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US 9,068,748 B2Granted
Axial stage combustor for gas turbine engines
AssigneeUnited Technologies Corporation
InventorJames B. Hoke
Life of this patent
Estimating the term…
Abstract
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.
Claims15
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;
- and
- 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.
5 more elements
A gas turbine engine, comprising:
- a compressor section;
- a turbine section;
- and
- a 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.
8 more elements
Dependencies are inferred from the claim text.