초록 ▼

A rotor blade assembly includes at least one rotor blade including an airfoil that has a leading edge internal cooling passage that extends through the rotor blade and is in communication with cooling holes along a leading edge of the airfoil. A compression portion includes a compression passage tha

A rotor blade assembly includes at least one rotor blade including an airfoil that has a leading edge internal cooling passage that extends through the rotor blade and is in communication with cooling holes along a leading edge of the airfoil. A compression portion includes a compression passage that is in communication with the leading edge internal cooling passage.

대표청구항 ▼

1. A rotor blade assembly comprising: at least one rotor blade including an airfoil having a leading edge internal cooling passage extending through the rotor blade in communication with cooling holes along a leading edge of the airfoil;a compression portion including a compression passage in commun

1. A rotor blade assembly comprising: at least one rotor blade including an airfoil having a leading edge internal cooling passage extending through the rotor blade in communication with cooling holes along a leading edge of the airfoil;a compression portion including a compression passage in communication with the leading edge internal cooling passage; anda trailing edge internal cooling passage in fluid communication with a passage at least partially defined by an axially forward surface of a rotor disk that is spaced from an axially downstream surface of the compression portion. 2. The rotor blade assembly of claim 1, wherein the compression passage includes an inlet having a first cross-sectional area and an outlet having a second cross-sectional area smaller than the first cross sectional area. 3. The rotor blade assembly of claim 1, wherein the leading edge internal cooling passage includes an aerodynamic shape and the aerodynamic shape tapers in a radially outward direction. 4. The rotor blade assembly of claim 1, wherein the compression portion includes a compressor wheel forming a ring. 5. The rotor blade assembly of claim 1, wherein the compression portion is integral with the at least one rotor blade and includes a scoop. 6. The rotor blade assembly of claim 1, wherein the leading edge internal cooling passage is configured to receive cooling air from an intermediate stage of a compressor section to provide bleed air to the leading edge internal cooling passage. 7. The rotor blade assembly of claim 1, wherein the trailing edge internal cooling passage is fluidly isolated from the compression portion. 8. The rotor blade assembly of claim 1, wherein the passage at least partially defined by an axially forward surface of a rotor disk is upstream of a disk slot passage and the disk slot passage extends in an axial direction is in fluid communication with an inlet to the trailing edge internal cooling passage. 9. A gas turbine engine assembly comprising: a compressor section for providing a bleed air source;a rotor blade assembly including: at least one rotor blade including an airfoil having a leading edge internal cooling passage extending through the rotor blade in communication with cooling holes along a leading edge of the airfoil;a compression portion including a compression passage in communication with the leading edge internal cooling passage and the bleed air source; anda trailing edge internal cooling passage in fluid communication with a passage at least partially defined by an axially forward surface of a rotor disk that is spaced from an axially downstream surface of the compression portion. 10. The gas turbine engine of claim 9, wherein the bleed air source is from an intermediate stage of the compressor section. 11. The gas turbine engine of claim 10, wherein the compressor section includes a high pressure compressor and the intermediate stage is located in the high pressure compressor. 12. The gas turbine engine of claim 9, wherein the compression passage includes an inlet having a first cross-sectional area and an outlet having a second cross-sectional area smaller than the first cross sectional area. 13. The gas turbine engine of claim 9, wherein the leading edge internal cooling passage includes an aerodynamic shape and the aerodynamic shape tapers in a radially outward direction. 14. The gas turbine engine of claim 9, wherein the compression portion includes a compressor wheel forming a ring. 15. The gas turbine engine of claim 9, wherein the compression portion is integral with the at least one rotor blade and includes a scoop. 16. The gas turbine engine of claim 9, wherein the passage at least partially defined by an axially forward surface of a rotor disk is upstream of a disk slot passage and the disk slot passage extends in an axial direction and is in fluid communication with an inlet to the trailing edge internal cooling passage. 17. A method of cooling a rotor blade comprising directing bleed air from a compressor section into a rotor cavity;directing a first portion of the bleed air at a first pressure into a compression portion for increasing a pressure of the bleed air entering a leading edge internal cooling passage of at least one rotor blade to a second pressure; anddirecting a second portion of the bleed air at the first pressure into a trailing edge internal cooling passage, wherein the trailing edge internal cooling passage in fluid communication with a passage at least partially defined by an axially forward surface of a rotor disk that is spaced from an axially downstream surface of the compression portion. 18. The method of claim 17, including rotating the compression portion with the at least one rotor blade. 19. The method of claim 18, including extracting the bleed air from an intermediate stage of a high pressure compressor in the compressor section. 20. The method of claim 17, wherein the first pressure is greater than the second pressure.