Turbofan with bleed supercharged auxiliary engine

1. An aircraft gas turbine engine system comprising:

a first gas turbine engine comprising:

a first spool with a first compressor section;

a second spool with a second compressor section downstream of and rotationally independent from the first compressor section;

a second gas turbine engine configured to provide power to at least one of the first and second spools of the first gas turbine engine;

an inter-engine gas path disposed to receive gas flow bled from a bleed location in the first gas turbine engine downstream of the first compressor section, and to supply this gas flow to an inlet of the second gas turbine engine; and

a switch valve disposed at the inlet of the second gas turbine engine, and toggleable between two states:

an inter-engine state fluidly connecting the inter-engine gas path to an inlet of the second gas turbine engine; and

an external state fluidly disconnecting the inter-engine gas path from the inlet of the second gas turbine engine, and opening the inlet of the second gas turbine engine to an ambient air source.

2. The aircraft gas turbine engine system of claim 1 , further comprising:

a generator connected to the second gas turbine engine, and disposed to be powered by the second gas turbine engine; and

a motor disposed to boost at least one of the first and second spools, and powered by the generator,

wherein providing power to at least one of the first and second spools of the first gas turbine engine comprises powering the generator to drive the motor.

3. The aircraft gas turbine engine system of claim 1 , wherein the first gas turbine engine is a propulsion engine, and the second gas turbine engine is an auxiliary power unit (APU).

4. The aircraft gas turbine engine system of claim 1 , further comprising an intercooler disposed along the inter-engine gas path so as to cool the gas flow, thereby supplying the gas flow to the inlet of the second gas turbine engine at a lower temperature than bled from the bleed location.

5. The aircraft gas turbine engine system of claim 1 , wherein the bleed location is situated between the first compressor section and the second compressor section.

6. The aircraft gas turbine engine system of claim 1 , wherein the bleed location is situated in the second compressor section.

7. The aircraft gas turbine engine system of claim 1 , wherein the bleed location extends downstream from between the first compressor section and the second compressor section, and encompasses a plurality of bleed outlets from compressor regions of the first gas turbine engine.

8. The gas turbine engine system of claim 1 , further comprising a linkage coupling the second gas turbine engine to the first spool of the first gas turbine engine.

9. The gas turbine engine system of claim 1 , wherein power provided to the first spool from the second gas turbine engine would create an overpressure at the intersection of the first and second compressor sections, but for the gas flow bled from the bleed location.

10. The gas turbine engine system of claim 9 , wherein power provided to the first spool from the second gas turbine engine exceeds power provided to the second spool from the second gas turbine engine.

11. The gas turbine engine system of claim 1 , wherein the gas flow bled from the bleed location is metered.

12. The aircraft gas turbine engine system of claim 1 , further comprising:

a third gas turbine engine comprising:

a third spool with a third compressor section;

a fourth spool with a fourth compressor section downstream of and rotationally independent from the third compressor section; and

a secondary inter-engine gas path disposed to receive gas flow bleed from a third engine bleed location in the third gas turbine engine downstream of the third compressor section, and to supply this gas flow to the inlet of the second gas turbine engine,

wherein the second gas turbine engine additionally provides power to at least one of the third or fourth spools.

13. The aircraft gas turbine engine system of claim 12 , further comprising a mixing valve disposed upstream of the inlet of the second gas turbine engine, to receive fluid from both the inter-engine gas path and the secondary inter-engine gas path.

14. The aircraft gas turbine engine system of claim 13 , further comprising an intercooler disposed between the mixing valve and the inlet of the second engine.

15. A method of operating a system of gas turbine engines including a first engine and a second engine, the method comprising:

boosting a first spool of the first engine using the second engine, thereby producing an overpressure between a first compressor on the first spool and an adjacent second compressor on a second spool of the first engine rotationally independent from and axially downstream of the first compressor;

bleeding a pressurized inter-engine airflow from the first engine, downstream of the first compressor;

supplying the pressurized inter-engine airflow to an intake of the second engine, thereby supercharging the second engine;

interrupting the supply of pressurized inter-engine airflow to the intake of the second engine; and

supplying an ambient airflow to the intake of the second engine after interrupting the supply of pressurized inter-engine airflow.

16. The method of claim 15 , wherein the bleed from the first engine alleviates the overpressure sufficiently to offset the possibility of stall due to the boosting of the first spool.

17. The method of claim 16 , further comprising cooling the pressurized inter-engine airflow before supplying the pressurized inter-engine airflow to the intake of the second engine.

18. The method of claim 16 , wherein boosting the first spool of the first engine comprises:

driving an electrical generator with the second engine;

powering a motor from the electrical generator; and

driving the first spool using the motor.

19. An aircraft gas turbine engine system comprising:

a first gas turbine engine comprising:

a first spool with a first compressor section;

a second spool with a second compressor section downstream of and rotationally independent from the first compressor section;

a second gas turbine engine configured to provide power to at least one of the first and second spools of the first gas turbine engine;

an inter-engine gas path disposed to receive gas flow bled from a bleed location in the first gas turbine engine downstream of the first compressor section, and to supply this gas flow to an inlet of the second gas turbine engine;

a third gas turbine engine comprising:

a third spool with a third compressor section;

a fourth spool with a fourth compressor section downstream of and rotationally independent from the third compressor section; and

a secondary inter-engine gas path disposed to receive gas flow bleed from a third engine bleed location in the third gas turbine engine downstream of the third compressor section, and to supply this gas flow to the inlet of the second gas turbine engine,

wherein the second gas turbine engine additionally provides power to at least one of the third or fourth spools.