System and method for impingement cooling of turbine system components

1. A combustor, comprising:

a combustor shell defining an outer liner;

an inner liner disposed inside the combustor shell and having an inner surface defining a cavity configured to receive hot combustion gases from a combustion chamber of the combustor, and an outer surface, the combustor shell and the inner liner defining an annular flow channel therebetween; and

a segment carrier operatively connected to the inner liner and operative to receive an upper portion of the inner liner, the segment carrier and the inner liner defining a purging cavity therebetween;

wherein the inner liner includes a plurality of impingement jet holes configured to direct a flow of cooling air from the annular flow channel to the purging cavity.

2. The combustor of claim 1 , further comprising:

a retaining ring coupled to the segment carrier and being configured to protect at least a portion of the segment carrier from the hot combustion gases.

3. The combustor of claim 2 , wherein:

the plurality of impingement jet holes are configured to direct the flow of cooling air to impinge on the retaining ring to provide impingement cooling of the retaining ring.

4. The combustor of claim 3 , wherein:

the inner liner includes a conical portion; and

the plurality of impingement jet holes are formed in the conical portion of the inner liner.

5. The combustor of claim 4 , wherein:

the plurality of impingement jet holes are located approximately every 1.8° throughout the conical portion of the inner liner.

6. The combustor of claim 5 , wherein:

the annular flow channel is configured to receive the cooling air from a compressor stage of a gas turbine.

7. The combustor of claim 5 , wherein:

the combustor is a silo combustor.

8. The combustor of claim 4 , wherein:

the plurality of impingement jet holes are located approximately every 1° to 2.6° throughout the conical portion of the inner liner.

9. The combustor of claim 5 , wherein:

the retaining ring is formed from steel.

10. The combustor of claim 5 , further comprising:

a plurality of segments carried on an inner periphery of the segment carrier, the plurality of segments and the segment carrier defining a segmented zone of a hot gas passage of the combustor.

11. A gas turbine system, comprising:

a compressor; and

a combustor downstream from the compressor and including:

a combustor shell defining an outer liner;

an inner liner disposed inside the combustor shell and having an inner surface defining a cavity configured to receive hot combustion gases from a combustion chamber of the combustor, and an outer surface, the combustor shell and the inner liner defining an annular flow channel therebetween;

a segment carrier operatively connected to the inner liner and operative to

receive an upper portion of the inner liner, the segment carrier and the inner liner defining a purging cavity therebetween; and

a plurality of impingement jet holes formed in the inner liner and providing a flow passage between the annular flow channel and the purging cavity;

wherein the compressor is configured to supply compressed air to the annular flow channel;

wherein a first portion of the compressed air is used by the combustor for combustion, producing the hot combustion gases; and

wherein a second portion of the compressed air is directed through the plurality of impingement jet holes to the purging cavity to purge the purging cavity of the hot combustion gases.

12. The gas turbine system of claim 11 , wherein:

the combustor further includes a retaining ring coupled to the segment carrier, the retaining ring being configured to protect at least a portion of the segment carrier from the hot combustion gases;

wherein the plurality of impingement jet holes are configured to direct the second portion of the compressed air to impinge on the retaining ring to provide impingement cooling of the retaining ring.

13. The gas turbine system of claim 12 , wherein:

the inner liner includes a conical portion; and

the plurality of impingement jet holes are formed in the conical portion of the inner liner.

14. The gas turbine system of claim 13 , wherein:

the plurality of impingement jet holes are located approximately every 1.8° throughout a circumference of the conical portion of the inner liner.

15. The gas turbine system of claim 14 , wherein:

the combustor is a silo combustor.

16. The gas turbine system of claim 14 , wherein:

the retaining ring is formed from steel.

17. The gas turbine system of claim 13 , wherein:

the plurality of impingement jet holes are located approximately every 1° to 2.6° throughout a circumference of the conical portion of the inner liner.

18. A method of cooling a component in a gas turbine system, comprising the steps of:

passing compressed air into an annular channel defined between an outer surface of an inner liner of a combustor of a gas turbine system and one of a middle liner and an outer shell of the combustor, the inner liner being configured to receive a flow of hot combustion gas from a combustion zone of the combustor there through;

passing a portion of the compressed air in the annular channel through a plurality of impingement jet holes in the inner liner such that the compressed air impinges on a component exposed to the flow of hot combustion gas to provide for impingement cooling of the component; and wherein:

the component is a retaining ring of the combustor, the retaining ring shielding a segment carrier of the combustor from the flow of hot combustion gas.

19. The method according to claim 18 , wherein:

the segment carrier receives an upper portion of the inner liner, the segment carrier and the inner liner defining a purging cavity therebetween;

wherein the plurality of impingement jet holes direct the portion of the compressed air into the purging cavity to clear the purging cavity of the hot combustion gas.