Gas turbine engine with rotor bore heating

Patent metadata

Field	Value
Publication number	US-10094296-B2
Application number	US-201414761073-A
Country	US
Kind code	B2
Filing date	Feb 11, 2014
Priority date	Feb 19, 2013
Publication date	Oct 9, 2018
Grant date	Oct 9, 2018

Abstract

A gas turbine engine has a compressor rotor with blades and a disk. A bore is defined radially inwardly of the disk. A combustor includes a burner nozzle. A tap taps air that has been combusted in the combustor section through a valve, and into the bore of the disk. A method is also disclosed.

First claim

The invention claimed is:

1. A gas turbine engine comprising:
a compressor rotor including blades and a disk, with a bore defined radially inwardly of said disk;
a combustor, including a burner nozzle;
a tap for tapping air that has been combusted in the combustor through a valve, and into said bore of said disk;
said valve is controlled to only be opened under certain periods of operation of the gas turbine engine; and
a control for said valve is programmed to be open when an aircraft associated with the gas turbine engine is at idle and to be closed at take-off.

2. The gas turbine engine as set forth in claim 1 , wherein a second valve is provided to provide for redundancy.

3. The gas turbine engine as set forth in claim 1 , wherein a supplemental compressor is provided on a flow path from said tap to said bore of said disk.

4. The gas turbine engine as set forth in claim 1 , wherein the tapped air is also delivered to a bore of a turbine section.

5. The gas turbine engine as set forth in claim 4 , wherein a labyrinth seal prevents leakage of the tapped air into a diffusor chamber in the turbine section.

6. The gas turbine engine as set forth in claim 1 , wherein a location of the tap is within the combustor.

7. The gas turbine engine as set forth in claim 1 , wherein said tapped air heats said disk to a temperature that exceeds a temperature of said blades while the gas turbine engine is at an idle point of operation.

8. A gas turbine engine comprising:
a compressor rotor including blades and a disk, with a bore defined radially inwardly of said disk;
a combustor, including a burner nozzle;
a tap for tapping air that has been combusted in the combustor through a valve, and into said bore of said disk; and
said tapped air is jacketed in a jacket of compressed air.

9. The gas turbine engine as set forth in claim 8 , wherein said valve is controlled to only be opened under certain periods of operation of the gas turbine engine.

10. The gas turbine engine as set forth in claim 9 , wherein a control for said valve is programmed to be open when an aircraft associated with the gas turbine engine is at idle and to be closed at take-off.

11. A method of operating a gas turbine engine comprising:
tapping air that has been combusted in a combustor section through a valve, and into a bore of a disk of a compressor rotor;
said valve is opened only under certain periods of operation of the gas turbine engine; and
said valve is opened when an aircraft associated with the gas turbine engine is at idle and is closed at take-off.

12. The method as set forth in claim 11 , wherein a second valve is provided for redundancy.

13. The method as set forth in claim 11 , wherein a supplemental compressor is provided on a flow path from a tap for tapping the air to said bore of said disk.

14. The method as set forth in claim 11 , wherein the tapped air is also delivered to a bore of a turbine section.

15. The method as set forth in claim 13 , wherein a labyrinth seal prevents leakage of the tapped air toward a vane in the turbine section.

16. The method as set forth in claim 11 , wherein a location of a tap is for tapping the air within the combustor.

17. The method as set forth in claim 11 , wherein said tapped air is jacketed in a jacket of compressed air, on a path for delivery into said bore of said disk.

18. The method as set forth in claim 11 , wherein said tapped air heats said disk to a temperature that exceeds a temperature of said blades while the gas turbine engine is at an idle point of operation.

Assignees

• United Technologies Corp

Inventors

• Schwarz Frederick M
• Duesler Paul W

Classifications

• F05D2260/232

  characterized by the cooling medium · CPC title

• Y02T50/676

  Cross-Sectional Technologies · mapped topic

• F02C3/06

  the compressor comprising only axial stages (F02C3/10 takes precedence) · CPC title

• F01D5/08 Primary

  Heating, heat-insulating or cooling means {(specially adapted for radial flow machines or engines F01D5/04)} · 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