RADIAL COOLING SYSTEM FOR GAS TURBINE ENGINE COMPRESSORS

• US 20190211683A1
• Filed: 01/08/2018
• Published: 07/11/2019
• Est. Priority Date: 01/08/2018
• Status: Active Grant

• Current Assignee: Rtx Corporation
• Original Assignee: United Technologies Corporation (Rtx Corporation)
• Inventors: Bintz, Matthew E., DiBenedetto, Enzo, Coderre, Paul E.
• Granted Patent: US 10,767,485 B2
• Time in Patent Office: Days
• CPC Class Codes: F01D 11/001 for sealing space between stator blade and rotor
• CPC Class Codes: F01D 11/02 by non-contact sealings, e.g. of labyrinth type for sealing space between rotor blade tips and stator F01D11/08
• CPC Class Codes: F01D 5/06 Rotors for more than one axial stage, e.g. of drum or multiple disc type; Details thereof, e.g. shafts, shaft connections F01D5/022, F01D5/023 take precedence
• CPC Class Codes: F01D 5/087 in the radial passages of the rotor disc
• CPC Class Codes: F02C 7/141 of working fluid
• CPC Class Codes: F02C 7/18 the medium being gaseous, e.g. air F02C7/125 takes precedence
• CPC Class Codes: F04D 29/321 for axial flow compressors
• CPC Class Codes: F04D 29/582 specially adapted for elastic fluid pumps
• CPC Class Codes: F05D 2210/12 gaseous, i.e. compressible
• CPC Class Codes: F05D 2220/3218 for an intermediate stage of a compressor
• CPC Class Codes: F05D 2220/3219 for the last stage of a compressor or a high pressure compressor
• CPC Class Codes: F05D 2240/11 Shroud seal segments
• CPC Class Codes: F05D 2240/12 Fluid guiding means, e.g. vanes
• CPC Class Codes: F05D 2240/24 for turbines
• CPC Class Codes: F05D 2240/55 Seals
• CPC Class Codes: F05D 2260/20 Heat transfer, e.g. cooling
• CPC Class Codes: F05D 2260/209 using vortex tubes
• CPC Class Codes: F05D 2260/213 by the provision of a heat exchanger within the cooling circuit

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.

Claims

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 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.

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 first 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 first 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, further comprising a third rotor disposed axially downstream of the second rotor and supporting a third plurality of circumferentially spaced rotor blades, a second bore cavity between the second rotor and the third rotor, a second fluid passageway configured to provide cooled air to the second bore cavity and a second anti-vortex component positioned proximate 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.

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

14. The gas turbine engine of claim 13, 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.

15. 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 and a second bore cavity between the second rotor and the third rotor;

a fluid passageway configured to provide cooled air to the first bore cavity and the second bore cavity;

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; and

a second anti-vortex component positioned proximate 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.

16. The compressor of claim 15, 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.

17. The compressor of claim 16, 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 second plurality of circumferentially spaced vanes to an axially upstream position of the second plurality of circumferentially spaced vanes via a second seal flow space between the second seal and the second radially inner surface of the second vane inner shroud;

a first seal inlet configured to introduce a first cooled air flow path from the first bore cavity into the first seal flow space and

a second seal inlet configured to introduce a second cooled air flow path from the second bore cavity into the second seal flow space.

18. The compressor of claim 17, wherein the first anti-vortex component comprises a first anti-vortex tube and the second anti-vortex component comprises a second anti-vortex tube.

19. The compressor of claim 18, further comprising a third anti-vortex tube positioned within a third bore cavity, in fluid communication with the fluid passageway and positioned downstream of the third rotor, and configured to increase pressure of the cooled air as the cooled air traverses radially outward from the third bore cavity, and a rim inlet configured to introduce a third cooled air flow path from the third bore cavity into a rim portion of the third rotor.

20. The compressor of claim 19, further comprising a heat exchanger positioned within the fluid passageway.