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US20080008599A1 - Integral main body-tip microcircuits for blades - Google Patents

Integral main body-tip microcircuits for blades Download PDF

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Publication number
US20080008599A1
US20080008599A1 US11/484,143 US48414306A US2008008599A1 US 20080008599 A1 US20080008599 A1 US 20080008599A1 US 48414306 A US48414306 A US 48414306A US 2008008599 A1 US2008008599 A1 US 2008008599A1
Authority
US
United States
Prior art keywords
leg
turbine engine
engine component
component according
cooling
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
US11/484,143
Other languages
English (en)
Inventor
Francisco J. Cunha
William Abdel-Messeh
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.)
RTX Corp
Original Assignee
United Technologies 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 United Technologies Corp filed Critical United Technologies Corp
Priority to US11/484,143 priority Critical patent/US20080008599A1/en
Assigned to UNITED TECHNOLOGIES CORPORATION reassignment UNITED TECHNOLOGIES CORPORATION ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: ABDEL-MESSEH, WILLIAM, CUNHA, FRANCISCO J.
Priority to JP2007168349A priority patent/JP2008019861A/ja
Priority to EP07252702A priority patent/EP1878874B1/en
Publication of US20080008599A1 publication Critical patent/US20080008599A1/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
    • F01D5/186—Film cooling
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2230/00—Manufacture
    • F05D2230/20—Manufacture essentially without removing material
    • F05D2230/21—Manufacture essentially without removing material by casting
    • 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
    • F05D2250/10—Two-dimensional
    • F05D2250/18—Two-dimensional patterned
    • F05D2250/185—Two-dimensional patterned serpentine-like
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

Definitions

  • the present invention relates to a cooling microcircuit for use in a turbine engine component such as a turbine blade.
  • prior art turbine blades have a plurality of cavities with each blade internal cavity feeding a cooling microcircuit located on a side of the airfoil, either on a pressure side or on a suction side.
  • a cooling microcircuit located on a side of the airfoil, either on a pressure side or on a suction side.
  • the flow in the internal supply cavities assume relatively low Mach number distribution when compared to more conventional serpentine cooling designs. Due to these low Mach numbers, the rotational forces, resulting from the angular speeds of the rotor, become predominant. This, in turn, induces a series of cavity vortices inside the supply cavities. Since the internal cavities supply coolant air to the microcircuits embedded in the wall, the location of these supply links, between cavity and microcircuit, become extremely important.
  • FIGS. 1 and 2 there is shown an airfoil with embedded wall circuits having cooling supply flows from a large internal cavity.
  • the embedded circuits on the pressure side would experience a relative lower flow compared to those on the suction side.
  • the cooling effectiveness is much lower on the pressure side than on the suction side.
  • the simpler option is the second option above where a blade is designed to have dedicated and independent cooling supplies to each microcircuit.
  • the microcircuit flow cooling characteristics are then de-sensitized from potential high rotational effects and other interferences.
  • seven internal large cavities were cored for the turbine blade of FIG. 1 , leading to six internal ribs within the airfoil portion.
  • the existence of several cold ribs is particularly significant for this application, as cold ribs provide increased creep capability for the airfoil.
  • the direct relationship of one microcircuit per supply cavity leads to potential assembly issues when microcircuit cores are tied-in with the main-body cores of the supply cavities during the pre-casting operation.
  • the cooling microcircuit is a separate and independent circuit.
  • a turbine engine component which broadly comprises an airfoil portion having a tip, a root portion, and a cooling microcircuit arrangement within the airfoil portion.
  • the cooling microcircuit arrangement comprises a multi-leg main body portion for allowing a flow of coolant to convectively cool the airfoil portion and at least one integrally formed tip cooling microcircuit for cooling the tip of the airfoil portion.
  • FIG. 1 is an illustration of a prior art turbine blade having internal blade cavities for separate imbedded airfoil wall microcircuits
  • FIG. 2 is an illustration of a prior art turbine blade having dedicated supply cavities per microcircuit to avoid cooling effectiveness debits
  • FIG. 3 is an illustration of a portion of a prior art turbine blade having three pressure side microcircuits
  • FIG. 4 is an illustration of a portion of a prior art turbine blade having two suction side microcircuits
  • FIG. 5 is an illustration of a prior art turbine blade having a trailing edge microcircuit
  • FIG. 6 is an illustration of a turbine blade having a serpentine airfoil microcircuit supplied from the blade root and integrated main mid-body with tip microcircuit as one unit.
  • a turbine engine component 10 such as a turbine blade, having an airfoil portion 12 , a platform 14 , and a root portion 16 .
  • the airfoil portion 12 has a tip 18 .
  • a cooling microcircuit 20 is imbedded within the airfoil portion 12 .
  • the imbedded cooling microcircuit 20 receives a coolant flow stream from an inlet 24 formed within the root portion 16 .
  • the inlet 24 is preferably positioned adjacent a leading edge of the root portion 16 .
  • the inlet 24 may communicate with any suitable source of cooling fluid such as engine bleed air.
  • the coolant flow stream is allowed to flow radially upward (in a direction away from the platform 14 ) through a first leg 26 of the cooling microcircuit 20 so as to take advantage of the natural pumping force.
  • the cooling microcircuit 20 may have a serpentine configuration.
  • the coolant flow stream reaches the vicinity of the tip 18 of the airfoil portion 12 , the coolant flow bends and proceeds to a second leg 28 .
  • the coolant flows radially downward (in a direction toward the platform 14 ).
  • some bypass coolant flow may be used to cool the tip 18 via tip cooling circuits 30 and 32 . As shown in FIG.
  • the tip cooling circuit 30 comprises a plurality of spaced apart flow passages 70 .
  • Each flow passage 70 has an inlet which may communicate with and receive coolant from the first leg 26 as well as from a U-shaped flow turn portion 34 connecting the legs 26 and 28 .
  • the cooling microcircuit 30 may be provided with a third leg 36 in which the coolant flows radially upward.
  • the tip circuit 32 also may comprise a plurality of spaced apart flow passages 72 .
  • Each flow passage 72 may have an inlet which communicates with the third leg 36 of the cooling microcircuit 20 so as to receive coolant therefrom.
  • Each cooling circuit passage 70 and 72 has a fluid outlet or exit 33 which allows cooling fluid to flow over a surface of the airfoil portion 12 .
  • the exits 33 are configured to allow the coolant to exit on the pressure side 35 of the airfoil portion 12 .
  • the tip cooling exits 33 from the circuits 30 and 32 may extend from a point near the leading edge 44 to a point near the trailing edge 50 of the airfoil portion 12 .
  • a root inlet refresher leg 38 may be fabricated within the root portion 16 .
  • the root inlet refresher leg 38 is in fluid communication with the third leg 36 and may be used to insure adequate cooling flow in the third leg 36 .
  • the root inlet refresher leg 38 may communicate with any suitable source (not shown) of cooling fluid such as engine bleed air.
  • exit tabs 40 forming film slots 42 may be provided in the legs 26 and/or 28 .
  • the exit tabs 40 and film slots 42 allow coolant fluid to flow from the legs 26 and/or 28 onto a surface of the airfoil portion.
  • the surface may be the pressure side surface 35 or the suction side surface 37 .
  • Fluid exiting the slots 42 helps form a cooling film over one or more of the exterior surfaces of the turbine engine component 10 .
  • Such film slots 42 may be useful in an open-cooling system.
  • the leading edge 44 of the airfoil portion 12 may be provided with a plurality of fluid outlets or exits 46 which allow a film of coolant to flow over the leading edge portions of the pressure side 35 and the suction side 37 of the airfoil portion 12 .
  • the outlets or exits 46 may be supplied with coolant from a supply cavity 48 .
  • the supply cavity 48 may communicate directly with a source (not shown) of cooling fluid, such as engine bleed air, or alternatively, the supply cavity 48 may be in fluid communication with the first leg 26 .
  • the cooling microcircuit of the present invention may also be used in a closed loop system without film cooling for industrial gas turbine applications where the external thermal load is not as high as that for aircraft engine applications.
  • the cooling microcircuit arrangement of the present invention may be formed using any suitable technique known in the art.
  • one or more sheets formed from a refractory metal material may be configured in the shape of the cooling microcircuit arrangement 20 including the inlet 24 and the root inlet refresher leg 38 , the legs 26 , 28 , and 36 , the tip cooling microcircuits 30 and 32 , the exits 33 , the tabs 40 , and the film slots 42 .
  • the refractory metal material sheets may be placed or positioned within a mold cavity.
  • the turbine engine component 10 including the airfoil portion 12 , the platform 14 , and the root portion 16 may be cast from any suitable metal known in the art such as a nickel based superalloy, a titanium based superalloy, or an iron based superalloy.
  • the refractory metal material sheets may be removed using any suitable means known in the art, leaving the cooling microcircuit arrangement 20 of the present invention.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
US11/484,143 2006-07-10 2006-07-10 Integral main body-tip microcircuits for blades Abandoned US20080008599A1 (en)

Priority Applications (3)

Application Number Priority Date Filing Date Title
US11/484,143 US20080008599A1 (en) 2006-07-10 2006-07-10 Integral main body-tip microcircuits for blades
JP2007168349A JP2008019861A (ja) 2006-07-10 2007-06-27 タービンエンジン構成要素
EP07252702A EP1878874B1 (en) 2006-07-10 2007-07-05 Integral main body-tip microcircuit for blades

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US11/484,143 US20080008599A1 (en) 2006-07-10 2006-07-10 Integral main body-tip microcircuits for blades

Publications (1)

Publication Number Publication Date
US20080008599A1 true US20080008599A1 (en) 2008-01-10

Family

ID=38461959

Family Applications (1)

Application Number Title Priority Date Filing Date
US11/484,143 Abandoned US20080008599A1 (en) 2006-07-10 2006-07-10 Integral main body-tip microcircuits for blades

Country Status (3)

Country Link
US (1) US20080008599A1 (ja)
EP (1) EP1878874B1 (ja)
JP (1) JP2008019861A (ja)

Cited By (15)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20080131285A1 (en) * 2006-11-30 2008-06-05 United Technologies Corporation RMC-defined tip blowing slots for turbine blades
US20100206512A1 (en) * 2009-02-17 2010-08-19 United Technologies Corporation Process and Refractory Metal Core For Creating Varying Thickness Microcircuits For Turbine Engine Components
US20110274559A1 (en) * 2010-05-06 2011-11-10 United Technologies Corporation Turbine Airfoil with Body Microcircuits Terminating in Platform
US8109725B2 (en) 2008-12-15 2012-02-07 United Technologies Corporation Airfoil with wrapped leading edge cooling passage
US8157527B2 (en) 2008-07-03 2012-04-17 United Technologies Corporation Airfoil with tapered radial cooling passage
US8303252B2 (en) 2008-10-16 2012-11-06 United Technologies Corporation Airfoil with cooling passage providing variable heat transfer rate
US8572844B2 (en) 2008-08-29 2013-11-05 United Technologies Corporation Airfoil with leading edge cooling passage
US9422817B2 (en) 2012-05-31 2016-08-23 United Technologies Corporation Turbine blade root with microcircuit cooling passages
US9429027B2 (en) 2012-04-05 2016-08-30 United Technologies Corporation Turbine airfoil tip shelf and squealer pocket cooling
US20170114648A1 (en) * 2015-10-27 2017-04-27 General Electric Company Turbine bucket having cooling passageway
US9885243B2 (en) 2015-10-27 2018-02-06 General Electric Company Turbine bucket having outlet path in shroud
US10005123B2 (en) * 2013-10-24 2018-06-26 United Technologies Corporation Lost core molding cores for forming cooling passages
US10370980B2 (en) * 2013-12-23 2019-08-06 United Technologies Corporation Lost core structural frame
US10508554B2 (en) 2015-10-27 2019-12-17 General Electric Company Turbine bucket having outlet path in shroud
EP3597859B1 (en) * 2018-07-13 2023-08-30 Honeywell International Inc. Turbine blade with dust tolerant cooling system

Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US6932571B2 (en) * 2003-02-05 2005-08-23 United Technologies Corporation Microcircuit cooling for a turbine blade tip
US20060153680A1 (en) * 2005-01-07 2006-07-13 Siemens Westinghouse Power Corporation Turbine blade tip cooling system
US7431562B2 (en) * 2005-12-21 2008-10-07 General Electric Company Method and apparatus for cooling gas turbine rotor blades

Family Cites Families (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US6280140B1 (en) * 1999-11-18 2001-08-28 United Technologies Corporation Method and apparatus for cooling an airfoil
FR2829174B1 (fr) * 2001-08-28 2006-01-20 Snecma Moteurs Perfectionnement apportes aux circuits de refroidissement pour aube de turbine a gaz
US7097425B2 (en) * 2003-08-08 2006-08-29 United Technologies Corporation Microcircuit cooling for a turbine airfoil

Patent Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US6932571B2 (en) * 2003-02-05 2005-08-23 United Technologies Corporation Microcircuit cooling for a turbine blade tip
US20060153680A1 (en) * 2005-01-07 2006-07-13 Siemens Westinghouse Power Corporation Turbine blade tip cooling system
US7431562B2 (en) * 2005-12-21 2008-10-07 General Electric Company Method and apparatus for cooling gas turbine rotor blades

Cited By (23)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20080131285A1 (en) * 2006-11-30 2008-06-05 United Technologies Corporation RMC-defined tip blowing slots for turbine blades
US8157527B2 (en) 2008-07-03 2012-04-17 United Technologies Corporation Airfoil with tapered radial cooling passage
US8572844B2 (en) 2008-08-29 2013-11-05 United Technologies Corporation Airfoil with leading edge cooling passage
US8303252B2 (en) 2008-10-16 2012-11-06 United Technologies Corporation Airfoil with cooling passage providing variable heat transfer rate
US8109725B2 (en) 2008-12-15 2012-02-07 United Technologies Corporation Airfoil with wrapped leading edge cooling passage
US8333233B2 (en) 2008-12-15 2012-12-18 United Technologies Corporation Airfoil with wrapped leading edge cooling passage
US9038700B2 (en) 2009-02-17 2015-05-26 United Technologies Corporation Process and refractory metal core for creating varying thickness microcircuits for turbine engine components
US20100206512A1 (en) * 2009-02-17 2010-08-19 United Technologies Corporation Process and Refractory Metal Core For Creating Varying Thickness Microcircuits For Turbine Engine Components
US8347947B2 (en) * 2009-02-17 2013-01-08 United Technologies Corporation Process and refractory metal core for creating varying thickness microcircuits for turbine engine components
US9121290B2 (en) * 2010-05-06 2015-09-01 United Technologies Corporation Turbine airfoil with body microcircuits terminating in platform
US20110274559A1 (en) * 2010-05-06 2011-11-10 United Technologies Corporation Turbine Airfoil with Body Microcircuits Terminating in Platform
US9429027B2 (en) 2012-04-05 2016-08-30 United Technologies Corporation Turbine airfoil tip shelf and squealer pocket cooling
US9422817B2 (en) 2012-05-31 2016-08-23 United Technologies Corporation Turbine blade root with microcircuit cooling passages
US10005123B2 (en) * 2013-10-24 2018-06-26 United Technologies Corporation Lost core molding cores for forming cooling passages
US10821500B2 (en) 2013-10-24 2020-11-03 Raytheon Technologies Corporation Lost core molding cores for forming cooling passages
US10370980B2 (en) * 2013-12-23 2019-08-06 United Technologies Corporation Lost core structural frame
US11085305B2 (en) 2013-12-23 2021-08-10 Raytheon Technologies Corporation Lost core structural frame
US20170114648A1 (en) * 2015-10-27 2017-04-27 General Electric Company Turbine bucket having cooling passageway
US10508554B2 (en) 2015-10-27 2019-12-17 General Electric Company Turbine bucket having outlet path in shroud
US10156145B2 (en) * 2015-10-27 2018-12-18 General Electric Company Turbine bucket having cooling passageway
US11078797B2 (en) 2015-10-27 2021-08-03 General Electric Company Turbine bucket having outlet path in shroud
US9885243B2 (en) 2015-10-27 2018-02-06 General Electric Company Turbine bucket having outlet path in shroud
EP3597859B1 (en) * 2018-07-13 2023-08-30 Honeywell International Inc. Turbine blade with dust tolerant cooling system

Also Published As

Publication number Publication date
EP1878874B1 (en) 2012-06-27
EP1878874A2 (en) 2008-01-16
JP2008019861A (ja) 2008-01-31
EP1878874A3 (en) 2011-05-25

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Legal Events

Date Code Title Description
AS Assignment

Owner name: UNITED TECHNOLOGIES CORPORATION, CONNECTICUT

Free format text: ASSIGNMENT OF ASSIGNORS INTEREST;ASSIGNORS:CUNHA, FRANCISCO J.;ABDEL-MESSEH, WILLIAM;REEL/FRAME:018051/0869

Effective date: 20060629

STCB Information on status: application discontinuation

Free format text: ABANDONED -- AFTER EXAMINER'S ANSWER OR BOARD OF APPEALS DECISION