Axial stage combustor for gas turbine engines
원문보기
IPC분류정보
국가/구분
United States(US) Patent
등록
국제특허분류(IPC7판)
F23R-003/34
F23R-003/06
F23R-003/50
출원번호
US-0012212
(2011-01-24)
등록번호
US-9068748
(2015-06-30)
발명자
/ 주소
Hoke, James B.
출원인 / 주소
United Technologies Corporation
대리인 / 주소
Miller, Matthias & Hull LLP
인용정보
피인용 횟수 :
5인용 특허 :
39
초록▼
A combustor for a gas turbine engine includes a radially inboard liner, a radially outboard liner, and a bulkhead that cooperatively define an annular combustion chamber, a plurality of first fuel injectors that are disposed in the bulkhead, and a plurality of second fuel injectors that are disposed
A combustor for a gas turbine engine includes a radially inboard liner, a radially outboard liner, and a bulkhead that cooperatively define an annular combustion chamber, a plurality of first fuel injectors that are disposed in the bulkhead, and a plurality of second fuel injectors that are disposed in at least one of the inboard liner and the outboard liner aftward of the bulkhead. A method is also provided for operating the combustor of the gas turbine engine wherein fuel distribution between the forward combustion zone and the downstream combustion zone is selectively varied in response to the power operating mode of the gas turbine engine with an objective to control NOx formation.
대표청구항▼
1. A combustor for a gas turbine engine comprising: a radially inboard liner extending circumferentially and fore to aft longitudinally;a radially outboard liner extending circumferentially and fore to aft longitudinally, the outboard liner disposed radially outward from and circumscribing the inboa
1. A combustor for a gas turbine engine comprising: a radially inboard liner extending circumferentially and fore to aft longitudinally;a radially outboard liner extending circumferentially and fore to aft longitudinally, the outboard liner disposed radially outward from and circumscribing the inboard liner;an annular bulkhead having circumferential expanse and radial expanse and extending between a forward end of the inboard liner and a forward end of the outboard liner, the inboard liner, the outboard liner and the bulkhead cooperatively defining an annular combustion chamber;a plurality of first fuel injectors disposed in the bulkhead;a forward combustion zone in a forward region of the annular combustion chamber, the forward combustion zone associated with the plurality of first fuel injectors, the forward combustion zone establishing a fuel-rich environment during a low power operation of the gas turbine engine;a plurality of second fuel injectors disposed in each of the inboard liner and the outboard liner aftward of the bulkhead, the plurality of second fuel injectors are arranged in first and second circumferential rings, fuel injectors in the second circumferential ring being disposed in staggered relationship relative to fuel injectors in the first circumferential ring, the plurality of second fuel injectors associated with a downstream combustion zone in a downstream region of the annular combustion chamber, both the downstream combustion zone and the forward combustion zone establishing respective fuel-lean environments during a high power operation of the gas turbine engine; anda plurality of aft combustion air admission holes penetrating at least one of the inboard liner and the outboard liner, the aft combustion air admission holes disposed in a circumferential ring located aft of the plurality of second fuel injectors. 2. The combustor as recited in claim 1 wherein the plurality of first fuel injectors are arranged in the bulkhead to inject fuel generally longitudinally into the forward combustion zone in the forward region of the combustion chamber and the plurality of second fuel injectors are arranged in each of said inboard liner and outboard liner to inject fuel generally radially inward into the downstream combustion zone in the downstream region of the combustion chamber. 3. The combustor as recited in claim 2 wherein the plurality of second fuel injectors are arranged to inject fuel generally radially with a tangential component. 4. The combustor as recited in claim 1 further comprising: a plurality of intermediate combustion air admission holes penetrating at least one of the inboard liner and the outboard liner, the intermediate combustion air admission holes disposed in a circumferential ring located forward of the plurality of second fuel injectors. 5. The combustor as recited in claim 4 further comprising a plurality of air admission swirlers disposed in the bulkhead in operative association with the plurality of first fuel injectors, each air admission swirler disposed about a respective one of the plurality of first fuel injectors. 6. The combustor as recited in claim 5 further comprising a plurality of main combustion air admission devices disposed in operative association with the plurality of second fuel injectors for admitting a second flow of combustion air in association with a flow of fuel admitted through the plurality of second fuel injectors. 7. The combustor as recited in claim 1 wherein the number of the plurality of second fuel injectors is between two to six times the number of the plurality of first fuel injectors. 8. A gas turbine engine, comprising: a compressor section;a turbine section; anda combustor having a radially inboard liner extending circumferentially and fore to aft longitudinally,a radially outboard liner extending circumferentially and fore to aft longitudinally, the outboard liner disposed radially outward from and circumscribing the inboard liner;an annular bulkhead having circumferential expanse and radial expanse and extending between a forward end of the inboard liner and a forward end of the outboard liner, the inboard liner, the outboard liner and the bulkhead cooperatively defining an annular combustion chamber,a plurality of first fuel injectors disposed in the bulkhead,a forward combustion zone in a forward region of the annular combustion chamber, the forward combustion zone associated with the plurality of first fuel injectors, the forward combustion zone establishing a fuel-rich environment during a low power operation of the gas turbine engine,a plurality of second fuel injectors disposed in each of the inboard liner and the outboard liner aftward of the bulkhead, the plurality of second fuel injectors are arranged in first and second circumferential rings, fuel injectors in the second circumferential ring being disposed in staggered relationship relative to fuel injectors in the first circumferential ring, the plurality of second fuel injectors associated with a downstream combustion zone in a downstream region of the annular combustion chamber, both the downstream combustion zone and the forward combustion zone establishing respective fuel-lean environments during a high power operation of the gas turbine engine,a plurality of aft combustion air admission holes penetrating at least one of the inboard liner and the outboard liner, the aft combustion air admission holes disposed in a circumferential ring located aft of the plurality of second fuel injectors. 9. The gas turbine engine as recited in claim 8 wherein the plurality of first fuel injectors are arranged in the bulkhead to inject fuel generally longitudinally into the forward combustion zone in the forward region of the combustion chamber and the plurality of second fuel injectors are arranged in each of said inboard liner and outboard liner to inject fuel generally radially inward into the downstream combustion zone in the downstream region of the combustion chamber. 10. The gas turbine engine as recited in claim 9 wherein the plurality of second fuel injectors are arranged to inject fuel generally radially with a tangential component. 11. The gas turbine engine as recited in claim 8 wherein the combustor further has a plurality of intermediate combustion air admission holes penetrating at least one of the inboard liner and the outboard liner, the intermediate combustion air admission holes disposed in a circumferential ring located forward of the plurality of second fuel injectors. 12. The gas turbine engine as recited in claim 11 wherein the combustor further has a plurality of air admission swirlers disposed in the bulkhead in operative association with the plurality of first fuel injectors, each air admission swirler disposed about a respective one of the plurality of first fuel injectors. 13. The gas turbine engine as recited in claim 12 wherein the combustor further has a plurality of main combustion air admission devices disposed in operative association with the plurality of second fuel injectors for admitting a second flow of combustion air in association with a flow of fuel admitted through the plurality of second fuel injectors. 14. The gas turbine engine as recited in claim 8 wherein the second fuel injectors are arranged in a circumferential ring in circumferentially spaced relationship. 15. The gas turbine engine as recited in claim 8 wherein the number of the plurality of second fuel injectors is between two to six times the number of the plurality of first fuel injectors.
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이 특허에 인용된 특허 (39)
Correa Sanjay M. (Schenectady NY) Warren ; Jr. Richard E. (Schenectady NY), Apparatus for supersonic combustion in a restricted length.
Jorgensen Robert A. (Clifton Park NY) Farrell Roger A. (Schenectady NY) Gerhold Bruce W. (Rexford NY), Dual stage-dual mode low emission gas turbine combustion system.
Burd,Steven W.; Cheung,Albert K.; Ols,John T.; Smith,Reid D.; Segalman,Irving, Gas turbine annular combustor having a first converging volume and a second converging volume, converging less gradually than the first converging volume.
Willis Jeffrey D,GBX, Gas turbine engine combustion chamber having premixed homogeneous combustion followed by catalytic combustion and a method of operation thereof.
Waslo Jennifer (Scotia NY) Kuwata Masayoshi (Ballston Lake NY) Washam Roy M. (Schenectady NY), Impingement cooled liner for dry low NOx venturi combustor.
Ansart Denis R. H. (Bois le Roi FRX) Commaret Patrice A. (Maincy FRX) David Etienne S. R. (Bois le Roi FRX) Desaulty Michel A. A. (Vert Saint Denis FRX) Quinquenneau Bruno M. M. (Massy FRX) Sandelis , Pre-mixing injection system for a turbojet engine.
Moreno Frederick E. (Los Altos CA) Joshi Narendra D. (Phoenix AZ), Staged lean premix low nox hot wall gas turbine combustor with improved turndown capability.
Ansart Denis Roger Henri,FRX ; James Bruno,FRX ; Desaulty Michel Andre Albert,FRX ; Staessen Richard Emile,FRX, Turbomachine combustion chamber with inner and outer injector rows.
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