WO2014113043A1 - Flanc de raccordement composite pour aube directrice - Google Patents

Flanc de raccordement composite pour aube directrice Download PDF

Info

Publication number
WO2014113043A1
WO2014113043A1 PCT/US2013/026553 US2013026553W WO2014113043A1 WO 2014113043 A1 WO2014113043 A1 WO 2014113043A1 US 2013026553 W US2013026553 W US 2013026553W WO 2014113043 A1 WO2014113043 A1 WO 2014113043A1
Authority
WO
WIPO (PCT)
Prior art keywords
radius
curvature
airfoil
platform
guide vane
Prior art date
Application number
PCT/US2013/026553
Other languages
English (en)
Inventor
Christopher S. McKaveney
Perry Bowes
John Joseph PAPALIA
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
Publication of WO2014113043A1 publication Critical patent/WO2014113043A1/fr

Links

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
    • F01D5/141—Shape, i.e. outer, aerodynamic form
    • 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
    • F01D5/147—Construction, i.e. structural features, e.g. of weight-saving hollow 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
    • F05D2220/00—Application
    • F05D2220/30—Application in turbines
    • F05D2220/36—Application in turbines specially adapted for the fan of turbofan engines
    • 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
    • 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
    • F05D2250/00—Geometry
    • F05D2250/70—Shape
    • F05D2250/71—Shape curved
    • 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
    • F05D2300/00—Materials; Properties thereof
    • F05D2300/10—Metals, alloys or intermetallic compounds
    • F05D2300/12—Light metals
    • F05D2300/121—Aluminium
    • 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

  • This application relates to a structural guide vane for use in a gas turbine engine.
  • Gas turbine engines are known and, typically, include a fan delivering air into a bypass duct, where it provides propulsion for an associated aircraft and further providing air into a compressor.
  • the air is compressed and the compressed air passes downstream into a combustor.
  • the air is mixed with fuel and ignited and products of the combustion pass downstream over turbine rotors, driving them to rotate.
  • the turbine rotors in turn, drive compressor rotors and the fan.
  • the increased size in the fan has also required an increase in size in associated structure.
  • structural guide vanes are included in the fan and are utilized to direct the air. As the fan itself is increased in diameter, so has the structural guide vanes. With this change, it has been proposed to form the structural guide vanes out of lighter weight material.
  • traditional fan guide vanes have been formed with titanium, aluminum has been proposed. The aluminum is lighter weight than titanium, but is also not as strong.
  • a guide vane has an airfoil and a platform at least at one end of the airfoil.
  • a compound fillet blends the airfoil into the platform.
  • the compound fillet has a first portion at a first radius of curvature adjacent to the platform.
  • a second portion at a second radius of curvature blends the first portion into the airfoil. The second radius of curvature is different than the first radius of curvature.
  • the structural guide vane is for use in a fan section of a gas turbine engine.
  • the airfoil has a platform at each of two ends.
  • the compound fillet blends the airfoil into a platform that will be a radially inward platform when the guide vane is mounted in a gas turbine engine.
  • the airfoil and platform are formed, at least in part, of aluminum.
  • the second radius of curvature is greater than the first radius of curvature.
  • a ratio of the first radius of curvature to the second radius of curvature is equal to, or between, .05 and .55.
  • a fan section has a fan rotor with fan blades. At least one guide vane is positioned adjacent to the fan blades, with the guide vane having an airfoil and a platform at least at one end of the airfoil.
  • a compound fillet blends the airfoil into the platform.
  • the compound fillet has a first portion at a first radius of curvature adjacent to the platform.
  • a second portion at a second radius of curvature blends the first portion into the airfoil. The second radius of curvature is different than the first radius of curvature.
  • the airfoil has a platform at each of two ends.
  • the compound fillet blends the airfoil into a platform that will be a radially inward platform when the guide vane is mounted in a gas turbine engine.
  • the airfoil and platform are formed, at least in part, of aluminum.
  • the second radius of curvature is greater than the first radius of curvature.
  • a ratio of the first radius of curvature to the second radius of curvature is equal to, or between, .05 and .55.
  • a gas turbine engine has at least one of a fan section, a compressor section and a turbine section.
  • the at least one fan section, compressor section and turbine section has rotating blades and a guide vane mounted adjacent to the rotating blades.
  • the guide vane has an airfoil and a platform at least at one end of the airfoil.
  • a compound fillet blends the airfoil into the platform.
  • the compound fillet has a first portion at a first radius of curvature adjacent to the platform.
  • a second portion at a second radius of curvature blends the first portion into the airfoil.
  • the second radius of curvature is different than the first radius of curvature.
  • the guide vane is in the fan section.
  • the airfoil has a platform at each of two ends.
  • the compound fillet blends the airfoil into a platform that will be a radially inward platform when the guide vane is mounted in a gas turbine engine.
  • the airfoil and platform are formed, at least in part, of aluminum.
  • the second radius of curvature is greater than the first radius of curvature.
  • a ratio of the first radius of curvature to the second radius of curvature is equal to, or between, .05 and .55.
  • Figure 1 schematically shows a gas turbine engine.
  • Figure 2 shows a fan guide vane
  • Figure 3 shows a method of generating a compound fillet for the Figure 2 fan guide vane.
  • 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, while 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.
  • 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
  • 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.
  • the low speed spool 30 generally includes an inner shaft 40 that interconnects a fan 42, a low pressure compressor 44 and a low pressure turbine 46.
  • the inner shaft 40 is connected to the fan 42 through 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 high pressure compressor 52 and high pressure turbine 54.
  • a combustor 56 is arranged 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.
  • the turbines 46, 54 rotationally drive the respective low speed spool 30 and high speed spool 32 in response to the expansion.
  • 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 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 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 2.5: 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.
  • a significant amount of thrust is provided by the bypass flow B due to the high bypass ratio.
  • 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.
  • 'TSFC' Thrust Specific Fuel Consumption
  • Low fan pressure ratio is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. As shown, the engine 20 does include a fan exit guide vane 200.
  • 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.
  • a gas turbine such as shown in Figure 1 has a fan diameter that is greatly increased over the prior art. As the fan diameter has increased the weight of structure within the fan would increase also, if weight reduction methods were not taken. One method that has been utilized is to use aluminum, rather than titanium, or other stronger materials. Further, as shown in Figure 2, there may be a solid structural guide vane 80, which may be used in place of the fan exit guide vane 200 of Figure 1. On the other hand, this application extends to hollow vanes.
  • an airfoil 82 extends between platforms 84 and 86. Fillets 88 and 90 blend the airfoil 82 into the platforms 84/86 at a radially inner end 91 and a radially outer end 93, respectively.
  • the fillets 88/90 may both be compound, and have at least two cylindrical portions.
  • the fillet 90 (which will be at an inner diameter) may be compound with the other fillet 88, a simple fillet (on a single radius of curvature).
  • the compound fillet is defined by two radii and an offset from either the airfoil (82) or platform face (84, 86).
  • the case depicted in Figure 3 is a compound fillet with airfoil offset; denoting the tertiary face used to construct the fillet is offset from the airfoil, not platform.
  • the small circular radius is positioned to be tangent to the platform surface (84,86) and tangent to the offset face (97).
  • the large circular radius is positioned to be tangent to the airfoil surface (82) and the small circular radius.
  • the combination of the adjoining arc segments from the small and large radius circles (92,94) at location (96) form the compound fillet bridging the airfoil (82) and platform surfaces (84, 86) [0042]
  • the compound could also vary along the chord of the airfoil 82.
  • the small and large radii can vary as a function of chord. This would be referred to as a variable compound fillet.
  • This compound blend significantly reduces stress in geometries where maintaining an appropriate stiffness and mass is necessary. Applicant has found it is desirable to have a large radius on the airfoil side of the blend (the second portion 94) so that the change in stiffness from a thin airfoil section to the thicker fillet is smooth and gradual. A smaller radius (forming section 92) reduces a stiffness adjacent to the platform runout so that the airfoil to fillet transition is not as abrupt, and stress at the airfoil to fillet runout is minimized.
  • the feature is particularly useful in vanes generally formed of aluminum.
  • a ratio of the radius of the first portion 92 (radius 1) to the radius of the second portion 94 (radius 2) is equal to, or between, five percent and fifty- five percent.
  • the radius 1 might be .35 in, while the radius 2 might be 2.0 in.
  • a compound radius where radius 1 is greater than radius 2 may also come within the scope of this application. Broadly, the compound radius simply requires the first and second radii be different.

Abstract

L'invention concerne une aube directrice comportant un profil aérodynamique et une plate-forme au moins à une extrémité du profil aérodynamique. Un flanc de raccordement composite fusionne le profil aérodynamique dans la plate-forme. Le flanc de raccordement composite possède une première partie à un premier rayon de courbure adjacent à la plate-forme, et une seconde partie à un second rayon de courbure fusionnant de la première partie dans le profil aérodynamique. Le second rayon de courbure est différent du premier rayon de courbure.
PCT/US2013/026553 2013-01-18 2013-02-17 Flanc de raccordement composite pour aube directrice WO2014113043A1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
US201361754096P 2013-01-18 2013-01-18
US61/754,096 2013-01-18

Publications (1)

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WO2014113043A1 true WO2014113043A1 (fr) 2014-07-24

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Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP3067518A1 (fr) * 2015-03-11 2016-09-14 Rolls-Royce Corporation Élément allongé et procédé associé de fabrication
US11118466B2 (en) 2018-10-19 2021-09-14 Pratt & Whiiney Canada Corp. Compressor stator with leading edge fillet
US11230934B2 (en) * 2017-02-07 2022-01-25 Ihi Corporation Airfoil of axial flow machine

Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4343593A (en) * 1980-01-25 1982-08-10 The United States Of America As Represented By The Secretary Of The Air Force Composite blade for turbofan engine fan
US20060275112A1 (en) * 2005-06-06 2006-12-07 General Electric Company Turbine airfoil with variable and compound fillet
EP1790831A2 (fr) * 2005-11-29 2007-05-30 General Electric Company Turboréacteur à double flux avec des aubes directrices à calage variable
EP2184442A1 (fr) * 2008-11-11 2010-05-12 ALSTOM Technology Ltd Raccord de profil d'aube
US20100284815A1 (en) * 2008-11-19 2010-11-11 Alstom Technologies Ltd. Llc Compound variable elliptical airfoil fillet

Patent Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4343593A (en) * 1980-01-25 1982-08-10 The United States Of America As Represented By The Secretary Of The Air Force Composite blade for turbofan engine fan
US20060275112A1 (en) * 2005-06-06 2006-12-07 General Electric Company Turbine airfoil with variable and compound fillet
EP1790831A2 (fr) * 2005-11-29 2007-05-30 General Electric Company Turboréacteur à double flux avec des aubes directrices à calage variable
EP2184442A1 (fr) * 2008-11-11 2010-05-12 ALSTOM Technology Ltd Raccord de profil d'aube
US20100284815A1 (en) * 2008-11-19 2010-11-11 Alstom Technologies Ltd. Llc Compound variable elliptical airfoil fillet

Cited By (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP3067518A1 (fr) * 2015-03-11 2016-09-14 Rolls-Royce Corporation Élément allongé et procédé associé de fabrication
US10309241B2 (en) 2015-03-11 2019-06-04 Rolls-Royce Corporation Compound fillet varying chordwise and method to manufacture
US11230934B2 (en) * 2017-02-07 2022-01-25 Ihi Corporation Airfoil of axial flow machine
US11118466B2 (en) 2018-10-19 2021-09-14 Pratt & Whiiney Canada Corp. Compressor stator with leading edge fillet

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