Gas turbine engine with vane having a cooling inlet

Patent metadata

Field	Value
Publication number	US-11359646-B2
Application number	US-201916555558-A
Country	US
Kind code	B2
Filing date	Aug 29, 2019
Priority date	Nov 16, 2015
Publication date	Jun 14, 2022
Grant date	Jun 14, 2022

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, the compressor comprising:
an outer casing;
a set of inner and outer rings disposed within the outer casing and 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 blades, 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 elongated inlet located on one of the pressure side or suction side of the vanes and an outlet at the rotor upstream of a corresponding seal for the vanes,
wherein the inlet is located along a span of the respective vane in an area of coolest air flow over the vane.

2. The compressor of claim 1 , wherein the inlet is located in a mid-span area of the vane.

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

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

5. The compressor of claim 4 , wherein the inlet comprises a scoop.

6. 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.

7. The compressor of claim 1 , further comprising a first inner ring of the set of inner rings, the first inner ring located within the casing and supporting the vanes of the respective compressor stage at a root of the corresponding vane and the first inner ring defines a circumferential channel forming part of the cooling air circuit.

8. The compressor of claim 7 , wherein the outlet of the cooling air circuit is formed in the ring.

9. The compressor of claim 8 , wherein the seal comprises a honeycomb element mounted to the ring and fingers extending from the rotor and abutting the honeycomb element.

10. The compressor of claim 1 , wherein the rotor comprises posts and the outlet emits the cooling air toward the post upstream of the vane.

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

12. A method of cooling a multi-stage compressor of a gas turbine engine, the method comprising routing compressor air through a vane having a pressure side and a suction side and located in of one of the stages by receiving the compressor air through an elongated inlet located on one of the pressure side or suction side, passing the routed compressor air through the vane, and emitting the routed compressor 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 inlet is located along a span of the vane in an area of coolest air flow over the vane.

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

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

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

16. The method of claim 12 , wherein the routed compressor air is drawn from a mid-span area of the vane.

17. The method of claim 12 , wherein the routed compressor air is drawn from the pressure side of the vane.

18. A vane assembly for a compressor of a gas turbine engine comprising:
a vane comprising a pressure side and a suction side and having 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 elongated inlet located on one of the pressure side or suction side of the vane and an outlet at a rotor, with the outlet located at least one of upstream or downstream of the seal,
wherein the inlet is located along a span of the vane in an area of coolest air flow over the vane.

19. The vane assembly of claim 18 wherein the inlet is located on one of a mid-span area of the vane or the pressure side of the vane.

Assignees

• Gen Electric

Inventors

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

Classifications

• Y02T50/60

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

• F05D2220/3216

  for a special compressor stage · CPC title

• F05D2240/81

  Cooled platforms · CPC title

• F04D29/584

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

• F01D25/12

  Cooling · CPC title