Circumferentially and axially staged can combustor for gas turbine engine
A combustor section for a gas turbine engine includes a can combustor with a combustion chamber. A pilot fuel injection system is in axial communication with the combustion chamber. A main fuel injection system is in radial communication with the combustion chamber. The main fuel injection system includes a multiple of first main fuel nozzles that circumferentially alternate with a multiple of second main fuel nozzles.
1. A combustor section for a gas turbine engine, comprising:
a can combustor including a combustion chamber;
a pilot fuel injection system in axial communication with the combustion chamber; and
a main fuel injection system in radial communication with the combustion chamber, the main fuel injection system comprising a main fuel flow path that delivers fuel to a plurality of fuel stems, where each of the plurality of fuel stems constantly delivers fuel to an associated first main fuel nozzle during operation of the gas turbine engine, and where each of the plurality of fuel stems delivers fuel to an associated second main fuel nozzle that includes a valve in fluid communication with the fuel stem to selectively communicate fuel to the associated second main fuel nozzle.
2. The combustor section as recited in claim 1 , wherein for each of the fuel stems the associated first main fuel nozzle and the associated second main fuel nozzle are fueled in pairs.
3. The combustor section as recited in claim 1 , wherein the pilot fuel injection system includes a multiple of forward fuel injectors, one of the forward fuel injectors within each of a multiple of can combustors.
4. A gas turbine engine comprising:
a compressor section;
a turbine section;
a combustor section between the compressor section and the turbine section, the combustor section including a multiple of can combustors each including a combustion chamber;
a pilot fuel injection system in axial communication with the combustion chamber of each of the can combustors; and
a main fuel injection system in radial communication with the combustion chamber of each of the can combustors, the main fuel injection system comprising a main fuel flow path that delivers fuel to a plurality of fuel stems, where each of the plurality of fuel stems constantly delivers fuel to an associated first main fuel nozzle during operation of the gas turbine engine, and where each of the plurality of fuel stems delivers fuel to an associated second main fuel nozzle that includes a valve in fluid communication with the fuel stem to selectively communicate fuel to the associated second main fuel nozzle.
5. The gas turbine engine as recited in claim 4 , wherein the multiple of can combustors communicate with a transition section in communication with the turbine section.
6. The gas turbine engine as recited in claim 4 , wherein the pilot fuel injection system includes a multiple of forward fuel injectors, one of the forward fuel injectors within each of the multiple of can combustors.
7. The gas turbine engine as recited in claim 6 , wherein for each of the fuel stems the associated first main fuel nozzle and the associated second main fuel nozzle fueled in pairs.
8. A method of communicating fuel to a combustor section of a gas turbine engine, the method comprising:
communicating pilot fuel axially into a combustion chamber;
communicating fuel radially inboard into the combustion chamber; and
circumferentially varying the fuel communicating radially inboard into the combustion chamber via a main fuel flow path that delivers fuel to a plurality of fuel stems, where each of the plurality of fuel stems constantly delivers fuel to an associated first main fuel nozzle during operation of the gas turbine engine, and where each of the plurality of fuel stems delivers fuel to an associated second main fuel nozzle that includes a valve in fluid communication with the fuel stem to selectively communicate fuel to the associated second main fuel nozzle.