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WO2014113105A2 - Injecteur de combustible pour moteur à turbine à gaz - Google Patents

Injecteur de combustible pour moteur à turbine à gaz Download PDF

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Publication number
WO2014113105A2
WO2014113105A2 PCT/US2013/065453 US2013065453W WO2014113105A2 WO 2014113105 A2 WO2014113105 A2 WO 2014113105A2 US 2013065453 W US2013065453 W US 2013065453W WO 2014113105 A2 WO2014113105 A2 WO 2014113105A2
Authority
WO
WIPO (PCT)
Prior art keywords
fuel
fuel nozzle
recited
helical
tip
Prior art date
Application number
PCT/US2013/065453
Other languages
English (en)
Other versions
WO2014113105A3 (fr
Inventor
Roger O. COFFEY
Original Assignee
United Technologies Corporation
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 United Technologies Corporation filed Critical United Technologies Corporation
Publication of WO2014113105A2 publication Critical patent/WO2014113105A2/fr
Publication of WO2014113105A3 publication Critical patent/WO2014113105A3/fr

Links

Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23D—BURNERS
    • F23D11/00—Burners using a direct spraying action of liquid droplets or vaporised liquid into the combustion space
    • F23D11/10—Burners using a direct spraying action of liquid droplets or vaporised liquid into the combustion space the spraying being induced by a gaseous medium, e.g. water vapour
    • F23D11/106—Burners using a direct spraying action of liquid droplets or vaporised liquid into the combustion space the spraying being induced by a gaseous medium, e.g. water vapour medium and fuel meeting at the burner outlet
    • F23D11/107—Burners using a direct spraying action of liquid droplets or vaporised liquid into the combustion space the spraying being induced by a gaseous medium, e.g. water vapour medium and fuel meeting at the burner outlet at least one of both being subjected to a swirling motion
    • 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
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23D—BURNERS
    • F23D2900/00—Special features of, or arrangements for burners using fluid fuels or solid fuels suspended in a carrier gas
    • F23D2900/00016—Preventing or reducing deposit build-up on burner parts, e.g. from carbon
    • 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
    • F23R2900/00004—Preventing formation of deposits on surfaces of gas turbine components, e.g. coke deposits

Definitions

  • the present disclosure relates to a gas turbine engine and, more particularly, to a fuel nozzle tip assembly.
  • Gas turbine engines such as those which power modern commercial and military aircraft, include a compressor for pressurizing a supply of air, a combustor for burning a hydrocarbon fuel in the presence of the pressurized air, and a turbine for extracting energy from the resultant combustion gases.
  • the combustor generally includes radially spaced apart inner and outer liners that define an annular combustion chamber therebetween.
  • a plurality of circumferentially distributed fuel injectors axially project into a forward section of the combustion chamber to supply the fuel to be mixed with the pressurized air.
  • a fuel nozzle tip assembly may need to be geometrically configured to inject/atomize the fuel into the combustor air stream
  • a fuel nozzle for a combustor of a gas turbine engine includes a swirler tip along an axis; and a helical adapter along the axis and within the swirler tip.
  • a further embodiment of the present disclosure includes, wherein the helical adapter defines a central passage for a primary fuel flow.
  • a further embodiment of any of the foregoing embodiments of the present disclosure includes, wherein the central passage directs the primary fuel flow into an atomizer tip.
  • a further embodiment of any of the foregoing embodiments of the present disclosure includes, wherein the helical adapter includes a multiple of helical channels for a secondary fuel flow.
  • a further embodiment of any of the foregoing embodiments of the present disclosure includes, wherein the multiple of helical channels are adjacent to the swirler tip.
  • a further embodiment of any of the foregoing embodiments of the present disclosure includes, wherein the multiple of helical channels are adjacent to the swirler tip.
  • a further embodiment of any of the foregoing embodiments of the present disclosure includes, wherein the swirler tip includes a multiple of radial orifices.
  • a further embodiment of any of the foregoing embodiments of the present disclosure includes a heat shield that at least partially surrounds the swirler tip.
  • a further embodiment of any of the foregoing embodiments of the present disclosure includes a wear sleeve that at least partially surrounds the heat shield.
  • a further embodiment of any of the foregoing embodiments of the present disclosure includes a valve housing around an entire external surfaces of a valve for a secondary fuel flow between the helical adapter and the swirler tip.
  • a further embodiment of any of the foregoing embodiments of the present disclosure includes, wherein the wherein the helical adapter defines a central passage for a primary fuel flow and the helical adapter includes a multiple of helical channels for the secondary fuel flow.
  • a method of directing a fuel gas and a liquid through a fuel nozzle and into a combustor of a gas turbine engine includes directing a primary fuel flow through a helical insert; and directing a secondary fuel flow around the helical insert.
  • a further embodiment of any of the foregoing embodiments of the present disclosure includes directing the primary fuel flow through an atomizer tip.
  • a further embodiment of any of the foregoing embodiments of the present disclosure includes directing the secondary fuel flow through a multiple of radial orifices.
  • a further embodiment of any of the foregoing embodiments of the present disclosure includes directing the secondary fuel flow through a multiple of helical channels.
  • Figure 1 is a schematic cross-section of an example gas turbine engine architecture
  • Figure 2 is a schematic cross-section of another example gas turbine engine architecture
  • Figure 3 is an expanded longitudinal schematic sectional view of a combustor section according to one non-limiting embodiment
  • Figure 4 is an isometric view of a fuel injector
  • Figure 5 is a sectional view of the fuel injector of Figure 4.
  • Figure 6 is an expanded sectional view of a fuel nozzle tip assembly according to one disclosed non-limiting embodiment
  • Figure 7 is a perspective sectional view of the fuel nozzle tip assembly according to another disclosed non-limiting embodiment.
  • Figure 8 is a schematic sectional view of a valve for the fuel nozzle according to one disclosed non-limiting embodiment.
  • FIG. 1 schematically illustrates a gas turbine engine 20.
  • the gas turbine engine 20 is disclosed herein as a two-spool turbo fan that generally includes a fan section 22, a compressor section 24, a combustor section 26 and a turbine section 28.
  • the fan section 22 drives air along a bypass flowpath and into the compressor section 24.
  • the compressor section 24 drives air along a core flowpath for compression and communication into the combustor section 26, which then expands and directs the air through the turbine section 28.
  • turbofan Although depicted as a turbofan in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with turbofans as the teachings may be applied to other types of turbine engines such as a low bypass augmented turbofan, turbojets, turboshafts, and three-spool (plus fan) turbofans with an intermediate spool. Still other aeroderivative engine architectures 20A are located within an enclosure 30 (Figure 2) typical of an industrial gas turbine (IGT).
  • IGT industrial gas turbine
  • the combustor section 26 generally includes a combustor 50 with an outer combustor wall assembly 60, an inner combustor wall assembly 62 and a diffuser case 64.
  • the outer combustor wall assembly 60 and the inner combustor wall assembly 62 are spaced apart such that a combustion chamber 66 is defined therebetween.
  • the combustion chamber 66 may be generally annular in shape.
  • the outer combustor liner assembly 60 is spaced radially inward from an outer diffuser case 64A of the diffuser case module 64 to define an outer annular plenum 76.
  • the inner combustor liner assembly 62 is spaced radially outward from an inner diffuser case 64B of the diffuser case module 64 to define an inner annular plenum 78. It should be understood that although a particular combustor is illustrated, other combustor types with various combustor liner arrangements will also benefit herefrom. It should be further understood that the disclosed cooling flow paths are but an illustrated embodiment and should not be limited only thereto.
  • the combustor wall assemblies 60, 62 contain the combustion products for direction toward the turbine section 28.
  • Each combustor wall assembly 60, 62 generally includes a respective support shell 68, 70 which supports one or more liner panels 72, 74 mounted to a hot side of the respective support shell 68, 70.
  • Each of the liner panels 72, 74 may be generally rectilinear and manufactured of, for example, a nickel based super alloy, ceramic or other temperature resistant material and are arranged to form a liner array.
  • the liner array includes a multiple of forward liner panels 72A and a multiple of aft liner panels 72B that are circumferentially staggered to line the hot side of the outer shell 68.
  • a multiple of forward liner panels 74A and a multiple of aft liner panels 74B are circumferentially staggered to line the hot side of the inner shell 70.
  • the combustor 50 further includes a forward assembly 80 immediately downstream of the compressor section 24 to receive compressed airflow therefrom.
  • the forward assembly 80 generally includes an annular hood 82 and a bulkhead assembly 84 which locate a multiple of fuel nozzles 86 (one shown) and a multiple of swirlers 90 (one shown).
  • Each of the swirlers 90 is mounted within an opening 92 of the bulkhead assembly 84 to be circumferentially aligned with one of a multiple of annular hood ports 94 along an axis F.
  • Each bulkhead assembly 84 generally includes a bulkhead support shell 96 secured to the combustor wall assembly 60, 62, and a multiple of circumferentially distributed bulkhead liner panels 98 secured to the bulkhead support shell 96.
  • the annular hood 82 extends radially between, and is secured to, the forwardmost ends of the combustor wall assemblies 60, 62.
  • the annular hood 82 forms the multiple of circumferentially distributed hood ports 94 that accommodate the respective fuel nozzle 86 and introduce air into the forward end of the combustion chamber 66.
  • Each fuel nozzle 86 may be secured to the diffuser case module 64 and project through one of the hood ports 94 and the respective swirler 90.
  • the forward assembly 80 introduces core combustion air into the forward section of the combustion chamber 66 while the remainder enters the outer annular plenum 76 and the inner annular plenum 78.
  • the multiple of fuel nozzles 86 and adjacent structure generate a blended fuel-air mixture that supports stable combustion in the combustion chamber 66.
  • NGVs 28A are static engine components which direct the combustion gases onto the turbine blades in the turbine section 28 to facilitate the conversion of pressure energy into kinetic energy.
  • the combustion gases are also accelerated by the NGVs 28 A because of their convergent shape and are typically given a "spin” or a "swirl" in the direction of turbine rotation.
  • each fuel injector 86 generally includes a first inlet 100 and a second inlet 102 defined by an inlet housing 104, a support housing 106 and a tip assembly 108.
  • the primary inlet 100 receives approximately twenty (20) pounds per hour (pph) of fuel while the secondary inlet 102 receives between approximately twenty-nine (29) to four hundred (400) pph of fuel depending on flight condition.
  • the fuel injector 86 in the disclosed non-limiting embodiment provides concentric passages for a liquid such as Jet-A, diesel, JP8, water and combinations thereof and alternatively, a gas such as natural gas.
  • the inlet housing 104 is received within the support housing 106 and a tube 110 extends through the housings 102, 104 ( Figure 5).
  • the tube 110 is secured within the inlet housing 104 with a seal such as an O-ring at one end section 112 and at the opposite end section 114 in the tip assembly 108 via a braze, weld, thread or other attachment.
  • the tube 110 defines an annular gas passage 116 within the housings 104, 106 that operates as a heat shield to minimize or prevent coking of the fluid through the tube 110.
  • the tube 110 may incorporate a spacer 118 that dampens the tube 110 to minimize high cycle fatigue. The helical shape of the spacer 118 further facilitates flow of the adjacent secondary fuel circuit.
  • the tip assembly 108 is at least partially received within the swirler 90 along the nozzle axis F to inject/atomize the fuel into the combustor chamber 66 ( Figure 2).
  • a heat shield 120 surrounds a swirler tip 122 and a wear sleeve 124 surrounds the heat shield 120 to thermally and mechanically protect the fuel circuits from the hot combustor environment.
  • This heat shield 120 thermally shields the fuel circuit walls from the high velocity combustor chamber environment to reduce the impact of thermal convection, minimize the thermal conduction path from the hot exterior to the interior fuel circuit and allows fuel to fill the void between heat shield and fuel circuit such that this heated fuel may eventually solidify and form an insulating fuel coke.
  • the fuel nozzle tip assembly 108 may be assembled together by a series of braze and/or weld joints. For example, these thin walled details may be welded to the fuel nozzle 86 at final assembly such that the shields extend from the support to the tip.
  • the primary fuel flow is communicated through a pressure atomizer 126 within the swirler tip 122 that is sized to meet the flow requirements of the particular fuel nozzle 86.
  • the secondary fuel flow is atomized through a multiple of radial orifices 128 (six (6) shown).
  • the multiple of radial orifices 128 within the swirler tip 122 are sized to and/or angled with respect to axis F provide the desired flow requirements. It should be appreciated that any number of orifices, angles and/or arrangements may alternatively or additionally be provided.
  • a helical insert 130 within the swirler tip 122 upstream of the atomizer tip 126 facilitates management and distribution of the fuel flow within the fuel nozzle tip 108.
  • the helical insert 130 includes a central passage 132 that communicates primary fuel into the atomizer tip 126 from the tube 110.
  • the helical insert 130 further includes an outer surface 134 with a multiple of helical channels 136 (three (3) shown).
  • the multiple of helical channels 136 increases fuel velocity and minimize dead zones (low velocity) regions prior to communication through the multiple of radial orifices 128. This facilitates maintenance of desired fuel circuit wetted wall temperatures within acceptable limits.
  • a simplex fuel nozzle 86' includes a solid helical adapter 13 OA as additional steams are not provided as in the duplex nozzle. That is, no central passage is required.
  • the second inlet 102 may include a valve 140 to manage the fuel flow.
  • the valve 140 includes a spring 142 for alignment (perpendicularity) relative to travel of a valve 144.
  • a series of grooves 146 on a piston 148 ensures concentricity of the piston 148.
  • a lock pin 150 is operable to receive a valve set screw 150. It should be understood that other beneficial features such a surface coating and piston shape may additionally or alternatively be provided to further facilitate operation.
  • a valve housing 160 is mounted around the valve 140 to thermally protect the fuel nozzle valve 140. The valve housing 160 thermally minimizes thermal conduction to the fuel circuits fuel may also be permitted to fill the void between housing 160 and valve 140 such that the fuel will eventually solidify and further provide an insulating coke.

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)
  • Spray-Type Burners (AREA)

Abstract

Injecteur de combustible pour chambre de combustion d'un moteur à turbine à gaz comprenant une pointe de coupelle de tourbillonnement le long d'un axe ; et un adaptateur hélicoïdal le long de l'axe et dans la pointe de coupelle de tourbillonnement. Un procédé d'acheminement d'un gaz combustible et d'un liquide dans un injecteur de combustible jusque dans un moteur à turbine à gaz consiste à acheminer un écoulement de combustible primaire dans un insert hélicoïdal ; et à acheminer un écoulement de combustible secondaire autour de l'insert hélicoïdal.
PCT/US2013/065453 2012-10-17 2013-10-17 Injecteur de combustible pour moteur à turbine à gaz WO2014113105A2 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
US201261715020P 2012-10-17 2012-10-17
US61/715,020 2012-10-17

Publications (2)

Publication Number Publication Date
WO2014113105A2 true WO2014113105A2 (fr) 2014-07-24
WO2014113105A3 WO2014113105A3 (fr) 2014-10-16

Family

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Family Applications (1)

Application Number Title Priority Date Filing Date
PCT/US2013/065453 WO2014113105A2 (fr) 2012-10-17 2013-10-17 Injecteur de combustible pour moteur à turbine à gaz

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Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US12072104B1 (en) 2023-09-22 2024-08-27 Pratt & Whitney Canada Corp. Fuel delivery apparatus for a gas turbine engine

Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US6021635A (en) * 1996-12-23 2000-02-08 Parker-Hannifin Corporation Dual orifice liquid fuel and aqueous flow atomizing nozzle having an internal mixing chamber
US6715292B1 (en) * 1999-04-15 2004-04-06 United Technologies Corporation Coke resistant fuel injector for a low emissions combustor
US20070074517A1 (en) * 2005-09-30 2007-04-05 Solar Turbines Incorporated Fuel nozzle having swirler-integrated radial fuel jet
US7926282B2 (en) * 2008-03-04 2011-04-19 Delavan Inc Pure air blast fuel injector
US20120186259A1 (en) * 2011-01-26 2012-07-26 United Technologies Corporation Fuel injector assembly

Patent Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US6021635A (en) * 1996-12-23 2000-02-08 Parker-Hannifin Corporation Dual orifice liquid fuel and aqueous flow atomizing nozzle having an internal mixing chamber
US6715292B1 (en) * 1999-04-15 2004-04-06 United Technologies Corporation Coke resistant fuel injector for a low emissions combustor
US20070074517A1 (en) * 2005-09-30 2007-04-05 Solar Turbines Incorporated Fuel nozzle having swirler-integrated radial fuel jet
US7926282B2 (en) * 2008-03-04 2011-04-19 Delavan Inc Pure air blast fuel injector
US20120186259A1 (en) * 2011-01-26 2012-07-26 United Technologies Corporation Fuel injector assembly

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US12072104B1 (en) 2023-09-22 2024-08-27 Pratt & Whitney Canada Corp. Fuel delivery apparatus for a gas turbine engine

Also Published As

Publication number Publication date
WO2014113105A3 (fr) 2014-10-16

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