High bypass ratio engine bypass duct nozzle with controlled nozzle area

1. A nacelle assembly of a gas turbine engine, comprising:

an annular structure defining a central axis, and having a radially inward surface and a radially outward surface, the radially inward surface at least partially defining a bypass duct;

an aft portion of the radially inward surface at least partially defining an axially extending convergent-divergent exit nozzle;

a secondary nozzle flap radially spaced from the aft portion of the radially inward surface, the secondary nozzle flap and the aft portion of the radially inward surface defining a secondary bypass duct therebetween, the secondary nozzle flap operably connected to the annular structure such that the secondary nozzle flap is selectably movable in a direction along the central axis relative to the aft portion of the radially inward surface, thereby changing a cross-sectional area of a secondary bypass duct;

wherein one or more of the radially inward surface or a radially outer flap surface include a convex protrusion such that when the nozzle flap is driven from an axially extended position toward an axially stowed position, the cross-sectional area increases.

2. The nacelle assembly of claim 1 , further comprising an actuator operably connected to the secondary nozzle flap and disposed in the annular structure configured to drive movement of the secondary nozzle flap.

3. The nacelle assembly of claim 2 , wherein the actuator is one of an electrical or hydraulic actuator.

4. The nacelle assembly of claim 1 , wherein a leading edge of the secondary nozzle flap is radially offset from a trailing edge of the secondary nozzle flap, relative to the central axis.

5. The nacelle assembly of claim 1 , wherein the secondary nozzle flap includes a radially outward flap surface having one or more undulations or protrusions.

6. A gas turbine engine comprising:

an engine core;

a fan operably connected to the engine core and driven by the engine core;

a nacelle surrounding the engine core and the fan and defining a bypass duct between the nacelle and the engine core, the nacelle including:

an annular structure defining a central axis, and having a radially inward surface and a radially outward surface, the radially inward surface and the engine core defining the bypass passage;

an aft portion of the radially inward surface at least partially defining an axially extending convergent-divergent exit nozzle; and

a secondary nozzle flap radially spaced from the aft portion of the radially inward surface, the secondary nozzle flap and the aft portion of the radially inward surface defining a secondary bypass duct therebetween, the secondary nozzle flap operably connected to the annular structure such that the secondary nozzle flap is selectably movable in a direction along the central axis relative to the aft portion of the radially inward surface, thereby changing a cross-sectional area of a secondary bypass duct exit;

wherein one or more of the radially inward surface or a radially outer flap surface include a convex protrusion such that when the nozzle flap is driven from an axially extended position toward an axially stowed position, the cross-sectional area increases.

7. The gas turbine engine of claim 6 , further comprising an actuator operably connected to the secondary nozzle flap and disposed in the annular structure configured to drive movement of the secondary nozzle flap.

8. The gas turbine engine of claim 7 , wherein the actuator is one of an electrical or hydraulic actuator.

9. The gas turbine engine of claim 6 , wherein a leading edge of the secondary nozzle flap is radially offset from a trailing edge of the secondary nozzle flap, relative to the central axis.

10. The gas turbine engine of claim 6 , wherein the secondary nozzle flap includes a radially outward flap surface having one or more undulations or protrusions.

11. A method of operating a gas turbine engine, comprising:

defining a bypass duct of the gas turbine engine between a core of the gas turbine engine and a nacelle disposed radially outboard of the core, an aft portion of the radially inward surface at least partially defining an axially extending convergent-divergent exit nozzle; and

positioning a secondary nozzle flap radially spaced from the aft portion of the radially inward surface, the secondary nozzle flap and the aft portion of the radially inward surface defining a secondary bypass duct therebetween,

selectably moving the secondary nozzle flap in a direction along a central longitudinal axis of the gas turbine engine relative to the aft portion of the radially inward surface, thereby changing a cross-sectional area of a secondary bypass duct exit;

wherein one or more of the radially inward surface or a radially outer flap surface include a convex protrusion such that when the nozzle flap is driven from an axially extended position toward an axially stowed position, the cross-sectional area increases.

12. The method of claim 11 , wherein the secondary nozzle flap is selectably movable on one or more of an axial direction or a radial direction.

13. The method of claim 11 , wherein the secondary nozzle flap is driven by an actuator disposed in the annular structure.

14. The method of claim 13 , wherein the actuator is one of an electrical or hydraulic actuator.

15. The method of claim 11 , wherein a leading edge of the secondary nozzle flap is radially offset from a trailing edge of the secondary nozzle flap, relative to the central axis.

16. The method of claim 11 , wherein the secondary nozzle flap includes a radially outward flap surface having one or more undulations or protrusions.