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A combustor for a gas turbine engine having an annular combustion chamber includes a plurality of main fuel injection and air swirler assemblies and a plurality of pilot fuel injection and air swirler assemblies disposed in a circumferential ring extending about the circumferential expanse of a forw

A combustor for a gas turbine engine having an annular combustion chamber includes a plurality of main fuel injection and air swirler assemblies and a plurality of pilot fuel injection and air swirler assemblies disposed in a circumferential ring extending about the circumferential expanse of a forward bulkhead. The plurality of pilot fuel injection and air swirler assemblies are interspersed amongst and disposed in the circumferential ring of main fuel injection and air swirler assemblies. Fuel being supplied to the combustor is selectively distributed between the plurality of main fuel injection and air swirler assemblies and the plurality of pilot fuel injection and air swirler assemblies in response to the level of power demand on the gas turbine engine.

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1. A combustor for a gas turbine engine having an annular combustion chamber, comprising: a bulkhead disposed at a forward end of the annular combustion chamber, the bulkhead having a radial expanse and a circumferential expanse;a plurality of main fuel injection and air swirler assemblies opening t

1. A combustor for a gas turbine engine having an annular combustion chamber, comprising: a bulkhead disposed at a forward end of the annular combustion chamber, the bulkhead having a radial expanse and a circumferential expanse;a plurality of main fuel injection and air swirler assemblies opening through the bulkhead for admitting fuel and air into the annular combustion chamber, the plurality of main fuel injection and air swirler assemblies disposed in a circumferential ring extending about the circumferential expanse of the bulkhead;a plurality of pilot fuel injection and air swirler assemblies opening through the bulkhead for admitting fuel and air into the annular combustion chamber, the plurality of pilot fuel injection and air swirler assemblies interspersed amongst and disposed in the circumferential ring of main fuel injection and air swirler assemblies, the air swirlers of the plurality of main fuel injection and air swirler assemblies adapted to impart a lower level of swirl to air passing therethrough relative the air swirlers of the plurality of pilot fuel injection and air swirler assemblies;an inboard liner extending longitudinally fore to aft from a radially inboard extent of the circumferential expanse of the bulkhead to a combustor exit;an outboard liner extending longitudinally fore to aft from a radially outboard extent of the circumferential expanse of the bulkhead to the combustor exit;a first plurality of relatively larger combustion air admission holes formed in the outboard liner, disposed in multiple subsets of multiple holes and each subset arranged in a circumferential ring aftward of the bulkhead in association with the plurality of main fuel injection and air swirler assemblies; anda first plurality of relatively smaller combustion air admission holes formed in the outboard liner, disposed in multiple subsets of multiple holes and each subset arranged in a circumferential ring aftward of the bulkhead in association with the plurality of pilot fuel injection and air swirler assemblies. 2. The combustor for a gas turbine engine as recited in claim 1 wherein the plurality of main fuel injection and air swirler assemblies and the plurality of pilot fuel injection and air swirler assemblies are disposed in alternating sequence in the circumferential ring of the plurality of main fuel injection and air swirler assemblies and the plurality of pilot fuel injection and air swirler assemblies. 3. The combustor for a gas turbine engine as recited in claim 1 further comprising: a main fuel manifold for supplying fuel to the plurality of main fuel injection and air swirler assemblies;a pilot fuel manifold for supplying fuel to the plurality of pilot fuel injection and air swirler assemblies; anda fuel control configured to selectively modulate a fuel flow between the main fuel manifold and the pilot fuel manifold. 4. The combustor for a gas turbine engine as recited in claim 1 wherein the first plurality of relatively larger combustion air admission holes and the first plurality of relatively smaller combustion air admission holes are interspersed in a common circumferential ring in the outboard liner. 5. The combustor for a gas turbine engine as recited in claim 1 wherein the first plurality of relatively larger combustion air admission holes are arranged in a first circumferential ring and the first plurality of relatively smaller combustion air admission holes are arranged in a second circumferential ring, the second circumferential ring being located forward of the first circumferential ring in the outboard liner. 6. The combustor for a gas turbine engine as recited in claim 1 further comprising: a second plurality of relatively larger combustion air admission holes formed in the inboard liner and arranged in a circumferential ring aftward of the bulkhead in association with the plurality of main fuel injection and air swirler assemblies; anda second plurality of relatively smaller combustion air admission holes formed in the inboard liner and arranged in a circumferential ring aftward of the bulkhead in association with the plurality of pilot fuel injection and air swirler assemblies. 7. The combustor for a gas turbine engine as recited in claim 6 wherein the second plurality of relatively larger combustion air admission holes and the second plurality of relatively smaller combustion air admission holes are interspersed in a common circumferential ring in the inboard liner. 8. The combustor for a gas turbine engine as recited in claim 6 wherein the second plurality of relatively larger combustion air admission holes are arranged in a first circumferential ring and the second plurality of relatively smaller combustion air admission holes are arranged in a second circumferential ring, the second circumferential ring being located forward of the first circumferential ring in the inboard liner. 9. The combustor for a gas turbine engine as recited in claim 6 wherein the second plurality of relatively larger combustion air admission holes and the second plurality of relatively smaller combustion air admission holes located in the inboard liner are staggered circumferentially relative to the first plurality of relatively larger combustion air admission holes and the first plurality of relatively smaller combustion air admission holes located in the outboard liner. 10. A method for operating a combustor of a gas turbine engine, comprising: providing a bulkhead disposed at a forward end of the annular combustion chamber, the bulkhead having a radial expanse and a circumferential expanse;providing a plurality of main fuel injection and air swirler assemblies opening through the bulkhead for admitting fuel and air into the annular combustion chamber, the plurality of main fuel injection and air swirler assemblies disposed in a circumferential ring extending about the circumferential expanse of the bulkhead;providing a plurality of pilot fuel injection and air swirler assemblies opening through the bulkhead for admitting fuel and air into the annular combustion chamber, the plurality of pilot fuel injection and air swirler assemblies interspersed amongst and disposed in the circumferential ring of main fuel injection and air swirler assemblies, the air swirlers of the plurality of main fuel injection and air swirler assemblies adapted to impart a lower level of swirl to air passing therethrough relative the air swirlers of the plurality of pilot fuel injection and air swirler assemblies;providing an inboard liner extending longitudinally fore to aft from a radially inboard extent of the circumferential expanse of the bulkhead to a combustor exit;providing an outboard liner extending longitudinally fore to aft from a radially outboard extent of the circumferential expanse of the bulkhead to the combustor exit;providing a first plurality of relatively larger combustion air admission holes formed in the outboard liner, disposed in multiple subsets of multiple holes and each subset arranged in a circumferential ring aftward of the bulkhead in association with the plurality of main fuel injection and air swirler assemblies;providing a first plurality of relatively smaller combustion air admission holes formed in the outboard liner, disposed in multiple subsets of multiple holes and each subset arranged in a circumferential ring aftward of the bulkhead in association with the plurality of pilot fuel injection and air swirler assemblies; anddistributing fuel being supplied to the combustor between the plurality of main fuel injection and air swirler assemblies and the plurality of pilot fuel injection and air swirler assemblies in response to a level of power demand on the gas turbine engine. 11. The method as recited in claim 10 further comprising the step of at lower power demand distributing fuel being supplied to the combustor relatively evenly between the plurality of main fuel injection and air swirler assemblies and the plurality of pilot fuel injection and air swirler assemblies. 12. The method as recited in claim 10 further comprising the step of at moderate (mid-range) power demand distributing fuel being supplied to the combustor predominately to the plurality of main fuel injection and air swirler assemblies.