EP0207799A2 - Canaux de refroidissement pour les aubes d'une turbine à gaz - Google Patents

Canaux de refroidissement pour les aubes d'une turbine à gaz Download PDF

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Publication number
EP0207799A2
EP0207799A2 EP86305182A EP86305182A EP0207799A2 EP 0207799 A2 EP0207799 A2 EP 0207799A2 EP 86305182 A EP86305182 A EP 86305182A EP 86305182 A EP86305182 A EP 86305182A EP 0207799 A2 EP0207799 A2 EP 0207799A2
Authority
EP
European Patent Office
Prior art keywords
flow area
blade
section
coolant
hub
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Withdrawn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
EP86305182A
Other languages
German (de)
English (en)
Other versions
EP0207799A3 (fr
Inventor
Paul Clarence Holden
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.)
Westinghouse Electric Corp
Original Assignee
Westinghouse Electric 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 Westinghouse Electric Corp filed Critical Westinghouse Electric Corp
Publication of EP0207799A2 publication Critical patent/EP0207799A2/fr
Publication of EP0207799A3 publication Critical patent/EP0207799A3/fr
Withdrawn legal-status Critical Current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
    • 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

Definitions

  • the present invention relates to rotor blades in a combustion turbine and more particularly to coolant systems therefor.
  • Rotating turbine blades are frequently cooled with air flowing radially outward through a plurality of holes which pass from the blade root to the blade tip.
  • the holes typically have either a constant diameter along the airfoil and root portions of the holes or a first constant diameter along the airfoil portion and a second constant diameter along the root portion of the holes.
  • the diameter of the hole along the root portion typically being larger to avoid pressure loss in a region that does not require appreciable cooling.
  • the airfoil portions of the coolant holes need to be relatively small in cross section to produce the high coolant velocity and heat transfer coefficient required there.
  • the critical design region of the blade for both stress and cooling is the center span portion of the blade, and the hole diameter, the number of holes and the coolant flow are normally set by design considerations. Since the coolant heats up appreciably as it flows outwardly along the blade, having received heat from the hot blade path gas, the coolant temperature at the center span region becomes considerably higher than the coolant temperature near the blade hub.
  • the lower coolant temperature near the blade hub tends to cool the blade to a lower temperature than is needed from the standpoint of stress design.
  • Overcoo'ling near the blade hub means the coolant absorbs more heat than necessary from the hub region which results in increased coolant temperature in the center span region and thus higher coolant flows and/or higher metal temperatures than if the overcooling did not occur.
  • overcooling since reduced centrifugal stress usually more than offsets the coolant heatup in the blade tip region, overcooling also tends to occur there.
  • An important consequence of overcooling in both the hub and tip regions is a higher level of pressure loss than would be encountered if the coolant flow per unit of flow area in overcooled regions were reduced to produce a heat transfer level matching the cooling requirements.
  • Reduction of pressure loss in overcooled regions would make possible higher cooling flow per unit of flow area in the mid-span region in order to provide increased cooling for a given supply pressure at the blade root. This can be translated into a design for higher turbine inlet temperature or one with reduced cooling flow for given turbine inlet temperatures.
  • cores are typically used to form the radial coolant holes in a blade. Structural weakness of the cores has resulted in core breakage and blade scrapping more frequently than desirable thereby adding to per unit manufacturing costs of blades.
  • the present invention resides in a combustion turbine rotor blade having a root portion and an airfoil portion extending therefrom, said airfoil portion having a plurality of coolant holes extending from said root portion along its span to its tip to provide flow of coolant therethrough from said root portion, each of said coolant holes having an inner hub portion and a second portion extending outwardly from said hub portion, characterized by said hub and second hole portions having relative flow areas which cause the coolant to have reduced flow per unit of flow area in said hub hole portion relative to the coolant flow per unit of flow area through said second hole portion over a mid-span region of said blade thereby avoiding hub blade region overcooling.
  • FIG. 1 a combustion turbine blade 10 having a root portion 12 and an airfoil portion 14 arranged in accordance with the principles of the invention.
  • Coolant holes or channels 16 extend radially outwardly along the span of the blade. Air flow supplied from complementary coolant holes in the blade root and directed outwardly along the coolant holes 16 cools the blade airfoil portion 14 and its surface which is exposed to the hot blade path gas.
  • a center span portion 18 of the blade airfoil is the critical design region for which the number of holes, the hole diameter and the coolant flow are set to meet its cooling needs.
  • the total blade coolant system structure is then designed to support the critical region needs.
  • the temperature of the coolant in the holes 16 is lower in a blade hub region 20 than in the mid-span region 18 because the coolant heats up as it flows outwardly along the blade.
  • the lower temperature coolant tends to overcool the blade hub 20 resulting in coolant temperature in the mid-span region 18 being higher than it would be if overcooling did not occur.
  • coolant flow and/or metal temperature is higher than it could be if coolant temperature rise were reduced with the elimination of overcooling.
  • blade hub overcooling is substantially reduced or eliminated to permit lower coolant flow and/or lower metal temperature resulting in greater cooling and turbine efficiency.
  • the coolant hole 16 includes a hub portion 22 and an outer portion 24 which extends from the hub portion 22 through the blade mid-span region to the blade tip 26.
  • the hub portion 22 of the coolant hole 16 is tapered from a first diameter at its inlet end to a smaller diameter where it joins the outer hole portion 24.
  • the diameter of the outer hole portion 24 is substantially constant along its length.
  • the hole flow area in the hub region 20 is increased relative to the hole flow area in the mid-span region 18.
  • Reduced coolant flow per unit of flow area in the hub region 20 reduces the hub region heat transfer coefficient and the amount of cooling.
  • Turbine and blade design parameters determine the amount and length of taper to reduce or substantially eliminate overcooling in the hub region 20.
  • Tapering of the coolant holes also reduces pressure loss in the coolant holes. Accordingly, higher flow per unit of flow area is obtained in the mid-span. region to obtain lower metal temperature for a given supply pressure at the blade root. Spanwise hole cooling technology can thus be employed with higher turbine inlet temperature levels. The higher flow per unit of flow area in the design section can also be used to reduce cooling flow for a given level of turbine inlet temperatures.
  • a coolant hole 16a includes a tapered hub portion 22a, a constant diameter mid-span portion 24a and a tip portion 25a which is tapered outwardly in the direction of coolant flow.
  • the constant diameter mid-span portion 24a can be eliminated from the hole 16a so that the hole 16a is formed by the oppositely tapered portions 22a and 25a in end-to- end relation.
  • the junction point of the hole portions 22a and 25a in this alternative would be determined by stress and heat transfer considerations.
  • the tapered hole portions be provided with continuous tapering, either linear as shown or nonlinear as warranted by design considerations.
  • the coolant holes have a circular cross-section but the invention can be implemented with non-circular cross- sections.
  • FIG 4 there is shown another embodiment of the invention in which a coolant hole 16b has a stepped configuration to provide a varying cross-section along its length for substantially reduced blade region overcooling.
  • the coolant hole 16b includes a hub portion 22b having an inner section 22bl having a first diameter and an outer section 22b2 having a lesser diameter.
  • a mid-span section 24b having a further reduced diameter adjoins the outer hub section 22b2.
  • a tip portion 25b of the coolant hole 16b has a first section 25bl and a second section 25b2 with successively greater diameters.
  • tapered blade coolant holes provides a further advantage by permitting the blades to be manufactured by casting processes using cores for coolant hole formation.
  • tapered cores can be employed to form the coolant holes 16 and 16a, and tapered cores with larger diameter in the hub and tip region are characteristically stronger than cores conventionally used to form constant diameter coolant holes in the airfoil. With stronger cores, there is reduced core breakage, reduced blade scrapping and reduced manufacturing costs.
  • the stronger tapered cores permit manufacture of smaller hole diameters in the mid-span portion of the coolant holes. In turn, improved blade cooling and reduced blade coolant flow can be realized.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP86305182A 1985-07-03 1986-07-03 Canaux de refroidissement pour les aubes d'une turbine à gaz Withdrawn EP0207799A3 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
US75165785A 1985-07-03 1985-07-03
US751657 1985-07-03

Publications (2)

Publication Number Publication Date
EP0207799A2 true EP0207799A2 (fr) 1987-01-07
EP0207799A3 EP0207799A3 (fr) 1988-09-14

Family

ID=25022941

Family Applications (1)

Application Number Title Priority Date Filing Date
EP86305182A Withdrawn EP0207799A3 (fr) 1985-07-03 1986-07-03 Canaux de refroidissement pour les aubes d'une turbine à gaz

Country Status (5)

Country Link
EP (1) EP0207799A3 (fr)
JP (1) JPS6210402A (fr)
CN (1) CN86104500A (fr)
CA (1) CA1262868A (fr)
IE (1) IE861475L (fr)

Cited By (24)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP0319758A1 (fr) * 1987-12-08 1989-06-14 General Electric Company Refroidissement de l'extrémité d'une aube par diffusion
EP0550184A1 (fr) * 1991-12-30 1993-07-07 General Electric Company Canaux de refroidissement avec promoteurs de turbulence pour des aubes de turbines à gaz
US6539627B2 (en) 2000-01-19 2003-04-01 General Electric Company Method of making turbulated cooling holes
WO2004057157A1 (fr) * 2002-12-21 2004-07-08 John Macdonald Aube de turbine
EP1541805A1 (fr) * 2003-12-12 2005-06-15 General Electric Company Aube avec trous de refroidissement
US6994514B2 (en) * 2002-11-20 2006-02-07 Mitsubishi Heavy Industries, Ltd. Turbine blade and gas turbine
EP1820937A2 (fr) 2006-02-15 2007-08-22 United Technologies Corporation Pale de turbine dotée de canaux radiaux de refroidissement
US20100329862A1 (en) * 2009-06-24 2010-12-30 General Electric Company Cooling Hole Exits for a Turbine Bucket Tip Shroud
EP2354453A1 (fr) * 2010-02-02 2011-08-10 Siemens Aktiengesellschaft Composant de moteur à turbine pour un refroidissement adaptatif
US8157527B2 (en) * 2008-07-03 2012-04-17 United Technologies Corporation Airfoil with tapered radial cooling passage
US20130052035A1 (en) * 2011-08-24 2013-02-28 General Electric Company Axially cooled airfoil
US8506251B2 (en) 2010-03-03 2013-08-13 Mitsubishi Heavy Industries, Ltd. Gas turbine blade, manufacturing method therefor, and gas turbine using turbine blade
US8511992B2 (en) * 2008-01-22 2013-08-20 United Technologies Corporation Minimization of fouling and fluid losses in turbine airfoils
EP2743454A1 (fr) * 2012-12-11 2014-06-18 General Electric Company Composant de turbine comprenant des passages de refroidissement à diamètre variable
US9200526B2 (en) 2010-12-21 2015-12-01 Kabushiki Kaisha Toshiba Transition piece between combustor liner and gas turbine
WO2017121689A1 (fr) 2016-01-15 2017-07-20 Siemens Aktiengesellschaft Profil aérodynamique de turbine à gaz
US10408079B2 (en) 2015-02-18 2019-09-10 Siemens Aktiengesellschaft Forming cooling passages in thermal barrier coated, combustion turbine superalloy components
US10686199B2 (en) 2012-08-14 2020-06-16 Loop Energy Inc. Fuel cell flow channels and flow fields
US10734661B2 (en) 2012-08-14 2020-08-04 Loop Energy Inc. Fuel cell components, stacks and modular fuel cell systems
US10930942B2 (en) 2016-03-22 2021-02-23 Loop Energy Inc. Fuel cell flow field design for thermal management
US11060195B2 (en) 2012-08-14 2021-07-13 Loop Energy Inc. Reactant flow channels for electrolyzer applications
CN113454322A (zh) * 2019-03-29 2021-09-28 三菱动力株式会社 高温部件以及高温部件的制造方法
US20220170376A1 (en) * 2019-06-05 2022-06-02 Mitsubishi Power Ltd. Turbine blade, manufacturing method for turbine blade, and gas turbine
US12586798B2 (en) 2021-07-03 2026-03-24 Cevizdere LLC Methods and apparatus for mold mitigation in fuel cell humidifiers

Families Citing this family (11)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE4430302A1 (de) * 1994-08-26 1996-02-29 Abb Management Ag Prallgekühltes Wandteil
EP1712739A1 (fr) * 2005-04-12 2006-10-18 Siemens Aktiengesellschaft Elément avec trou de refroidissement par film
CN1318735C (zh) * 2005-12-26 2007-05-30 北京航空航天大学 一种适用于燃气涡轮发动机的脉动冲击冷却叶片
JP2010053749A (ja) * 2008-08-27 2010-03-11 Mitsubishi Heavy Ind Ltd タービン用翼
US8201621B2 (en) * 2008-12-08 2012-06-19 General Electric Company Heat exchanging hollow passages with helicoidal grooves
WO2011108164A1 (fr) * 2010-03-03 2011-09-09 三菱重工業株式会社 Pale de rotor pour turbine à gaz, son procédé de fabrication, et turbine à gaz utilisant la pale de rotor
US8727724B2 (en) * 2010-04-12 2014-05-20 General Electric Company Turbine bucket having a radial cooling hole
US9181819B2 (en) * 2010-06-11 2015-11-10 Siemens Energy, Inc. Component wall having diffusion sections for cooling in a turbine engine
US10577943B2 (en) * 2017-05-11 2020-03-03 General Electric Company Turbine engine airfoil insert
CN110159357B (zh) * 2019-06-04 2021-01-29 北京航空航天大学 一种提升被动安全的航空发动机涡轮叶片缩扩型供气通道
JP6637630B1 (ja) * 2019-06-05 2020-01-29 三菱日立パワーシステムズ株式会社 タービン翼およびタービン翼の製造方法並びにガスタービン

Family Cites Families (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB1018747A (en) * 1964-11-13 1966-02-02 Rolls Royce Aerofoil shaped blade for fluid flow machines
JPS4825103B1 (fr) * 1967-06-05 1973-07-26
JPS5520041A (en) * 1978-07-29 1980-02-13 Noto Denshi Kogyo Kk Piezoelectric device
JPS5669423A (en) * 1979-11-09 1981-06-10 Hitachi Ltd Air-cooled blade of gas turbine

Cited By (38)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4893987A (en) * 1987-12-08 1990-01-16 General Electric Company Diffusion-cooled blade tip cap
EP0319758A1 (fr) * 1987-12-08 1989-06-14 General Electric Company Refroidissement de l'extrémité d'une aube par diffusion
EP0550184A1 (fr) * 1991-12-30 1993-07-07 General Electric Company Canaux de refroidissement avec promoteurs de turbulence pour des aubes de turbines à gaz
US5413463A (en) * 1991-12-30 1995-05-09 General Electric Company Turbulated cooling passages in gas turbine buckets
US6824360B2 (en) 2000-01-19 2004-11-30 General Electric Company Turbulated cooling holes
US6539627B2 (en) 2000-01-19 2003-04-01 General Electric Company Method of making turbulated cooling holes
US6994514B2 (en) * 2002-11-20 2006-02-07 Mitsubishi Heavy Industries, Ltd. Turbine blade and gas turbine
EP1422383A3 (fr) * 2002-11-20 2006-05-31 Mitsubishi Heavy Industries, Ltd. Refroidissement d'une aube de turbine à gaz
WO2004057157A1 (fr) * 2002-12-21 2004-07-08 John Macdonald Aube de turbine
EP1541805A1 (fr) * 2003-12-12 2005-06-15 General Electric Company Aube avec trous de refroidissement
US6997679B2 (en) * 2003-12-12 2006-02-14 General Electric Company Airfoil cooling holes
EP1820937A2 (fr) 2006-02-15 2007-08-22 United Technologies Corporation Pale de turbine dotée de canaux radiaux de refroidissement
EP1820937A3 (fr) * 2006-02-15 2011-02-23 United Technologies Corporation Pale de turbine dotée de canaux radiaux de refroidissement
US8511992B2 (en) * 2008-01-22 2013-08-20 United Technologies Corporation Minimization of fouling and fluid losses in turbine airfoils
US8157527B2 (en) * 2008-07-03 2012-04-17 United Technologies Corporation Airfoil with tapered radial cooling passage
US8511990B2 (en) * 2009-06-24 2013-08-20 General Electric Company Cooling hole exits for a turbine bucket tip shroud
US20100329862A1 (en) * 2009-06-24 2010-12-30 General Electric Company Cooling Hole Exits for a Turbine Bucket Tip Shroud
JP2011007181A (ja) * 2009-06-24 2011-01-13 General Electric Co <Ge> タービンバケット先端シュラウド用の冷却孔出口
EP2354453A1 (fr) * 2010-02-02 2011-08-10 Siemens Aktiengesellschaft Composant de moteur à turbine pour un refroidissement adaptatif
US8506251B2 (en) 2010-03-03 2013-08-13 Mitsubishi Heavy Industries, Ltd. Gas turbine blade, manufacturing method therefor, and gas turbine using turbine blade
US8827646B2 (en) 2010-03-03 2014-09-09 Mitsubishi Heavy Industries, Ltd. Gas turbine blade, manufacturing method therefor, and gas turbine using turbine blade
US9200526B2 (en) 2010-12-21 2015-12-01 Kabushiki Kaisha Toshiba Transition piece between combustor liner and gas turbine
US20130052035A1 (en) * 2011-08-24 2013-02-28 General Electric Company Axially cooled airfoil
US11489175B2 (en) 2012-08-14 2022-11-01 Loop Energy Inc. Fuel cell flow channels and flow fields
US10686199B2 (en) 2012-08-14 2020-06-16 Loop Energy Inc. Fuel cell flow channels and flow fields
US10734661B2 (en) 2012-08-14 2020-08-04 Loop Energy Inc. Fuel cell components, stacks and modular fuel cell systems
US12227855B2 (en) 2012-08-14 2025-02-18 Loop Energy Inc. Reactant flow channels for electrolyzer applications
US11060195B2 (en) 2012-08-14 2021-07-13 Loop Energy Inc. Reactant flow channels for electrolyzer applications
EP2743454A1 (fr) * 2012-12-11 2014-06-18 General Electric Company Composant de turbine comprenant des passages de refroidissement à diamètre variable
US10408079B2 (en) 2015-02-18 2019-09-10 Siemens Aktiengesellschaft Forming cooling passages in thermal barrier coated, combustion turbine superalloy components
WO2017121689A1 (fr) 2016-01-15 2017-07-20 Siemens Aktiengesellschaft Profil aérodynamique de turbine à gaz
US11901591B2 (en) 2016-03-22 2024-02-13 Loop Energy Inc. Fuel cell flow field design for thermal management
US10930942B2 (en) 2016-03-22 2021-02-23 Loop Energy Inc. Fuel cell flow field design for thermal management
CN113454322A (zh) * 2019-03-29 2021-09-28 三菱动力株式会社 高温部件以及高温部件的制造方法
CN113454322B (zh) * 2019-03-29 2023-10-31 三菱动力株式会社 高温部件以及高温部件的制造方法
US20220170376A1 (en) * 2019-06-05 2022-06-02 Mitsubishi Power Ltd. Turbine blade, manufacturing method for turbine blade, and gas turbine
US11905848B2 (en) * 2019-06-05 2024-02-20 Mitsubishi Heavy Industries, Ltd. Turbine blade, manufacturing method for turbine blade, and gas turbine
US12586798B2 (en) 2021-07-03 2026-03-24 Cevizdere LLC Methods and apparatus for mold mitigation in fuel cell humidifiers

Also Published As

Publication number Publication date
IE861475L (en) 1987-01-03
JPS6210402A (ja) 1987-01-19
EP0207799A3 (fr) 1988-09-14
CN86104500A (zh) 1987-02-04
CA1262868A (fr) 1989-11-14

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