Radial cooling system for gas turbine engine compressors

Patent metadata

Field	Value
Publication number	US-10767485-B2
Application number	US-201815864925-A
Country	US
Kind code	B2
Filing date	Jan 8, 2018
Priority date	Jan 8, 2018
Publication date	Sep 8, 2020
Grant date	Sep 8, 2020

Abstract

A gas turbine engine is disclosed. The gas turbine engine includes a first rotor supporting a first plurality of circumferentially spaced rotor blades and a second rotor disposed axially downstream of the first rotor and supporting a second plurality of circumferentially spaced rotor blades, a first bore cavity between the first rotor and the second rotor, a first fluid passageway configured to provide cooled air to the first bore cavity and a first anti-vortex component positioned proximate the first bore cavity and configured to increase pressure of the cooled air as the cooled air traverses radially outward from the first bore cavity.

First claim

What is claimed:

1. A gas turbine engine, comprising:
a first rotor supporting a first plurality of circumferentially spaced rotor blades and a second rotor disposed axially downstream of the first rotor and supporting a second plurality of circumferentially spaced rotor blades;
a first bore cavity between the first rotor and the second rotor;
a second bore cavity between the second rotor and an aft hub;
a fluid passageway configured to provide a cooled air to the first bore cavity and to the second bore cavity;
a first anti-vortex component positioned within the first bore cavity and configured to increase pressure of the cooled air as the cooled air traverses radially outward from the first bore cavity;
a second anti-vortex component positioned within the second bore cavity and configured to increase pressure of the cooled air as the cooled air traverses radially outward from the second bore cavity; and
an orifice through the aft hub configured to introduce an aft cooled air flow path from the second bore cavity onto a downstream rim portion of the second rotor,
wherein the downstream rim portion of the second rotor is disposed axially upstream and adjacent a plurality of exit guide vanes.

2. The gas turbine engine of claim 1 , further comprising a vane inner shroud disposed radially outward of the first bore cavity and axially intermediate the first rotor and the second rotor, the vane inner shroud supporting a plurality of circumferentially spaced vanes on a radially outer surface and having a sealing surface on a radially inner surface, the first plurality of circumferentially spaced rotor blades, the second plurality of circumferentially spaced rotor blades and the plurality of circumferentially spaced vanes defining a portion of a core flow path.

3. The gas turbine engine of claim 2 , further comprising a seal disposed between the radially inner surface of the vane inner shroud and the first bore cavity, the seal configured to obstruct passage of air from the core flow path from an axially downstream position of the plurality of circumferentially spaced vanes to an axially upstream position of the plurality of circumferentially spaced vanes via a seal flow space between the seal and the radially inner surface of the vane inner shroud.

4. The gas turbine engine of claim 3 , further comprising a first gap disposed downstream of the plurality of circumferentially spaced vanes that provides an inlet passage for air from the core flow path to flow between the seal and the radially inner surface of the vane inner shroud.

5. The gas turbine engine of claim 4 , further comprising a second gap disposed upstream of the plurality of circumferentially spaced vanes that provides an exit passage for air between the seal and the radially inner surface of the vane inner shroud to return to the core flow path.

6. The gas turbine engine of claim 5 , further comprising a seal inlet configured to introduce a cooled air flow path from the first bore cavity into the seal flow space.

7. The gas turbine engine of claim 6 , wherein the seal comprises a labyrinth seal having one or more knife edges.

8. The gas turbine engine of claim 6 , wherein the first anti-vortex component comprises a first anti-vortex tube.

9. The gas turbine engine of claim 6 , wherein the fluid passageway includes a removal orifice extending into the core flow path at a removal location.

10. The gas turbine engine of claim 9 , wherein the removal location is positioned upstream of the first rotor.

11. The gas turbine engine of claim 10 , further comprising a heat exchanger positioned within the fluid passageway intermediate the removal orifice and an inlet orifice positioned between the first rotor and the second rotor.

12. The gas turbine engine of claim 11 , wherein the first anti-vortex component comprises a first anti-vortex tube and the second anti-vortex component comprises a second anti-vortex tube.

13. The gas turbine engine of claim 12 , wherein the first anti-vortex tube has a first length and the second anti-vortex tube has a second length greater than the first length.

14. A compressor for a gas turbine engine, comprising:
a first rotor supporting a first plurality of circumferentially spaced rotor blades, a second rotor disposed axially downstream of the first rotor and supporting a second plurality of circumferentially spaced rotor blades and a third rotor disposed axially downstream of the second rotor and supporting a third plurality of circumferentially spaced rotor blades;
a first bore cavity between the first rotor and the second rotor, a second bore cavity between the second rotor and the third rotor and a third bore cavity between the third rotor and an aft hub;
a fluid passageway configured to provide cooled air to the first bore cavity, the second bore cavity and the third bore cavity;
a first anti-vortex component positioned within the first bore cavity and configured to increase pressure of the cooled air as the cooled air traverses radially outward from the first bore cavity;
a second anti-vortex component positioned within the second bore cavity and configured to increase pressure of the cooled air as the cooled air traverses radially outward from the second bore cavity; and
a third anti-vortex component positioned within the third bore cavity and configured to increase pressure of the cooled air as the cooled air traverses radially outward from the third bore cavity; and
an orifice through the aft hub configured to introduce an aft cooled air flow path from the third bore cavity onto a downstream rim portion of the third rotor,
wherein the downstream rim portion of the third rotor is disposed axially upstream and adjacent a plurality of exit guide vanes.

15. The compressor of claim 14 , further comprising:
a first vane inner shroud disposed radially outward of the first bore cavity and axially intermediate the first rotor and the second rotor, the first vane inner shroud supporting a first plurality of circumferentially spaced vanes on a first radially outer surface and having a first sealing surface on a first radially inner surface,
a second vane inner shroud disposed radially outward of the second bore cavity and axially intermediate the second rotor and the third rotor, the second vane inner shroud supporting a second plurality of circumferentially spaced vanes on a second radially outer surface and having a second sealing surface on a second radially inner surface,
the first plurality of circumferentially spaced rotor blades, the second plurality of circumferentially spaced rotor blades and the third plurality of circumferentially spaced rotor blades, and the first plurality of circumferentially spaced vanes and the second plurality of circumferentially spaced vanes defining a portion of a core flow path.

16. The compressor of claim 15 , further comprising:
a first seal disposed between the first radially inner surface of the first vane inner shroud and the first bore cavity, the first seal configured to obstruct passage of air from the core flow path from an axially downstream position of the first plurality of circumferentially spaced vanes to an axially upstream position of the first plurality of circumferentially spaced vanes via a first seal flow space between the first seal and the first radially inner surface of the first vane inner shroud and
a second seal disposed between the second radially inner surface of the second vane inner shroud and the second bore cavity, the second seal configured to obstruct passage of air from the core flow path from an axially downstream position of the se

Assignees

• United Technologies Corp
• Raytheon Tech Corp

Inventors

• Bintz Matthew E
• Dibenedetto Enzo
• Coderre Paul E

Classifications

• F01D11/02

  by non-contact sealings, e.g. of labyrinth type (for sealing space between rotor blade tips and stator F01D11/08) · CPC title

• F05D2220/3219

  for the last stage of a compressor or a high pressure compressor · CPC title

• F05D2260/209

  using vortex tubes · CPC title

• F05D2260/213

  by the provision of a heat exchanger within the cooling circuit · CPC title

• F01D11/001

  for sealing space between stator blade and rotor · CPC title