Abstract: A method for fabricating a composite structure. A first section for the composite structure is formed in which the first section has a first end with a chevron shape, wherein first composite layers in the first section has a first step pattern at the first end. A second section for the composite structure is formed in which the second section has a second end with a counterpart shape to the chevron shape and in which second composite layers in the second section have a second step pattern at the second end. The first end the second end are positioned such that a first composite layer in the first composite layers in the first step pattern overlap the second composite layers in the second step pattern at a splice location.

Abstract: A variable area nozzle assembly for a gas turbine engine includes a fixed structure surrounding an exhaust duct extending along a nozzle centerline. The fixed structure includes an upper side and a lower side opposite the upper side. The variable area nozzle assembly further includes a nozzle disposed about the nozzle centerline. The nozzle includes a nozzle throat cross-sectional area and a nozzle outlet cross-sectional area downstream of the nozzle throat cross-sectional area. The nozzle includes an upper panel and a lower panel. The upper panel includes an upper downstream end and the lower panel including a lower downstream end. The upper downstream end and the lower downstream end define a portion of the nozzle throat cross-sectional area. The variable area nozzle assembly further includes a nozzle actuation system including an upper shaft connected to the upper panel and a lower shaft connected to the lower panel.

Abstract: A ducted fan housing for directing a duct flow includes an annular cowling and an adjustable fan duct nozzle. The nozzle includes a flexible sleeve having rigid areas arranged circumferentially around the nozzle orifice and connected by flexible areas so as to form a unitary sleeve structure. The rigid areas are radially moveable between a normal configuration in which the orifice is smaller and a dilated configuration in which the orifice is larger. A drive mechanism uses drive elements to move at least some of the rigid areas between the configurations to adjust the size of the orifice. The rigid areas may be constructed from laminated graphite and epoxy, and the flexible areas may be constructed from laminated graphite and soft resin. If the housing includes a thrust reverser, then a flexible joint area extends circumferentially around the housing and connects the flexible sleeve to a thrust reverser cowl.

Abstract: A two-dimensional variable area plug (2D VAP) nozzle assembly for a high-speed flight vehicle. In one embodiment, a 2D VAP nozzle assembly comprises a nozzle including a plurality of sidewalls; a plug body within the nozzle, the plug body abutting at least two of the plurality of sidewalls; a first convergent flap hingedly connected to at least one of the plurality of sidewalls; and a second convergent flap hingedly connected to at least one of the plurality of sidewalls. In one embodiment, the nozzle assembly includes only a first convergent flap and a second convergent flap, without diverging flaps. The 2D VAP nozzle assembly has a simplified design with reduced sidewall length, which results in reduced manufacturing and maintenance costs.

Abstract: A gas turbine engine includes a front center body case structure. A geared architecture is at least partially supported by the front center body case structure. A bearing structure is mounted to the front center body case structure to rotationally support a shaft driven by the geared architecture, the shaft drive a fan. A bearing compartment passage structure is in communication with the bearing structure through the front center body case structure. A method is also disclosed.

Abstract: A heat exchanger system for a gas turbine engine is disclosed. The heat exchanger system may include a first structure at least partially defining a first plenum configured to receive a first air stream, a second structure at least partially defining a second plenum configured to receive a second air stream having lower pressure than the first air stream, a third structure at least partially defining a third plenum configured to receive a third air stream having lower pressure than the second air stream, and a heat exchanger configured for operative communication with the first air stream, the second air stream, and the third air stream while disposed between the second air stream and the third air stream. The heat exchanger may be configured to transfer heat from the first air stream to the third air stream.

Abstract: Methods of replacing flex seals in exhaust frames of turbine systems are disclosed. The method may include removing a section of an outer casing of an exhaust frame to form an opening within the outer casing. The section may be formed within a flow path of the exhaust frame. The method may also include removing the flex seal from the exhaust frame via the opening formed within the outer casing, inserting a distinct flex seal within the exhaust frame via the opening formed within the outer casing, and covering the opening formed within the outer casing of the exhaust frame.

Abstract: Tail cone apparatus and methods for reducing nozzle surface temperatures of aircraft engines are disclosed. An example apparatus includes a tail cone to be coupled to an aircraft engine. The tail cone includes a central axis, a cone section, and a plurality of fins. The fins are spaced about the central axis and extend outwardly from an outer surface of the cone section.

Abstract: A rocket motor includes a case and first and second nozzles in the case. The first nozzle is disposed in the second nozzle. The first nozzle includes a forward leg, a rear leg, and an intermediate leg. The intermediate leg has a convex conical geometry, and the forward leg has a forward lip that is spaced from the case. The rear leg has a rear lip that is spaced from the case. The forward leg and the rear leg at least partially define a flow passage through the first nozzle. The first nozzle is exclusively secured by the intermediate leg to at least one of the case or the second nozzle. At least a portion of a fragmentation system is disposed between the first and second nozzles.

Abstract: The disclosure relates to a turbofan nacelle including a thrust reverser, the thrust reverser including a movable cover moving back from a closed position in which the thrust is not reversed to an open position for uncovering cascades that reverse the direction of the flow of cold air which is diverted from the annular stream of secondary air, the movable cover including a radially outer portion intended for being adjacent to a leading edge of a wing of an aircraft. The movable cover includes on the radially outer portion at least one cut-out intended for avoiding interference with a movable slat of the leading edge of the wing of the aircraft, as well as a panel for closing the cut-out, the closing panel including a stationary portion at least partially covered by an upstream portion of the movable cover when the movable cover is in the open position.

Abstract: A thrust reverser includes: a thrust-reversing element movable between a stowed position and a deployed position; at least one hydraulic actuator operably coupled to move the thrust-reversing element between the stowed and deployed positions; at least one hydraulic primary lock configured to transition, in response to a first activation pressure, between an engaged state, where movement of the thrust-reversing element is inhibited, and a released state, where movement of the thrust-reversing element is uninhibited; and a directional control unit fluidly coupled to the hydraulic actuator and the hydraulic primary lock, the directional control unit configured to transition from a first stage to a second stage in response to a second activation pressure that is greater than the first activation pressure, and where a transition from the first stage to the second stage by the directional control unit causes the hydraulic actuator to move the thrust-reversing element to the deployed position.

Abstract: A heat exchanger apparatus including a surface cooler and a passive automatic retraction and extension system coupled to the surface cooler. The surface cooler having disposed therein one or more fluid flow channels configured for the passage therethrough of a heat transfer fluid to be cooled. The heat transfer fluid in a heat transfer relation on an interior side of said one or more fluid flow channels. The surface cooler including a plurality of fins projecting from an outer surface thereof. The passive automatic retraction and extension system including a thermal actuation component responsive to a change in temperature of at least one of the heat transfer fluid and a cooling fluid flow so as to actuate a change in a geometry of the surface cooler. Further disclosed is an engine including the heat exchanger apparatus.

Abstract: A gas turbine engine system includes an airframe support, a gas turbine engine mounted to the airframe support, a nozzle exhaust mounted to the airframe support, and a flexible annular seal having a first end that is attached to the gas turbine engine and a second end that is attached to an exhaust nozzle. The flexible annular seal may include a flexible annular wall that permits relative deflection between the gas turbine engine and the exhaust nozzle to facilitate the reduction of undesirable load transfer.

Abstract: A turbofan engine (10) employs a flow control device (41) that changes an effective exit nozzle area (40) associated with a bypass flow path (B) of the turbofan engine. A spool (14) couples a fan (20) to a generator (52). The turbofan emergency power system includes a controller (50) that communicates with the flow control device (41). Upon sensing an emergency condition, the controller manipulates the flow control device to reduce the effective nozzle exit area (40) of the bypass flow path, which chokes the flow through the bypass flow path thereby increasing the rotational speed of the fan. In this manner, the generator is driven at a higher rotational speed than if the flow through the bypass flow path was not choked, which enables a smaller generator to be utilized.

Abstract: A nacelle assembly for a high-bypass gas turbine engine includes a core nacelle defined about an engine centerline axis. A fan nacelle is mounted at least partially around the core nacelle to define a fan bypass flow path. A variable area fan nozzle is in communication with the fan bypass flow path. The variable area fan nozzle has a first fan nacelle section and a second fan nacelle section. The second fan nacelle section is axially movable relative to the first fan nacelle section to define an auxiliary port at a non-closed position to vary a fan nozzle exit area and adjust fan bypass airflow. The second fan nacelle section includes an acoustic system that has an acoustic impedance located on a radially outer surface.

Abstract: An apparatus installed on an aircraft, comprising: a sleeve or duct having a trailing lip area; a plurality of petals arranged side by side with gaps therebetween, one end of each petal being attached or pivotably coupled to the lip area; and a plurality of elastomeric seals configured and disposed to close the gaps between adjacent petals. Each elastomeric seal comprises a first portion that moves with a portion of a first petal that is in contact therewith, a second portion that moves with a portion of a second petal that is in contact there, and a third portion which is stretched as the first and second petals move further apart from each other. Petal deflection is actuated by a system comprising a flexible member, a motor, a shaft driven by the motor, and an arm projecting from the shaft. One end of the flexible member is attached to the arm, the flexible member being movable to deflect the petals inward in response to a shaft rotation.

Abstract: A turbofan engine includes a variable area fan nozzle which effectively changes the physical area and geometry within a fan bypass flow path to manipulate the pressure ratio of the bypass flow with a multitude of bladders circumferentially located about a core cowl.

Abstract: A bypass gas turbine engine includes a variable area fan nozzle with a second fan nacelle section axially movable relative to a first fan nacelle section to define an auxiliary port to vary a fan nozzle exit area and adjust fan bypass airflow. The second fan nacelle section includes a leading edge region that defines a concave external contour and the first fan nacelle section includes a trailing edge region that defines a convex internal contour.

Abstract: A high-bypass gas turbine engine includes a variable area fan nozzle with an acoustic system having an acoustic impedance.

Abstract: A method of varying a fan duct nozzle throat area of a gas turbine engine includes pivoting a fan nozzle outwardly relative to a longitudinal axis of the gas turbine engine. The fan nozzle is configured to move axially non-contemporaneously with the pivoting of the fan nozzle.

Abstract: The invention relates to a linear telescopic actuator for moving a first (10b) and a second (10a) element relative to a stationary element (102). Said actuator comprises a base (101) that is to be connected to the stationary element (102) and is used as a cavity for a first rotationally locked rod (106) which can be translated by a drive shaft (104) that is to be connected to rotational driving means (107). One end (108) of said first rod is to be connected to the first element that is to be moved. The actuator is characterized in that the first rod (106) supports a second rod (117) which is aligned therewith and one end (118) of which is to be connected to the second element that is to be moved. Said second rod (117) can be rotationally locked and can be translated by a second drive shaft (112, 115) which extends through the base and is connected to rotational driving means (113, 111).

Abstract: An engine apparatus is provided for a flying object, the engine apparatus including a combustion chamber having a throat region and a nozzle region, the nozzle region having a nozzle wall, wherein the nozzle region expands from the throat region towards an exit end relative to a combustion chamber axis, wherein the nozzle region has associated therewith a skirt having a skirt wall, the skirt being positioned downstream relative to the exit end and surrounding the exit end of the nozzle region, and wherein the skirt wall is at an acute angle away from the combustion chamber axis with respect to the nozzle wall, at least at the exit end of the nozzle region.

Abstract: A thrust reverser for a turbojet engine includes a telescopic linear actuator for a moving cowl and a nozzle-forming end part. The telescopic linear actuator has a base attached to a fixed frame, acting as a housing for a first rod prevented from rotating but able to be translationally driven by a first drive shaft connected to a rotational-driver. The first rod is attached by one end to the moving cowl and supports a second rod, which is attached by one end to the nozzle-forming end part that is to be moved. The second rod is prevented from rotating and translationally driven by a second drive shaft passing through the base and connected to a rotational-driver. In particular, the rotational-driver of the rods includes a motor to drive an input shaft of a differential having two output shafts.

Abstract: An external flap connection assembly for a turbine engine nozzle has a track frame, a slider track, and an external flap. The external flap is removably connected to the slider track via a retaining member which has a slider block.

Abstract: A seal assembly for a turbojet nacelle includes a first end fastened to a first structure, and a second end contacting against a bearing zone of a second structure. The seal assembly further includes a plurality of adjacent blades arranged along the first end and extending longitudinally and perpendicularly thereto. In particular, a portion of the blades has an accordion-shaped structure.

Abstract: A method of varying a fan duct throat area of a gas turbine engine may include non-pivotably moving a fan nozzle outwardly relative to a longitudinal axis of the gas turbine engine during axially aft translation of the fan nozzle without activating a thrust reverser.

Abstract: An apparatus and method comprising a fan nozzle sleeve, a plurality of slider mechanisms, and a fan nozzle actuator system. A flow of gases generated by an engine moves through and exits the engine at an aft end of the fan nozzle sleeve. The plurality of slider mechanisms is configured to connect the fan nozzle sleeve to a translating sleeve for a thrust reverser for the engine. The fan nozzle actuator system is configured to activate the plurality of slider mechanisms to move the fan nozzle sleeve in an aft direction to change a direction of the flow of gases exiting the engine at the aft end of the fan nozzle sleeve.

Abstract: A turbofan gas turbine engine (10) comprises a variable area exhaust nozzle (12) arranged at the downstream end of a casing (17). A control unit (66) analyzes the power produced by the gas turbine engine (10), the flight speed of the gas turbine engine (1) and/or the altitude of the gas turbine engine (10). The control unit (66) configures the variable area nozzle (12) at a first cross-sectional area (70A) when the flight speed of the gas turbine engine (10) is less than a first predetermined value. The control unit (66) configures the variable area nozzle (12) at a second, smaller, cross-sectional area (70B) when the flight speed of the gas turbine engine (10) is greater than the first predetermined value and the power produced by the gas turbine engine (10) is greater than a second predetermined value.

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: The present invention relates to a jet engine nacelle (1) comprising an aft section (5) having an internal structure (9) intended to surround an aft portion of an engine compartment and to define, with an exhaust nozzle (6), a calibrated outlet section for the ventilation of the engine compartment by way of spacing means arranged in the outlet section, characterized in that the spacing means are divided into rigid spacing means (11) designed to provide a constant spacing and into compensating means (13) designed so as to be able to adapt to the relative movements of the jet engine with respect to the nacelle.

Abstract: A core nacelle for a gas turbine engine, according to an exemplary aspect of the present disclosure includes, among other things, a core cowl positioned adjacent to an inner duct boundary of a fan bypass passage having an associated cross-sectional area that radially extends between a fan exhaust nozzle and the inner duct boundary. The core cowl includes at least one groove that is selectively exposed to change the cross-sectional area at an axial location of the fan exhaust nozzle.

Abstract: The present invention relates to an aircraft gas turbine with a core engine which is surrounded by a bypass duct enclosed radially outwards by a bypass wall, with a radially inner wall of the bypass duct forming with the radially outer bypass wall a bypass nozzle, and with the radially inner wall including an adjusting element extending along the circumference of the inner wall and being deformable radially outwards.

Abstract: A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a nacelle assembly that includes an inlet lip section and an inlet internal diffuser section downstream of the inlet lip section. A variable area fan nozzle is positioned near an aft segment of the nacelle assembly, the variable area fan nozzle adaptable to move 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. At least one boundary layer control device is positioned near one of the inlet lip section and the inlet internal diffuser section. A controller is configured to move the variable area fan nozzle from the first position to the second position and to actuate the at least one boundary layer control device to introduce an airflow in response to an operability condition.

Abstract: A gas turbine engine is provided having an offtake passage that in one form is capable of extracting a bypass flow from the engine. The airflow traversing the offtake passage is introduced to an exhaust flow of the gas turbine engine through an offtake outlet. The offtake outlet includes an airflow member that is moveable. A nozzle is also provided for exhaust from the gas turbine engine. In one form the nozzle includes moveable duct members. Flows exiting the offtake outlet and the nozzle can be combined after passing the airflow member and the moveable duct members, respectively.

Abstract: The present disclosure relates to a double-acting linear actuator for moving an inner part and an outer part of a cowl relative to a fixed frame. The linear actuator includes a first tubular body housing a first drive shaft and a second tubular body housing a second drive shaft. In particular, the first and second tubular bodies are mounted in series by the second drive shaft which is mounted on the first tubular body. The second drive shaft translates the second tubular body relative to the first tubular body when a lock of the first tubular body is in a locked position, and the first drive shaft translates both the first tubular and second tubular bodies when the lock is in an unlocked position.

Abstract: A high-bypass gas turbine engine includes a variable area fan nozzle with an acoustic system having an acoustic impedance.

Abstract: A fan variable area nozzle (FVAN) includes a flap driven through a cable which circumscribes the fan nacelle. The cable is strung through a multiple of flaps to define a flap set of each circumferential sector of the EVAN. An actuator system includes a compact high power density electromechanical actuator which rotates a spool to deploy and retract the cable and effectively increase or decrease the length thereof between the spool and a fixed attachment to increase and decrease the fan nozzle exit area.

Abstract: A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a nacelle assembly that includes an inlet lip section and an inlet internal diffuser section downstream of the inlet lip section. A variable area fan nozzle is positioned near an aft segment of the nacelle assembly, the variable area fan nozzle adaptable to move 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. At least one boundary layer control device is positioned near one of the inlet lip section and the inlet internal diffuser section. A controller is configured to move the variable area fan nozzle from the first position to the second position and to actuate the at least one boundary layer control device to introduce an airflow in response to an operability 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 regulator system includes a slidable ramp intermediate a secondary flow path and a primary flow path of a gas turbine engine nozzle section.

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 nacelle for an aircraft engine is provided that includes a thrust reverser cowling slideably mounted between a direct jet position and a reversed jet position, an outlet jet pipe nozzle with variable cross-section positioned in a downstream extension of the reverser cowling, and means for respectively actuating the cowling and the nozzle. The nozzle is slideably mounted on the thrust reverser cowling, and the means includes at least one actuator for actuating said thrust reverser cowling, at least one driving pinion rotatably mounted on a fixed structure of said nacelle, and at least one rack for actuating the nozzle, secured to the nozzle and meshing with the driving pinion when the thrust reverser cowling is in the direct jet position, and escaping from this pinion when said reverser cowling is in the reversed jet position.

Abstract: A method of providing a slim-line nacelle for a gas turbine engine includes the steps of detecting a windmilling condition and increasing a discharge airflow area of a variable area fan nozzle in response to the detection of the windmilling condition.

Abstract: A variable area fan nozzle for use with a gas turbine engine system includes a nozzle having a cross-sectional area associated with a discharge air flow from a fan bypass passage. The nozzle is selectively moveable between multiple static positions for varying the cross-sectional area and multiple thrust reverse positions that are different from the static positions for reversing a direction of the discharge flow from the fan bypass passage to produce a thrust reversing force.

Abstract: A thrust reverser assembly for use in a turbofan engine assembly. The engine assembly includes a core gas turbine engine and a core cowl which circumscribes the core gas turbine engine. A nacelle is positioned radially outward from the core cowl to define a fan nozzle duct between the core cowl and a portion of the nacelle. The nacelle includes a stationary cowl. The thrust reverser assembly includes a first translating cowl that is slidably coupled to the nacelle. The first translating cowl is positionable with respect to the stationary cowl. A second translating cowl is slidably coupled to the nacelle such that the first translating cowl is positioned between the stationary cowl and the second translating cowl. The second translating cowl is positionable with respect to the first translating cowl. A positioning assembly is coupled to the first translating cowl. An actuator assembly is operatively coupled to the second translating cowl for selectively moving the second translating cowl.

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 apparatus for controlling a fluid flow is disclosed herein. The apparatus includes first and second spaced walls defining opposite sides of a first fluid passageway extending along a first axis. At least one of the first and second walls is cantilevered along the first axis and includes a distal end moveable relative to the first axis. The distal end of the at least one wall is moveable in response to changes in the flow temperature to passively vary a size of the first fluid passageway.

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: A multi-layered honeycomb structure adapted to reduce and/or eliminate thermal deformation is disclosed herein. In some embodiments, walls of the honeycomb structure comprise a first layer, optionally, a second layer, and a core layer adjacent to the first layer or between the first and second layers. The first and second layers may be compositions of Inconel or other high strength materials. The core layer may be copper or another thermally conductive material. The core layer is adapted to transmit heat between a first region of a structure and a second region. In this manner, heat can be transferred from the heated region to an unheated region, thereby reducing the temperature difference between the regions and thus the amount of thermal deformation.

Abstract: A turbomachine comprises a flow duct with coaxially arranged radially inner and outer endwalls, first and second sets of axially spaced rotor stages arranged between the inner and outer endwalls, and a plurality of variable endwall segments arranged along the inner endwall. The first set of rotor stages rotates in a first direction, and the second set rotates in a second direction. The first and second sets alternate in axial series along the flow duct, such that axially adjacent rotor stages rotate in different directions. The variable endwall segments are radially positionable, in order to regulate loading on the first and second sets of rotor stages by changing a cross-sectional flow area between the inner and outer endwalls.