Abstract: Provided is a variable-throat exhaust turbocharger in which a nozzle assembly including a nozzle mount and nozzle vanes has a firm support structure without affection by a thermal deformation of a turbine casing and an external force exerted to the turbine casing, and a scroll portion is formed in a substantially opened configuration so that the nozzle assembly can be simply formed in order to reduce the number of mold cores for a scroll during casting of the turbine casing, thereby it is possible to enhance the productivity of the turbine casing.

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: An axisymmetric nozzle for a gas turbine engine has divergent flaps and seals disposed about a central longitudinal axis thereof. The divergent flaps and seals each have an inner surface defining cooling air inlet holes at an upstream portion, cooling air exit holes at a downstream portion, and cooling air channels disposed within the divergent flap and seal, and communicating at a first end with the inlet holes and at a second end with the exit holes for conducting cooling air therethrough. Some of the exit holes of the divergent flap are disposed along lateral edge regions thereof. The divergent seals are interposed between adjacent divergent flaps, and each include lateral edge regions extending laterally beyond the exit holes. The lateral edge regions of each divergent seal have an outer surface overlying at least a portion of the exit holes of an adjacent divergent flap.

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: A thermal management system locates a separate engine heat exchanger and generator heat exchanger into an engine nacelle pylon that is split to provide independent control of both a generator cooling circuit and an engine cooling circuit. A variable exhaust nozzle is provided downstream of each heat exchanger as each has different thermal and heat rejection requirements such that intake air is minimized and maximum thrust is produced in response to the current generator load and engine heat rejection.

Abstract: A method facilitates assembling a flap system for a gas turbine engine exhaust nozzle including at least one backbone assembly. The method comprises providing a basesheet including a pair of circumferentially-spaced sides coupled together by an upstream side and a downstream side, forming at least one relief cut in the basesheet that extends at least partially across the basesheet from at least one of the circumferentially-spaced sides, and coupling the basesheet to the backbone assembly.

Abstract: An exhaust nozzle bell for a rocket engine including a first part arranged on a motor of the rocket engine, and a second part coupled to the first part. The second part having a stowed position in which the second part surrounds the first part and is positioned closer to the motor, and an operating position in which the first part and the second part form a continuous shape and the second part is arranged farther from the motor. Moreover, the nozzle can include an extension mechanism structured and arranged to extend the second part from the stowed position to the operating position, the extension mechanism including a plurality of swiveling extension arms, wherein the extension arms have first and second ends. Furthermore, each of the extension arms comprise one of a sliding and rolling element on an end facing the second part of the exhaust nozzle bell.

Abstract: A method for extending a nozzle for a rocket engine is described, in which at least one second region of the contour of the nozzle is arranged as an extension of a first region of the contour of the nozzle. This extension of the first region by the second region occurs at an altitude at which the contour of the open jet of the rocket engine substantially corresponds to the contour of the second region. Also described is a corresponding extendible nozzle for a rocket engine.

Abstract: A turbine engine nozzle assembly has an upstream flap and a downstream flap pivotally coupled thereto for relative rotation about a hinge axis. An actuator linkage is coupled to the downstream flap along a forward portion thereof for actuating the upstream flaps and downstream flaps between a number of throat area conditions while permitting mode changes.

Abstract: A device for supplying cooling air to a hollow flap of an exhaust nozzle in a turbojet engine. The device includes a tube connecting the hollow flap to a cooling air source. The tube comprises at least one telescopic portion and at least one ball-joint connection.

Abstract: An exhaust nozzle assembly for a gas turbine engine, the assembly comprises a main axis, an inner nozzle, an outer nozzle, a translatable centre-body and an actuator for translating the centre-body between a forward position and a rearward position. The centre-body is disposed radially inwardly of the inner nozzle thereby partly defining an inner duct for a core engine gas flow, an outer duct for a bypass gas flow is defined by the inner nozzle and the outer nozzle. The outer nozzle extends downstream of the inner nozzle and with the centre-body defines a final mixing duct having a final exhaust exit area. When the centre-body is in the forward position the final exhaust exit area is at a maximum area, and when the centre-body is in a rearward position the final exhaust exit area is at a minimum area.

Abstract: A controlled hot flap shutter for a turbojet nozzle includes a friction surface made of metal, and including straight lateral flanks to reduce the gap between each controlled hot flap shutter and the tracking flap shutters adjacent thereto.

Abstract: A thrust control valve includes a nozzle having a gas supply chamber, a gas jetting chamber and a connecting passage therebetween. The nozzle has an upstream inner surface defining the gas supply chamber and a downstream inner surface defining the gas jetting chamber. The valve also includes a plug axially movably inserted into the gas supply chamber, the connecting passage and the gas jetting chamber. The plug has an upstream outer surface facing the upstream inner surface to define an upstream passage and a downstream outer surface facing the downstream inner surface to define a downstream passage. The plug can control thrust by simultaneously changing sectional areas of the upstream and downstream passages by an axial movement of the plug.

Abstract: An exhaust nozzle includes an exhaust duct with an outlet and a row of radial apertures upstream therefrom. A radial frame surrounds the duct upstream from the apertures. A row of flaps are hinged to the frame to selectively cover and uncover the apertures for controlling exhaust flow discharged therethrough. An arcuate unison bar surrounds the duct adjacent to the frame and includes circumferentially spaced apart cam followers engaging corresponding cams affixed to the flaps. An actuator is joined to the bar for selective rotation thereof between opposite first and second directions to pivot open and closed the flaps atop the apertures.

Abstract: The disclosed device is directed toward a periodic combustion propulsion system. The periodic combustion propulsion system comprises at least one periodic combustion chamber configured to contain periodic combustion and at least one supersonic exhaust nozzle coupled to the at least one periodic combustion chamber. The supersonic exhaust nozzle is configured with a variable exhaust expansion ratio.

Abstract: In a variable-section turbomachine nozzle, a plurality of flexible follower flaps are provided that are interposed between adjacent controlled flaps against which the follower flaps are held by fastener means, each follower flap comprising a plurality of independent transverse boxes mounted side by side in a support sheet, the boxes contacting one another via their longitudinal edges and being held in the support sheet via their lateral edges which are engaged in the longitudinal edges of the support sheet, which longitudinal edges define respective bearing zones for bearing against the adjacent controlled flaps.

Abstract: A convergent/divergent nozzle for a gas turbine engine has a throat portion of non-constant radius of curvature varying from an upstream high value to an intermediate low value and then to a downstream high value.

Abstract: The present invention relates to a convergent-divergent turbojet nozzle (1) comprising driven divergent flaps (4a), follower divergent flaps (4b) interposed between the driven flaps, and means (32) for supplying cooling air to the follower flaps (4b), said follower flaps (4b) having a box structure and having lateral openings for delivering cooling air towards the inner face of said driven flaps (4a), in such a way as to limit the heating up of these flaps when the turbojet is in operation.

Abstract: A side thruster valve of an aerospace craft is improved to reduce torque of a servo-motor, etc. needed for nozzle opening and closing to thereby realize a compact and light weight device. The side thruster valve comprises a valve plug having its back directed to an axis of the aerospace craft and independently movable between a fully opened position and a fully closed position of the valve plug in a plane orthogonal to the axis of the aerospace craft, an actuating means for moving the valve plug in an axial direction of the valve plug and an elastic member for activating the valve plug in the axial direction of the valve plug. In a side thruster device comprising a plurality of the side thruster valves, the side thruster valves are arranged independently of each other to thereby broaden freedom of combustion control and improve fuel consumption.

Abstract: A gas turbine engine exhaust nozzle includes a divergent section located aft of a convergent section and a throat therebetween. An exterior fairing is spaced radially outwardly of the divergent section. An ejector cooling air flowpath leads from an ejector cooling air inlet in an aft portion of the fairing to a cooling air ejector in the nozzle. An annular nozzle plenum may be disposed between the divergent section of the nozzle and the external fairing and be part of the ejector cooling air flowpath between the ejector cooling air inlet and the ejector. A plurality of divergent flaps and divergent seals in the divergent section may employ cooling air passages, such as slots, to serve as the ejector. The fairing may include a plurality of circumferentially adjacent exterior flaps and exterior seals and employ truncated ends of or apertures in the exterior seals as the ejector cooling air inlet.

Abstract: The disclosed device is directed toward a periodic combustion propulsion system. The periodic combustion propulsion system comprises at least one periodic combustion chamber configured to contain periodic combustion and at least one supersonic exhaust nozzle coupled to the at least one periodic combustion chamber. The supersonic exhaust nozzle is configured with a variable exhaust expansion ratio.

Abstract: A turbomachine comprising, downstream from the turbine, a post-combustion chamber defined by a thermal protection lining inside a casing defining an annular channel for a secondary flow, an annular diaphragm secured at the downstream end of said channel, said nozzle comprising a plurality of flaps hinged to the upstream end of said casing, each flap fitted on its inside face with a thermal protection plate to define a passage for cooling air delivered by said diaphragm, wherein the cooling air is provided by an annular duct defined by a first flexible annular gasket in sliding contact against the downstream inside face of the casing and against the upstream inside faces of the flaps under the pressure of the secondary flow, and defined on the inside by a second flexible annular gasket whose upstream end is fixed to the radially inner zone of the diaphragm, and whose downstream end is in sliding contact against the upstream inside face of the protection plates.

Abstract: The invention proposes a long life nozzle flap for aircraft turbojets. This flap is remarkable in that it is composed of a hollow tapered body (10) flattened in the transverse direction along straight geometric generating lines (11), the body (10) forming a thin wall (12) with an approximately constant thickness E, this wall (12) comprising a continuous inner surface with a radius of curvature equal to at least 2×E, the wall (12) being made of a refractory composite material composed of reinforcing fibers (18) embedded in a matrix also made from a refractory material, the reinforcing fibers (18) being arranged in a plurality of continuous layers of fibers (18) stacked on each other, the fibers (18a) of a layer crossing the fibers (18b) of any adjacent layer.

Abstract: The present invention reveals a method and apparatus for controlling the effective area and thrust vector angle of a fluid flow. In one embodiment, the fluid flow is controlled in an advanced, high aspect ratio, complex aperture geometry nozzle using asymmetric injection into the subsonic portion of the fluid flow. The present invention vectors the primary flow by partially blocking the flow with an opposed flow across the flow field. A fluidic flow field is defined in a flow container that directs a pressurized, primary fluidic flow from the container towards an exit of the container. A nozzle may cooperate with the exit of the flow container to control the fluidic flow as it exits the flow container. One or more injectors associated with the container are proximate to the effect throat of the primary flow while other are located downstream of to introduce an opposing fluidic flow that interacts with the primary fluidic flow.

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

Abstract: A basesheet for use in an aircraft gas turbine engine exhaust nozzle includes a plurality of basesheet segments extending between longitudinally extending spaced apart basesheet leading and trailing edges, respectively, of the basesheet. Right and left hand basesheet side edges, respectively, extend longitudinally between the basesheet leading and trailing edges. Each basesheet segment has a panel body defined between a pair of segment side edges longitudinally extending between segment leading and trailing edges, respectively. Stiffeners having leading and trailing edge ribs at the segment leading and trailing edges of aft and forward ones of the basesheet segments, respectively, extend widthwise across the panel body between the segment side edges. The leading and trailing edge ribs are joined together only at intermediate sections of the stiffeners which are centered between left and right hand expansion sections, respectively, of the stiffener.

Abstract: A center plug is configured for variable area use in an exhaust nozzle. A forward portion of the plug includes an inlet having a closing flap therein, and an aft portion of the plug includes an outlet. The flap is selectively open during aircraft takeoff operation and up to a preset climb altitude for bypassing a portion of core exhaust flow through the plug, with the flap being closed from preset altitude and cruise operation. Opening the flap correspondingly increases the available flow area for discharging the exhaust flow for reducing velocity thereof, and attenuating noise.

Abstract: The actuator assembly comprises a master actuator and at least one slave actuator having a cylinder containing a piston defining two chambers in the cylinder. A first chamber of the slave actuator is connected to an outlet of a directional control valve which is connected to a hydraulic control circuit and to a control member whose position determines the pressure in the first chamber of the slave actuator. A displacement of the piston of the master actuator is transferred to the control member of the control valve associated with the slave actuator by means of a mechanical transmission such as a flexible linear rolling bearing. Thus, the displacement of the piston of the slave actuator is servo-controlled to the displacement of the piston of the master actuator by controlling the position of the control member of the control valve associated with the slave actuator.

Abstract: An engine nozzle is described in which the nozzle is deformable from a first state, typically a round shape, to a second state which may approach a polygon or have a fluted edge. The first and the second state defining different nozzle cross-sections whereby the nozzle can be adjusted between the first and second state for a nozzle cross-section optimised for engine efficiency. Deformation in the nozzle may be achieved through use of shape memory materials, pressure chambers, piezo-electric element deformation or other technique.

Abstract: A method for extending a nozzle for a rocket engine is described, in which at least one second region of the contour of the nozzle is arranged as an extension of a first region of the contour of the nozzle. This extension of the first region by the second region occurs at an altitude at which the contour of the open jet of the rocket engine substantially corresponds to the contour of the second region. Also described is a corresponding extendible nozzle for a rocket engine.

Abstract: A turbofan exhaust nozzle includes a fan duct defined between a fan nacelle and a core engine cowling. The fan duct includes a longitudinal endwall and a partition spaced therefrom to define a secondary flow duct. The fan duct also includes a primary outlet, and the secondary duct includes a secondary outlet. The partition between the two ducts includes an aperture covered by a selectively movable flap.

Abstract: A method for assembling a flap system for a gas turbine engine exhaust nozzle including a plurality of backbone assemblies facilitates extending a useful life of the exhaust nozzle. The method includes providing a flap basesheet having a width defined between a pair of side edges that are coupled together by a leading edge and a trailing edge, and including at least one stiffener that extends between the basesheet side edges and includes an intermediate portion that has a width that is smaller than that of the basesheet and is at least one of bonded to and formed integrally with the basesheet, and coupling the basesheet to the gas turbine engine with a backbone assembly.

Abstract: The present invention comprises a rocket motor nozzle that replaces the fixed-wall throat with a “wall” created by injecting gas radially into the nozzle. By injecting gas into the flow of combustion products that are going through the nozzle, this will deflect the combustion product flow, restricting and accelerating such flow, just as a fixed-wall does in a standard nozzle. However, by using the “gas-wall” described herein, no erosion will result at the restriction area.

Abstract: A method for assembling a flap and seal system for a gas turbine engine exhaust nozzle including a plurality of backbone assemblies facilitates attaching a basesheet to a backbone. The method includes attaching an attachment system including at least one strap to a basesheet, and coupling the basesheet to a backbone using the attachment system strap.

Abstract: Control apparatus for a nozzle for a rocket or other vehicle having a combustion chamber with a propellant therein in communication with the nozzle, the nozzle having a throat and a pair of bores extending into the throat The bores may be in opposed or any other suitable relation. A pair of plungers are slidably mounted in the bores and are movable between an open position wherein they are disposed outside of the throat and a closed position wherein they extend into the throat to substantially close it. The plungers have a width or diameter substantially the same as the width or diameter of the throat. Actuator devices are provided for moving the plungers independently of each other to control thrust by controlling the flow of combustion gases through the throat and the pressure in the combustion chamber. The movement of the plungers can also be used for thrust vector control.

Abstract: An exhaust nozzle assembly (18) for a gas turbine engine (10), the assembly (18) comprises a main axis (20), an inner nozzle (19), an outer nozzle (17), a translatable centre-body (44) and a means for translating the centre-body (49) between a forward position and a rearward position. The centre-body (44) is disposed radially inwardly of the inner nozzle (19) thereby partly defining an inner duct (35) for a core engine gas flow (28), an outer duct (17) for a bypass gas flow (30) is defined by the inner nozzle (19) and the outer nozzle (17). The outer nozzle (17) extends downstream of the inner nozzle (19) and with the centre-body (44) defines a final mixing duct (36) having a final exhaust exit area (38). When the centre-body (44) is in the forward position the final exhaust exit area (38) is at a maximum area, and when the centre-body (44) is in a rearward position the final exhaust exit area (38) is at a minimum area.

Abstract: A variable geometry exhaust nozzle 10 for a turbine engine includes convergent and divergent flaps 12, 18 that circumscribe a gaspath 26 and define convergent and divergent nozzle sections 30, 32 with a throat 34 therebetween. A projection 56, which resides on the divergent flaps, extends radially inwardly from the nozzle. The nozzle also includes a coolant flowpath comprising an interior space 44 in the nozzle, an intake 58 for admitting coolant C into the interior space, and a coolant outlet 60 aft of the intake for discharging the coolant from the interior space. The intake resides forward of the throat. During operation, a coolant film flows along the radially inner surface of the nozzle. The projection encourages a portion of the coolant film to enter and flow through the the coolant flowpath to convectively cool the nozzle.

Abstract: A method for assembling a flap and seal system for a gas turbine engine exhaust nozzle including a plurality of backbone assemblies facilitates attaching a basesheet to a backbone. The method includes attaching an attachment system including at least one strap to a basesheet, and coupling the basesheet to a backbone using the attachment system strap.

Abstract: According to the present invention, the pointed end (11) of the central spike (7) consists of a deformable structure (13) that can adopt either a folded position or a deployed position.

Abstract: A rocket nozzle assembly is provided for operatively coupling to a rocket motor. The assembly includes a nozzle insert structure having a converging-diverging passageway with a throat region. First and second thrust control cylinders are parallel to each other and rotatable about respective axes transverse to the passageway central axis. The first and second thrust control cylinders intersect the throat region and include respective grooves. A thrust control cylinder-rotating subassembly is operatively associated with the first and second thrust control cylinders to rotate the cylinders about their respective axes and move the grooves relative to the throat region. Movement of the grooves changes the effective cross-sectional throat area of the converging/diverging passageway at the throat region for controlling the thrust when the rocket nozzle assembly is operatively engaged with an operating rocket motor.

Abstract: A center plug is configured for variable area use in an exhaust nozzle. A forward portion of the plug includes an inlet having a closing flap therein, and an aft portion of the plug includes an outlet. The flap is selectively open during aircraft takeoff operation and up to a preset climb altitude for bypassing a portion of core exhaust flow through the plug, with the flap being closed from preset altitude and cruise operation. Opening the flap correspondingly increases the available flow area for discharging the exhaust flow for reducing velocity thereof, and attenuating noise.

Abstract: A method for assembling a flap and seal system for a gas turbine engine exhaust nozzle including a plurality of backbone assemblies facilitates attaching a basesheet to a backbone. The method includes attaching an attachment system including at least one strap to a basesheet, and coupling the basesheet to a backbone using the attachment system strap.

Abstract: A method for assembling a flap system for a gas turbine engine exhaust nozzle including a plurality of backbone assemblies facilitates extending a useful life of the exhaust nozzle. The method includes providing a flap basesheet having a width defined between a pair of side edges that are coupled together by a leading edge and a trailing edge, and including at least one stiffener that extends between the basesheet side edges and includes an intermediate portion that has a width that is smaller than that of the basesheet and is at least one of bonded to and formed integrally with the basesheet, and coupling the basesheet to the gas turbine engine with a backbone assembly.

Abstract: A nozzle extension including flange elements of C/SiC is able to be manufactured relatively simply and provides sufficient physical properties. The carbon-fiber structure of the nozzle extension is wound, and a flange for fastening the nozzle extension to the nozzle outlet of the combustion chamber is integrated in the nozzle extension by a continuous fiber flow of the wound carbon-fiber structure. The foregoing may be intended for use in rocket propulsion units.

Abstract: The actuator assembly comprises a master actuator and at least one slave actuator having a cylinder containing a piston defining two chambers in the cylinder. A first chamber of the slave actuator is connected to an outlet of a directional control valve which is connected to a hydraulic control circuit and to a control member whose position determines the pressure in the first chamber of the slave actuator. A displacement of the piston of the master actuator its transferred to the control member of the control valve associated with the slave actuator by means of a mechanical transmission such as a flexible linear rolling bearing. Thus, the displacement of the piston of the slave actuator is servo-controlled to the displacement of the piston of the master actuator by controlling the position of the control member of the control valve associated with the slave actuator.

Abstract: Control apparatus for a nozzle for a rocket or other vehicle having a combustion chamber with a propellant therein in communication with the nozzle, the nozzle having a throat and a pair of bores extending into the throat The bores may be in opposed or any other suitable relation. A pair of plungers are slidably mounted in the bores and are movable between an open position wherein they are disposed outside of the throat and a closed position wherein they extend into the throat to substantially close it. The plungers have a width or diameter substantially the same as the width or diameter of the throat. Actuator devices are provided for moving the plungers independently of each other to control thrust by controlling the flow of combustion gases through the throat and the pressure in the combustion chamber. The movement of the plungers can also be used for thrust vector control.

Abstract: To provide auxiliary steam, a low pressure valve is opened in a combined cycle system to divert low pressure steam from the heat recovery steam generator to a header for supplying steam to a second combined cycle's steam turbine seals, sparging devices and cooling steam for the steam turbine if the steam turbine and gas turbine lie on a common shaft with the generator. Cooling steam is supplied the gas turbine in the combined cycle system from the high pressure steam turbine. Spent gas turbine cooling steam may augment the low pressure steam supplied to the header by opening a high pressure valve whereby high and low pressure steam flows are combined. An attemperator is used to reduce the temperature of the combined steam in response to auxiliary steam flows above a predetermined flow and a steam header temperature above a predetermined temperature.

Abstract: A gas turbine engine includes a variable area nozzle having a plurality of flaps. The flaps are actuated by a plurality of actuating mechanisms driven by shape memory alloy (SMA) actuators to vary fan exist nozzle area. The SMA actuator has a deformed shape in its martensitic state and a parent shape in its austenitic state. The SMA actuator is heated to transform from martensitic state to austenitic state generating a force output to actuate the flaps. The variable area nozzle also includes a plurality of return mechanisms deforming the SMA actuator when the SMA actuator is in its martensitic state.