Hybrid compressor bleed air for aircraft use

1. A gas turbine engine and an air use destination on a fuselage of an associated aircraft 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;

a high tap to tap air downstream of said first stage of said higher pressure compressor rotor;

an air handling system and an air use destination, wherein said air handling system selectively communicates both said low tap and said high tap to said air use destination, wherein said air use destination includes a fuselage of an associated aircraft;

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; and

at least one valve controls the selective supply of the high tap to the air handling system;

wherein said air use destination includes pressurizing a fuselage of an associated aircraft; and

said low power operation including at least cruise and idle, and said high power operation including at least take-off of the gas turbine engine.

2. The gas turbine engine and an air use destination on a fuselage of an associated aircraft as set forth in claim 1 , wherein said high tap taps air from an intermediate location in said higher pressure compressor rotor.

3. The gas turbine engine and an air use destination on a fuselage of an associated aircraft as set forth in claim 2 , wherein said high tap also taps air from a location downstream of a downstream most stage of said higher pressure compressor rotor.

4. The gas turbine engine and an air use destination on a fuselage of an associated aircraft as set forth in claim 1 , wherein said high tap taps air from a location downstream of a downstream most stage of said higher pressure compressor rotor.

5. The gas turbine engine and an air use destination on a fuselage of an associated aircraft as set forth in claim 1 , wherein said supply duct passes through a heat exchanger before being delivered to said use.

6. The gas turbine engine and an air use destination on a fuselage of an associated aircraft as set forth in claim 5 , wherein a fan delivers air into a bypass duct and to said lower pressure compressor rotor, and air from said bypass duct cooling said heat exchanger.

7. The gas turbine engine and an air use destination on a fuselage of an associated aircraft as set forth in claim 1 , wherein said lower pressure compressor rotor has at least four stages.

8. The gas turbine engine and an air use destination on a fuselage of an associated aircraft as set forth in claim 7 , wherein said lower pressure compressor rotor has four or five stages.

9. The gas turbine engine and an air use destination on a fuselage of an associated aircraft as set forth in claim 8 , wherein said higher pressure compressor rotor has between 6 and 13 stages.

10. The gas turbine engine and an air use destination on a fuselage of an associated aircraft as set forth in claim 9 , wherein said use includes an air conditioning pack for conditioning air for use in a cabin of an associated aircraft.

11. The gas turbine engine and an air use destination on a fuselage of an associated aircraft as set forth in claim 1 , wherein said higher pressure compressor rotor has between 6 and 13 stages.

12. The gas turbine engine and an air use destination on a fuselage of an associated aircraft as set forth in claim 1 , 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 and an air use destination on a fuselage of an associated aircraft 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.

14. The gas turbine engine and an air use destination on a fuselage of an associated aircraft as set forth in claim 13 , wherein said bypass ratio is greater than or equal to 12.0.

15. The gas turbine engine and an air use destination on a fuselage of an associated aircraft as set forth in claim 1 , wherein a fan rotor is driven through a gear reduction by a fan drive turbine.

16. The gas turbine engine and an air use destination on a fuselage of an associated aircraft as set forth in claim 15 , wherein said fan drive turbine further driving said lower pressure compressor rotor.

17. The gas turbine engine and an air use destination on a fuselage of an associated aircraft as set forth in claim 16 , wherein a gear ratio of said gear reduction is greater than or equal to 2.6.

18. The gas turbine engine and an air use destination on a fuselage of an associated aircraft as set forth in claim 17 , wherein said use includes an air conditioning pack for conditioning air for use in a cabin of an associated aircraft.

19. The gas turbine engine and an air use destination on a fuselage of an associated aircraft 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.