Multi-stage axial fuel injection system with discrete air supplies

A combustor for a gas turbine engine, the combustor comprising:

• a head end section containing a fuel nozzle assembly;
• a liner extending downstream from the head end section to an aft frame and defining a combustion chamber therein;
• an outer sleeve having a forward impingement sleeve and an aft impingement sleeve;
• a first plurality of injectors disposed at a first axial location spaced apart from the head end section to direct a first fuel/air mixture through the liner, wherein the forward impingement sleeve circumferentially surrounds the first plurality of injectors, wherein a first annulus is defined between the forward impingement sleeve and the liner;
• a second plurality of injectors disposed at a second axial location downstream from the first plurality of injectors to direct a second fuel/air mixture through the liner, wherein the aft impingement sleeve circumferentially surrounds the second plurality of injectors, wherein a second annulus is defined between the aft impingement sleeve and the liner, and wherein the first annulus is fluidly isolated from the second annulus;
• wherein each of the head end section, the first plurality of injectors, and the second plurality of injectors receives a respective air supply from a compressor discharge plenum that is defined by a compressor discharge casing that at least partially surrounds the combustor, the respective air supplies being directed to only one of the fuel nozzle assembly, the first plurality of injectors, and the second plurality of injectors;
• and
• wherein the respective air supply to the second plurality of injectors is greater than each of the respective air supply to the first plurality of injectors and the respective air supply to the head end section.