US20100247328A1 - Microcircuit cooling for blades - Google Patents

Microcircuit cooling for blades Download PDF

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Publication number
US20100247328A1
US20100247328A1 US11/447,463 US44746306A US2010247328A1 US 20100247328 A1 US20100247328 A1 US 20100247328A1 US 44746306 A US44746306 A US 44746306A US 2010247328 A1 US2010247328 A1 US 2010247328A1
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US
United States
Prior art keywords
cooling
leg
turbine engine
engine component
microcircuit
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
US11/447,463
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English (en)
Inventor
Francisco J. Cunha
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.)
Raytheon Technologies 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 US11/447,463 priority Critical patent/US20100247328A1/en
Assigned to UNITED TECHNOLOGIES CORPORATION reassignment UNITED TECHNOLOGIES CORPORATION ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: CUNHA, FRANCISCO J.
Priority to JP2007149913A priority patent/JP2007327494A/ja
Priority to EP07252282.4A priority patent/EP1865151A3/fr
Publication of US20100247328A1 publication Critical patent/US20100247328A1/en
Abandoned legal-status Critical Current

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    • 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
    • F05D2250/00—Geometry
    • F05D2250/10—Two-dimensional
    • F05D2250/11—Two-dimensional triangular
    • 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
    • F05D2250/12—Two-dimensional rectangular
    • F05D2250/121—Two-dimensional rectangular square
    • 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
    • F05D2260/204—Heat transfer, e.g. cooling by the use of microcircuits
    • 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
    • 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 turbine engine component, such as a turbine blade, having a cooling microcircuit which is easy to fabricate and which has a plurality of cooling devices for effecting heat pick-up.
  • each blade internal cavity feeds a microcircuit located on a side of the airfoil, either on a pressure side or on a suction side.
  • this design is desirable to de-sensitize the cooling design from rotational effects and sink pressure interferences in microcircuit supply flows, it makes the assembly of the numerous microcircuit cores complex.
  • a turbine engine component such as a turbine blade, having a cooling microcircuit whose assembly is not complex.
  • a turbine engine component which broadly comprises at least one cooling circuit having a plurality of legs through which a cooling fluid flows, and a plurality of cooling devices in at least one of the legs.
  • Each of the cooling devices has a heat transfer multiplier in the range of from 1.8 to 2.4 and a reattachment length in the range of from 1.9 to 2.5.
  • a cooling microcircuit for use in a turbine engine component which broadly comprises a first leg for receiving a cooling fluid, a second leg for receiving the cooling fluid from the first leg, and a third leg for receiving the cooling fluid from the second leg. At least one of the first and second legs contains a plurality of cooling devices. Each of the cooling devices has a heat transfer multiplier in the range of from'1.8 to 2.4 and a reattachment length in the range of from 1.9 to 2.5.
  • FIG. 1 illustrates the two dimensional span of a prior art blade
  • FIG. 2 illustrates the heat transfer characteristics of a cylinder shaped cooling device
  • FIG. 3 illustrates the heat transfer characteristics of a cube shaped cooling device
  • FIG. 4 illustrates the heat transfer characteristics of a diamond shaped cooling device
  • FIG. 5 is a schematic representation of a turbine blade having a cooling microcircuit in accordance with the present invention.
  • cooling microcircuits have banks of pedestals as cooling devices to enhance heat pick-up. While employing these cooling devices in the cooling microcircuits, it is also desirable to minimize their number for the same heat pick-up capability.
  • FIGS. 2-4 illustrate the heat transfer characteristics of different shaped pedestals.
  • FIG. 2 illustrates the heat transfer characteristics for a cylinder shaped pedestal.
  • FIG. 3 illustrates the heat transfer characteristics for a cube shaped pedestal.
  • FIG. 4 illustrates the heat transfer characteristics for a diamond-shaped pedestal configuration.
  • the spanwise domain of influence for a cylinder type pedestal is on the lowest of all three configurations, with the diamond-shaped pedestal being the greatest.
  • the spanwise domain of influence of the diamond-shaped pedestal is about 32% greater than a cylinder shaped pedestal.
  • the preferred inter-element spacing for the diamond shaped pedestal would be two-fold greater than that for the cylinder type of pedestal with even higher heat transfer enhancement. It can be concluded that for a given surface size, the number of elements needed to achieve effective heat transfer enhancement can be minimized using a diamond geometry.
  • the value of the heat transfer multiplier recovers downstream to reach a maximum of about 2.4 heat transfer enhancement (reference being the flat plate heat transfer) at an x/d of 2.5. Not only is this the farthest location of reattachment induced enhancement downstream to the obstacle, but also it has the highest value of maximum heat transfer multiplier.
  • the key factor responsible for this effect is the special flow characteristics related to diamond shaped pedestals. It is dominated by highly turbulent delta-wing vortices as opposed to the commonly observed, recirculating bubble. These vortices substantially elevate the surface heat transfer underneath their tracks. It is expected that such influence persists further downstream as the shear layer reattached to the endwall.
  • a cooling device having a reattachment length in the range of 1.9 to 2.5 and a heat transfer multiplier relative to flat plate heat transfer in the range of from 1.8 to 2.2, preferably from 2.2 to 2.4.
  • the diamond shaped pedestal has the strongest reattachment-induced enhancement with the widest spread in the wake region. In addition, its reattachment length is also the longest.
  • the turbine engine component 10 such as a turbine blade, has 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 .
  • each of the legs 26 and 28 has a plurality of cooling devices 80 .
  • the cooling devices 80 may have any desired shape. While it is preferred that the cooling devices be diamond shaped, they may also be cylindrical or cubed shaped. If a diamond shaped cooling device is used, it is preferred that the tip 86 of the diamond shape be aligned with the direction of the cooling fluid flowing through the respective one of the legs 24 and 26 . The angle between the surfaces forming the tip 86 is important and should preferably be in the range of from 30 to 60 degrees.
  • each cooling device 80 could have a cube shape.
  • one of the sides of the cube should be oriented substantially normal to the direction of flow of the cooling fluid in the leg in which the cooling device 80 is located.
  • a plurality of cooling devices 80 may be positioned within each of the legs 24 and 26 .
  • the cooling devices 80 in each leg are arranged in a staggered configuration.
  • the cooling microcircuit 20 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)
US11/447,463 2006-06-06 2006-06-06 Microcircuit cooling for blades Abandoned US20100247328A1 (en)

Priority Applications (3)

Application Number Priority Date Filing Date Title
US11/447,463 US20100247328A1 (en) 2006-06-06 2006-06-06 Microcircuit cooling for blades
JP2007149913A JP2007327494A (ja) 2006-06-06 2007-06-06 タービンエンジン部品および冷却用微細回路
EP07252282.4A EP1865151A3 (fr) 2006-06-06 2007-06-06 Refroidissement de microcircuit pour pales

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US11/447,463 US20100247328A1 (en) 2006-06-06 2006-06-06 Microcircuit cooling for blades

Publications (1)

Publication Number Publication Date
US20100247328A1 true US20100247328A1 (en) 2010-09-30

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US11/447,463 Abandoned US20100247328A1 (en) 2006-06-06 2006-06-06 Microcircuit cooling for blades

Country Status (3)

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US (1) US20100247328A1 (fr)
EP (1) EP1865151A3 (fr)
JP (1) JP2007327494A (fr)

Cited By (12)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20120034102A1 (en) * 2010-08-09 2012-02-09 General Electric Company Bucket assembly cooling apparatus and method for forming the bucket assembly
US9850762B2 (en) 2013-03-13 2017-12-26 General Electric Company Dust mitigation for turbine blade tip turns
US9957816B2 (en) 2014-05-29 2018-05-01 General Electric Company Angled impingement insert
US9995148B2 (en) 2012-10-04 2018-06-12 General Electric Company Method and apparatus for cooling gas turbine and rotor blades
US20180371941A1 (en) * 2017-06-22 2018-12-27 United Technologies Corporation Gaspath component including minicore plenums
US10233775B2 (en) 2014-10-31 2019-03-19 General Electric Company Engine component for a gas turbine engine
US10280785B2 (en) 2014-10-31 2019-05-07 General Electric Company Shroud assembly for a turbine engine
US10364684B2 (en) 2014-05-29 2019-07-30 General Electric Company Fastback vorticor pin
US10422235B2 (en) 2014-05-29 2019-09-24 General Electric Company Angled impingement inserts with cooling features
US10563514B2 (en) 2014-05-29 2020-02-18 General Electric Company Fastback turbulator
US10690055B2 (en) 2014-05-29 2020-06-23 General Electric Company Engine components with impingement cooling features
US11333023B2 (en) 2018-11-09 2022-05-17 Raytheon Technologies Corporation Article having cooling passage network with inter-row sub-passages

Families Citing this family (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
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
US11180998B2 (en) 2018-11-09 2021-11-23 Raytheon Technologies Corporation Airfoil with skincore passage resupply

Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4407632A (en) * 1981-06-26 1983-10-04 United Technologies Corporation Airfoil pedestaled trailing edge region cooling configuration
US4587700A (en) * 1984-06-08 1986-05-13 The Garrett Corporation Method for manufacturing a dual alloy cooled turbine wheel
US20040151587A1 (en) * 2003-02-05 2004-08-05 Cunha Frank J. Microcircuit cooling for a turbine blade tip
US6832889B1 (en) * 2003-07-09 2004-12-21 General Electric Company Integrated bridge turbine blade
US20050169752A1 (en) * 2003-10-24 2005-08-04 Ching-Pang Lee Converging pin cooled airfoil
US7207775B2 (en) * 2004-06-03 2007-04-24 General Electric Company Turbine bucket with optimized cooling circuit
US20070128034A1 (en) * 2005-12-05 2007-06-07 General Electric Company Zigzag cooled turbine airfoil

Family Cites Families (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JP3377563B2 (ja) * 1993-09-08 2003-02-17 三菱重工業株式会社 ガスタービンの空気冷却動翼
US7478994B2 (en) * 2004-11-23 2009-01-20 United Technologies Corporation Airfoil with supplemental cooling channel adjacent leading edge
US7438527B2 (en) * 2005-04-22 2008-10-21 United Technologies Corporation Airfoil trailing edge cooling

Patent Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4407632A (en) * 1981-06-26 1983-10-04 United Technologies Corporation Airfoil pedestaled trailing edge region cooling configuration
US4587700A (en) * 1984-06-08 1986-05-13 The Garrett Corporation Method for manufacturing a dual alloy cooled turbine wheel
US20040151587A1 (en) * 2003-02-05 2004-08-05 Cunha Frank J. Microcircuit cooling for a turbine blade tip
US6832889B1 (en) * 2003-07-09 2004-12-21 General Electric Company Integrated bridge turbine blade
US20050169752A1 (en) * 2003-10-24 2005-08-04 Ching-Pang Lee Converging pin cooled airfoil
US7207775B2 (en) * 2004-06-03 2007-04-24 General Electric Company Turbine bucket with optimized cooling circuit
US20070128034A1 (en) * 2005-12-05 2007-06-07 General Electric Company Zigzag cooled turbine airfoil

Cited By (14)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US8647064B2 (en) * 2010-08-09 2014-02-11 General Electric Company Bucket assembly cooling apparatus and method for forming the bucket assembly
US20120034102A1 (en) * 2010-08-09 2012-02-09 General Electric Company Bucket assembly cooling apparatus and method for forming the bucket assembly
US9995148B2 (en) 2012-10-04 2018-06-12 General Electric Company Method and apparatus for cooling gas turbine and rotor blades
US9850762B2 (en) 2013-03-13 2017-12-26 General Electric Company Dust mitigation for turbine blade tip turns
US10422235B2 (en) 2014-05-29 2019-09-24 General Electric Company Angled impingement inserts with cooling features
US10364684B2 (en) 2014-05-29 2019-07-30 General Electric Company Fastback vorticor pin
US9957816B2 (en) 2014-05-29 2018-05-01 General Electric Company Angled impingement insert
US10563514B2 (en) 2014-05-29 2020-02-18 General Electric Company Fastback turbulator
US10690055B2 (en) 2014-05-29 2020-06-23 General Electric Company Engine components with impingement cooling features
US10233775B2 (en) 2014-10-31 2019-03-19 General Electric Company Engine component for a gas turbine engine
US10280785B2 (en) 2014-10-31 2019-05-07 General Electric Company Shroud assembly for a turbine engine
US20180371941A1 (en) * 2017-06-22 2018-12-27 United Technologies Corporation Gaspath component including minicore plenums
US10808571B2 (en) * 2017-06-22 2020-10-20 Raytheon Technologies Corporation Gaspath component including minicore plenums
US11333023B2 (en) 2018-11-09 2022-05-17 Raytheon Technologies Corporation Article having cooling passage network with inter-row sub-passages

Also Published As

Publication number Publication date
EP1865151A3 (fr) 2014-06-25
JP2007327494A (ja) 2007-12-20
EP1865151A2 (fr) 2007-12-12

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AS Assignment

Owner name: UNITED TECHNOLOGIES CORPORATION, CONNECTICUT

Free format text: ASSIGNMENT OF ASSIGNORS INTEREST;ASSIGNOR:CUNHA, FRANCISCO J.;REEL/FRAME:017962/0422

Effective date: 20060605

STCB Information on status: application discontinuation

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