EP1746254B1 - Dispositif et méthode de refroidissement d'une virole de turbine et de l'anneau externe d'une aube statorique de turbine - Google Patents

Dispositif et méthode de refroidissement d'une virole de turbine et de l'anneau externe d'une aube statorique de turbine Download PDF

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Publication number
EP1746254B1
EP1746254B1 EP06253774.1A EP06253774A EP1746254B1 EP 1746254 B1 EP1746254 B1 EP 1746254B1 EP 06253774 A EP06253774 A EP 06253774A EP 1746254 B1 EP1746254 B1 EP 1746254B1
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EP
European Patent Office
Prior art keywords
shroud
turbine
cooling air
platform
cooling
Prior art date
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Application number
EP06253774.1A
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German (de)
English (en)
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EP1746254A3 (fr
EP1746254A2 (fr
Inventor
Eric Durocher
Assaf Farah
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Pratt and Whitney Canada Corp
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Pratt and Whitney Canada Corp
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Filing date
Publication date
Priority claimed from US11/183,741 external-priority patent/US7520715B2/en
Priority claimed from US11/183,922 external-priority patent/US7374395B2/en
Priority claimed from US11/184,843 external-priority patent/US20070020088A1/en
Application filed by Pratt and Whitney Canada Corp filed Critical Pratt and Whitney Canada Corp
Publication of EP1746254A2 publication Critical patent/EP1746254A2/fr
Publication of EP1746254A3 publication Critical patent/EP1746254A3/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
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/08—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator
    • 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
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/08—Cooling; Heating; Heat-insulation
    • F01D25/12—Cooling
    • 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
    • F01D9/00—Stators
    • F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
    • 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
    • F05D2240/11—Shroud seal segments
    • 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
    • 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
    • F05D2260/205—Cooling fluid recirculation, i.e. after cooling one or more components is the cooling fluid recovered and used elsewhere for other purposes

Definitions

  • the invention relates generally to gas turbine engines and more particularly to turbine shroud and downstream vane outer shroud cooling.
  • a gas turbine engine usually includes a hot section, i.e., a turbine section which includes at least one rotor stage, for example, having a plurality of shroud segments disposed circumferentially one adjacent to another to form a shroud ring surrounding a turbine rotor, and at least one stator vane stage disposed immediately downstream and/or upstream of the rotor stage, formed with outer and inner shrouds and a plurality of radial stator vanes extending therebetween.
  • the rotor stage and the stator vane stage need to be cooled. Since flowing coolant through the rotor and stator vane stages diminishes overall engine performance, it is typically desirable to minimize cooling flow consumption without degrading durability of components of the turbine section.
  • a turbine shroud segment having the features of the preamble of claim 1 is disclosed in US 6899518 B2 .
  • the present invention therefore provides a turbine shroud segment as set forth in claim 1.
  • the present invention also provides gas turbine engine structure as set forth in claim 6.
  • the present invention also provides a method of reusing turbine shroud cooling air for impingement cooling on a downstream turbine vane outer shroud as set forth in claim 10.
  • a turbofan gas turbine engine incorporates an embodiment of the present invention, presented as an example of the application of the present invention, and includes a housing or a nacelle 10, a core casing 13, a low pressure spool assembly seen generally at 12 which includes a fan 14, low pressure compressor 16 and low pressure turbine 18, and a high pressure spool assembly seen generally at 20 which includes a high pressure compressor 22 and a high pressure turbine 24. There is provided a burner 25 for generating combustion gases.
  • the low pressure turbine 18 and high pressure turbine 24 include a plurality of rotor stages 28 and stator vane stages 30.
  • each of the rotor stages 28 has a plurality of rotor blades 33 encircled by a turbine shroud assembly 32 and each of the stator vane stages 30 includes a stator vane assembly 34 which is positioned upstream and/or downstream of a rotor stage 31, for directing combustion gases into or out of an annular gas path 36 within a corresponding turbine shroud assembly 32, and through the corresponding rotor stage 31.
  • the stator vane assembly 34 for example a first stage of a low pressure turbine (LPT) vane assembly, is disposed, for example, downstream of the shroud assembly 32 of one rotor stage 28, and includes, for example a plurality of stator vane segments (not indicated) joined one to another in a circumferential direction to form a turbine vane outer shroud 38 which comprises a plurality of axial stator vanes 40 (only a portion of one is shown) which divide a downstream section of the annular gas path 36 relative to the rotor stage 28, into sectoral gas passages for directing combustion gas flow out of the rotor stage 28.
  • LPT low pressure turbine
  • the shroud assembly 32 in the rotor stage 28 includes a plurality of shroud segments 42 (only one shown) each of which includes a platform 44 having front and rear radial legs 46, 48 with respective hooks (not indicated).
  • the shroud segments 42 are joined one to another in a circumferential direction and thereby form the shroud assembly 32.
  • each shroud segment 42 has outer and inner surfaces 50, 52 and is defined axially between leading and trailing ends 54, 56, and circumferentially between opposite sides 58, 60 thereof.
  • the platforms 44 of the segments collectively form a turbine shroud ring (not indicated) which encircles the rotor blades 33 and in combination with the rotor stage 28, defines a section of the annular gas path 36.
  • the turbine shroud ring is disposed immediately upstream of and abuts the turbine vane outer shroud 38, to thereby form a portion of an outer wall (not indicated) of the annular gas path 36.
  • the front and rear radial legs 46, 48 are axially spaced apart and integrally extend from the outer surface 50 radially and outwardly such that the hooks of the front a rear radial legs 46, 48 are conventionally connected with an annular shroud support structure 62 which is formed with a plurality of shroud support segments (not indicated) and is in turn supported within the core casing 13.
  • An annular cavity 64 is thus defined axially between the front and rear legs 46, 48 and radially between the platforms 44 of the shroud segments 42 and the annular shroud support structure 62.
  • the annular middle cavity is in fluid communication with a cooling air source, for example bleed air from the low or high pressure compressors 16, 22 and thus the cooling air under pressure is introduced into and accommodated within the annular cavity 64.
  • each shroud segment 42 preferably includes a passage, for example a plurality of holes 66 extending axially within the platform 44 for directing cooling air therethrough for transpiration cooling of the platform 44.
  • a groove 68 extending in a circumferential direction with opposite ends closed is provided, for example, on the outer surface 50 of the platform 44 such that holes 66 can be drilled from the trailing end 56 of the platform straightly and axially towards and terminate at the groove 68.
  • the groove 68 forms a common inlet of the holes 66 for intake of cooling air accommodated within the cavity 64.
  • other types of outlets can be made to achieve the convenience of the hole drilling process.
  • outlets of the holes 66 in order to adequately discharge the cooling air from the holes 66 and reduce the contact surface of the trailing end 56 of the platform 44 of the shroud segments 42 with respect to the turbine vane outer shroud 38.
  • an elongate recess 70 is provided in the trailing end 56 of the platform 44 with an opening on the inner surface 52 of the platform 44, thereby forming a common outlet of the holes 66 to discharge the cooling air, for example to the gas path 36.
  • Other types of outlets can be used for adequately discharging the cooling air from the holes 66.
  • the groove 68 is in fluid communication with the middle cavity 64 and thus cooling air introduced into the cavity 64 is directed into and through the axial holes 66 for effectively cooling the platform 44 of the shroud segments 42.
  • the cooling air is then discharged through the elongate recess 70 at the trailing end 56 of the platform 42, impinging on a downstream engine part such as the turbine vane outer shroud 38, before entering the gas path 36.
  • the groove 68 which functions as the common inlet of the holes 66 is preferably located close to the front leg 46 such that the holes 66 extend through a major section of the entire axial length of the platform 44 of the shroud segment 42, thereby efficiently cooling the platform 44 of the shroud segment 42.
  • the holes 66 are preferably substantially evenly spaced apart in a circumferential direction and are preferably aligned with the turbine vane outer shroud. Thus, the cooling air impinges on the leading end of the turbine vane outer shroud 38.
  • the number of holes 66 in each shroud segment 42 is determined such that the cooling air discharged from the holes 66 effectively cools the entire circumference of the leading end of the turbine vane outer shroud 38.
  • cooling air introduced into the cavity 64 is directed within and through the platforms 44 of the shroud segments 42 via the holes 66, in order to cool the turbine shroud ring.
  • cooling air flowing through the substantially axial straight holes 66 forms a plurality of substantially straight axial cooling air streams directed towards the leading end of the turbine vane outer shroud 38, preferably with a high velocity thereof, for impingement cooling on the turbine vane outer shroud 38.
  • the substantially axial straight holes 66 direct the cooling air through the entire length thereof, thereby forming substantially straight cooling air streams with relatively high directionality.
  • the substantially straight cooling air streams are more individually focused and interfere less with adjacent air streams when approaching the leading end of the turbine vane outer shroud 38, which results in greater impingement effects on the leading end of the turbine vane outer shroud 38.

Landscapes

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

Claims (14)

  1. Segment d'enveloppe de turbine (42) pour un agencement de refroidissement dans une section de turbine d'un moteur à turbine à gaz pour un refroidissement dans une relation d'écoulement en série, à la fois le segment d'enveloppe de turbine (42) et une enveloppe externe d'aube de turbine (38) étant disposée en aval du segment d'enveloppe de turbine (42), le segment d'enveloppe de turbine comprenant une pluralité de passages (66) s'étendant dans une plate-forme d'enveloppe (44) pour diriger l'air de refroidissement au travers pour refroidir le segment d'enveloppe de turbine (42) et pour évacuer l'air de refroidissement au niveau d'une extrémité de fuite (56) de la plate-forme d'enveloppe (44) pour impact sur l'enveloppe externe d'aube de turbine (38), la plate-forme d'enveloppe (44) comprenant un refoulement commun (70) pour ladite pluralité de passages (66) sur l'extrémité de fuite de ladite plate-forme (44) ; caractérisé en ce que ledit refoulement commun (70) a une ouverture sur la surface radialement interne (52) de la plate-forme (44).
  2. Segment d'enveloppe selon la revendication 1, dans lequel la pluralité de passages (66) s'étend à travers une section majeure d'une longueur axiale entière de la plate-forme d'enveloppe (44).
  3. Segment d'enveloppe selon la revendication 2, dans lequel la pluralité de passages (66) est en communication fluidique avec une cavité (64) définie entre des pattes avant et arrière (46, 48) du segment d'enveloppe de turbine (42).
  4. Pluralité de passages (66) selon une quelconque revendication précédente, dans laquelle les passages (66) comprennent une pluralité de trous droits sensiblement axiaux.
  5. Structure de moteur de turbine à gaz destinée à définir une portion d'une paroi externe d'un trajet de gaz annulaire d'une section de turbine, comprenant une enveloppe de turbine (32) comprenant une pluralité de segments d'enveloppe de turbine (42) tels que revendiqués dans la revendication 1, et une enveloppe externe d'aube de turbine (38) dotée d'une pluralité d'aubes (40) disposée immédiatement en aval de l'enveloppe de turbine, les passages (66) s'alignant sensiblement avec l'enveloppe externe d'aube de turbine (38), moyennant quoi de l'air de refroidissement délivré desdits passages vient impacter sensiblement l'étendue entière d'une circonférence d'une extrémité d'attaque de l'enveloppe externe d'aube de turbine (38).
  6. Structure de moteur de turbine à gaz selon la revendication 5, dans laquelle les passages (66) comprennent une pluralité de trous ayant au moins une admission (68) de ceux-ci sur une surface externe (50) de la plate-forme (44).
  7. Structure de moteur de turbine à gaz selon la revendication 6, dans laquelle la au moins une admission (68) des trous (66) est située dans une position proche et en aval d'une patte avant (46) de l'enveloppe.
  8. Structure de moteur de turbine à gaz selon la revendication 6 ou 7, dans laquelle les trous (66) s'étendent dans une direction sensiblement droite sur une section majeure de longueur axiale entière de la plate-forme (44).
  9. Procédé de réutilisation d'air de refroidissement d'enveloppe de turbine pour un refroidissement par impact sur une enveloppe externe d'aube de turbine aval (38), le procédé comprenant les étapes de :
    (a) direction de l'air de refroidissement à l'intérieur de et à travers une plate-forme (44) d'un segment d'enveloppe (42) tel que revendiqué dans la revendication 4 d'une enveloppe de turbine pour refroidir l'enveloppe de turbine ; et
    (b) utilisation de l'air de refroidissement de l'étape (a) pour former une pluralité de courants d'air de refroidissement sensiblement droits axialement vers une extrémité d'attaque de l'enveloppe externe d'aube de turbine (38) pour un refroidissement par impact sur l'enveloppe externe d'aube de turbine.
  10. Procédé selon la revendication 9, dans lequel les étapes (a) et (b) sont conduites sensiblement simultanément.
  11. Procédé selon la revendication 10, dans lequel les étapes (a) et (b) sont pratiquées en dirigeant l'air de refroidissement à travers une pluralité de passages sensiblement axiaux et droits (66) s'étendant à l'intérieur de la plate-forme (44) du segment d'enveloppe (42) de l'enveloppe de turbine pour délivrer les courants d'air de refroidissement sensiblement droits à une vitesse élevée.
  12. Procédé selon la revendication 11, dans lequel l'air de refroidissement est dirigé à partir d'une cavité (64) définie entre des pattes avant et arrière du segment d'enveloppe (42), dans les passages sensiblement axiaux et droits (66).
  13. Procédé selon l'une quelconque des revendications 9 à 12, comprenant en outre une étape (c) d'évacuation de l'air de refroidissement dans un trajet de gaz lors du refroidissement par impact de celui-ci sur l'enveloppe externe d'aube de turbine (38).
  14. Procédé selon l'une quelconque des revendications 9 à 13, comprenant la direction des courants d'air de refroidissement sensiblement droits d'une façon permettant un refroidissement par impact sur une circonférence sensiblement entière de l'extrémité d'attaque de l'enveloppe externe d'aube de turbine (38).
EP06253774.1A 2005-07-19 2006-07-19 Dispositif et méthode de refroidissement d'une virole de turbine et de l'anneau externe d'une aube statorique de turbine Active EP1746254B1 (fr)

Applications Claiming Priority (3)

Application Number Priority Date Filing Date Title
US11/183,741 US7520715B2 (en) 2005-07-19 2005-07-19 Turbine shroud segment transpiration cooling with individual cast inlet and outlet cavities
US11/183,922 US7374395B2 (en) 2005-07-19 2005-07-19 Turbine shroud segment feather seal located in radial shroud legs
US11/184,843 US20070020088A1 (en) 2005-07-20 2005-07-20 Turbine shroud segment impingement cooling on vane outer shroud

Publications (3)

Publication Number Publication Date
EP1746254A2 EP1746254A2 (fr) 2007-01-24
EP1746254A3 EP1746254A3 (fr) 2010-03-10
EP1746254B1 true EP1746254B1 (fr) 2016-03-23

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EP06253774.1A Active EP1746254B1 (fr) 2005-07-19 2006-07-19 Dispositif et méthode de refroidissement d'une virole de turbine et de l'anneau externe d'une aube statorique de turbine

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Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
KR101670618B1 (ko) 2010-04-20 2016-10-28 미츠비시 쥬고교 가부시키가이샤 분할 링 냉각 구조

Families Citing this family (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US7377742B2 (en) * 2005-10-14 2008-05-27 General Electric Company Turbine shroud assembly and method for assembling a gas turbine engine
US8998572B2 (en) * 2012-06-04 2015-04-07 United Technologies Corporation Blade outer air seal for a gas turbine engine

Family Cites Families (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2519374B1 (fr) * 1982-01-07 1986-01-24 Snecma Dispositif de refroidissement des talons d'aubes mobiles d'une turbine
JPH0234731U (fr) * 1988-08-30 1990-03-06
JP3779517B2 (ja) * 2000-01-21 2006-05-31 株式会社日立製作所 ガスタービン
JP2002201913A (ja) * 2001-01-09 2002-07-19 Mitsubishi Heavy Ind Ltd ガスタービンの分割壁およびシュラウド
JP4698847B2 (ja) * 2001-01-19 2011-06-08 三菱重工業株式会社 ガスタービン分割環
US6899518B2 (en) * 2002-12-23 2005-05-31 Pratt & Whitney Canada Corp. Turbine shroud segment apparatus for reusing cooling air

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
KR101670618B1 (ko) 2010-04-20 2016-10-28 미츠비시 쥬고교 가부시키가이샤 분할 링 냉각 구조

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Publication number Publication date
EP1746254A3 (fr) 2010-03-10
EP1746254A2 (fr) 2007-01-24

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