Gas turbine engine compressor rotor vaporization cooling

1. A method of cooling a gas turbine engine rotor assembly, the gas turbine engine rotor assembly comprising a rotor having a disc, the method comprising the steps of:

rotating the rotor with a phase change material contained within a cavity in the rotor;

changing a phase of the phase change material at a radially outer chamber of the cavity to cool a radially outer portion of the rotor; and

returning the changed phase of the phase change material to a radially inner chamber of the cavity via a passageway having an axial width that is less than an axial width of either of the inner and outer chambers,

wherein the radially outer chamber, the radially inner chamber, and the passageway are each defined within the disc of the rotor.

2. The method according to claim 1 , comprising the step of transferring heat from a gas flow path of a compressor section to the radially outer portion of the rotor, and performing the phase changing step in response to the heat transferring step.

3. The method according to claim 1 , wherein blades are supported on the rotor, and the phase change material is isolated from the blades.

4. A gas turbine engine rotor assembly comprising:

a rotor having a disc, the rotor providing a cooling cavity, the cooling cavity has a first chamber and a second chamber fluidly connected to one another by a passageway, a portion of the rotor is configured to support a blade that is fluidly isolated from the cooling cavity; and

a phase change material is arranged in the cooling cavity, the phase change material is configured to be arranged in the first chamber in a first state and in the second chamber in a second state, the passageway is configured to carry the phase change material between the second and first chambers, wherein the passageway includes an axial width that is less than an axial width of either of the first and second chambers,

and wherein the first chamber, the second chamber, and the passageway are each defined within the disc of the rotor.

5. The gas turbine engine rotor assembly according to claim 4 , comprising an array of circumferentially spaced blades supported relative to the rotor.

6. The gas turbine engine rotor assembly according to claim 5 , wherein each of the circumferentially spaced blades includes a root, and the rotor includes a circumferential array of slots that receive the roots.

7. The gas turbine engine rotor assembly according to claim 5 , wherein each of the circumferentially spaced blades is integrally formed with the rotor.

8. The gas turbine engine rotor assembly according to claim 5 , wherein the first chamber is arranged radially inward of the blade.

9. The gas turbine engine rotor assembly according to claim 4 , wherein the rotor includes first and second portions arranged axially relative to one another and secured at axial mate faces.

10. The gas turbine engine rotor assembly according to claim 9 , wherein the axial mate faces are friction welded to one another.

11. The gas turbine engine rotor assembly according to claim 4 , wherein the first chamber includes axially extending ribs.

12. The gas turbine engine rotor assembly according to claim 4 , wherein the second chamber is arranged in a bell-shaped portion of the disc of the rotor.

13. The gas turbine engine rotor assembly according to claim 4 , wherein the phase change material is phosphorus-based.

14. The gas turbine engine rotor assembly according to claim 1 , wherein the phase change material is configured to transition from a liquid to a gas in the first chamber at a temperature of substantially 1000° F., and the phase change material is configured to transition from a gas to a liquid in the second chamber at a temperature of substantially 1000° F.

15. The gas turbine engine rotor assembly according to claim 4 , wherein the rotor includes a plug sealing the cooling cavity from an exterior of the rotor.

16. The gas turbine engine rotor assembly according to claim 4 , wherein the first and second chambers are arranged radially inward of the blade.

17. A gas turbine engine comprising:

a rotor having a disc, the rotor providing a cooling cavity, the cooling cavity has a first chamber and a second chamber that is arranged radially inward from the first chamber, the first and second chambers are fluidly connected to one another by a passageway;

an array of circumferentially spaced blades supported relative to the rotor, each of the circumferentially spaced blades are fluidly isolated from the cooling cavity; and

a phase change material is arranged in the cooling cavity, the phase change material is configured to be arranged in the first chamber in a first state and in the second chamber in a second state, the passageway is configured to carry the phase change material between the second and first chambers, wherein the passageway includes an axial width that is less than an axial width of either of the first and second chambers,

and wherein the first chamber, the second chamber, and the passageway are each defined within the disc of the rotor.

18. The gas turbine engine according to claim 17 , comprising a combustor section arranged axially between a compressor section and a turbine section, the rotor arranged in the compressor section.

19. The gas turbine engine according to claim 17 , wherein each of the circumferentially spaced blades includes a root, and the rotor includes a circumferential array of slots that receive the roots.

20. The gas turbine engine according to claim 17 , wherein each of the circumferentially spaced blades are integrally formed with the rotor.