US20050265840A1 - Cooled rotor blade with leading edge impingement cooling - Google Patents

Cooled rotor blade with leading edge impingement cooling Download PDF

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Publication number
US20050265840A1
US20050265840A1 US10/855,076 US85507604A US2005265840A1 US 20050265840 A1 US20050265840 A1 US 20050265840A1 US 85507604 A US85507604 A US 85507604A US 2005265840 A1 US2005265840 A1 US 2005265840A1
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US
United States
Prior art keywords
rib
radial passage
crossover
oblong
rotor blade
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
US10/855,076
Other languages
English (en)
Inventor
Jeffrey Levine
Edward Pietraszkiewicz
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 US10/855,076 priority Critical patent/US20050265840A1/en
Assigned to UNITED TECHNOLOGIES CORPORATION reassignment UNITED TECHNOLOGIES CORPORATION ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: LEVINE, JEFFREY R., PIETRASZKIEWICZ, EDWARD
Priority to DE602005022018T priority patent/DE602005022018D1/de
Priority to EP05253261A priority patent/EP1605138B1/fr
Priority to JP2005154978A priority patent/JP2005337257A/ja
Publication of US20050265840A1 publication Critical patent/US20050265840A1/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
    • F01D5/187—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
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/201—Heat transfer, e.g. cooling by impingement of a fluid
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/202—Heat transfer, e.g. cooling by film cooling

Definitions

  • This invention applies to gas turbine rotor blades in general, and to cooled gas turbine rotor blades in particular.
  • Turbine sections within an axial flow turbine engine include rotor assemblies that include a rotating disc and a number of rotor blades circumferentially disposed around the disk.
  • Rotor blades include an airfoil portion for positioning within the gas path through the engine. Because the temperature within the gas path very often negatively affects the durability of the airfoil, it is known to cool an airfoil by passing cooling air through the airfoil. The cooled air helps decrease the temperature of the airfoil material and thereby increase its durability.
  • Prior art cooled rotor blades very often utilize internal passage configurations that include a first radial passage extending contiguous with the leading edge, a second radial passage, and a rib disposed between and separating the passages.
  • a plurality of crossover apertures is disposed within the rib, typically oriented perpendicular to the airfoil wall along the leading edge.
  • a pressure difference across the rib causes a portion of the cooling air traveling within the second radial passage to pass through the crossover apertures and impinge on the leading edge wall.
  • Prior art leading edge impingement configurations typically employed circular crossover apertures uniformly spaced along the rib.
  • the cooling air impinging from each circular crossover aperture creates a region of relatively high heat transfer, albeit a small one.
  • the circular crossover apertures create a line of discrete regions of high heat transfer separated by larger areas of relatively low heat transfer. The variations in heat transfer make the leading edge increase the possibility of undesirable fatigue, distress, oxidation, etc. within the leading edge wall.
  • a rotor blade having a hollow airfoil and a root.
  • the hollow airfoil has a cavity defined by a suction side wall, a pressure side wall, a leading edge, a trailing edge, a base, and a tip.
  • An internal passage configuration is disposed within the cavity.
  • the configuration includes a first radial passage, a second radial passage, and a rib disposed between and separating the first radial passage and second radial passage.
  • a plurality of crossover apertures are disposed within the rib.
  • a portion of the plurality of crossover apertures are oblong having a length extending through the rib, and a height and a width. The height of each oblong aperture is greater than the width.
  • the oblong crossover apertures are aligned heightwise along the rib.
  • the root includes a conduit that is operable to permit airflow through the root and into the first passage.
  • One of the advantages of the present rotor blade and method is that airflow pressure losses within the airfoil are decreased relative to prior art airfoils having impingement cooling of which we are aware.
  • FIG. 1 is a diagrammatic perspective view of the rotor assembly section.
  • FIG. 2 is a diagrammatic sectional view of a rotor blade having an embodiment of the internal passage configuration.
  • FIG. 3 is a diagrammatic sectional view of a portion of a rotor blade having an embodiment of the internal passage configuration.
  • FIG. 4 is a diagrammatic partial view of a rib with oblong crossover apertures disposed therein.
  • a rotor blade assembly 10 for a gas turbine engine having a disk 12 and a plurality of rotor blades 14 .
  • the disk 12 includes a plurality of recesses 16 circumferentially disposed around the disk 12 and a rotational centerline 18 about which the disk 12 may rotate.
  • Each blade 14 includes a root 20 , an airfoil 22 , a platform 24 , and a radial centerline 25 .
  • the root 20 includes a geometry (e.g., a fir tree configuration) that mates with that of one of the recesses 16 within the disk 12 .
  • the root 20 further includes conduits 26 through which cooling air may enter the root 20 and pass through into the airfoil 22 .
  • the airfoil 22 includes a base 28 , a tip 30 , a leading edge 32 , a trailing edge 34 , a pressure side wall 36 (see FIG. 1 ), and a suction side wall 38 (see FIG. 1 ), and an internal passage configuration 40 .
  • FIG. 2 diagrammatically illustrates an airfoil 22 sectioned between the leading edge 32 and the trailing edge 34 .
  • the pressure side wall 36 and the suction side wall 38 extend between the base 28 and the tip 30 and meet at the leading edge 32 and the trailing edge 34 .
  • the internal passage configuration includes a first conduit 42 , a second conduit 44 , and a third conduit 46 extending through the root 20 into the airfoil 22 . Fewer or more conduits may be used alternatively.
  • the first conduit 42 is in fluid communication with a first radial passage 48 .
  • a second radial passage 50 is disposed forward of the first radial passage 48 , contiguous with the leading edge 32 , and is connected to the first radial passage 48 by a plurality of crossover apertures 52 .
  • the crossover apertures 52 are disposed in a rib 53 that extends between and separates the first radial passage 48 and the second radial passage 50 .
  • the second radial passage 50 is connected to the exterior of the airfoil 22 by a plurality of cooling apertures 54 disposed along the leading edge 32 .
  • the second radial passage 50 comprises one or more cavities.
  • the second radial passage 50 may be in direct fluid communication with the first conduit 42 .
  • the first radial passage 48 is connected to an axially extending passage 56 that extends to the trailing edge 34 of the airfoil 22 , adjacent the tip 30 of the airfoil 22 .
  • a portion of the crossover apertures 52 disposed in the rib 53 are oblong, each having a length 70 , width 72 , and height 74 .
  • substantially all of the crossover apertures 52 are oblong.
  • the length 70 of each crossover aperture 52 extends through the rib 53 .
  • the height 74 and width 72 are substantially perpendicular to each other and to the length 70 .
  • the height 74 of each oblong crossover aperture 52 is greater than the width 72 . In a preferred embodiment, the height 74 is approximately twice the width 72 in magnitude.
  • the oblong crossover apertures 52 are aligned heightwise along the rib 53 , such that the heights 74 of the oblong crossover apertures 52 are substantially collinear. In the embodiment shown in FIGS. 3 and 4 , the oblong crossover apertures 52 are shown as having a constant width 72 and circular ends. The oblong crossover apertures 52 are not limited to this embodiment.
  • the rib 53 is separated from the interior surface of the leading edge wall 78 by a distance “L”.
  • the oblong crossover apertures 52 may be described as having a hydraulic diameter “D”.
  • the separation of the rib 53 from the leading edge wall 78 , and the size of the oblong crossover apertures 53 are such that the ratio of L/D is on average in the approximate range of 2.8 to 3.0. It is our experience that an L/D in this approximate range provides desirable impingement cooling.
  • the first radial passage 48 includes a plurality of trip strips 58 attached to the interior surface of one or both of the pressure side wall 36 and the suction side wall 38 .
  • the trip strips 58 are disposed within the passage 48 at an angle ⁇ that is skewed relative to the cooling airflow direction 60 within passage 48 ; i.e., at an angle between perpendicular and parallel to the airflow direction 60 .
  • the trip strips 58 are oriented at angle of approximately 45° to the cooling airflow direction 60 .
  • the orientation of each trip strip 58 within the passage 48 is such that the trip strip 58 converges toward the rib 53 containing the crossover apertures 52 , when viewed in the airflow direction 60 .
  • Each of the trip strips 58 has an end disposed adjacent the rib 53 (i.e., a “rib end”). At least a portion of the trip strips 58 have a rib end radially located between a pair of crossover apertures 52 , preferably approximately midway between the pair of crossover apertures 52 .
  • the second conduit 44 is in fluid communication with a serpentine passage 64 disposed immediately aft of the first and second radial passages 50 , 48 , in the mid-body region of the airfoil 22 .
  • the serpentine passage 64 has an odd number of radial segments 66 , which number is greater than one; e.g., 3, 5, etc.
  • the odd number of radial segments 66 ensures that the last radial segment in the serpentine 64 ends adjacent the axially extending passage 56 .
  • Passage configurations other than the aforesaid serpentine passage 64 may be used within the mid-body region alternatively.
  • the third conduit 46 is in fluid communication with one or more passages 68 disposed between the serpentine passage 64 and the trailing edge 34 of the airfoil 22 .
  • the rotor blade airfoil 22 is disposed within the core gas path of the turbine engine.
  • the airfoil 22 is subject to high temperature core gas passing by the airfoil 22 .
  • Cooling air that is substantially lower in temperature than the core gas, is fed into the airfoil 22 through the conduits 42 , 44 , 46 disposed in the root 20 .
  • Cooling air traveling through the first conduit 42 passes directly into the first radial passage 48 , and subsequently into the axially extending passage 56 adjacent the tip 30 of the airfoil 22 .
  • a portion of the cooling air traveling within the first radial passage 48 encounters the trip strips 58 disposed within the passage 48 .
  • the trip strips 58 converging toward the rib 53 direct the portion of cooling airflow toward the rib 53 .
  • the position of the trip strips 58 relative to the crossover apertures 52 are such that the portion of cooling airflow directed toward the rib 53 is also directed toward the crossover apertures 52 .
  • the portion of cooling airflow travels through the crossover apertures 52 and into the second radial passage 50 .
  • the cooling air subsequently exits the second radial passage 50 via the cooling apertures 52 disposed in the leading edge 32 and impinges on the interior surface of the leading edge wall.
  • prior art circular crossover apertures typically create a line of discrete regions of high heat transfer separated by larger areas of relatively low heat transfer.
  • the oblong crossover apertures 52 of the present invention provide a more uniform radial heat transfer profile along the leading edge 32 that the aforesaid prior art.
  • the regions of desirable relatively high heat transfer are larger, and the regions of undesirable relatively low heat transfer are smaller.
  • the heat transfer within the regions of relatively low heat transfer appears to be increased by cooling air showering radially outward from the oblong crossover apertures 52 .

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
US10/855,076 2004-05-27 2004-05-27 Cooled rotor blade with leading edge impingement cooling Abandoned US20050265840A1 (en)

Priority Applications (4)

Application Number Priority Date Filing Date Title
US10/855,076 US20050265840A1 (en) 2004-05-27 2004-05-27 Cooled rotor blade with leading edge impingement cooling
DE602005022018T DE602005022018D1 (de) 2004-05-27 2005-05-27 Gekühlte Rotorschaufel mit Prallkühlung im Bereich der Eintrittskante
EP05253261A EP1605138B1 (fr) 2004-05-27 2005-05-27 Aube de rotor refroidie ayant un refroidissement par impact au niveau du bord d'attaque
JP2005154978A JP2005337257A (ja) 2004-05-27 2005-05-27 ロータブレード

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US10/855,076 US20050265840A1 (en) 2004-05-27 2004-05-27 Cooled rotor blade with leading edge impingement cooling

Publications (1)

Publication Number Publication Date
US20050265840A1 true US20050265840A1 (en) 2005-12-01

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Family Applications (1)

Application Number Title Priority Date Filing Date
US10/855,076 Abandoned US20050265840A1 (en) 2004-05-27 2004-05-27 Cooled rotor blade with leading edge impingement cooling

Country Status (4)

Country Link
US (1) US20050265840A1 (fr)
EP (1) EP1605138B1 (fr)
JP (1) JP2005337257A (fr)
DE (1) DE602005022018D1 (fr)

Cited By (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20060083614A1 (en) * 2004-10-18 2006-04-20 United Technologies Corporation Airfoil with large fillet and micro-circuit cooling
US20130280080A1 (en) * 2012-04-23 2013-10-24 Jeffrey R. Levine Gas turbine engine airfoil with dirt purge feature and core for making same
US20140238028A1 (en) * 2011-11-08 2014-08-28 Ihi Corporation Impingement cooling mechanism, turbine blade, and combustor
US8840370B2 (en) 2011-11-04 2014-09-23 General Electric Company Bucket assembly for turbine system
US20160024938A1 (en) * 2014-07-25 2016-01-28 United Technologies Corporation Airfoil cooling apparatus
US20160115796A1 (en) * 2013-05-20 2016-04-28 Kawasaki Jukogyo Kabushiki Kaisha Turbine blade cooling structure
US9932836B2 (en) 2012-03-22 2018-04-03 Ansaldo Energia Ip Uk Limited Turbine blade
US20180298763A1 (en) * 2014-11-11 2018-10-18 Siemens Aktiengesellschaft Turbine blade with axial tip cooling circuit
EP3597859B1 (fr) * 2018-07-13 2023-08-30 Honeywell International Inc. Aube de turbine avec système de refroidissement tolérant à la poussière

Families Citing this family (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2918105B1 (fr) * 2007-06-27 2013-12-27 Snecma Aube refroidie de turbomachine comprenant des trous de refroidissement a distance d'impact variable.
US9151173B2 (en) * 2011-12-15 2015-10-06 General Electric Company Use of multi-faceted impingement openings for increasing heat transfer characteristics on gas turbine components
KR101906701B1 (ko) * 2017-01-03 2018-10-10 두산중공업 주식회사 가스터빈 블레이드
US10895168B2 (en) 2019-05-30 2021-01-19 Solar Turbines Incorporated Turbine blade with serpentine channels

Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3619082A (en) * 1968-07-05 1971-11-09 Gen Motors Corp Turbine blade
US3767322A (en) * 1971-07-30 1973-10-23 Westinghouse Electric Corp Internal cooling for turbine vanes
US5387086A (en) * 1993-07-19 1995-02-07 General Electric Company Gas turbine blade with improved cooling
US6705836B2 (en) * 2001-08-28 2004-03-16 Snecma Moteurs Gas turbine blade cooling circuits

Family Cites Families (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5688104A (en) * 1993-11-24 1997-11-18 United Technologies Corporation Airfoil having expanded wall portions to accommodate film cooling holes
JP2000265802A (ja) * 1999-01-25 2000-09-26 General Electric Co <Ge> ガスタービン動翼冷却通路連絡路
US6174134B1 (en) * 1999-03-05 2001-01-16 General Electric Company Multiple impingement airfoil cooling
US6290463B1 (en) * 1999-09-30 2001-09-18 General Electric Company Slotted impingement cooling of airfoil leading edge
EP1213442B1 (fr) * 2000-12-05 2009-03-11 United Technologies Corporation Aube rotorique
DE10332563A1 (de) * 2003-07-11 2005-01-27 Rolls-Royce Deutschland Ltd & Co Kg Turbinenschaufel mit Prallkühlung

Patent Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3619082A (en) * 1968-07-05 1971-11-09 Gen Motors Corp Turbine blade
US3767322A (en) * 1971-07-30 1973-10-23 Westinghouse Electric Corp Internal cooling for turbine vanes
US5387086A (en) * 1993-07-19 1995-02-07 General Electric Company Gas turbine blade with improved cooling
US6705836B2 (en) * 2001-08-28 2004-03-16 Snecma Moteurs Gas turbine blade cooling circuits

Cited By (14)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20060083614A1 (en) * 2004-10-18 2006-04-20 United Technologies Corporation Airfoil with large fillet and micro-circuit cooling
US7217094B2 (en) * 2004-10-18 2007-05-15 United Technologies Corporation Airfoil with large fillet and micro-circuit cooling
US8840370B2 (en) 2011-11-04 2014-09-23 General Electric Company Bucket assembly for turbine system
US20140238028A1 (en) * 2011-11-08 2014-08-28 Ihi Corporation Impingement cooling mechanism, turbine blade, and combustor
US9932836B2 (en) 2012-03-22 2018-04-03 Ansaldo Energia Ip Uk Limited Turbine blade
US9279331B2 (en) * 2012-04-23 2016-03-08 United Technologies Corporation Gas turbine engine airfoil with dirt purge feature and core for making same
US20130280080A1 (en) * 2012-04-23 2013-10-24 Jeffrey R. Levine Gas turbine engine airfoil with dirt purge feature and core for making same
US9938837B2 (en) 2012-04-23 2018-04-10 United Technologies Corporation Gas turbine engine airfoil trailing edge passage and core for making same
US20160115796A1 (en) * 2013-05-20 2016-04-28 Kawasaki Jukogyo Kabushiki Kaisha Turbine blade cooling structure
US10018053B2 (en) * 2013-05-20 2018-07-10 Kawasaki Jukogyo Kabushiki Kaisha Turbine blade cooling structure
US20160024938A1 (en) * 2014-07-25 2016-01-28 United Technologies Corporation Airfoil cooling apparatus
US10012090B2 (en) * 2014-07-25 2018-07-03 United Technologies Corporation Airfoil cooling apparatus
US20180298763A1 (en) * 2014-11-11 2018-10-18 Siemens Aktiengesellschaft Turbine blade with axial tip cooling circuit
EP3597859B1 (fr) * 2018-07-13 2023-08-30 Honeywell International Inc. Aube de turbine avec système de refroidissement tolérant à la poussière

Also Published As

Publication number Publication date
EP1605138A3 (fr) 2007-10-03
DE602005022018D1 (de) 2010-08-12
EP1605138A2 (fr) 2005-12-14
EP1605138B1 (fr) 2010-06-30
JP2005337257A (ja) 2005-12-08

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Legal Events

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AS Assignment

Owner name: UNITED TECHNOLOGIES CORPORATION, CONNECTICUT

Free format text: ASSIGNMENT OF ASSIGNORS INTEREST;ASSIGNORS:LEVINE, JEFFREY R.;PIETRASZKIEWICZ, EDWARD;REEL/FRAME:014951/0305

Effective date: 20040630

STCB Information on status: application discontinuation

Free format text: ABANDONED -- FAILURE TO RESPOND TO AN OFFICE ACTION