US20090016871A1 - Systems and Methods Involving Variable Vanes - Google Patents

Systems and Methods Involving Variable Vanes Download PDF

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Publication number
US20090016871A1
US20090016871A1 US11/775,523 US77552307A US2009016871A1 US 20090016871 A1 US20090016871 A1 US 20090016871A1 US 77552307 A US77552307 A US 77552307A US 2009016871 A1 US2009016871 A1 US 2009016871A1
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United States
Prior art keywords
vane
pressurized air
operative
throat area
turbine engine
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
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US11/775,523
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English (en)
Inventor
Michael G. McCaffrey
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RTX Corp
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United Technologies Corp
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Filing date
Publication date
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Priority to US11/775,523 priority Critical patent/US20090016871A1/en
Assigned to UNITED TECHNOLOGIES CORP. reassignment UNITED TECHNOLOGIES CORP. ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: MCCAFFREY, MICHAEL G
Assigned to UNITED TECHNOLOGIES CORP. reassignment UNITED TECHNOLOGIES CORP. CORRECTIVE ASSIGNMENT TO CORRECT THE ASSIGNEE ADDRESS PREVIOUSLY RECORDED ON REEL 019538 FRAME 0268. ASSIGNOR(S) HEREBY CONFIRMS THE 400 MAIN STREET EAST HARTFORD, CT 06108. Assignors: MCCAFFREY, MICHAEL G.
Priority to EP08252364A priority patent/EP2014871B1/de
Publication of US20090016871A1 publication Critical patent/US20090016871A1/en
Assigned to RAYTHEON TECHNOLOGIES CORPORATION reassignment RAYTHEON TECHNOLOGIES CORPORATION CHANGE OF NAME (SEE DOCUMENT FOR DETAILS). Assignors: UNITED TECHNOLOGIES CORPORATION
Assigned to RAYTHEON TECHNOLOGIES CORPORATION reassignment RAYTHEON TECHNOLOGIES CORPORATION CORRECTIVE ASSIGNMENT TO CORRECT THE AND REMOVE PATENT APPLICATION NUMBER 11886281 AND ADD PATENT APPLICATION NUMBER 14846874. TO CORRECT THE RECEIVING PARTY ADDRESS PREVIOUSLY RECORDED AT REEL: 054062 FRAME: 0001. ASSIGNOR(S) HEREBY CONFIRMS THE CHANGE OF ADDRESS. Assignors: UNITED TECHNOLOGIES CORPORATION
Assigned to RTX CORPORATION reassignment RTX CORPORATION CHANGE OF NAME (SEE DOCUMENT FOR DETAILS). Assignors: RAYTHEON TECHNOLOGIES CORPORATION
Abandoned legal-status Critical Current

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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/141—Shape, i.e. outer, aerodynamic form
    • F01D5/145—Means for influencing boundary layers or secondary circulations
    • 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
    • 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
    • F01D5/148—Blades with variable camber, e.g. by ejection of fluid
    • 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
    • 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
    • F05D2270/00—Control
    • F05D2270/01—Purpose of the control system
    • F05D2270/17—Purpose of the control system to control boundary layer

Definitions

  • the invention relates to gas turbine engines.
  • Gas turbine engines use compressors to compress gas for combustion.
  • a compressor typically uses alternating sets of rotating blades and stationary vanes to compress gas. Gas flowing through such a compressor is forced between the sets and between adjacent blades and vanes of a given set. Similarly, after combustion, hot expanding gas drives a turbine that has sets of rotating blades and stationary vanes.
  • an exemplary embodiment of a gas turbine engine defining a gas flow path comprises: a first vane extending into the gas flow path and having: an interior operative to receive pressurized air; an outer surface; and outlet ports communicating between the outer surface and the interior of the first vane, the outlet ports being operative to receive the pressurized air from the interior and emit the pressurized air into the gas flow path such that a throat area defined, at least in part, by the first vane is modified.
  • An exemplary embodiment of a vane assembly comprises: a first vane having: an outer surface; an interior defining a cavity operative to receive pressurized air; and outlet ports communicating between the outer surface and the cavity, the outlet ports being operative to receive the pressurized air from the cavity and emit the pressurized air through the outer surface a valve assembly operative to regulate the pressurized air emitted by the first vane.
  • An exemplary embodiment of a method for modifying the throat area between vanes of a gas turbine engine comprises: directing a gas flow path of the gas turbine engine between a first vane and a second vane, wherein each of the first vane and the second vane has an outer surface and an interior; and emitting pressurized air from outlet ports communicating between the outer surface and the interior of the first vane, wherein the emitted pressurized air from the first vane modifies a throat area between the first vane and the second vane.
  • FIG. 1 is a schematic side cutaway view illustrating an exemplary embodiment of a turbine section of a gas turbine engine.
  • FIG. 2 is a side cutaway view of an exemplary embodiment of a vane.
  • FIG. 3 is a top cutaway view of an exemplary embodiment of vanes in a gas flow path.
  • FIG. 4 is a top cutaway view of another exemplary embodiment of vanes in a gas flow path.
  • FIG. 5 is a top cutaway view of another exemplary embodiment of vanes in a gas flow path.
  • gas passing through a gas turbine engine enters a turbine that includes rotating blades and stationary vanes.
  • the gas, following the gas flow path is forced between adjacent vanes.
  • the vanes are often shaped like airfoils and, therefore, have aerodynamic properties similar to airfoils.
  • the flow of gas between adjacent vanes results in a throat area determined by, for example, the shape and relative position of the vanes.
  • the angle of the vanes relative to the gas flow path may be mechanically changed to vary the location and/or size of the throat area and alter the efficiency of the engine.
  • the gas turbine engine is configured as a turbofan.
  • FIG. 1 is a schematic side view illustrating an exemplary embodiment of a turbine section 100 of a gas turbine engine.
  • rotating blades 104 are attached to a disk that is rotated by a shaft 106 .
  • Stationary vanes 108 are attached to the casing of the engine between the blades 104 .
  • gas enters the turbine section along gas flow path 102 and drives the blades 104 .
  • the gas exits the turbine section 100 along gas flow path 102 .
  • FIG. 2 is a simplified, side cutaway view of vane assembly 200 that includes a vane airfoil 202 and a valve assembly 208 .
  • vane airfoil 202 typically is mounted to and spans between an outer diameter vane platform and an inner diameter vane platform, neither of which is depicted in FIG. 2 .
  • valve assembly 208 includes a piston 204 and solenoid 220 , which is used to actuate the piston.
  • Inlet ports 218 provide gas to the valve assembly so that actuation of the piston pressurizes the received gas.
  • the pressurized gas from the outlet ports 216 urges the gas flow path, which flows about the vane airfoil during operation of the gas turbine engine, away from the exterior surface of the vane airfoil to a greater extent than that caused by pressurized gas involved in film cooling.
  • the boundary layer formed by the film-cooling air also is urged away from the exterior of the vane airfoil.
  • the pressure of the gas required to alter the throat is not available from the compressor alone.
  • piston 204 is used in the embodiment of FIG. 2 to increase the pressure of the gas provided to the outlet ports. In other embodiments, various other mechanisms could be used to increase the gas pressure.
  • the shape of the vane assembly 200 illustrated in FIG. 2 is merely an illustration of but one possible embodiment.
  • the shape of the vane assembly 200 may vary depending on a variety of factors including, but not limited to, the component to which the vane assembly 200 is attached, the location of the vane assembly 200 in the gas turbine engine, the gas flow path around the vane assembly 200 at particular gas flow velocities, desired design characteristics of the gas turbine engine, and materials used in the fabrication of the gas turbine engine.
  • a controller 212 also is provided.
  • the controller 212 is used to open and close the valve assembly 208 .
  • the valve assembly 208 is left open such that the outlet ports 216 emit a constant flow of pressurized air.
  • the valve assembly 208 may be opened and closed intermittently.
  • the pressurized air may be emitted from the outlet ports 216 in pulses.
  • operation in a pulsed mode allows the pressure of the pressurized air to increase prior to being emitted into a gas flow path.
  • the controller 212 may be set to control the frequency of the pulses of emitted pressurized air. Controlling the frequency of the pulses may be desirable because a change in the throat area based on a frequency of pulses may allow the aerodynamic characteristics of the engine to be adjusted.
  • the frequencies of the pulses may be controlled to modify one or more throat areas in a specific region of an engine to control local pressure ratios and/or local temperatures.
  • the pulse frequencies may also be timed to adjust for resonance in the engine that may result in vane and blade vibrations. These pulses may be used to add a canceling frequency that may effectively cancel engine resonance, for example.
  • FIG. 3 is a top cutaway view of a pair of vanes in an embodiment of a gas turbine engine. As shown in FIG. 3 , gas is forced between the vanes 300 along gas flow path 302 , forming a throat area 304 .
  • the shape of the adjacent vanes 300 , their proximity to each other, and the angle of incidence to the gas flow path 302 are possible factors that can influence the location and size of the throat area 304 .
  • FIG. 4 depicts a top cutaway view of another embodiment of a vane assembly.
  • vanes 406 and 412 are adjacent vanes.
  • Vane 406 has an interior cavity 404 that is connected to a pressurized air source (not shown).
  • Outlet ports 410 are located on the surface of vane 406 and are in communication with interior cavity 404 .
  • Pressurized air emitted from the outlet ports 410 in vane 406 defines a boundary layer 408 that has an aerodynamic effect on the gas flow path 402 .
  • the boundary layer 408 associated with the pressurized air from the outlet ports modifies the location and/or size of the throat area 416 .
  • the outlet ports of this embodiment are oriented such that the flow from the outlet ports is generally in a direction of the gas flow path. In other embodiments, however, the orientation can be different, such as by providing a perpendicular (see FIG. 5 ) or counter flow (not shown).
  • Modifying the throat area of an engine may affect the flow of gasses through the engine. For instance, such modifying can affect the pressure ratio of the compressor and change the relationship between the flow and the pressure ratio. For example, a lower flow rate can increase the pressure ratio.
  • FIG. 5 depicts a top cutaway view of another embodiment of a vane assembly.
  • vane assembly 500 incorporates two adjacent vanes, a first vane 501 and a second vane 503 .
  • the first vane 501 and the second vane 503 are spaced from each other to define a gas flow path 502 .
  • the first vane 501 includes three chambers—a film-cooling chamber 504 , a suction side chamber 505 and a pressure side chamber 507 .
  • the film-cooling chamber 504 , suction side chamber 505 and the pressure side chamber 507 include ports, such as ports 506 , 509 and 511 , respectively.
  • the film-cooling chamber 504 receives cooling pressurized air that is emitted from the associated ports, e.g., port 506 .
  • This air creates a relatively thin boundary layer 530 that is located adjacent to the exterior of the vane 501 to serve as a barrier against the hot gas flowpath 502 .
  • the suction side chamber 505 and the pressure side chamber 507 also receive pressurized air, which is at a higher pressure than that provided to chamber 504 , that is emitted from associated ports, e.g., ports 509 and 511 .
  • the pressurized air emitted from chamber 507 creates a boundary layer 513 along the pressure surface 515 of the first vane 501 that affects the throat area 550 .
  • the boundary layer 513 tends to urge the boundary layer 530 away from the pressure surface 515 , thereby causing the boundary layer 530 to dissipate and mix with the gas of the gas flow path 502 .
  • the second vane 503 also includes three chambers—a film-cooling chamber 532 , a suction side chamber 510 and a pressure side chamber 512 .
  • the film-cooling chamber 532 , suction side chamber 510 and the pressure side chamber 512 include ports, such as ports 534 , 522 and 514 , respectively.
  • the film-cooling chamber 532 receives cooling pressurized air that is emitted from the associated ports, e.g., port 534 .
  • This air creates a relatively thin boundary layer 536 that is located adjacent to the exterior of the vane 503 .
  • the suction side chamber 510 and the pressure side chamber 512 also receive pressurized air, which is at a higher pressure than that provided to chamber 534 , that is emitted from associated ports, e.g., ports 522 and 514 .
  • the pressurized air emitted from chamber 510 creates a boundary layer 525 along the suction surface 506 of the vane 503 that affects the throat area 550 .
  • the boundary layer 525 tends to urge the boundary layer 536 away from the suction surface 506 , thereby causing the boundary layer 536 to dissipate and mix with the gas of the gas flow path 502 .

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Physics & Mathematics (AREA)
  • Fluid Mechanics (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Supercharger (AREA)
  • Control Of Turbines (AREA)
US11/775,523 2007-07-10 2007-07-10 Systems and Methods Involving Variable Vanes Abandoned US20090016871A1 (en)

Priority Applications (2)

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US11/775,523 US20090016871A1 (en) 2007-07-10 2007-07-10 Systems and Methods Involving Variable Vanes
EP08252364A EP2014871B1 (de) 2007-07-10 2008-07-10 Systeme und Verfahren mit variablen Schaufeln

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US11/775,523 US20090016871A1 (en) 2007-07-10 2007-07-10 Systems and Methods Involving Variable Vanes

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Cited By (13)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20090162189A1 (en) * 2007-12-19 2009-06-25 United Technologies Corp. Systems and Methods Involving Variable Throat Area Vanes
EP2532857A1 (de) 2011-06-06 2012-12-12 United Technologies Corporation Turbomaschinenanordnung mit Brennkammern von unterschiedlichen Strömungsrichtungen und zugehöriges Betriebsverfahren
US20160061054A1 (en) * 2014-09-03 2016-03-03 Honeywell International Inc. Structural frame integrated with variable-vectoring flow control for use in turbine systems
US20160298470A1 (en) * 2015-04-08 2016-10-13 United Technologies Corporation Airfoils
JP2016211569A (ja) * 2015-05-11 2016-12-15 ゼネラル・エレクトリック・カンパニイ タービンにおける流れ制御のためのシステムおよび方法
EP3133246A1 (de) * 2015-08-18 2017-02-22 General Electric Company Luftstromeinspritzdüse für einen gasturbinenmotor
US20170070973A1 (en) * 2011-08-12 2017-03-09 Qualcomm Incorporated Devices for reduced overhead paging
US9617868B2 (en) 2013-02-26 2017-04-11 Rolls-Royce North American Technologies, Inc. Gas turbine engine variable geometry flow component
US20180355738A1 (en) * 2017-06-13 2018-12-13 General Electric Company Turbine engine with variable effective throat
US10578028B2 (en) 2015-08-18 2020-03-03 General Electric Company Compressor bleed auxiliary turbine
US10711702B2 (en) 2015-08-18 2020-07-14 General Electric Company Mixed flow turbocore
US20210301684A1 (en) * 2020-03-30 2021-09-30 General Electric Company Fluidic flow control device
US11692448B1 (en) 2022-03-04 2023-07-04 General Electric Company Passive valve assembly for a nozzle of a gas turbine engine

Families Citing this family (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB0910647D0 (en) 2009-06-22 2009-08-05 Rolls Royce Plc A compressor blade
EP3477075B1 (de) * 2016-12-21 2022-04-13 Mitsubishi Heavy Industries Engine & Turbocharger, Ltd. Turbolader, turboladerleitschaufel und turbine
US20200362704A1 (en) * 2019-05-17 2020-11-19 Solar Turbines Incorporated Nozzle segment

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US2825532A (en) * 1951-01-04 1958-03-04 Snecma Device for controlling the flow of fluid between cambered blades
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US4624104A (en) * 1984-05-15 1986-11-25 A/S Kongsberg Vapenfabrikk Variable flow gas turbine engine
US4707981A (en) * 1986-01-27 1987-11-24 Rockwell International Corporation Variable expansion ratio reaction engine
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US4741667A (en) * 1986-05-28 1988-05-03 United Technologies Corporation Stator vane
US5207556A (en) * 1992-04-27 1993-05-04 General Electric Company Airfoil having multi-passage baffle
US5645397A (en) * 1995-10-10 1997-07-08 United Technologies Corporation Turbine vane assembly with multiple passage cooled vanes
US5833433A (en) * 1997-01-07 1998-11-10 Mcdonnell Douglas Corporation Rotating machinery noise control device
US6026791A (en) * 1997-03-03 2000-02-22 Alliedsignal Inc. Exhaust gas recirculation valve with integral feedback proportional to volumetric flow
US6164903A (en) * 1998-12-22 2000-12-26 United Technologies Corporation Turbine vane mounting arrangement
US6565313B2 (en) * 2001-10-04 2003-05-20 United Technologies Corporation Bleed deflector for a gas turbine engine
US6929446B2 (en) * 2003-10-22 2005-08-16 General Electric Company Counterbalanced flow turbine nozzle
US20090162189A1 (en) * 2007-12-19 2009-06-25 United Technologies Corp. Systems and Methods Involving Variable Throat Area Vanes

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US2801790A (en) * 1950-06-21 1957-08-06 United Aircraft Corp Compressor blading
US2746672A (en) * 1950-07-27 1956-05-22 United Aircraft Corp Compressor blading
US2825532A (en) * 1951-01-04 1958-03-04 Snecma Device for controlling the flow of fluid between cambered blades
US4072008A (en) * 1976-05-04 1978-02-07 General Electric Company Variable area bypass injector system
US4504189A (en) * 1982-11-10 1985-03-12 Rolls-Royce Limited Stator vane for a gas turbine engine
US4624104A (en) * 1984-05-15 1986-11-25 A/S Kongsberg Vapenfabrikk Variable flow gas turbine engine
US4740138A (en) * 1985-12-04 1988-04-26 MTU Motoren-und Turbinen-Munchen GmbH Device for controlling the throat areas between the diffusor guide vanes of a centrifugal compressor of a gas turbine engine
US4707981A (en) * 1986-01-27 1987-11-24 Rockwell International Corporation Variable expansion ratio reaction engine
US4741667A (en) * 1986-05-28 1988-05-03 United Technologies Corporation Stator vane
US5207556A (en) * 1992-04-27 1993-05-04 General Electric Company Airfoil having multi-passage baffle
US5645397A (en) * 1995-10-10 1997-07-08 United Technologies Corporation Turbine vane assembly with multiple passage cooled vanes
US5833433A (en) * 1997-01-07 1998-11-10 Mcdonnell Douglas Corporation Rotating machinery noise control device
US6026791A (en) * 1997-03-03 2000-02-22 Alliedsignal Inc. Exhaust gas recirculation valve with integral feedback proportional to volumetric flow
US6164903A (en) * 1998-12-22 2000-12-26 United Technologies Corporation Turbine vane mounting arrangement
US6565313B2 (en) * 2001-10-04 2003-05-20 United Technologies Corporation Bleed deflector for a gas turbine engine
US6929446B2 (en) * 2003-10-22 2005-08-16 General Electric Company Counterbalanced flow turbine nozzle
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Cited By (20)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US8197209B2 (en) * 2007-12-19 2012-06-12 United Technologies Corp. Systems and methods involving variable throat area vanes
US20090162189A1 (en) * 2007-12-19 2009-06-25 United Technologies Corp. Systems and Methods Involving Variable Throat Area Vanes
EP2532857A1 (de) 2011-06-06 2012-12-12 United Technologies Corporation Turbomaschinenanordnung mit Brennkammern von unterschiedlichen Strömungsrichtungen und zugehöriges Betriebsverfahren
US20170070973A1 (en) * 2011-08-12 2017-03-09 Qualcomm Incorporated Devices for reduced overhead paging
US9617868B2 (en) 2013-02-26 2017-04-11 Rolls-Royce North American Technologies, Inc. Gas turbine engine variable geometry flow component
US10221720B2 (en) * 2014-09-03 2019-03-05 Honeywell International Inc. Structural frame integrated with variable-vectoring flow control for use in turbine systems
US20160061054A1 (en) * 2014-09-03 2016-03-03 Honeywell International Inc. Structural frame integrated with variable-vectoring flow control for use in turbine systems
US20190078466A1 (en) * 2014-09-03 2019-03-14 Honeywell International Inc. Structural frame integrated with variable-vectoring flow control for use in turbine systems
US20160298470A1 (en) * 2015-04-08 2016-10-13 United Technologies Corporation Airfoils
US10641113B2 (en) * 2015-04-08 2020-05-05 United Technologies Corporation Airfoils
JP2016211569A (ja) * 2015-05-11 2016-12-15 ゼネラル・エレクトリック・カンパニイ タービンにおける流れ制御のためのシステムおよび方法
US20170051680A1 (en) * 2015-08-18 2017-02-23 General Electric Company Airflow injection nozzle for a gas turbine engine
US10578028B2 (en) 2015-08-18 2020-03-03 General Electric Company Compressor bleed auxiliary turbine
EP3133246A1 (de) * 2015-08-18 2017-02-22 General Electric Company Luftstromeinspritzdüse für einen gasturbinenmotor
US10711702B2 (en) 2015-08-18 2020-07-14 General Electric Company Mixed flow turbocore
CN109083690A (zh) * 2017-06-13 2018-12-25 通用电气公司 具有可变有效喉道的涡轮发动机
US20180355738A1 (en) * 2017-06-13 2018-12-13 General Electric Company Turbine engine with variable effective throat
US10760426B2 (en) * 2017-06-13 2020-09-01 General Electric Company Turbine engine with variable effective throat
US20210301684A1 (en) * 2020-03-30 2021-09-30 General Electric Company Fluidic flow control device
US11692448B1 (en) 2022-03-04 2023-07-04 General Electric Company Passive valve assembly for a nozzle of a gas turbine engine

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Publication number Publication date
EP2014871A3 (de) 2011-08-31
EP2014871B1 (de) 2012-11-14
EP2014871A2 (de) 2009-01-14

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