Abstract: A turbofan engine deicing system includes a core nacelle (12) housing a turbine. A turbofan (20) is arranged upstream from the core nacelle. A controller (50) manipulates the turbofan in response to detecting an icing condition for avoiding undesired ice buildup on the turbofan engine (10) and nacelle parts. In one example, a variable area nozzle (40) is actuated to generate pressure pulses or a surge condition to break up any ice buildup. The icing condition can be determined by at least one sensor (52) and/or predicted based upon icing conditions schedules.

Abstract: A joint including a first component and a second component. The second component including a shape memory material component. The first component has a half dovetail shaped slot and the shape memory material second component has a half dovetail shaped attachment feature disposed in the half dovetail shaped slot in the first component.

Abstract: A nozzle system includes a multitude of circumferentially distributed divergent seals that circumscribe an engine centerline. Each divergent seal includes a multiple of divergent seal intakes adjacent to a joint structure to receive cooling airflow. Each divergent seal body is manufactured of a metallic hot sheet inner skin and a metallic cold sheet outer skin. The skins form a multiple of longitudinal channels which communicate with a multiple of edge channels formed within the first longitudinal side and the second longitudinal side of each divergent seal. The multiple of edge channels are located transverse to the longitudinal axis and are raked aft to facilitate cooling of the gas path surface of each divergent seal and adjacent divergent flaps.

Abstract: A method of controlling mixing of a flow exiting a downstream end of a primary nozzle associated with a jet engine. The method may involve coupling a shape memory alloy (SMA) element to a mixing structure disposed at the downstream edge of the primary nozzle. An electrical signal may be applied to the SMA element to heat the SMA element and induce a phase change in the SMA element. The phase change may cause an axial length of the SMA element to constrict, to cause movement of the mixing structure into a path of the flow exiting the primary nozzle.

Abstract: An exhaust gas cooling apparatus and method for cooling an exhaust gas is provided. The exhaust cooling apparatus having a first fluid conduit having a variable nozzle opening disposed at one end of the first fluid conduit, the variable nozzle opening being defined by a shape memory alloy extending from the first fluid conduit; and a second fluid conduit having an inlet end and an outlet end, the variable nozzle opening being disposed in the second fluid conduit proximate to the inlet end of the second fluid conduit such that a fluid inlet opening is disposed about an outer periphery of the variable nozzle opening and an inner surface of the inlet end of the second fluid conduit, the shape memory alloy varying the size of the variable nozzle opening by moving toward or away from a center line of the first fluid conduit in response to a temperature of an exhaust gas flowing through the first fluid conduit.

Abstract: Varying the fan nozzle flow area in a gas turbine engine allows adjustment of the bypass flow for varying operational conditions such as cruise and take off. However, care must be taken with regard to variation in order not to introduce drag or operational susceptibility to failure. By provision of a deformable member having a contiguous surface, which is deformed laterally from a splitter such as a pylon variation in flow area is achieved. A deformable member will be secured through seals to ensure a smooth surface. Actuation may be through a shape memory alloy material or provision of mechanical actuators to cause deformation of the contiguous surface.

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: The present invention relates to a deployable diverging bell (2) for a thruster, the bell comprising: a stationary first diverging bell portion (3) suitable for being connected to a stationary support of the thruster (1); a movable second diverging bell portion (4) that is movable between a retracted position and a deployed position in which it is connected to the downstream end of the first diverging bell portion (3) in order to extend it (3); and at least one deployment mechanism (6, 8) including means (6) for moving the movable second diverging bell portion (4) relative to the stationary first diverging bell portion (3) in order to deploy the movable second diverging bell portion (4) from its retracted position to its deployed position; the bell being characterized in that it includes at least one damper means (13) for damping vibration coming from the movable second diverging bell portion (4) and/or the deployment mechanism (6, 8).

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: A fluid flow control device (26, 28) comprising a guide member (30) to guide a fluid passing through a duct (22, 24, 76), the guide member (30) being movable between first and second positions; and an urging arrangement (32) capable of providing an urging force (F1) to urge the guide member (30) towards the first position, characterised in that the urging arrangement (32) is a resilient torsion bar that is formed so as to allow the guide member (30) to be moved towards the second position by a pressure force (F2) exceeding and opposite to the urging force (F1), the pressure force (F2) being provided by a pressure difference across the guide member (30).

Abstract: A propulsion system, such as for use in a missile, includes a multinozzle grid having a pair of plates that are separably mechanically coupled together. When coupled together in a first configuration, the plates provide multiple nozzles in a first nozzle configuration (geometry). Separation of the plates, such as by separating an aft plate from a forward plate that remains with the missile, reconfigures the multinozzle grid to a second configuration that has nozzles in a second nozzle configuration (geometry). The nozzle configurations may be suitable for different types of propulsion mechanisms. The hybrid propulsion system utilizing the multinozzle grid may include a pair of pressurized gas sources, for example a solid rocket fuel and a combustion chamber for a ramjet.

Abstract: A gas turbine engine system includes a fan (14), a housing (28) arranged about the fan, a gas turbine engine core having a compressor (16) at least partially within the housing, and a fan bypass passage (30) downstream of the fan for conveying a bypass airflow (D) between the housing and the gas turbine engine core. A nozzle (40) associated with the fan bypass passage includes a first nozzle section (54a) that is operative to move in a generally axial direction to influence the bypass airflow, and a second nozzle section that is operative to also move in a generally axial direction between a stowed position and a thrust reverse position that diverts the bypass airflow in a thrust reversing direction. An actuator (42) selectively moves the first nozzle section and the second nozzle section to influence the bypass airflow and provide thrust reversal.

Abstract: An inner core cowl baffle assembly for a turbofan engine assembly is provided. The engine assembly includes a core gas turbine engine, a core cowl which circumscribes the core gas turbine engine, a nacelle positioned radially outward from the core cowl, and a fan nozzle duct defined between the core cowl and the nacelle. The inner core cowl baffle assembly includes an inner core cowl baffle, and an actuator assembly configured to vary the throat area of the fan nozzle duct by selectively repositioning the inner core cowl baffle with respect to the core cowl.

Abstract: An exhaust nozzle includes a conical duct terminating in an annular outlet. A row of vortex generating duplex tabs are mounted in the outlet. The tabs have compound radial and circumferential aft inclination inside the outlet for generating streamwise vortices for attenuating exhaust noise while reducing performance loss.

Abstract: An aircraft nacelle comprising an outer cowl and an inner cowl located inward of the outer cowl with an air flow path formed therebetween. The inner cowl may comprise a forward portion and an aft portion slidable in a forward-to-aft direction via one or more actuators. When translated aftward, the aft portion may decrease the area of the nacelle's fan duct nozzle. The greatest radial distance between the inner cowl and a center axis of the nacelle may occur at or forward of a location in which the forward and aft portions overlap each other. The forward portion may be comprised of a different material than the aft portion.

Abstract: An exhaust diffuser includes a shroud and a wall radially separated from the shroud to define a fluid passage between the shroud and the wall. A strut extends between the shroud and the wall, and the strut includes a first surface having an adjustable camber. A method for adjusting air flow across a strut having a first side camber and a second side camber includes determining an incidence angle between the air flow and the strut and adjusting the first side camber of the strut.

Abstract: A mechanism included in a convergent-divergent nozzle connected to an aftward portion of a gas turbine engine, which mechanism includes at least one actuator configured to be connected to the aftward portion of the engine and to move through one or more positions, a first link pivotally connected to the at least one actuator, a convergent flap pivotally connected to the first link, a divergent flap rotatably connected to the convergent flap and configured to be kinematically connected to an annular ring connected to the at least one actuator, and at least one track configured to extend from the aftward portion of the engine and to movably and pivotally receive the pivotal connection between the convergent flap and the first link.

Abstract: A propulsion system with a divergent-convergent nozzle is provided with variable thrust by a pintle whose position relative to the nozzle throat is controlled by an actuator that includes a rotor that external to, and encircling, the nozzle, and that is coupled to either the pintle or to the nozzle shell by a linkage that translates the rotary movement of the rotor into linear movement of either the pintle or the shell.

Abstract: A jet engine exhaust nozzle flow effector is a chevron formed with a radius of curvature with surfaces of the flow effector being defined and opposing one another. At least one shape memory alloy (SMA) member is embedded in the chevron closer to one of the chevron's opposing surfaces and substantially spanning from at least a portion of the chevron's root to the chevron's tip.

Abstract: An exhaust duct assembly includes a cooling liner spaced apart from an exhaust duct case that articulate for use in a short take off vertical landing (STOVL) type of aircraft. The cooling liner assembly is attached to the exhaust duct case through a foldable attachment hanger system. The foldable attachment hanger system provides a low profile (foldable up/down) for a limited access installation envelope typical of a three bearing swivel duct (3BSD) which rotates about three bearing planes to permit transition between a cruise configuration and a hover configuration. In this way, each cooling liner segment may be formed as a complete cylindrical member requiring joints only between the swivelable duct sections.

Abstract: A variable area fan nozzle for use with a gas turbine engine system includes a nozzle section that is attachable to a gas turbine engine for influencing flow through the fan bypass passage of the gas turbine engine. The nozzle section is movable between a first length and a second length that is larger than the first length to influence the flow. For example, the nozzle section includes members that are woven together to form collapsible openings that are open when the nozzle section is moved to the first length and that are closed when the nozzle section is moved to the second length.

Abstract: A jet engine exhaust nozzle flow effector is a chevron formed with a radius of curvature with surfaces of the flow effector being defined and opposing one another. At least one shape memory alloy (SMA) member is embedded in the chevron closer to one of the chevron's opposing surfaces and substantially spanning from at least a portion of the chevron's root to the chevron's tip.

Abstract: A turbine engine provides a spool supporting a turbine. The spool is arranged in a core nacelle and includes a thrust bearing. A fan is arranged upstream from the core nacelle and is coupled to the spool. A fan nacelle surrounds the fan and core nacelle and provides a bypass flow path that includes a fan nozzle exit area. A flow control device is adapted to effectively change the fan nozzle exit area. A controller is programmed to monitor the thrust bearing and command the flow control device in response to an undesired load on the thrust bearing. Effectively changing the fan nozzle exit area with the flow control device actively manages the bearing thrust load to desired levels.

Abstract: A mechanism included in a convergent-divergent nozzle connected to an aft portion of a gas turbine engine, which mechanism includes a convergent flap configured to be kinematically connected to the aft portion of the gas turbine engine, a divergent flap pivotally connected to the convergent flap and configured to be kinematically connected to the aft portion of the gas turbine engine, at least one actuator configured to be connected to the aft portion of the gas turbine engine and to mechanically prescribe the position of the convergent flap and the divergent flap by moving through one or more positions, and at least one biasing apparatus. The biasing apparatus is configured to position at least one of the convergent flap and the divergent flap independent of the positions of the actuator by substantially balancing a force on the at least one of the convergent flap and the divergent flap produced by a gas pressure on the nozzle.

Abstract: An actuator arrangement comprises a plurality of linearly extendable actuators arranged to be driven by a common electrically driven motor, each actuator, being provided with limit stops to limit extension and/or retraction thereof, wherein the limit stops of the actuators are positioned such that a first one of the actuators has a smaller range of permitted extension than at least a second one of the actuators.

Abstract: An inner core cowl baffle assembly for a turbofan engine assembly is provided. The engine assembly includes a core gas turbine engine, a core cowl which circumscribes the core gas turbine engine, a nacelle positioned radially outward from the core cowl, and a fan nozzle duct defined between the core cowl and the nacelle. The inner core cowl baffle assembly includes an inner core cowl baffle, and an actuator assembly configured to vary the throat area of the fan nozzle duct by selectively repositioning the inner core cowl baffle with respect to the core cowl.

Abstract: A shape memory alloy (SMA) actuated aerostructure operable to dynamically change shape according to flight conditions is disclosed. Deformable structures are actuated by SMA actuators that are coupled to face sheets of the deformable structures. Actuating the SMA actuators produces complex shape changes of the deformable structures by activating shape changes of the SMA actuators. The SMA actuators are actuated via an active or passive temperature change based on operating conditions. The SMA actuated aerostructure can be used for morphable nozzles such as a variable area fan nozzle and/or a variable geometry chevron of a jet engine to reduce engine noise during takeoff without degrading fuel burn during cruise.

Abstract: A mechanism included in a convergent-divergent nozzle connected to an aft portion of a gas turbine engine, which mechanism includes a convergent flap configured to be kinematically connected to the aft portion of the gas turbine engine, a divergent flap pivotally connected to the convergent flap and configured to be kinematically connected to the aft portion of the gas turbine engine, at least one actuator configured to be connected to the aft portion of the gas turbine engine and to mechanically prescribe the position of the convergent flap and the divergent flap by moving through one or more positions, and at least one biasing apparatus. The biasing apparatus is configured to position at least one of the convergent flap and the divergent flap independent of the positions of the actuator by substantially balancing a force on the at least one of the convergent flap and the divergent flap produced by a gas pressure on the nozzle.

Abstract: A turbine engine provides a spool supporting a turbine. The spool is arranged in a core nacelle and includes a thrust bearing. A fan is arranged upstream from the core nacelle and is coupled to the spool. A fan nacelle surrounds the fan and core nacelle and provides a bypass flow path that includes a fan nozzle exit area. A flow control device is adapted to effectively change the fan nozzle exit area. A controller is programmed to monitor the thrust bearing and command the flow control device in response to an undesired load on the thrust bearing. Effectively changing the fan nozzle exit area with the flow control device actively manages the bearing thrust load to desired levels.

Abstract: A transition duct (112) is provided at the outlet of a turbine engine. The transition duct transitions the outlet of the tip turbine engine (10) from round to rectangular. The transition duct also provides a plurality of variable vanes at the outlet of the transition duct. The transition duct incorporates a perimeter slot (124) providing cooling to the duct outer wall and attached components. The variable vanes (120) are installed to vector the exhaust gases. This allows the aircraft to decelerate, hover or accelerate in the forward direction by commanding the position of the variable vanes. One potential application of the tip turbine engine is for vertical installations in aircraft.

Abstract: A balance pressure control is provided for flaps which pivot in a rear of a gas turbine engine nozzle to change the cross-sectional area of the nozzle. An actuator drives a sync ring to move the flaps through a linkage. A supply of pressurized air is also provided to the sync ring to assist the actuator in resisting forces from high pressure gases within the nozzle. When those forces are lower than normal the flow of air to the rear of the sync ring is reduced or blocked. A ring rotates with another ring to control both this air flow, and a supply of cooling air to an interior of the nozzle.

Abstract: A rocket thruster of this invention includes a nozzle and a valve assembly. The valve subassembly includes a pintle with a head portion, which has a truncated substantially conical surface region facing and concentric with converging and throat regions of the nozzle to provide a variable effective throat area therebetween. The truncated portion of the head portion has an outer edge defining a bleed passageway opening communicating with a primary bleed passage leading through the head portion to a bleed cavity located on an opposite side of the head portion. During activation of a rocket motor to which the rocket thruster is coupled, gases imparting a load on the head portion pass through the primary bleed passage to the opposite side of the head portion for counterbalancing loads acting on the head portion. A thrust control subassembly moves the pintle axially for changing the effective throat area.

Abstract: This invention is directed to nacelles and nacelle components for use in aircraft engines. The nacelles and components comprise composite material containing carbon fiber. The carbon fiber comprises from 0.1 to 20 percent by weight nanoreinforcement material.

Abstract: A gas turbine engine system includes nacelle, a first boundary layer control device, a second boundary layer control device, at least one sensor, and a controller. The nacelle is defined about an axis and includes an inlet lip section, an inlet internal diffuser section, and a variable area fan nozzle moveable to influence a discharge airflow area of the nacelle. The first boundary layer control device is positioned near the inlet lip section and is actuable to introduce a first airflow at the inlet lip section. The second boundary layer control device is positioned near the inlet internal diffuser section and is actuable to introduce a second airflow at the inlet internal diffuser section. The sensor produces a signal that represents an operability condition.

Abstract: The invention relates to a nacelle for an aircraft engine that comprises a front cowling (13) and a rear cowling (1a), the rear cowling (1a) being mounted so as to slide between an upstream position defining a reduced nozzle (9) section and a downstream position defining an enlarged nozzle (9) section. The nacelle includes an intermediate member (25) arranged edge-to-edge with said front cowling (13), said member defining a housing (27) for receiving the upstream edge (11) of said rear cowling (1a) when the latter is in the upstream position.

Abstract: A nacelle for an aircraft high-bypass jet engine, in which a jet engine having a longitudinal axis is mounted, the nacelle including a wall concentrically and at least partially surrounding the jet engine and defining with the latter an annular duct for fluid inner flow, including at the downstream end of the nacelle wall a passage section of a flow outlet. The nacelle includes a displacement mechanism displacing on request a portion of the nacelle wall to modify the passage section of the flow outlet through which a major portion of the flow escapes, the displacement forming in the nacelle wall at least one opening through which a small portion of a leak flow, naturally escapes. The nacelle further includes a fluid device that uses a fluid for compelling the leak flow to flow along the outer face of the portion of the nacelle wall located downstream relative to the at least one opening.

Abstract: Nozzles having variable geometry are disclosed that are comprised of a plurality of vanes that contact one another and can slide with respect to one another. A control element forces the vanes to slide, with contacting sides of the vanes defining an inner flow path through which a fluid may pass. The vanes may be self-similar or different from one another to define the desired geometries of the opening in the nozzles. The vanes may be configured to provide convergent, divergent and convergent-divergent nozzles for a range of applications.

Abstract: Integrated engine exhaust systems and techniques for operating integrated engine exhaust systems are disclosed. In one embodiment, a propulsion system includes an engine installation configured to be mounted on a wing assembly of an aircraft. The engine installation includes an engine, and an exhaust system operatively coupled to the engine. The exhaust system includes at least one nozzle to exhaust an exhaust flow from the engine. The nozzle includes a variable portion configured to vary an exit aperture of the nozzle from a first shape to a second shape to change a flowfield shape of at least a portion of the nozzle flowfield proximate the wing assembly, thereby reducing at least one of drag and thermal loading on the wing assembly. In a further embodiment, the exhaust system includes an inner nozzle that exhausts a core exhaust flow, and an outer nozzle that exhausts a secondary exhaust flow, the outer nozzle having the variable portion configured to vary the exit aperture of the outer nozzle.

Abstract: A nacelle assembly for a gas turbine engine includes a nacelle, a variable area fan nozzle, a sensor that detects a windmilling condition and a controller that communicates with the sensor. The variable area fan nozzle is moveable between a first position having a first discharge airflow area and a second position having a second discharge airflow area greater than the first discharge airflow area in response to detecting the windmilling condition.

Abstract: One embodiment of the present invention includes a nozzle defining a passage to receive and discharge working fluid to produce thrust. The nozzle includes the first wall structure opposite a second wall structure. The first wall structure includes a first convergent flap pivotally connected to a first divergent flap. The second wall structure includes a second convergent flap pivotally connected to a second divergent flap. The first wall structure and the second wall structure define the throat along the passage and are reconfigurable to adjust dimensional area of the throat. One or more control valves modulate flow of the pressurized fluid into the passage through a first opening in the first wall structure and a second opening in the second wall structure approximate to the throat to change effective area of the throat by fluidic control.

Abstract: A gas turbine engine has the convergent flaps and seals of an adjustable cross-sectional area nozzle connected to associated liners with a thermally compliant bracket. The bracket includes spaced feet with an intermediate notch spaced away from the liner. In this manner, a temperature gradient along the liner can be easily compensated at the connection of the bracket to the liner by adjustment within the notch. In other features, the bracket is riveted to the liner, providing further resistance to thermal expansion due to temperature gradiations.

Abstract: Nozzles that offer shape variability to maintain or purposely change the pressure drop across the throat are obtained by constructing the nozzles with components that change their shape, angle, or curvature in response to temperature changes that occur during the flow of combustion products through the nozzle. The temperature change may be the gradual heating of the nozzle wall from hot combustion gases, and the shape change may result in a decrease in the throat diameter or an expansion of the throat diameter. A decrease in throat diameter will be useful when the depletion of propellant as burning proceeds causes a drop in the pressure or flow rate of the combustion gas and there is a need to compensate for this drop to maintain the pressure drop across the throat. An increase in throat diameter will be useful when an initial high thrust is no longer needed and depletion of the fuel by itself is insufficient to lower the thrust to its desired reduced level.

Abstract: A turbofan engine includes an engine core and a ducted fan assembly, with the ducted fan assembly including an annular cowling and a dilating fan duct nozzle. The nozzle includes continuous nozzle sections with intermeshing tiles. The tiles include drive tiles that are pivotal between nominal and dilated positions and driven tiles that intermesh with the drive tiles and shift with the drive tiles between the positions. The intermeshing tiles cooperatively adjust an orifice size of the nozzle by shifting between the positions and thereby affect the thrust and noise developed by the ducted fan assembly.

Abstract: In order to provide noise suppression, bumps or undulations are provided on a nozzle surface in order to vary the convergent-divergent ratio between that surface and an opposed nozzle surface. By such an approach, a circumferential variation in the shock cell pattern is created and the flow is deflected so as to enhance turbulent mixing thereby suppressing noise.

Abstract: A gas turbine engine variable area exhaust nozzle has a plurality of first and second movable members arranged circumferentially around a fan casing extending in downstream. The first and second movable members alternate around the variable area nozzle. During cruise conditions the first and second movable members are in un-actuated positions. At take off conditions the first movable members are in an actuated position and the second movable members are in the un-actuated position. At top of climb conditions the first movable members are in the un-actuated position and the second movable members are in an actuated position. The first movable members move radially outwardly from the un-actuated position to the actuated position and the second movable members move radially inwardly from the un-actuated position to the actuated position.

Abstract: A thrust vectorable fan variable area nozzle (FVAN) includes a synchronizing ring, a static ring, and a flap assembly mounted within a fan nacelle. An actuator assembly selectively rotates synchronizing ring segments relative the static ring to adjust segments of the flap assembly to vary the annular fan exit area and vector the thrust through asymmetrical movement of the thrust vectorable FVAN segments. In operation, adjustment of the entire periphery of the thrust vectorable FVAN in which all segments are moved simultaneously to maximize engine thrust and fuel economy during each flight regime. By separately adjusting the segments of the thrust vectorable FVAN, engine trust is selectively vectored to provide, for example only, trim balance or thrust controlled maneuvering.

Abstract: A variable fan nozzle for use in a gas turbine engine includes a nozzle section, such as an inflatable bladder, associated with a fan bypass passage for conveying a bypass airflow. The nozzle section has an internal fluid pressure that is selectively variable to influence the bypass airflow.

Abstract: An aircraft engine nozzle assembly includes an adjustable flap portion that is movable about a pivot for changing the size of an exhaust exit area. A bracket having a curved slot establishes a range of possible movement of the flap portion. A link includes a guide member portion that is received in the curved slot for guiding the movement of the flap portion. The link is connected to a static engine structure and the bracket is movable with the flap portion. The curved slot establishes a range of possible movement of the flap portion along a direction of the curved slot.

Abstract: A fan variable area nozzle (FVAN) selectively rotates a synchronizing ring relative the static ring to adjust a flap assembly through a linkage to vary an annular fan exit area through which fan air is discharged. By adjusting the FVAN, engine thrust and fuel economy are maximized during each flight regime.

Abstract: A turbofan engine includes an engine core and a ducted fan assembly, with the ducted fan assembly including an annular cowling and a dilating fan duct nozzle. The nozzle includes continuous nozzle sections with intermeshing tiles. The tiles include drive tiles that are pivotal between nominal and dilated positions and driven tiles that intermesh with the drive tiles and shift with the drive tiles between the positions. The intermeshing tiles cooperatively adjust an orifice size of the nozzle by shifting between the positions and thereby affect the thrust and noise developed by the ducted fan assembly.