Methods and apparatus for controlling cooling air temperature in gas turbine engines

1. A cooling system for a gas turbine engine including a compressor, a combustor, a high-pressure turbine, and a high-pressure turbine cooling air supply duct, coupled together in serial flow arrangement, said cooling system comprising: a storage tank; a working fluid stored within the storage tank; a manifold coupled in flow communication between the combustor and the high-pressure turbine, said cooling system configured to discharge the working fluid from said storage tank into the high-pressure turbine cooling air supply duct using said manifold to facilitate reducing an operating temperature of the compressed airflow channeled to the high pressure turbine and a return conduit comprising a first end coupled to an outlet of said manifold and a second end coupled to an inlet of said storage tank, said return conduit configured to channel an unused portion of said working fluid to the inlet of said storage tank.

2. A cooling system in accordance with claim 1 wherein said storage tank, said manifold, and the gas turbine engine are installed on an aircraft.

3. A cooling system in accordance with claim 1 wherein said manifold comprises a toroidal manifold.

4. A cooling system in accordance with claim 1 wherein the gas turbine engine includes an off-take duct and a high-pressure turbine inducer, said manifold positioned proximate to an entrance of the off-take duct feeding the high-pressure turbine inducer.

5. A cooling system in accordance with claim 1 further comprising a pump coupled in flow communication with said storage tank and said manifold, said pump selectively energized to channel the working fluid from said storage tank into said manifold.

6. A cooling system in accordance with claim 5 wherein said storage tank and said pump are each placed in a selected core cowl space proximate to a plane of a high-pressure turbine vane inlet to facilitate maintaining the working fluid at a predetermined operational temperature.

7. A cooling system in accordance with claim 1 further comprising a plurality of nozzles coupled to said manifold, said nozzles configured to inject a predetermined quantity of working fluid into the airflow stream entering the high-pressure turbine.

8. A gas turbine engine assembly comprising:

a core gas turbine engine comprising

a compressor;

a combustor downstream from said compressor; and

a high-pressure turbine downstream from said combustor; and

a cooling system comprising

a storage tank;

a working fluid stored within said storage tank;

a manifold coupled to a duct that supplies cooling air to the high pressure turbine, discharging the fluid into the turbine cooling air supply duct, and thereby reducing the temperature of the compressed cooling airflow channeled to the high pressure turbine; and

a return conduit comprising a first end coupled to an outlet of said manifold and a second end coupled to an inlet of said storage tank, said return conduit configured to channel an unused portion of said working fluid to the inlet of said storage tank.

9. A gas turbine engine assembly in accordance with claim 8 wherein said cooling system and said core gas turbine engine are installed on an aircraft.

10. A gas turbine engine assembly in accordance with claim 8 wherein said manifold comprises a toroidal manifold.

11. A gas turbine engine assembly in accordance with claim 8 wherein said core gas turbine engine further comprises an off-take duct and a high-pressure turbine inducer, said manifold positioned proximate to an entrance of said off-take duct feeding said high-pressure turbine inducer.

12. A gas turbine engine assembly in accordance with claim 8 further comprising a pump coupled in flow communication with said storage tank and said manifold, said pump selectively energized to channel the working fluid from said storage tank into said manifold.

13. A gas turbine engine assembly in accordance with claim 11 wherein said storage tank and said pump are each placed in a selected core cowl space proximate to a plane of a high-pressure turbine vane inlet to facilitate maintaining the working fluid at a predetermined operational temperature.

14. A gas turbine engine assembly in accordance with claim 8 further comprising a plurality of nozzles coupled to said manifold, said nozzles configured to inject a predetermined quantity of working fluid into the airflow stream entering said high-pressure turbine.