EP2434096A2 - Gasturbinenschaufel mit einem Leitungssockel - Google Patents

Gasturbinenschaufel mit einem Leitungssockel Download PDF

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Publication number
EP2434096A2
EP2434096A2 EP11182897A EP11182897A EP2434096A2 EP 2434096 A2 EP2434096 A2 EP 2434096A2 EP 11182897 A EP11182897 A EP 11182897A EP 11182897 A EP11182897 A EP 11182897A EP 2434096 A2 EP2434096 A2 EP 2434096A2
Authority
EP
European Patent Office
Prior art keywords
leading edge
airfoil
recited
side wall
rib
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.)
Granted
Application number
EP11182897A
Other languages
English (en)
French (fr)
Other versions
EP2434096B1 (de
EP2434096A3 (de
Inventor
Brandon W. Spangler
Amanda Jean Learned
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 EP2434096A2 publication Critical patent/EP2434096A2/de
Publication of EP2434096A3 publication Critical patent/EP2434096A3/de
Application granted granted Critical
Publication of EP2434096B1 publication Critical patent/EP2434096B1/de
Active legal-status Critical Current
Anticipated expiration legal-status Critical

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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
    • 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
    • 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
    • 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

Definitions

  • the present disclosure relates to a gas turbine engine, and more particularly to an airfoil cooling arrangement.
  • a gas turbine engine includes a compressor section that compresses air then channels the compressed air to a combustor section wherein the compressed airflow is mixed with fuel and ignited to generate high temperature combustion gases.
  • the combustion core gases flow downstream through a turbine section which extracts energy therefrom to power the compressor section and a fan section. Since the combustion core gases are at a high temperature, turbine vanes and turbine blades within the turbine section may have relatively high heat loads at the leading edges.
  • An airfoil for a gas turbine engine includes a pressure side wall and a suction side wall which define a leading edge cavity and a forward cavity between the pressure side wall and the suction side wall, with the leading edge cavity at least partially defined by a leading edge wall which extends between the pressure side wall and the suction side wall.
  • a rib between the pressure side wall and the suction side wall separates the forward cavity and the leading edge cavity.
  • a pedestal extends between the leading edge wall and the rib.
  • An airfoil for a gas turbine engine includes a multiple of pedestals which extend between a leading edge and a rib, the multiple of pedestals arrayed along a length of the airfoil between a first end portion and a second end portion.
  • Figure 1 schematically illustrates a gas turbine engine 10 which generally includes a fan section 12, a compressor section 14, a combustor section 16, and a turbine section 18. Within and aft of the combustor section 16, engine components are typically cooled due to intense temperature of the combustion core gases. While a two spool high bypass turbofan engine is schematically illustrated in the disclosed non-limiting embodiment, it should be understood that the disclosure is applicable to other gas turbine engine configurations.
  • the cooling airflow passes through at least one cooling circuit flow path 26 ( Figure 2 ) to transfer thermal energy from the component to the cooling airflow.
  • Each cooling circuit flow path 26 may be disposed in any component that requires cooling, and in most cases the component receives cooling airflow therethrough as the external surface thereof is exposed to combustion core gases.
  • the cooling circuit flow path 26 will be described herein as being disposed within a portion of an airfoil 32 such as that of a stator vane 24 or rotor blade 22. It should be understood, however, that the cooling circuit flow path 26 is not limited to these applications and may be utilized within other areas such as liners, seals, and other structures with stagnation regions exposed to high temperature core gas flow.
  • the cooling circuit flow path 26 communicates with a multiple of cavities, for example 34A-34B shown in Figure 3 , formed within the airfoil 32.
  • the multiple of cavities 34A-34B direct cooling airflow which may include air received from the compressor section into high temperature areas of the airfoil 32.
  • the airfoil 32 is defined by an outer airfoil wall surface 40 between a leading edge 36 and a trailing edge 42.
  • the outer airfoil wall surface 40 typically has a generally concave shaped portion forming a pressure side 40P and a generally convex shaped portion forming a suction side 40S which are connected by a leading edge wall 40L at the leading edge 36.
  • the outer airfoil wall surface 40 is longitudinally defined to span a first end portion 46 and a second end portion 48.
  • the end portions 46, 48 may include features to mount the airfoil to other structures such as engine static structure or rotor disk.
  • the end portions 46, 48 for a vane may include outer vane platforms and for a blade may include an attachment section and a blade tip. It should be understood that various component arrangement may likewise be utilized with the present invention.
  • the forward cavity 34A is generally defined by a first rib 54 just aft of the leading edge 36.
  • the first rib 54 separates the forward cavity 34A from a leading edge cavity 56 defined at least partially by the outer airfoil wall surface 40 and often referred to as a "peanut" cavity.
  • the first rib 54 may, for example, at least partially define an impingement leading edge 62 ( Figure 4 ) or a radial flow leading edge 64 ( Figure 5 ) which may span a portion of or the entire length of the airfoil 32. That is, the pedestals 60 may be specifically located along the entire airfoil 32 span or a select portion or portions thereof
  • the leading edge cavity 56 includes the multiple of pedestals 60 which are transverse to and extend between the leading edge 36 and the first rib 54. It should be understood that any number of pedestals 60 may be so positioned.
  • the pedestals 60 provide an additional thermal conductive path along a conduction path axis H ( Figure 6 ) from the leading edge 36 to the first rib 54 to reduce the temperature of the leading edge 36 as the leading edge 36 may otherwise be hundreds of degrees hotter than the pressure side 40P and suction side 40S of the airfoil 32 due to higher external heat transfer coefficients at the stagnation region S ( Figure 7 ). It should be understood that the stagnation region S is a region within which the combustion gas flow Mach number may be relatively low such that a temperature concentration occurs.
  • the first rib 54 may define a multiple of cooling holes 66 which communicate a cooling flow from the forward cavity 34A into the leading edge cavity 56 through the first rib 54 then out through a multiple of leading edge cooling holes 68. That is, the cooling flow is communicated generally along the pedestals 60.
  • the cooling flow from within the leading edge cavity 56 passes transverse to the pedestals 60 and out through a multiple of leading edge cooling holes 70. It should be understood that various such cooling schemes will benefit from the pedestals 60.
  • the pedestals 60 reduce leading edge 36 temperatures mainly from the enhanced conduction effects of the pedestals 60 from the leading edge 36 to the first rib 54 ( Figures 8 and 9 ).
  • a portion of the metal temperature reduction is achieved by the enhancement of the internal heat transfer coefficient as coolant flow passes over the pedestals 60.
  • the lower temperature at the stagnation region beneficially results in, for example, a higher oxidation, local creep, and Thermal Mechanical Fatigue (TMF) capability.
  • TMF Thermal Mechanical Fatigue
  • the pedestals 60 may be selectively oriented at a multiple of different angles in the leading edge cavity 56 to achieve the desired thermal reduction effect. That is, the pedestals 60-1, 60-2 may be aligned along conduction path axes H1, H2 ( Figure 10 ) which extend into the highest temperature areas in the stagnation region of the leading edge 36 ( Figure 11 ) to facilitate a more direct heat transfer from the leading edge 36 to the first rib 54. It should be understood that the axes H1, H2 may change along the span of the airfoil 32. The relative positions of the pedestals 60-1, 60-2 may thereby also change along the span to con-espand therewith.
  • the manufacture of the pedestals 60 may be achieved by a proprietary Fugitive Core Process which uses thermoplastic inserts to create a one piece core with multiple pull angles as developed by Alcoa Howmet of Cleveland Ohio USA.
  • a proprietary Fugitive Core Process which uses thermoplastic inserts to create a one piece core with multiple pull angles as developed by Alcoa Howmet of Cleveland Ohio USA.
  • sacrificial thermoplastic pieces make up the rib and leading edge pedestals; the thermoplastic pieces are assembled into the core die and core material is injected around the thermoplastic pieces; the thermoplastic pieces are melted, leaving voids in finished core; and metal fill voids in core to form pedestals in the finished part.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP11182897.6A 2010-09-28 2011-09-27 Gasturbinenschaufel mit einem Leitungssockel Active EP2434096B1 (de)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US12/892,056 US20120076660A1 (en) 2010-09-28 2010-09-28 Conduction pedestals for a gas turbine engine airfoil

Publications (3)

Publication Number Publication Date
EP2434096A2 true EP2434096A2 (de) 2012-03-28
EP2434096A3 EP2434096A3 (de) 2015-04-29
EP2434096B1 EP2434096B1 (de) 2016-08-03

Family

ID=44719513

Family Applications (1)

Application Number Title Priority Date Filing Date
EP11182897.6A Active EP2434096B1 (de) 2010-09-28 2011-09-27 Gasturbinenschaufel mit einem Leitungssockel

Country Status (2)

Country Link
US (1) US20120076660A1 (de)
EP (1) EP2434096B1 (de)

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2018153796A1 (en) * 2017-02-24 2018-08-30 Siemens Aktiengesellschaft A turbomachine blade or vane having a cooling channel with a criss-cross arrangement of pins
EP3594448B1 (de) * 2018-07-13 2022-11-02 Honeywell International Inc. Schaufelblatt mit konvektiver vorderkantenkühlung

Families Citing this family (11)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2436884A1 (de) * 2010-09-29 2012-04-04 Siemens Aktiengesellschaft Turbinenanordnung und Gasturbinenmotor
US9759072B2 (en) 2012-08-30 2017-09-12 United Technologies Corporation Gas turbine engine airfoil cooling circuit arrangement
US9115590B2 (en) 2012-09-26 2015-08-25 United Technologies Corporation Gas turbine engine airfoil cooling circuit
US10427213B2 (en) 2013-07-31 2019-10-01 General Electric Company Turbine blade with sectioned pins and method of making same
US9695696B2 (en) 2013-07-31 2017-07-04 General Electric Company Turbine blade with sectioned pins
US20150196489A1 (en) * 2013-12-18 2015-07-16 Massachusetts Institute Of Technology Polymer matrices for controlling crystallization
US20160230566A1 (en) * 2015-02-11 2016-08-11 United Technologies Corporation Angled pedestals for cooling channels
FR3057295B1 (fr) * 2016-10-12 2020-12-11 Safran Aircraft Engines Aube comprenant une plate-forme et une pale assemblees
US10787932B2 (en) 2018-07-13 2020-09-29 Honeywell International Inc. Turbine blade with dust tolerant cooling system
US10989067B2 (en) * 2018-07-13 2021-04-27 Honeywell International Inc. Turbine vane with dust tolerant cooling system
US11230929B2 (en) 2019-11-05 2022-01-25 Honeywell International Inc. Turbine component with dust tolerant cooling system

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Publication number Priority date Publication date Assignee Title
US4257737A (en) * 1978-07-10 1981-03-24 United Technologies Corporation Cooled rotor blade
US5271715A (en) * 1992-12-21 1993-12-21 United Technologies Corporation Cooled turbine blade
FR2765265B1 (fr) * 1997-06-26 1999-08-20 Snecma Aubage refroidi par rampe helicoidale, par impact en cascade et par systeme a pontets dans une double peau
US7018176B2 (en) * 2004-05-06 2006-03-28 United Technologies Corporation Cooled turbine airfoil
US7195458B2 (en) * 2004-07-02 2007-03-27 Siemens Power Generation, Inc. Impingement cooling system for a turbine blade
EP1921269A1 (de) * 2006-11-09 2008-05-14 Siemens Aktiengesellschaft Turbinenschaufel
GB0813839D0 (en) * 2008-07-30 2008-09-03 Rolls Royce Plc An aerofoil and method for making an aerofoil
US9353631B2 (en) * 2011-08-22 2016-05-31 United Technologies Corporation Gas turbine engine airfoil baffle

Non-Patent Citations (1)

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Title
None

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2018153796A1 (en) * 2017-02-24 2018-08-30 Siemens Aktiengesellschaft A turbomachine blade or vane having a cooling channel with a criss-cross arrangement of pins
EP3594448B1 (de) * 2018-07-13 2022-11-02 Honeywell International Inc. Schaufelblatt mit konvektiver vorderkantenkühlung

Also Published As

Publication number Publication date
EP2434096B1 (de) 2016-08-03
EP2434096A3 (de) 2015-04-29
US20120076660A1 (en) 2012-03-29

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