WO2002088604A1 - Turbine a gaz pourvue d'une combinaison de chambres tubulaires et annulaires et procede permettant de faire fonctionner une turbine a gaz - Google Patents

Turbine a gaz pourvue d'une combinaison de chambres tubulaires et annulaires et procede permettant de faire fonctionner une turbine a gaz Download PDF

Info

Publication number
WO2002088604A1
WO2002088604A1 PCT/EP2002/004310 EP0204310W WO02088604A1 WO 2002088604 A1 WO2002088604 A1 WO 2002088604A1 EP 0204310 W EP0204310 W EP 0204310W WO 02088604 A1 WO02088604 A1 WO 02088604A1
Authority
WO
WIPO (PCT)
Prior art keywords
gas turbine
combustion
hot gas
turbine according
type
Prior art date
Application number
PCT/EP2002/004310
Other languages
English (en)
Inventor
Gerhard Simon
Wilhelm Schulten
David Amos
Peter Tiemann
Original Assignee
Siemens Aktiengesellschaft
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 Siemens Aktiengesellschaft filed Critical Siemens Aktiengesellschaft
Priority to JP2002585864A priority Critical patent/JP2004524508A/ja
Priority to EP02730176A priority patent/EP1381813A1/fr
Publication of WO2002088604A1 publication Critical patent/WO2002088604A1/fr

Links

Classifications

    • 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
    • F23R3/42—Continuous combustion chambers using liquid or gaseous fuel characterised by the arrangement or form of the flame tubes or combustion chambers
    • F23R3/46—Combustion chambers comprising an annular arrangement of several essentially tubular flame tubes within a common annular casing or within individual casings
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23D—BURNERS
    • F23D2210/00—Noise abatement

Definitions

  • the invention is related to the field of gas turbines and, more particularly, to operating conditions of gas turbines, where combustion oscillations may occur.
  • Gas turbines are widely used for power generation as well as in jet engine aircrafts.
  • air and fuel are incinerated in a combustor, leading to the generation of a hot gas stream which expands in a turbine part, thereby rotating a shaft.
  • a part of the so generated rotational energy is used for driving a compressor and providing compressed air for the combustor.
  • Combustion oscillations are acoustical modes that arise from a positive coupling between pressure fluctuations caused by flame instabilities and acoustical reflections from combustor walls.
  • the combustion oscillations produce high noise emissions and may lead to extensive damage to gas turbine parts.
  • German published patent application DE 43 39 094 Al sets out a method of suppressing combustion oscillations in a gas turbine by injecting an inert fluid into the combustion zone, thereby redistributing the combustion zone and disturbing the mechanism that establishes the combustion oscillation.
  • This method is applied to an annular combustor which is specifically susceptible to combustion oscillations because of the large space usually defined by this geometry.
  • the method belongs to the category of active instability control, because combustion oscillations are monitored and suppressed with a closed loop active control cycle.
  • active instability control is a complex and expensive method, introducing a significant cost increase for the equipment needed to operate the gas turbine.
  • WO 00/12939 sets out a gas turbine with an annular combustor, to which a plurality of burners are connected.
  • a passive measure Cylindrical means are installed around the orifices of several of the burners, which leads to a shift of the flame position and therefore to a detuning of acoustical modes occurring in the combustor.
  • Passive measures are generally less expensive than active instability control, however, only certain frequency bands are normally suppressed by the static means.
  • US 5309710 sets out a can-type combustor for a gas turbine.
  • Can-type combustors are less prone to combustion oscillations.
  • the can-type construction is more complex and expensive and not as easy to service as an annular combustor.
  • an annular combustor provides a more homogenous temperature and pressure distribution at the turbine inlet, which is unavailable in the can-type combustors.
  • the present invention provides a gas turbine and a method of operating a gas turbine which substantially avoids combustion oscillations.
  • the present invention also advantageously provides a gas turbine and a method of operating a gas turbine which leads to a homogenous pressure and temperature distribution of a hot gas stream at a turbine inlet. Additionally, the present invention provides a gas turbine and a method of operating a gas turbine which is easy to service. It is another advantage of the invention to provide a method which can be implemented so as to enable mass production at reasonable engineering effort and expense and with maximally replicable component characteristics. As a further advantage, the present invention provides a gas turbine and a method of operating a gas turbine which requires comparably low cooling.
  • the present invention further advantageously provides a gas turbine and a method of operating a gas turbine which requires comparably low complexity of construction and therefore leads to lower costs. Additionally, the present invention provides a gas turbine and a method of operating a gas turbine which is comparably stable in operation. As a further advantage, the invention provides a gas turbine and a method of operating a gas turbine which lowers sealing requirements for sealing against the hot gas.
  • the present invention provides a gas turbine, comprising a combustor for combustion of an air/fuel-mixture, thereby generating a hot gas stream, and a turbine with a turbine inlet through which the hot gas stream is conducted, wherein the combustor comprises an annular part, located next to the turbine inlet and a can-type part, comprising a plurality of can-type combustors, each connected to the annular part and opened to the annular part through a respective can outlet.
  • the combination of a can-type combustor with an annular combustor combines advantages of the respective single concepts, thereby avoiding disadvantages of the respective single concepts.
  • the can-type combustors each have a head-part and a sidewall extending from said head-part to said can outlet. Through the head part fuel and air is introduced into the can-type combustors.
  • a main combustion zone is defined as being the zone of a visible flame.
  • the extension of the sidewall of each can-type combustor is preferably sized such that each can-type combustor encloses essentially a complete main combustion zone.
  • the extension of the sidewall of each can-type combustor is sized such that it does not significantly exceed the main combustion zone.
  • the annular part has an annular part axial extension, measured along the direction of said hot gas stream and the can-type part has a can-type part axial extension, measured along the direction of the hot gas stream and the annular part axial extension is preferably at least a half of the can-type part axial extension. This assures a sufficient merger of the individual hot gas streams from the can-type combustors, resulting in a high degree of homogenization of temperature and pressure.
  • the turbine inlet is at least partially formed by a vane support structure, carrying vanes located in the turbine and the annular part is preferably connected to and thereby supported by the vane support structure.
  • Each of the can-type combustors is preferably connected to the annular part and thereby supported by the annular part. Accordingly, sealing of clearances between the can- type combustors and the annular part come out to be comparably easy because of lack of relative movement of these parts against one another. Moreover, the support of the can-type combustors is less complex compared to combustors which have only cans, because those usually require support from an outer casing. Accordingly, costs are reduced.
  • each of the connections of the can-type combustors to the annular part is designed in such a manner that each of the can- type combustors is individually easily removable from the annular part, thereby leading to a construction that is easy to service by simply exchanging single can-type combustors .
  • the can-type combustors are preferably equally distributed along the circumference of the annular part, thereby defining hollow spaces between respective sidewalls of adjacent can-type combustors. The hollow spaces can be used for introducing a cooling fluid for cooling of the can-type combustors.
  • the annular part is preferably designed with a radically outward located outer shell and a radially inward located inner shell, wherein ribs, extending from the outer shell to the inner shell, are located next to the can-type part and wherein each of said can-type combustors is connected with its side wall to at least one of said ribs.
  • the sidewalls are diffuser-shaped; thereby diffusing the hot gas stream discharged from the can-type combustors into the annular part .
  • the present invention provides a gas turbine with a combustor for combustion of an air/fuel-mixture, thereby generating a main hot gas stream and a turbine with a turbine inlet through which the main hot gas stream is conducted, wherein the combustor comprises a first stage in which individual hot gas streams are generated in distinct main combustion zones, which are defined to be the region of a visible flame, each combustion zone being separated from another, wherein the first region of the combustor is followed by a second region in which the individual hot gas streams are brought together to form the main hot gas stream.
  • the second region is preferably designed to assure appropriate intermixing of the individual hot gas streams, thereby homogenizing the temperature distribution in the main hot gas stream.
  • the present invention provides a method for reducing combustion oscillations in a gas turbine, the method comprising the following steps: conducting air and fuel to a plurality of distinct combustion zones, incinerating the air and fuel in the distinct combustion zones which isolated from one another, thereby generating individual hot gas streams and thereby substantially avoiding acoustical coupling of flame instabilities occurring in the distinct combustion zones, merging the individual hot gas streams to a main hot gas stream thereby substantially homogenizing the temperature and pressure distribution and feeding the main hot gas stream through a turbine inlet.
  • the advantages of this method correspond to the foregoing described advantages of the gas turbine .
  • FIG. 1 is a schematic view on a gas turbine
  • FIG. 2 illustrates combustion in an annular combustor
  • FIG. 3 illustrates schematically an improvement of the present invention over that shown in FIG. 2
  • FIG 4 illustrates a plurality of can-type combustors connected to an annular combustor
  • FIG. 5 illustrates in a longitudinal section a gas turbine with a can-type combustor connected to an annular combustor.
  • FIG. 6 and 7 illustrate in a three dimensional view can- type combustors connected to an annular combustor.
  • FIG. 1 illustrates schematically a gas turbine 1.
  • the gas turbine 1 comprises of a compressor 3, a combustor 5 and a turbine 7.
  • the turbine 7 is mounted on a shaft 9 together with the compressor 3.
  • air 11 is compressed and conducted to the combustor 5.
  • fuel 13 is conducted to the compressor 5, being mixed with the air 11 to give an air/fuel mixture 15.
  • the air/fuel mixture 15 is incinerated, thereby generating a hot gas stream 17.
  • the hot gas stream 17 is conducted through a turbine inlet 19 into the turbine 7, where through an arrangement of vanes and blades (not shown) energy is transferred to the shaft 9 which rotates.
  • the rotational energy is partly used for driving the compressor 3. Another part of the rotational energy could be used, for instance, for driving an electric generator.
  • FIG. 2 illustrates an annular combustor 5.
  • Burners 21 function as suppliers of air/fuel mixtures 15 which are incinerated in a visible flame 25, defining a main combustion zone 23. Flame instabilities can occur, resulting in pressure pulses which in turn can induce more flame instabilities. In particular, adjacent burners 21 can have mutually acoustic coupling, thereby mutually inducing flame instabilities indicated by arrows 27. By reflection of pressure pulses from the combustor wall, stable combustion oscillations can arise from the flame instabilities .
  • FIG. 3 illustrates a novel concept for the combustor 5.
  • the combustor 5 comprises a can-type part 33 and an annular part 35.
  • the can-type part 33 comprises of a plurality of can-type combustors 31, each connected to the annular part 35 and opened to the annular part 35 through respective can outlets 37. Through respective head-parts 39 of each can-type combustor 31, air 11 and fuel 13 is introduced, thereby forming an air/fuel mixture 15.
  • a sidewall 41 extends from the head part of each can-type combustor 31 to the respective can outlet 37.
  • the air/fuel mixture 15 is incinerated in the main combustion zones 23, defined by visible flames 25, in each of the can-type combustors 31, thereby each generating an individual hot gas stream 42.
  • the individual or component hot gas streams 42 are merged in the annular part 35 to form the main hot gas stream 17.
  • the temperature and pressure of the main hot gas stream 17 turns out to be well homogenized before entering the turbine 7.
  • cross- ignition between the different flames 23 is established which results in improved stability of operation, even under extreme lean combustion conditions.
  • the air 11 and fuel 13 undergo at least in part a premixing to form the air/fuel mixture 15, thereby achieving a lean combustion which is, if appropriate, stabilized by a diffusion flame.
  • FIG. 4 illustrates in a three-dimensional view more particularly the assembly of the annular part 35. It is formed of an outer shell 51 and an inner shell 53 which enclose an annular shaped space 54. The inner surface of the outer shell 51 as well as the inner shell 53 is covered by heat shield elements 55 for protection against hot gas.
  • the can-type combustors 31 of the can-type part 33 may be distributed substantially equally along a circumferential direction 57 of the annular part 35.
  • FIG. 5 illustrates a longitudinal section of the combustor 5 supported in the gas turbine 1.
  • the compressor 3 has blades 61 and vanes 63 which, from a rotation of the shaft 9, compress and conduct air 11 through a compressor diffuser 65 to a space 67 that is formed from an outer casing 69.
  • the compressed air 11 is mainly conducted to the combustor 5 but a part of the air 11 is used for cooling the combustor 5 and turbine 7.
  • the outer casing 69 also encloses a vane support structure 71 of the turbine 7. This vane support structure 71 supports vanes 73. Between two respective rows of vanes 73, blades 75 are mounted on the shaft 9.
  • the vane support structure 71 also supports the annular part 35 of the combustor 5 which is attached to the vane support structure 71. Accordingly, very little relative movement between the annular part 35 and the turbine 7 occurs which leads to an easy sealing against discharge of hot gas out of clearances between the annular part 35 and the turbine 7. Moreover, the can-type combustors 31 are attached to the annular part 35 which again leads to very little relative movement between the annular part 35 and the can-type combustors 31 and, accordingly, to an easy sealing against discharge of hot gas out of clearances. Furthermore, without the need of supporting the combustor 5 from the outer casing 69, a less complex and less expensive construction is achieved. In a direction of the hot gas stream, the can-type part 33 has an can-type part extension C and the annular part 35 has an annular part extension A which is at least one half of the can-type part extension C.
  • FIG. 6 and FIG. 7 illustrate in a three-dimensional view, the connection of the can-type combustors 31 to the annular part 35.
  • the annular part 35 has ribs 91, extending from the outer shell 51 to the inner shell 53.
  • the can-type combustors 31 are connected to these ribs 91 with their sidewalls 41.
  • the sidewalls 41 are diffuser shaped and have flanges 85 with which they are connected to the annular part 35. Every single can-type combustor 31 is easily removable for service.
  • the head parts 39 have openings 83 in which a burner can be mounted.
  • FIG. 6 illustrates furthermore the can-type combustors 31 having a double wall structure which comprises an outer wall 101 and an inner wall 103, enclosed by the outer wall 101. Between the outer wall 101 and the inner wall 103 hollow space or opening 105 is formed which provides for conduction of a cooling fluid.
  • a cooling fluid In particular, steam could be used as a cooling fluid, separated in a closed cooling cycle from the hot gas inside the combustor.
  • the inner wall 103 may be manufactured as a cylinder 103a with an on welded flange 103b .
  • the inner wall 103 is mounted within the outer wall 101 by guidance elements 107 and fixation in a fork-like support 109. Between adjacent can-type combustors 31, the space 105 for cooling fluid is narrowed.

Landscapes

  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Abstract

L'invention concerne une turbine à gaz pourvue d'une chambre de combustion qui comprend une pluralité d'éléments de chambres tubulaires et un élément annulaire. Une combustion isolée se produit dans la chambre tubulaire, générant ainsi des flux individuels constitutifs de gaz chaud. Ces flux constitutifs de gaz chaud se mélangent dans la partie annulaire. Pour éviter sensiblement toute oscillation de combustion, on découple les flammes de combustion, ce qui permet de réaliser l'homogénéisation de la température et de la pression ainsi que la stabilisation de la combustion par intercommunication d'allumage.
PCT/EP2002/004310 2001-04-27 2002-04-18 Turbine a gaz pourvue d'une combinaison de chambres tubulaires et annulaires et procede permettant de faire fonctionner une turbine a gaz WO2002088604A1 (fr)

Priority Applications (2)

Application Number Priority Date Filing Date Title
JP2002585864A JP2004524508A (ja) 2001-04-27 2002-04-18 ガスタービンとその運転方法
EP02730176A EP1381813A1 (fr) 2001-04-27 2002-04-18 Turbine a gaz pourvue d'une combinaison de chambres tubulaires et annulaires et procede permettant de faire fonctionner une turbine a gaz

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
US09/842,881 2001-04-27
US09/842,881 US20020157400A1 (en) 2001-04-27 2001-04-27 Gas turbine with combined can-type and annular combustor and method of operating a gas turbine

Publications (1)

Publication Number Publication Date
WO2002088604A1 true WO2002088604A1 (fr) 2002-11-07

Family

ID=25288477

Family Applications (1)

Application Number Title Priority Date Filing Date
PCT/EP2002/004310 WO2002088604A1 (fr) 2001-04-27 2002-04-18 Turbine a gaz pourvue d'une combinaison de chambres tubulaires et annulaires et procede permettant de faire fonctionner une turbine a gaz

Country Status (5)

Country Link
US (1) US20020157400A1 (fr)
EP (1) EP1381813A1 (fr)
JP (1) JP2004524508A (fr)
CN (1) CN1246639C (fr)
WO (1) WO2002088604A1 (fr)

Families Citing this family (20)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE50108163D1 (de) * 2001-05-01 2005-12-29 Alstom Technology Ltd Baden Schwingungsreduktion in einer Brennkammer
EP1284391A1 (fr) * 2001-08-14 2003-02-19 Siemens Aktiengesellschaft Chambre de combustion pour turbines à gaz
EP1398569A1 (fr) * 2002-09-13 2004-03-17 Siemens Aktiengesellschaft Turbine à gaz
JP2008064405A (ja) * 2006-09-08 2008-03-21 Mitsubishi Heavy Ind Ltd ガスタービン燃焼器
FR2911669B1 (fr) * 2007-01-23 2011-09-16 Snecma Carenage pour chambre de combustion, chambre de combustion en etant equipee et turboreacteur les comportant.
US20100192578A1 (en) * 2009-01-30 2010-08-05 General Electric Company System and method for suppressing combustion instability in a turbomachine
US20110072826A1 (en) * 2009-09-25 2011-03-31 General Electric Company Can to can modal decoupling using can-level fuel splits
US20110088379A1 (en) * 2009-10-15 2011-04-21 General Electric Company Exhaust gas diffuser
JP5156066B2 (ja) * 2010-08-27 2013-03-06 株式会社日立製作所 ガスタービン燃焼器
US9249687B2 (en) 2010-10-27 2016-02-02 General Electric Company Turbine exhaust diffusion system and method
CH704829A2 (de) * 2011-04-08 2012-11-15 Alstom Technology Ltd Gasturbogruppe und zugehöriges Betriebsverfahren.
US20130081407A1 (en) * 2011-10-04 2013-04-04 David J. Wiebe Aero-derivative gas turbine engine with an advanced transition duct combustion assembly
US9709279B2 (en) 2014-02-27 2017-07-18 General Electric Company System and method for control of combustion dynamics in combustion system
US9845956B2 (en) * 2014-04-09 2017-12-19 General Electric Company System and method for control of combustion dynamics in combustion system
US9909507B2 (en) 2015-01-27 2018-03-06 General Electric Company Control system for can-to-can variation in combustor system and related method
JP6563312B2 (ja) * 2015-11-05 2019-08-21 川崎重工業株式会社 ガスタービンエンジンの抽気構造
JP6625427B2 (ja) * 2015-12-25 2019-12-25 川崎重工業株式会社 ガスタービンエンジン
US11143108B2 (en) * 2019-03-07 2021-10-12 Pratt & Whitney Canada Corp. Annular heat shield assembly for combustor
CN113416583B (zh) * 2021-07-29 2022-06-24 北京工商大学 一种生物质原位控氮气化联产热炭装置
EP4394253A1 (fr) * 2022-12-30 2024-07-03 Ansaldo Energia Switzerland AG Moteur à turbine à gaz

Citations (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2447482A (en) * 1945-04-25 1948-08-24 Westinghouse Electric Corp Turbine apparatus
FR1401071A (fr) * 1963-07-18 1965-05-28 Westinghouse Electric Corp Générateur à combustion
US5309710A (en) 1992-11-20 1994-05-10 General Electric Company Gas turbine combustor having poppet valves for air distribution control
DE4339094A1 (de) 1993-11-16 1995-05-18 Abb Management Ag Verfahren zur Dämpfung von thermoakustischen Schwingungen sowie Vorrichtung zur Durchführung des Verfahrens
US5974781A (en) * 1995-12-26 1999-11-02 General Electric Company Hybrid can-annular combustor for axial staging in low NOx combustors
WO2000012939A1 (fr) 1998-08-31 2000-03-09 Siemens Aktiengesellschaft Ensemble bruleur
US6182451B1 (en) * 1994-09-14 2001-02-06 Alliedsignal Inc. Gas turbine combustor waving ceramic combustor cans and an annular metallic combustor
US6374594B1 (en) * 2000-07-12 2002-04-23 Power Systems Mfg., Llc Silo/can-annular low emissions combustor

Patent Citations (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2447482A (en) * 1945-04-25 1948-08-24 Westinghouse Electric Corp Turbine apparatus
FR1401071A (fr) * 1963-07-18 1965-05-28 Westinghouse Electric Corp Générateur à combustion
US5309710A (en) 1992-11-20 1994-05-10 General Electric Company Gas turbine combustor having poppet valves for air distribution control
DE4339094A1 (de) 1993-11-16 1995-05-18 Abb Management Ag Verfahren zur Dämpfung von thermoakustischen Schwingungen sowie Vorrichtung zur Durchführung des Verfahrens
US6182451B1 (en) * 1994-09-14 2001-02-06 Alliedsignal Inc. Gas turbine combustor waving ceramic combustor cans and an annular metallic combustor
US5974781A (en) * 1995-12-26 1999-11-02 General Electric Company Hybrid can-annular combustor for axial staging in low NOx combustors
WO2000012939A1 (fr) 1998-08-31 2000-03-09 Siemens Aktiengesellschaft Ensemble bruleur
US6374594B1 (en) * 2000-07-12 2002-04-23 Power Systems Mfg., Llc Silo/can-annular low emissions combustor

Also Published As

Publication number Publication date
CN1505745A (zh) 2004-06-16
CN1246639C (zh) 2006-03-22
JP2004524508A (ja) 2004-08-12
EP1381813A1 (fr) 2004-01-21
US20020157400A1 (en) 2002-10-31

Similar Documents

Publication Publication Date Title
US20020157400A1 (en) Gas turbine with combined can-type and annular combustor and method of operating a gas turbine
US6983600B1 (en) Multi-venturi tube fuel injector for gas turbine combustors
JP4570136B2 (ja) ガスタービン用燃焼器とガスタービンエンジン
EP2211111B1 (fr) Agencement des buses d'injection à multi-tubes regroupés pour turbomachine
EP2613083B1 (fr) Ensemble de buse à combustible à utiliser avec un moteur à turbine et son procédé d'assemblage
EP3052786B1 (fr) Panneau de protection contre la chaleur doté d'assemblages par recouvrement pour chambre de combustion de moteur de turbine
JP4641648B2 (ja) モジュール式燃焼器ドーム
US4413470A (en) Catalytic combustion system for a stationary combustion turbine having a transition duct mounted catalytic element
EP2669580B1 (fr) Ensemble d'injection de carburant à utiliser dans les moteurs à turbine et procédé d'assemblage correspondant
EP1486732B1 (fr) Chambre de combustion avec un revêtement flottant
US5117624A (en) Fuel injector nozzle support
JP2005098678A (ja) ガスタービンエンジンのエミッションを低減するための方法及び装置
CA2672502C (fr) Corps central d'injecteur de carburant et methode d'assemblage
US5996352A (en) Thermally decoupled swirler for a gas turbine combustor
US11280495B2 (en) Gas turbine combustor fuel injector flow device including vanes
JP2005037122A (ja) ガスタービンエンジンの燃焼器を冷却するための方法及び装置
EP1058061B1 (fr) Chambre de combustion de turbines à gaz
EP0767868B1 (fr) Support d'injecteur de carburant d'un recipient sous pression, destine a un moteur industriel a turbine a gaz
US20200041129A1 (en) Mixer Assembly for a Combustor
US11953205B2 (en) Swirler with fuel manifold, and combustor and gas turbine including the same
JP2017516007A (ja) ガスタービン燃焼器の固有振動数を変化させる予混合器アセンブリ及び機構
CN110736108A (zh) 用于热发动机的燃烧器组件
KR102146564B1 (ko) 연소공진 완화구조를 구비한 연소기 및 가스터빈
JPH06213446A (ja) チップを冷却する燃料噴射ノズル
WO2023014343A1 (fr) Chambre de combustion dans un moteur à turbine à gaz

Legal Events

Date Code Title Description
AK Designated states

Kind code of ref document: A1

Designated state(s): CN JP

AL Designated countries for regional patents

Kind code of ref document: A1

Designated state(s): AT BE CH CY DE DK ES FI FR GB GR IE IT LU MC NL PT SE TR

121 Ep: the epo has been informed by wipo that ep was designated in this application
DFPE Request for preliminary examination filed prior to expiration of 19th month from priority date (pct application filed before 20040101)
WWE Wipo information: entry into national phase

Ref document number: 2002730176

Country of ref document: EP

WWE Wipo information: entry into national phase

Ref document number: 2002585864

Country of ref document: JP

WWE Wipo information: entry into national phase

Ref document number: 028087917

Country of ref document: CN

WWP Wipo information: published in national office

Ref document number: 2002730176

Country of ref document: EP

WWW Wipo information: withdrawn in national office

Ref document number: 2002730176

Country of ref document: EP