US20160033130A1 - Fuel nozzle for a gas turbine engine - Google Patents

Fuel nozzle for a gas turbine engine Download PDF

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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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US
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
Application number
US14/775,971
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English (en)
Inventor
Kevin J. Low
Joey Wong
Chris J. Niggemeier
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
RTX Corp
Original Assignee
United Technologies Corp
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 Corp filed Critical United Technologies Corp
Priority to US14/775,971 priority Critical patent/US20160033130A1/en
Assigned to UNITED TECHNOLOGIES CORPORATION reassignment UNITED TECHNOLOGIES CORPORATION ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: LOW, KEVIN J., Niggemeier, Chris J., WONG, JOEY
Publication of US20160033130A1 publication Critical patent/US20160033130A1/en
Assigned to RAYTHEON TECHNOLOGIES CORPORATION reassignment RAYTHEON TECHNOLOGIES CORPORATION CHANGE OF NAME (SEE DOCUMENT FOR DETAILS). Assignors: UNITED TECHNOLOGIES CORPORATION
Assigned to RAYTHEON TECHNOLOGIES CORPORATION reassignment RAYTHEON TECHNOLOGIES CORPORATION CORRECTIVE ASSIGNMENT TO CORRECT THE AND REMOVE PATENT APPLICATION NUMBER 11886281 AND ADD PATENT APPLICATION NUMBER 14846874. TO CORRECT THE RECEIVING PARTY ADDRESS PREVIOUSLY RECORDED AT REEL: 054062 FRAME: 0001. ASSIGNOR(S) HEREBY CONFIRMS THE CHANGE OF ADDRESS. Assignors: UNITED TECHNOLOGIES CORPORATION
Abandoned legal-status Critical Current

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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
    • F23R3/283—Attaching or cooling of fuel injecting means including supports for fuel injectors, stems, or lances
    • 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
    • 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
    • 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
    • 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
    • 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
    • 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
    • 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
    • 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.

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)
  • Nozzles For Spraying Of Liquid Fuel (AREA)
US14/775,971 2013-03-15 2014-03-13 Fuel nozzle for a gas turbine engine Abandoned US20160033130A1 (en)

Priority Applications (1)

Application Number Priority Date Filing Date Title
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
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 (fr) 2013-03-15 2014-03-13 Buse de carburant pour un moteur à turbine à gaz

Related Parent Applications (1)

Application Number Title Priority Date Filing Date
PCT/US2014/026306 A-371-Of-International WO2014197072A2 (fr) 2013-03-15 2014-03-13 Buse de carburant pour un moteur à turbine à gaz

Related Child Applications (1)

Application Number Title Priority Date Filing Date
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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US14/775,971 Abandoned US20160033130A1 (en) 2013-03-15 2014-03-13 Fuel nozzle for a gas turbine engine
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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EP (1) EP2971685B1 (fr)
WO (1) WO2014197072A2 (fr)

Cited By (3)

* Cited by examiner, † Cited by third party
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)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR3116567B1 (fr) * 2020-11-20 2023-06-16 Safran Aircraft Engines Procede de montage d'injecteurs de turbomachine et dispositif de detrompage associe

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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
US20140271144A1 (en) * 2013-03-13 2014-09-18 Rolls-Royce North American Technologies, Inc. Turbine shroud

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US5295773A (en) * 1992-11-10 1994-03-22 United Technologies Corporation Self-locking pressed in insert arrangement
US7014258B2 (en) * 2003-09-18 2006-03-21 Newfrey Llc Error proof anti-chucking wedge assembly
US7540141B2 (en) * 2005-12-13 2009-06-02 Hamilton Sundstrand Corporation Smart fuel control system
US7805948B2 (en) * 2005-12-15 2010-10-05 Pratt & Whitney Canada Corp. Internally mounted device for a pressure vessel
US7703287B2 (en) * 2006-10-31 2010-04-27 Delavan Inc Dynamic sealing assembly to accommodate differential thermal growth of fuel injector components
US20090077973A1 (en) * 2007-09-20 2009-03-26 Hamilton Sundstrand Corporation Gas Turbine Fuel System for High Altitude Starting and Operation
US8607571B2 (en) * 2009-09-18 2013-12-17 Delavan Inc Lean burn injectors having a main fuel circuit and one of multiple pilot fuel circuits with prefiliming air-blast atomizers
FR2935777B1 (fr) * 2008-09-09 2013-03-08 Snecma Chambre de combustion de turbomachine
US8141368B2 (en) * 2008-11-11 2012-03-27 Delavan Inc Thermal management for fuel injectors
FR2939171B1 (fr) * 2008-11-28 2010-12-31 Snecma Turbomachine a systemes d'injection de carburant distincts, utilisant des joints d'etancheite identiques.
GB0918169D0 (en) * 2009-10-19 2009-12-02 Rolls Royce Plc Fuel injector mounting system
US20160033130A1 (en) * 2013-03-15 2016-02-04 United Technologies Corporation Fuel nozzle for a gas turbine engine

Patent Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
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
US20140271144A1 (en) * 2013-03-13 2014-09-18 Rolls-Royce North American Technologies, Inc. Turbine shroud

Cited By (4)

* Cited by examiner, † Cited by third party
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
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 (fr) 2016-01-20
EP2971685A4 (fr) 2016-03-30
US11226102B2 (en) 2022-01-18
EP2971685B1 (fr) 2021-06-23
WO2014197072A2 (fr) 2014-12-11
WO2014197072A3 (fr) 2015-02-26
US20180209652A1 (en) 2018-07-26

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Free format text: CORRECTIVE ASSIGNMENT TO CORRECT THE AND REMOVE PATENT APPLICATION NUMBER 11886281 AND ADD PATENT APPLICATION NUMBER 14846874. TO CORRECT THE RECEIVING PARTY ADDRESS PREVIOUSLY RECORDED AT REEL: 054062 FRAME: 0001. ASSIGNOR(S) HEREBY CONFIRMS THE CHANGE OF ADDRESS;ASSIGNOR:UNITED TECHNOLOGIES CORPORATION;REEL/FRAME:055659/0001

Effective date: 20200403