pass through the cap of the blade in a radial direction, whereby the second cooling holes are in the form of long cylindrical bores which run obliquely upwards and include an angle of 0° to 45°, preferably of approximately 30°, with the longitudinal axis of the blade.
  • the second cooling holes are arranged along the suction face in a row with a predetermined second periodicity, and the ratio between said second periodicity and said first diameter is in the range of 3 to 8, and is preferably about 6.
  • said first cooling holes exit into the environment of the blade at a predetermined height below the upper end of the blade tip, and the ratio between said height and said first diameter is in a range between 5 and 10, and is preferably about 6.5.
  • dust holes arranged along the cap between said leading edge and trailing edge, and said dust holes have a second diameter, such that the ratio between said second diameter and said first diameter is between 1.2 and 4.5.
  • the cap of the blade is bounded at the edge on its upper face by a circumferential blade crown, and the second cooling holes open into the outside area within the blade crown.
  • said blade has a blade crown at the blade tip, which is bounded by a circumferential rail having a predetermined thickness, whereby the width between the opposing rails varies with the distance along the chord line, such that t/W is between 0.05 and 0.15 for k/ k 0 between 0 and 0.3, and that t/W is between 0.15 and 0.3 for k/ k 0 larger than 0.3 and up to 1.0, k 0 being the overall chord line length.
  • D/W 0.1 to 0.3 for k/ k 0 0 to 0.3
  • D/W 0.1 to 0.8 for k/ k 0 >0 to 1.0
  • D means the depth of the tip crown
  • W means the width, according Fig. 3a .
  • the invention relates to a cooled gas turbine blade which is particularly suitable for implementation of the invention.
  • the blade (10 in Figures 1, 2 ), which is a rotor blade, has an airfoil section (12 in Figure 2 ), which extends in the radial direction of the turbine and extends in the radial direction between a platform (not shown), which bounds the hot gas channel, and a blade tip (11 in Figure 2 ).
  • a platform not shown
  • blade tip 11 in Figure 2
  • the airfoil section 12 has a leading edge 15 and a trailing edge 16 ( Figure 1 ), and has a (concave) pressure face 17 and a (convex) suction face 18 in the form of an airfoil profile.
  • a blade root (not shown) is formed underneath the platform, and is used to mount the blade 10 in a groove provided for this purpose in the rotor (or, in the case of a stator blade, in the housing surrounding the rotor).
  • Cooling channels 19a, 19b, 19c and 20 ( Figure 1 ) through which cooling air flows run in the radial direction in the interior of the airfoil section 12, and this cooling air enters the blade 10 as a cooling air flow through appropriate cooling air inlets (not shown) in the blade root.
  • the cooling channels 19a, 19b and 19c are connected to one another by means of a serpentine-like channel structure.
  • the cooling air flowing through the cooling channels 19a, 19b and 19c cools the blade 10 from the inside and emerges to the outside at different points through cooling holes or cooling openings.
  • the cooling channel 20 is specifically used to cool the leading edge 15.
  • turbulators in the form of obliquely positioned ribs can be provided in the cooling channels 19a, b, c and 20 and lead to swirling of the cooling air, and therefore to an improvement in the heat transfer.
  • first, comparatively long cooling holes 25 for convection cooling are provided, distributed over the blade width, from the cooling channels 19 and 19a, b, c in the blade tip 11, passing to the outside on the pressure face 17 of the blade 10.
  • Second cooling holes 27 are passed to the outside through the cap 33 of the blade 10, for film cooling on the suction face 18 of the blade 10.
  • a particularly advantageous cooling effect is achieved by the combination of convection cooling on the pressure face 17 and film cooling on the suction face 18 of the blade.
  • the first and second cooling holes 25 and 27, respectively, may have the form of cylindrical bores in the simplest embodiment ( Figure 3a ) and can be introduced into the blade 10 by appropriate drilling methods (EDM, laser drilling).
  • the first cooling holes 25 are advantageously in the form of holes or bores which run obliquely upwards, in order to achieve the necessary hole length. They preferably include a first angle ⁇ 1 of between 25° and 35°, preferably of approximately 30°, with the outer surface 17 of the blade 10.
  • the first and second cooling holes (25a,b in Figure 2 and Figure 3b ) comprise only sections in the form of cylindrical bores with a predetermined first diameter d. They therefore open advantageously into the environment of the blade 10 with a fan-shaped section (29, 30 in Figures 4 a-c , 5a+b ) of the bore.
  • first cooling holes 25a there are two different kinds 25a (see Figure 4a ) and 25b (see Figure 5a ) of first cooling holes provided at the pressure side (17) of the blade 10: Those of the first cooling holes arranged outside the trailing edge 16 of the blade 10, i.e. first cooling holes 25a, preferably open into the environment of the blade 10 with a 3D (3-dimensional) symmetric fan-shaped section 29 of the bore, which is shown in Figures 4a, 4b and 4c .
  • Said 3D symmetric fan-shaped section 29 has a first aperture angle 2 ⁇ 1 ( Figure 4b ) having a range of 10° to 50°, and being preferably about 24°, and a second aperture angle ⁇ 2 ( Figure 4a ) perpendicular to said first aperture angle 2 ⁇ 1 .
  • Said second aperture angle ⁇ 2 has a range of 5° to 25°, and is preferably about 12°. Furthermore, these first cooling holes 25a arranged outside the trailing edge 16 of the blade 10 include a second angle ⁇ 2 of between 15° and 45°, preferably of approximately 30°, with the outer surface 17 of the blade 10 ( Figure 4a ).
  • first cooling holes 25b arranged at the trailing edge 16 of the blade 10 preferably open into the environment of the blade 10 with a 2D (2-dimensional) symmetric fan-shaped section 30 of the bore ( Figures 5a, 5b and 5c ).
  • Said 2D symmetric fan-shaped section 30 has a third aperture angle 2 ⁇ 3 ( Figure 5a ), which has a range of 10° to 40°, and is preferably about 20°.
  • These first cooling holes 25b arranged at the trailing edge 16 of the blade 10 include a third angle ⁇ 3 ( Figure 5c ) of between 5° and 45°, preferably of approximately 30°, with the outer surface 17 of the blade 10.
  • those of the first cooling holes 25b arranged at the trailing edge 16 of the blade 10 have a bore of a predetermined overall length L.
  • This overall length L is subdivided into the aforementioned 2D symmetric fan-shaped section 30 and a cylindrical section of a second length L 1 .
  • the ratio L 1 /L of both lengths lies in the range of 0.2 to 0.7, and is preferably about 0.5.
  • Figure 2 shows, that the first cooling holes 25a and 25b are arranged along the pressure face 17 in a row with a (first) periodicity P 1 . It is advantageous to choose a certain ratio P 1 /d between this periodicity P 1 and the diameter d (see Figure 3a ) of the cooling hole bores. This ratio is chosen to be in the range of 3 to 8, and is preferably about 6.
  • the second cooling holes 27 are arranged along the suction face 18 in a row with a (second) periodicity P 2 .
  • the ratio P 2 /d 1 between the second periodicity P 2 and the diameter d lies in the range of 5 to 8, and is preferably about 6.
  • the blade 10 is closed at the blade tip 11 at the top by a flat cap 33, which is surrounded on the upper face by a circumferential rail-like blade crown 32.
  • the second cooling holes 27 pass through the cap 33 of the blade 10 in a radial direction. They are in the form of long cylindrical bores which run obliquely upwards and include an angle Y of 0° to 45°, preferably of approximately 30°, with the longitudinal axis of the blade 10 ( Figure 3c ).
  • the first cooling holes 25 open into the outside area underneath the cap 33 of the blade 10. They exit into the environment of the blade 10 at a predetermined height H below the upper end of the blade tip 11 ( Figure 3a ).
  • the ratio H/d between said height H and the diameter d is in a range between 5 and 10, and is preferably about 6.5.
  • the second cooling holes 27 are arranged on the opposite face and pass through the cap 33 of the blade 10 in the radial direction, opening into the outside area within the blade crown 32.
  • dust holes 26 are provided and arranged along the cap 33 between the leading edge 15 and trailing edge 16 ( Figure 2 ). These dust holes 26, which are used to remove dust particles from the interior cooling channels, each have a diameter d 1 , such that the ratio d 1 /d between the diameter d 1 and the bore diameter d (see Figure 3a ) is between 1.2 and 4.5.
  • the blade 10 is provided with a blade crown 32 at the blade tip 11, which blade crown 32 is bounded by a circumferential rail having a predetermined thickness t ( Figure 3a ).
  • the width W between the opposing rails varies with the distance k along the chord line ( Figure 2 ), such that t/W is between 0.05 and 0.15 for k/ k 0 between 0 and 0.3, and that t/W is between 0.15 and 0.3 for k/ k 0 larger than 0.3 and up to 1.0, k 0 being the overall chord line length;
  • the surfaces of the pressure face 17 and suction face 18 as well as the upper face of the cap 33 are provided with a thermal protection layer (Thermal Barrier Coating TBC) 28.
  • Thermal Protection layer Thermal Barrier Coating TBC

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP09155437A 2009-03-18 2009-03-18 Aube de turbine avec refroidissement de l'extrémité Withdrawn EP2230383A1 (fr)

Priority Applications (4)

Application Number Priority Date Filing Date Title
EP09155437A EP2230383A1 (fr) 2009-03-18 2009-03-18 Aube de turbine avec refroidissement de l'extrémité
RU2011141997/06A RU2531712C2 (ru) 2009-03-18 2010-03-15 Лопатка для газовой турбины с охлаждаемой законцовкой периферической части лопатки
PCT/EP2010/053286 WO2010108809A1 (fr) 2009-03-18 2010-03-15 Pale pour turbine à gaz avec capuchon d'extrémité refroidi
US13/234,592 US20120070308A1 (en) 2009-03-18 2011-09-16 Cooled blade for a gas turbine

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP09155437A EP2230383A1 (fr) 2009-03-18 2009-03-18 Aube de turbine avec refroidissement de l'extrémité

Publications (1)

Publication Number Publication Date
EP2230383A1 true EP2230383A1 (fr) 2010-09-22

Family

ID=41343214

Family Applications (1)

Application Number Title Priority Date Filing Date
EP09155437A Withdrawn EP2230383A1 (fr) 2009-03-18 2009-03-18 Aube de turbine avec refroidissement de l'extrémité

Country Status (4)

Country Link
US (1) US20120070308A1 (fr)
EP (1) EP2230383A1 (fr)
RU (1) RU2531712C2 (fr)
WO (1) WO2010108809A1 (fr)

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EP3034793A1 (fr) * 2014-12-15 2016-06-22 United Technologies Corporation Composant de moteur à turbine à gaz avec capacité de refroidissement accrue
EP3225782A1 (fr) * 2016-03-29 2017-10-04 Ansaldo Energia Switzerland AG Profil d'aube et élément aubagé associé
US10539026B2 (en) 2017-09-21 2020-01-21 United Technologies Corporation Gas turbine engine component with cooling holes having variable roughness
EP2547871B1 (fr) * 2010-03-19 2020-04-29 Ansaldo Energia IP UK Limited Profil de turbine à gaz avec trous d'éjection de réfrigérant sur le bord de fuite

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US10036319B2 (en) 2014-10-31 2018-07-31 General Electric Company Separator assembly for a gas turbine engine
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US10428664B2 (en) 2015-10-15 2019-10-01 General Electric Company Nozzle for a gas turbine engine
US9988936B2 (en) 2015-10-15 2018-06-05 General Electric Company Shroud assembly for a gas turbine engine
US10196904B2 (en) * 2016-01-24 2019-02-05 Rolls-Royce North American Technologies Inc. Turbine endwall and tip cooling for dual wall airfoils
US10704425B2 (en) 2016-07-14 2020-07-07 General Electric Company Assembly for a gas turbine engine
FR3062675B1 (fr) * 2017-02-07 2021-01-15 Safran Helicopter Engines Aube haute pression ventilee de turbine d'helicoptere comprenant un conduit amont et une cavite centrale de refroidissement
US10400610B2 (en) * 2017-02-14 2019-09-03 General Electric Company Turbine blade having a tip shroud notch
EP3669054B1 (fr) * 2017-08-14 2022-02-09 Siemens Energy Global GmbH & Co. KG Aube de turbine et procédé de maintenance correspondant
US10641106B2 (en) 2017-11-13 2020-05-05 Honeywell International Inc. Gas turbine engines with improved airfoil dust removal
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CN112682106B (zh) * 2020-12-20 2022-11-11 中国航发四川燃气涡轮研究院 带有异形微群气膜冷却孔的涡轮叶片端壁结构及方法和燃气涡轮
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EP0319758A1 (fr) * 1987-12-08 1989-06-14 General Electric Company Refroidissement de l'extrémité d'une aube par diffusion
EP0684364A1 (fr) * 1994-04-21 1995-11-29 Mitsubishi Jukogyo Kabushiki Kaisha Refroidissement des extremités des aubes de turbine
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EP1059419A1 (fr) * 1999-06-09 2000-12-13 General Electric Company Aube avec trois nervures sur l'extrémité de l'aube
DE19944923A1 (de) 1999-09-20 2001-03-22 Asea Brown Boveri Turbinenschaufel für den Rotor einer Gasturbine
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US20050042074A1 (en) * 2002-09-05 2005-02-24 Siemens Westinghouse Power Corporation Combustion turbine with airfoil having multi-section diffusion cooling holes and methods of making same
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WO2005068783A1 (fr) 2004-01-16 2005-07-28 Alstom Technology Ltd Aube refroidie pour une turbine a gaz
EP1621731A1 (fr) * 2004-07-26 2006-02-01 General Electric Company Aube ayant une chambre commune en extrémité
US20080118367A1 (en) * 2006-11-21 2008-05-22 Siemens Power Generation, Inc. Cooling of turbine blade suction tip rail

Cited By (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2547871B1 (fr) * 2010-03-19 2020-04-29 Ansaldo Energia IP UK Limited Profil de turbine à gaz avec trous d'éjection de réfrigérant sur le bord de fuite
EP3009600A1 (fr) * 2014-10-14 2016-04-20 United Technologies Corporation Aube de turbine à bout refroidi de moteur à turbine à gaz
EP3034793A1 (fr) * 2014-12-15 2016-06-22 United Technologies Corporation Composant de moteur à turbine à gaz avec capacité de refroidissement accrue
US10247011B2 (en) 2014-12-15 2019-04-02 United Technologies Corporation Gas turbine engine component with increased cooling capacity
EP3225782A1 (fr) * 2016-03-29 2017-10-04 Ansaldo Energia Switzerland AG Profil d'aube et élément aubagé associé
CN107237653A (zh) * 2016-03-29 2017-10-10 安萨尔多能源瑞士股份公司 翼型
US11035234B2 (en) 2016-03-29 2021-06-15 Ansaldo Energia Switzerland AG Airfoil having a tip capacity
CN107237653B (zh) * 2016-03-29 2021-12-14 安萨尔多能源瑞士股份公司 翼型
US10539026B2 (en) 2017-09-21 2020-01-21 United Technologies Corporation Gas turbine engine component with cooling holes having variable roughness

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RU2011141997A (ru) 2013-04-27
US20120070308A1 (en) 2012-03-22
WO2010108809A1 (fr) 2010-09-30
RU2531712C2 (ru) 2014-10-27

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