EP2631544A1 - Pilote de prémélange annulaire de buse d'injection de carburant - Google Patents

Pilote de prémélange annulaire de buse d'injection de carburant Download PDF

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Publication number
EP2631544A1
EP2631544A1 EP20130156828 EP13156828A EP2631544A1 EP 2631544 A1 EP2631544 A1 EP 2631544A1 EP 20130156828 EP20130156828 EP 20130156828 EP 13156828 A EP13156828 A EP 13156828A EP 2631544 A1 EP2631544 A1 EP 2631544A1
Authority
EP
European Patent Office
Prior art keywords
fuel
air
nozzle
centerbody
premix
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.)
Granted
Application number
EP20130156828
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German (de)
English (en)
Other versions
EP2631544B1 (fr
Inventor
Jong Ho Uhm
Derrick Walter Simons
Gregory Allen Boardman
Bryan Wesley Romig
Kara Edwards
Michael John Hughes
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General Electric Co
Original Assignee
General Electric Co
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Filing date
Publication date
Application filed by General Electric Co filed Critical General Electric Co
Publication of EP2631544A1 publication Critical patent/EP2631544A1/fr
Application granted granted Critical
Publication of EP2631544B1 publication Critical patent/EP2631544B1/fr
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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
    • F23R3/28—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
    • F23R3/286—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply having fuel-air premixing devices
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23C—METHODS OR APPARATUS FOR COMBUSTION USING FLUID FUEL OR SOLID FUEL SUSPENDED IN  A CARRIER GAS OR AIR 
    • F23C7/00—Combustion apparatus characterised by arrangements for air supply
    • F23C7/002—Combustion apparatus characterised by arrangements for air supply the air being submitted to a rotary or spinning motion
    • F23C7/004—Combustion apparatus characterised by arrangements for air supply the air being submitted to a rotary or spinning motion using vanes
    • 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
    • F23C—METHODS OR APPARATUS FOR COMBUSTION USING FLUID FUEL OR SOLID FUEL SUSPENDED IN  A CARRIER GAS OR AIR 
    • F23C2900/00—Special features of, or arrangements for combustion apparatus using fluid fuels or solid fuels suspended in air; Combustion processes therefor
    • F23C2900/07001—Air swirling vanes incorporating fuel injectors
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23D—BURNERS
    • F23D14/00—Burners for combustion of a gas, e.g. of a gas stored under pressure as a liquid
    • F23D14/02—Premix gas burners, i.e. in which gaseous fuel is mixed with combustion air upstream of the combustion zone
    • F23D14/04—Premix gas burners, i.e. in which gaseous fuel is mixed with combustion air upstream of the combustion zone induction type, e.g. Bunsen burner
    • F23D14/08—Premix gas burners, i.e. in which gaseous fuel is mixed with combustion air upstream of the combustion zone induction type, e.g. Bunsen burner with axial outlets at the burner head
    • 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
    • F23R2900/03343—Pilot burners operating in premixed mode
    • 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
    • F23R3/10—Air inlet arrangements for primary air
    • F23R3/12—Air inlet arrangements for primary air inducing a vortex
    • F23R3/14—Air inlet arrangements for primary air inducing a vortex by using swirl vanes

Definitions

  • the present invention generally involves a gas turbine engine that combusts a hydrocarbon fuel mixed with air to generate a high temperature gas stream that drives turbine blades to rotate a shaft attached to the blades and more particularly to the engine's fuel nozzle having a pilot nozzle that premixes fuel and air while achieving lower nitrogen oxides.
  • Gas turbine engines are widely used to generate power for numerous applications.
  • a conventional gas turbine engine includes a compressor, a combustor, and a turbine.
  • the compressor provides compressed air to the combustor.
  • the air entering the combustor is mixed with fuel and combusted.
  • Hot gases of combustion are exhausted from the combustor and flow into the blades of the turbine so as to rotate the shaft of the turbine connected to the blades. Some of that mechanical energy of the rotating shaft drives the compressor and/or other mechanical systems.
  • a gas turbine engine may employ one or more fuel nozzles to intake air and fuel to facilitate fuel/air mixing in the combustor.
  • the fuel nozzles may be located in a head end portion of the combustor, and may be configured to intake an air flow to be mixed with a fuel input.
  • each fuel nozzle may be internally supported by a center body located inside of the fuel nozzle, and a pilot can be mounted at the downstream end of the center body.
  • a so-called swozzle can be mounted to the exterior of the center body and located upstream from the pilot.
  • the swozzle has curved vanes that extend radially from the center body across an annular flow passage and from which fuel is introduced into the annular flow passage to be entrained into a flow of air that is swirled by the vanes of the swozzle.
  • a premix pilot also can be used as a pilot to stabilize the pilot flame, even in low fuel to air ratio to prevent an increase in NOx.
  • a fuel nozzle includes premix conduits that are configured with concentric axes that are parallel to the burner tube axis to direct the fuel/air mixture axially from the premixed pilot nozzle, there is at least one air jet beside each of the premix conduits wherein each of the air jets is directed radially outwardly from the premix conduits so that each of the air jets can entrain some portion of fuel/air mixture to direct the mixture radially outward from the premix conduits to form a more uniform fuel/air mixture in the burn exit plane of the nozzle and into the combustion chamber of a gas turbine.
  • the air jets desirably are disposed at the downstream end of an annular channel that is disposed radially outwardly of the premix conduits.
  • each of the premix conduits is itself configured concentrically about a central longitudinal axis disposed at an acute angle with respect to the axis of burner tube to direct the fuel/air mixture radially outwardly about the axis of burner tube to form a more uniform fuel/air mixture in the burn exit plane of the nozzle and into the combustion chamber of a gas turbine.
  • each of the premix conduits is itself configured concentrically about a bi-directed central longitudinal axis that has a first leg disposed parallel to the axis of burner tube and a second leg disposed at an acute angle with respect to the axis of burner tube to direct the fuel/air mixture radially outwardly about the axis of burner tube to form a more uniform fuel/air mixture in the burn exit plane of the nozzle and into the combustion chamber of a gas turbine.
  • the fuel nozzle is itself configured concentrically about a central longitudinal axis disposed at an acute angle with respect to the axis of burner tube to direct the fuel/air mixture radially outwardly about the axis of burner tube, there is at least one air jet beside each of the premix conduits so that each of the air jets can entrain some portion of fuel/air mixture to further direct the mixture radially outwardly from the premix conduits to form a more uniform fuel/air mixture in the burn exit plane of the nozzle and into the combustion chamber of a gas turbine.
  • one or more of the air jets desirably is directed radially outwardly from the premix conduits to entrain some portion of fuel/air mixture to further direct the mixture radially outwardly from the premix conduits to form a more uniform fuel/air mixture in the burn exit plane of the nozzle and into the combustion chamber of a gas turbine.
  • ranges and limits mentioned herein include all sub-ranges located within the prescribed limits, inclusive of the limits themselves unless otherwise stated.
  • a range from 100 to 200 also includes all possible sub-ranges, examples of which are from 100 to 150, 170 to 190, 153 to 162, 145.3 to 149.6, and 187 to 200.
  • a limit of up to 7 also includes a limit of up to 5, up to 3, and up to 4.5, as well as all sub-ranges within the limit, such as from about 0 to 5, which includes 0 and includes 5 and from 5.2 to 7, which includes 5.2 and includes 7.
  • the turbine system 10 may use liquid or gas fuel, such as natural gas and/or a hydrogen rich synthetic gas, to run the turbine system 10.
  • liquid or gas fuel such as natural gas and/or a hydrogen rich synthetic gas
  • a plurality of fuel nozzles 12 of the type described more fully below intakes a fuel supply 14, mixes the fuel with air, and distributes the air-fuel mixture into a combustor 16.
  • the air-fuel mixture combusts in a chamber within the combustor 16, thereby creating hot pressurized exhaust gases.
  • the combustor 16 directs the exhaust gases through a turbine 18 toward an exhaust outlet 20.
  • the shaft 22 may be connected to various components of the turbine system 10, including a compressor 24.
  • the compressor 24 also includes blades that may be coupled to the shaft 22.
  • the blades within the compressor 24 also rotate, thereby compressing air from an air intake 26 through the compressor 24 and into the fuel nozzles 12 and/or combustor 16.
  • the shaft 22 also may be connected to a load 28, which may be a vehicle or a stationary load, such as an electrical generator in a power plant or a propeller on an aircraft, for example.
  • the load 28 may include any suitable device capable of being powered by the rotational output of turbine system 10.
  • FIG. 2 is a simplified drawing of cross sectional views of several portions of the gas turbine system 10 schematically depicted in FIG. 1 .
  • the turbine system 10 includes one or more fuel/air nozzles 12 located in a head end portion 27 of one or more combustors 16 in the gas turbine engine.
  • Each illustrated fuel nozzle 12 may include multiple fuel nozzles integrated together in a group and/or a standalone fuel nozzle, wherein each illustrated fuel nozzle 12 relies at least substantially or entirely on internal structural support (e.g., load bearing fluid passages).
  • the system 10 comprises a compressor section 24 for pressurizing a gas, such as air, flowing into the system 10 via air intake 26.
  • air enters the turbine system 10 through the air intake 26 and may be pressurized in the compressor 24.
  • the gas may be any gas suitable for use in a gas turbine system 10.
  • Pressurized air discharged from the compressor section 24 flows into a combustor section 16, which is generally characterized by a plurality of combustors 16 (only one of which is illustrated in FIGs. 1 and 2 ) disposed in an annular array about an axis of the system 10.
  • the air entering the combustor section 16 is mixed with fuel and combusted within the combustion chamber 32 of the combustor 16.
  • the fuel nozzles 12 may inject a fuel/air mixture into the combustor 16 in a suitable fuel-air ratio for optimal combustion, emissions, fuel consumption, and power output.
  • the combustion generates hot pressurized exhaust gases, which then flow from each combustor 16 to a turbine section 18 ( FIG. 1 ) to drive the system 10 and generate power.
  • the hot gases drive one or more blades (not shown) within the turbine 18 to rotate the shaft 22 and, thus, the compressor 24 and the load 28.
  • the rotation of the shaft 22 causes blades 30 within the compressor 24 to rotate and draw in and pressurize the air received by the intake 26.
  • a combustor 16 need not be configured as described above and illustrated herein and in general may have any configuration that permits pressurized air to be mixed with fuel, combusted and transferred to a turbine section 18 of the system 10.
  • FIGs. 3 - 9 schematically illustrate various embodiments of fuel/air nozzles 12 in accordance with exemplary embodiments of the present invention.
  • at least each upstream end of each premix conduit 41 defines a central axis 41d configured and disposed so that the flow of fluid that enters the entrance opening 66a of each premix conduit 41 is directed parallel to the central axis 36 of the centerbody 52.
  • FIGs. 4 , 8 and 9 for example, at least each upstream end of each premix conduit 41 defines a central axis 41d configured and disposed so that the flow of fluid that enters the entrance opening 66a of each premix conduit 41 is directed parallel to the central axis 36 of the centerbody 52.
  • each upstream end of each premix conduit 41 defines a central axis 41d configured and disposed so that the flow of fluid that enters the entrance opening 66a of each premix conduit 41 is directed at an acute angle away from the central axis 36 of the centerbody 52 in the range of 0.1 degrees to twenty degrees.
  • the fuel/air nozzle 12 includes premix conduits 41 that are configured with concentric axes that are parallel to the burner tube axis 36 to direct the fuel/air mixture axially from the premixed pilot nozzle 40. In the embodiments that are schematically shown in FIGs.
  • each of the air jets 42 can entrain some portion of fuel/air mixture to direct the mixture radially outwardly from the premix conduits 41 to form a more uniform fuel/air mixture in the burn exit plane 44 of the nozzle 12 and into the combustion chamber 32 of the combustor 16 ( FIGs. 1 and 2 ).
  • each of the air jets 42 can entrain some portion of fuel/air mixture to direct the mixture radially outwardly from the premix conduits 41 to form a more uniform fuel/air mixture in the burn exit plane 44 of the nozzle 12 and into the combustion chamber 32 of the combustor 16 ( FIGs. 1 and 2 ).
  • each of the air jets 42 desirably is directed radially outwardly from the premix conduits 41 so that each of the air jets 42 more readily can entrain more of the fuel/air mixture to direct more of the mixture radially outwardly from the premix conduits 41 to form a more uniform fuel/air mixture in the burn exit plane 44 of the nozzle 12 and into the combustion chamber 32 of the combustor 16 ( FIGs. 1 and 2 ).
  • a fuel/air nozzle 12 for a gas turbine engine desirably includes an axially elongating peripheral wall 50 defining an outer envelope of the nozzle 12.
  • the peripheral wall 50 of fuel/air nozzle 12 has an outer surface 50a and an inner surface 50b facing opposite the outer surface 50a and defining an axially elongating inner cavity 50c.
  • a fuel/air nozzle 12 for a gas turbine engine desirably includes a hollow, axially elongating centerbody 52 disposed within the inner cavity 50c of the fuel/air nozzle 12 and defining a central axis 36, which is the same as the central axis of the burner tube.
  • the centerbody 52 is defined by a centerbody wall 52a that defines an upstream end 52b and a downstream end 52c disposed axially opposite the upstream end 52b.
  • the centerbody wall 52a is defined by an exterior surface 52d and an interior surface 52e facing opposite the exterior surface 52d.
  • the interior surface 52d of the centerbody wall 52a defines an axially elongating interior passage 53 that is disposed concentrically about the central axis 36 of the centerbody 52.
  • a primary air flow channel 51 is defined in the annular space between the inner surface 50b of the peripheral wall 50 and the exterior surface 52d of the centerbody wall 52a.
  • a fuel/air nozzle 12 for a gas turbine engine desirably includes an axially elongated, hollow fuel supply line 54 extending axially through the interior passage 53 of the centerbody 52.
  • the fuel supply line 54 has an upstream end 54a that is disposed at the upstream end 52b of the centerbody 52 and that is configured for connection to a source of fuel (not shown).
  • the fuel supply line 54 has a downstream end 54b that is disposed at the downstream end 52c of the centerbody 52.
  • a secondary air flow channel is defined by an annular space between the interior surface 52e of the centerbody 52 and the exterior surface of the fuel supply line 54, and that annular space defines the axially elongating interior passage 53 schematically shown in FIG. 3 .
  • the primary fuel may be supplied to the combustion chamber 32 of the combustor 16 ( FIG. 2 ) through a plurality of air swirler vanes 56 that are fixed and extend across the flow path of the primary air flow channel 51.
  • These air swirler vanes 56 define a so-called swozzle that extends radially from the exterior surface of centerbody wall 52a.
  • each of the air swirler vanes 56 of the swozzle desirably is provided with internal fuel conduits 57 that terminate in fuel injection ports or holes 58 from which primary fuel (indicated by the arrows designated by the numeral 57a) flowing from the conduits 57 can be injected into the primary air (indicated by the arrows designated by the numeral 51a) that flows past the fuel injection ports 58 in the air swirler vanes 56.
  • a swirling pattern is imparted to the primary air flow 51a that facilitates the mixing of the primary air flow 51a with the primary fuel that is ejected from the holes 58 of the air swirler vanes 56 into the passing primary air flow 51a.
  • the primary air flow 51a mixed with the primary fuel then may flow into the pre-mixing annulus 51 that is defined between the peripheral wall 50 and the inner centerbody 52, wherein the primary air flow 51a and the primary fuel continue to mix together prior to entering the combustion chamber 32.
  • a fuel/air nozzle 12 for a gas turbine engine desirably includes a premixed pilot nozzle 40 that has an upstream end 40a connected to the downstream end 52c of the centerbody 52.
  • the downstream end 52c of the centerbody 52 and the upstream end 40a of the premixed pilot nozzle 40 are defined in part by different sections of the metal cylinder that forms the centerbody 52 and the outermost wall that defines the premixed pilot nozzle 40.
  • the premixed pilot nozzle 40 has a downstream end 40b disposed axially opposite the upstream end 40a of the premixed pilot nozzle 40.
  • the premixed pilot nozzle 40 defines a pilot fuel nozzle 60, which defines an upstream end 60a and a downstream end 60b.
  • the upstream end 60a of the pilot fuel nozzle 60 is connected in fluid communication with the downstream end 54b of the fuel supply line 54.
  • the downstream end 60b of the pilot fuel nozzle 60 defines at least one fuel jet 61 configured in fluid communication with the upstream end 60a of the pilot fuel nozzle 60.
  • fuel 62 entering the pilot fuel nozzle 60 via the downstream end 54b of the fuel supply line 54 exits the pilot fuel nozzle 60 via a plurality of fuel jets 61 (shown in dashed line in FIG. 5 ).
  • the premixed pilot nozzle 40 further defines an annular-shaped fuel plenum wall 63 disposed radially outwardly from the pilot fuel nozzle 60 and desirably concentrically with respect to the burner tube axis 36 shown in FIG. 3 .
  • the fuel plenum wall 63 defines a fuel plenum 64 between the pilot fuel nozzle 60 and the fuel plenum wall 63.
  • the fuel plenum wall 63 further defines a plurality of fuel holes 63a through which fuel is discharged from the fuel plenum 64. As schematically shown in FIGs.
  • At least one of the fuel holes 63a is connected in fluid communication with at least one of the fuel jets 61 via the fuel plenum 64.
  • each of the fuel holes 63a is connected in fluid communication with each of the fuel jets 61 via the fuel plenum 64.
  • the premixed pilot nozzle 40 further defines a plurality of axially elongated, hollow premix conduits 41 disposed radially outwardly from the fuel plenum wall 63.
  • each of the premix conduits 41 desirably is defined in part by the plenum wall 63.
  • each premix conduit 41 has an upstream end 41a that is disposed near the downstream end 52c of the centerbody 52.
  • FIGs. 3 schematically shown in FIGs.
  • each premix conduit 41 defines at its upstream end 41a an entrance opening 66a that communicates fluidly with the interior passage 53 of the centerbody 52.
  • the arrows designated 53a in FIGs. 4 and 9 schematically indicate the flow of air 53a that enters the entrance opening 66a of each premix conduit 41 from the interior passage 53 of the centerbody 52.
  • each premix conduit 41 is connected in fluid communication with at least one of the fuel holes 63a that is defined in the fuel plenum wall 63 of the fuel plenum 64.
  • the arrows designated 62a in FIGs. 4 and 9 schematically indicate the flow of fuel 62a that exits from the fuel plenum 64 through the fuel holes 63a defined in the fuel plenum wall 63 and passes into each premix conduit 41.
  • the arrows designated 62b in FIGs. 4 and 9 schematically indicate the flow of the fuel-air mix 62b that travels downstream in the premix conduits 41 of the premixed pilot nozzle 40.
  • each premix conduit 41 has a downstream end 41b that is disposed axially opposite the upstream end 41a of the premix conduit 41 and that is disposed near the downstream end 40b of the premixed pilot nozzle 40.
  • each downstream end 41b of each premix conduit 41 defines an exit opening 66b that allows fluid, i.e., the fuel-air mix 62b, to discharge from the hollow premix conduit 41.
  • each downstream end 41b of each premix conduit 41 defines a central axis 41c, and the walls that define each premix conduit 41 are disposed concentrically about this central axis 41c.
  • each central axis 41c is a straight line that is disposed so that the flow of the fluid that is the mix of fuel and air that discharges from the exit opening 66b of each premix conduit 41 is directed parallel to the central axis 36 of the centerbody 52.
  • Various embodiments of the present invention include features that counteract this much higher equivalence ratio that exists at the section of the burn exit plane 44 ( FIGs. 4 and 9 ) that is located immediately downstream of the exit opening 66b of each premix conduit 41.
  • the premixed pilot nozzle 40 further defines an annular channel 70 that is disposed radially outwardly from the premix conduits 41.
  • the inner wall 43 of the annular channel 70 also serves to define the outer wall 43 of the premix conduits 41.
  • the centerbody wall 52a also serves to define the outer wall 52a of the annular channel 70.
  • the upstream end of the annular channel 70 is configured to communicate fluidly with the interior passage 53 of the centerbody 52 and thus receives a flow of air from the interior passage 53 of the centerbody 52.
  • a plurality of air jets 42 is defined. At least one of the air jets 42 is disposed nearby at least one of the exit openings 66b of at least one of the premix conduits 41. Desirably, as shown in FIGs. 4, 5 and 9 for example, at least one of the air jets 42 is disposed nearby each of the exit openings 66b of each one of the premix conduits 41.
  • each of the passages that defines one of the air jets 42 is defined concentrically around a central axis 42a that is disposed parallel to the central axis 41c of the downstream end 41b of the nearby premix conduit 41.
  • the higher velocity air leaving each air jets 42 entrains some of the fuel/air mixture leaving the premix conduits 41 and serves to direct the fuel/air mixture that exists at the section of the burn exit plane 44 ( FIG. 9 ) that is located immediately downstream of the exit opening 66b of each premix conduit 41.
  • the overall result is a more uniform equivalence ratio at the burn exit plane 44 ( FIG. 9 ) of the fuel-air nozzle 12, and this also can be used by lighting source as a pilot.
  • each of the passages that defines one of the air jets 42 is defined concentrically around a central axis 42a that is disposed at an acute angle with respect to the central axis 41c of the downstream end 41b of the nearby premix conduit 41.
  • the magnitude of this acute angle desirably is in the range of 0.1 degrees to twenty degrees.
  • the air jet 42 in the FIG. 4 embodiment desirably is configured and disposed so that the respective air jet 42 directs the flow of air exiting the air jet 42 in a direction away from the nearby one of the exit openings 66b of the premix conduits 41.
  • each air jet 42 is pointed so that the air leaving that air jet 42 moves in a direction that is both downstream and radially away from the central axis 36 of the centerbody 52 as well as radially away from the central axis 41c of downstream end 41b of the respective nearby premix conduit 41.
  • the air leaving each air jets 42 entrains some of the fuel/air mixture leaving the premix conduits 41 and serves to direct the fuel/air mixture that exists at the section of the burn exit plane 44 ( FIG. 4 ) that is located immediately downstream of the exit opening 66b of each premix conduit 41 radially outwardly toward the peripheral wall 50 ( FIG. 3 ).
  • the overall result is a more uniform equivalence ratio at the burn exit plane 44 ( FIG. 4 ) of the fuel-air nozzle 12, and this also can be used by lighting source as a pilot.
  • each of the premix conduits 41 is itself configured concentrically about a central longitudinal axis 41c that is disposed at an acute angle with respect to the central axis 36 of the centerbody 52.
  • the magnitude of this acute angle desirably is in the range of 0.1 degrees to twenty degrees.
  • each of the premix conduits 41 is configured and disposed so as to direct the fuel/air mixture exiting from the exit opening 66b of each premix conduit 41 radially outwardly away from the central axis 36 of the centerbody 52 so as to form a more uniform fuel/air mixture in the burn exit plane 44 of the fuel-air nozzle 12 and into the combustion chamber of a gas turbine 10 and so as to continuously and effectively light the fuel/air mixture that is supplied through an axially elongating inner cavity 50c ( FIG. 3 ) of the fuel-air nozzle 12. Though not specifically illustrated in FIG.
  • the axial length of one or more of the premix conduits 41 can extend beyond the downstream end 40b of the premixed pilot nozzle 40 in a manner concentrically around the central axis 41c of the downstream end 41b of premix conduit 41.
  • each of the premix conduits 41 is itself configured concentrically about a bi-directed central longitudinal axis having a first leg 41d and a second leg 41c.
  • FIG. 8 As schematically in FIG.
  • the first leg 41d of the bi-directed central longitudinal axis is disposed parallel to the central axis 36 of the centerbody 52 and extends between the upstream end 41a and the downstream end 41b of each premix conduit 41.
  • a second leg 41c of the bi-directed central longitudinal axis is disposed at an acute angle with respect to the central axis 36 of the centerbody 52 and extends through the downstream end 41b of each premix conduit 41.
  • the magnitude of this acute angle desirably is in the range of 0.1 degrees to twenty degrees.
  • the fuel-air nozzle 8 is configured and disposed so as to direct the fuel/air mixture exiting from the exit opening 66b radially outwardly away from the central axis 36 of the centerbody 52 so as to form a more uniform fuel/air mixture in the burn exit plane 44 of the fuel-air nozzle 12 and into the combustion chamber of a gas turbine 12 and so as to continuously and effectively light the fuel/air mixture that is supplied through an axially elongating inner cavity 50c ( FIG. 3 ) of the fuel-air nozzle 12.
  • each of the premix conduits 41 is itself configured concentrically about a central longitudinal axis 41c that is disposed at an acute angle with respect to the central axis 36 of the centerbody 52.
  • the magnitude of this acute angle desirably is in the range of 0.1 degrees to twenty degrees.
  • each of the premix conduits 41 is configured and disposed so as to direct the fuel/air mixture exiting from the exit opening 66b of each premix conduit 41 radially outwardly away from the central axis 36 of the centerbody 52 so as to form a more uniform fuel/air mixture in the burn exit plane 44 of the fuel-air nozzle 12 and into the combustion chamber of a gas turbine 10 and so as to continuously and effectively light the fuel/air mixture that is supplied through an axially elongating inner cavity 50c ( FIG. 3 ) of the fuel-air nozzle 12.
  • the upstream end of the annular channel 70 is configured to taper as it proceeds in the downstream direction toward the air jets 42 that are disposed beside each of the exit openings of premix conduits 41. As described above with regard to the embodiment of FIGs.
  • each of the air jets 42 is directed radially outwardly from the premix conduits 41 so that each of the air jets 42 can entrain some portion of fuel/air mixture to further direct the mixture radially outwardly from the premix conduits 41 to form a more uniform fuel/air mixture in the burn exit plane 44 of the fuel-air nozzle 12 and into the combustion chamber 32 of a combustor 16 ( FIGs. 1 and 2 ) and to continuously and effectively light the fuel/air mixture that is supplied through an axially elongating inner cavity 50c ( FIG. 3 ) of the fuel-air nozzle 12.
  • Each embodiment of the premixed pilot nozzle 40 provides a small well anchored premixed flame near the base of the fuel-air nozzle 12, thus anchoring the swirling fuel air mixture exiting the fuel-air nozzle 12.
  • the improved flame stability enables lower fuel/air operations, thus extending LBO and the emissions operability window.
  • One embodiment of such a method of operating a fuel/air nozzle 12 for a gas turbine engine 10 desirably includes the following steps schematically shown in FIG. 10 : the step 81 of delivering a primary flow of air 51a (e.g., FIG. 3 ) downstream past the swozzle to swirl the primary flow of air; the step 82 of delivering a primary flow of fuel 57a (e.g., FIG.
  • FIG. 11 Another embodiment of such a method of operating a fuel/air nozzle 12 for a gas turbine engine 10 desirably includes the following steps schematically shown in FIG. 11 : the step 81 of delivering a primary flow of air 51a (e.g., FIG. 3 ) downstream past the swozzle to swirl the primary flow of air; the step 82 of delivering a primary flow of fuel 57a (e.g., FIG. 3 ) through the swozzle to mix with the swirled primary flow of air 51a downstream of the swozzle; the step 83 of delivering a flow of pilot fuel 62 (e.g., FIG.
  • the pressure of the air expelled from the annular channel 70 exceeds the pressure of the air entering the annular channel 70.
  • the annular channel 70 includes an upstream end that is configured to taper as it proceeds in the downstream direction.
  • a further embodiment of the method comprises the step of expelling the fuel/air mixture from the premix conduits 41 of the premixed pilot nozzle 40 away from the central axis 36 of the centerbody 52.
  • a further embodiment of the method comprises the step of expelling the fuel/air mixture 62b from the premix conduits 41 of the premixed pilot nozzle 40 in a direction that is parallel to the central axis 36 of the centerbody 52.
  • a further embodiment of the method comprises the step of expelling the fuel/air mixture 62b from the annular channel 70 in a direction that is radially away from the central axis 36 of the centerbody 52.

Landscapes

  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Nozzles For Spraying Of Liquid Fuel (AREA)
EP13156828.9A 2012-02-27 2013-02-26 Pilote de prémélange annulaire de buse d'injection de carburant Active EP2631544B1 (fr)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US13/405,550 US20130219899A1 (en) 2012-02-27 2012-02-27 Annular premixed pilot in fuel nozzle

Publications (2)

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EP2631544A1 true EP2631544A1 (fr) 2013-08-28
EP2631544B1 EP2631544B1 (fr) 2018-05-09

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Country Status (5)

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US (2) US20130219899A1 (fr)
EP (1) EP2631544B1 (fr)
JP (1) JP6186132B2 (fr)
CN (1) CN103438480B (fr)
RU (1) RU2621566C2 (fr)

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GB2539536A (en) * 2015-04-16 2016-12-21 Gen Electric Fuel nozzle assembly including a pilot nozzle
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EP3150918A4 (fr) * 2014-05-30 2018-01-10 Kawasaki Jukogyo Kabushiki Kaisha Dispositif de combustion pour turbine à gaz
US10775047B2 (en) 2014-05-30 2020-09-15 Kawasaki Jukogyo Kabushiki Kaisha Combustor for gas turbine engine
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Also Published As

Publication number Publication date
RU2621566C2 (ru) 2017-06-06
US20130219899A1 (en) 2013-08-29
JP2013174431A (ja) 2013-09-05
CN103438480A (zh) 2013-12-11
JP6186132B2 (ja) 2017-08-23
CN103438480B (zh) 2016-08-24
EP2631544B1 (fr) 2018-05-09
US20150253011A1 (en) 2015-09-10
RU2013108313A (ru) 2014-09-10

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