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WO2008060195A1 - Ensemble d'aubes configurées pour faire tourner un écoulement dans un moteur de turbine à gaz, un composant de stator comprenant l'ensemble d'aubes, une turbine à gaz et un moteur à réaction d'avion - Google Patents

Ensemble d'aubes configurées pour faire tourner un écoulement dans un moteur de turbine à gaz, un composant de stator comprenant l'ensemble d'aubes, une turbine à gaz et un moteur à réaction d'avion Download PDF

Info

Publication number
WO2008060195A1
WO2008060195A1 PCT/SE2006/001292 SE2006001292W WO2008060195A1 WO 2008060195 A1 WO2008060195 A1 WO 2008060195A1 SE 2006001292 W SE2006001292 W SE 2006001292W WO 2008060195 A1 WO2008060195 A1 WO 2008060195A1
Authority
WO
WIPO (PCT)
Prior art keywords
guide vane
vane
assembly according
additional guide
main guide
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.)
Ceased
Application number
PCT/SE2006/001292
Other languages
English (en)
Inventor
Stéphane BARALON
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.)
GKN Aerospace Sweden AB
Original Assignee
Volvo Aero AB
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 Volvo Aero AB filed Critical Volvo Aero AB
Priority to EP06813013.7A priority Critical patent/EP2092163A4/fr
Priority to PCT/SE2006/001292 priority patent/WO2008060195A1/fr
Priority to US12/514,800 priority patent/US20100158684A1/en
Publication of WO2008060195A1 publication Critical patent/WO2008060195A1/fr
Anticipated expiration legal-status Critical
Ceased legal-status Critical Current

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
    • 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
    • F01D5/141—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
    • 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
    • F01D9/06—Fluid supply conduits to nozzles or the like
    • F01D9/065—Fluid supply or removal conduits traversing the working fluid flow, e.g. for lubrication-, cooling-, or sealing fluids
    • 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

  • Vane assembly configured for turning a flow in a gas turbine engine, a stator component comprising the vane assembly, a gas turbine and an aircraft jet engine
  • the present invention relates to a vane assembly configured for turning a flow in a gas turbine engine.
  • the invention is also related to a stator component comprising the vane assembly.
  • Jet engine is meant to include various types of engines, which admit air at relatively low velocity, heat it by combustion and shoot it out at a much higher velocity.
  • Accommodated within the term jet engine are, for example, turbojet engines and turbo-fan engines.
  • stationary guide vane assemblies are used to turn the flow from one angle to another.
  • the stationary guide vane assembly may be applied in a stator component of a turbo-fan engine at a fan outlet, in a Turbine Exhaust Case (TEC) and even in an InterMediate Case (IMC) .
  • TEC Turbine Exhaust Case
  • IMC InterMediate Case
  • Typical Fan OGV configurations which turn the flow by about 40-50 degrees, with structural loads but no servicing through may result in 58 vanes in the bypass duct.
  • Typical TEC have about 14 vanes to turn the flow by about 30 degrees. TEC are however much thicker and have a max thickness to chord ratio of about 14 %.
  • One purpose of the invention is to achieve a vane assembly, which creates conditions for a large amount of flow turning while minimizing the upstream influence of the vane pressure field. Further, the axial extension of the vane assembly should be kept to a minimum.
  • a vane assembly configured for turning a flow in a gas turbine engine comprising a stationary main guide vane and an additional guide vane, wherein a leading edge of the additional guide vane is positioned upstream of a leading edge of the main guide vane and wherein the additional guide vane extends a distance along the main guide vane towards a trailing edge of the main guide vane forming a passageway between the additional guide vane and the main guide vane.
  • the main guide vane is preferably configured to turn an incoming flow and the additional guide vane is configured to assist the main guide vane in turning the incoming flow.
  • the additional stationary guide vane is arranged in the vicinity of the leading edge of the main guide vane and the additional guide vane is aerodynamicalIy coupled to the main guide vane.
  • This design creates conditions for achieving a larger flow turning with a smaller number of main guide vanes
  • struts For example, for a TEC, about 50% more flow turning may be achieved with about 30% less main guide vanes . Further, a less complicated manufacturing (for example casting and forging) may be used compared to classical high lift devices .
  • this type of vane assembly may lead to more loading on upstream stages, a shorter engine length, reduced engine weigth and reduced part count .
  • the additional guide vane is positioned relative to the main guide vane so that the passageway becomes more narrow in a downstream direction.
  • the additional guide vane is positioned relative to the main guide vane so that the passageway continuously narrows down from an upstream opening of the passageway to a downstream opening.
  • the additional guide vane is positioned relative to the main guide vane so that the passageway is shaped as a nozzle.
  • the additional guide vane has an at least partly curved shape.
  • a first, upstream portion of a suction side of the additional guide vane extending from a leading edge of the additional guide vane is substantially straight and that a second, downstream portion of the suction side of the additional guide vane is curved.
  • the first, upstream portion of the suction side of the additional guide vane extends over at least 50% of the chord of the additional guide vane.
  • a leading edge of the additional guide vane has an elliptic cross sectional shape. In this way, upstream forcing may be reduced.
  • FIG 1 is a schematic side view of the engine cut along a plane in parallel with the rotational axis of the engine
  • FIG 2 is a perspective view of a stator component comprising an inventive guide vane assembly
  • FIG 3 shows the vane assembly of figure 2 in cross section.
  • the invention will below be described for a high bypass ratio aircraft engine 100, see figure 1.
  • the engine 100 comprises an outer housing 102, an inner housing 104 and an intermediate housing 106 which is concentric to the first two housings and divides the gap between them into an inner primary gas channel 108 for the compression of the propulsion gases and a secondary channel 110 in which the engine bypass circulates.
  • each of the gas channels 108,110 is annular in a cross section perpendicular to an axial direction 112 of the engine 100.
  • the engine 100 comprises a fan 114 which receives ambient air 115, a booster or low pressure compressor
  • LPC low pressure compressor
  • HPC high pressure compressor
  • HPT high pressure turbine
  • LPT low pressure turbine
  • a first or high pressure shaft joins the high pressure turbine 122 to the high pressure compressor 118 to substantially form a first or high pressure rotor.
  • a second or low pressure shaft joins the low pressure turbine 124 to the low pressure compressor 116 to substantially form a second or low pressure rotor.
  • the high pressure compressor 118, combustor 120 and high pressure turbine 122 are collectively referred to as a core engine.
  • the second or low pressure shaft is at least in part rotatably disposed co-axially with and radially inwardly of the first or high pressure rotor.
  • the housings 102,104,106 are supported by structures 126 which connect the housings by radial arms. These arms are generally known as struts.
  • the struts must be sufficiently resistant to provide this support and not to break or buckle in the event of a fan blade coming loose and colliding with them.
  • the struts are designed for transmission of loads in the engine.
  • the struts are hollow in order to house service components such as means for the intake and outtake of oil and/or air, for housing instruments, such as electrical and metallic cables for transfer of information concerning measured pressure and/or temperature, a drive shaft for a start engine etc.
  • the struts can also be used to conduct a coolant.
  • the compressor structure 126 connecting the intermediate housing 106 and the inner housing 104 is conventionally referred to as an Intermediate Case
  • the compressor structure 126 is designed for guiding the gas flow from the low pressure compressor 116 radially inwards toward to the high pressure compressor 118 inlet.
  • the compressor structure 126 connecting the intermediate housing 106 and the inner housing 102 comprises a plurality of radial struts 208 see figure 2 and 3, at mutual distances in the circumferential direction of the compressor structure. These struts 208 are structural parts, designed for transmission of both axial and radial loads and at least some are hollow in order to house service components.
  • Figure 2 shows a perspective view of a stator component in the form of the compressor structure 126.
  • the compressor structure 126 comprises an inner ring 202, an outer ring 204 encompassing the inner ring 202, and a plurality of vane assemblies 206 extending radially between the inner ring 202 and the outer ring 204.
  • the vane assemblies 206 are circumferentially spaced and rigidly connected to the rings 202,204.
  • Each of the vane assemblies 206 comprises the strut 208 and an additional guide vane 210, see also figure 3.
  • FIG. 3 shows one vane assembly 206 of figure 2 in an enlarged cross section view.
  • the vane assembly 206 comprises a stationary main guide vane 208 (the strut) and the additional stationary guide vane 210.
  • the main guide vane 208 is structurally bearing.
  • the main guide vane 208 has a first sidewall 306 and a second sidewall 308 that are connected at a leading edge 310 and a trailing edge 312.
  • the main guide vane 208 is configured to turn an incoming flow.
  • the main guide vane 208 has an at least partly curved shape.
  • the first side wall 306 of the main guide vane 208 is convex and defines a suction side.
  • the second side wall 308 of the main guide vane 208 is concave and defines a pressure side.
  • the additional guide vane 210 has a first sidewall 314 and a second sidewall 316 that are connected at a leading edge 318 and a trailing edge 320.
  • the additional guide vane 210 is configured to turn an incoming flow.
  • the additional guide vane 210 has an at least partly curved shape.
  • the first side wall 314 of the additional guide vane 210 is convex and defines a suction side.
  • the second side wall 316 of the additional guide vane 210 is concave and defines a pressure side.
  • the magnitude of the turning of the gas flow in the stator component 126 depends on several parameters.
  • the main guide vane 208 has a cambered airfoil shape, see figure 3.
  • the main guide vanes are designed with a sufficient curvature for a substantial turning of the gas flow.
  • the main vane 208 is not only structural, but also has an aerodynamic function. More specifically, the direction of a mean camber line M at the leading edge 310 is inclined with an angle in relation to the direction of the mean camber line M at the trailing edge 312 corresponding to the desired turning angle.
  • the direction of the mean camber line M at the leading edge 310 of the cambered main vane 208 is therefore inclined with at least 20°, suitably at least 30°, especially at least 40°, and preferably at least 50° in relation to the direction of the mean camber line M at the trailing edge 312.
  • chord is defined as the distance between the leading edge 310 and the trailing edge 312 of the main vane 208 along the chord line C, see figure 3.
  • chord line C is defined as a straight line connecting the leading edge 310 and the trailing edge 312.
  • the thickness of the main vane 208 is defined as the maximum distance between the two opposing strut surfaces 306,308 in a direction perpendicular to a mean chamber line M.
  • the mean camber line M is defined as the locus of points halfway between the upper and lower surfaces 306,308 of the main vane as measured perpendicular to the mean camber line itself.
  • the camber A is defined as the maximum distance between the mean chamber line M and the chord line C measured perpendicular to the chord line.
  • the main guide vane 208 has the shape of an airfoil in cross section. In other words, the mean camber line M is curved.
  • the maximum thickness to chord ratio is another measure for the gas flow turning capacity of the struts.
  • the maximum thickness is preferably less than 20%, especially less than 15% and more specifically about 10% of the chord according to the example shown in the drawings .
  • the leading edge 318 of the additional guide vane 210 is positioned upstream of the leading edge 310 of the main guide vane 208.
  • the additional guide vane 210 extends a distance along the main guide vane 208 forming a passageway 322 between the additional guide vane 210 and the main guide vane 208.
  • the additional guide vane 210 at least partly overlaps the main guide vane 208.
  • the additional guide vane 210 is positioned relative to the main guide vane 208 so that the passageway 322 becomes more narrow in a downstream direction.
  • the additional guide vane 210 is positioned relative to the main guide vane 208 so that the passageway 322 continuously narrows down from an upstream opening 324 of the passageway to a downstream opening 326.
  • the leading edge 318 of the additional guide vane 210 is at a greater distance from a periphery of the main guide vane 208 than a trailing edge 320 of the additional guide vane 210 is from a periphery of the main guide vane 208.
  • the additional guide vane 210 is positioned relative to the main guide vane 208 so that the passageway 322 is shaped as a nozzle.
  • the trailing edge 320 of the additional guide vane 210 is arranged upstream of the trailing edge 312 of the main guide vane 208. More specifically, the trailing edge 326 of the additional guide vane 210 is arranged upstream of a point halfway of the chord of the main guide vane 208.
  • a pressure side of the additional guide vane 210 faces a suction side of the main guide vane 208, wherein the passageway 322 is defined therebetween.
  • the additional guide vane 210 has the shape of an airfoil in cross section. In other words, the mean camber line is curved.
  • the additional guide vane 210 has a substantially smaller thickness to chord ratio than the main guide vane 208.
  • the vane assembly 206 is designed for turning a swirling gas flow.
  • the swirling gas normally flows with an angle of 40-60° relative to the axial direction 112 of the engine. In this case the turning of the gas flow is in the combined axial-tangential and axial-radial directions .
  • a first, upstream portion 328 of a suction side 314 of the additional guide vane 210 extending from a leading edge 318 is substantially straight and a second, downstream portion 330 of the suction side 314 of the additional guide vane is curved. More specifically, the first, upstream portion 328 of the suction side of the additional guide vane extends over at least 50% of the chord of the additional guide vane. Especially, the first, upstream portion 328 of the suction side of the additional guide vane extends over about 70% of the chord of the additional guide vane.
  • a concave part of the additional guide vane 210 defines one side of the passageway 322. More specifically, the concave part of the additional guide vane 210 is arranged along a convex part of the main guide vane 208, wherein the passageway 322 is defined therebetween .
  • the vane assembly 206 is adapted to turn an incoming flow with a flow angle of about -45 degrees to an outgoing flow with a flow angle of about -5 degrees. It is expected that this type of vane assembly may turn a flow with about 60 degrees.
  • the struts are often hollow in order to house service components such as means for the intake and outtake of oil and/or air, for housing instruments, such as electrical and metallic cables for transfer of information concerning measured pressure and/or temperature etc.
  • the struts normally have a symmetric airfoil shape in cross section in order to effect the gas flow as little as possible.
  • the servicing requirement usually governs the number of struts required.
  • the main guide vane 208 described above has an at least partly curved shape.
  • the mean camber line M is curved.
  • the main guide vane may be symmetrical in that its mean camber line may be straight.

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

Abstract

L'invention concerne un ensemble d'aubes configurées pour faire tourner un écoulement dans un moteur de turbine à gaz comprenant une aube de guidage principale stationnaire (208) et une aube de guidage supplémentaire (210). Un bord d'attaque (318) de l'aube de guidage supplémentaire (210) est positionné en amont d'un bord d'attaque (70) de l'aube de guidage principale (208); l'aube de guidage supplémentaire (210) s'étend sur une distance le long de l'aube de guidage principale (208) vers un bord de fuite (312) de l'aube de guidage principale (208) en formant un passage (322) entre l'aube de guidage supplémentaire (210) et l'aube de guidage principale (208).
PCT/SE2006/001292 2006-11-14 2006-11-14 Ensemble d'aubes configurées pour faire tourner un écoulement dans un moteur de turbine à gaz, un composant de stator comprenant l'ensemble d'aubes, une turbine à gaz et un moteur à réaction d'avion Ceased WO2008060195A1 (fr)

Priority Applications (3)

Application Number Priority Date Filing Date Title
EP06813013.7A EP2092163A4 (fr) 2006-11-14 2006-11-14 Ensemble d'aubes configurées pour faire tourner un écoulement dans un moteur de turbine à gaz, un composant de stator comprenant l'ensemble d'aubes, une turbine à gaz et un moteur à réaction d'avion
PCT/SE2006/001292 WO2008060195A1 (fr) 2006-11-14 2006-11-14 Ensemble d'aubes configurées pour faire tourner un écoulement dans un moteur de turbine à gaz, un composant de stator comprenant l'ensemble d'aubes, une turbine à gaz et un moteur à réaction d'avion
US12/514,800 US20100158684A1 (en) 2006-11-14 2006-11-14 Vane assembly configured for turning a flow in a gas turbine engine, a stator component comprising the vane assembly, a gas turbine and an aircraft jet engine

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
PCT/SE2006/001292 WO2008060195A1 (fr) 2006-11-14 2006-11-14 Ensemble d'aubes configurées pour faire tourner un écoulement dans un moteur de turbine à gaz, un composant de stator comprenant l'ensemble d'aubes, une turbine à gaz et un moteur à réaction d'avion

Publications (1)

Publication Number Publication Date
WO2008060195A1 true WO2008060195A1 (fr) 2008-05-22

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PCT/SE2006/001292 Ceased WO2008060195A1 (fr) 2006-11-14 2006-11-14 Ensemble d'aubes configurées pour faire tourner un écoulement dans un moteur de turbine à gaz, un composant de stator comprenant l'ensemble d'aubes, une turbine à gaz et un moteur à réaction d'avion

Country Status (3)

Country Link
US (1) US20100158684A1 (fr)
EP (1) EP2092163A4 (fr)
WO (1) WO2008060195A1 (fr)

Cited By (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE102010001059A1 (de) * 2010-01-20 2011-07-21 Rolls-Royce Deutschland Ltd & Co KG, 15827 Zwischengehäuse für ein Gasturbinentriebwerk
EP2626515A1 (fr) * 2012-02-10 2013-08-14 MTU Aero Engines GmbH Agencement de groupes d'aubes en tandem
EP2746535A1 (fr) * 2012-12-20 2014-06-25 General Electric Company Conception de profil aérodynamique, à double rangée étagée, pour trame d'échappement de turbine à gaz
EP2397652A3 (fr) * 2010-06-20 2014-12-17 Honeywell International Inc. Aube de turbocompresseur à plusieurs profils aérodynamiques
WO2015108602A1 (fr) * 2013-10-29 2015-07-23 General Electric Company Ensemble entretoise de moteur d'aéronef et procédés d'assemblage de celui-ci
WO2018084902A1 (fr) * 2016-07-15 2018-05-11 General Electric Company Moteur à double flux et procédé de fonctionnement correspondant
GB2515961B (en) * 2012-04-05 2018-06-27 Snecma Stator vane formed by a set of vane parts
GB2568109A (en) * 2017-11-07 2019-05-08 Gkn Aerospace Sweden Ab Splitter vane

Families Citing this family (22)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE102010014900A1 (de) * 2010-04-14 2011-10-20 Rolls-Royce Deutschland Ltd & Co Kg Nebenstromkanal eines Turbofantriebwerkes
US9062559B2 (en) 2011-08-02 2015-06-23 Siemens Energy, Inc. Movable strut cover for exhaust diffuser
US20130170969A1 (en) * 2012-01-04 2013-07-04 General Electric Company Turbine Diffuser
US8968437B2 (en) 2012-05-02 2015-03-03 Michael J Kline Jet engine with deflector
US8657895B2 (en) 2012-05-02 2014-02-25 Michael J. Kline Jet engine deflector
JP6039059B2 (ja) * 2012-05-02 2016-12-07 ゲーコーエヌ エアロスペース スウェーデン アーベー ガスタービンエンジンの支持構造
US9593691B2 (en) 2013-07-19 2017-03-14 General Electric Company Systems and methods for directing a flow within a shroud cavity of a compressor
EP2752320B1 (fr) * 2013-09-16 2018-02-14 Weidplas GmbH Réservoir d'eau pour un véhicule automobile
WO2015142200A1 (fr) * 2014-03-18 2015-09-24 General Electric Company Diffuseur de gaz d'échappement avec entretoises principales et petites entretoises
FR3027053B1 (fr) * 2014-10-10 2019-09-13 Safran Aircraft Engines Stator de turbomachine d'aeronef
US10704418B2 (en) 2016-08-11 2020-07-07 General Electric Company Inlet assembly for an aircraft aft fan
US10252790B2 (en) 2016-08-11 2019-04-09 General Electric Company Inlet assembly for an aircraft aft fan
US10259565B2 (en) 2016-08-11 2019-04-16 General Electric Company Inlet assembly for an aircraft aft fan
US10253779B2 (en) 2016-08-11 2019-04-09 General Electric Company Inlet guide vane assembly for reducing airflow swirl distortion of an aircraft aft fan
FR3061519B1 (fr) * 2016-12-30 2019-01-25 Safran Aircraft Engines Moyeu de carter intermediaire comprenant des canaux de guidage du flux de decharge formes par les ailettes de decharge
US10578055B2 (en) * 2017-08-10 2020-03-03 Mra Systems, Llc Turbine engine thrust reverser stop
US20190107046A1 (en) * 2017-10-05 2019-04-11 General Electric Company Turbine engine with struts
US11149552B2 (en) * 2019-12-13 2021-10-19 General Electric Company Shroud for splitter and rotor airfoils of a fan for a gas turbine engine
US11781506B2 (en) 2020-06-03 2023-10-10 Rtx Corporation Splitter and guide vane arrangement for gas turbine engines
US11702995B2 (en) * 2020-07-15 2023-07-18 Pratt & Whitney Canada Corp. Devices and methods for guiding bleed air in a turbofan engine
CN115434759B (zh) * 2022-09-20 2025-08-12 中国航发贵阳发动机设计研究所 一种航空发动机涡轮后支板叶片及涡轮后机匣
FR3155026A1 (fr) * 2023-11-03 2025-05-09 Safran Aircraft Engines Aube de redresseur avec une fente

Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB875411A (en) * 1959-05-20 1961-08-16 Bristol Siddeley Engines Ltd Improvements in gas turbine engines
JP2000337101A (ja) * 1999-05-28 2000-12-05 Mitsubishi Heavy Ind Ltd 軸流タービン翼列
WO2005119028A1 (fr) * 2004-06-01 2005-12-15 Volvo Aero Corporation Systeme de compression a turbine a gaz et structure de compresseur

Family Cites Families (11)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2331076A (en) * 1939-05-11 1943-10-05 Bbc Brown Boveri & Cie Turbine nozzle ring
GB628263A (en) * 1943-06-01 1949-08-25 Louis Breguet Improvements in or relating to axial flow compressors
US2576700A (en) * 1947-06-02 1951-11-27 Schneider Brothers Company Blading for fluid flow devices
US2935246A (en) * 1949-06-02 1960-05-03 Onera (Off Nat Aerospatiale) Shock wave compressors, especially for use in connection with continuous flow engines for aircraft
US3195807A (en) * 1958-10-20 1965-07-20 Gen Dynamics Corp Turbo-machine with slotted blades
DE1628237C3 (de) * 1966-07-21 1973-11-22 Wilhelm Prof. Dr.-Ing. 5100 Aachen Dettmering Stromungsmaschinen Umlenk schaufel gitter
US3572960A (en) * 1969-01-02 1971-03-30 Gen Electric Reduction of sound in gas turbine engines
US4652208A (en) * 1985-06-03 1987-03-24 General Electric Company Actuating lever for variable stator vanes
US4859145A (en) * 1987-10-19 1989-08-22 Sundstrand Corporation Compressor with supercritical diffuser
US6715983B2 (en) * 2001-09-27 2004-04-06 General Electric Company Method and apparatus for reducing distortion losses induced to gas turbine engine airflow
ITBA20030052A1 (it) * 2003-10-17 2005-04-18 Paolo Pietricola Pale rotoriche e statoriche a profili multipli

Patent Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB875411A (en) * 1959-05-20 1961-08-16 Bristol Siddeley Engines Ltd Improvements in gas turbine engines
JP2000337101A (ja) * 1999-05-28 2000-12-05 Mitsubishi Heavy Ind Ltd 軸流タービン翼列
WO2005119028A1 (fr) * 2004-06-01 2005-12-15 Volvo Aero Corporation Systeme de compression a turbine a gaz et structure de compresseur

Non-Patent Citations (1)

* Cited by examiner, † Cited by third party
Title
See also references of EP2092163A4 *

Cited By (16)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US8857193B2 (en) 2010-01-20 2014-10-14 Rolls-Royce Deutschland Ltd & Co Kg Intermediate casing for a gas-turbine engine
DE102010001059A1 (de) * 2010-01-20 2011-07-21 Rolls-Royce Deutschland Ltd & Co KG, 15827 Zwischengehäuse für ein Gasturbinentriebwerk
EP2397652A3 (fr) * 2010-06-20 2014-12-17 Honeywell International Inc. Aube de turbocompresseur à plusieurs profils aérodynamiques
EP2626515A1 (fr) * 2012-02-10 2013-08-14 MTU Aero Engines GmbH Agencement de groupes d'aubes en tandem
US9470091B2 (en) 2012-02-10 2016-10-18 Mtu Aero Engines Gmbh Blade group arrangement as well as turbomachine
GB2515961B (en) * 2012-04-05 2018-06-27 Snecma Stator vane formed by a set of vane parts
EP2746535A1 (fr) * 2012-12-20 2014-06-25 General Electric Company Conception de profil aérodynamique, à double rangée étagée, pour trame d'échappement de turbine à gaz
US9422864B2 (en) 2012-12-20 2016-08-23 General Electric Company Staggered double row, slotted airfoil design for gas turbine exhaust frame
WO2015108602A1 (fr) * 2013-10-29 2015-07-23 General Electric Company Ensemble entretoise de moteur d'aéronef et procédés d'assemblage de celui-ci
WO2018084902A1 (fr) * 2016-07-15 2018-05-11 General Electric Company Moteur à double flux et procédé de fonctionnement correspondant
GB2568109A (en) * 2017-11-07 2019-05-08 Gkn Aerospace Sweden Ab Splitter vane
WO2019091965A1 (fr) * 2017-11-07 2019-05-16 Gkn Aerospace Sweden Ab Structures arrière de turbine, moteur à turbine à gaz correspondant, aéronef et procédé de fabrication
CN111465750A (zh) * 2017-11-07 2020-07-28 Gkn航空公司 涡轮后部结构、对应的燃气涡轮发动机、飞机和制造方法
GB2568109B (en) * 2017-11-07 2021-06-09 Gkn Aerospace Sweden Ab Splitter vane
US11230943B2 (en) 2017-11-07 2022-01-25 Gkn Aerospace Sweden Ab Aircraft turbine rear structures
CN111465750B (zh) * 2017-11-07 2023-09-29 Gkn航空公司 涡轮后部结构、对应的燃气涡轮发动机、飞机和制造方法

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EP2092163A1 (fr) 2009-08-26
US20100158684A1 (en) 2010-06-24

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