EP1870561B1 - Refroidissement du bord d'attaque d'un composant de turbine à gaz par générateurs de turbulence - Google Patents

Refroidissement du bord d'attaque d'un composant de turbine à gaz par générateurs de turbulence Download PDF

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Publication number
EP1870561B1
EP1870561B1 EP07252545.4A EP07252545A EP1870561B1 EP 1870561 B1 EP1870561 B1 EP 1870561B1 EP 07252545 A EP07252545 A EP 07252545A EP 1870561 B1 EP1870561 B1 EP 1870561B1
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EP
European Patent Office
Prior art keywords
leading edge
trip strips
trip
turbine engine
engine component
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.)
Revoked
Application number
EP07252545.4A
Other languages
German (de)
English (en)
Other versions
EP1870561A2 (fr
EP1870561A3 (fr
Inventor
Jeffrey R. Levine
Eleanor Kaufman
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
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Filing date
Publication date
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Application filed by United Technologies Corp filed Critical United Technologies Corp
Publication of EP1870561A2 publication Critical patent/EP1870561A2/fr
Publication of EP1870561A3 publication Critical patent/EP1870561A3/fr
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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
    • F05D2240/00—Components
    • F05D2240/10—Stators
    • F05D2240/12—Fluid guiding means, e.g. vanes
    • 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
    • F05D2240/00—Components
    • F05D2240/10—Stators
    • F05D2240/12—Fluid guiding means, e.g. vanes
    • F05D2240/121—Fluid guiding means, e.g. vanes related to the leading edge of a stator vane
    • 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
    • F05D2240/00—Components
    • F05D2240/20—Rotors
    • F05D2240/30—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor
    • F05D2240/303—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor related to the leading edge of a rotor blade
    • 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 enhanced cooling of the leading edge of airfoil portions of turbine engine components using trip strips that are preferably staggered and that are wrapped around the nose of the leading edge cavity.
  • FIG. 1 where there is shown an airfoil portion 10 of a turbine engine component 12. As can be seen from the figure, a radial flow leading edge cavity 14 is used to effect cooling of the leading edge region.
  • FIG. 1 Another exemplary prior art cooling arrangement is shown in US 6,068,445 .
  • the present invention provides a turbine engine component as set forth in claim 1.
  • FIG. 2 illustrates the leading edge 30 of an airfoil portion 32 of a turbine engine component.
  • the leading edge 30 has a leading edge cavity 34 in which a cooling fluid, such as engine bleed air, flows in a radial direction.
  • the leading edge 30 also has a nose portion 36 and an external stagnation region 38.
  • trip strips are desirable to provide adequate cooling of the leading edge, especially at the nose portion 36 of the airfoil portion 32 adjacent to the external stagnation region 38.
  • the trip strip arrangement which will be discussed hereinafter provides high heat transfer to the leading edge 30 of the airfoil portion 32.
  • a plurality of trip strips 40 are positioned on the pressure side 42 of the airfoil portion 32, while, as shown in FIGS. 2 , 3, and 6 , a plurality of trip strips 44 are placed on the suction side 46 of the airfoil portion 32.
  • the trip strips 40 on the pressure side 42 are wrapped around the leading edge nose portion 36.
  • the curvature of the leading edge nose portion 36 causes the trip strips 40 to be oriented more or less normal to the direction of flow 48 (see FIG. 6 ).
  • the flow is tripped and generates a large vortex 49 at the leading edge (see FIG. 7 ). This large vortex generates very high heat transfer coefficients at the leading edge nose 36.
  • the trip strips 40 and the trip strips 44 are preferably staggered approximately one half pitch apart between the suction side 46 and the pressure side 42 of the airfoil portion 32. As shown in FIGS. 2 and 7 , there is also a gap 47 between adjacent ones of the trip strips 40 and the trip strips 44. Each gap 47 is located along a parting line 70 of the airfoil portion 32.
  • the orientation of the trip strips 40 and 44 in the cavity 34 also increases heat transfer at the leading edge 30 of the airfoil portion 32.
  • the trip strips 40 and 44 may be oriented at an angle ⁇ of approximately 45 degrees relative to the flow direction 48.
  • the leading edges 54 and 56 of the trip strips 40 and 44 are positioned in the region of highest heat load, in this case the leading edge nose 36.
  • This trip strip orientation permits the creation of a turbulent vortex 49 in the cavity 34.
  • the cooling fluid initially hits the leading edges 54 and 56 of the trip strip and separates from the airfoil surface. The flow then re-attaches downstream of the trip strip leading edges 54 and 56 and moves toward the divider rib 60 between the leading edge cavity 34 and the adjacent cavity 62.
  • trip strip configuration allows for cooling flow to impinge on the leading edge nose 36, further enhancing heat transfer.
  • the leading edge of the trip strips 40 and 44 is located near the nose 36 of the leading edge cavity 34.
  • trip strips 40 although skewed at an angle ⁇ with respect to the direction of flow 48 along the pressure-side wall 42, become normal to the direction of flow 48 as they wrap around the nose 36 of the leading edge cavity 34, increasing the turbulent vortex 49 generated by the trip strips 40 and 44, and subsequently increasing the heat transfer coefficient.
  • the staggered and 45 degree angled trip strips generate a vortex that impinges flow onto the nose 36 of the leading edge cavity.
  • the trip strip configuration of the present invention maintains a P/E ratio between 3 and 25 where P is the radial pitch in between trip strips and E is trip strip height. Further, the trip strip configuration described herein maintains an E/H ratio of between 0.15 and 1.50 where E is trip strip height and H is the height of the cavity 34.
  • Airflow testing has shown that the heat transfer coefficients at the leading edge of the airfoil adjacent to the external stagnation region when using the staggered trip strips of the present invention are enhanced by approximately two times, greatly increasing airfoil oxidation and thermo-mechanical fatigue cracking life.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Claims (8)

  1. Composant de moteur à turbine comprenant :
    une portion de profil aérodynamique (32) ayant un bord d'attaque (30), un côté aspiration (46), un côté pression (42) et un bord de fuite ;
    une cavité de bord d'attaque à écoulement radial (34) à travers laquelle s'écoule un fluide de refroidissement pour refroidir ledit bord d'attaque et une nervure de diviseur (60) entre ladite cavité de bord d'attaque (34) et une cavité adjacente (62) ; et
    un moyen pour générer un vortex (49) dans ladite cavité de bord d'attaque (34) qui impacte une portion de nez (36) de ladite cavité de bord d'attaque (34) ;
    dans lequel ledit moyen de génération de vortex comprend une pluralité de premières bandes de déclenchement (40) enroulées autour de ladite portion de nez (36) de ladite cavité de bord d'attaque (34) et une pluralité de secondes bandes de déclenchement (44), dans lequel lesdites premières bandes de déclenchement (40) sont montées sur le côté pression (42) de ladite portion de profil aérodynamique (32) et lesdites secondes bandes de déclenchement (44) sont montées sur ledit côté aspiration (46) de ladite portion de profil aérodynamique (32), et dans lequel lesdites premières bandes de déclenchement (40) sont obliques à un angle α par rapport à la direction d'écoulement (48) dudit fluide de refroidissement le long du côté pression (42) de la portion de profil aérodynamique (32) et deviennent normales à la direction d'écoulement (48) lorsqu'elles sont enroulées autour de ladite portion de nez (36) de ladite cavité de bord d'attaque (34) ; dans lequel il y a une pluralité d'écartements (47) entre des bandes adjacentes desdites premières bandes de déclenchement (40 ; 44) et desdites secondes bandes de déclenchement (44 ; 40) ;
    caractérisé en ce que ladite pluralité d'écartements (47) est située le long d'une ligne de jonction (70) de ladite portion de profil aérodynamique (32).
  2. Composant de moteur à turbine selon la revendication 1, dans lequel lesdites premières bandes de déclenchement (40 ; 44) et lesdites secondes bandes de déclenchement (44 ; 40) sont décalées.
  3. Composant de moteur à turbine selon la revendication 1 ou 2, dans lequel lesdites premières bandes de déclenchement (40 ; 44) et secondes bandes de déclenchement (44 ; 40) sont positionnées le long d'une direction d'écoulement (48) dudit fluide de refroidissement.
  4. Composant de moteur à turbine selon la revendication 3, dans lequel lesdites premières et secondes bandes de déclenchement (40 ; 44) sont chacune orientées à un angle de 45 degrés par rapport à la direction d'écoulement dudit fluide de refroidissement.
  5. Composant de moteur à turbine selon la revendication 3, dans lequel chacune desdites premières et secondes bandes de déclenchement (40 ; 44) a un bord d'attaque et ledit bord d'attaque de chacune desdites bandes de déclenchement est positionné dans une région de charge calorifique la plus élevée.
  6. Composant de moteur à turbine selon l'une quelconque des revendications 3 à 5, dans lequel lesdites bandes de déclenchement (40 ; 44) ont un rapport P/E dans la plage de 3 à 25 où P est un pas radial entre des bandes de déclenchement et E est une hauteur de bande de déclenchement.
  7. Composant de moteur à turbine selon l'une quelconque des revendications 3 à 6, dans lequel lesdites bandes de déclenchement (40 ; 44) ont un rapport E/H entre 0,15 et 1,50 où E est une hauteur de bande de déclenchement et H est une hauteur de la cavité.
  8. Composant de moteur à turbine selon l'une quelconque des revendications 3 à 6, dans lequel lesdites premières bandes de déclenchement (40 ; 44) et lesdites secondes bandes de déclenchement (44 ; 40) sont décalées approximativement d'un demi-pas.
EP07252545.4A 2006-06-22 2007-06-22 Refroidissement du bord d'attaque d'un composant de turbine à gaz par générateurs de turbulence Revoked EP1870561B1 (fr)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US11/473,893 US20070297916A1 (en) 2006-06-22 2006-06-22 Leading edge cooling using wrapped staggered-chevron trip strips

Publications (3)

Publication Number Publication Date
EP1870561A2 EP1870561A2 (fr) 2007-12-26
EP1870561A3 EP1870561A3 (fr) 2010-12-22
EP1870561B1 true EP1870561B1 (fr) 2017-04-05

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EP07252545.4A Revoked EP1870561B1 (fr) 2006-06-22 2007-06-22 Refroidissement du bord d'attaque d'un composant de turbine à gaz par générateurs de turbulence

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US (1) US20070297916A1 (fr)
EP (1) EP1870561B1 (fr)
JP (1) JP2008002464A (fr)

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US8690538B2 (en) * 2006-06-22 2014-04-08 United Technologies Corporation Leading edge cooling using chevron trip strips
EP1921269A1 (fr) * 2006-11-09 2008-05-14 Siemens Aktiengesellschaft Aube de turbine
US8376706B2 (en) * 2007-09-28 2013-02-19 General Electric Company Turbine airfoil concave cooling passage using dual-swirl flow mechanism and method
US8348613B2 (en) 2009-03-30 2013-01-08 United Technologies Corporation Airflow influencing airfoil feature array
US8821111B2 (en) * 2010-12-14 2014-09-02 Siemens Energy, Inc. Gas turbine vane with cooling channel end turn structure
US8757961B1 (en) * 2011-05-21 2014-06-24 Florida Turbine Technologies, Inc. Industrial turbine stator vane
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
EP2971544B1 (fr) 2013-03-14 2019-08-21 United Technologies Corporation Refroidissement d'un composant d'un moteur à turbine à gaz avec des bandes de déclenchement en regard imbriquées
WO2014159800A1 (fr) * 2013-03-14 2014-10-02 United Technologies Corporation Barrette perturbatrice en forme de chevron à angle obtus
WO2015026430A1 (fr) 2013-08-20 2015-02-26 United Technologies Corporation Plaque de revêtement de plateforme de canalisation
US10247099B2 (en) * 2013-10-29 2019-04-02 United Technologies Corporation Pedestals with heat transfer augmenter
EP3084182B8 (fr) 2013-12-20 2021-04-07 Raytheon Technologies Corporation Cavite de refroidissement de composants de turbine a gaz avec elements favorisant la generation de tourbillons
WO2015184294A1 (fr) 2014-05-29 2015-12-03 General Electric Company Générateur de turbulence fastback
US10422235B2 (en) 2014-05-29 2019-09-24 General Electric Company Angled impingement inserts with cooling features
US9957816B2 (en) 2014-05-29 2018-05-01 General Electric Company Angled impingement insert
CA2949539A1 (fr) 2014-05-29 2016-02-18 General Electric Company Elements de turbine a gaz ayant des caracteristiques de refroidissement
US10364684B2 (en) 2014-05-29 2019-07-30 General Electric Company Fastback vorticor pin
US10119404B2 (en) * 2014-10-15 2018-11-06 Honeywell International Inc. Gas turbine engines with improved leading edge airfoil cooling
US10280785B2 (en) 2014-10-31 2019-05-07 General Electric Company Shroud assembly for a turbine engine
US10233775B2 (en) 2014-10-31 2019-03-19 General Electric Company Engine component for a gas turbine engine
US10406596B2 (en) 2015-05-01 2019-09-10 United Technologies Corporation Core arrangement for turbine engine component
US10422233B2 (en) * 2015-12-07 2019-09-24 United Technologies Corporation Baffle insert for a gas turbine engine component and component with baffle insert
US10280841B2 (en) 2015-12-07 2019-05-07 United Technologies Corporation Baffle insert for a gas turbine engine component and method of cooling
US10337334B2 (en) 2015-12-07 2019-07-02 United Technologies Corporation Gas turbine engine component with a baffle insert
US10577947B2 (en) 2015-12-07 2020-03-03 United Technologies Corporation Baffle insert for a gas turbine engine component
US10352177B2 (en) 2016-02-16 2019-07-16 General Electric Company Airfoil having impingement openings
US10577944B2 (en) 2017-08-03 2020-03-03 General Electric Company Engine component with hollow turbulators
US10590778B2 (en) 2017-08-03 2020-03-17 General Electric Company Engine component with non-uniform chevron pins
US11788416B2 (en) 2019-01-30 2023-10-17 Rtx Corporation Gas turbine engine components having interlaced trip strip arrays
CN115182787A (zh) * 2022-04-27 2022-10-14 上海交通大学 改善前缘旋流冷却能力的涡轮叶片及发动机

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US5232343A (en) 1984-05-24 1993-08-03 General Electric Company Turbine blade
US5052889A (en) 1990-05-17 1991-10-01 Pratt & Whintey Canada Offset ribs for heat transfer surface
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EP1473439A2 (fr) 2003-04-29 2004-11-03 General Electric Company Aube crénelée refroidie pour turbine
EP1503038A1 (fr) 2003-08-01 2005-02-02 Snecma Moteurs Circuit de refroidissement pour aube de turbine

Also Published As

Publication number Publication date
US20070297916A1 (en) 2007-12-27
EP1870561A2 (fr) 2007-12-26
JP2008002464A (ja) 2008-01-10
EP1870561A3 (fr) 2010-12-22

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