WO2018190926A1 - Chambre de combustion à tourbillon piégé à cavité unique - Google Patents

Chambre de combustion à tourbillon piégé à cavité unique Download PDF

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Publication number
WO2018190926A1
WO2018190926A1 PCT/US2018/013460 US2018013460W WO2018190926A1 WO 2018190926 A1 WO2018190926 A1 WO 2018190926A1 US 2018013460 W US2018013460 W US 2018013460W WO 2018190926 A1 WO2018190926 A1 WO 2018190926A1
Authority
WO
WIPO (PCT)
Prior art keywords
wall
fuel nozzle
combustion chamber
outer liner
annular cavity
Prior art date
Application number
PCT/US2018/013460
Other languages
English (en)
Inventor
Beverly Stephenson Duncan
Eric John STEVENS
Walter Byron HOUCHENS
Arthur Wesley JOHNSON
John Herman MUELLER
Thomas George HOLLAND
Original Assignee
General Electric Company
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 General Electric Company filed Critical General Electric Company
Priority to CN201880024588.0A priority Critical patent/CN110494693B/zh
Publication of WO2018190926A1 publication Critical patent/WO2018190926A1/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
    • F23R3/02—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
    • 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
    • F23R3/28—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
    • F23R3/34—Feeding into different combustion zones
    • F23R3/343—Pilot flames, i.e. fuel nozzles or injectors using only a very small proportion of the total fuel to insure continuous combustion
    • 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
    • F23R3/50—Combustion chambers comprising an annular flame tube within an annular casing
    • 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
    • F23R2900/00—Special features of, or arrangements for continuous combustion chambers; Combustion processes therefor
    • F23R2900/00015—Trapped vortex combustion chambers

Definitions

  • the present subject matter relates generally to propulsion system combustion assemblies. More particularly, the present subject matter relates to trapped vortex combustor assemblies.
  • Propulsion systems such as gas turbine engines generally include combustion sections in which compressed air is mixed with a fuel and ignited to generate high pressure, high temperature combustion gases that then flow downstream and expand to drive a turbine section coupled to a compressor section, a fan section and/or a load device.
  • Conventional combustion sections are challenged to burn a variety of fuels of various caloric values.
  • Conventional combustion sections are also challenged to reduce emissions, such as nitric oxides, unburned hydrocarbons, and smoke, while also maintaining or improving combustion stability across a wider range of fuel/air ratios, air flow rates, and inlet pressures.
  • conventional combustion sections are challenged to achieve any or all of these criteria while maintaining or reducing longitudinal and/or radial dimensions and/or part quantities.
  • the present disclosure is directed to a combustor assembly for a propulsion system.
  • the combustor assembly includes an annular inner liner extended generally along the longitudinal direction.
  • the combustor assembly further include an annular outer liner defining a first wall at a first radius extended at least partially along the longitudinal direction, a second wall at a second radius extended at least partially along the longitudinal direction, and a transition wall therebetween coupling the first wall and the second wall between the first radius and the second radius.
  • the combustor assembly further includes a bulkhead wall disposed at the upstream end of the inner and outer liners. The bulkhead wall generally extends in the radial direction and couples the inner liner and the outer liner.
  • the bulkhead wall, the inner liner, and the outer liner together define a combustion chamber therebetween.
  • the bulkhead wall, the first wall of the outer liner, and the transition wall of the outer liner together define a single annular cavity of the combustion chamber.
  • the bulkhead wall defines a primary airflow opening radially inward of the single annular cavity.
  • the bulkhead wall and/or the outer liner defines one or more secondary airflow openings.
  • the combustor assembly further includes a fuel nozzle defining a fuel nozzle exit disposed adjacent to the single annular cavity of the combustion chamber.
  • the fuel nozzle is disposed through the bulkhead wall and adjacent to the single annular cavity of the combustion chamber.
  • the fuel nozzle is disposed through the first wall of the outer liner, in which the fuel nozzle exit is adjacent to the single annular cavity of the combustion chamber.
  • the fuel nozzle is disposed through the transition wall of the outer liner, in which the fuel nozzle exit is adjacent to the single annular cavity of the combustion chamber.
  • the combustor assembly further includes a secondary fuel nozzle defining a secondary fuel nozzle exit adjacent to the combustion chamber and radially inward of the single annular cavity.
  • the secondary fuel nozzle exit is disposed in circumferentially adjacent arrangement with the primary airflow opening of the bulkhead wall.
  • At least one of the secondary airflow openings is disposed on the first wall of the outer liner.
  • At least one of the secondary airflow openings is disposed on the transition wall of the outer liner.
  • At least one of the secondary airflow openings is disposed on the bulkhead wall adjacent to the single annular cavity of the combustion chamber.
  • the combustor assembly further includes a primary airflow tube extended generally along the longitudinal direction and coupled to the bulkhead wall.
  • the primary airflow tube is radially inward of the single annular cavity.
  • the primary airflow tube defines an inlet opening toward the upstream end of the primary airflow tube and the primary airflow opening of the bulkhead wall is defined through the bulkhead wall and the downstream end of the primary airflow tube.
  • the primary airflow tube extends at least partially along the circumferential direction to induce a circumferential swirl of air through the primary airflow tube into the combustion chamber.
  • the primary airflow tube defines a straight or non-swirl flow of air through the primary airflow tube into the combustion chamber.
  • the secondary airflow opening defines an angled or serpentine passage through the bulkhead wall or the outer liner at least partially along the circumferential direction to induce a circumferential swirl of air into the single annular cavity.
  • the present disclosure is further directed to a propulsion system defining a longitudinal direction, a radial direction, and a circumferential direction, and an upstream end and a downstream end opposite of the upstream end along the longitudinal direction.
  • the propulsion system further includes the combustor assembly.
  • the fuel nozzle is disposed through the bulkhead wall and adjacent to the single annular cavity of the combustion chamber. In another embodiment, the fuel nozzle is disposed through the first wall of the outer liner, and wherein the fuel nozzle exit is adjacent to the single annular cavity of the combustion chamber. In yet another embodiment, the fuel nozzle is disposed through the transition wall of the outer liner, and wherein the fuel nozzle exit is adjacent to the single annular cavity of the combustion chamber.
  • the engine further includes a secondary fuel nozzle defining a secondary fuel nozzle exit adjacent to the combustion chamber and radially inward of the single annular cavity.
  • the secondary fuel nozzle exit is disposed in circumferentially adjacent arrangement with the primary airflow opening of the bulkhead wall.
  • At least one of the secondary airflow openings is disposed on the first wall of the outer liner. In another embodiment, at least one of the secondary airflow openings is disposed on the transition wall of the outer liner. In still another embodiment, at least one of the secondary airflow openings is disposed on the bulkhead wall adjacent to the single annular cavity of the combustion chamber.
  • FIG. 1 is a schematic cross sectional view of an exemplary propulsion system incorporating an exemplary embodiment of a combustor assembly
  • FIG. 2 is an axial cross sectional view of an exemplary embodiment of a combustor assembly of the exemplary engine shown in FIG. 1 ;
  • FIG. 3 is an axial cross sectional view of another exemplary embodiment of a combustor assembly of the exemplary engine shown in FIG. 1 ;
  • FIG. 4 is an axial cross sectional view of yet another exemplary embodiment of a combustor assembly of the exemplary engine shown in FIG. 1; and
  • FIG. 5 is an axial cross sectional view of still another exemplary embodiment of a combustor assembly of the exemplary engine shown in FIG. 1.
  • first, second, and third may be used interchangeably to distinguish one component from another and are not intended to signify location or importance of the individual components.
  • upstream and downstream refer to the relative direction with respect to fluid flow in a fluid pathway.
  • upstream refers to the direction from which the fluid flows
  • downstream refers to the direction to which the fluid flows.
  • a single cavity trapped vortex combustor (TVC) for a propulsion system may improve fuel/air mixing, emissions output and improve combustion stability across a wider range of fuel/air ratios, air flow rates, and inlet pressures while also reducing combustion section dimensions.
  • the single cavity TVC shown and described herein may provide high combustor heat release in a short, compact package (i.e., reduced longitudinal and/or radial dimensions).
  • the single cavity TVC may provide a wide range of fuel/air ratios with single sheltered cavity fuel/air mixing and with or without bulk swirl introduction.
  • the single cavity TVC described herein may improve high altitude relight and lean blowout (LBO) with the single cavity structure.
  • Manufacturability of the single cavity TVC may further be improved over conventional TVC, annular, can-annular, or can combustors, thereby improving cost and maintainability. Still further, the single cavity TVC provided herein may simplify fuel introduction and mixing into the combustion chamber.
  • FIG. 1 is a schematic partially cross- sectioned side view of an exemplary propulsion system defining a high by -pass turbofan engine 10 herein referred to as "engine 10" as may incorporate various embodiments of the present disclosure.
  • engine 10 generally defines a longitudinal direction L, a radial direction R, and a circumferential direction C.
  • the engine 10 has a longitudinal or axial centerline axis 12 that extends there through for reference purposes.
  • the engine 10 may include a fan assembly 14 and a core engine 16 disposed downstream from the fan assembly 14.
  • the core engine 16 may generally include a substantially tubular outer casing 18 that defines an annular inlet 20.
  • the outer casing 18 encases or at least partially forms, in serial flow relationship, a compressor section having a booster or low pressure (LP) compressor 22, a high pressure (HP) compressor 24, a combustion section 26, a turbine section including a high pressure (HP) turbine 28, a low pressure (LP) turbine 30 and a jet exhaust nozzle section 32.
  • a high pressure (HP) rotor shaft 34 drivingly connects the HP turbine 28 to the HP compressor 24.
  • a low pressure (LP) rotor shaft 36 drivingly connects the LP turbine 30 to the LP compressor 22.
  • the LP rotor shaft 36 may also be connected to a fan shaft 38 of the fan assembly 14.
  • the LP rotor shaft 36 may be connected to the fan shaft 38 by way of a reduction gear 40 such as in an indirect- drive or geared-drive configuration.
  • the engine 10 may further include an intermediate pressure (IP) compressor and turbine rotatable with an intermediate pressure shaft.
  • IP intermediate pressure
  • the fan assembly 14 includes a plurality of fan blades 42 that are coupled to and that extend radially outwardly from the fan shaft 38.
  • An annular fan casing or nacelle 44 circumferentially surrounds the fan assembly 14 and/or at least a portion of the core engine 16.
  • the nacelle 44 may be supported relative to the core engine 16 by a plurality of circumferentially- spaced outlet guide vanes or struts 46.
  • at least a portion of the nacelle 44 may extend over an outer portion of the core engine 16 so as to define a bypass airflow passage 48 therebetween.
  • the engine 10 further includes a combustor assembly 100 including an annular inner liner 110, an annular outer liner 120, and a bulkhead wall 130.
  • the inner liner 110 extends generally along the longitudinal direction 1.
  • the outer liner 120 includes a first wall 122 at a first radius 121 and a second wall 124 at a second radius 123.
  • the first wall 122 extends at least partially along the longitudinal direction L.
  • the second wall 124 extends at least partially along the longitudinal direction L.
  • the outer liner 120 further includes a transition wall 126 extended between and coupling the first wall 122 and the second wall 124 from the first radius 121 and the second radius 123, respectively.
  • the bulkhead wall 130 is disposed at the upstream end 99 of the inner and outer liners 110, 120.
  • the bulkhead wall 130 generally extends in the radial direction R and couples the inner liner 1 10 and the outer liner 120.
  • the bulkhead wall 130, the inner liner 110, and the outer liner 120 together define a combustion chamber 140 therebetween.
  • the bulkhead wall 130, the first wall 122 of the outer liner 120, and the transition wall 126 of the outer liner 120 together define a single annular cavity 145 of the combustion chamber 140.
  • the bulkhead wall 130 further defines a primary airflow opening 135 inward along the radial direction R of the single annular cavity 145.
  • a plurality of primary airflow opening 135 may be arranged in adjacent arrangement along the circumferential direction C and/or the radial direction R.
  • the primary airflow opening 135 defines a generally round orifice, such as, but not limited to, a circular, ovular, or generally oblong orifice, and/or one or more polygonal orifices, or combinations thereof.
  • the bulkhead wall 130 and/or the outer liner 120 defines one or more secondary airflow openings 125.
  • the secondary airflow opening 125 may define a straight passage through the bulkhead wall 130 and/or the outer liner 120.
  • the secondary airflow opening 125 defines an angled or serpentine passage through the bulkhead wall 130 and/or the outer liner 120 that may induce a flow of air 83 at least partially along the circumferential direction C into the single annular cavity 145.
  • the combustor assembly 100 further includes a fuel nozzle 150 defining a fuel nozzle exit 151 disposed adjacent to the single annular cavity 145 of the combustion chamber 140.
  • the fuel nozzle 150 is disposed through the bulkhead wall 130 and is adjacent to the single annular cavity 145 of the combustion chamber 140.
  • the fuel nozzle exit 151 is disposed generally adjacent along the longitudinal direction L to the single annular cavity 145.
  • the fuel nozzle 150 is disposed through the first wall 122 of the outer liner 120.
  • the fuel nozzle exit 151 of the fuel nozzle 150 is adjacent generally along the radial direction R to the single annular cavity 145 of the combustion chamber 140.
  • the fuel nozzle 100 is disposed through the transition wall 126 of the outer liner 120.
  • the fuel nozzle exit 151 of the fuel nozzle 150 is adjacent generally along the longitudinal direction L to the single annular cavity 145 of the combustion chamber 140.
  • a flow of fuel from the fuel nozzle 150 generally egresses through the fuel nozzle exit 151 generally upstream toward the single annular cavity 145.
  • the fuel nozzle 150 may generally be arranged as shown and described in regard to FIGS. 2-4. However, the fuel nozzle 150 further includes a secondary fuel nozzle 155 defining a secondary fuel nozzle exit 156 adjacent generally along the longitudinal direction L to the combustion chamber 140 and inward along the radial direction R of the single annular cavity 145. In one embodiment, the secondary fuel nozzle exit 156 is disposed in adjacent arrangement along the circumferential direction with the primary airflow opening 135 through the bulkhead wall 130.
  • the embodiments of the combustion assembly 100 shown and described in regard to FIGS. 2-5 may be arranged in any combination of the embodiments.
  • several fuel nozzles 150 may be disposed adjacent to the combustion chamber 140 as shown and described in the several embodiments in regard to FIGS. 2-5.
  • the fuel nozzle 150 is disposed through the bulkhead wall 130 and another fuel nozzle 150 is disposed through the first wall 122 and/or the transition wall 126.
  • the fuel nozzle 150 is disposed through the first wall 122 and another fuel nozzle 150 is disposed through the transition wall 126 and/or the bulkhead wall 130.
  • the fuel nozzle 150 is disposed through the transition wall 126 and another fuel nozzle 150 is disposed through the bulkhead wall 130 and/or the first wall 122.
  • the secondary fuel nozzle 155 may further be included in any of the aforementioned combinations.
  • one or more of the secondary airflow openings 125 is defined through the first wall 122 of the outer liner 120. In another embodiment, one or more of the secondary airflow openings 125 is defined through the transition wall 126 of the outer liner 120. In still another embodiment, at least one of the secondary airflow openings 125 is disposed on the bulkhead wall 130 adjacent to the single annular cavity 145 of the combustion chamber 140.
  • the combustor assembly 100 may further include a first prediffuser passage 160.
  • the first prediffuser passage 160 is generally defined by a first inner prediffuser wall 162 and a first outer prediffuser wall 164.
  • Each first prediffuser wall 162, 164 extends generally along the longitudinal direction L at an approximately similar radius as the primary airflow opening 135 through the bulkhead wall 130.
  • the combustor assembly 100 may define a plurality of prediffuser passages.
  • the combustor assembly 100 may define the first prediffuser passage 160 and a second prediffuser passage 170 generally defined by a second inner prediffuser wall 172 and a second outer prediffuser wall 174.
  • Each second prediffuser wall 172, 174 extends generally along the longitudinal direction L and at least partially along the outward radial direction R.
  • prediffuser passages 160, 170 may be defined in the various embodiments shown and described in regard to FIGS. 2-5.
  • various embodiments may define the first prediffuser passage 160 or the first prediffuser passage 160 and the second prediffuser passage 170.
  • the combustor assembly 100 may define a plurality of prediffuser passages (e.g., a third passage, a fourth passage, etc.).
  • the combustor assembly 100 further includes a primary airflow tube 137 extended generally along the longitudinal direction L and coupled to the bulkhead wall 130.
  • the primary airflow tube 137 is inward along the radial direction R of the single annular cavity 145 of the combustion chamber 140.
  • the primary airflow tube 137 comprises walls defining an inlet opening 136 toward the upstream end of the primary airflow tube 137.
  • the primary airflow opening 135 of the bulkhead wall 130 is defined through the bulkhead wall 130 and the downstream end of the primary airflow tube 137.
  • the primary airflow tube 137 extends at least partially along the circumferential direction C (e.g., at an angle or as a serpentine structure) to induce a circumferential swirl of air through the primary airflow tube 137 into the combustion chamber 140.
  • the primary airflow tube 137 defines a generally straight or longitudinal passage to induce a straight flow or non-swirl of air through the primary airflow tube 137 into the combustion chamber 140.
  • a volume of air as indicated schematically by arrows 74 enters the engine 10 through an associated inlet 76 of the nacelle 44 and/or fan assembly 14.
  • Air 80 is progressively compressed as it flows through the LP and HP compressors 22, 24 towards the combustion section 26.
  • the now compressed air as indicated schematically by arrows 82 flows through the first prediffuser passage 160 and the second prediffuser passage 170 and into a pressure plenum 84 generally surrounding the combustion chamber 140 of the combustion section 26.
  • the compressed air 82 flows around and through the pressure plenum 84 and into the combustion chamber 140 through the primary airflow tube 137, as shown schematically by arrows 81 , and through the secondary airflow openings 125, as shown schematically by arrows 83.
  • a fuel such as a liquid or gaseous fuel, shown schematically by arrows 71 , flows through the fuel nozzle 150 and into the single annular cavity 145 of the combustion chamber 140.
  • the fuel 71 and the air 83 mix and ignite within the single annular cavity 145 of the combustion chamber 140.
  • the fuel 71 through the fuel nozzle 150 and air 83 through the secondary airflow openings 125 generally mix and generate a vortex within the single annular cavity 145 in which the fuel 71 and air 83 ignite, expand, and generally recirculate within the single annular cavity 145 as a generally uniform fuel/air mixture, thereby reducing undesired emissions in the combustion gases 86.
  • the air 81 through the primary airflow tube 137 may then flow the combustion gases from the fuel/air mixture within the single annular cavity 145 through the combustion chamber 140 and further downstream into the turbine section, such as shown schematically by combustion gases 86.
  • a second fuel may flow through the secondary fuel nozzle 155 and into the combustion chamber 140.
  • the secondary fuel nozzle 155 may generally provide a secondary fuel and air mixture, such as by mixing the second fuel 72 and the air 81 through the primary airflow tube 137 and in the combustion chamber 140.
  • the secondary fuel nozzle 155 may generally enable added combustion or engine operation range, or enable light-off across larger operating conditions, such as at altitude (e.g., up to approximately 16,000 meters or higher), or mitigate lean blowout, or enable further combustion stability off of an aero design point or equivalent.
  • the combustion gases 86 generated in the combustion chamber 62 flow from the combustor assembly 100 into the HP turbine 28, thus causing the HP rotor shaft 34 to rotate, thereby supporting operation of the HP compressor 24.
  • the combustion gases 86 are then routed through the LP turbine 30, thus causing the LP rotor shaft 36 to rotate, thereby supporting operation of the LP compressor 22 and/or rotation of the fan shaft 38.
  • the combustion gases 86 are then exhausted through the jet exhaust nozzle section 32 of the core engine 16 to provide propulsive thrust.
  • All or part of the combustor assembly may be part of a single, unitary component and may be manufactured from any number of processes commonly known by one skilled in the art. These manufacturing processes include, but are not limited to, those referred to as “additive manufacturing” or “3D printing”.
  • any number of casting, machining, welding, brazing, or sintering processes, or any combination thereof may be utilized to construct the combustor 100 separately or integral to one or more other portions of the combustion section 26.
  • the combustor assembly may constitute one or more individual components that are mechanically joined (e.g. by use of bolts, nuts, rivets, or screws, or welding or brazing processes, or combinations thereof) or are positioned in space to achieve a substantially similar geometric, aerodynamic, or thermodynamic results as if manufactured or assembled as one or more components.
  • suitable materials include high-strength steels, nickel and cobalt-based alloys, and/or metal or ceramic matrix composites, or combinations thereof.

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Pre-Mixing And Non-Premixing Gas Burner (AREA)

Abstract

La présente invention concerne un ensemble chambre de combustion pour un système de propulsion. L'ensemble chambre de combustion comprend un chemisage interne annulaire s'étendant globalement selon la direction longitudinale. L'ensemble chambre de combustion comprend en outre un chemisage externe annulaire définissant une première paroi avec un premier rayon s'étendant au moins partiellement selon la direction longitudinale, une seconde paroi avec un second rayon s'étendant au moins partiellement selon la direction longitudinale, et une paroi de transition entre celles-ci, accouplant la première paroi et la seconde paroi entre le premier rayon et le second rayon. L'ensemble chambre de combustion comprend en outre une paroi de cloison disposée à l'extrémité amont des chemisages interne et externe. La paroi de cloison s'étend globalement dans la direction radiale et accouple le chemisage interne et le chemisage externe. La paroi de cloison, le chemisage interne et le chemisage externe définissent ensemble une chambre de combustion entre ceux-ci. La paroi de cloison, la première paroi du chemisage externe et la paroi de transition du chemisage externe définissent ensemble une cavité annulaire unique de la chambre de combustion. La paroi de cloison définit une ouverture d'écoulement d'air primaire radialement vers l'intérieur de la cavité annulaire unique. La paroi de cloison et/ou le chemisage externe définit une ou plusieurs ouvertures d'écoulement d'air secondaires. L'ensemble chambre de combustion comprend en outre une buse de carburant définissant une sortie de buse de carburant disposée adjacente à la cavité annulaire unique de la chambre de combustion.
PCT/US2018/013460 2017-04-13 2018-01-12 Chambre de combustion à tourbillon piégé à cavité unique WO2018190926A1 (fr)

Priority Applications (1)

Application Number Priority Date Filing Date Title
CN201880024588.0A CN110494693B (zh) 2017-04-13 2018-01-12 单腔捕获涡流燃烧器

Applications Claiming Priority (2)

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US201715486526A 2017-04-13 2017-04-13
US15/486,526 2017-04-13

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EP3739265A1 (fr) * 2019-05-17 2020-11-18 Raytheon Technologies Corporation Chambre de combustion monolithique pour applications de moteur pouvant résister à une attrition
US11578869B2 (en) 2021-05-20 2023-02-14 General Electric Company Active boundary layer control in diffuser

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US11725817B2 (en) * 2021-06-30 2023-08-15 General Electric Company Combustor assembly with moveable interface dilution opening

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US20130318991A1 (en) * 2012-05-31 2013-12-05 General Electric Company Combustor With Multiple Combustion Zones With Injector Placement for Component Durability

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JPH11264540A (ja) * 1997-12-18 1999-09-28 General Electric Co <Ge> ベンチュリレススワールカップ
US9068751B2 (en) * 2010-01-29 2015-06-30 United Technologies Corporation Gas turbine combustor with staged combustion
US20120131924A1 (en) * 2010-11-30 2012-05-31 Hitachi, Ltd. Gas Turbine Combustor and Fuel Supply Method Used for the Same
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Cited By (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP3739265A1 (fr) * 2019-05-17 2020-11-18 Raytheon Technologies Corporation Chambre de combustion monolithique pour applications de moteur pouvant résister à une attrition
US11136901B2 (en) 2019-05-17 2021-10-05 Raytheon Technologies Corporation Monolithic combustor for attritiable engine applications
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US11578614B2 (en) 2019-05-17 2023-02-14 Raytheon Technologies Corporation Monolithic combustor for attritiable engine applications
US11578869B2 (en) 2021-05-20 2023-02-14 General Electric Company Active boundary layer control in diffuser

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CN110494693A (zh) 2019-11-22
CN110494693B (zh) 2020-12-29

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