Abstract: A rocket engine nozzle manufacturable and applicable to tactical missile designs includes an aerospike having a plurality of airfoil fins distributed around a central longitudinal axis of a rocket engine combustion chamber. The aerospike is integrated on an exit plane at an exit end of the combustion chamber. The airfoil fins and an inner perimeter of the combustion chamber define a plurality of apertures which choke an airflow exiting the combustion chamber and cause the airflow to expand supersonically along the airfoil fins. The aerospike rocket engine nozzle requires less machine precision and achieves packing benefits over conventional bell and aerospike nozzle geometries. The configuration of the aerospike rocket engine nozzle also removes the producibility and heating constraints typically encountered with conventional aerospike nozzles in tactical missile applications while improving thrust performance of the rocket engine across a wide range of altitudes.

Abstract: A rocket engine nozzle includes an aerospike having a plurality of adjustable airfoil vanes distributed around a central longitudinal axis of a rocket engine combustion chamber. The aerospike is integrated on an exit plane at an exit end of the combustion chamber. The adjustable airfoil vanes and an inner perimeter of the combustion chamber define a plurality of apertures which choke an exhaust exiting the combustion chamber and cause the exhaust to expand supersonically along the adjustable airfoil vanes, creating a supersonic jet. An actuator is configured to adjust a position of each of the adjustable airfoil vane relative to each other so as to direct the exhaust exiting the rocket engine combustion chamber as the exhaust expands supersonically over the airfoil vanes without causing a shockwave to be imparted on the supersonic jet that is created. Accordingly, performance of the rocket engine is improved over conventional systems.

Abstract: An exhaust nozzle for a gas turbine engine according to an example of the present disclosure includes, among other things, a duct having a first surface and a second surface extending about a duct axis to define an exhaust flow path, and at least one effector positioned along the first surface. The at least one effector is pivotable about an effector axis to vary a throat area of the exhaust flow path. The at least one effector tapers along the effector axis. A method of exhaust control for a gas turbine engine is also disclosed.

Abstract: A convergent-divergent nozzle for a turbofan engine of a supersonic aircraft, wherein the nozzle has an inner wall and forms a flow channel with a nozzle throat surface and a nozzle exit surface. A trim ring that is displaceable in the axial direction between a first position and a second position is provided, wherein the inner wall of the nozzle and the trim ring are embodied and positioned with respect to each other in such a manner that the trim ring extends at a radial distance to the inner wall in the first position, wherein a ring-shaped bypass channel is provided, extending between the trim ring and the inner wall of the nozzle, and the trim ring abuts the inner wall in the second position.

Abstract: An actuation system for a gas turbine engine oriented about a centerline includes a translating ring, at least one hydraulic actuator, and at least one electric actuator. The translating ring is oriented about the centerline and configured to move axially along the centerline. The at least one hydraulic actuator is configured to provide a first mechanical force to move the translating ring along the centerline. The at least one electric actuator is configured to provide a second mechanical force to move the translating ring in the axial direction. The at least one electric actuator is controlled to provide the second mechanical force upon determination of an operating condition.

Abstract: A thrust reverser device for an aircraft turbojet engine includes: a movable cowl, translating cascades connected to the movable cowl, a mast; and a front suspension to suspend the turbojet engine. The cascades translates between a direct jet position where the cascades are retracted in a fan casing, and an indirect jet position where the cascades are brought out of the fan casing. In particular, the cascades have a sliding connection with the mast by means of at least one spacer connected to the mast downstream of the front suspension.

Abstract: Example folding door thrust reversers for aircraft engines are disclosed herein. An example apparatus includes a nacelle of a turbofan engine, where a fan duct is defined between the nacelle and a core of the turbofan engine. The example apparatus includes an opening in the nacelle between an outside of the nacelle and the fan duct and an inner door and an outer door disposed within the opening and pivotably coupled to the nacelle along aft edges thereof. The example inner and outer doors are pivotable between a first position in which the inner door and the outer door are disposed within the opening and oriented substantially parallel to each other, and a second position in which the inner door is disposed in the fan duct and oriented substantially perpendicular to an outer surface of the core and the outer door extends outward from the nacelle.

Abstract: A nozzle having a movable diverging section, the nozzle including a stationary throat forming a ball joint and a diverging section having a top portion that is movably mounted around the throat; an annular lubrication chamber being interposed between the stationary throat and the top portion of the diverging section, the annular lubrication chamber containing a grease seal; an annular throttling element arranged between the stationary throat and the diverging section downstream from the annular lubrication chamber, the annular throttling element being made of refractory material; and an annular thermal protection element arranged between the stationary throat and the diverging section downstream from the annular lubrication chamber and upstream from the annular throttling element, the annular thermal protection element closing the lubrication chamber in its bottom portion.

Abstract: Methods and apparatus to vary reverse thrust of aircraft engines are disclosed. An example cascade apparatus disclosed herein includes a cascade frame having a first end, a second end, and a fixed structure extending between the first end and second end, where the cascade frame defines a slot. A cascade forms a reverse thrust flow path and at least a portion of the cascade is slidably coupled to the frame via the slot. The cascade slides relative the frame between a first position to produce a first reverse thrust and a second position to produce a second reverse thrust, where the first reverse thrust is different than the second reverse thrust.

Abstract: A thrust-vectoring rocket motor nozzle includes a forward assembly having a forward shell with a flange configured for connection to a motor and a throat portion opposite the flange. A ball joint sleeve may be disposed proximate the throat portion, and an exit cone assembly may include a ball joint socket configured to mate with the ball joint sleeve to allow movement of the exit cone assembly about one or more axes relative to the forward assembly. A thermal barrier may be disposed in a gap between the forward assembly and the exit cone assembly. The forward assembly may include a throat insulator mechanically locked within the forward shell. Related methods include forming thrust-vectorable rocket motor nozzles. Rocket motors may include such nozzles.

Abstract: A thrust reverser unit (TRU) 100 for a gas turbine engine 10 is provided that has first and second cascade elements 110, 120. In a stowed configuration, both the first and second cascade elements, and the operating mechanism, are located inside the nacelle 40, meaning that the TRU has no detrimental impact on the flow through the bypass duct 22. In the deployed configuration, the first cascade element 110 extends across the bypass duct 22. The first cascade element 110 has flow passages 112 that allow the flow to pass through, redirecting the flow towards the second cascade element 120. The second cascade element 120 further turns the flow so as to provide decelerating reverse thrust.

Abstract: An exhaust centerbody for a turbine engine is provided. The centerbody includes a truncated downstream part, which is connected to an upstream part by an annular ridge marking a discontinuity between the outer surfaces of the upstream and downstream parts. The outer surface of the downstream part has a substantially conical general shape, of which the tip is oriented downstream and is positioned in the region of the axis A, the axial half-section of this outer surface defining a line of which the upstream end part is substantially tangential to a straight line passing through the ridge and forming a non-zero angle ? with a tangent to the outer surface of the upstream part, in the region of the ridge, and of which the downstream end part is substantially tangential to a straight line passing through the tip and forming a non-zero angle ? with the axis A.

Abstract: A gas turbine engine (30) comprising a rear cowl (38) defining an exhaust aperture (40) and a motive system. The rear cowl (38) comprises at least one panel (42) and the motive system is operable to selectively move the panel (42) between deployed and stowed configurations by rotation of the panel (42) about an axis substantially parallel to the main rotational axis of the engine (30). This alters the area of the exhaust aperture (40).

Abstract: A turbomachine for an aircraft including at least one turbine driving rotation of a fan, and a jet nozzle coaxially extending a turbine and creating, with an outlet cone that terminates the turbine, a passage of annular cross section for combustion gases that pass through the turbine. The jet nozzle can be mounted so that it can move between two extreme positions for which under action of a controller, an annular cross section of the passage for the gases is respectively at its minimum or maximum, making it possible to vary the cross section according to phases of operation of the aircraft and distribute thrust between the fan and the jet nozzle.

Abstract: The present invention relates to an aircraft gas turbine thrust-reversing device with an engine having an engine cowling whose rear area can be moved in the axial direction of the engine from a closed forward-thrust position into a rearward moved thrust-reversing position in which a free, substantially annular space to a front and fixed area of the engine cowling is present, where on the engine cowling guide rails are arranged, inside which mounting elements can be moved by means of rolling or sliding elements, with which mounting elements the rear area of the engine cowling is mounted, characterized in that the guide rails have an oval and C-shaped cross-section, provided on a longitudinal side with a slot-like opening, and are mounted using mounting projections arranged centrally and opposite to one another and designed in one piece with the C-shaped cross-section.

Abstract: A deployable nozzle for a rocket engine, the nozzle including at least a stationary divergent segment and a movable divergent segment that is coaxial about the stationary divergent segment and configured to move along the stationary divergent segment from a retracted position towards a deployed position. The deployable nozzle further includes a transverse stiffener that is prestressed in tension and that extends transversely relative to the movable divergent segment in a vicinity of a downstream end of the movable divergent segment between at least two points at a periphery of an inside wall of the movable divergent segment.

Abstract: A turbofan engine nacelle has a downstream section which includes at least one front frame. The front frame is attached to a stationary portion of the nacelle, and the downstream section is provided with at least one rail or slide which extends in a longitudinal direction of the nacelle and is capable of engaging with at least one corresponding slide or rail of an attachment pylon of the turbofan engine. The front frame is connected to the rail or slide of the downstream section by means of at least one swiveled connecting rod.

Abstract: The present invention relates to an aircraft gas turbine thrust-reversing device with an engine having an engine cowling, the rear area of which can be displaced in the axial direction of the engine from a closed forward thrust position into a rearwardly displaced thrust reversal position, resulting in an essentially annular free space towards a forward stationary area of the engine cowling, with the rear area of the engine cowling being functionally coupled to deflecting elements arranged in the forward thrust position within the front area of the engine cowling, with the deflecting elements during displacement of the rear area of the engine cowling being moveable on a partial-circular path facing the central axis of the engine and contactable by their rear end areas with a cowling of the core engine.

Abstract: A propulsion system of an aircraft includes an inner fixed structure (IFS) and the outer sleeve that may be separately coupled to the pylon. For instance, the inner fixed structure and the outer sleeve may move independently with respect to each other. A single latch of the latching system disclosed herein, may be configured to latch all 4 panels, (e.g. both halves of the IFS and both doors of the outer sleeve) together and/or in a closed and retained position.

Abstract: A system for enhancing movement of a multi-engine flight vehicle about either of pitch or yaw axes. The engines are oriented relative the centerline of the flight vehicle and a portion of the engines have a positive tangential cant and the remaining engines have a negative tangential cant. The attitude of the flight vehicle about the pitch, yaw and roll axes is controlled by differentially throttling the multi-engine flight vehicle.

Abstract: The invention relates to a method for the self-calibration of electric jacks (6a, 6b) capable of actuating a mobile member (2) of a turbojet nacelle (1) and connected to at least one position measuring member, characterized in that it comprises the steps of: moving the jack(s) into a retracted position corresponding to a first position of the mobile member, storing one or more position values fed back by the position measuring member for the jack(s) in said retracted position; moving the jack(s) into an extended position corresponding to a second position of the mobile member; storing one or mom position values fed back by the position measuring member for the jack(s) in said extended position.

Abstract: The invention relates to a method for controlling a plurality of actuators (9a, 9b, 9c) for a movable thrust reverser cowl (1), characterized remarkable in that the position deviation between adjacent actuators is measured in real time, and in that the speed profile of the actuator or actuators in question is adjusted by a position deviation exceeding a predetermined threshold.

Abstract: An air discharging device (20) for an aircraft turbine engine, comprising at least one door (24) displaceable between an open and a closed position of a corresponding orifice (30) and comprising two valves (26, 28), which delimit between them a conduit (68) for guiding a portion (74) of the secondary flow (76) outwards in the downstream direction, and which are integral with each other and are hinged around a pivot axis (58), so that in said open position, the upstream end of said internal valve (26) protrudes from the inner side relative to the internal surface (12), the downstream end of said external valve (28) protrudes from the external side relative to the external surface (14), and said internal valve (26) is spaced away from the fixed structure (22) so that an air passage exists downstream of said internal valve (26), between the latter and said fixed structure (22).

Abstract: A control system for controlling a plurality of distinct functions of a turbojet engine, each function being associated with a respective actuation device. The system includes: an electric motor adapted to supply mechanical energy to each of the actuation devices; an electronic control unit for the electric motor; and at least one switching device interposed between the motor and the actuation devices, the switching device(s) serving to distribute the mechanical energy supplied by the electric motor selectively to one of the actuation devices.

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: Disclosed is a structure for a grid-type thrust reverser, the reverser including an upper beam, a lower beam, and two half-frames fastened to the upper beam and the lower beam, and hollow structures defined by the two half-frames, the hollow structures each having upper ends and lower ends, wherein at least one of the upper and lower beams includes integral parts delimited by walls, the walls forming casings fitted into and blocking corresponding ends of the half-frames.

Abstract: A variable area fan nozzles comprising an array of rigid petals hinged to a lip area at the downstream end of a thrust reverser sleeve (or half-sleeve). In one embodiment, the actuation system comprises a rotatable ring segment, a drive system for rotating the ring segment, a plurality of cams attached or integrally formed with the ring segment, a plurality of cam followers, and a plurality of petal linkages which operatively couple the petals to the cam followers. This actuation system controls the deflection of the petals, thereby controlling the amount of opening and the rate at which the fan nozzle throat area changes. Each cam may have forward- and rearward-facing camming surfaces that interact with the cam followers to force the petals to pivot outward or inward respectively.

Abstract: A propulsion thrust control system and method for controlling thrust in a rocket motor includes configuring valves of an energized rocket motor to an initial total valve area according to a total thrust command. The total thrust command is converted into a commanded propellant mass flow discharge rate. A varying total valve area is computed from an error between the commanded propellant mass flow discharge rate and a calculated propellant mass flow discharge rate. The valves are reconfigured according to a distribution of the varying total valve area. The propulsion system includes a pressure vessel with valves and a controller for regulating the valve area according to a propellant mass flow discharge rate from the pressure vessel.

Abstract: This leaf spring (17) for a thrust reverser shutter (13) is noteworthy in that it has two branches (19, 21) defining a U shape, the ends (19a, 19b, 21a, 21b) of these branches being able to cooperate respectively with a shutter (13) of said reverser and with a link (15) for actuating this shutter.

Abstract: A nacelle assembly includes a first portion having an outer fairing and a trailing edge, and a translatable variable area fan nozzle. The fan nozzle includes two or more nozzle segments, each nozzle segment having first and second opposed ends and a leading edge. The nozzle segments are selectively movable between a stowed position and one or more deployed positions. In the deployed position, an upstream bypass flow exit is formed between the trailing edge and the leading edge. The nacelle assembly further includes a guide mechanism for guiding the nozzle segments between the stowed position and the deployed position. A split beavertail failing shields the guide mechanism against air flow when the nozzle segments are in the stowed position.

Abstract: A method and aerodynamic device for modifying at least one vane of a cascade thrust reverser. One or more of the aerodynamic devices may be formed and attached to forward-facing surfaces of vanes of the thrust reverser. The aerodynamic device may include a first side shaped to rest flush against at least a portion of the forward-facing surface of the vane, a second side opposite of the first side, a ledge portion protruding from the first side and configured to rest flush against a blunt leading edge of the vane, a rounded first end portion extending from the ledge portion to the second side of the aerodynamic device, and a second end opposite of the first end portion at which the first side and the second side converge. The aerodynamic device changes the size and contour of airflow channels between the vanes, increase the efficiency of the thrust reverser.

Abstract: A by-pass turbojet including a thrust reverser is disclosed. The thrust reverse includes a deflector device which deflects all or part of the primary stream gas in such a manner that the deflected primary stream gas encounters the secondary stream, thereby reducing the injection speed of the secondary stream in an aft direction, and thus generating the thrust-reversal effect. The secondary stream escapes from the aft end of the nacelle. In the thrust-reversal position, the deflector device is inscribed radially substantially within the section of the primary nozzle cowl at the aft end of the nacelle. Such a thrust reverser is particularly simple in design, inexpensive, and makes it possible to avoid having moving parts present on the outside portion of the nacelle, thus simplifying the design of the turbojet.

Abstract: An assembly for an aircraft gas turbine propulsion engine includes an engine nacelle and a thrust reverser. The engine nacelle is configured to have at least a portion of an aircraft gas turbine propulsion engine mounted therein, and includes an intake end, an exhaust end, and an outer surface that extends between the intake end and the exhaust end. At least the engine nacelle outer surface proximate the exhaust end has a semicircular section and a noncircular section. The thrust reverser includes a forward end, an aft end, and an outer surface that extends between the forward and aft ends. The thrust reverser forward end is coupled to the engine nacelle exhaust end. At least the thrust reverser outer surface proximate the forward end has an arc section and a noncircular section. The arc section subtends an angle greater than ? radians.

Abstract: The invention relates to an exhaust system (100) for channeling the streams from a by-pass gas turbine, the system comprising a stream channeling nozzle and an exhaust casing (110) for connecting the channeling nozzle to the outlet of the gas turbine, said channeling nozzle comprising a primary nozzle (120) fastened to the exhaust casing (110) and a secondary nozzle (130) placed around the primary nozzle. The exhaust system further comprises means for fastening the secondary nozzle (130) directly to the exhaust casing (110), said secondary nozzle being supported by the exhaust casing independently of the primary nozzle (120).

Abstract: A turbomachine including at least one thrust reverser and a device for actuating the thrust reverser, which includes at least one hydraulic engine supplied with pressurized fluid by a fuel circuit of the turbomachine, is disclosed.

Abstract: A device is provided for pivoting at least one pivotable element in a gas turbine engine between a first and a second position in order to influence a gas flow in an annular gas duct in at least one of the positions. The device includes a moveable annular member which is arranged externally around the gas duct and is connected to the pivotable element in order to effect the pivoting of the pivotable element. The annular member is more specifically arranged to be displaced in a substantially axial direction and arranged to pivot the pivotable element when it is displaced axially.

Abstract: An example nozzle for use in a gas turbine engine includes a nozzle door having a first end, a second end opposed from the first end, and a first pivot on the door between the first and second ends. A linkage connects to the nozzle door and to an actuator. The actuator is selectively operative to move the linkage about a second pivot to in turn move the nozzle door about the first pivot between a plurality of positions, such as stowed, intermediate, and thrust reverse positions to influence bypass airflow through a fan bypass passage. The fan bypass passage has a radially outward side defined by a nacelle in at least one position. At least a portion of the nozzle door is radially outward of the nacelle and includes a lip extending from the nozzle door between the first pivot and one of the first end and second end.

Abstract: Latch beams and hinge beams of an aircraft thrust reverser and a method for manufacturing the beams. The latch and hinge beams may comprise a plurality of hollow composite tubes joined with a plurality of machined metal fittings in alternating succession. The metal fittings may comprise at least one of hinge fittings and latching fittings. The latch and hinge beams may further comprise slider tracks configured to slidably attach to a translating sleeve of the thrust reverser. The hinge beams may each be rotatably attached to an engine strut or pylon of the aircraft, and the latch beams may each be mechanically latched with each other. Each of the latch and hinge beams may also be fixedly attached to an inner engine cowling of the thrust reverser.

Abstract: A rear fan case for a gas turbine engine includes a main body and a discrete joining arrangement mounted on the outside of the main body towards its rear end. The joining arrangement provides a substantially radial groove around the main body, which can receive a joining blade on a thrust reverser unit to form a joint between the rear fan case and the thrust reverser unit. The joining arrangement may be made of metal while the main body is made of composite material. Features may be provided to ensure proper alignment and location of the joining arrangement.

Abstract: A gas turbine engine system includes a fan bypass passage (27), a core nacelle (28) having an inner fixed structure (40) within the fan bypass passage, a passage (42) extending through the inner fixed structure, and a duct nozzle (48). The passage includes an inlet (44) for receiving a fan airflow (F2) from the fan bypass passage and an outlet (46) for discharging the fan airflow. The duct nozzle includes a variable cross-sectional exit area (50) for controlling the fan airflow within the passage and is selectively moveable to influence the variable cross-sectional exit area.

Abstract: A turbofan engine (10) includes a fan variable area nozzle (FVAN) (50) which effectively changes the physical area and geometry within a fan bypass flow path (40) to manipulate the pressure ratio of the bypass flow. The FVAN generally includes a multitude of aerodynamically shaped inserts (52) circumferentially located about the core nacelle (12). The FVAN at a fully stowed position (takeoff/landing) takes up a minimum of area within the fan bypass flow path to effectively maximize the fan nozzle exit area (44) while in the fully deployed position (cruise) takes up a maximum of area within the bypass flow path to effectively minimize the fan nozzle exit area. By separately adjusting each of the multiple of inserts relative the other inserts the FVAN provides an asymmetrical fan nozzle exit area to selectively vector the fan bypass flow.

Abstract: The rocket engine nozzle system comprises a set of individual nozzles distributed in a ring around a central core of axially symmetrical shape that presents a central axis. Each individual nozzle placed at the periphery of the central core comprises a circular section throat for receiving gas from a combustion chamber, and a diverging portion that is tangential to the central core. The diverging portion has an outlet section that presents first and second lateral sides that converge radially towards the central axis of the central core, a curved outer side having its convex side facing outwards, and an inner side situated in the vicinity of the central core.

Abstract: A gas turbine engine includes a fan, a bypass duct having a nozzle, through which fluid from the fan flows, and a variable area nozzle device located in the nozzle. The variable area nozzle device has an arcuate fairing having one end in slidable cooperation with the bypass duct. In a deployed position the area of the nozzle is reduced from that when stowed, thereby improving engine operability and efficiency.

Abstract: A method for exhausting gas from an aircraft engine assembly is provided. The method includes coupling a first exhaust duct in fluid communication only to a first engine. The first exhaust duct includes a primary outlet and a secondary outlet. The method further includes coupling a second exhaust duct in fluid communication only to a second engine. The second exhaust duct includes a primary outlet and a secondary outlet. The method also includes aligning a portion of the first engine secondary outlet concentrically within a portion of the second engine secondary outlet.

Abstract: A missile in which a nose portion is rotatably mounted on a main body portion of the missile and is subjected to thrust to bring it to a demanded roll attitude and to apply to it a demanded lateral steering force. The thrust is produced by a propellant gas and supplied to discharge ducts in the nose portion which provide for the discharge of the propellant gas tangentially with respect to the nose portion. The discharge ducts are selectively opened and closed by relative axial displacement of the nose portion and a valve spool which controls the flow of propellant gas. An actuator responsive to guidance control signals causes a predetermined roll torque to bring the nose portion to a demanded roll attitude and a predetermined lateral steering force on the missile.

Abstract: In a bypass turbomachine, at least one air takeoff orifice is provided in the primary channel, said orifice leading to an air takeoff pipe housed inside an air inlet sleeve of the nacelle. The air takeoff pipe opens out in the vicinity of the pylon into two air diffusion pipes each secured to a respective maintenance cover, each air diffusion pipe opening out into an air injection pipe secured to a thrust reverser cover and itself opening out to the outside of the nacelle via the trailing edge thereof, each air injection pipe being suitable for uncoupling from the corresponding air diffusion pipe when the corresponding thrust reverser cover slides downstream.

Abstract: A gas turbine engine comprising a compressor, a combustion chamber, and at least two turbines mounted oppositely to the combustion chamber, such that the gas turbine engine is capable of generating multidirectional thrust.

Abstract: A flow reverser comprising a sleeve having a first part and a second part to define a flow path and axially separable from each other about a conjunction formed by a respective profile edges for the first part and the second part, the profile edges in a stowed position overlapping, the reverser characterised in that in a deployed position a nozzle part of each profile edge define together a nozzle jet and a reverser part of one edge profile is adjacent to a core to provide an effective flow deflector.

Abstract: A gas turbine engine comprising a core engine a fan and a fan casing surrounding the fan and extending rearwardly to attach to an array of outlet guide vanes characterised in that a cascade is rigidly attached to the fan casing and is connected to the core engine via a second support.

Abstract: A turbofan exhaust system includes at least one exhaust nozzle with a trailing edge having an upper edge, a lower edge, an inner lateral edge-and an outer lateral edge. The upper edge is upstream with respect to the lower edge. The outer lateral edge of the exhaust nozzle is also upstream of the inner lateral edge. This arrangement makes it possible to reduce the noise perceived in the cabin of the aircraft at the same time as the noise perceived on the ground.