Gas turbine engine component cooling assembly

Patent metadata

Field	Value
Publication number	US-10107109-B2
Application number	US-201514965227-A
Country	US
Kind code	B2
Filing date	Dec 10, 2015
Priority date	Dec 10, 2015
Publication date	Oct 23, 2018
Grant date	Oct 23, 2018

Abstract

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.

First claim

What is claimed is:

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; and
a 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; and
a 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; and
directing 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.

Assignees

• United Technologies Corp

Inventors

• Mccaffrey Michael G

Classifications

• F01D25/24

  Casings (modified for heating or cooling F01D25/14); Casing parts, e.g. diaphragms, casing fastenings (casings for rotary machines or engines in general F16M {; special arrangements in stators dealing with breaking-off of part of rotor F01D21/045}) · CPC title

• F01D5/187 Primary

  Convection cooling · CPC title

• F05D2260/202

  by film cooling · CPC title

• F05D2220/32

  in gas turbines · CPC title

• Y02T50/60

  Efficient propulsion technologies, e.g. for aircraft · CPC title