EP2659112A1 - Moteur à turbine à gaz et système d'ailettes à cambrure variable - Google Patents

Moteur à turbine à gaz et système d'ailettes à cambrure variable

Info

Publication number
EP2659112A1
EP2659112A1 EP11853198.7A EP11853198A EP2659112A1 EP 2659112 A1 EP2659112 A1 EP 2659112A1 EP 11853198 A EP11853198 A EP 11853198A EP 2659112 A1 EP2659112 A1 EP 2659112A1
Authority
EP
European Patent Office
Prior art keywords
airfoil
crown
gas turbine
turbine engine
groove
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
EP11853198.7A
Other languages
German (de)
English (en)
Other versions
EP2659112A4 (fr
EP2659112B1 (fr
Inventor
Robert A. Ress, Jr.
James Morton
Dan MOLNAR
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.)
Rolls Royce North American Technologies Inc
Original Assignee
Rolls Royce North American Technologies Inc
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 Rolls Royce North American Technologies Inc filed Critical Rolls Royce North American Technologies Inc
Publication of EP2659112A1 publication Critical patent/EP2659112A1/fr
Publication of EP2659112A4 publication Critical patent/EP2659112A4/fr
Application granted granted Critical
Publication of EP2659112B1 publication Critical patent/EP2659112B1/fr
Active legal-status Critical Current
Anticipated expiration legal-status Critical

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
    • F01D17/00—Regulating or controlling by varying flow
    • F01D17/10—Final actuators
    • F01D17/12—Final actuators arranged in stator parts
    • F01D17/14—Final actuators arranged in stator parts varying effective cross-sectional area of nozzles or guide conduits
    • F01D17/16—Final actuators arranged in stator parts varying effective cross-sectional area of nozzles or guide conduits by means of nozzle vanes
    • F01D17/162—Final actuators arranged in stator parts varying effective cross-sectional area of nozzles or guide conduits by means of nozzle vanes for axial flow, i.e. the vanes turning around axes which are essentially perpendicular to the rotor centre line
    • 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
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/005—Sealing means between non relatively rotating elements
    • 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
    • F01D5/146—Shape, i.e. outer, aerodynamic form of blades with tandem configuration, split blades or slotted blades
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/08—Sealings
    • F04D29/083—Sealings especially adapted for elastic fluid pumps
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/40—Casings; Connections of working fluid
    • F04D29/52—Casings; Connections of working fluid for axial pumps
    • F04D29/54—Fluid-guiding means, e.g. diffusers
    • F04D29/541—Specially adapted for elastic fluid pumps
    • F04D29/542—Bladed diffusers
    • F04D29/544—Blade shapes
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/40—Casings; Connections of working fluid
    • F04D29/52—Casings; Connections of working fluid for axial pumps
    • F04D29/54—Fluid-guiding means, e.g. diffusers
    • F04D29/56—Fluid-guiding means, e.g. diffusers adjustable
    • F04D29/563—Fluid-guiding means, e.g. diffusers adjustable specially adapted for elastic fluid pumps

Definitions

  • the present invention relates to gas turbine engines, and more particularly, to gas turbine engines with variable camber vane systems.
  • One embodiment of the present invention is a unique variable camber vane system for a gas turbine engine. Another embodiment is a unique gas turbine engine. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for gas turbine engines and variable camber vane systems. Further embodiments, forms, features, aspects, benefits, and advantages of the present application will become apparent from the description and figures provided herewith.
  • FIG. 1 schematically depicts some aspects of a non-limiting example of a gas turbine engine in accordance with an embodiment of the present invention.
  • FIG. 2 schematically depicts some aspects of a non-limiting example of a fan system for a gas turbine engine in accordance with an embodiment of the present invention.
  • FIG. 3 depicts some aspects of a non-limiting example of a variable camber guide vane system in accordance with an embodiment of the present invention.
  • FIG. 4 depicts some aspects of the variable camber guide vane system of FIG. 3.
  • FIG. 5 depicts some aspects of a non-limiting example of a seal strip in accordance with an embodiment of the present invention.
  • gas turbine engine 10 is an aircraft propulsion power plant.
  • gas turbine engine 10 may be a land-based or marine engine.
  • gas turbine engine 10 is a multi-spool turbofan engine. In other words,
  • gas turbine engine 10 may be a single or multi-spool turbofan, turboshaft, turbojet, turboprop gas turbine or combined cycle engine.
  • Gas turbine engine 10 includes a fan system 12, a compressor system 14, a diffuser 16, a combustion system 18 and a turbine system 20.
  • Compressor system 14 is in fluid communication with fan system 12.
  • Diffuser 16 is in fluid communication with compressor system 14.
  • Combustion system 18 is flu idly disposed between compressor system 14 and turbine system 20.
  • Fan system 2 includes a fan rotor system 22.
  • fan rotor system 22 includes one or more rotors (not shown) that are powered by turbine system 20.
  • Compressor system 14 includes a compressor rotor system 24.
  • compressor rotor system 24 includes one or more rotors (not shown) that are powered by turbine system 20.
  • Turbine system 20 includes a turbine rotor system 26.
  • turbine rotor system 26 includes one or more rotors (not shown) operative to drive fan rotor system 22 and compressor rotor system 24.
  • Turbine rotor system 26 is driving coupled to compressor rotor system 24 and fan rotor system 22 via a shafting system 28.
  • shafting system 28 includes a plurality of shafts that may rotate at the same or different speeds and directions. In some embodiments, only a single shaft may be employed.
  • combustion system 1 8 includes a combustion liner (not shown) that contains a continuous combustion process.
  • combustion system 18 may take other forms, and may be, for example, a wave rotor combustion system, a rotary valve combustion system, or a siinger combustion system, and may employ deflagration and/or detonation combustion processes.
  • the hot gases exiting combustor 1 8 are directed into turbine system 20, which extracts energy in the form of mechanical shaft power to drive fan system 12 and compressor system 14 via shafting system 28.
  • the hot gases exiting turbine system 20 are directed into a nozzle (not shown), and provide a component of the thrust output by gas turbine engine 10.
  • Fan system 12 includes a variable guide vane system 40 having a variable inlet guide vane stage 42 and a variable outlet guide vane stage 44 disposed on either side of a rotating fan stage 46.
  • Variable inlet guide vane stage 42 is operative to guide air into rotating fan stage 46, and to selectively vary the incidence angle of the air flow entering rotating fan stage 46.
  • Variable outlet guide vane stage 44 is operative to guide air exiting rotating fan stage 46, and to selectively vary the incidence angle of the air flow exiting rotating fan stage 46.
  • Variable inlet guide vane stage 42 and variable outlet guide vane stage 44 are actuated by an actuation system (not shown).
  • variable guide vane system 40 may also or alternatively be employed as part of compressor system 14.
  • variable guide vane system 40 includes both variable inlet and outlet guide vane stages, other embodiments may include only a variable inlet guide vane stage or a variable outlet guide vane stage.
  • variable inlet guide vane stage 42 in accordance with an embodiment of the present invention is illustrated. It will be understood that some embodiments of variable outlet guide vane stage 44 may be similar to variable inlet guide vane stage 42, and hence, the following description of variable inlet guide vane stage 42 is also applicable to aspects of some embodiments of variable outlet guide vane stage 44.
  • Variable inlet guide vane stage 42 includes an outer band 50, an inner band 52 and plurality of vanes 54. Outer band 50 defines an outer flowpath wall of variable inlet guide vane stage 42.
  • Inner band 52 defines an inner flowpath wall of variable inlet guide vane stage 42.
  • Vanes 54 are airfoils that extend between outer band 50 and inner band 52, and are spaced apart circumferentially. in one form, vanes 54 extend in the radial direction between outer band 50 and inner band 52. In other embodiments, vanes 54 may extend between outer band 50 and inner band 52 at other angles.
  • Each vane 54 includes an airfoil portion 56 and an airfoil portion 58.
  • Airfoil portion 56 extends between a tip portion 60 and a root portion 62.
  • airfoil portion 56 includes the trailing edge 64 of vane 54.
  • airfoil portion 56 may be formed with a leading edge of vane 54 instead of trailing edge 64, e.g., for use in variable outlet guide vane 44.
  • Airfoil portion 58 extends between a tip portion 66 and a root portion 68.
  • airfoil portion 58 includes the leading edge 70 of vane 54.
  • airfoil portion 58 may be formed with a trailing edge instead of leading edge 70, e.g., for use in variable outlet guide vane 44.
  • airfoil portion 56 is fixed, i.e., stationary.
  • airfoil portion 56 may be movable, e.g., pivotable about an axis so as to be able to vary the angle of the trailing edge of vane 54.
  • airfoil portion 58 is variable, being configured to pivot about a pivot axis 72 with respect to airfoil portion 56, to provide a variable camber for vane 54.
  • airfoil portion 58 may be fixed.
  • airfoil portion 58 is coupled to an actuation system (not shown) that is operative to selectively position airfoil portion 58 at a desired incidence angle.
  • airfoii portion 56 may also or alternatively be coupled to an actuation system (not shown) that is operative to selectively position airfoil portion 56 at a desired incidence angle.
  • Outer band 50 includes a plurality of spaced apart openings 78.
  • Inner band 52 includes a plurality of spaced apart openings 80. Openings 78 and 80 are operative to receive pivot shafts 74 and 76, respectively, and retain airfoil portions 58 in the engine axial, circumferential and radial direction.
  • pivot shafts 74 and 76 retain airfoil portion 58 in outer band 50 and inner band 52 via anti-friction bushings 82 and 84.
  • Anti-friction bushings 82 and 84 are operative to provide bearing surfaces for pivot shafts 74 and 76. Other embodiments may not include anti-friction bushings 82 and 84.
  • Airfoil portion 58 is operative to rotate in rotation directions 86 about pivot axis 72.
  • Vane 54 has a pressure side 90 and a suction side 92, wherein the pressure on pressure side 90 exceeds that of suction side 92.
  • the pressure differential between pressure side 90 and suction side 92 may vary, e.g., depending upon vane 54 camber and engine operating conditions.
  • the pressure differential between pressure side 90 and suction side 92 provides an impetus to flow from pressure side 90 to suction side 92, e.g., between airfoil portion 56 and airfoil portion 58.
  • vanes 54 include a sealing arrangement 94 operative to seal between airfoil portion 56 and airfoil portion 58.
  • Sealing arrangement 94 includes a seal strip 96 arranged to seal against fluid flow between airfoil portion 56 and airfoil portion 58 during the operation of engine 10, and to accommodate movement of one or both of airfoil portions 56 and 58, e.g., rotation of airfoil portion 58 about pivot axis 72, while sealing against fluid flow.
  • seal strip 96 is a rigid structure that does not substantially deform in use or installation. In other embodiments, seal strip 96 may be a flexible structure. In one form, seal strip 96 is formed of a polymeric material, such as Vespel (commercially available from DuPont Engineering Polymers, located in Newark, Delaware, U.S.A.) and/or Torlon polyamide-imide (commercially available from Solvay Advanced
  • seal strip 96 may be formed of other materials.
  • seal strip 96 is disposed in a groove 98.
  • groove 98 is disposed in a face 1 00 of airfoil portion 56 that faces airfoil portion 58.
  • seal strip 96, groove 98 and face 100 extend between tip portion 60 and root portion 62 of airfoil portion 56.
  • seal strip 96, groove 98 and/or face 100 may extend only partially between tip portion 60 and root portion 62.
  • Face 100 is formed with a radius 02 centered on pivot axis 72. In one form, face 100 is formed integrally with airfoil portion 56.
  • face 100 may be formed separately and affixed to airfoil portion 56.
  • seal strip 96 is partially installed in groove 98, that is, leaving a portion 108 of seal strip 96 extending beyond face 100 of airfoil portion 56.
  • Seal strip 96 has a width 104 greater than a width 106 of groove 98, and is installed into groove 98 with an interference fit, e.g., 0.001 -0.002 inch. The amount of interference may vary with the needs of the application.
  • Airfoil portion 58 includes a crown 1 10 facing face 00 of airfoil portion 56. In one form, crown 1 10 is formed integrally with airfoil portion 58.
  • crown 1 10 may be formed separately and affixed to airfoil portion 58. Crown 1 10 is formed with a radius 1 12 centered on pivot axis 72. In one form, crown 1 10 extends between tip portion 66 and root portion 68 of airfoil portion 58, and is positioned opposite groove 98. In other embodiments, crown 110 may extend only partially between tip portion 66 and root portion 68. In one form, face 100 of airfoil portion 56 is concave, and is operative to receive therein crown 1 10 opposite groove 98 in a nested arrangement. In other embodiments, face 100 may be convex.
  • crown 1 10 of airfoil portion 58 is convex, and is operative to be received into face 100 in a nested arrangement.
  • crown 1 10 may be convex, e.g., an inverted crown.
  • the depicted embodiment includes groove 98 and seal strip 96 being located in face 100, in other embodiments, groove 98 and seal strip 96 may be located in crown 1 10.
  • Seal strip 96 includes a rubbing surface 1 14.
  • rubbing surface 1 14 is disposed opposite radius 1 12 of crown 1 10, and is operative to contact and seal against radius 1 12 of crown 110 of airfoil portion 58.
  • rubbing surface 1 14 may rub against crown 1 10, e.g., until wear of seal strip 96 resulting from rotation of airfoil portion 58 reduces or eliminates contact between seal strip 96 and crown 1 10.
  • rubbing surface 1 14 may be configured to be in close proximity to crown 1 10, but without any rubbing contact.
  • seal strip 96 may be installed in crown 1 10, and rubbing surface 1 14 may be configured to seal against face 100.
  • Rubbing surface 1 14 is preformed prior to installation into airfoil portion 56 , e.g., machined.
  • rubbing surface 1 14 is configured as a radius 1 16 centered about pivot axis 72, e.g., the same radius as radius 1 12 of crown 1 0.
  • radius 1 16 may be the same radius as radius 102 of face 100 or any other radius suitable for the application, in still other embodiments, other shapes for rubbing surface 1 14 may be employed.
  • rubbing surface 1 14 is concave. In other embodiments, rubbing surface 1 14 may take other forms, and may be, for example, convex.
  • Embodiments of the present invention include a variable camber vane system for a gas turbine engine, comprising: a first airfoil portion having a first tip portion, a first root portion, a face extending at least partially between the first tip portion and the first root portion, and a groove in the face extending at least partially between the first tip portion and the first root portion, wherein the groove has a groove width; a second airfoil portion arranged to rotate with respect to the first airfoil portion about a pivot axis, wherein the second airfoil portion includes a second tip portion; a second root portion; and a crown extending at least partially between the second tip portion and the second root portion, wherein the crown includes a crown radius centered about the pivot axis and positioned opposite the groove; and a seal strip having a seal width greater than the groove width and a rubbing surface opposite the crown radius, wherein the seal strip is at least partially disposed in the groove with an interference fit; and wherein the seal strip is arranged to seal against fluid
  • the seal strip is a rigid structure.
  • the rubbing surface has a rubbing surface radius the same as the crown radius.
  • the crown is formed integrally with the second airfoil portion.
  • the face is formed integrally with the first airfoil portion.
  • the face is concave and operative to receive the crown therein.
  • the first airfoil portion is stationary.
  • first airfoil portion and the second airfoil portion form at least part of an inlet guide vane having a fixed leading edge and a variable trailing edge; wherein the first airfoil portion includes the leading edge; and wherein the second airfoil portion includes the trailing edge.
  • first airfoil portion and the second airfoil portion form at least part of an outlet guide vane having a variable leading edge and a fixed trailing edge; wherein the first airfoil portion includes the leading edge; and wherein the second airfoil portion includes the trailing edge.
  • Embodiments of the present invention include a gas turbine engine, comprising: at least one of a fan and a compressor having a variable camber vane system, the variable camber vane system including: at least two airfoil portions adapted to vary a camber of the variable camber vane system , wherein a first of the airfoil portions includes a groove and a second of the airfoil portions includes a crown having a crown radius; and a seal strip at least partially disposed in the groove with an interference fit, wherein the seal strip includes a rubbing surface opposite the crown radius and operative to seal against fluid flow between the first of the airfoil portions and the second of the airfoil portions.
  • the rubbing surface contacts the crown at the crown radius.
  • the seal strip is formed of a polymer material.
  • the seal strip is formed of at least one of Vespel and
  • the at least two airfoil portions form an inlet guide vane.
  • the at least two airfoil portions form an outlet guide vane.
  • Embodiments include a gas turbine engine, comprising: at least one of a fan and a compressor having a variable camber vane system, the variable camber vane system including: at least two airfoil portions adapted to vary a camber of the variable camber vane system, wherein a first of the airfoil portions includes a groove; and wherein a second of the airfoil portions includes a crown having a crown radius; and a seal strip disposed in the groove; wherein the seal strip has a rubbing surface radius preformed thereon and configured for sealing engagement with the crown.
  • the seal strip is a rigid structure formed of a polymer.
  • the crown radius is convex, and the rubbing surface radius is concave.
  • the seal strip is fitted in the groove with an interference fit.
  • the crown is nested within the first of the airfoil portions opposite the groove.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Physics & Mathematics (AREA)
  • Geometry (AREA)
  • Fluid Mechanics (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)

Abstract

Selon un mode de réalisation, la présente invention concerne un système unique d'ailettes à cambrure variable pour un moteur à turbine à gaz. Selon un autre mode de réalisation, l'invention concerne un moteur à turbine à gaz unique. D'autres modes de réalisation comprennent des appareils, des systèmes, des dispositifs, du matériel, des procédés, et des combinaisons pour moteurs à turbine à gaz et des systèmes d'ailettes à cambrure variable. D'autres modes de réalisation, formes, caractéristiques, aspects, bénéfices, et avantages sont décrits dans la description et les dessins qui y sont associés.
EP11853198.7A 2010-12-27 2011-12-27 Moteur à turbine à gaz et système d'ailettes à cambrure variable Active EP2659112B1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
US12/978,843 US20120163960A1 (en) 2010-12-27 2010-12-27 Gas turbine engine and variable camber vane system
PCT/US2011/067393 WO2012092277A1 (fr) 2010-12-27 2011-12-27 Moteur à turbine à gaz et système d'ailettes à cambrure variable

Publications (3)

Publication Number Publication Date
EP2659112A1 true EP2659112A1 (fr) 2013-11-06
EP2659112A4 EP2659112A4 (fr) 2018-03-07
EP2659112B1 EP2659112B1 (fr) 2020-10-07

Family

ID=46317012

Family Applications (1)

Application Number Title Priority Date Filing Date
EP11853198.7A Active EP2659112B1 (fr) 2010-12-27 2011-12-27 Moteur à turbine à gaz et système d'ailettes à cambrure variable

Country Status (4)

Country Link
US (1) US20120163960A1 (fr)
EP (1) EP2659112B1 (fr)
CA (1) CA2822965C (fr)
WO (1) WO2012092277A1 (fr)

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RU217010U1 (ru) * 2022-12-30 2023-03-14 Федеральное государственное автономное образовательное учреждение высшего образования "Омский государственный технический университет" Устройство входного направляющего аппарата центробежного компрессора
US12292056B2 (en) 2023-03-17 2025-05-06 Rolls-Royce North American Technologies Inc. Segmented variable fan outlet guide vane with gear assembly
US12398648B2 (en) 2023-03-17 2025-08-26 Rolls-Royce North American Technologies Inc. Segmented variable fan outlet guide vane with cam assembly and pass through actuation mechanisms
US12398655B2 (en) 2023-03-17 2025-08-26 Rolls-Royce North American Technologies Inc. Segmented variable fan outlet guide vane with segment interface components
US12428974B2 (en) 2023-03-17 2025-09-30 Rolls-Royce North American Technologies Inc. Segmented variable fan outlet guide vane with unique actuation mechanisms
US12480422B2 (en) 2023-03-17 2025-11-25 Rolls-Royce North American Technologies Inc. Segmented variable fan outlet guide vane with pass through actuation mechanisms

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RU217010U1 (ru) * 2022-12-30 2023-03-14 Федеральное государственное автономное образовательное учреждение высшего образования "Омский государственный технический университет" Устройство входного направляющего аппарата центробежного компрессора
US12292056B2 (en) 2023-03-17 2025-05-06 Rolls-Royce North American Technologies Inc. Segmented variable fan outlet guide vane with gear assembly
US12398648B2 (en) 2023-03-17 2025-08-26 Rolls-Royce North American Technologies Inc. Segmented variable fan outlet guide vane with cam assembly and pass through actuation mechanisms
US12398655B2 (en) 2023-03-17 2025-08-26 Rolls-Royce North American Technologies Inc. Segmented variable fan outlet guide vane with segment interface components
US12428974B2 (en) 2023-03-17 2025-09-30 Rolls-Royce North American Technologies Inc. Segmented variable fan outlet guide vane with unique actuation mechanisms
US12480422B2 (en) 2023-03-17 2025-11-25 Rolls-Royce North American Technologies Inc. Segmented variable fan outlet guide vane with pass through actuation mechanisms

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US20120163960A1 (en) 2012-06-28
WO2012092277A1 (fr) 2012-07-05
CA2822965A1 (fr) 2012-07-05
EP2659112A4 (fr) 2018-03-07
CA2822965C (fr) 2020-02-11
EP2659112B1 (fr) 2020-10-07

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