Gas turbine engine with vane having a cooling inlet

Patent metadata

Field	Value
Publication number	US-10451084-B2
Application number	US-201514941995-A
Country	US
Kind code	B2
Filing date	Nov 16, 2015
Priority date	Nov 16, 2015
Publication date	Oct 22, 2019
Grant date	Oct 22, 2019

Abstract

An apparatus and method of cooling a hot portion of a gas turbine engine, such as a multi-stage compressor of a gas turbine engine, by reducing an operating air temperature in a space between a seal and a blade post of adjacent stages by routing cooling air through an inlet in the vane, passing the routed cooling air through the vane, and emitting the routed cooling air into the space.

First claim

What is claimed is:

1. A compressor for a gas turbine engine comprising:
an outer casing having circumferentially spaced vanes arranged in axially spaced groups of vanes, wherein each vane comprises a pressure side and a suction side and extends axially between a leading edge and a trailing edge; and
a rotor located within the outer casing and having circumferentially spaced blades arranged in axially spaced groups of blades in alternating axially arrangement with the groups of vanes to define axially arranged pairs of vanes and blade, with each pair forming a compressor stage;
the compressor stages having a circumferential seal extending between the rotor and the vanes to fluidly seal axially adjacent compressor stages; and
a cooling air circuit passing through the vanes and having an inlet located on one of the pressure or suction side of the vanes and an outlet at the rotor upstream of a corresponding seal for the vanes, wherein the inlet extends between the leading edge and the trailing edge of the vanes, wherein the inlet is located in a mid-span area of the vane.

2. The compressor of claim 1 , wherein the inlet is elongated in the flow direction.

3. The compressor of claim 1 , wherein the inlet is located on a pressure side of the vane.

4. The compressor of claim 1 , wherein the cooling air circuit is provided in at least some of the vanes in the most downstream compressor stage.

5. The compressor of claim 1 , further comprising an inner ring located within the outer casing and supporting the vanes of the compressor stage at a root of the vane and the inner ring defines a circumferential channel forming part of the cooling air circuit.

6. The compressor of claim 5 , wherein the outlet of the cooling air circuit is formed in the inner ring.

7. The compressor of claim 6 , wherein the seal comprises a honeycomb element mounted to the inner ring and fingers extending from the rotor abut the honeycomb element.

8. A compressor for a gas turbine engine comprising:
an outer casing having circumferentially spaced vanes arranged in axially spaced groups of vanes, wherein each vane comprises a pressure side and a suction side and extends axially between a leading edge and a trailing edge; and
a rotor located within the outer casing and having circumferentially spaced blades arranged in axially spaced groups of blades in alternating axially arrangement with the groups of vanes to define axially arranged pairs of vanes and blade, with each pair forming a compressor stage;
the compressor stages having a circumferential seal extending between the rotor and the vanes to fluidly seal axially adjacent compressor stages; and
a cooling air circuit passing through the vanes and having an inlet located on one of the pressure or suction side of the vanes and an outlet at the rotor upstream of a corresponding seal for the vanes, wherein the inlet extends between the leading edge and the trailing edge of the vanes, wherein the inlet is located along the span where the coolest air flows over the vane.

9. A compressor for a gas turbine engine comprising:
an outer casing having circumferentially spaced vanes arranged in axially spaced groups of vanes, wherein each vane comprises a pressure side and a suction side and extends axially between a leading edge and a trailing edge; and
a rotor located within the outer casing and having circumferentially spaced blades arranged in axially spaced groups of blades in alternating axially arrangement with the groups of vanes to define axially arranged pairs of vanes and blade, with each pair forming a compressor stage;
the compressor stages having a circumferential seal extending between the rotor and the vanes to fluidly seal axially adjacent compressor stages; and
a cooling air circuit passing through the vanes and having an inlet located on one of the pressure or suction side of the vanes and an outlet at the rotor upstream of a corresponding seal for the vanes, wherein the inlet extends between the leading edge and the trailing edge of the vanes, wherein the rotor comprises posts and the cooling air circuit outlet emits the cooling air toward the post upstream of the vane.

10. The compressor of claim 9 , wherein a space between the posts of one compressor stage and seal for a downstream compressor stage define a seal cavity and the cooling air circuit outlet emits cooling air into the seal cavity.

11. A method of cooling a multi-stage compressor of a gas turbine engine, the method comprising routing compressor air through an inlet located on one of a pressure or suction side of a vane and extending between a leading edge and a trailing edge of the vane of one of the stages, passing the routed compressor air through the vane, and emitting the routed compressor air into a space between the vane and a blade of at least one of an upstream stage and downstream stage of the compressor, wherein the routed compressor air is drawn from a mid-span area of the vane.

12. The method of claim 11 , wherein the space is upstream of a seal for the vane.

13. The method of claim 12 , wherein the space is radially inward of the blade.

14. The method of claim 13 , wherein the space is between the seal and a post mounting the blade.

15. A vane assembly for a compressor of a gas turbine engine comprising:
a vane having a pressure side and a suction side and extending axially between a leading edge and a trailing edge and a span extending from a root to a tip;
a seal located on the root; and
a cooling air circuit passing through the vane and having an inlet on one of the pressure or suction side of the vane and an outlet at a rotor, with outlet located at least one of upstream or downstream of the seal and wherein the inlet extends between the leading edge and the trailing edge of the vane, wherein the inlet is located on a mid-span area of the vane.

16. The vane assembly of claim 15 , wherein the inlet is located on the pressure side of the vane.

17. A method of cooling a multi-stage compressor of a gas turbine engine, the method comprising reducing an operating air temperature in a space between a seal and a blade post of adjacent stages below a creep temperature of the blade post by routing compressor air through an inlet located on one of a pressure or suction side of a vane and extending between a leading edge and a trailing edge of the vane, passing the routed compressor air through the vane, and emitting the routed compressor air into the space, which is upstream of the vane, wherein the routed compressor air is drawn from a mid-span area of the vane.

18. The method of claim 17 , wherein the temperature is reduced at least 50 degrees Fahrenheit.

19. The method of claim 17 , wherein the routed compressor air is drawn from the pressure side of the vane.

20. A method of cooling a multi-stage compressor of a gas turbine engine, the method comprising reducing an operating air temperature in a space between a seal and a blade post of adjacent stages at least 50 degrees Fahrenheit by routing compressor air through an inlet located along a pressure side of a vane and extending between a leading edge and a trailing edge of a vane as compared to without the cooling by routing compressor air through an inlet in the vane, passing the routed compressor air through the vane, and emitting the routed compressor into the space, which is upstream of the vane, wherein the routed compressor air is drawn from a mid-span area of the vane.

21. The method of claim 20 , wherein the inlet located along a pressure sid

Assignees

• Gen Electric

Inventors

• Subramanian Sesha
• Shelton Monty Lee
• Proctor Robert
• Thoppil Mohamed Musthafa
• Gurram Venkateswararao
• Joiner Josiah Warren
• Sak Wojciech

Classifications

• F01D25/12

  Cooling · CPC title

• F01D25/00

  Component parts, details, or accessories, not provided for in, or of interest apart from, other groups · CPC title

• F01D9/041

  using blades (F01D5/148 takes precedence) · CPC title

• F01D5/187 Primary

  Convection cooling · CPC title

• F04D29/584

  cooling or heating the machine (F04D29/5846, F04D29/5853 take precedence) · CPC title