EP0924384A2 - Refroidissement de l'arête amont d'une aube pour une turbomachine - Google Patents

Refroidissement de l'arête amont d'une aube pour une turbomachine Download PDF

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Publication number
EP0924384A2
EP0924384A2 EP98310191A EP98310191A EP0924384A2 EP 0924384 A2 EP0924384 A2 EP 0924384A2 EP 98310191 A EP98310191 A EP 98310191A EP 98310191 A EP98310191 A EP 98310191A EP 0924384 A2 EP0924384 A2 EP 0924384A2
Authority
EP
European Patent Office
Prior art keywords
airfoil
trench
leading edge
cooling
cooling air
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.)
Withdrawn
Application number
EP98310191A
Other languages
German (de)
English (en)
Other versions
EP0924384A3 (fr
Inventor
George P. Liang
Thomas A. Auxier
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
Publication of EP0924384A2 publication Critical patent/EP0924384A2/fr
Publication of EP0924384A3 publication Critical patent/EP0924384A3/fr
Withdrawn legal-status Critical Current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
    • 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
    • F05D2260/202—Heat transfer, e.g. cooling by film cooling

Definitions

  • This invention relates to cooled rotor blades and/or stator vanes for gas turbines in general, and to apparatus that cools the leading edge and establishes film cooling along the surface of the rotor blade or stator vane in particular.
  • stator vane and rotor blade stages In the turbine section of a gas turbine engine, core gas travels through a plurality of stator vane and rotor blade stages.
  • Each stator vane or rotor blade has an airfoil with one or more internal cavities surrounded by an exterior wall. The suction and pressure sides of the exterior wall extend between the leading and trailing edges of the airfoil.
  • Stator vane airfoils extend spanwise between inner and outer platforms and the rotor blade airfoils extend spanwise between a platform and a blade tip.
  • High temperature core gas (which includes air and combustion products) encountering the leading edge of an airfoil will diverge around the suction and pressure sides of the airfoil, or impinge on the leading edge.
  • the point along the leading edge where the velocity of the core gas flow goes to zero i.e., the impingement point
  • the stagnation point There is a stagnation point at every spanwise position along the leading edge of the airfoil, and collectively those points are referred to as the stagnation line. Air impinging on the leading edge of the airfoil is subsequently diverted around either side of the airfoil.
  • Cooling air typically bled off of a compressor stage at a temperature lower and pressure higher than the core gas passing through the turbine section, is used to cool the airfoils.
  • the cooler compressor air provides the medium for heat transfer and the difference in pressure provides the energy required to pass the cooling air through the stator or rotor stage.
  • film cooling In many cases, it is desirable to establish film cooling along the surface of the stator or rotor airfoil.
  • a film of cooling air traveling along the surface of the airfoil transfers thermal energy away from the airfoil, increases the uniformity of the cooling, and insulates the airfoil from the passing hot core gas.
  • film cooling is difficult to establish and maintain in the turbulent environment of a gas turbine.
  • film cooling air is bled out of cooling apertures extending through the external wall of the airfoil. The term "bled" reflects the small difference in pressure motivating the cooling air out of the internal cavity of the airfoil.
  • One of the problems associated with using apertures to establish a cooling air film is the films sensitivity to pressure difference across the apertures. Too great a pressure difference across an aperture will cause the air to jet out into the passing core gas rather than aid in the formation of a film of cooling air. Too small a pressure difference will result in negligible cooling air flow through the aperture, or an in-flow of hot core gas. Both cases adversely affect film cooling effectiveness.
  • Another problem associated with using apertures to establish film cooling is that cooling air is dispensed from discrete points along the span of the airfoil, rather than along a continuous line. The gaps between the apertures, and areas immediately downstream of those gaps, are exposed to less cooling air than are the apertures and the spaces immediately downstream of the apertures, and are therefore more susceptible to thermal degradation.
  • What is needed is an apparatus that provides adequate cooling along the leading edge of an airfoil, one that creates a uniform and durable cooling air film downstream of the leading edge on both sides of the airfoil, and one that creates minimal stress concentrations within the airfoil wall.
  • a hollow airfoil which includes a body, a trench, and a plurality of cooling apertures disposed within the trench.
  • the body extends chordwise between a leading edge and a trailing edge, and includes an exterior wall surrounding a cavity.
  • the trench is disposed in the exterior wall along the leading edge, extending in a spanwise direction.
  • An advantage of the present invention is that uniform and durable film cooling downstream of the leading edge is provided on both sides of the airfoil.
  • the cooling air bleeds out of the trench on both sides and creates continuous film cooling downstream of the leading edge.
  • the trench minimizes cooling losses characteristic of cooling apertures, and thereby provides more cooling air for film development and maintenance.
  • Another advantage of the present invention is that stress is minimized along the leading edge and areas immediately downstream of the leading edge.
  • One characteristic responsible for minimizing stress is the trench of cooling air extending continuously along the leading edge.
  • the trench substantially eliminates discrete cooling points separted by uncooled areas, and thereby eliminates the thermally induced stress associated therewith.
  • the trench also minimizes stress by distributing cooling air along the leading edge.
  • the cooling air bleeds out of the trench on both sides and creates continuous film cooling downstream of the leading edge.
  • the continuous film eliminates uncooled zones between and downstream of cooling apertures, and thereby eliminates the thermally induced stress associated therewith.
  • a gas turbine engine turbine rotor blade 10 includes a root portion 12, a platform 14, an airfoil 16, and a blade tip 18.
  • the airfoil 16 comprises one or more internal cavities 20 (see FIGS. 2 and 3) surrounded by a external wall 22, at least one of which is proximate the leading edge 24 of the airfoil 16.
  • the suction side 26 and the pressure side 28 of the external wall 22 extend chordwise 27 between the leading edge 24 and the trailing edge 29 of the airfoil 16, and spanwise 31 between the platform 14 and the blade tip 18.
  • the leading edge 24 has a smoothly curved contour which blends with the suction side 26 and the pressure side 28 of the airfoil 16.
  • FIG. 3 shows an embodiment having a plurality of trenches 30.
  • Each trench 30 extends substantially the entire span 31 (see FIG. 1) of the airfoil 16 leading edge 24.
  • a plurality of cooling apertures 36 are disposed in the trench 30, extending between an internal cavity 20 and the trench 30.
  • the shape of the cooling apertures 36 and their position within the trench 30 will vary depending upon the application. In most cases, however, the cooling apertures 36 are uniformly distributed in the base 32 of the trench 30 throughout the span 31.
  • the cooling apertures 36 include a diffusion portion 38.
  • cooling air typically bled off of a compressor stage is routed into the airfoil 16 of the rotor blade 10 (or stator vane) by means well known in the art. Cooling air disposed within the internal cavity 20 proximate the leading edge 24 of the airfoil 16 is at a lower temperature and higher pressure than the core gas flowing past the external wall 22 of the airfoil 16. The pressure difference across the airfoil external wall 22 forces the internal cooling air to enter the cooling apertures 36 and subsequently pass into the trench(es) 30 located in the external wall 22 along the leading edge 24. The cooling air exiting the cooling apertures 36 diffuses into the cooling air already in the trench 30 and distributes within the trench 30. In the preferred embodiment where the cooling apertures 36 include diffusion portions 38, the diffusion portions 38 increase cooling air diffusion and distribution and therefore uniformity within the trench 30.
  • the difference in pressure across a cooling aperture 36 is a function of the local internal cavity 22 pressure and the local core gas pressure adjacent the aperture 36. Both of these pressures vary as a function of time. If the core gas pressure is high and the internal cavity pressure is low adjacent a particular cooling aperture in a conventional scheme, undesirable hot core gas in-flow can occur. The present invention minimizes the opportunity for the undesirable in-flow because the cooling air from all apertures 36 collectively distributes within the trench 30, thereby decreasing the opportunity for any low pressure zones to occur. Likewise, the distribution of cooling air within the trench 30 also avoids cooling air pressure spikes which, in a conventional scheme, would jet the cooling air into the core gas rather than add it to the film of cooling air downstream.
  • Cooling air subsequently exits the trench 30 in a uniform manner along both spanwise sides of the trench 30.
  • the exiting flow forms a film of cooling air on both sides of the trench 30 that extends downstream.
  • the cooling air exiting a trench 30 positioned downstream of the stagnation point 40 of the airfoil 16 may exit predominantly on the downstream side of the trench 30.
  • film cooling emanating from a upstream trench 30 predominantly cools the external wall 22 of the airfoil 16 between the two adjacent trenches 30.
  • FIGS. 2 and 3 show a partial sectional view of an airfoil.
  • the airfoil 16 may be that of a stator vane or a rotor blade.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP98310191A 1997-12-17 1998-12-11 Refroidissement de l'arête amont d'une aube pour une turbomachine Withdrawn EP0924384A3 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
US99232397A 1997-12-17 1997-12-17
US992323 1997-12-17

Publications (2)

Publication Number Publication Date
EP0924384A2 true EP0924384A2 (fr) 1999-06-23
EP0924384A3 EP0924384A3 (fr) 2000-08-23

Family

ID=25538194

Family Applications (1)

Application Number Title Priority Date Filing Date
EP98310191A Withdrawn EP0924384A3 (fr) 1997-12-17 1998-12-11 Refroidissement de l'arête amont d'une aube pour une turbomachine

Country Status (3)

Country Link
EP (1) EP0924384A3 (fr)
JP (1) JPH11247611A (fr)
KR (1) KR19990063131A (fr)

Cited By (21)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP1262631A2 (fr) * 2001-05-21 2002-12-04 United Technologies Corporation Aube de turbine refroidie par couche d'air
EP1645721A2 (fr) 2004-10-04 2006-04-12 ALSTOM Technology Ltd Aube de turbine à gaz avec refroidissement du bord d'attaque
EP1898051A2 (fr) 2006-08-25 2008-03-12 ALSTOM Technology Ltd Aube de turbine à gaz avec refroidissement du bord d'attaque
EP1972396A1 (fr) * 2007-03-14 2008-09-24 United Technologies Corporation Fonctions coulées pour profil aérodynamique d'un moteur de turbine
EP2075409A2 (fr) * 2007-12-10 2009-07-01 United Technologies Corporation Bord d'attaque de profil aérodynamique
EP2154333A2 (fr) * 2008-08-14 2010-02-17 United Technologies Corporation Aube refroidie et ensemble de turbine associé
CN102482944A (zh) * 2009-09-02 2012-05-30 西门子公司 构成为转子盘或者涡轮叶片的燃气轮机构件的冷却
US8371814B2 (en) 2009-06-24 2013-02-12 Honeywell International Inc. Turbine engine components
US8529193B2 (en) 2009-11-25 2013-09-10 Honeywell International Inc. Gas turbine engine components with improved film cooling
US8628293B2 (en) 2010-06-17 2014-01-14 Honeywell International Inc. Gas turbine engine components with cooling hole trenches
CN103806952A (zh) * 2014-01-20 2014-05-21 北京航空航天大学 一种具有前缘凹腔的涡轮叶片
US20140356188A1 (en) * 2013-04-26 2014-12-04 Honeywell International, Inc. Turbine blade airfoils including film cooling systems, and methods for forming an improved film cooled airfoil of a turbine blade
US9022737B2 (en) 2011-08-08 2015-05-05 United Technologies Corporation Airfoil including trench with contoured surface
US9228440B2 (en) 2012-12-03 2016-01-05 Honeywell International Inc. Turbine blade airfoils including showerhead film cooling systems, and methods for forming an improved showerhead film cooled airfoil of a turbine blade
US9650900B2 (en) 2012-05-07 2017-05-16 Honeywell International Inc. Gas turbine engine components with film cooling holes having cylindrical to multi-lobe configurations
EP3179039A1 (fr) * 2015-12-11 2017-06-14 Rolls-Royce plc Composant pour moteur à turbine à gaz
US9719357B2 (en) 2013-03-13 2017-08-01 Rolls-Royce Corporation Trenched cooling hole arrangement for a ceramic matrix composite vane
US10113433B2 (en) 2012-10-04 2018-10-30 Honeywell International Inc. Gas turbine engine components with lateral and forward sweep film cooling holes
EP3502419A1 (fr) * 2017-12-21 2019-06-26 Rolls-Royce plc Profil aérodynamique comportant une agencement de refroidissement
US11021965B2 (en) 2016-05-19 2021-06-01 Honeywell International Inc. Engine components with cooling holes having tailored metering and diffuser portions
US11293352B2 (en) * 2018-11-23 2022-04-05 Rolls-Royce Plc Aerofoil stagnation zone cooling

Families Citing this family (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JP7232035B2 (ja) * 2018-12-18 2023-03-02 三菱重工業株式会社 ガスタービンの静翼及びガスタービン

Family Cites Families (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
BE407456A (fr) * 1934-01-29
GB2127105B (en) * 1982-09-16 1985-06-05 Rolls Royce Improvements in cooled gas turbine engine aerofoils
US5152667A (en) * 1991-07-16 1992-10-06 General Motors Corporation Cooled wall structure especially for gas turbine engines

Non-Patent Citations (1)

* Cited by examiner, † Cited by third party
Title
None

Cited By (36)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP1262631A3 (fr) * 2001-05-21 2004-05-26 United Technologies Corporation Aube de turbine refroidié par couche d'air
US6932572B2 (en) 2001-05-21 2005-08-23 United Technologies Corporation Film cooled article with improved temperature tolerance
EP1262631A2 (fr) * 2001-05-21 2002-12-04 United Technologies Corporation Aube de turbine refroidie par couche d'air
EP1645721A2 (fr) 2004-10-04 2006-04-12 ALSTOM Technology Ltd Aube de turbine à gaz avec refroidissement du bord d'attaque
US7300252B2 (en) 2004-10-04 2007-11-27 Alstom Technology Ltd Gas turbine airfoil leading edge cooling construction
US7997866B2 (en) 2006-08-25 2011-08-16 Alstom Technology Ltd. Gas turbine airfoil with leading edge cooling
EP1898051A2 (fr) 2006-08-25 2008-03-12 ALSTOM Technology Ltd Aube de turbine à gaz avec refroidissement du bord d'attaque
US8695683B2 (en) 2007-03-14 2014-04-15 United Technologies Corporation Cast features for a turbine engine airfoil
EP1972396A1 (fr) * 2007-03-14 2008-09-24 United Technologies Corporation Fonctions coulées pour profil aérodynamique d'un moteur de turbine
US7980819B2 (en) 2007-03-14 2011-07-19 United Technologies Corporation Cast features for a turbine engine airfoil
US8955576B2 (en) 2007-03-14 2015-02-17 United Technologies Corporation Cast features for a turbine engine airfoil
EP2075409A2 (fr) * 2007-12-10 2009-07-01 United Technologies Corporation Bord d'attaque de profil aérodynamique
EP2075409A3 (fr) * 2007-12-10 2012-04-25 United Technologies Corporation Bord d'attaque de profil aérodynamique
US8439644B2 (en) 2007-12-10 2013-05-14 United Technologies Corporation Airfoil leading edge shape tailoring to reduce heat load
EP2154333A2 (fr) * 2008-08-14 2010-02-17 United Technologies Corporation Aube refroidie et ensemble de turbine associé
EP2154333A3 (fr) * 2008-08-14 2012-11-07 United Technologies Corporation Aube refroidie et ensemble de turbine associé
US8371814B2 (en) 2009-06-24 2013-02-12 Honeywell International Inc. Turbine engine components
US8956116B2 (en) 2009-09-02 2015-02-17 Siemens Aktiengesellschaft Cooling of a gas turbine component designed as a rotor disk or turbine blade
CN102482944B (zh) * 2009-09-02 2016-01-27 西门子公司 构成为转子盘或者涡轮叶片的燃气轮机构件的冷却
CN102482944A (zh) * 2009-09-02 2012-05-30 西门子公司 构成为转子盘或者涡轮叶片的燃气轮机构件的冷却
US8529193B2 (en) 2009-11-25 2013-09-10 Honeywell International Inc. Gas turbine engine components with improved film cooling
US8628293B2 (en) 2010-06-17 2014-01-14 Honeywell International Inc. Gas turbine engine components with cooling hole trenches
US9022737B2 (en) 2011-08-08 2015-05-05 United Technologies Corporation Airfoil including trench with contoured surface
US9650900B2 (en) 2012-05-07 2017-05-16 Honeywell International Inc. Gas turbine engine components with film cooling holes having cylindrical to multi-lobe configurations
US10113433B2 (en) 2012-10-04 2018-10-30 Honeywell International Inc. Gas turbine engine components with lateral and forward sweep film cooling holes
US9228440B2 (en) 2012-12-03 2016-01-05 Honeywell International Inc. Turbine blade airfoils including showerhead film cooling systems, and methods for forming an improved showerhead film cooled airfoil of a turbine blade
US9719357B2 (en) 2013-03-13 2017-08-01 Rolls-Royce Corporation Trenched cooling hole arrangement for a ceramic matrix composite vane
US9562437B2 (en) * 2013-04-26 2017-02-07 Honeywell International Inc. Turbine blade airfoils including film cooling systems, and methods for forming an improved film cooled airfoil of a turbine blade
US20140356188A1 (en) * 2013-04-26 2014-12-04 Honeywell International, Inc. Turbine blade airfoils including film cooling systems, and methods for forming an improved film cooled airfoil of a turbine blade
CN103806952A (zh) * 2014-01-20 2014-05-21 北京航空航天大学 一种具有前缘凹腔的涡轮叶片
EP3179039A1 (fr) * 2015-12-11 2017-06-14 Rolls-Royce plc Composant pour moteur à turbine à gaz
US11021965B2 (en) 2016-05-19 2021-06-01 Honeywell International Inc. Engine components with cooling holes having tailored metering and diffuser portions
US11286791B2 (en) 2016-05-19 2022-03-29 Honeywell International Inc. Engine components with cooling holes having tailored metering and diffuser portions
EP3502419A1 (fr) * 2017-12-21 2019-06-26 Rolls-Royce plc Profil aérodynamique comportant une agencement de refroidissement
US20200024961A1 (en) * 2017-12-21 2020-01-23 Rolls-Royce Plc Aerofoil cooling arrangement
US11293352B2 (en) * 2018-11-23 2022-04-05 Rolls-Royce Plc Aerofoil stagnation zone cooling

Also Published As

Publication number Publication date
EP0924384A3 (fr) 2000-08-23
KR19990063131A (ko) 1999-07-26
JPH11247611A (ja) 1999-09-14

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