US20160160652A1 - Cooled pocket in a turbine vane platform - Google Patents

Cooled pocket in a turbine vane platform Download PDF

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Publication number
US20160160652A1
US20160160652A1 US14/751,192 US201514751192A US2016160652A1 US 20160160652 A1 US20160160652 A1 US 20160160652A1 US 201514751192 A US201514751192 A US 201514751192A US 2016160652 A1 US2016160652 A1 US 2016160652A1
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US
United States
Prior art keywords
platform
pocket
cover plate
airfoil
apertures
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.)
Abandoned
Application number
US14/751,192
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English (en)
Inventor
James P. Bangerter
Corneil S. Paauwe
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 Corporation
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 Corporation filed Critical United Technologies Corporation
Priority to US14/751,192 priority Critical patent/US20160160652A1/en
Publication of US20160160652A1 publication Critical patent/US20160160652A1/en
Abandoned legal-status Critical Current

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    • 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
    • F01D9/041—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector using blades
    • 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
    • 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
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

Definitions

  • a gas turbine engine typically includes a fan section, a compressor section, a combustor section and a turbine section. Air entering the compressor section is compressed and delivered into the combustion section where it is mixed with fuel and ignited to generate a high-speed exhaust gas flow. The high-speed exhaust gas flow expands through the turbine section to drive the compressor and the fan section.
  • the compressor section typically includes low and high pressure compressors, and the turbine section includes low and high pressure turbines.
  • the compressor or turbine sections may include vanes mounted on vane platforms.
  • the vanes may require cooling. Some areas that require cooling cannot be reached by impingement cooling. Therefore, the vane platforms may include “cavities” to provide cooling air to the vanes.
  • the cavities may be created in the platforms during casting of the platforms.
  • an airfoil in a featured embodiment, includes a platform including a pocket, a continuous solid cover plate bonded over the pocket, and a plurality of apertures in the platform in fluid communication with the pocket and adjacent the continuous solid cover plate.
  • the cover plate includes a contour extending into the pocket to define a height of the pocket.
  • the cover plate is bonded over the pocket by being welded to the platform.
  • a total cross-section of the plurality of apertures is less than approximately 30% of a cross-section of the pocket.
  • the plurality of apertures comprises a first set of apertures adjacent a first side of the cover plate and a second set of apertures adjacent a second side of the cover plate, the second side being opposite the first side.
  • the pocket includes a heat-transfer-enhancing surface feature.
  • a first side of the platform is adjacent a gas path, and where in the pocket is in a second side opposite the first side.
  • a height of the pocket is substantially uniform.
  • a gas turbine engine in another featured embodiment, includes a compressor section, a turbine section downstream from the compressor section, and a combustor section in communication with the compressor section and the turbine section.
  • At least one of the compressor and turbine sections includes a platform including a pocket in a first side of the platform, an airfoil extending from a second side of the platform opposite the first side, a continuous solid cover plate bonded over the pocket, and a plurality of apertures in the platform adjacent the continuous solid cover plate in fluid communication with the pocket.
  • the plurality of apertures comprises a first set of apertures adjacent a first side of the cover plate and a second set of apertures adjacent a second side of the cover plate, the second side being opposite the first side.
  • cooling air is supplied to the first set of apertures from the compressor section.
  • the first side of the platform is radially inward from the second side with respect to a longitudinal axis of the gas turbine engine.
  • the platform is a first platform, and further comprising a second platform, the airfoil supported between the first and second platforms.
  • the first platform is radially inward from the second platform with respect to a longitudinal axis of the gas turbine engine.
  • the airfoil is cantilever-mounted on the platform.
  • the platform and airfoil are in the turbine section.
  • a method of cooling an article of a gas turbine engine includes the steps of supplying cooling air to a pocket in a platform covered by a continuous solid cover plate via a first set of apertures adjacent to a first side of the continuous solid cover plate and removing cooling air from the pocket via a second set of apertures adjacent to a second side of the continuous solid cover plate, the second side being opposite the first side.
  • Another featured embodiment according to any of the previous embodiments further comprises the step of cooling an article with the cooling air prior to the supplying step.
  • the cooling air is supplied from a compressor section.
  • the article is an airfoil supported on the platform.
  • FIG. 1 schematically illustrates an example gas turbine engine.
  • FIG. 2 schematically illustrates a vane assembly for the gas turbine engine of FIG. 1 .
  • FIG. 3 schematically illustrates a cross-section of a platform of the vane assembly of FIG. 2 including a pocket.
  • FIG. 4A schematically illustrates a cross-section of the platform of the vane assembly of FIG. 2 including an open pocket.
  • FIG. 4B schematically illustrates a cross-section of the platform of the vane assembly of FIG. 2 including a pocket covered by a cover plate.
  • FIG. 5 schematically illustrates another cross-section of the platform and airfoil of the vane assembly of FIGS. 4A-4B .
  • FIG. 1 schematically illustrates a gas turbine engine 20 .
  • the gas turbine engine 20 is disclosed herein as a two-spool turbofan that generally incorporates a fan section 22 , a compressor section 24 , a combustor section 26 and a turbine section 28 .
  • Alternative engines might include an augmentor section (not shown) among other systems or features.
  • the fan section 22 drives air along a bypass flow path B in a bypass duct defined within a nacelle 15
  • the compressor section 24 drives air along a core flow path C for compression and communication into the combustor section 26 then expansion through the turbine section 28 .
  • the engine 20 generally includes a low speed spool 30 and a high speed spool 32 mounted for rotation about an engine central longitudinal axis A relative to an engine static structure 36 via several bearing systems 38 . It should be understood that various bearing systems 38 at various locations may alternatively or additionally be provided, and the location of bearing systems 38 may be varied as appropriate to the application.
  • the low speed spool 30 generally includes an inner shaft 40 that interconnects a fan 42 , a first (or low) pressure compressor 44 and a first (or low) pressure turbine 46 .
  • the inner shaft 40 is connected to the fan 42 through a speed change mechanism, which in exemplary gas turbine engine 20 is illustrated as a geared architecture 48 to drive the fan 42 at a lower speed than the low speed spool 30 .
  • the high speed spool 32 includes an outer shaft 50 that interconnects a second (or high) pressure compressor 52 and a second (or high) pressure turbine 54 .
  • a combustor 56 is arranged in exemplary gas turbine 20 between the high pressure compressor 52 and the high pressure turbine 54 .
  • a mid-turbine frame 57 of the engine static structure 36 is arranged generally between the high pressure turbine 54 and the low pressure turbine 46 .
  • the mid-turbine frame 57 further supports bearing systems 38 in the turbine section 28 .
  • the inner shaft 40 and the outer shaft 50 are concentric and rotate via bearing systems 38 about the engine central longitudinal axis A which is collinear with their longitudinal axes.
  • the core airflow is compressed by the low pressure compressor 44 then the high pressure compressor 52 , mixed and burned with fuel in the combustor 56 , then expanded over the high pressure turbine 54 and low pressure turbine 46 .
  • the mid-turbine frame 57 includes airfoils 59 which are in the core airflow path C.
  • the turbines 46 , 54 rotationally drive the respective low speed spool 30 and high speed spool 32 in response to the expansion.
  • gear system 48 may be located aft of combustor section 26 or even aft of turbine section 28
  • fan section 22 may be positioned forward or aft of the location of gear system 48 .
  • the engine 20 in one example is a high-bypass geared aircraft engine.
  • the engine 20 bypass ratio is greater than about six (6), with an example embodiment being greater than about ten (10)
  • the geared architecture 48 is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3 and the low pressure turbine 46 has a pressure ratio that is greater than about five.
  • the engine 20 bypass ratio is greater than about ten (10:1)
  • the fan diameter is significantly larger than that of the low pressure compressor 44
  • the low pressure turbine 46 has a pressure ratio that is greater than about five (5:1).
  • Low pressure turbine 46 pressure ratio is pressure measured prior to inlet of low pressure turbine 46 as related to the pressure at the outlet of the low pressure turbine 46 prior to an exhaust nozzle.
  • the geared architecture 48 may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans.
  • the fan section 22 of the engine 20 is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet.
  • “Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system.
  • the low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.45.
  • Low corrected fan tip speed is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram ° R.)/(518.7° R)] ⁇ 0.5.
  • the “Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150 ft/second.
  • the high and low pressure compressors 44 , 52 and the high and low pressure turbines 46 , 54 may include alternating rows of rotating blade assemblies 58 and stationary vane assemblies 60 .
  • the designs of the blade assemblies 58 and vane assemblies 60 can differ between the compressor 44 , 52 and the turbine 46 , 54 .
  • a representative vane assembly 60 is shown in FIG. 2 .
  • the vane assembly 60 of FIG. 2 includes an airfoil 62 supported between first and second platforms 64 , 66 .
  • the airfoil 62 may be cantilever-mounted on one of the first and second platforms 64 , 66 .
  • the first platform 64 is radially inward from the second platform 66 with respect to the engine axis A ( FIG. 1 ).
  • FIGS. 2-5 the first platform 64 is shown. However, it is to be understood that the below description is applicable to the second platform 66 as well.
  • FIG. 3 shows a cross-section of the platform 64 along the section line 3 - 3 in FIG. 2 .
  • the platform 64 includes a pocket 72 adjacent a footprint of the airfoil 62 .
  • a cover plate 74 covers the pocket 72 .
  • the cover plate 74 is continuous and solid. That is, the cover plate is free of any openings through which gas from the pocket 72 can escape.
  • the cover plate 74 can be bonded to the platform 64 over the pocket 72 so as to seal the pocket 72 , for example, by welding, brazing, or the like. Therefore, the pressure inside the pocket 72 when covered by the cover plate 74 can be higher than the pressure outside the pocket 72 .
  • the pocket 72 is substantially rectangular. However, in another example, the pocket can be another shape.
  • the pocket 72 can be formed in the platform 64 during the process of casting the platform 64 .
  • the pocket 72 can be formed or by machining, for example by electrical discharge machining (EDM).
  • EDM electrical discharge machining
  • the pocket 72 is formed by tooling used to cast the platform 64 .
  • the inlet apertures 80 are adjacent a first side 83 a of the pocket 72 and cover plate 74
  • the outlet apertures 82 are adjacent a second side 83 b of the pocket 72 and cover plate 74 opposite the first side 83 a .
  • the cooling air can be cooling air that is used to cool the airfoil 62 .
  • the apertures 80 , 82 may be, for example film cooling holes.
  • the outlet apertures 82 are open to the exterior of the platform 62 .
  • the pocket 72 is shown open.
  • the pocket 72 includes one or more heat-transfer-enhancing surface features 76 .
  • the heat-transfer-enhancing surface features 76 are trip strips 76 shown schematically in the internal surface of the pocket 72 .
  • the trip strips 76 enhance heat transfer from cooling air in the pocket 72 .
  • the trip strips 76 may be ridges that extend obliquely to flow direction through the pocket 72 , to turbulate the flow.
  • the heat-transfer-enhancing surface features 76 can be pedestals or other types of features.
  • the heat-transfer-enhancing surface features 76 can be excluded.
  • FIG. 5 shows a cross-section of the platform 64 and airfoil 62 along the section line 5 - 5 shown in FIG. 4B .
  • the cover plate 74 is between approximately 12 and 20 mils (0.30 to 0.51 mm) thick and includes a contour 78 that extends a depth D into the pocket 72 from a surface of the platform 64 .
  • the size of the contour 78 defines a height H of the pocket 72 .
  • the height H of the pocket 72 is between about 40 and 60 mils (1.02 to 1.52 mm), and is substantially uniform along a length L of the pocket 72 ( FIG. 5 ).
  • FIG. 5 shows a cross-section of the the platform 64 and airfoil 62 along the line 5 - 5 ( FIG. 4B ).
  • the core flow gas path C is adjacent a radially outward side 85 a of the platform 64 with respect to the engine axis A, and the pocket 72 is in on a radially inward side 85 b of the platform 64 with respect to the engine axis A.
  • the inlet aperture 80 opens at an inlet on the radially inward side 85 b of the platform 64 and extends across a rail 84 of the platform 64 .
  • the aperture 80 opens into the pocket 72 .
  • a total cross-section of apertures 80 , 82 opening into the pocket 72 is less than approximately 30% of a cross-section of the pocket 72 in order to maintain a high pressure in the pocket 72 .
  • the vane assembly 60 is in one of the low pressure and high pressure turbines 46 , 54 and cooling air for impingement cooling the airfoil 62 and for cooling the platform 64 via the pocket 72 is supplied from one of the high or low pressure compressors 44 , 52 , respectively.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
US14/751,192 2014-07-14 2015-06-26 Cooled pocket in a turbine vane platform Abandoned US20160160652A1 (en)

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US14/751,192 US20160160652A1 (en) 2014-07-14 2015-06-26 Cooled pocket in a turbine vane platform

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Application Number Priority Date Filing Date Title
US201462023991P 2014-07-14 2014-07-14
US14/751,192 US20160160652A1 (en) 2014-07-14 2015-06-26 Cooled pocket in a turbine vane platform

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US14/751,192 Abandoned US20160160652A1 (en) 2014-07-14 2015-06-26 Cooled pocket in a turbine vane platform

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EP (1) EP2975222A1 (fr)

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20200340362A1 (en) * 2019-04-24 2020-10-29 United Technologies Corporation Vane core assemblies and methods

Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5344283A (en) * 1993-01-21 1994-09-06 United Technologies Corporation Turbine vane having dedicated inner platform cooling
US6254333B1 (en) * 1999-08-02 2001-07-03 United Technologies Corporation Method for forming a cooling passage and for cooling a turbine section of a rotary machine
US20060269409A1 (en) * 2005-05-27 2006-11-30 Mitsubishi Heavy Industries, Ltd. Gas turbine moving blade having a platform, a method of forming the moving blade, a sealing plate, and a gas turbine having these elements

Family Cites Families (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US6508620B2 (en) * 2001-05-17 2003-01-21 Pratt & Whitney Canada Corp. Inner platform impingement cooling by supply air from outside
US8851845B2 (en) * 2010-11-17 2014-10-07 General Electric Company Turbomachine vane and method of cooling a turbomachine vane
US8714909B2 (en) * 2010-12-22 2014-05-06 United Technologies Corporation Platform with cooling circuit

Patent Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5344283A (en) * 1993-01-21 1994-09-06 United Technologies Corporation Turbine vane having dedicated inner platform cooling
US6254333B1 (en) * 1999-08-02 2001-07-03 United Technologies Corporation Method for forming a cooling passage and for cooling a turbine section of a rotary machine
US20060269409A1 (en) * 2005-05-27 2006-11-30 Mitsubishi Heavy Industries, Ltd. Gas turbine moving blade having a platform, a method of forming the moving blade, a sealing plate, and a gas turbine having these elements

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20200340362A1 (en) * 2019-04-24 2020-10-29 United Technologies Corporation Vane core assemblies and methods
US11021966B2 (en) * 2019-04-24 2021-06-01 Raytheon Technologies Corporation Vane core assemblies and methods

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EP2975222A1 (fr) 2016-01-20

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