Multi-source turbine cooling air

Patent metadata

Field	Value
Publication number	US-10823071-B2
Application number	US-201816158581-A
Country	US
Kind code	B2
Filing date	Oct 12, 2018
Priority date	Dec 10, 2015
Publication date	Nov 3, 2020
Grant date	Nov 3, 2020

Abstract

A gas turbine engine comprises a compressor section and a turbine section, with the turbine section having a first stage blade row and a downstream blade row. A higher pressure tap is tapped from a higher pressure first location in the compressor. A lower pressure tap is tapped from a lower pressure location in the compressor which is at a lower pressure than the first location. The higher pressure tap passes through a heat exchanger, and then is delivered to cool the first stage blade row in the turbine section. The lower pressure tap is delivered to at least partially cool the downstream blade row.

First claim

The invention claimed is:

1. A gas turbine engine comprising:
a compressor section and a turbine section, with said turbine section having a first stage blade row and a downstream blade row;
a higher pressure tap tapped from a higher pressure first location in said compressor;
a lower pressure tap tapped from a lower pressure location in said compressor which is at a lower pressure than said first location, said higher pressure tap passing through a heat exchanger, and then being delivered to cool said first stage blade row in said turbine section, and said lower pressure tap being delivered to at least partially cool said downstream blade row; and
wherein said downstream stage blade row is a second stage, and said first stage blade row and said second stage rotating together as a single rotor.

2. The gas turbine engine as set forth in claim 1 , wherein radially outwardly extending air from said higher pressure tap also cooling a vane mounted intermediate said first stage blade row and said downstream blade row.

3. The gas turbine engine as set forth in claim 2 , wherein said radially outwardly extending air from said higher pressure tap also cooling an upstream end of said downstream blade row.

4. The gas turbine engine as set forth in claim 3 , wherein said lower pressure tap passing radially inwardly of said first stage blade row, and axially beyond said downstream blade row and then radially outwardly to cool a downstream end of said downstream stage blade row.

5. The gas turbine engine as set forth in claim 4 , wherein said downstream stage blade row is a second stage, and said first stage blade row and said second stage rotating together as a single rotor.

6. A gas turbine engine comprising:
a compressor section and a turbine section, with said turbine section having a first stage blade row and a downstream blade row;
a higher pressure tap tapped from a higher pressure first location in said compressor; and a lower pressure tap tapped from a lower pressure location in said compressor which is at a lower pressure than said first location, said higher pressure tap passing through a heat exchanger, and then being delivered to cool said first stage blade row in said turbine section, and said lower pressure tap being delivered to at least partially cool said downstream blade row;
wherein a fan drive turbine rotor is positioned downstream of a turbine rotor including said first stage blade row and said downstream blade row, with said fan drive turbine driving said fan through a gear reduction;
wherein a gear ratio of said gear reduction is greater than or equal to about 2.3:1.

7. The gas turbine engine as set forth in claim 1 , wherein a fan is positioned upstream of said compressor section and said fan delivering air into a bypass duct as propulsion air, and into said compressor section with a bypass ratio defined as the volume ratio of air delivered into said bypass duct compared to the air delivered into said compressor, with said bypass ratio being greater than or equal to about 6.0.

8. The gas turbine engine as set forth in claim 7 , wherein said bypass ratio is greater than or equal to about 10.0.

9. The gas turbine engine as set forth in claim 6 , wherein said downstream stage blade row is a second stage, and said first stage blade row and said second stage rotating together as a single rotor.

10. The gas turbine engine as set forth in claim 1 , wherein said lower pressure tap passing radially inwardly of said first stage blade row, and axially beyond said downstream blade row and then radially outwardly to cool a downstream end of said downstream stage blade row.

11. The gas turbine engine as set forth in claim 9 , wherein a fan is positioned upstream of said compressor section and said fan delivering air into a bypass duct as propulsion air, and into said compressor section with a bypass ratio defined as the volume ratio of air delivered into said bypass duct compared to the air delivered into said compressor, with said bypass ratio being greater than or equal to about 6.0.

12. The gas turbine engine as set forth in claim 11 , wherein said higher pressure tap passing from said heat exchanger toward said turbine section, and split into a first path heading radially outwardly to cool an upstream end of said first stage blade row, and a second path moving radially inwardly of a hub mounting said first stage blade row and then moving radially outwardly to cool a downstream end of said first stage blade row.

13. The gas turbine engine as set forth in claim 12 , wherein radially outwardly extending air from said higher pressure tap also cooling a vane mounted intermediate said first stage blade row and said downstream blade row.

14. The gas turbine engine as set forth in claim 13 , wherein said radially outwardly extending air from said higher pressure tap also cooling an upstream end of said downstream blade row.

15. The gas turbine engine as set forth in claim 14 , wherein said lower pressure tap passing radially inwardly of said first stage blade row, and axially beyond said downstream blade row and then radially outwardly to cool a downstream end of said downstream stage blade row.

16. The gas turbine engine as set forth in claim 15 , wherein a fan is positioned upstream of said compressor section and said fan delivering air into a bypass duct as propulsion air, and into said compressor section with a bypass ratio defined as the volume ratio of air delivered into said bypass duct compared to the volume of air delivered into said compressor, with said bypass ratio being greater than or equal to about 6.0.

17. The gas turbine engine as set forth in claim 9 , wherein a fan drive turbine rotor is positioned downstream of a turbine rotor including said first stage blade row and said downstream blade row, with said fan drive turbine driving said fan through a gear reduction.

Assignees

• Raytheon Tech Corp

Inventors

• Merry Brian D
• Suciu Gabriel L
• Ackermann William K

Classifications

• F01D5/081 Primary

  Cooling fluid being directed on the side of the rotor disc or at the roots of the blades (F01D5/087 takes precedence) · CPC title

• F02C6/08

  the gas being bled from the gas-turbine compressor · CPC title

• Y02T50/60

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

• F02C7/36

  Power transmission arrangements between the different shafts of the gas turbine plant, or between the gas-turbine plant and the power user ({F02C3/107 - F02C3/13 and} F02C7/32 take precedence; couplings for transmitting rotation F16D; gearing in general F16H) · CPC title

• F02K3/06

  with front fan · CPC title