Airflow distribution to a low emissions combustor

1. A gas turbine combustor having increased combustion stability, said combustor comprising:

A flow sleeve having a first end, a second end, and a plurality of first holes in a plurality of axially spaced rows located proximate said second end;

A combustion liner located radially within said flow sleeve thereby forming a first passage therebetween;

At least one fuel nozzle for injecting a fuel to mix with air in said combustion liner; and,

A plurality of vanes, said vanes fixed to said flow sleeve proximate said plurality of first holes and extending radially inward towards said combustion liner into said first passage such that said plurality of vanes significantly remove the tangential velocity component from air entering said first passage through said plurality of first holes, thereby directing said air in a substantially axial direction towards said flow sleeve first end, wherein said plurality of vanes are equal in number to a quantity of first holes in a row and wherein said vanes are circumferentially offset from said holes in at least one of said axially spaced rows while bisecting remaining holes in said flow sleeve.

2. The gas turbine combustor of claim 1 wherein said plurality of vanes are equally spaced circumferentially about said flow sleeve.

3. The gas turbine combustor of claim 1 wherein said vanes have an axial length, a first wall and a second wall, thereby establishing a vane thickness, said first wall and second wall terminating in an edge opposite said flow sleeve.

4. The gas turbine combustor of claim 3 wherein said vane edge is rounded.

5. The gas turbine combustor of claim 3 wherein said vane edge is spaced a radial distance from said combustion liner.

6. The gas turbine combustor of claim 5 wherein said radial distance is up to 0.350 inches.

7. The gas turbine combustor of claim 1 wherein said plurality of first holes are spaced axially in circumferential rows about said flow sleeve.

8. The gas turbine combustor of claim 7 wherein said plurality of first holes have a diameter of up to 2.00 inches.

9. The method for reducing pressure drop across a gas turbine combustor, said method comprising the steps:

Providing a gas turbine combustor comprising a flow sleeve having a first end, a second end, and a plurality of first holes in a plurality of axially spaced rows located proximate said second end, a combustion liner located radially within said flow sleeve thereby forming a first passage therebetween, at least one fuel nozzle for injecting a fuel to mix with air in said combustion liner, and a plurality of vanes, said vanes fixed to said flow sleeve proximate said plurality of first holes and extending radially inward towards said combustion liner into said first passage, wherein said plurality of vanes are equal in number to a quantity of first holes in a row and wherein said vanes are circumferential offset from said holes in at least one of said axially spaced rows while bisecting remaining holes in said flow sleeve;

Directing a flow of compressed air through said plurality of first holes, into said first passage, and between said plurality of vanes;

Straightening said flow of compressed air by way of said plurality of vanes to significantly remove the tangential velocity component from said flow of compressed air and then directing said flow of compressed air in a substantially axial direction towards said flow sleeve first end, wherein pressure drop across said combustor from said flow sleeve second end to said flow sleeve first end is reduced by mechanically straightening said flow of compressed air through said plurality of vanes.

10. The method of claim 9 wherein said plurality of vanes are equally spaced circumferentially about said flow sleeve.

11. The method of claim 9 wherein said vanes have an axial length, a first wall and a second wall, thereby establishing a vane thickness, said first wall and second wall terminating in an edge opposite said flow sleeve.

12. The method of claim 11 wherein said vane edge is rounded.

13. The method of claim 11 wherein said vane edge is spaced a radial distance from said combustion liner.

14. The method of claim 13 wherein said radial distance is up to 0.350 inches.

15. A gas turbine combustor having a more uniform circumferential air flow distribution, said combustor comprising:

A flow sleeve having a first end, a second end, and a plurality of first holes in a plurality of axially spaced rows located proximate said second end;

A combustion liner located radially within said flow sleeve thereby forming a first passage therebetween;

At least one fuel nozzle for injecting a fuel to mix with air in said combustion liner; and

A plurality of vanes fixed to said flow sleeve proximate said plurality of first holes and extending radially inward into said first passage towards said combustion liner, wherein said plurality of vanes are equal in number to a quantity of first holes in a row and wherein said vanes are circumferentially offset from said holes in at least one of said axially spaced rows while bisecting remaining holes in said flow sleeve so as to provide substantially uniform air flow to areas between each of said vanes.

16. The gas turbine of claim 15 wherein said plurality of vanes are equally spaced circumferentially about said flow sleeve.

17. The gas turbine combustor of claim 16 wherein said vanes have an axial length, a first wall and a second wall, thereby establishing a van thickness, said first wall and second wall terminating in an edge opposite said flow sleeve.

18. The gas turbine combustor of claim 17 wherein said vane edge is rounded.

19. The gas turbine combustor of claim 15 wherein said radial distance is up to 0.350 inches.