Gas turbine engine with high and intermediate temperature compressed air zones

Patent metadata

Field	Value
Publication number	US-9175604-B2
Application number	US-201113227769-A
Country	US
Kind code	B2
Filing date	Sep 8, 2011
Priority date	Sep 8, 2011
Publication date	Nov 3, 2015
Grant date	Nov 3, 2015

Abstract

A gas turbine combustor ( 26 ) disposed in a combustion air plenum ( 65 ) and a transition piece ( 28 ) for the combustor disposed in a separate cooling air plenum ( 58, 67 ). The combustion air plenum may receive combustion air ( 50 ) from a high-pressure compressor stage ( 22 A). The cooling air plenum may receive cooling air ( 52 ) from an intermediate-pressure compressor stage ( 22 B) at lower temperature and pressure than the combustion air. This cools the transition piece using less air than prior systems, thus making the gas turbine engine ( 20 ) more efficient and less expensive, because less expensive materials are needed and/or higher combustion temperatures are allowed. The cooling air may exit the cooling air plenum through holes ( 62 ) in a downstream portion ( 61 ) of the transition piece. An outer wall ( 72 ) on the transition piece may provide forced convection along the transition piece.

First claim

The invention claimed is:

1. A gas turbine engine, comprising:
a combustor within a combustion air plenum, wherein the combustor intakes combustion air from the combustion air plenum;
a compressor comprising a final stage that supplies the combustion air to the combustion air plenum; and
a transition piece within a cooling air plenum, wherein the cooling, air plenum receives cooling air bled from a bleed port of an intermediate stage of the compressor;
wherein the transition piece channels a combustion gas from the combustor.

2. The gas turbine engine of claim 1 , further comprising:
cooling air exit holes in a downstream 5-30% of a length of the transition piece relative to a combustion gas flow direction therein;
wherein the cooling air passes into the transition piece through the cooling air exit holes.

3. The gas turbine engine of claim 1 , wherein the cooling air plenum is separated from the combustion air plenum by a partition there between.

4. The gas turbine engine of claim 1 , wherein:
the combustion air plenum is annular about a turbine axis;
the cooling air plenum is annular about the axis; and
the combustion air plenum is surrounded by the cooling air plenum.

5. The gas turbine engine of claim 1 , wherein:
the cooling air plenum is annular about a turbine axis;
said combustor is one of a plurality of combustors in an annular array within the combustion air plenum which is disposed within the cooling air plenum.

6. The gas turbine engine of claim 5 , wherein each of the combustors is enclosed by an inner shell that defines a discrete combustion air plenum for the combustor within the cooling air plenum.

7. The gas turbine engine of claim 1 , further comprising:
an upstream enclosure around only an upstream portion of the transition piece relative to a combustion gas flow direction therein, wherein the upstream enclosure is open to the combustion air plenum; and
combustion air exit holes only in the upstream portion of the transition piece for the combustion air to exit the upstream enclosure into the transition piece.

8. The gas turbine engine of claim 7 , wherein the upstream portion of the transition piece is selected within an upstream 5-30% of a length of the transition piece relative to the combustion gas flow direction therein, and a downstream portion of the transition piece is selected within a downstream 5-30% of a length of the transition piece relative to the combustion gas flow direction therein.

9. The gas turbine engine of claim 1 , further comprising:
an outer wall around the transition piece;
cooling air entry holes only in an upstream portion of the outer wall relative to a combustion gas flow direction within the transition piece; and
cooling air exit holes only in a downstream portion of the transition piece relative to the combustion gas flow direction within the transition piece;
wherein the cooling air enters the cooling air entry holes, flows within the outer wall along the transition piece, then passes through the cooling air exit holes into the transition piece.

10. The gas turbine engine of claim 9 , wherein the upstream portion of the outer wall is selected within an upstream 0-30% of a length of the outer wall relative to the combustion gas flow direction within the transition piece, and the downstream portion of the transition piece is selected within a downstream 5-30% of a length of the transition piece relative to the combustion gas flow direction within the transition piece.

11. A gas turbine engine comprising:
a combustor within a combustion air plenum, wherein the combustion air plenum receives compressed combustion air from a final stage of a compressor of the gas turbine engine and the combustor receives the combustion air from the combustion air plenum;
a transition piece within a cooling air plenum; wherein the cooling air plenum receives compressed cooling air bled from a bleed port of an intermediate stage of the compressor of the gas turbine engine;
wherein the transition piece channels combustion gas from the combustor to an inlet of a turbine;
cooling air exit holes in a downstream 5-30% of a length of the transition piece relative to a combustion gas flow direction therein;
wherein the cooling air passes into the transition piece through the cooling air exit holes; and
wherein the cooling air plenum is separated from the combustion air plenum by a partition there between.

12. A gas turbine engine, comprising:
an outer shell around a combustion system, the outer shell defining a cooling air plenum;
a combustion air plenum defined by an inner shell within the cooling air plenum;
a combustor within the combustion air plenum, wherein the combustor receives combustion air from the combustion air plenum;
a transition piece within the cooling air plenum, wherein the transition piece channels combustion gas from the combustor to an inlet of a turbine;
wherein the cooling air plenum receives compressed cooling air bled from a bleed port of an intermediate stage of a compressor; and
cooling air exit holes in a downstream portion of a length of the transition piece relative to a combustion gas flow direction therein;
wherein the cooling air passes into the transition piece through the cooling air exit holes; and wherein the inner shell is surrounded by the cooling air.

13. The gas turbine engine of claim 12 , wherein:
the combustion air plenum is annular about an axis of the turbine;
the cooling air plenum is annular about the axis of the turbine.

14. The gas turbine engine of claim 13 , wherein:
the cooling air plenum is annular about an axis of the turbine;
said combustor is one of a plurality of combustors in a circular array within the combustion air plenum.

15. The gas turbine engine of claim 14 , wherein each of the combustors is enclosed by a respective inner shell that defines a discrete combustion air plenum for the combustor within the cooling air plenum.

16. The gas turbine engine of claim 12 , further comprising:
an upstream enclosure around only an upstream 5-30% of the transition piece relative to the combustion gas flow direction therein, wherein the upstream enclosure is open to the combustion air plenum; and
combustion air exit holes in the upstream 5-30% of the transition piece for the combustion air to exit the upstream enclosure into the transition piece.

17. The gas turbine engine of claim 12 , further comprising:
an outer wall around the transition piece;
cooling air entry holes in an upstream 5-30% of the outer wall relative to the combustion gas flow direction in the transition piece; and
the cooling air exit holes in a downstream 5-30% of the transition piece relative to the combustion gas flow direction therein;
wherein the cooling air enters the cooling air entry holes, flows within the outer wall along the transition piece, and then passes through the cooling air exit holes into the transition piece.

Assignees

• Charron Richard C
• Little David A
• Siemens Energy Inc

Inventors

• Charron Richard C
• Little David A

Classifications

• F23R3/26

  Controlling the air flow · CPC title

• F02C3/14 Primary

  characterised by the arrangement of the combustion chamber in the plant (combustion chambers per se F23R; F02C3/205 takes precedence) · CPC title

• F23R3/04

  Air inlet arrangements · CPC title

• Y02E20/16

  Combined cycle power plant [CCPP], or combined cycle gas turbine [CCGT] · CPC title