EP1882819A1 - Microcircuits pour pales de turbine intégrés dans la plate-forme, la pointe, et l'aube - Google Patents

Microcircuits pour pales de turbine intégrés dans la plate-forme, la pointe, et l'aube Download PDF

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Publication number
EP1882819A1
EP1882819A1 EP07252838A EP07252838A EP1882819A1 EP 1882819 A1 EP1882819 A1 EP 1882819A1 EP 07252838 A EP07252838 A EP 07252838A EP 07252838 A EP07252838 A EP 07252838A EP 1882819 A1 EP1882819 A1 EP 1882819A1
Authority
EP
European Patent Office
Prior art keywords
cooling
turbine engine
engine component
component according
microcircuit
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
EP07252838A
Other languages
German (de)
English (en)
Other versions
EP1882819B1 (fr
Inventor
Francisco J. Cunha
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.)
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
Priority claimed from US11/489,155 external-priority patent/US7513744B2/en
Priority claimed from US11/491,405 external-priority patent/US7553131B2/en
Application filed by United Technologies Corp filed Critical United Technologies Corp
Publication of EP1882819A1 publication Critical patent/EP1882819A1/fr
Application granted granted Critical
Publication of EP1882819B1 publication Critical patent/EP1882819B1/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
    • 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
    • F01D5/22—Blade-to-blade connections, e.g. for damping vibrations
    • F01D5/225—Blade-to-blade connections, e.g. for damping vibrations by shrouding
    • 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
    • F05D2240/80—Platforms for stationary or moving blades
    • F05D2240/81—Cooled platforms
    • 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

  • the present invention relates to a turbine engine component having an integrated system for cooling the platform, the tip, and the main body of an airfoil portion of the component.
  • FIG. 1 depicts an engine arrangement 10 illustrating the relative location of a high pressure turbine blade 12.
  • FIGS. 2 and 3 depict the main design characteristics of a typical conventionally cooled high-pressure blade 12.
  • cooling flow passes through these blades by means of internal cooling channels 14 that are turbulated with trip strips 16 for enhancing heat transfer inside the blade.
  • the cooling effectiveness of these blades is around 0.50 with a convective efficiency of around 0.40.
  • cooling effectiveness is a dimensionless ratio of metal temperature ranging from zero to unity as the minimum and maximum values.
  • the convective efficiency is also a dimensionless ratio and denotes the ability for heat pick-up by the coolant, with zero and unity denoting no heat pick-up and maximum heat pick-up respectively.
  • the blade cooling flow should not increase and if possible, even decrease for turbine efficiency improvements. That objective is extremely difficult to achieve with current cooling technology which is shown schematically in FIGS. 2 and 3.
  • the cooling flow would have to increase more than 5% of the engine core flow.
  • the metal temperature in the embodiment of FIG. 3 is about 2180 degrees Fahrenheit (1193°C). This level of temperature is considered above the target limit.
  • the airfoil with a thermal barrier coating is a first requirement.
  • the other requirements are: (1) improved film cooling in terms of slots for increased film coverage; (2) improved heat pick-up; and (3) improved heat transfer coefficients in the blade cooling passages.
  • the overall cooling effectiveness will approach 0.8 with a convective efficiency approaching 0.5, allowing for a lower cooling flow of no more than 3.5% of the engine core flow.
  • a turbine engine component having an airfoil portion with a pressure side and a suction side.
  • the turbine engine component broadly comprises means for cooling the suction side of the airfoil portion, which cooling means comprises a first cooling microcircuit embedded within a first wall forming the suction side.
  • the first cooling microcircuit has means for allowing a cooling fluid in the first cooling microcircuit to exit at a tip of the airfoil portion.
  • the turbine engine component preferably further has a second cooling microcircuit in the pressure side of the airfoil portion and integrated means for cooling a platform portion of the turbine engine component.
  • a turbine engine component 90 such as a high pressure turbine blade, is cooled using the cooling design scheme of the present invention.
  • the cooling design scheme encompasses two serpentine microcircuits 100 and 102 located peripherally in the airfoil walls 104 and 106 respectively for cooling the main body 108 of the airfoil portion 110 of the turbine engine component.
  • Separate cooling microcircuits 96 and 98 may be used to cool the leading and trailing edges 112 and 114 respectively of the airfoil main body 108.
  • the coolant inside the turbine engine component may be used to feed the leading and trailing edge regions 112 and 114.
  • the coolant may be ejected out of the turbine engine component by means of film cooling.
  • the microcircuit 102 has a fluid inlet 126 for supplying cooling fluid to a first leg 128.
  • the inlet 126 receives the cooling fluid from one of the feed cavities 142 in the turbine engine component. Fluid flowing through the first leg 128 travels to an intermediate leg 130 and from there to an outlet leg 132. Fluid supplied by one of the feed cavities 142 may also be introduced into the cooling microcircuit 96 and used to cool the leading edge 112 of the airfoil portion 110.
  • Tne cooling microcircuit 96 may include fluid passageway 131 having fluid outlets 133.
  • fluid from the outlet leg 132 may be used to cool the leading edge 112 via an outlet passage 135.
  • the thermal load to the turbine engine component may not require film cooling from each of the legs that form the serpentine peripheral cooling microcircuit 102.
  • the flow of cooling fluid may be allowed to exit from the outlet leg 132 at the tip 134 by means of film blowing from the pressure side 116 to the suction side 118 of the turbine engine component.
  • the outlet leg 132 may communicate with a passageway 136 in the tip 134 having fluid outlets 138.
  • the serpentine cooling microcircuit 100 for the pressure side 116 of the airfoil portion 110.
  • the microcircuit 100 has an inlet 141 which communicates with one of the feed cavities 142 and a first leg 144 which receives cooling fluid from the inlet 141.
  • the cooling fluid in the first leg 144 flows through the intermediate leg 146 and through the outlet leg 148.
  • fluid from the feed cavity 142 may also be supplied to the trailing edge cooling microcircuit 98.
  • the cooling microcircuit 98 may have a plurality of fluid passageways 150 which have outlets 152 for distributing cooling fluid over the trailing edge 114 of the airfoil portion 110.
  • the outlet leg 148 may have one or more fluid outlets 153 for supplying a film of cooling fluid over the pressure side 116 of the airfoil portion 110 in the region of the trailing edge 114.
  • cooling microcircuit scheme of FIGS. 4 - 6 is completely different from existing designs where a dedicated cooling passage, denoted as a tip flag is employed for cooling the tip 134.
  • the pressure side 116 of the airfoil main body 108 is cooled with a serpentine microcircuit 100 located peripherally in the airfoil wall 104.
  • a flow exits in a series of film cooling slots 153 close to the aft side of the airfoil 110 to protect the airfoil trailing edge 114.
  • each leg 128, 130, 132, 144, 146, and 148 of the serpentine cooling microcircuits 100 and 102 may be provided with one or more internal features (not shown), such as pedestals and/or trip strips, to enhance the heat pick-up and increase the heat transfer coefficients characteristics inside the cooling blade passage(s).
  • cooling microcircuits may be located around and imbedded in a platform portion 170 of the turbine blade.
  • the cooling microcircuits may include a leading edge or forward cooling microcircuit 172 having an inlet portion A and an outlet portion B.
  • the inlet portion A may receive fluid from one of the feed cavities 142. Fluid from the outlet portion B flows back into the cooling microcircuit 96.
  • the platform cooling microcircuits may include a trailing edge or aft cooling microcircuit 180 having an inlet portion C and an outlet portion D.
  • the inlet portion C may receive fluid from one of the feed cavities 142. Fluid from the outlet portion D flows into the cooling microcircuit 98.
  • the platform cooling is independent of the serpentine cooling microcircuits 100 and 102 used for the airfoil portion 100.
  • the inlet coolant flow to either of the leading and trailing edge cooling microcircuits 172 and 180 comes from a lower radii. This coolant flow is allowed to pass through the platform walls before discharging into the cooling microcircuit 96 or 98 at a higher radii.
  • the rotational pumping which is created, along with the ejector-type action of the main flow, will ensure circulation in the peripheral platform cooling microcircuits 172 and 180.
  • an integrated cooling system has been devised to cool the platform 170, the main body 108 of the airfoil portion 110, and the tip 134 of the airfoil portion 110 by taking advantage of the microcircuit cooling characteristics.
  • the platform cooling microcircuits 172 and 180 may be provided with one or more internal features (not shown), such as pedestals, to enhance heat pick-up and increase the heat transfer coefficient characteristics inside the cooling passage(s) of the cooling microcircuits.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP20070252838 2006-07-18 2007-07-18 Microcircuits pour pales de turbine intégrés dans la plate-forme, la pointe, et l'aube Active EP1882819B1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
US11/489,155 US7513744B2 (en) 2006-07-18 2006-07-18 Microcircuit cooling and tip blowing
US11/491,405 US7553131B2 (en) 2006-07-21 2006-07-21 Integrated platform, tip, and main body microcircuits for turbine blades

Publications (2)

Publication Number Publication Date
EP1882819A1 true EP1882819A1 (fr) 2008-01-30
EP1882819B1 EP1882819B1 (fr) 2010-09-08

Family

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EP20070252838 Active EP1882819B1 (fr) 2006-07-18 2007-07-18 Microcircuits pour pales de turbine intégrés dans la plate-forme, la pointe, et l'aube

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EP (1) EP1882819B1 (fr)
DE (1) DE602007008996D1 (fr)

Cited By (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2385216A3 (fr) * 2010-05-06 2014-02-19 United Technologies Corporation Surface portante de turbine dotée de microcircuits de corps aboutissant à une plateforme
WO2014055811A1 (fr) * 2012-10-05 2014-04-10 General Electric Company Pales de turbine ayant un circuit de refroidissement de plate-forme et turbine à gaz correspondante
CN104520538A (zh) * 2012-08-09 2015-04-15 通用电气公司 涡轮叶片
WO2015181488A1 (fr) * 2014-05-28 2015-12-03 Snecma Aube de turbine a refroidissement optimise au niveau de son bord de fuite comprenant des conduits amont et aval et des cavités latérales internes
WO2015181497A1 (fr) * 2014-05-28 2015-12-03 Snecma Aube de turbine a refroidissement optimise
EP3508696A1 (fr) * 2018-01-09 2019-07-10 United Technologies Corporation Configuration de refroidissement de plate-forme d'aube de moteur de turbine à gaz à double paroi dotée d'un déflecteur d'impact
EP3508697A1 (fr) * 2018-01-09 2019-07-10 United Technologies Corporation Configuration de refroidissement de plate-forme d'aube de turbine de moteur à turbine à gaz à double paroi comportant un réapprovisionnement du noyau principal
EP3670841A1 (fr) * 2018-11-09 2020-06-24 United Technologies Corporation Profil aérodynamique ayant un réapprovisionnement hybride du passage central de l'enveloppe

Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5813835A (en) * 1991-08-19 1998-09-29 The United States Of America As Represented By The Secretary Of The Air Force Air-cooled turbine blade
US20040197190A1 (en) * 2003-04-07 2004-10-07 Stec Philip Francis Turbine blade with recessed squealer tip and shelf
EP1561900A2 (fr) * 2004-02-03 2005-08-10 United Technologies Corporation Circuit de refroidissement pour plate-forme d'aube de turbine
EP1586738A2 (fr) * 2004-04-14 2005-10-19 General Electric Company Méthode et dispositif de refroidissement des aubes des turbine
US20060056970A1 (en) * 2004-09-15 2006-03-16 General Electric Company Apparatus and methods for cooling turbine bucket platforms

Patent Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5813835A (en) * 1991-08-19 1998-09-29 The United States Of America As Represented By The Secretary Of The Air Force Air-cooled turbine blade
US20040197190A1 (en) * 2003-04-07 2004-10-07 Stec Philip Francis Turbine blade with recessed squealer tip and shelf
EP1561900A2 (fr) * 2004-02-03 2005-08-10 United Technologies Corporation Circuit de refroidissement pour plate-forme d'aube de turbine
EP1586738A2 (fr) * 2004-04-14 2005-10-19 General Electric Company Méthode et dispositif de refroidissement des aubes des turbine
US20060056970A1 (en) * 2004-09-15 2006-03-16 General Electric Company Apparatus and methods for cooling turbine bucket platforms

Cited By (21)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US9121290B2 (en) 2010-05-06 2015-09-01 United Technologies Corporation Turbine airfoil with body microcircuits terminating in platform
EP2385216A3 (fr) * 2010-05-06 2014-02-19 United Technologies Corporation Surface portante de turbine dotée de microcircuits de corps aboutissant à une plateforme
CN104520538B (zh) * 2012-08-09 2016-06-01 通用电气公司 涡轮叶片
CN104520538A (zh) * 2012-08-09 2015-04-15 通用电气公司 涡轮叶片
WO2014055811A1 (fr) * 2012-10-05 2014-04-10 General Electric Company Pales de turbine ayant un circuit de refroidissement de plate-forme et turbine à gaz correspondante
CN104704202A (zh) * 2012-10-05 2015-06-10 通用电气公司 涡轮机叶片组件的平台冷却
CN106470782A (zh) * 2014-05-28 2017-03-01 赛峰飞机发动机公司 包括上游管道和下游管道以及内侧部空腔的在后缘处具有优化的冷却的涡轮叶片
US10662789B2 (en) 2014-05-28 2020-05-26 Safran Aircraft Engines Turbine blade with optimised cooling at the trailing edge of same comprising upstream and downstream ducts and inner side cavities
FR3021699A1 (fr) * 2014-05-28 2015-12-04 Snecma Aube de turbine a refroidissement optimise au niveau de son bord de fuite
WO2015181497A1 (fr) * 2014-05-28 2015-12-03 Snecma Aube de turbine a refroidissement optimise
WO2015181488A1 (fr) * 2014-05-28 2015-12-03 Snecma Aube de turbine a refroidissement optimise au niveau de son bord de fuite comprenant des conduits amont et aval et des cavités latérales internes
US20170183969A1 (en) * 2014-05-28 2017-06-29 Safran Aircraft Engines Turbine blade with optimised cooling
US10689985B2 (en) 2014-05-28 2020-06-23 Safran Aircraft Engines Turbine blade with optimised cooling
FR3021697A1 (fr) * 2014-05-28 2015-12-04 Snecma Aube de turbine a refroidissement optimise
CN106470782B (zh) * 2014-05-28 2020-05-08 赛峰飞机发动机公司 叶片以及相关的模制装置、涡轮和涡轮机
US10648343B2 (en) 2018-01-09 2020-05-12 United Technologies Corporation Double wall turbine gas turbine engine vane platform cooling configuration with main core resupply
EP3508697A1 (fr) * 2018-01-09 2019-07-10 United Technologies Corporation Configuration de refroidissement de plate-forme d'aube de turbine de moteur à turbine à gaz à double paroi comportant un réapprovisionnement du noyau principal
US10662780B2 (en) 2018-01-09 2020-05-26 United Technologies Corporation Double wall turbine gas turbine engine vane platform cooling configuration with baffle impingement
EP3508696A1 (fr) * 2018-01-09 2019-07-10 United Technologies Corporation Configuration de refroidissement de plate-forme d'aube de moteur de turbine à gaz à double paroi dotée d'un déflecteur d'impact
EP3670841A1 (fr) * 2018-11-09 2020-06-24 United Technologies Corporation Profil aérodynamique ayant un réapprovisionnement hybride du passage central de l'enveloppe
US11180998B2 (en) 2018-11-09 2021-11-23 Raytheon Technologies Corporation Airfoil with skincore passage resupply

Also Published As

Publication number Publication date
DE602007008996D1 (de) 2010-10-21
EP1882819B1 (fr) 2010-09-08

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