Hybrid compressor bleed air for aircraft use

Patent metadata

Field	Value
Publication number	US-2016009399-A1
Application number	US-201514737580-A
Country	US
Kind code	A1
Filing date	Jun 12, 2015
Priority date	Jul 9, 2014
Publication date	Jan 14, 2016
Grant date	—

Abstract

A gas turbine engine comprises at least two compressor rotors, including a first lower pressure compressor rotor and a second higher pressure compressor rotor. At least two corresponding air taps include a low tap for tapping low pressure compressor air from a location downstream of a first stage of the lower pressure compressor rotor, and upstream of a first stage of the higher pressure compressor rotor, and a high tap to tap air downstream of the first stage of the higher pressure compressor rotor. an air handling system selectively communicates both the low tap and the high tap to an air use destination. Air is selectively supplied from the low tap to the air handling system at a high power operation and from the high tap to the air handling system at a low power operation.

First claim

1 . A gas turbine engine comprising:
at least two compressor rotors, including a first lower pressure compressor rotor and a second higher pressure compressor rotor; at least two corresponding air taps, including: a low tap for tapping low pressure compressor air from a location downstream of a first stage of the lower pressure compressor rotor, and upstream of a first stage of the higher pressure compressor rotor; and a high tap to tap air downstream of said first stage of said higher pressure compressor rotor; an air handling system to selectively communicate both said low tap and said high tap to an air use destination; and
wherein air is selectively supplied from the low tap to the air handling system at a high power operation and from said high tap to the air handling system at a low power operation.

2 . The gas turbine engine as set forth in claim 1 , wherein at least one valve controls the selective supply of the low tap and high tap to the air handling system.

3 . The gas turbine engine as set forth in claim 2 , wherein said high tap taps air from an intermediate location in said higher pressure compressor rotor.

4 . The gas turbine engine as set forth in claim 3 , wherein said high tap also taps air from a location downstream of a downstream most stage of said higher pressure compressor rotor.

5 . The gas turbine engine as set forth in claim 2 , wherein said high tap taps air from a location downstream of a downstream most stage of said higher pressure compressor rotor.

6 . The gas turbine engine as set forth in claim 2 , wherein said supply duct passes through a heat exchanger before being delivered to said use.

7 . The gas turbine engine as set forth in claim 6 , wherein a fan delivers air into a bypass duct and to said lower pressure compressor rotor, and air from said bypass duct cooling said air heat exchanger.

8 . The gas turbine engine as set forth in claim 2 , wherein said lower pressure compressor rotor has at least four stages.

9 . The gas turbine engine as set forth in claim 8 , wherein said lower pressure compressor rotor has four or five stages.

10 . The gas turbine engine as set forth in claim 9 , wherein said higher pressure compressor rotor has between 6 and 13 stages.

11 . The gas turbine engine as set forth in claim 10 , wherein said use includes pressurizing a fuselage of an associated aircraft.

12 . The gas turbine engine as set forth in claim 11 , wherein said use includes an air conditioning pack for conditioning air for use in a cabin of an associated aircraft.

13 . The gas turbine engine as set forth in claim 2 , wherein said higher pressure compressor rotor has between 6 and 13 stages.

14 . The gas turbine engine as set forth in claim 2 , wherein said use includes pressurizing a fuselage of an associated aircraft.

15 . The gas turbine engine as set forth in claim 14 , wherein said use includes an air conditioning pack for conditioning air for use in a cabin of an associated aircraft.

16 . The gas turbine engine as set forth in claim 21 , wherein said use includes an air conditioning pack for conditioning air for use in a cabin of an associated aircraft.

17 . The gas turbine engine as set forth in claim 1 , wherein a bypass ratio is defined by the volume of air delivered into said bypass duct, compared to the volume of air delivered into said lower pressure compressor rotor and wherein said bypass ratio is greater than or equal to 10.

18 . The gas turbine engine as set forth in claim 17 , wherein said bypass ratio is greater than or equal to 12.0.

19 . The gas turbine engine as set forth in claim 1 , wherein a fan rotor is driven through a gear reduction by a fan drive turbine.

20 . The gas turbine engine as set forth in claim 19 , wherein said fan drive turbine further driving said lower pressure compressor rotor.

21 . The gas turbine engine as set forth in claim 20 , wherein a gear ratio of said gear reduction is greater than or equal to 2.6.

22 . The gas turbine engine as set forth in claim 1 , wherein said high tap includes at least two different tap locations, and the two different tap locations are selectively utilized at different operational conditions.

Assignees

• United Technologies Corp

Inventors

• Schwarz Frederick M
• Pixton Stephen G

Classifications

• F05D2220/3217

  for the first stage of a compressor or a low pressure compressor · CPC title

• F05D2260/213

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

• F01D5/06

  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 title

• F05D2220/32

  in gas turbines · CPC title

• F02C6/08

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