EP1849960A2 - Turbinenschaufel mit innerem Kühlkanal - Google Patents

Turbinenschaufel mit innerem Kühlkanal Download PDF

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Publication number
EP1849960A2
EP1849960A2 EP07008241A EP07008241A EP1849960A2 EP 1849960 A2 EP1849960 A2 EP 1849960A2 EP 07008241 A EP07008241 A EP 07008241A EP 07008241 A EP07008241 A EP 07008241A EP 1849960 A2 EP1849960 A2 EP 1849960A2
Authority
EP
European Patent Office
Prior art keywords
blade
trailing edge
cooling
wall surface
turbine
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Withdrawn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
EP07008241A
Other languages
English (en)
French (fr)
Other versions
EP1849960A3 (de
Inventor
Ryou c/o Hitachi Ltd Intellectual Property Group Akiyama
Yasuhiro c/o Hitachi Ltd Intellectual Property Group Horiuchi
Shinya c/o Hitachi Ltd Intellectual Property Group Marushima
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.)
Hitachi Ltd
Original Assignee
Hitachi Ltd
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 Hitachi Ltd filed Critical Hitachi Ltd
Publication of EP1849960A2 publication Critical patent/EP1849960A2/de
Publication of EP1849960A3 publication Critical patent/EP1849960A3/de
Withdrawn legal-status Critical Current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
    • F01D5/18—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
    • 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
    • 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
    • F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
    • F01D9/041—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector using blades
    • 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
    • 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
    • F05D2260/2214—Improvement of heat transfer by increasing the heat transfer surface
    • F05D2260/22141—Improvement of heat transfer by increasing the heat transfer surface using fins or ribs

Definitions

  • the present invention relates to a turbine blade having an internal cooling passage.
  • a gas turbine is an apparatus in which fuel and air compressed by a compressor are mixedly burned by a burner to obtain a high temperature and high pressure working gas, which drives a turbine, thereby providing rotational energy.
  • This rotational energy is used as power source to provide electric energy using a generator or to drive a pump or the like.
  • HAT humid air turbine
  • the heat load of a turbine high-temperature portion tends to increase from year to year because of increased temperature of a gas turbine working gas or of high-temperature gas added with moisture.
  • a turbine blade is one of the high-temperature components of a turbine.
  • the turbine blade is provided with an internal hollow cooling passage, through which a cooling medium is allowed to flow, thereby cooling the blade material to a level not higher than a permissible temperature.
  • the bleed air or discharged air of the compressor is frequently used as the cooling medium. If the air from the compressor is used, since an increase in the flow rate of cooling air implies a reduction in amount of burning air, the efficiency of a gas turbine is lowered. Thus, it is desirable that the turbine blade be efficiently cooled with a smaller amount of air.
  • a gas turbine moving blade has pin fins provided on a to-be-cooled portion on the trailing side thereof. To enhance an effect of cooling the vicinity of the trailing edge side tip of the moving blade, the moving blade has a plurality of cooling air passages whose outlets are radially arranged in the trailing edge portion thereof.
  • JP-A-5-156901 discloses a configuration in which to cool the trailing edge side of a gas turbine moving blade having pin fins in a to-be-cooled portion on the trailing edge of the blade, cooling air is fed from the blade-tip side and root side of the a pin fin installed area.
  • the configuration of radially arranging the outlets of the plurality of cooling passages in the trailing edge portion of the blade insufficiently cools to-be-cooled portions other than the trailing edge portion of the blade.
  • Increasing the number of the cooling passages increases the necessary amount of cooling air, lowering the efficiency of the gas turbine. Reducing the number of the cooling passages enlarges the diameters thereof, which leads to a slower flow rate.
  • an effect of cooling portions other than the pin fin installed portion is reduced.
  • the configuration where the partition plate is provided to direct a portion of cooling air to the tip portion of the blade can slightly increase the amount of cooling air fed to the trailing edge side tip portion of the blade.
  • the cooling air does not reach the tip portion of the blade, that is, it cannot cool the vicinity of the trailing edge side tip of the blade.
  • the configuration where the cooling air is fed from the tip side and root side of the blade makes it difficult for the cooling air to flow in the vicinity of the intermediate area between the tip side and root side of the blade. This area cannot be sufficiently cooled. That is to say, the configuration of a cooling passage of a blade is desired which sufficiently cool the entire to-be-cooled portion of the trailing edge portion of the blade.
  • a turbine blade which includes an internal cooling passage and a plurality of projections on a blade ventral side wall surface and a rear side wall surface of the cooling passage on a blade trailing edge side thereof or a plurality of portions connecting the blade ventral side wall surface with the rear side wall surface, wherein the projections or the portions are arranged so as to have different passage resistances depending on a position on the wall surfaces on the blade trailing edge side.
  • the present invention can provide a turbine blade that can enhance efficiency of a gas turbine by being efficiently cooled and has a high degree of reliability.
  • the present invention can be applied to the moving blade or stationary blade of a gas turbine.
  • the following embodiments use compressor bleed air or discharged air is used as a cooling medium by way of example.
  • Fig. 1 is a longitudinal cross-sectional view of a turbine moving blade according to the embodiment of the present invention.
  • the turbine moving blade 1 includes a shank section 2 located on the root side thereof and a blade section 3 located on the tip side thereof.
  • Hollow cooling passages 4, 5 are provided to extend from the inside of the shank section 2 to the inside of the blade section 3.
  • Compressor bleed air or discharged air is fed to the cooling passages 4 and 5 through cooling medium supply holes 14 and 15, respectively.
  • direction 12 denotes a blade trailing edge side.
  • the cooling passage 4 is partitioned by partition walls 6a, 6b into cooling passages 7a, 7b, 7c in the blade section 3.
  • the cooling passages 7a, 7b, 7c together with a tip bent portion 8a and a lower end bent portion 9a form a serpentine cooling passage which is a fold passage.
  • the cooling passage 5 is partitioned by partition walls 6c, 6d into cooling passages 7d, 7e, 7f in the blade section 3.
  • the cooling passages 7d, 7e, 7f together with a tip bent portion 8b and a lower end bent portion 9b form a serpentine cooling passage.
  • a blowout section 13 is provided on the trailing edge side of the cooling passage 7f to allow the cooling air flowing in the cooling passage 5 to flow to the outside of the blade.
  • a plurality of pin fins 16 is provided in the blowout section 13 so as to be integral with the blade.
  • the present embodiment uses the pin fins 16 which are each formed almost cylindrical.
  • Cooling air is fed from a rotor disk (not shown) carrying the turbine moving blade 1 to the supply holes 14, 15 and internally cools the moving blade 1 while passing through the cooling passages 4, 5.
  • Most of the cooling air flowing into the supply hole 14 flows out through a blowout hole 11 provided at the tip of the moving blade.
  • Most of the cooling air flowing into the supply hole 15 flows out to the outside from a blowout section 13 provided at the trailing edge of the moving blade.
  • the turbine moving blade 1 of the present embodiment includes six cooling passages in total to cool the blade section 3: three passages extending from the blade root side toward the blade tip side and three passages extending from the blade tip side toward the blade root side.
  • six cooling passages in total to cool the blade section 3 three passages extending from the blade root side toward the blade tip side and three passages extending from the blade tip side toward the blade root side.
  • the turbine moving blade 1 of the present embodiment includes two cooling systems in total of cooling passages of cooling the blade section 3: the cooling passage 4 adapted to feed cooling air from the supply hole 14 and the cooling passage 5 adapted to feed cooling air from the supply hole 15.
  • a sufficient amount of cooling air must be supplied to the supply hole of each system. Therefore, if the number of the cooling systems of the cooling passage is increased, the necessary amount of cooling air to be supplied is increased.
  • the compressor bleed air or discharged air, which is a portion of a working medium, is often used as cooling air.
  • the increased necessary amount of cooling air leads to a reduction in amount of burning air to be supplied to a burner, which lowers the efficiency of the gas turbine.
  • the present embodiment can suppress the total number of the cooling systems to as small as two, an amount of supply cooling air is small, which can enhance the efficiency of the gas turbine. Taking into consideration the efficiency of the gas turbine, it is desired that the total number of cooling systems be two or less in total for the turbine blade used in the same application as in that of the present embodiment.
  • Fig. 2 is a cross-sectional view of the moving blade taken along line A-A of Fig. 1.
  • Direction 22 denotes the blade trailing edge side.
  • the pin fin 16 is an almost-circular constituent element which extends from the blade ventral side wall 28 to the blade rear side wall 29.
  • Fig. 3 is an enlarged view of the cooling passage 7f and blowout section 13 of the turbine moving blade 1 depicted in Fig. 1.
  • Flows 18, 19a-19e denote the flows of cooling air.
  • a portion of cooling air is led as flow 18 to the blade trailing edge side tip by the cooling passage 7f. While the cooling air is led as flow 18 to the blade tip, a portion of flow 18 gradually flows to the outside as flows 19a-19d from the blowout section 13.
  • the cooling passage 7f is provided on the blade leading edge side of the blowout section 13 provided with the pin fins.
  • the blade root portion is provided with large pin fins 16b whereas the blade tip portion is provided with small pin fins 16a.
  • the passage resistance of the cooling medium encountered when the cooling medium blows out to the outside from the area provided with the large pin fins 16b, which is the blade root side area of the blowout section 13 can be made larger than that from the area provided with the small pin fins 16a.
  • the flow volume of flows 19a-19c can be made small whereas the flow volume of flows 19d, 19e can be made large.
  • this configuration can provide two effects, an improvement in reliability of the turbine blade and an improvement in efficiency of the gas turbine, similar to the effects obtained by providing the cooling passage 7f.
  • the passage resistance of the present embodiment means how difficult it is for the cooling air to flow in the blowout section 13 when the cooling air flows from the cooling passage 7f through the blowout section 13, which is the area provided with the pin fins 16, to the outside.
  • a large pin fin means that the cross-section area of its surface almost parallel to the surface (the blade ventral side wall 28 or the blade rear side wall 29) provided with the pin fin 16 is large.
  • a small pin fin means that the cross-section area of the pin fin 16 is small.
  • a ratio of the cross-section area of the large pin fin to that of the small pin fin should be optimally designed based on the cooling capacity required by the upstream side and downstream side of arrow 18. The ratio is generally selected from about 1.4 to 4.0.
  • the cross-sectional shape of the pin fin is circular as in the embodiment, about 1.4 to 4.0 corresponds to about 1.2 to 2.0 in a ratio between circular diameters.
  • the small pins 16a are provided on the blade tip side. If the cooling capacity of other portions such as the blade root side and the like is intended to be increased, it is conceivable that the pin fins 16 at appropriate portions are made larger or smaller.
  • FIG. 4 is a longitudinal cross-sectional view of a turbine moving blade 1 according to the second embodiment of the present invention.
  • Fig. 5 is an enlarged view of a cooling passage 7f and a blowout section 13 of the turbine moving blade 1 depicted in Fig. 4.
  • the number of pin fins 16 per unit area i.e., the installation density of the pin fins 16, of a part in a blowout section 3 is made different from that of another part in the blowout section 3.
  • the installation density of pin fins 16c in an area (on the blade tip side) corresponding to passages of flaws 19d, 19e in the blowout section 13 is made lower than that of pin fins 16d in an area (on the blade tip side) corresponding to passages of flows 19a-19c in the blowout section 13.
  • the present embodiment changes intervals between the pin fins 16 to be arranged.
  • the interval ratio should be optimally designed based on the cooling capacity required by the blade tip side and the blade root side.
  • the interval ratio is generally selected from about 1.1 to 2.0 by way of example.
  • the installation density of the pin fins 16c installed on the blade tip side is made small. If the cooling capacity of other portions such as the blade root side and the like is intended to be increased, it is conceivable that the density of the pin fins 16 at appropriate portions are made larger or smaller.
  • FIG. 6 is a longitudinal cross-sectional view of a turbine moving blade according to the third embodiment of the invention.
  • Fig. 7 is an enlarged view of a cooling passage 7f and a blowout section of the turbine moving blade 1 depicted in Fig. 6.
  • the installation density and size of the pin fins 16 of a part in a blowout section 13 is made different from those of another part in the blowout section 13.
  • small pin fins 16e and large pin fins 16f are provided on the blade tip side and blade root side, respectively, of the blowout section 13 and the installation density of the small pin fins 16e is made smaller than that of the large pin fins 16f.
  • the configuration described above can provide an improvement in reliability of the turbine blade and an improvement in efficiency of the gas turbine as described in the first and second embodiments.
  • the turbine blade of the present embodiment includes the pin fins having the different sizes and further different installation densities. Therefore, the turbine blade of this embodiment can provide the larger effects than those of the first and second embodiments.
  • the gas turbine moving blade 1 has been thus far described taking the turbine blade as an example. However, the application of the present invention is not limited to the moving blade.
  • a fourth embodiment describes a gas turbine stationary blade embodying the invention.
  • FIG. 8 is a longitudinal cross-sectional view of a turbine stationary blade according to the fourth embodiment of the present invention.
  • a turbine stationary blade 61 includes cooling passages 67a, 67b inside a blade section 63.
  • Direction 72 denotes a blade trailing edge side
  • direction 85 denotes a blade root side
  • direction 86 denotes a blade tip side.
  • the cooling passages 67a and 67b are partitioned by a partition wall 66.
  • a plurality of pin fins 76 are provided in a blowout section 73 on the blade trailing edge side of the cooling passage 67b.
  • Flows 64, 65, 79a-79e denotes flows of cooling air.
  • the cooling air fed from a cooing air supply hole 75 flows in the cooling passage 67b toward the blade tip side.
  • Flow 65 toward the blade tip side branches into flows 79a, 79b, 79c, 79d, 79e, which cool the blowout section 73 and flow out to the outside of the blade.
  • small pin fins 76a are installed in the area 87 and large pin fins 76b are installed in the other areas in the blowout section 73.
  • the passage resistance of cooling air flowing out as flow 79c from the area 87 can be made lower than that as flow 79a, 79b, 79d, 79e from the other areas, thereby feeding a sufficient amount of cooling air.
  • this embodiment can provide two effects, an improvement in reliability of the turbine blade and an improvement in efficiency of the gas turbine.
  • the size of the pin fin 76 in an area of the blowout section 73 is made different from that in the other areas, whereby the passage resistance of cooling air flowing out through the area from the blowout section 73 is made different from that through the other areas.
  • the passage resistance in an area of the blowout section 73 is made different from that in other areas by other methods of, e.g., making the installation density in an area different from that in other areas.
  • the sizes or installation densities of the pin fins 16 are made different from each other depending on the areas. This makes it easy for cooling air to reach an area where it is otherwise difficult for the cooling air to reach, such as the blade trailing edge side tip portion or the like, or an area which particularly needs to be cooled. Thus, the turbine blade is effectively cooled.
  • the effective cooling is achieved by providing areas in the blowout section provided with the pin fins, the areas being high and low in passage resistance of the cooling air flowing out from the blowout section. That is to say, if the passage resistances can be made different from each other depending on flowing-out portions of the cooling air in the blowout section, the two effects can be provided, an improvement in reliability of the turbine blade and an improvement in efficiency of the gas turbine.
  • Allowing the pin fins 16 to have different sizes and/or different installation densities depending on areas is to allow them to have different pin fin installation areas per unit area on the blade ventral side wall or on the blade rear side wall.
  • the turbine blade configured as described above can provide the two effects mentioned above.
  • the pin fin 16 used in the embodiments described above is formed almost cylindrically to be integral with the blade but the present invention is not limited to this pin fin.
  • a triangular fin which extends from the blade ventral side wall 28 to the blade rear side wall 29 may be used as a structure which plays the same role as the pin fin 16.
  • projections may be provided on the blade ventral side wall 28 or the blade rear side wall 29.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP07008241A 2006-04-27 2007-04-23 Turbinenschaufel mit innerem Kühlkanal Withdrawn EP1849960A3 (de)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
JP2006122884A JP2007292006A (ja) 2006-04-27 2006-04-27 内部に冷却通路を有するタービン翼

Publications (2)

Publication Number Publication Date
EP1849960A2 true EP1849960A2 (de) 2007-10-31
EP1849960A3 EP1849960A3 (de) 2010-03-10

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EP07008241A Withdrawn EP1849960A3 (de) 2006-04-27 2007-04-23 Turbinenschaufel mit innerem Kühlkanal

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JP (1) JP2007292006A (de)

Cited By (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2009109462A1 (de) * 2008-03-07 2009-09-11 Alstom Technology Ltd Schaufel für eine gasturbine
WO2013076109A1 (de) * 2011-11-21 2013-05-30 Siemens Aktiengesellschaft Kühlbares heissgasbauteil für eine gasturbine
CN104791019A (zh) * 2014-01-17 2015-07-22 通用电气公司 涡轮叶片及用于延长涡轮叶片寿命的方法
WO2015116338A1 (en) 2014-01-30 2015-08-06 United Technologies Corporation Trailing edge cooling pedestal configuration for a gas turbine engine airfoil
CN107366555A (zh) * 2016-05-12 2017-11-21 通用电气公司 叶片以及涡轮转子叶片
US10208604B2 (en) 2016-04-27 2019-02-19 United Technologies Corporation Cooling features with three dimensional chevron geometry
US11313238B2 (en) * 2018-09-21 2022-04-26 Doosan Heavy Industries & Construction Co., Ltd. Turbine blade including pin-fin array

Families Citing this family (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
ES2542064T3 (es) * 2008-03-28 2015-07-30 Alstom Technology Ltd Álabe de guía para una turbina de gas y turbina de gas con un álabe de guía de esta clase
JP5189406B2 (ja) 2008-05-14 2013-04-24 三菱重工業株式会社 ガスタービン翼およびこれを備えたガスタービン
US20140064983A1 (en) * 2012-08-31 2014-03-06 General Electric Company Airfoil and method for manufacturing an airfoil

Citations (13)

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Publication number Priority date Publication date Assignee Title
US4278400A (en) * 1978-09-05 1981-07-14 United Technologies Corporation Coolable rotor blade
EP0032646A1 (de) * 1980-01-10 1981-07-29 Societe Nationale D'etude Et De Construction De Moteurs D'aviation, "S.N.E.C.M.A." Gasturbinenleitschaufel
GB2112467A (en) * 1981-12-28 1983-07-20 United Technologies Corp Coolable airfoil for a rotary machine
JPH05156901A (ja) * 1991-12-02 1993-06-22 Hitachi Ltd ガスタービン冷却静翼
JPH06137102A (ja) * 1992-10-26 1994-05-17 Mitsubishi Heavy Ind Ltd ガスタービン中空動翼
US5538394A (en) * 1993-12-28 1996-07-23 Kabushiki Kaisha Toshiba Cooled turbine blade for a gas turbine
US5772397A (en) * 1996-05-08 1998-06-30 Alliedsignal Inc. Gas turbine airfoil with aft internal cooling
DE19807563A1 (de) * 1997-02-25 1998-09-24 Mitsubishi Heavy Ind Ltd Kühlkonstruktion zum Kühlen einer Montageplatte für Antriebsschaufeln einer Gasturbine
GB2349920A (en) * 1999-05-10 2000-11-15 Abb Alstom Power Ch Ag Cooling arrangement for turbine blade
EP1467065A2 (de) * 2003-04-08 2004-10-13 United Technologies Corporation Turbinenschaufel
EP1538305A2 (de) * 2003-11-19 2005-06-08 United Technologies Corporation Schaufel mit Stegenanordnung von variabler Dichte an der Abströmkante
EP1674661A2 (de) * 2004-12-23 2006-06-28 United Technologies Corporation Kühlluftkanal für eine Turbinenschaufel
EP1715139A2 (de) * 2005-04-22 2006-10-25 United Technologies Corporation Kühlung der Abströmkante einer Turbinenschaufel

Patent Citations (13)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4278400A (en) * 1978-09-05 1981-07-14 United Technologies Corporation Coolable rotor blade
EP0032646A1 (de) * 1980-01-10 1981-07-29 Societe Nationale D'etude Et De Construction De Moteurs D'aviation, "S.N.E.C.M.A." Gasturbinenleitschaufel
GB2112467A (en) * 1981-12-28 1983-07-20 United Technologies Corp Coolable airfoil for a rotary machine
JPH05156901A (ja) * 1991-12-02 1993-06-22 Hitachi Ltd ガスタービン冷却静翼
JPH06137102A (ja) * 1992-10-26 1994-05-17 Mitsubishi Heavy Ind Ltd ガスタービン中空動翼
US5538394A (en) * 1993-12-28 1996-07-23 Kabushiki Kaisha Toshiba Cooled turbine blade for a gas turbine
US5772397A (en) * 1996-05-08 1998-06-30 Alliedsignal Inc. Gas turbine airfoil with aft internal cooling
DE19807563A1 (de) * 1997-02-25 1998-09-24 Mitsubishi Heavy Ind Ltd Kühlkonstruktion zum Kühlen einer Montageplatte für Antriebsschaufeln einer Gasturbine
GB2349920A (en) * 1999-05-10 2000-11-15 Abb Alstom Power Ch Ag Cooling arrangement for turbine blade
EP1467065A2 (de) * 2003-04-08 2004-10-13 United Technologies Corporation Turbinenschaufel
EP1538305A2 (de) * 2003-11-19 2005-06-08 United Technologies Corporation Schaufel mit Stegenanordnung von variabler Dichte an der Abströmkante
EP1674661A2 (de) * 2004-12-23 2006-06-28 United Technologies Corporation Kühlluftkanal für eine Turbinenschaufel
EP1715139A2 (de) * 2005-04-22 2006-10-25 United Technologies Corporation Kühlung der Abströmkante einer Turbinenschaufel

Cited By (11)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2009109462A1 (de) * 2008-03-07 2009-09-11 Alstom Technology Ltd Schaufel für eine gasturbine
US8182225B2 (en) 2008-03-07 2012-05-22 Alstomtechnology Ltd Blade for a gas turbine
WO2013076109A1 (de) * 2011-11-21 2013-05-30 Siemens Aktiengesellschaft Kühlbares heissgasbauteil für eine gasturbine
EP2602439A1 (de) * 2011-11-21 2013-06-12 Siemens Aktiengesellschaft Kühlbares Heißgasbauteil für eine Gasturbine
CN104791019A (zh) * 2014-01-17 2015-07-22 通用电气公司 涡轮叶片及用于延长涡轮叶片寿命的方法
WO2015116338A1 (en) 2014-01-30 2015-08-06 United Technologies Corporation Trailing edge cooling pedestal configuration for a gas turbine engine airfoil
US20160333699A1 (en) * 2014-01-30 2016-11-17 United Technologies Corporation Trailing edge cooling pedestal configuration for a gas turbine engine airfoil
EP3099901A4 (de) * 2014-01-30 2017-02-01 United Technologies Corporation Austrittskantenkühlsockelkonfiguration für eine gasturbinenmotorschaufel
US10208604B2 (en) 2016-04-27 2019-02-19 United Technologies Corporation Cooling features with three dimensional chevron geometry
CN107366555A (zh) * 2016-05-12 2017-11-21 通用电气公司 叶片以及涡轮转子叶片
US11313238B2 (en) * 2018-09-21 2022-04-26 Doosan Heavy Industries & Construction Co., Ltd. Turbine blade including pin-fin array

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Publication number Publication date
JP2007292006A (ja) 2007-11-08
EP1849960A3 (de) 2010-03-10

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