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 coaxial injection device for injecting and dispersing reagents into a reactor, including an exterior duct for high-velocity gas injection; an outer-middle injector with at least one nozzle for liquid injection; an inner-middle duct for low-velocity gas injection; and an interior injector with nozzle for liquid injection; wherein, the exterior duct is formed by the internal wall of an insert and the external wall of the outer-middle injector; and is located externally to and circumferentially surrounds all other injectors and ducts; the outer-middle injector is formed by two concentric cylinders with end plate and injector nozzles; the inner-middle duct is formed by interior wall of the outer-middle injector and the exterior wall of the interior injector; the interior injector is formed by a cylinder with an endplate, the endplate having a nozzle; thereby ensuring the mixing and dispersion of the liquids and gases into the reactor to increase reaction homogeneity, reaction efficiency, reactor efficiency and

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 integrated, single piece mixer-center body ventilation apparatus is disclosed for use with a turbofan jet engine. The apparatus may incorporate a circumferential forward body portion adapted to be coupled to an aft end of a core engine turbine case of the jet engine, and a center body tube portion integrally formed with the forward body portion and having an axially opening vent exit. The forward body portion may have a plurality of inner mixer flow paths in communication with scalloped projecting portions. The inner mixer flow paths direct a pressurized core exhaust flow through the mixer device and mix the pressurized core exhaust flow with a portion of a pressurized fan exhaust flow, to thus significantly cool the pressurized core exhaust flow.

Abstract: The present invention relates to an aircraft gas-turbine engine with a core engine surrounded by a bypass duct, with a radially outer engine cowling enclosing the bypass duct and being provided at its rear region with a thrust-reversing device which is moveable relative to the engine cowling, with at least one cooler element extending over at least part of the circumference being arranged in the intermediate area between the engine cowling and the thrust-reversing device.

Abstract: One embodiment of the present invention is a unique flight vehicle. Another embodiment is a unique propulsion system. Another embodiment is a unique thrust vectoring system. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for flight vehicles, propulsion systems and thrust vectoring systems. Further embodiments, forms, features, aspects, benefits, and advantages of the present application will become apparent from the description and figures provided herewith.

Abstract: The invention relates to a method for controlling at least one actuator for actuating the cowlings of a thrust inverter in a turbojet engine, the actuator being driven by an electric motor including a relative position sensor providing information on the evolution of the movement thereof, wherein the motor is controlled based on the instantaneous position of the cowling in at least one portion of the movement thereof between an open position and a closed position, the instantaneous position of the cowling being determined from at least one reference position absolute data and relative position data relative to said reference position provided by the relative position sensor of the motor, wherein in case the actuation is resumed after an interruption, a new determination of the reference position is initiated.

Abstract: The present application provides a component for use in streams of hot gas, which comprises a first region of fiber-ceramic material, a second region of fiber-ceramic material and a middle layer of fiber-ceramic material which is arranged between the first region and the second region, wherein the fiber-ceramic material of the middle layer has a lower ceramic content than the fiber-ceramic material of the first region and the second region, and wherein at least one acute-angled in-flow edge and/or away-flow edge of the component is formed on the middle layer.

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: Track beams for the cowling of a jet engine for aircraft are formed of a base body with at least one slide for displaceably mounting a thrust reverser of the jet engine. A connection is provided for an inner engine cowling and fittings for detachably and pivotally attaching to a supporting structure or for connecting to another track beam. To provide such a track beam which has a particularly low weight yet, nevertheless, is sufficiently stiff, the base body is formed by a hollow profile section with a substantially closed cross-section. The hollow section is produced of a carbon-fiber-reinforced synthetic material by way of a resin-infusion process. Two flanges are provided on the hollow section of the base body of the track beam for providing a connection to the inner engine cowling.

Abstract: A variable area nozzle system for a gas turbine engine includes a fan duct inner wall, a fan duct outer wall disposed in radially spaced relation to the fan duct inner wall, and a fan nozzle. The fan nozzle defines at least a portion of the fan duct outer wall and includes a nozzle aft edge. The fan duct inner wall and the nozzle aft edge collectively define a fan duct nozzle throat area. The fan nozzle is configured to pivot about a pivot axis that may be oriented transversely relative to a longitudinal axis of the gas turbine engine. The fan nozzle may be pivoted from a stowed position to a deployed position in order to vary the fan duct nozzle throat area.

Abstract: A variable area nozzle system for a turbofan gas turbine engine comprises a fan duct inner wall, a fan duct outer wall and a fan nozzle. The fan duct outer wall is disposed in radially-spaced relation to the fan duct inner wall. The fan nozzle defines at least a portion of the fan duct outer wall and has a nozzle aft edge defining a fan duct throat area relative to the fan duct inner wall. The fan nozzle is configured to move outwardly relative to the longitudinal axis during axial aft translation thereof in order to vary the fan duct throat area.

Abstract: A 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 pivot between the first end and the second end. A linkage connects to the nozzle door and to an actuator. The actuator is selectively operative to move the linkage to in turn move the nozzle door about the pivot between a plurality of positions, such as a stowed position, an intermediate position, and a thrust reverse position, to influence a bypass airflow through a fan bypass passage.

Abstract: A transcowl for a gas turbine engine thrust reverser includes an arcuate outer wall; an arcuate inner wall; and an arcuate baffle positioned between the inner and outer walls at a forward end of the transcowl. The baffle has an arcuate cross-sectional shape which defines a forward-facing interior area. An arcuate forward seal is carried at a forward end of the inner wall. The forward seal includes a plurality of axially-extending, spaced-apart arcuate seal teeth which collectively define a labyrinth seal.

Abstract: A gas turbine engine system includes a nozzle having a plurality of positions for altering a discharge flow received through the nozzle from a gas turbine engine fan bypass passage. The nozzle is integrated with a thrust reverser having a stowed position and a deployed position to divert the discharge flow and generate a reverse thrust force. At least one actuator is coupled with the nozzle and the thrust reverser to selectively move the nozzle between the plurality of positions and to move the thrust reverser between the stowed position and the deployed position.

Abstract: A swept fan ramp for a pivot door thrust reverser includes a cylindrical unit with an elliptical flared portion, a side portion, and a rectangular flared portion. The elliptical flared portion is operable to reduce drag on airflow in a reverse direction and is coupled to the top center and bottom center of the circumference of the cylindrical unit. The side portion is operable to reduce side spillage airflow and is coupled to the central left side and central right side of the aft circumference of the cylindrical unit. The rectangular flared portion is operable to promote separation of airflow into an upper airflow path and a lower airflow path, is coupled with the aft circumference of the cylindrical unit, and is connected to the elliptical flared portion and the side portion.

Abstract: The exit area of a nozzle assembly is varied by translating a ring assembly located at a rear of the engine nacelle. The ring may be axially translatable along the axis of the engine. As the ring translates, the trailing edge of the ring defines a variable nozzle exit area. Translation of the ring creates an upstream exit at a leading edge of the ring assembly. The upstream exit can be used to bleed or otherwise spill flow excess from the engine bypass duct. As the engine operates in various flight conditions, the ring can be translated to obtain lower fan pressure ratios and thereby increase the efficiency of the engine. Fairings partially enclose actuator components for reduced drag.

Abstract: The thrust reverser is used with an aircraft gas turbine engine and includes upper and lower doors pivotable between a stowed position and a deployed position. When deployed, the doors redirect a portion of the efflux to generate a nose-down pitching moment on an aircraft to improve handling during thrust reversing on a runway.

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: An exhaust vane disposed in a divergent section of an aircraft gas turbine engine exhaust nozzle is sideways pivotable about a vane pivot axis. The nozzle vane pivot axis may be centrally located at an unvectored nozzle throat. Transversely spaced apart upper and lower tips of the exhaust vane may be incorporated to sealingly engage a nozzle outer wall along upper and lower surfaces of the outer wall of the nozzle. The exhaust vane has flat or contoured vane sidewalls and contoured vane sidewalls may be concave. The exhaust vane may have transversely biased apart upper and lower vane sections extending transversely inwardly from the upper and lower tips of the exhaust vane respectively. The exhaust vane may be articulated having upstream and downstream sections separately sideways pivotable about the vane pivot axis and a second pivot axis downstream of and parallel to the vane pivot axis respectively.

Abstract: A variable area flow duct and method. In one embodiment the flow duct includes a plurality of helical vanes arranged around an interior surface wall of the flow duct. The vanes cause secondary flow vortices to be developed in the vicinity of each of the vanes that effectively reduce the interior cross-sectional area of the duct that a primary flow sees as it flows through the duct. In various embodiments fluidic injection is employed to suppress the formation of the secondary flow vortices during certain phases of operation, for example, during an afterburn phase of operation of a jet aircraft engine which the flow duct is being used with. In another embodiment, an ablative coating is used over the vanes to suppress the formation of the secondary flow vortices. The ablative material is removed by the hot fluid flow during an afterburn phase of operation, thus exposing the vanes and enabling the subsequent formation of secondary flow vortices to narrow the cross-sectional area of the throat.

Abstract: A gas turbine engine system includes a first nozzle section associated with a gas turbine engine bypass passage and a second nozzle section that includes a plurality of positions relative to the first nozzle section. In at least one of the positions, there is a gap between the first nozzle section and the second nozzle section. A movable door between the first nozzle section and the second nozzle section selectively opens or closes the gap.

Abstract: An exhaust nozzle (10) for a gas turbine engine is described for vectoring a flow of exhaust gases issuing therefrom at an angle to a centre line (8) of the engine. The exhaust nozzle (10) comprises a plurality of radially outer flaps (25) and radially inner flaps (26) which are circumferentially disposed around the engine. The radially outer flaps (25) and radially inner flaps (26) alternately overlap each other to define a path through which the exhaust gases flow. The radially outer flaps (25) can be moved asymmetrically about the engine centre line (8). The radially inner flaps (26) are maintained in sealing contact with the outer flaps (25) by the flow of exhaust gases passing therethrough. The radially inner flaps (26) are divided into a number of triangular sections (29) which renders the flaps (26) torsionally flexible so that they can twist along their lengths to maintain sealing contact with adjacent radially outer flaps (25) when the outer flaps (25) are moved asymmetrically.

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

Abstract: A gas turbine engine air valve assembly has first and second valving elements. The second element is rotatable about a first axis relative to the first element. The rotation controls a flow of air through the first and second elements. An actuator is coupled by a linkage to the second element. The linkage includes a spindle having a first portion coupled to the actuator to rotate the spindle about a second axis. A guided spherical bearing couples a second portion of the spindle to the second element.

Abstract: A bi-directional locking ring assembly is provided comprising a housing and a gear having a plurality of teeth. The gear is coupled to the housing to rotate in opposite directions about a rotational axis. A pin is also coupled to the housing and has an apex configured to travel along a line that does not pass through the rotational axis of the gear. The pin is configured to translate between (1) a disengaged position wherein the gear is free to rotate, and (2) an engaged position wherein the apex resides between two of the plurality of teeth to prevent rotation of the gear in both rotational directions. A pin actuator is coupled to the housing and configured to engage the pin to selectively lock the pin in the engaged position.

Abstract: In order to provide a flying object with an engine comprising a combustion chamber with a combustion space and a nozzle with a nozzle space following the combustion chamber, the transition between combustion chamber and nozzle lying in a plane in which a cross-section of the combustion space converges, in which thrust control in terms of magnitude and/or thrust vector can be carried out in a simple manner, it is provided that a centre body adapted to have combustion products flowing around it is at least partially arranged in the nozzle space, outside of the combustion space, and at least three jet vanes are arranged at the centre body.

Abstract: An internal combustion engine exhaust gas system includes an EGR arrangement adapted to lead a first flow of exhaust gas, the EGR flow, from an outlet side of the engine to an inlet side of the engine, an exhaust gas conduit adapted to lead away a second flow of exhaust gas, the exhaust flow, from the outlet side of the engine, and an energy recovering unit, such as a turbo, associated with the exhaust gas conduit. The unit is adapted to recover exhaust gas energy from the exhaust flow. The system includes a heat exchanger adapted to allow heat exchange between at least a part of the EGR flow and at least a part of the exhaust flow, the heat exchanger being associated with the exhaust gas conduit at a position downstream the energy recovering unit.

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

Abstract: An outlet device for a jet engine includes an outlet nozzle and a plurality of guide vanes arranged movably in the outlet nozzle for guiding a gas from the jet engine for the purpose of steering a craft equipped with the jet engine. At least one of the guide vanes can be adjusted into such a position that hot parts of the engine located inside the outlet nozzle are at least substantially concealed seen from the outlet side of the outlet nozzle.

Abstract: A propulsive system for aircraft that has a jet engine and a nacelle surrounding the aforementioned jet engine. The nacelle has a front section fixed relative to the jet engine, and a rear extremity behind the front section, and has a rear section with two petals hinged relative to the front section. Each of the two petals in a position called zero-setting forms an extension of the front section of the nacelle in which the inside and outside surfaces of the petals and, respectively, determine the aerodynamic forms of the rear section, which are in geometric continuity with the inside and outside surfaces and, respectively, that determine the aerodynamic forms of the front section. The petals can be set individually at a positive or negative setting angle to put into play the reverser, APU, air brake, vectorial, and taxiing modes of the propulsive system.

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: The present application provides a component for use in streams of hot gas, which comprises a first region of fiber-ceramic material, a second region of fiber-ceramic material and a middle layer of fiber-ceramic material which is arranged between the first region and the second region, wherein the fiber-ceramic material of the middle layer has a lower ceramic content than the fiber-ceramic material of the first region and the second region, and wherein at least one acute-angled in-flow edge and/or away-flow edge of the component is formed on the middle layer.

Abstract: The thrust reverser is used with a gas turbine engine and includes first and second doors pivotable between a stowed position and a deployed position. When deployed, pivoting fairings on the first door are moved so as to give room to the second door as its trailing edge moves within the first door. When stowed, the second door preferably provides the locking mechanism to the pivoting fairings of the first door.

Abstract: Aircraft systems including cascade thrust reversers are disclosed herein. An aircraft system in accordance with one embodiment includes a cascade thrust reverser having a fixed reverser ramp and a nozzle outer wall section at least partially aft of the fixed reverser ramp. The nozzle outer wall section is movable between a deployed position and a stowed position. The nozzle outer wall section includes a forward portion with a leading edge section. The fixed reverser ramp has a portion forward of and adjacent to the nozzle outer wall section when the nozzle outer wall section is in the stowed position. The portion of the fixed reverser ramp has a first slope. The forward portion of the nozzle outer wall section that is aft of the leading edge section has a second slope different than the first slope.

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

Abstract: A device for controlling fluid flow. The device includes an arc generator coupled to electrodes. The electrodes are placed adjacent a fluid flowpath such that upon being energized by the arc generator, an arc filament plasma adjacent the electrodes is formed. In turn, this plasma forms a localized high temperature, high pressure perturbation in the adjacent fluid flowpath. The perturbations can be arranged to produce vortices, such as streamwise vortices, in the flowing fluid to control mixing and noise in such flows. The electrodes can further be arranged within a conduit configured to contain the flowing fluid such that when energized in a particular frequency and sequence, can excite flow instabilities in the flowing fluid. The placement of the electrodes is such that they are unobtrusive relative to the fluid flowpath being controlled.

Abstract: A low noise aircraft which can reduce engine noise on the ground sharply at the time of take-off and landing by utilizing a well-known mechanism. By utilizing a thrust deflection means constituted of a well-known mechanism for making the direction of thrust variable, exhaust direction of an engine is deflected to the upper side (upward) of the course direction (flight direction) of the aircraft so that the maximum propagation direction of the engine jet exhaust noise is kept away from the ground, thus reducing the engine noise level on the ground at the time of aircraft take-off and landing. In particular, when the thrust deflection type exhaust nozzle (20) is constituted of an upper deflection nozzle (21) and a lower deflection nozzle (22), the engine can generate reverse thrust during a landing run.

Abstract: A ducted air power plant, comprising a motor driven fan (7) situated in a duct (4), the fan (7) having an air intake side and in operation providing a high pressure air stream in the duct, and the fan being located adjacent air splitter mechanism (18), the air splitter mechanism (18) being arranged to divert the air stream into two or more subsidiary streams for delivery to respective jet nozzles (9) of the plant. The plant may be used in a vehicle such as an aircraft in order to provide a vertical take-off and hover capability as well a level flight power source.

Abstract: A swept fan ramp for a pivot door thrust reverser includes a cylindrical unit with an elliptical flared portion, a side portion, and a rectangular flared portion. The elliptical flared portion is operable to reduce drag on airflow in a reverse direction and is coupled to the top center and bottom center of the circumference of the cylindrical unit. The side portion is operable to reduce side spillage airflow and is coupled to the central left side and central right side of the aft circumference of the cylindrical unit. The rectangular flared portion is operable to promote separation of airflow into an upper airflow path and a lower airflow path, is coupled with the aft circumference of the cylindrical unit, and is connected to the elliptical flared portion and the side portion.

Abstract: The invention concerns a method for opening a mobile cowl (2) actuated by a motor (7) and equipping a turbojet thrust reverser, characterized in that it comprises, at the start of the opening phase, a bridling step (103) whereby the speed of the mobile cowl is maintained lower or equal to a predetermined threshold speed.

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: The invention relates to a device comprising gratings, which are used to form a reverse flow from a jet engine thrust flow. The inventive device comprises a plurality of flow diverter gratings which are disposed side-by-side at the outer edge of an annular thrust flow circulation channel. Each of the gratings comprises a plurality of intersecting inner transverse blades and inner longitudinal blades, two adjacent gratings being separated by a side clearance which generates a leakage flow. The device also comprises means of re-directing the aforementioned leakage flow in a direction such as to increase the effectiveness of the reverse flow, said re-direction means comprising aerodynamic appendages which are mounted to at least one outer longitudinal edge of each grating.

Abstract: A translating cowl composed of two sub-structures forms the thrust reverser for a turbofan engine. The two sub-structures form the rear part of a nacelle, are translatable, and have operative and inoperative modes of operation. The operative mode is used for direct or reverse thrust operation of the engine and the inoperative mode is used for access to the engine.

Abstract: A method and system for thrust vectoring a primary fluid flow from an exhaust nozzle of a jet engine that significantly increases the non-axial force able to be generated by a flight control surface associated with the nozzle. In one implementation the method involves placing a flight control element having a movable portion adjacent a downstream edge of the nozzle. A secondary fluid flow is created adjacent a surface of the flight control element that influences a boundary layer of the primary fluid flow over the flight control element. This causes the primary fluid flow to generate a force that is directed non-parallel (i.e., non-axial) to a longitudinal axis of the nozzle. In one specific implementation a plurality of slots are formed in the flight control surface, and the flight control surface is formed by an airfoil. In another implementation the flight control surface is formed on an interior wall of the nozzle at a downstream edge of the nozzle.

Abstract: The braking method for an aircraft propelled by at least one ducted fan jet engine, the latter comprising a fan, a jet engine core in which the primary flow circulates, a secondary flow exhaust nozzle, encased in a nacelle, comprises the step, in low-power operation, of unloading at least part of the secondary flow so as to reduce the residual thrust of the jet engine without producing any thrust. With the method of the invention, aircraft braking can be achieved without any thrust reverser device using the braking force generated by the drag of the jet engine or engines completed by reduction of residual thrust by unloading the secondary flow.

Abstract: An aircraft includes at least one turbofan engine assembly having a shrouded core engine, a short nacelle surrounding a fan and a forward portion of the core engine, and a fan exhaust duct through the nacelle. A mixer duct shell is positioned substantially coaxial with the engine shroud and extends forwardly into the fan duct to provide an interstitial mixer duct between the mixer duct shell and the core engine shroud. The aft portion of the mixer duct shell extends over a turbine exhaust frame, an attached mixer (if included), and a tail cone exhaust plug. The mixer duct shell can reduce noise and plume exhaust heat radiated from aircraft turbofan engines.

Abstract: A nacelle structure for a gas turbine engine includes an inlet aft of an exit guide vane for supplying air flow to a forward plenum. Air is exhausted from the plenum into an air intake opening and over an inner surface of the nacelle for controlling air flow characteristics within the nacelle.

Abstract: A retractable thrust vector control system (10) for a rocket motor (26) that can generate an exhaust plume comprises at least one control vane (12) connectable to an attitude control assembly (20) that rotates the vane (12) about a control axis (44). The system also includes a retraction mechanism (14) for withdrawing the control vane (12) along the control axis (44) from an extended position at least partially within a path of a rocket exhaust plume and a retracted position substantially out of a path of a rocket exhaust plume.

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.