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A multi stream aircraft fixed geometry nozzle includes an inner nozzle, an outer nozzle, and a supersonic ejector. The outer nozzle is configured to channel a third stream from an aft end of a third stream duct surrounding a bypass duct of a multi stream aircraft engine to the supersonic ejector to

A multi stream aircraft fixed geometry nozzle includes an inner nozzle, an outer nozzle, and a supersonic ejector. The outer nozzle is configured to channel a third stream from an aft end of a third stream duct surrounding a bypass duct of a multi stream aircraft engine to the supersonic ejector to merge the third stream with the primary stream. The fixed geometry nozzle is configured to operate between an SFC mode and a thrust mode such that, when the inner nozzle accelerates the primary stream supersonically to the supersonic ejector, at which the primary stream is merged with the third stream, in the SFC mode the total pressure of the primary stream is substantially the same as the total pressure of the third stream, and in the thrust mode the total pressure of the primary stream is substantially greater than the total pressure of the third stream.

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1. A multi stream aircraft fixed geometry nozzle comprising: an inner nozzle;an outer nozzle disposed radially outward of the inner nozzle; anda supersonic ejector disposed axially aft of the inner nozzle and outer nozzle;a fan supplying motive fluid to form a bypass stream and a third stream;the in

1. A multi stream aircraft fixed geometry nozzle comprising: an inner nozzle;an outer nozzle disposed radially outward of the inner nozzle; anda supersonic ejector disposed axially aft of the inner nozzle and outer nozzle;a fan supplying motive fluid to form a bypass stream and a third stream;the inner nozzle being configured to channel a primary stream of a mixture of a propulsive core stream from an engine core and the bypass stream from a bypass duct surrounding the engine core, from an aft end of the engine core to the supersonic ejector; andthe outer nozzle being configured to channel the third stream from an aft end of a third stream duct surrounding the bypass duct to the supersonic ejector to merge the third stream with the primary stream;the fixed geometry nozzle including a flow control device comprising a plurality of flaps attached to a radially outer wall of the third stream duct which deploy radially inward from the radially outer wall, the flow control device configured to operate the fixed geometry nozzle between an SFC mode and a thrust mode such that, when the inner nozzle accelerates the primary stream supersonically to the supersonic ejector, at which the primary stream is merged with the third stream, in the SFC mode a total pressure of the primary stream is substantially the same as a total pressure of the third stream, and in the thrust mode the total pressure of the primary stream is substantially greater than the total pressure of the third stream. 2. The multi stream aircraft fixed geometry nozzle of claim 1 in which the inner nozzle has an internal fixed throat at its aft end and the primary stream is accelerated to the internal fixed throat before being accelerated supersonically to the supersonic ejector. 3. The multi stream aircraft fixed geometry nozzle of claim 1 in which in the SFC mode slip lines between the primary stream and the third stream are non radially expanding as viewed in cross section through an axis of the supersonic ejector. 4. The multi stream aircraft fixed geometry nozzle of claim 1 in which in the thrust mode slip lines between the primary stream and the third stream are radially outwardly expanding as viewed in cross section through an axis of the supersonic ejector. 5. The multi stream aircraft fixed geometry nozzle of claim 1 in which the supersonic ejector is configured so that an area ratio experienced by the primary stream in the supersonic ejector depends on a pressure differential between the primary stream and the third stream. 6. The multi stream aircraft fixed geometry nozzle of claim 1 in which the supersonic ejector is configured so that an area ratio experienced by the primary stream in the supersonic ejector depends on an amount of third stream passing through the outer nozzle. 7. The multi stream aircraft fixed geometry nozzle of claim 1 in which the supersonic ejector is configured so that the primary stream experiences an area ratio in the thrust mode about two times greater than the area ratio in the SFC mode. 8. The multi stream aircraft fixed geometry nozzle of claim 1 in which the outer nozzle comprises upper and lower outer nozzle portions, and the lower outer nozzle portion extends upward and around the inner nozzle to and into fluid communication with the upper outer nozzle portion, and the upper outer nozzle portion channels the third stream to an upper portion of the primary stream. 9. The multi stream aircraft fixed geometry nozzle of claim 8 in which the length of the inner nozzle is substantially the same as the length of the upper outer nozzle portion such that the distal end of the upper outer nozzle portion is substantially alongside the distal end of the inner nozzle, and further wherein the inner nozzle and the outer nozzle have an annular shape in axial cross section. 10. The multi stream aircraft fixed geometry nozzle of claim 1 further comprising an aircraft engine having the engine core configured to generate the propulsive core stream;the bypass duct surrounding the engine core and configured to channel the bypass stream from an inlet of the aircraft engine to the aft end of the engine core for mixture with the propulsive core stream to form the primary stream; andthe third stream duct surrounding the bypass duct and configured to channel the third stream from the inlet of the aircraft engine to a forward end of the fixed geometry nozzle. 11. A multi stream aircraft engine comprising: an exhaust nozzle;an engine core configured to generate a propulsive core flow;a fan supplying motive fluid to form a bypass flow and a third flow;a bypass flow duct surrounding the engine core and configured to channel the bypass flow to an aft end of the engine core to mix the bypass flow with the propulsive core flow to form a primary flow;a third flow duct surrounding the bypass duct and configured to channel the third flow to a forward end of the exhaust nozzle;the exhaust nozzle including a primary nozzle configured to accelerate the primary flow from the aft end of the engine core to an ejector of the exhaust nozzle, and an outer nozzle that channels the third flow from the forward end of the exhaust nozzle downstream to the ejector; anda flow control member configured to selectively control the amount of the third flow through the outer nozzle, the flow control member disposed in the third flow duct at the aft end of the third flow duct axially adjacent to a forward end of the exhaust nozzle and comprising a plurality of flaps attached to a radially outer wall of the third stream duct which deploy radially inward from the radially outer wall. 12. The multi stream aircraft engine of claim 11 in which the flow control member is configured to selectively control the amount of the third flow through the outer nozzle by directing a portion of the third flow channeled by the third flow duct. 13. The multi stream aircraft engine of claim 11 in which the flow control member comprises a composite material. 14. The multi stream aircraft engine of claim 11 in which the flow control member comprises one or more flaps. 15. A method comprising using an engine core to generate a propulsive core flow;channeling a bypass flow from a fan through a bypass flow duct surrounding the engine core to an aft end of the engine core and mixing the bypass flow with the propulsive core flow to form a primary flow;channeling a third flow from the fan through a third flow duct surrounding the bypass duct to a forward end of an exhaust nozzle;accelerating the primary flow through a primary nozzle of the exhaust nozzle from an aft end of the engine core to an ejector of the exhaust nozzle;channeling the third flow through a third flow nozzle of the exhaust nozzle from the forward end of the exhaust nozzle downstream to the ejector; andselectively changing the position of a flow control member that is disposed in the third flow duct at an aft end of the third flow duct axially adjacent to a forward end of the exhaust nozzle to control the amount of the third flow through the third flow nozzle, the flow control member comprising a plurality of flaps attached to a radially outer wall of the third stream duct which deploy radially inward from the radially outer wall. 16. The method of claim 15 in which the position of the flow control member is selectively changed to arrange the primary flow to have a total pressure that is substantially equal to a total pressure of the third flow to continue accelerating the primary flow through the primary nozzle. 17. The method of claim 15 in which the position of the flow control member is selectively changed to arrange the primary flow to have a total pressure that is relatively greater than a total pressure of the third flow so that the primary flow expands in the ejector to narrow an effective throat of the third flow, wherein a degree of expansion corresponds to an increase in thrust. 18. The method of claim 15 in which the position of the flow control member is selectively changed to change a pressure differential between the primary flow and the third flow to effect a corresponding area ratio in the primary flow in the ejector. 19. The method of claim 15 in which the position of the flow control member is selectively changed to change the amount of third flow passing through the exhaust nozzle to effect a corresponding area ratio in the primary flow in the ejector.