Modulating flow through gas turbine engine cooling system

1. A gas turbine engine system comprising:

an annular fan bypass passage extending between an inner flow surface and an outer flow surface relative to an engine centerline axis;

a cooling passage having an inlet for receiving a bleed flow from the annular fan bypass passage and an outlet at the inner flow surface for discharging the bleed flow into the annular fan bypass passage; and

a nozzle, disposed at an aft end of the annular fan bypass passage, having a variable cross-sectional area for controlling an airflow within the annular fan bypass passage, the nozzle including at least one flap located at the outer flow surface that pivots about a hinge that is selectively moveable to change the variable cross-sectional area thereby controlling an amount of the bleed flow through the cooling passage, wherein the outlet of the cooling passage is axially aligned with the hinge relative to the engine centerline axis.

2. The system as recited in claim 1 , wherein the annular fan bypass passage includes a pressure control plane that is controlled by movement of the at least one flap of the nozzle, wherein the outlet is axially aligned with the pressure control plane.

3. The system as recited in claim 1 , wherein the cooling passage includes an air-oil heat exchanger.

4. The system of claim 3 , wherein the air-oil heat exchanger is in thermal communication with a gas turbine engine component.

5. The system of claim 4 , wherein the gas turbine engine component is a fan drive gearbox.

6. The system of claim 4 , wherein the gas turbine engine component is an electric generator.

7. The system as recited in claim 1 , further comprising an actuator for selectively moving the at least one flap of the nozzle.

8. The system as recited in claim 7 , further comprising a controller for selectively commanding the actuator.

9. A gas turbine engine system comprising:

a nacelle;

a gas turbine engine within the nacelle;

an annular fan bypass passage between the nacelle and the gas turbine engine;

a cooling passage having an inlet for receiving a bleed flow from the annular fan bypass passage and an outlet for discharging the bleed flow into the annular fan bypass passage; and

a nozzle, disposed at an aft end of the annular fan bypass passage, having a variable cross-sectional area for controlling an airflow within the annular fan bypass passage, the nozzle including at least one flap located at the outer flow surface that pivots about a hinge that is selectively moveable to change the variable cross-sectional area to control an amount of the bleed flow through the cooling passage, wherein the outlet of the cooling passage is axially aligned with the hinge relative to an centerline axis of the gas turbine engine.

10. The system as recited in claim 9 , wherein the cooling passage includes an air-oil heat exchanger.

11. The system as recited in claim 10 , further comprising an actuator for selectively moving the at least one flap of the nozzle.

12. The system as recited in claim 11 , further comprising a controller for selectively commanding the actuator.

13. The system as recited in claim 9 , wherein the annular fan bypass passage includes a pressure control plane that is controlled by movement of the nozzle, wherein the outlet is axially aligned with the pressure control plane.

14. A method of controlling a gas turbine engine system including an annular fan bypass passage and a cooling passage having an inlet for receiving a bleed flow from the annular fan bypass passage and an outlet for discharging the bleed flow into the annular fan bypass passage, the method comprising:

controlling a flap at an exit of an exhaust nozzle disposed at an aft end of the annular fan bypass passage; and

varying a pressure of a control plane that is axially aligned with the outlet and a hinge of the flap of the exhaust nozzle of the annular fan bypass passage by selectively varying a cross-sectional area of the exhaust nozzle, to change an amount of bleed flow through the cooling passage.

15. The method as recited in claim 14 , further including decreasing the cross-sectional area to increase the pressure and decrease the amount of bleed flow.

16. The method as recited in claim 15 , further including decreasing the cross-sectional area in response to a cruise period of generally constant airspeed and generally constant altitude of an aircraft.

17. The method as recited in claim 14 , further including increasing the cross-sectional area to decrease the pressure and increase the amount of bleed flow.

18. The method as recited in claim 17 , further including increasing the cross-sectional area in response to a takeoff period of an aircraft.

19. The method as recited in claim 14 , further including selectively reducing the amount of bleed flow through the cooling passage in response to a cruise flight condition.

20. The method as recited in claim 14 , further including selectively increasing the amount of bleed flow through the cooling passage in response to a takeoff flight condition.

21. The method as recited in claim 14 , further including receiving 2% or less of the bleed flow in the cooling passage.