Compressor areas for high overall pressure ratio gas turbine engine

Patent metadata

Field	Value
Publication number	US-9897001-B2
Application number	US-201514620395-A
Country	US
Kind code	B2
Filing date	Feb 12, 2015
Priority date	Mar 4, 2014
Publication date	Feb 20, 2018
Grant date	Feb 20, 2018

Abstract

A gas turbine engine comprises a high pressure turbine rotor, an intermediate pressure turbine rotor and a fan drive turbine rotor. The fan drive turbine rotor drives a fan rotor through a gear reduction. The intermediate pressure rotor drives a low pressure compressor rotor and the high pressure turbine rotor drives a high pressure compressor rotor. A first flow cross-sectional area is between an outer periphery of a hub in the low pressure compressor rotor, and an outer tip of an upstream most blade row of the low pressure compressor rotor. A second flow cross-sectional area is between an outer periphery of a hub in the high pressure compressor rotor, and an outer tip of an upstream most blade row of the high pressure compressor rotor. A ratio of the first and second flow cross-sectional areas is greater than or equal to about 0.12 and less than or equal to about 0.33.

First claim

The invention claimed is:

1. A gas turbine engine comprising:
a high pressure turbine rotor, an intermediate pressure turbine rotor and a fan drive turbine rotor, said fan drive turbine rotor driving a fan rotor through a gear reduction, said intermediate pressure rotor driving a low pressure compressor rotor and said high pressure turbine rotor driving a high pressure compressor rotor,
a first flow cross-sectional area between an outer periphery of a hub in said low pressure compressor rotor, and an outer tip of an upstream most blade row of said low pressure compressor rotor; a second flow cross-sectional area between an outer periphery of a hub in said high pressure compressor rotor, and an outer tip of an upstream most blade row of said high pressure compressor rotor; and
a ratio of said second flow cross-sectional area to said first flow cross-sectional area being greater than or equal to 0.12 and less than or equal to 0.33.

2. The gas turbine engine as set forth in claim 1 , wherein a fan flow cross-sectional area is defined between an outer tip of fan blades and an outer periphery of a fan hub, and a ratio of said fan flow cross-sectional area to said first flow cross-sectional area being greater than or equal to 11 and less than or equal to 20.

3. The gas turbine engine as set forth in claim 2 , wherein said high pressure compressor having a downstream most vane row and an exit cross-sectional area defined between said outer periphery of said hub, and an inner periphery of a housing surrounding said high pressure compressor rotor at said downstream most vane row, and a ratio of said exit cross-sectional area to said second cross-sectional flow area being greater than or equal to 0.25 and less than or equal to 0.6.

4. The gas turbine engine as set forth in claim 3 , wherein a ratio of said exit cross-sectional area to said first flow cross-sectional area being greater than or equal to 0.0667 and less than or equal to 0.125.

5. The gas turbine engine as set forth in claim 4 , wherein a gear ratio of said gear reduction being greater than or equal to 2.6.

6. The gas turbine engine as set forth in claim 5 , wherein said low pressure compressor rotor having 6 to 14 stages.

7. The gas turbine engine as set forth in claim 6 , wherein said high pressure compressor rotor having 4 to 7 stages.

8. The gas turbine engine as set forth in claim 7 , wherein a total stage count across the high and low pressure compressor rotors being 10 to 21.

9. The gas turbine engine as set forth in claim 1 , wherein said high pressure compressor having a downstream most vane row and an exit cross-sectional area defined between said outer periphery of said hub, and an inner periphery of a housing surrounding said high pressure compressor rotor at said downstream most vane row, and a ratio of said exit cross-sectional area to said second cross-sectional area being greater than or equal to 0.25 and less than or equal to 0.6.

10. The gas turbine engine as set forth in claim 9 , wherein a ratio of said exit cross-sectional area to said first flow cross-sectional area being greater than or equal to 0.0667 and less than or equal to 0.125.

11. The gas turbine engine as set forth in claim 10 , wherein a gear ratio of said gear reduction being greater than or equal to 2.6.

12. The gas turbine engine as set forth in claim 11 , wherein said low pressure compressor rotor having 6 to 14 stages.

13. The gas turbine engine as set forth in claim 12 , wherein said high pressure compressor rotor having 4 to 7 stages.

14. The gas turbine engine as set forth in claim 13 , wherein a total stage count across the high and low pressure compressor rotors being 10 to 21.

15. The gas turbine engine as set forth in claim 1 , wherein said high pressure compressor having a downstream most vane row and an exit cross-sectional area defined between said outer periphery of said hub, and an inner periphery of a housing surrounding said high pressure compressor rotor at said downstream most vane row, and a ratio of said exit cross-sectional area to said first flow cross-sectional area being greater than or equal to 0.0667 and less than or equal to 0.125.

16. The gas turbine engine as set forth in claim 1 , wherein a gear ratio of said gear reduction being greater than or equal to 2.6.

17. The gas turbine engine as set forth in claim 1 , wherein said low pressure compressor rotor having 6 to 14 stages.

18. The gas turbine engine as set forth in claim 1 , wherein said high pressure compressor rotor having 4 to 7 stages.

19. The gas turbine engine as set forth in claim 18 , wherein a total stage count across the high and low pressure compressor rotors being 10 to 21.

20. The gas turbine engine as set forth in claim 1 , wherein a total stage count across the high and low pressure compressor rotors being 10 to 21.

Assignees

• United Technologies Corp

Inventors

• Schwarz Frederick M
• Pixton Stephen G

Classifications

• F02C3/067

  having counter-rotating rotors (F02C3/073 takes precedence) · CPC title

• F05D2250/44

  by counter rotation · 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

• F02C3/107 Primary

  with two or more rotors connected by power transmission · CPC title

• F05D2260/4031

  as in toothed gearing · CPC title