US20160033130A1 - Fuel nozzle for a gas turbine engine - Google Patents
Fuel nozzle for a gas turbine engine Download PDFInfo
- Publication number
- US20160033130A1 US20160033130A1 US14/775,971 US201414775971A US2016033130A1 US 20160033130 A1 US20160033130 A1 US 20160033130A1 US 201414775971 A US201414775971 A US 201414775971A US 2016033130 A1 US2016033130 A1 US 2016033130A1
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- United States
- Prior art keywords
- locator
- flange
- recited
- fuel nozzle
- fuel
- 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.)
- Abandoned
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Classifications
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/28—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
- F23R3/283—Attaching or cooling of fuel injecting means including supports for fuel injectors, stems, or lances
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C3/00—Gas-turbine plants characterised by the use of combustion products as the working fluid
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
- F02C7/22—Fuel supply systems
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- 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
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- 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
- F23D11/36—Details, e.g. burner cooling means, noise reduction means
- F23D11/38—Nozzles; Cleaning devices therefor
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2220/00—Application
- F05D2220/30—Application in turbines
- F05D2220/32—Application in turbines in gas turbines
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2230/00—Manufacture
- F05D2230/60—Assembly methods
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2240/00—Components
- F05D2240/35—Combustors or associated equipment
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R2900/00—Special features of, or arrangements for continuous combustion chambers; Combustion processes therefor
- F23R2900/00017—Assembling combustion chamber liners or subparts
-
- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T50/00—Aeronautics or air transport
- Y02T50/60—Efficient propulsion technologies, e.g. for aircraft
Definitions
- the pin is fixedly mounted within an aperture in the flange.
- the flange is symmetrical about an X-axis and a Y-axis.
- a combustor section for a gas turbine engine includes a diffuser case with a fuel injector mount pad having a second locator; and a fuel injector with a flange having a first locator, wherein the first locator is indexed with the second locator.
- the first locator is a pin.
- the first locator is a hole.
- FIG. 2 is an expanded longitudinal schematic sectional view of a combustor section according to one non-limiting embodiment
- FIG. 8 is a bottom view of a fuel injector flange according to another disclosed non-limiting embodiment
- FIG. 9 is a bottom view of a fuel injector flange according to another disclosed non-limiting embodiment.
- FIG. 10 is a bottom view of a fuel injector flange according to another 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 incorporates a fan section 22 , a compressor section 24 , a combustor section 26 and a turbine section 28 .
- Alternative engine architectures 200 might include an augmentor section and exhaust duct section (not shown) among other systems or features.
- the fan section 22 drives air along a bypass flowpath while the compressor section 24 drives air along a core flowpath for compression and communication into the combustor section 26 then expansion through the turbine section 28 .
- 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 wherein an intermediate spool includes an intermediate pressure compressor (“IPC”) between a Low Pressure Compressor (“LPC”) and a High Pressure Compressor (“HPC”), and an intermediate pressure turbine (“IPT”) between the high pressure turbine (“HPT”) and the Low pressure Turbine (“LPT”).
- IPC intermediate pressure compressor
- LPC Low Pressure Compressor
- HPC High Pressure Compressor
- IPT intermediate pressure turbine
- the engine 20 generally includes a low spool 30 and a high spool 32 mounted for rotation about an engine central longitudinal axis A relative to an engine static structure 36 via several bearing structures 38 .
- the low spool 30 generally includes an inner shaft 40 that interconnects a fan 42 , a low pressure compressor 44 (“LPC”) and a low pressure turbine 46 (“LPT”).
- the inner shaft 40 drives the fan 42 directly or through a geared architecture 48 to drive the fan 42 at a lower speed than the low spool 30 .
- An exemplary reduction transmission is an epicyclic transmission, namely a planetary or star gear system.
- the high spool 32 includes an outer shaft 50 that interconnects a high pressure compressor 52 (“HPC”) and high pressure turbine 54 (“HPT”).
- a combustor 56 is arranged between the high pressure compressor 52 and the high pressure turbine 54 .
- the inner shaft 40 and the outer shaft 50 are concentric and rotate about the engine central longitudinal axis A which is collinear with their longitudinal axes.
- Core airflow is compressed by the LPC 44 then the HPC 52 , mixed with the fuel and burned in the combustor 56 , then expanded over the HPT 54 and the LPT 46 .
- the turbines 54 , 46 rotationally drive the respective low spool 30 and high spool 32 in response to the expansion.
- the main engine shafts 40 , 50 are supported at a plurality of points by bearing structures 38 within the static structure 36 . It should be understood that various bearing structures 38 at various locations may alternatively or additionally be provided.
- the gas turbine engine 20 is a high-bypass geared aircraft engine.
- the gas turbine engine 20 bypass ratio is greater than about six (6:1).
- the geared architecture 48 can include an epicyclic gear train, such as a planetary gear system or other gear system.
- the example epicyclic gear train has a gear reduction ratio of greater than about 2.3, and in another example is greater than about 2.5:1.
- the geared turbofan enables operation of the low spool 30 at higher speeds which can increase the operational efficiency of the low pressure compressor 44 and low pressure turbine 46 and render increased pressure in a fewer number of stages.
- a pressure ratio associated with the low pressure turbine 46 is pressure measured prior to the inlet of the low pressure turbine 46 as related to the pressure at the outlet of the low pressure turbine 46 prior to an exhaust nozzle of the gas turbine engine 20 .
- the bypass ratio of the gas turbine engine 20 is greater than about ten (10:1)
- the fan diameter is significantly larger than that of the low pressure compressor 44
- the low pressure turbine 46 has a pressure ratio that is greater than about five (5:1). It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present disclosure is applicable to other gas turbine engines including direct drive turbofans.
- a significant amount of thrust is provided by the bypass flow path due to the high bypass ratio.
- the fan section 22 of the gas turbine engine 20 is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet. This flight condition, with the gas turbine engine 20 at its best fuel consumption, is also known as bucket cruise Thrust Specific Fuel Consumption (TSFC).
- TSFC Thrust Specific Fuel Consumption
- the outer combustor liner assembly 60 is spaced radially inward from an outer diffuser case 64 -O 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 64 -I 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.
- Each combustor liner assembly 60 , 62 contain the combustion products for direction toward the turbine section 28 .
- Each combustor liner 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 72 A and a multiple of aft liner panels 72 B that are circumferentially staggered to line the hot side of the outer shell 68 .
- a multiple of forward liner panels 74 A and a multiple of aft liner panels 74 B are circumferentially staggered to line the hot side of the inner shell 70 .
- the combustor 56 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 , a bulkhead assembly 84 , a multiple of fuel nozzles 86 (one shown) and a multiple of swirlers 90 (one shown).
- Each of the swirlers 90 is circumferentially aligned with one of the annular hood ports 94 to project through the bulkhead assembly 84 .
- Each bulkhead assembly 84 generally includes a bulkhead support shell 96 secured to the combustor liner assembly 60 , 62 , and a multiple of circumferentially distributed bulkhead liner panels 98 secured to the bulkhead support shell 96 around the central opening 92 .
- 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 .
- the mount flange 106 may be symmetrical along an X-axis and a Y-axis.
- a first aperture 110 and a second aperture 112 are located along the X-axis that flank the support 104 .
- An extended area 114 , 116 is located along a respective first side 118 and second side 120 .
- One of the extended areas 114 , 116 includes a first locator 122 in the Z-direction.
- the first locator 122 in one disclosed non-limiting embodiment may be a pin 124 that is fitted into on of a multiple of apertures 126 ( FIG. 6 ).
- the first locator 122 in another disclosed non-limiting embodiment is an integral raised area.
- the first locator 122 in another disclosed non-limiting embodiment is an aperture or a slot.
- the first locator 122 provides a mistake-proofing feature on the underside of the mount flange 106 that interfaces with a second locator 128 such as a hole, slot or pin in a fuel nozzle pad 130 on the diffuser case 64 ( FIG. 7 ) that indexes with the first locator 122 .
- Each fuel nozzle 86 is thereby associated with a particular fuel nozzle pad 130 on the diffuser case 64 ( FIG. 6 ) that has an associated positioned second locator 128 .
- the second locator 128 being a slot or hole—may be readily machined into a common fuel nozzle pad 130 . That is, all of the multiple of fuel nozzle pads 130 on the diffuser case 64 may be identical with only a particular second locator 128 being different to permit installation of only the proper fuel nozzle 86 in a mistake-proof manner. Mistake-proofing beneficially protects against backwards installation, simplex/duplex mix-ups, and cross-engine mix-ups through the physical prevention of mis-installation.
Abstract
Description
- This application claims priority to U.S. patent application. No. 61/787,469 filed Mar. 15, 2013.
- The present disclosure relates to a gas turbine engine and, more particularly, to a combustor section therefor.
- Gas turbine engines, such as those which power modern commercial and military aircraft, include a compressor section to pressurize a supply of air, a combustor section to burn a hydrocarbon fuel in the presence of the pressurized air, and a turbine section to extract energy from the resultant combustion gases.
- The combustor section generally includes a multiple of circumferentially distributed fuel nozzles mounted to an engine case to axially project into a forward section of a combustion chamber to supply the fuel to mix with the pressurized air. Different types of fuel nozzle are often located at particular circumferential locations. The fuel nozzles typically have multiple fasteners that utilize an offset fastener pattern to provide mistake proofing. On a lighter weight two-fastener design, however, an offset pattern may be structurally undesirable.
- A fuel nozzle for a combustor of a gas turbine engine according to one disclosed non-limiting embodiment of the present disclosure includes a flange with a first locator.
- In a further embodiment of the present disclosure, the first locator extends from the flange in a Z-direction.
- In a further embodiment of any of the foregoing embodiments of the present disclosure, the first locator is an integral extension.
- In a further embodiment of any of the foregoing embodiments of the present disclosure, the first locator is a pin.
- In a further embodiment of any of the foregoing embodiments of the present disclosure, the pin is fixedly mounted within an aperture in the flange.
- In a further embodiment of any of the foregoing embodiments of the present disclosure, the flange is symmetrical about an X-axis and a Y-axis.
- In a further embodiment of any of the foregoing embodiments of the present disclosure, the first locator is positioned within an extension in the flange.
- In a further embodiment of any of the foregoing embodiments of the present disclosure, the flange includes a first and second aperture.
- In a further embodiment of any of the foregoing embodiments of the present disclosure, the flange includes a first and second aperture along a common axis.
- A combustor section for a gas turbine engine according to another disclosed non-limiting embodiment of the present disclosure includes a diffuser case with a fuel injector mount pad having a second locator; and a fuel injector with a flange having a first locator, wherein the first locator is indexed with the second locator.
- In a further embodiment of any of the foregoing embodiments of the present disclosure, the first locator extends from the flange.
- In a further embodiment of any of the foregoing embodiments of the present disclosure, the first locator is a pin.
- In a further embodiment of any of the foregoing embodiments of the present disclosure, the second locator is a slot.
- In a further embodiment of any of the foregoing embodiments of the present disclosure, the first locator is a hole.
- In a further embodiment of any of the foregoing embodiments of the present disclosure, the second locator is a pin.
- In a further embodiment of any of the foregoing embodiments of the present disclosure, the flange includes a first aperture and a second aperture.
- A method of mounting a fuel injector into a combustor of a gas turbine engine according to another disclosed non-limiting embodiment of the present disclosure includes bolting a fuel injector flange to a fuel injector pad; and indexing a first locator that extends from the flange with respect to a second locator defined by the fuel injector pad.
- A further embodiment of any of the foregoing embodiments of the present disclosure includes bolting the fuel injector flange to the fuel injector pad with only two (2) bolts.
- The foregoing features and elements may be combined in various combinations without exclusivity, unless expressly indicated otherwise. These features and elements as well as the operation thereof will become more apparent in light of the following description and the accompanying drawings. It should be understood, however, the following description and drawings are intended to be exemplary in nature and non-limiting.
- Various features will become apparent to those skilled in the art from the following detailed description of the disclosed non-limiting embodiment. The drawings that accompany the detailed description can be briefly described as follows:
-
FIG. 1 is a schematic cross-section of an example gas turbine engine architecture; -
FIG. 2 is an expanded longitudinal schematic sectional view of a combustor section according to one non-limiting embodiment; -
FIG. 3 is an isometric view of a duplex fuel injector; -
FIG. 4 is an isometric view of a simplex fuel injector; -
FIG. 5 is an expanded bottom view of a fuel injector flange; -
FIG. 6 is a bottom view of a fuel injector flange according to one disclosed non-limiting embodiment; -
FIG. 7 is an expanded isometric view of a fuel injector pad; -
FIG. 8 is a bottom view of a fuel injector flange according to another disclosed non-limiting embodiment; -
FIG. 9 is a bottom view of a fuel injector flange according to another disclosed non-limiting embodiment; and -
FIG. 10 is a bottom view of a fuel injector flange according to another disclosed non-limiting embodiment. -
FIG. 1 schematically illustrates agas turbine engine 20. Thegas turbine engine 20 is disclosed herein as a two-spool turbo fan that generally incorporates afan section 22, acompressor section 24, acombustor section 26 and aturbine section 28. Alternative engine architectures 200 might include an augmentor section and exhaust duct section (not shown) among other systems or features. Thefan section 22 drives air along a bypass flowpath while thecompressor section 24 drives air along a core flowpath for compression and communication into thecombustor section 26 then expansion through theturbine section 28. 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 wherein an intermediate spool includes an intermediate pressure compressor (“IPC”) between a Low Pressure Compressor (“LPC”) and a High Pressure Compressor (“HPC”), and an intermediate pressure turbine (“IPT”) between the high pressure turbine (“HPT”) and the Low pressure Turbine (“LPT”). - The
engine 20 generally includes alow spool 30 and ahigh spool 32 mounted for rotation about an engine central longitudinal axis A relative to an enginestatic structure 36 viaseveral bearing structures 38. Thelow spool 30 generally includes aninner shaft 40 that interconnects afan 42, a low pressure compressor 44 (“LPC”) and a low pressure turbine 46 (“LPT”). Theinner shaft 40 drives thefan 42 directly or through a gearedarchitecture 48 to drive thefan 42 at a lower speed than thelow spool 30. An exemplary reduction transmission is an epicyclic transmission, namely a planetary or star gear system. - The
high spool 32 includes anouter shaft 50 that interconnects a high pressure compressor 52 (“HPC”) and high pressure turbine 54 (“HPT”). Acombustor 56 is arranged between thehigh pressure compressor 52 and thehigh pressure turbine 54. Theinner shaft 40 and theouter shaft 50 are concentric and rotate about the engine central longitudinal axis A which is collinear with their longitudinal axes. - Core airflow is compressed by the
LPC 44 then the HPC 52, mixed with the fuel and burned in thecombustor 56, then expanded over the HPT 54 and theLPT 46. Theturbines low spool 30 andhigh spool 32 in response to the expansion. Themain engine shafts structures 38 within thestatic structure 36. It should be understood thatvarious bearing structures 38 at various locations may alternatively or additionally be provided. - In one non-limiting example, the
gas turbine engine 20 is a high-bypass geared aircraft engine. In a further example, thegas turbine engine 20 bypass ratio is greater than about six (6:1). The gearedarchitecture 48 can include an epicyclic gear train, such as a planetary gear system or other gear system. The example epicyclic gear train has a gear reduction ratio of greater than about 2.3, and in another example is greater than about 2.5:1. The geared turbofan enables operation of thelow spool 30 at higher speeds which can increase the operational efficiency of thelow pressure compressor 44 andlow pressure turbine 46 and render increased pressure in a fewer number of stages. - A pressure ratio associated with the
low pressure turbine 46 is pressure measured prior to the inlet of thelow pressure turbine 46 as related to the pressure at the outlet of thelow pressure turbine 46 prior to an exhaust nozzle of thegas turbine engine 20. In one non-limiting embodiment, the bypass ratio of thegas turbine engine 20 is greater than about ten (10:1), the fan diameter is significantly larger than that of thelow pressure compressor 44, and thelow pressure turbine 46 has a pressure ratio that is greater than about five (5:1). It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present disclosure is applicable to other gas turbine engines including direct drive turbofans. - In one embodiment, a significant amount of thrust is provided by the bypass flow path due to the high bypass ratio. The
fan section 22 of thegas turbine engine 20 is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet. This flight condition, with thegas turbine engine 20 at its best fuel consumption, is also known as bucket cruise Thrust Specific Fuel Consumption (TSFC). TSFC is an industry standard parameter of fuel consumption per unit of thrust. - Fan Pressure Ratio is the pressure ratio across a blade of the
fan section 22 without the use of a Fan Exit Guide Vane system. The low Fan Pressure Ratio according to one non-limiting embodiment of the examplegas turbine engine 20 is less than 1.45. Low Corrected Fan Tip Speed is the actual fan tip speed divided by an industry standard temperature correction of (“T”/518.7)0.5. The Low Corrected Fan Tip Speed according to one non-limiting embodiment of the examplegas turbine engine 20 is less than about 1150 fps (351 m/s). - With reference to
FIG. 2 , thecombustor 56 generally includes an outercombustor liner assembly 60, an innercombustor liner assembly 62 and adiffuser case 64. The outercombustor liner assembly 60 and the innercombustor liner assembly 62 are spaced apart such that acombustion chamber 66 is defined therebetween. Thecombustion chamber 66 may be generally annular in shape. - The outer
combustor liner assembly 60 is spaced radially inward from an outer diffuser case 64-O of thediffuser case module 64 to define an outerannular plenum 76. The innercombustor liner assembly 62 is spaced radially outward from an inner diffuser case 64-I of thediffuser case module 64 to define an innerannular 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 liner assemblies turbine section 28. Eachcombustor liner assembly respective support shell respective support shell forward liner panels 72A and a multiple ofaft liner panels 72B that are circumferentially staggered to line the hot side of theouter shell 68. A multiple offorward liner panels 74A and a multiple ofaft liner panels 74B are circumferentially staggered to line the hot side of theinner shell 70. - The
combustor 56 further includes aforward assembly 80 immediately downstream of thecompressor section 24 to receive compressed airflow therefrom. Theforward assembly 80 generally includes anannular hood 82, abulkhead assembly 84, a multiple of fuel nozzles 86 (one shown) and a multiple of swirlers 90 (one shown). Each of theswirlers 90 is circumferentially aligned with one of theannular hood ports 94 to project through thebulkhead assembly 84. Eachbulkhead assembly 84 generally includes abulkhead support shell 96 secured to thecombustor liner assembly bulkhead liner panels 98 secured to thebulkhead support shell 96 around thecentral opening 92. - The
annular hood 82 extends radially between, and is secured to, the forward most ends of thecombustor liner assemblies annular hood 82 includes a multiple of circumferentially distributedhood ports 94 that accommodate therespective fuel nozzle 86 and introduce air into the forward end of thecombustion chamber 66 through acentral opening 92. Eachfuel nozzle 86 may be secured to thediffuser case module 64 and project through one of thehood ports 94 and through thecentral opening 92 within therespective swirler 90. - The
forward assembly 80 introduces core combustion air into the forward section of thecombustion chamber 66 while the remainder enters the outerannular plenum 76 and the innerannular plenum 78. The multiple offuel nozzles 86 and adjacent structure generate a blended fuel-air mixture that supports stable combustion in thecombustion chamber 66. - Opposite the
forward assembly 80, the outer andinner support shells HPT 54. TheNGVs 54A are static engine components which direct the combustion gases onto the turbine blades of the first turbine rotor in theturbine section 28 to facilitate the conversion of pressure energy into kinetic energy. The combustion gases are also accelerated by theNGVs 54A because of their convergent shape and are typically given a “spin” or a “swirl” in the direction of turbine rotor rotation. The turbine rotor blades absorb this energy to drive the turbine rotor at high speed. - With reference to
FIG. 3 , each of the multiple offuel nozzles 86—illustrated inFIG. 3 as a duplex fuel nozzle—may include aprimary inlet 100, asecondary inlet 102, asupport 104, amount flange 106 and atip 108. Theprimary inlet 100 receives approximately twenty (20) pounds per hour (pph) of fuel while thesecondary inlet 102 receives between approximately twenty-nine (29) to four hundred (400) pph of fuel depending on flight condition. Asimplex fuel nozzle 86′ only includes the secondary inlet 102 (FIG. 4 ) as additional steams are not provided as in the duplex nozzle. - With reference to
FIG. 5 , themount flange 106 may be symmetrical along an X-axis and a Y-axis. Afirst aperture 110 and asecond aperture 112 are located along the X-axis that flank thesupport 104. Anextended area first side 118 andsecond side 120. Although themount flange 106 defines a particular symmetrical geometry in the illustrated non-limiting embodiment, it should be appreciated that other geometries will benefit herefrom. - One of the
extended areas first locator 122 in the Z-direction. Thefirst locator 122 in one disclosed non-limiting embodiment may be apin 124 that is fitted into on of a multiple of apertures 126 (FIG. 6 ). Thefirst locator 122 in another disclosed non-limiting embodiment is an integral raised area. Thefirst locator 122 in another disclosed non-limiting embodiment is an aperture or a slot. Thefirst locator 122 provides a mistake-proofing feature on the underside of themount flange 106 that interfaces with asecond locator 128 such as a hole, slot or pin in afuel nozzle pad 130 on the diffuser case 64 (FIG. 7 ) that indexes with thefirst locator 122. - The combination of the
second locator 128 in thefuel nozzle pad 130 and the associatedfirst locator 122 match only when thecorrect fuel nozzle 86. For example, anotherfuel nozzle 86″ may include afirst locator 122 at a first predefined angle along the second extended area 116 (FIG. 8 ); anotherfuel nozzle 86″′ may include afirst locator 122 at a second predefined angle along the first extended area 114 (FIG. 9 ); and yet anotherfuel nozzle 86″″ may include afirst locator 122 at a second predefined angle along the second extended area 114 (FIG. 10 ). - Each
fuel nozzle 86 is thereby associated with a particularfuel nozzle pad 130 on the diffuser case 64 (FIG. 6 ) that has an associated positionedsecond locator 128. Thesecond locator 128—being a slot or hole—may be readily machined into a commonfuel nozzle pad 130. That is, all of the multiple offuel nozzle pads 130 on thediffuser case 64 may be identical with only a particularsecond locator 128 being different to permit installation of only theproper fuel nozzle 86 in a mistake-proof manner. Mistake-proofing beneficially protects against backwards installation, simplex/duplex mix-ups, and cross-engine mix-ups through the physical prevention of mis-installation. - It should be understood that relative positional terms such as “forward,” “aft,” “upper,” “lower,” “above,” “below,” and the like are with reference to the normal operational attitude of the vehicle and should not be considered otherwise limiting.
- It should be understood that like reference numerals identify corresponding or similar elements throughout the several drawings. It should also be understood that although a particular component arrangement is disclosed in the illustrated embodiment, other arrangements will benefit herefrom.
- Although particular step sequences are shown, described, and claimed, it should be understood that steps may be performed in any order, separated or combined unless otherwise indicated and will still benefit from the present disclosure.
- The foregoing description is exemplary rather than defined by the limitations within. Various non-limiting embodiments are disclosed herein, however, one of ordinary skill in the art would recognize that various modifications and variations in light of the above teachings will fall within the scope of the appended claims. It is therefore to be understood that within the scope of the appended claims, the disclosure may be practiced other than as specifically described. For that reason the appended claims should be studied to determine true scope and content.
Claims (18)
Priority Applications (1)
Application Number | Priority Date | Filing Date | Title |
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US14/775,971 US20160033130A1 (en) | 2013-03-15 | 2014-03-13 | Fuel nozzle for a gas turbine engine |
Applications Claiming Priority (3)
Application Number | Priority Date | Filing Date | Title |
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US201361787469P | 2013-03-15 | 2013-03-15 | |
US14/775,971 US20160033130A1 (en) | 2013-03-15 | 2014-03-13 | Fuel nozzle for a gas turbine engine |
PCT/US2014/026306 WO2014197072A2 (en) | 2013-03-15 | 2014-03-13 | Fuel nozzle for a gas turbine engine |
Related Parent Applications (1)
Application Number | Title | Priority Date | Filing Date |
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PCT/US2014/026306 A-371-Of-International WO2014197072A2 (en) | 2013-03-15 | 2014-03-13 | Fuel nozzle for a gas turbine engine |
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US15/920,993 Division US11226102B2 (en) | 2013-03-15 | 2018-03-14 | Fuel nozzle for a gas turbine engine |
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US20160033130A1 true US20160033130A1 (en) | 2016-02-04 |
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US15/920,993 Active 2035-08-14 US11226102B2 (en) | 2013-03-15 | 2018-03-14 | Fuel nozzle for a gas turbine engine |
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US15/920,993 Active 2035-08-14 US11226102B2 (en) | 2013-03-15 | 2018-03-14 | Fuel nozzle for a gas turbine engine |
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US (2) | US20160033130A1 (en) |
EP (1) | EP2971685B1 (en) |
WO (1) | WO2014197072A2 (en) |
Cited By (3)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US20180209652A1 (en) * | 2013-03-15 | 2018-07-26 | United Technologies Corporation | Fuel nozzle for a gas turbine engine |
US10408456B2 (en) * | 2015-10-29 | 2019-09-10 | Rolls-Royce Plc | Combustion chamber assembly |
US11378275B2 (en) | 2019-12-06 | 2022-07-05 | Raytheon Technologies Corporation | High shear swirler with recessed fuel filmer for a gas turbine engine |
Families Citing this family (1)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
FR3116567B1 (en) * | 2020-11-20 | 2023-06-16 | Safran Aircraft Engines | METHOD FOR ASSEMBLING TURBOMACHINE INJECTORS AND ASSOCIATED KEYING DEVICE |
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2014
- 2014-03-13 WO PCT/US2014/026306 patent/WO2014197072A2/en active Application Filing
- 2014-03-13 US US14/775,971 patent/US20160033130A1/en not_active Abandoned
- 2014-03-13 EP EP14807624.3A patent/EP2971685B1/en active Active
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2018
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US20040040310A1 (en) * | 2002-09-03 | 2004-03-04 | Prociw Lev Alexander | Stress relief feature for aerated gas turbine fuel injector |
US20050223709A1 (en) * | 2004-04-09 | 2005-10-13 | Delavan Inc. | Alignment and positioning system for installing a fuel injector in a gas turbine engine |
US20080105237A1 (en) * | 2006-11-03 | 2008-05-08 | Pratt & Whitney Canada Corp. | Fuel nozzle flange with reduced heat transfer |
US20110005231A1 (en) * | 2009-07-13 | 2011-01-13 | United Technologies Corporation | Fuel nozzle guide plate mistake proofing |
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US20180209652A1 (en) * | 2013-03-15 | 2018-07-26 | United Technologies Corporation | Fuel nozzle for a gas turbine engine |
US11226102B2 (en) * | 2013-03-15 | 2022-01-18 | Raytheon Technologies Corporation | Fuel nozzle for a gas turbine engine |
US10408456B2 (en) * | 2015-10-29 | 2019-09-10 | Rolls-Royce Plc | Combustion chamber assembly |
US11378275B2 (en) | 2019-12-06 | 2022-07-05 | Raytheon Technologies Corporation | High shear swirler with recessed fuel filmer for a gas turbine engine |
Also Published As
Publication number | Publication date |
---|---|
EP2971685A2 (en) | 2016-01-20 |
EP2971685A4 (en) | 2016-03-30 |
WO2014197072A2 (en) | 2014-12-11 |
EP2971685B1 (en) | 2021-06-23 |
WO2014197072A3 (en) | 2015-02-26 |
US20180209652A1 (en) | 2018-07-26 |
US11226102B2 (en) | 2022-01-18 |
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