Gas turbine engine turbine impeller pressurization

Patent metadata

Field	Value
Publication number	US-2016003166-A1
Application number	US-201414769591-A
Country	US
Kind code	A1
Filing date	Mar 10, 2014
Priority date	Mar 14, 2013
Publication date	Jan 7, 2016
Grant date	—

Abstract

A cooling system for a gas turbine engine turbine section includes a rotor supporting a blade having a cooling passage. A disc is secured relative to the rotor and it forms a cavity between the rotor and the disc. A bleed air source is in fluid communication with the cavity. An impeller is arranged in the cavity. The impeller is configured to increase a fluid pressure within the cavity to drive bleed air from the bleed air source and thereby provide a pressurized cooling fluid to the cooling passage.

First claim

What is claimed is:

1 . A cooling system for a gas turbine engine turbine section, comprising:
a rotor supporting a blade having a cooling passage; a disc secured relative to the rotor and forming a cavity between the rotor and the disc; a bleed air source in fluid communication with the cavity; and an impeller is arranged in the cavity and configured to increase a fluid pressure within the cavity to drive bleed air from the bleed air source and thereby provide a pressurized cooling fluid to the cooling passage.

2 . The cooling system according to claim 1 , wherein the blade is in a last stage of a high pressure turbine section.

3 . The cooling system according to claim 2 , comprising a spool, the rotor and the disc affixed to the spool for rotation therewith.

4 . The cooling system according to claim 1 , wherein the bleed air source is a stage of a high pressure compressor section.

5 . The cooling system according to claim 4 , wherein the high pressure compressor section includes an aft hub having an aft hub leak path, the aft hub leak path in fluid communication with the cavity and configured to provide aft hub fluid to the cavity.

6 . The cooling system according to claim 1 , comprising a tangential on board injector having a TOBI leak path, the TOBI leak path in fluid communication with the cavity and configured to provide a TOBI fluid to the cavity.

7 . The cooling system according to claim 1 , wherein the impeller is mounted on the disc.

8 . The cooling system according to claim 7 , wherein the impeller includes circumferentially spaced paddles integral with disc.

9 . The cooling system according to claim 1 , comprising static structure, and the disc includes a seal configured to seal relative to the static structure.

10 . A turbine stage for a gas turbine engine, comprising:
a rotor; and a disc secured relative to the rotor to provide a cavity there between, and an impeller arranged in the cavity.

11 . The turbine stage according to claim 10 , wherein the impeller includes a set of first paddles and a set of second paddles, the first and second paddles interleaved relative to one another, the first paddles larger than the second paddles.

12 . The turbine stage according to claim 10 , wherein rotor supports turbine blades having a cooling passage in fluid communication with the cavity, and the disc includes a seal in engagement with turbine blades.

13 . The turbine stage according to claim 12 , wherein the disc includes an annular flange providing the seal, the annular flange extending in an axial direction and spaced radially from the sets of first and second paddles to provide an annular channel radially between the annular flange and the sets of first and second paddles.

14 . A disc for a turbine stage, comprising:
a disc-shaped wall supporting paddles that extend from an inlet radially outward to an outlet, an annular flange extending axially from the wall to provide an annular channel arranged radially between the outlet and the annular wall.

15 . The disc according to claim 14 , wherein the paddles include a set of first paddles and a set of second paddles, the first and second paddles interleaved relative to one another, the first paddles larger than the second paddles.

16 . The disc according to claim 14 , wherein the annular flange includes a first seal.

17 . The disc according to claim 16 , comprising a second seal supported by the wall on a side opposite the paddles.

Assignees

• United Technologies Corp

Inventors

• Mosley John H
• Chrisikos James P
• Gates Roger
• O'Connor John J
• Wu Charles C
• Glasspoole David

Classifications

• F02C7/18

  the medium being gaseous, e.g. air {(F02C7/125 takes precedence)} · CPC title

• F02C9/18 Primary

  by bleeding, bypassing or acting on variable working fluid interconnections between turbines or compressors or their stages {(F02C3/113 takes precedence)} · CPC title

• F01D5/082 Primary

  on the side of the rotor disc · CPC title

• Y02T50/60

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

• F05D2220/3215

  the last stage of the turbine · CPC title