Abstract: An aircraft engine exhaust nozzle basesheet includes longitudinally extending plurality of basesheet segments having panel bodies between segment leading and trailing edges and slidable sealing joints with slidingly un-restrained center surfaces between adjacent segment leading and trailing edges. Segmented first and second basesheet side edges include first and second segment side edges of the basesheet segments respectively. Slidably sealingly engaged overlapping flanges at the segment leading and trailing edges include tacked together transversely spaced apart first and second distal ends of the overlapping flanges. A basesheet assembly includes a frame includes first and second side rails supporting the first and second basesheet side edges. An aircraft gas turbine engine exhaust nozzle includes divergent seals with the basesheet assemblies in sealing engagement against divergent flaps along the first and second side rails.

Abstract: An aircraft gas turbine engine exhaust nozzle basesheet includes longitudinally extending plurality of basesheet segments. Each of the basesheet segments includes a panel body between segment leading and trailing edges and slidable sealing joints with slidingly un-restrained center surfaces between adjacent segment leading and trailing edges. Segmented first and second basesheet side edges first and second segment side edges of the basesheet segments respectively. Slidably sealingly engaged overlapping flanges at the segment leading and trailing edges include tacked together transversely spaced apart first and second distal ends of the overlapping flanges. Leading and trailing edge ribs supporting the overlapping flanges are clipped together with longitudinally spaced apart and transversely extending forward and aft slits in retainers.

Abstract: A gas turbine engine exhaust nozzle includes a fluted shell terminating in a row of chevrons. The nozzle is radially serpentine circumferentially around the shells and has a circumferentially serpentine trailing edge.

Abstract: A method is provided for operating a gas turbine engine including a core engine having a core stream duct, an inner bypass duct, an outer bypass duct, and a nozzle assembly downstream of the core engine and including a core engine nozzle and a bypass nozzle separated by a liner. The method includes channeling a first airflow discharged from the core gas turbine engine to the core engine nozzle, and channeling a second airflow through the inner bypass duct such that the second airflow bypasses the core gas turbine engine. The second airflow is channeled to a plurality of fairings that are positioned upstream from a plurality of support struts coupled to the nozzle assembly liner. The method also includes channeling a third airflow through the outer bypass duct such that the third airflow bypasses the core gas turbine engine, wherein the third airflow is channeled through the support struts to the bypass nozzle.

Abstract: A variable geometry convergent-divergent nozzle for a gas turbine engine includes a centerbody extending rearward along a longitudinal axis of the engine which has a throat section of increased diameter. An inner shroud surrounds the centerbody and cooperates with the centerbody to define the throat of the nozzle. An outer shroud surrounds the inner shroud and cooperates with the centerbody to define the exit area of the nozzle. Both shrouds are independently translatable to provide independent control of the nozzle throat area and the nozzle expansion ratio.

Abstract: A method for adjusting a throat area of a jet aircraft exhaust nozzle assembly includes positioning a lower structure within a substantially rectangular nozzle assembly, coupling a ramp flap to the lower structure, and coupling an outer flap to the nozzle assembly such that movement of at least one of the ramp flap and the outer flap adjusts the throat area of the nozzle assembly.