Methods and systems for controlling thermal differential in turbine systems

1. A method for controlling a thermal differential within a compressor of a gas turbine system, said method comprising:

selecting one or more surfaces of a plurality of rotors of the compressor, wherein at least one of the surfaces selected is proximate to a wheel rim of a rotor of the plurality of rotors and at least one of within a loading location defined within the rotor and radially inward from the wheel rim; and

applying a thermal barrier coating to the one or more surfaces of the plurality of rotors to facilitate reducing a transient heat transfer rate during startup.

2. A method in accordance with claim 1 , wherein applying a thermal barrier coating further comprises applying the thermal barrier coating with a thickness of less than approximately 0.04 inches.

3. A method in accordance with claim 1 , wherein selecting a surface further comprises selecting a disc, and wherein applying a thermal barrier coating further comprises applying the thermal barrier coating to at least the disc.

4. A method in accordance with claim 1 , wherein selecting a surface further comprises selecting a slot defined within a compressor, and wherein applying a thermal barrier coating further comprises applying the thermal barrier coating within at least the slot.

5. A method in accordance with claim 1 , wherein selecting a surface further comprises selecting a plurality of discs and an aft section positioned downstream of the plurality of discs, and wherein applying a thermal barrier coating further comprises applying the thermal barrier coating to at least the aft section.

6. A method in accordance with claim 1 , wherein selecting a surface further comprises selecting the surface within a portion of a compressor.

7. A method in accordance with claim 1 , wherein the surface selected is not directly exposed to a primary hot gas path defined within the turbine system.

8. A plurality of rotors for use within a compressor of a gas turbine system, or each rotor of said plurality of rotors comprising:

one or more surfaces proximate to a wheel rim of said compressor rotor, wherein said surface is at least one of within a loading location defined within the compressor rotor and radially inward from the wheel rim; and

a thermal barrier coating applied to said one or more surfaces to facilitate reducing a transient heat transfer rate during startup.

9. A plurality of rotors in accordance with claim 8 , wherein said thermal barrier coating has a thickness of less than approximately 0.04 inches.

10. A plurality of rotors in accordance with claim 8 further comprising a disc, wherein said surface is positioned on said disc.

11. A plurality of rotors in accordance with claim 8 further comprising a circumferential slot defined within said turbine rotor, wherein said surface is positioned within said circumferential slot.

12. A plurality of rotors in accordance with claim 8 further comprising a plurality of discs and an aft section positioned downstream of said plurality of discs, wherein said surface is positioned on said aft section.

13. A plurality of rotors in accordance with claim 8 further comprising a radial slot defined within said turbine rotor, wherein said surface is positioned within said radial slot.

14. A turbine rotor in accordance with claim 8 , wherein said surface is not directly exposed to a primary hot gas path defined within the turbine system.

15. A gas turbine system comprising a compressor, said compressor comprising:

a plurality of rotors, each rotor comprising one or more surfaces, wherein at least one of the surfaces is proximate to a wheel rim of said rotor and is at least one of within a loading location defined within the rotor and radially inward from the wheel rim; and

a thermal barrier coating applied to the one or more surfaces of the plurality of rotors to facilitate reducing a transient heat transfer rate during startup.

16. A turbine system in accordance with claim 15 , wherein said thermal barrier coating has a thickness of less than approximately 0.04 inches.

17. A turbine system in accordance with claim 15 , wherein said turbine rotor further comprises a disc, wherein said surface is positioned on said disc.

18. A turbine system in accordance with claim 15 , wherein said turbine rotor further comprises a slot defined within said turbine rotor, wherein said surface is positioned within said slot.

19. A turbine system in accordance with claim 15 , wherein said turbine rotor further comprises an aft section positioned downstream of a plurality of discs of said turbine rotor, wherein said surface is positioned on said aft section.

20. A turbine system in accordance with claim 15 , wherein the surface is not directly exposed to a primary hot gas path defined within the turbine system.