US3608310A - Turbine stator-combustor structure - Google Patents

Turbine stator-combustor structure Download PDF

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Publication number
US3608310A
US3608310A US560658A US3608310DA US3608310A US 3608310 A US3608310 A US 3608310A US 560658 A US560658 A US 560658A US 3608310D A US3608310D A US 3608310DA US 3608310 A US3608310 A US 3608310A
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United States
Prior art keywords
combustor
air
vanes
stator
turbine
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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.)
Expired - Lifetime
Application number
US560658A
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English (en)
Inventor
John M Vaught
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Motors Liquidation Co
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General Motors Corp
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Publication date
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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
    • F23R3/16—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration with devices inside the flame tube or the combustion chamber to influence the air or gas flow
    • 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
    • F01D9/023—Transition ducts between combustor cans and first stage of the turbine in gas-turbine engines; their cooling or sealings
    • 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
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
    • F23R3/04—Air inlet arrangements
    • 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
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/201—Heat transfer, e.g. cooling by impingement of a fluid
    • 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

  • My invention relates generally to combustor equipment for a gas turbine engine or the like and more specifically to a turbine stator-combustor structure which is especially suitable for, but not limited to, use with a lift type gas turbine engine because of its reduced axial length.
  • VTOL aircraft For some time now, the feasibility of VTOL aircraft has been under study.
  • One basic concept under study is the inclusion of lift engines vertically mounted in the aircraft which are used during take-off and landing only and are inoperable during normal flight. With the lift engine mounted vertically, it is highly desirable to reduce its axial length as much as practicable.
  • my invention I have integrated a turbine inlet stator into the dilution zone of an annular combustor to decrease the axial length normally required by these components.
  • the material now available for the turbine stator or nozzle inlet vane row has a temperture limit which is far below the temperatures available from the combustion of liquid hydrocarbon fuels in air at or near the chemically correct fuel/ air ratio (stoichiometric ratio). For this reason, it is necessary to include a section in the combustor in which the combustion gases are diluted with a large excess of secondary air to reduce their temperature before they reach the turbine stator.
  • the area of the combustor in which dilution or secondary air is admitted is known as the dilution zone which heretofore has added length to the combustor and turbine inlet stator structure.
  • the dilution zone which heretofore has added length to the combustor and turbine inlet stator structure.
  • no dilution zone would be required and the added length of the dilution zone eliminated.
  • the temperature limit problem of the stator exists, however, and it increases as the distance between the primary combustion zone and the inlet stator decreases. Therefore, any decrease in this distance (and consequently the axial length of the structure) must take into account the temperature limit of the inlet nozzle vane stator.
  • the object of my invention in its broadest aspect is to provide an operable turbine stator-combustor structure for a gas turbine or the like having an axial length shorter than any available comparable structure.
  • Another object of my invention is to provide a relatively short combustor structure having a turbine stator ice incorporated in its dilution zone, in which the hydrocarbon fuel may be efliciently burned, in which the combustion gases are cooled to a temperature which is capable of being withstood by the turbine stator and in which does not appreciably affect engine performance.
  • Another object of my invention is to provide turbine stator-combustor structure having a relatively short axial length and including means for convection cooling the turbine stator vanes, film cooling of their outer surfaces, and impingement cooling of their leading edges so that they can withstand the environmental temperature produced by the combustion of a stoichiometric mixture of fuel and air in the combustor.
  • FIG. 1 is a section taken on a plane containing the longitudinal axis of a portion of a gas turbine engine showing an integrated turbine stator-combustor in accordance with my invention and its relationship to the engine.
  • FIG. 2 is a section taken along the line 2--2 of FIG. 1 and looking in the direction of the arrows.
  • the gas turbine engine combustion portion is indicated generally at 12.
  • the outer casing 14 of the engine and a wall 16 mounted within it define an annular air casing which receives compressed air at inlet 17.
  • An annular combustor 18 mounted in the air casing and spaced from it provides outer and inner annular air passages 20 and 22.
  • Mounted on the forward end of the combustor 18 is a deflector 23 which proportions the air flow from inlet 17 between the air passages 19 and the annular air passages 20 and 22.
  • a fuel ring 24 mounted inside the deflector 23 includes a number of circumferentially spaced fuel spray nozzles 26 which protrude into the forward end of the combustor 18.
  • nozzle vanes 28 Mounted radially across the aft end of the combustor 18 are a number of circumferentially spaced nozzle vanes 28. These nozzle vanes are hollow and have open radially outer and inner ends 30 and 32 which are in fluid communication with the annular air passages 20 and 22, respectively.
  • the walls 33 providing the aerodynamic contour of the vanes have a number of apertures 35 of various size, larger ones being provided in the midportions of the vanes.
  • Apertures 37 in the combustor inner wall admit dilution or cooling air to the area between the vanes.
  • the circumferentially spaced apertures 34 provided in the combustor 18 adjacent the leading edges 36 of the vanes 28 have short truncated tubes 38 mounted in each aperture 34 which direct secondary air from the passages 20 and 22 into the combustor and onto the leading edge 36 of the vanes 28.
  • the primary combustion section 40 is the area within the combustor between the tip of the fuel nozzles 26 and the point of admitting secondary air at the leading edges 36 of the vanes 28 through apertures 34.
  • the forward portion or primary combustion section 40 is fabricated from known techniques so that adequate film cooling of this area is provided by air flowing through the slots 42.
  • the particular combustor illustrated is of the progressive burning type so that primary air is admitted through the larger apertures 44 and smaller aperture 45 in addition to the primary air admitted through passages 19.
  • the progressive type burner forms no part of this invention but suflice it to say that the fuel admitted through nozzles 26 is completely burned by the time it reaches the dilution zone-sufiicient air for a stoichiometeric mixture with a little excess for practical reasons being admitted through passages 19 and apertures 44 and 45.
  • compressed air is proportioned by deflector 23 bet-ween passages 19 and the annular passages 20 and 22.
  • the air admitted through passages 19 is mixed with the fuel spray from nozzles 26 and ignited and burned.
  • Additional air from passages 20 and 22 flows through apertures 44 and 45 to completely burn the fuel and through slots 42 to film cool the combustor inner walls.
  • secondary air flows radially through the apertures 34 and is directed by the tubes 38 into the combustor and onto the vane leading edges 36. This flow is substantial and dilutes the combustion gases.
  • the flow has an impingement cooling elfect on the leading edges 36 of the vanes 28.
  • Air from passages 20 and 22 also flows radially into the hollow vanes from both directions and out the apertures 35 and along their radial walls toward the aft end of the combustor structure 12.
  • This last-mentioned flow provides convection cooling of the vanes 28 and a film of cool air adjacent the outer surface of the vanes 28. Dilution air may also be admitted through the apertures 37 between the vanes 28.
  • the flow of primary or combustion air is indicated by open arrows and the flow of secondary air is indicated by black arrows in FIG. 1. Note the vanes are convection cooled and have their exposed surface film cooled while their most critical areas, that is, their leading edges are impingement cooled by the dilution air flowing through aperture 34.
  • This cooling combination I has found provides adequate cooling of the vanes 28 made of presently available material with approximately 50% of the air flow while the remaining 50% of the compressed air available as primary air for I combustion is adequate to support efiicient combustion and yield satisfactory engine performance.
  • the cooling'of the blade in this manner also aids in presenting a temperature profile at the stator exit which is capable of being withstood by the first stage turbine 46.
  • Combustors are designed from empirical data and normally the primary combustion zone 40 has a length-toheight ratio of approximately 2.5 whereas the integrated turbine stator-combustor of my invention has a primary combustion zone with a length-to-height ratio on the order of 1.5.
  • my invention provides a turbine-stator-combustor structure capable of efl'icient combustion and with standing the combustion temperatures which is considerably shonter than comparable structures now available.
  • a turbine-stator-combustor for a gas turbine engine or the like comprising in combination: an annular casing, a combustor mounted in said casing and spaced radially therefrom to provide inner and outer annular air passages, a plurality of circumferentially spaced stator vanes mounted at the aft end of said combustor, said combustor having a plurality of circumferentially spaced apertures adjacent the leading edges of said vanes, truncated tubes mounted in said combustor apertures with their maximum height remote from the leading edges of said vanes, said truncated tubes extending into said combustor to direct secondary air onto said leading edges of said vanes to cool them by impingement of the secondary air.
  • stator vanes are hollow, having open inner and outer ends in fluid communication with said annular air passages, respectively, and have a plurality of apertures in their radial walls whereby secondary air is directed into said hollow blades and outwardly along the exposed surfaces of said radial walls to convection cool the said vanes and to provide a film of cool air adjacent .said radial walls.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
US560658A 1966-06-27 1966-06-27 Turbine stator-combustor structure Expired - Lifetime US3608310A (en)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US56065866A 1966-06-27 1966-06-27

Publications (1)

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US3608310A true US3608310A (en) 1971-09-28

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Family Applications (1)

Application Number Title Priority Date Filing Date
US560658A Expired - Lifetime US3608310A (en) 1966-06-27 1966-06-27 Turbine stator-combustor structure

Country Status (2)

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US (1) US3608310A (enrdf_load_stackoverflow)
GB (1) GB1239559A (enrdf_load_stackoverflow)

Cited By (30)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4733538A (en) * 1978-10-02 1988-03-29 General Electric Company Combustion selective temperature dilution
US4825640A (en) * 1987-06-22 1989-05-02 Sundstrand Corporation Combustor with enhanced turbine nozzle cooling
US4926630A (en) * 1988-12-12 1990-05-22 Sundstrand Corporation Jet air cooled turbine shroud for improved swirl cooling and mixing
US5101620A (en) * 1988-12-28 1992-04-07 Sundstrand Corporation Annular combustor for a turbine engine without film cooling
US5303543A (en) * 1990-02-08 1994-04-19 Sundstrand Corporation Annular combustor for a turbine engine with tangential passages sized to provide only combustion air
EP0564183B1 (en) * 1992-03-30 1997-07-23 General Electric Company Dilution pole combustor and method
US5953919A (en) * 1996-12-13 1999-09-21 Asea Brown Boveri Ag Combustion chamber having integrated guide blades
US6101814A (en) * 1999-04-15 2000-08-15 United Technologies Corporation Low emissions can combustor with dilution hole arrangement for a turbine engine
DE10214574A1 (de) * 2002-04-02 2003-10-16 Rolls Royce Deutschland Brennkammer für ein Luftstrahltriebwerk mit Sekundärluftzuführung
GB2391297A (en) * 2002-07-24 2004-02-04 Rolls Royce Plc Gas supply assembly
WO2004038181A1 (en) * 2002-10-23 2004-05-06 Pratt & Whitney Canada Corp. Aerodynamic method to reduce noise level in gas turbines
FR2871847A1 (fr) * 2004-06-17 2005-12-23 Snecma Moteurs Sa Montage d'un distributeur de turbine sur une chambre de combustion a parois en cmc dans une turbine a gaz
WO2006053825A1 (de) * 2004-11-16 2006-05-26 Alstom Technology Ltd Gasturbinenanlage und zugehörige brennkammer
EP2053312A2 (fr) * 2007-10-22 2009-04-29 Snecma Chambre de combustion à dilution optimisée et turbomachine en étant munie
US20090139239A1 (en) * 2007-11-29 2009-06-04 Honeywell International, Inc. Quench jet arrangement for annular rich-quench-lean gas turbine combustors
US20090266080A1 (en) * 2008-04-24 2009-10-29 Snecma Optimizing the angular positioning of a turbine nozzle at the outlet from a turbomachine combustion chamber
US20100077719A1 (en) * 2008-09-29 2010-04-01 Siemens Energy, Inc. Modular Transvane Assembly
US20100122538A1 (en) * 2008-11-20 2010-05-20 Wei Ning Methods, apparatus and systems concerning the circumferential clocking of turbine airfoils in relation to combustor cans and the flow of cooling air through the turbine hot gas flowpath
US20100162712A1 (en) * 2007-11-29 2010-07-01 Honeywell International Inc. Quench jet arrangement for annular rich-quench-lean gas turbine combustors
US20100229563A1 (en) * 2006-01-25 2010-09-16 Woolford James R Wall elements for gas turbine engine combustors
US20100313571A1 (en) * 2007-12-29 2010-12-16 Alstom Technology Ltd Gas turbine
US20110203282A1 (en) * 2008-09-29 2011-08-25 Charron Richard C Assembly for directing combustion gas
US20130086908A1 (en) * 2010-06-15 2013-04-11 Rolls-Royce Deutschland Ltd & Co Kg Gas turbine combustion chamber arrangement of axial type of construction
EP2613080A1 (en) * 2012-01-05 2013-07-10 Siemens Aktiengesellschaft Combustion chamber of an annular combustor for a gas turbine
US20140338336A1 (en) * 2012-09-26 2014-11-20 United Technologies Corporation Gas turbine engine combustor with integrated combustor vane
US9151501B2 (en) 2011-07-28 2015-10-06 Rolls-Royce Deutschland Ltd & Co Kg Gas turbine centripetal annular combustion chamber and method for flow guidance
US9322335B2 (en) 2013-03-15 2016-04-26 Siemens Energy, Inc. Gas turbine combustor exit piece with hinged connections
EP3076079A1 (en) * 2015-03-31 2016-10-05 Rolls-Royce plc Combustion equipment
US9590472B2 (en) 2013-02-15 2017-03-07 Siemens Aktiengesellschaft Through flow ventilation system for a power generation turbine package
EP3848556A1 (en) * 2020-01-13 2021-07-14 Ansaldo Energia Switzerland AG Gas turbine engine having a transition piece with inclined cooling holes

Families Citing this family (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB2030653B (en) * 1978-10-02 1983-05-05 Gen Electric Gas turbine engine combustion gas temperature variation
US4739621A (en) * 1984-10-11 1988-04-26 United Technologies Corporation Cooling scheme for combustor vane interface
FR2674317B1 (fr) * 1991-03-20 1993-05-28 Snecma Chambre de combustion de turbomachine comportant un reglage du debit de comburant.
FR2686683B1 (fr) * 1992-01-28 1994-04-01 Snecma Turbomachine a chambre de combustion demontable.

Cited By (54)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4733538A (en) * 1978-10-02 1988-03-29 General Electric Company Combustion selective temperature dilution
US4825640A (en) * 1987-06-22 1989-05-02 Sundstrand Corporation Combustor with enhanced turbine nozzle cooling
US4926630A (en) * 1988-12-12 1990-05-22 Sundstrand Corporation Jet air cooled turbine shroud for improved swirl cooling and mixing
WO1990007087A1 (en) * 1988-12-12 1990-06-28 Sundstrand Corporation Jet air cooled turbine shroud for improved swirl cooling and mixing
US5101620A (en) * 1988-12-28 1992-04-07 Sundstrand Corporation Annular combustor for a turbine engine without film cooling
US5303543A (en) * 1990-02-08 1994-04-19 Sundstrand Corporation Annular combustor for a turbine engine with tangential passages sized to provide only combustion air
EP0564183B1 (en) * 1992-03-30 1997-07-23 General Electric Company Dilution pole combustor and method
US5953919A (en) * 1996-12-13 1999-09-21 Asea Brown Boveri Ag Combustion chamber having integrated guide blades
EP0848210A3 (de) * 1996-12-13 1999-11-17 Asea Brown Boveri AG Brennkammer mit integrierten Leitschaufeln
US6101814A (en) * 1999-04-15 2000-08-15 United Technologies Corporation Low emissions can combustor with dilution hole arrangement for a turbine engine
DE10214574A1 (de) * 2002-04-02 2003-10-16 Rolls Royce Deutschland Brennkammer für ein Luftstrahltriebwerk mit Sekundärluftzuführung
GB2391297A (en) * 2002-07-24 2004-02-04 Rolls Royce Plc Gas supply assembly
WO2004038181A1 (en) * 2002-10-23 2004-05-06 Pratt & Whitney Canada Corp. Aerodynamic method to reduce noise level in gas turbines
US7234304B2 (en) 2002-10-23 2007-06-26 Pratt & Whitney Canada Corp Aerodynamic trip to improve acoustic transmission loss and reduce noise level for gas turbine engine
US20070227119A1 (en) * 2002-10-23 2007-10-04 Pratt & Whitney Canada Corp. HPT aerodynamic trip to improve acoustic transmission loss and reduce noise level for auxiliary power unit
US7533534B2 (en) 2002-10-23 2009-05-19 Pratt & Whitney Canada Corp. HPT aerodynamic trip to improve acoustic transmission loss and reduce noise level for auxiliary power unit
FR2871847A1 (fr) * 2004-06-17 2005-12-23 Snecma Moteurs Sa Montage d'un distributeur de turbine sur une chambre de combustion a parois en cmc dans une turbine a gaz
US20060010879A1 (en) * 2004-06-17 2006-01-19 Snecma Moteurs Mounting a turbine nozzle on a combustion chamber having CMC walls in a gas turbine
US7249462B2 (en) 2004-06-17 2007-07-31 Snecma Mounting a turbine nozzle on a combustion chamber having CMC walls in a gas turbine
RU2392447C2 (ru) * 2004-06-17 2010-06-20 Снекма Турбомашина, сопловой аппарат которой установлен на камере сгорания со стенками из композитного материала
WO2006053825A1 (de) * 2004-11-16 2006-05-26 Alstom Technology Ltd Gasturbinenanlage und zugehörige brennkammer
US8024933B2 (en) * 2006-01-25 2011-09-27 Rolls-Royce Plc Wall elements for gas turbine engine combustors
US20100229563A1 (en) * 2006-01-25 2010-09-16 Woolford James R Wall elements for gas turbine engine combustors
EP2053312A2 (fr) * 2007-10-22 2009-04-29 Snecma Chambre de combustion à dilution optimisée et turbomachine en étant munie
US20100162712A1 (en) * 2007-11-29 2010-07-01 Honeywell International Inc. Quench jet arrangement for annular rich-quench-lean gas turbine combustors
US8616004B2 (en) * 2007-11-29 2013-12-31 Honeywell International Inc. Quench jet arrangement for annular rich-quench-lean gas turbine combustors
US20090139239A1 (en) * 2007-11-29 2009-06-04 Honeywell International, Inc. Quench jet arrangement for annular rich-quench-lean gas turbine combustors
US20100313571A1 (en) * 2007-12-29 2010-12-16 Alstom Technology Ltd Gas turbine
US8783044B2 (en) * 2007-12-29 2014-07-22 Alstom Technology Ltd Turbine stator nozzle cooling structure
FR2930591A1 (fr) * 2008-04-24 2009-10-30 Snecma Sa Optimisation du positionnement angulaire d'un distributeur de turbine en sortie d'une chambre de combustion de turbomachine
US20090266080A1 (en) * 2008-04-24 2009-10-29 Snecma Optimizing the angular positioning of a turbine nozzle at the outlet from a turbomachine combustion chamber
US8215118B2 (en) 2008-04-24 2012-07-10 Snecma Optimizing the angular positioning of a turbine nozzle at the outlet from a turbomachine combustion chamber
US20100077719A1 (en) * 2008-09-29 2010-04-01 Siemens Energy, Inc. Modular Transvane Assembly
US20110203282A1 (en) * 2008-09-29 2011-08-25 Charron Richard C Assembly for directing combustion gas
US8230688B2 (en) 2008-09-29 2012-07-31 Siemens Energy, Inc. Modular transvane assembly
US8276389B2 (en) 2008-09-29 2012-10-02 Siemens Energy, Inc. Assembly for directing combustion gas
US8087253B2 (en) * 2008-11-20 2012-01-03 General Electric Company Methods, apparatus and systems concerning the circumferential clocking of turbine airfoils in relation to combustor cans and the flow of cooling air through the turbine hot gas flowpath
US20100122538A1 (en) * 2008-11-20 2010-05-20 Wei Ning Methods, apparatus and systems concerning the circumferential clocking of turbine airfoils in relation to combustor cans and the flow of cooling air through the turbine hot gas flowpath
US20130086908A1 (en) * 2010-06-15 2013-04-11 Rolls-Royce Deutschland Ltd & Co Kg Gas turbine combustion chamber arrangement of axial type of construction
US9151223B2 (en) * 2010-06-15 2015-10-06 Rolls-Royce Deutschland Ltd & Co Kg Gas turbine combustion chamber arrangement of axial type of construction
US9151501B2 (en) 2011-07-28 2015-10-06 Rolls-Royce Deutschland Ltd & Co Kg Gas turbine centripetal annular combustion chamber and method for flow guidance
WO2013102584A1 (en) * 2012-01-05 2013-07-11 Siemens Aktiengesellschaft Combustion chamber of a combustor for a gas turbine
CN104040259A (zh) * 2012-01-05 2014-09-10 西门子公司 用于燃气涡轮机的燃烧器的燃烧室
US9885480B2 (en) 2012-01-05 2018-02-06 Siemens Aktiengesellschaft Combustion chamber of a combustor for a gas turbine
CN104040259B (zh) * 2012-01-05 2016-07-06 西门子公司 用于燃气涡轮机的燃烧器的燃烧室
EP2613080A1 (en) * 2012-01-05 2013-07-10 Siemens Aktiengesellschaft Combustion chamber of an annular combustor for a gas turbine
US9482432B2 (en) * 2012-09-26 2016-11-01 United Technologies Corporation Gas turbine engine combustor with integrated combustor vane having swirler
US20140338336A1 (en) * 2012-09-26 2014-11-20 United Technologies Corporation Gas turbine engine combustor with integrated combustor vane
US9590472B2 (en) 2013-02-15 2017-03-07 Siemens Aktiengesellschaft Through flow ventilation system for a power generation turbine package
US9322335B2 (en) 2013-03-15 2016-04-26 Siemens Energy, Inc. Gas turbine combustor exit piece with hinged connections
EP3076079A1 (en) * 2015-03-31 2016-10-05 Rolls-Royce plc Combustion equipment
US10208664B2 (en) 2015-03-31 2019-02-19 Rolls-Royce Plc Combustion equipment
US11175042B2 (en) 2015-03-31 2021-11-16 Rolls-Royce Plc Combustion equipment
EP3848556A1 (en) * 2020-01-13 2021-07-14 Ansaldo Energia Switzerland AG Gas turbine engine having a transition piece with inclined cooling holes

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Publication number Publication date
GB1239559A (enrdf_load_stackoverflow) 1971-07-21

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