Staged fuel and air injection in combustion systems of gas turbines

1. A gas turbine that comprises:

a combustor coupled to a turbine that together define a working fluid flowpath, the working fluid flowpath extending aftward along a longitudinal axis from a forward end defined by a forward injector in the combustor, through an interface at which the combustor connects to the turbine, and then through the turbine to an aftward end defined therein;

a compressor discharge cavity formed about the working fluid flowpath for receiving a combustor air supply delivered thereto by a compressor;

a staged injection system that includes the forward injector and, axially spaced aftward therefrom along the longitudinal axis of the working fluid flowpath, a staged injector;

a stator blade positioned within a row of circumferentially spaced stator blades in the turbine, the stator blade comprising an airfoil extending across the working fluid flowpath between an inboard sidewall and an outboard sidewall;

fuel directing structure configured to apportion a combustor fuel supply between the forward injector and the staged injector; and

air directing structure for apportioning the combustor air supply between the forward injector and the staged injector;

wherein the air directing structure includes a one-way continuous coolant flowpath, the coolant flowpath comprising:

an intake section that comprises an upstream port that is fluidly coupled to the compressor discharge cavity and a downstream port that is formed through one of the inboard sidewall and the outboard sidewall;

an outtake section that comprises a downstream port that is fluidly coupled to the staged injector and an upstream port formed through whichever of the inboard sidewall and the outboard sidewall is the same as the one that the downstream port of the intake section is formed through; and

a cooling circuit extending through an interior of the stator blade connecting the intake section to the outtake section, wherein the cooling circuit comprises an upstream end that connects to the downstream port of the intake section and a downstream end that connects to the upstream port of the outtake section;

further comprising:

a flow annulus surrounding the working fluid flowpath; and

an axial partition positioned within the flow annulus;

wherein:

the row of stator blades comprises a forward most row of stator blades in the turbine; and

the intake section and the outtake section of the coolant flowpath comprise adjacent axial sections of the flow annulus that reside to each side of the axial partition, the axial partition configured to fluidly seal each from the other.

2. The gas turbine according to claim 1 , wherein the interior of the stator blade through which the cooling circuit extends comprises at least one of the inboard sidewall and the outboard sidewall.

3. The gas turbine according to claim 1 , wherein the interior of the stator blade through which the cooling circuit extends comprises the airfoil.

4. The gas turbine according to claim 1 , wherein the downstream port of the intake section and the upstream port of the outtake section are each formed through the inboard sidewall.

5. The gas turbine according to claim 1 , wherein the downstream port of the intake section and the upstream port of the outtake section are each formed through the outboard sidewall.

6. The gas turbine according to claim 1 , wherein the staged injector intersects the flow annulus so to attain an injection point within the working fluid flowpath; and

wherein, relative an axial position of the injection point, a forward annulus section is defined to a forward side of the injection point, and an aftward annulus section is defined to an aftward side of the injection point.

7. The gas turbine according to claim 6 , wherein the axial partition is positioned in the flow annulus so to axially coincide with the staged injector;

wherein the axial partition is configured for fluidly sealing the forward annulus section from the aftward annulus section such that:

substantially all of the air of the combustor air supply flowing into the forward annulus section is directed to the forward injector; and

substantially all of the air of the combustor air supply flowing into the aftward annulus section is directed to the staged injector.

8. The gas turbine according to claim 1 ,

wherein a flowpath wall defines the working fluid flowpath, the flowpath wall having a hot side, which faces the working fluid flowpath, and, opposite the hot side, a cold side, which faces the flow annulus; and

wherein an annulus wall surrounds and is offset from the flowpath wall so to form the flow annulus therebetween, the annulus wall having a hot side, which faces the flow annulus, and a cold side, which fluidly communicates with the compressor discharge cavity.

9. The gas turbine according to claim 8 , wherein the upstream port of the intake section of the coolant flowpath comprises openings formed through a corresponding section of the annulus wall for allowing a portion of the combustor air supply entry into the intake section; and

wherein the openings are configured for metering the portion of the combustor air supply flowing to the staged injector relative a remaining portion of the combustor air supply flowing to the forward injector.

10. The gas turbine according to claim 9 , wherein the openings formed through the annulus wall comprises impingement ports configured for cooling the cold side of the flowpath wall.

11. The gas turbine according to claim 9 , wherein the metering the combustor air supply includes directing at least 30% of the combustor air supply to the staged injector.

12. The gas turbine according to claim 9 , wherein the metering the combustor air supply includes directing at least 50% of the combustor air supply to the staged injector.

13. The gas turbine according to claim 8 , wherein:

within the turbine:

the flowpath wall comprises an inboard flowpath wall that defines an inboard boundary of the working fluid flowpath and an outboard flowpath wall that defines an outboard boundary of the working fluid flowpath; and

the flow annulus comprises an inboard flow annulus and an outboard flow annulus, and the annulus wall comprises an inboard annulus wall and an outboard annulus wall, wherein the inboard sidewall and the outboard sidewall comprise, respectively, axial sections of the inboard flowpath wall and the outboard flowpath wall;

the inboard annulus wall is offset from the inboard flowpath wall so to form the inboard flow annulus therebetween; and

the outboard annulus wall is offset from the outboard flowpath wall so to form the outboard flow annulus therebetween; and

within the combustor:

the flowpath wall comprises an inner radially wall; and

the annulus wall comprises an outer radial wall formed about the inner radial wall, the inner radial wall and the outer radial wall concentrically arranged about the longitudinal axis of the working fluid flowpath.

14. The gas turbine according to claim 13 , wherein the axial partition is positioned within the inboard flow annulus; and

wherein:

the intake section and the outtake section of the coolant flowpath comprise adjacent axial sections of the inboard flow annulus;

the axial partition is positioned within an axial range defined between a leading edge and a trailing edge of the stator blade; and

the intake section of the coolant flowpath resides to an aftward side of the axial partition and the outtake section of the coolant flowpath resides to a forward side of the axial partition.

15. The gas turbine according to claim 13 , wherein the axial partition is positioned within the outboard flow annulus; and

wherein:

the intake section and the outtake section of the coolant flowpath comprise adjacent axial sections of the outboard flow annulus;

the axial partition is positioned within an axial range defined between a leading edge and a trailing edge of the stator blade; and

the intake section of the coolant flowpath resides to an aftward side of the axial partition and the outtake section of the coolant flowpath resides to a forward side of the axial partition.

16. The gas turbine according to claim 13 , wherein the staged injector is positioned within the combustor, the staged injector being formed through the inner radial wall for enabling injection of a fuel and air mixture into the working fluid flowpath.

17. The gas turbine according to claim 13 , wherein the staged injector is positioned within the turbine; and

wherein the cooling circuit the coolant flowpath comprises a serpentine path through the stator blade.

18. The gas turbine according to claim 13 , wherein the staged injector comprises a nozzle extending through the flow annulus between the flow annulus wall and the flowpath wall, and the downstream port of the outtake section of the coolant flowpath comprises an opening in the nozzle; and

wherein the outboard annulus wall corresponding to the outtake section of the coolant flowpath comprises separating structure that fluidly isolates flow therethrough from flow in the compressor discharge cavity.

19. The gas turbine according to claim 13 , further comprising:

surface ports formed through an outer surface of the airfoil of the stator blade; and

connector channels branching from the cooling circuit that fluidly coupling the surface ports to the coolant flowpath.

20. The gas turbine according to claim 13 , wherein the staged injector comprises one of: a first injector type; a second injector type; and a third injector type; and

wherein:

the first injector type comprises a fuel and air mixing chamber that resides exterior to the flow annulus;

the second injector type comprises a jutting nozzle that extends from the flowpath wall into the working fluid flowpath; and

the third injector type comprises an injection port formed substantially flush with the hot side of the flowpath wall.

21. The gas turbine according to claim 13 , wherein the combustion zone extends from the forward injector to the interface at which the combustor connects to the turbine; and

wherein the combustor comprises reference planes including: a forward reference plane, a mid reference plane, and an aftward reference plane, each of which comprising planes aligned substantially perpendicular to a central axis of the working fluid flowpath;

wherein:

the forward reference plane aligns with a forward end of the combustion zone;

the mid reference plane aligns with an axial midpoint of the combustion zone; and

the aftward reference plane aligns with an aftward end of the combustion zone.

22. The gas turbine according to claim 21 , wherein the staged injector is positioned between the mid reference plane and the aftward reference plane.

23. The gas turbine according to claim 21 , wherein the staged injector is positioned between the aftward reference plane and the row of stator blades.