Abstract: Secondary air flow is provided for a ducted fan having a reverse flow turbine engine core driving a fan blisk. The fan blisk incorporates a set of thrust fan blades extending from an outer hub and a set of integral secondary flow blades extending intermediate an inner hub and the outer hub. A nacelle provides an outer flow duct for the thrust fan blades and a secondary flow duct carries flow from the integral secondary flow blades as cooling air for components of the reverse flow turbine engine.

Abstract: A bypass aeronautical engine including a fan, an LP compressor, an HP compressor, a combustion chamber, an HP turbine, and at least one LP turbine stage is disclosed. The LP turbine stage includes a first movable bladed wheel carried by a first turbine shaft and a second movable bladed wheel turning in the opposite direction to the first wheel. The fan is driven by the second wheel via a second turbine shaft and by the first shaft via a module for reversing its direction of rotation. The reversing module includes two bevel pinions carried respectively by the first and second LP turbine shafts, and a reversing pinion.

Abstract: Secondary air flow is provided for a ducted fan having an engine core driving a fan blisk. The fan blisk incorporates a set of thrust fan blades extending from an outer hub and a set of integral secondary flow blades extending intermediate an inner hub and the outer hub. A nacelle provides a first flow duct for the thrust fan blades and a secondary flow duct carries flow from the integral secondary flow blades.

Abstract: An example nacelle assembly for a gas turbine engine includes a nacelle defined about an axis and defining a boundary of a fan bypass flow path. A fan variable area nozzle includes a plurality of inserts movably mounted to said nacelle. Each of the multiple of inserts is located at a circumferential position about the nacelle. The multiple of inserts are each independently moveable into the fan bypass flow path relative the nacelle to selectively vary a fan nozzle exit area.

Abstract: An air vehicle propulsion system incorporates an engine core with a power shaft to drive an outer blade row. The power shaft extends through and is supported by a counter rotation transmission unit which drives an inner blade row in counter rotational motion to the outer blade row. The counter rotation transmission unit exchanges power from the engine core with the shaft. An actuator engages the shaft for translation from a first retracted position to a second extended position.

Abstract: A scoop inlet for use in a gas turbine engine nacelle has a scoop inlet, and a tab extending forwardly of the scoop inlet. The scoop communicates with a downstream flowpath. The tab has at least one opening at a location upstream of the scoop inlet. A nacelle and a gas turbine engine are also disclosed.

Abstract: A turbofan engine includes a core nacelle housing a spool having a turbine. A turbofan is arranged upstream from the core nacelle and is coupled to the spool. A fan nacelle surrounds the turbofan and core nacelle. A bifurcation supports the core nacelle relative to the fan nacelle and extends radially between the nacelles. The core and fan nacelles provide a bypass flow path receiving bypass flow from the turbofan. In the example, the nacelles provide a fixed nozzle exit area. The bifurcation includes an inlet that is in communication with a passage. The passage is selectively opened and closed by a valve to provide air from the inlet to a vent opening near the nozzle exit area. By opening the valve, the nozzle exit area is effectively increased to provide increased bypass flow through the turbine engine.

Abstract: A noise reduction assembly located at a junction plane between an air intake structure and a fan casing of an aircraft turbojet has improved acoustic properties since the alternation of acoustic and non-acoustic zones is displaced or eliminated. In particular, the inner wall of the air intake structure includes a projecting part that extends towards the aft direction beyond the junction plane, in other words the flange of the fan casing. The projecting part may be sufficiently long so that its inner skin is continuous as far as the limit between the acoustic zone and the non-acoustic zone of the nacelle.

Abstract: An exemplary gas turbine engine assembly includes a fan casing within a nacelle, a variable area fan nozzle. A controller is operable to move the variable area fan nozzle to influence a discharge airflow area associated with the variable area fan nozzle in response to an airfoil flutter condition. A gear train reduces a rotational speed of a fan in the gas turbine engine relative to another portion of the gas turbine engine.

Abstract: Systems and methods involving multiple torque paths of gas turbine engines are provided. In this regard, a representative method for reducing overspeed potential of a turbine of a gas turbine engine includes: providing a first load to the turbine via a first torque path; providing a second load to the turbine via a second torque path; and operating the turbine such that: mechanical failure of a component defining at least a portion of the first torque path does not inhibit the second load from being applied to the turbine via the second torque path; and mechanical failure of a component defining the second torque path does not inhibit the first load from being applied to the turbine via the first torque path.

Abstract: A gas turbine engine oil system has an air-oil separator for removing air from an air/oil mixture. A breather tube is connected to an exhaust of the air-oil separator for receiving hot air removed from the air/oil mixture in the air-oil separator. The gas turbine engine oil separator exhaust is directed in a cooled oil collector to cause the oil mist remaining in the air at the exit from the engine air-oil separator to condensate. The oil condensate is returned back into the engine oil system.

Abstract: A coupling assembly for an engine core of a turbomachine is provided. The coupling assembly comprises a duct having a substantially circular shape and surrounding the engine core, and a plurality of coupling struts, each coupling strut extending only in a substantially vertical direction relative to the direction of gravity when the engine core is at a rest position with zero pitch and zero roll, each of the plurality of coupling struts coupling the engine core to the duct.

Abstract: An example auxiliary power unit (APU) exhaust silencer includes cooling features to protect the outer skin and other components from heat generated by gases passing through an exhaust duct. Cooling air flow through a cooling air passage in thermal contact with the exhaust silencer carries heat away from other nearby components and the aircraft skin.

Abstract: A gas turbine engine includes a spool, a turbine coupled with the spool, a propulsor coupled to be rotated about an axis by the turbine through the spool and a gear assembly coupled between the propulsor and the spool such that rotation of the spool results in rotation of the propulsor at a different speed than the spool. The propulsor includes a hub and a row of propulsor blades that extends from the hub. Each of the propulsor blades has a span between a root at the hub and a tip, and a chord between a leading edge and a trailing edge such that the chord forms a stagger angle ? with the axis. The stagger angle ? is less than 62° at all positions along the span, with said hub being at 0% of the span and the tip being at 100% of the span.

Abstract: An airfoil includes an airfoil body that extends between a leading edge and a trailing edge, a suction side and a pressure side, and a first end and a second end. At least one fitting is located on at least one of the first end and the second end. The fitting includes at least one mounting lug. At least one shroud is adhesively bonded to the at least one fitting.

Abstract: An improved bypass turbojet engine includes a particularly shaped hot flow nozzle arranged about a particularly shaped type of boss. The boss is bulbous in shape and is connected by a first connecting surface to a turbine casing. The hot flow nozzle is formed as a single skin that extends in a rearward direction from an engine cowl, with the hot flow nozzle being connected to the engine cowl along a second connecting surface. The hot flow nozzle and the bulbous boss are positioned relative to one another to delimit a nominal nozzle throat section and a nominal outlet section for hot flow from a hot flow generator. The hot flow nozzle single skin is also approximately bulbous in shape and widens from the second connecting surface in a direction toward the engine cowl's external wall.

Abstract: Setting in place of one or several coolers in the wall of a secondary flow of a bypass turbomachine. The wall extends from an intermediate casing toward a leading edge of a separator nose between a primary flow and the secondary flow. The wall includes a series of support arms attached to an intermediate casing, distributed over the perimeter of the wall and directed upstream. A series of surface air-oil heat exchangers forming wall segments are arranged end-to-end on the support arms, so as to form an annular wall. A shroud having a leading edge is arranged and fixed in the area of the upstream edges of the heat exchangers, so as to complete the wall. The support arms include hydraulic connectors connected to one another on each arm, adapted to cooperate with corresponding connectors in the area of the heat exchangers and in the area of the intermediate casing.

Abstract: A control system of a hydraulic type for a thrust reversal device with which a jet engine nacelle is equipped, and associated with a nozzle variation device, is provided by the present disclosure. The control system includes at least one additional control unit hydraulically connecting actuators together, wherein hydraulic supply lines from the control unit in a nozzle mode and from the control unit in a reverser mode are then able to communicate hydraulically with each other. A plurality of control valves are laid out so that when the unit for driving the actuators in a nozzle mode is active, a corresponding hydraulic signal controls the valves of the control unit in a position such that the hydraulic supply lines of the actuators from the driving unit according to a reverser mode are cut off.

Abstract: A nacelle, for an aircraft jet engine having a high bypass ratio, in which a jet engine having a longitudinal axis is mounted. The nacelle includes a wall that concentrically, and at least partially, surrounds the jet engine and together defines an annular duct. The nacelle further includes a displacement device that controllably displaces a portion of the nacelle wall in order to modify the section of the flow outlet passage and form at least one longitudinally extending opening. A device for forming a fluid barrier along the at least one longitudinally extending opening is provided to counteract the natural exhaust.

Abstract: A hybrid propulsive technique, comprises providing at least some first thrust associated with a flow of a working fluid through at least a portion of an at least one axial flow jet engine. The hybrid propulsive technique includes extracting energy at least partially in the form of electrical power from the working fluid, and converting at least a portion of the electrical power to torque. The hybrid propulsive technique further includes rotating an at least one substantially axial-flow independently rotatable compressor rotor at least partially responsive to the converting the at least a portion of the electrical power to torque.

Abstract: A gas turbine engine has a compressor section, a low spool, and a fan. The fan delivers air into the compressor section and into a bypass duct having a variable area nozzle. The compressor section compresses air and delivers it into a combustion section. The combustion section mixes air with fuel, igniting the fuel, and driving the products of the combustion across a turbine. The turbine drives the low spool. A control for the gas turbine engine is programmed to position the nozzle at startup of the engine to increase airflow across the fan. A variable inlet guide vane is positioned upstream of the compressor section. The control also positions the variable inlet guide vane at start-up to increase air flow across the compressor section.

Abstract: A fan control apparatus includes a fan, two engine+compressor combinations, two air supply systems, and an FCC. When an abnormality occurs in one of the air supply systems and one of the engine+compressor combinations, the FCC maintains the flow rate of the normally operating air supply system and then increases the flow rate. As a result, the normally operating drive source is prevented from overloading. In another embodiment, a fan control apparatus includes a fan, an air source, two air supply systems, and a bypass channel. The air is caused to flow through the bypass channel when an abnormality occurs in one of the air supply systems. As a result, the time that elapses till the fluid can be supplied at a necessary flow rate is shortened.

Abstract: Gas turbine engine systems involving tip fans are provided. In this regard, a representative gas turbine engine system includes: a multi-stage fan having a first rotatable set of blades and a second counter-rotatable set of blades, the first rotatable set of blades defining an inner fan and a tip fan; and an epicyclic differential gear assembly operative to receive a torque input and differentially apply the torque input to the first set of blades and the second set of blades.

Abstract: A combined engine includes a turbopump including a pump injecting hydrogen into a heater arranged in an outer casing downstream from a central body, and a subsonic turbine driving the pump, which turbine receives partially-expanded hydrogen collected at an outlet from the heater to apply the hydrogen to a supersonic turbine to operate the engine as a turbojet. The hydrogen from the supersonic turbine is collected in tubes inside the central body to be sent to a combustion chamber defined downstream from the central body, while the hydrogen that is partially expanded in the subsonic turbine is sent directly to the combustion chamber via injectors to operate the engine as a ramjet.

Abstract: A composite component such as a fan casing liner panel for a gas turbine engine comprises an abradable layer and a layer. The layer is provided with a septum layer, to which the abradable layer is bonded. The septum layer is perforated by holes, which enable air to be extracted from the region between the septum layer and the abradable layer during the process of bonding the components together, for example by a vacuum bonding process.

Abstract: The present inventions include a boundary layer ejector fluidically connecting boundary layer bleed slots from an external surface of an aircraft to reduce aircraft/nacelle/pylon drag, reduce jet noise and decrease thrust specific fuel consumption. In one embodiment a boundary layer withdrawn through the boundary layer bleed slots is entrained with an exhaust flow of a gas turbine engine. In another embodiment a boundary layer withdrawn through the boundary layer bleed slots is entrained with a flow stream internal to the gas turbine engine, such as a fan stream of a turbofan. Members can be provided near an outlet of a passageway conveying the withdrawn boundary layer air to locally reduce the pressure of the fluid in which the withdrawn boundary layer air is to be entrained. A lobed mixer can be used in some embodiments to effect mixing between the boundary layer and a primary fluid of the ejector.

Abstract: A cam-type apparatus is included in a support link between outer and inner cases of gas turbine engines for centering the cases one to another. The cam-type apparatus is lockable to allow locking an adjusted position of the cam-type apparatus.

Abstract: A gas turbine engine (10) provides a differential gear system (58) coupling the turbine (20) to the bypass fan (14) and the compressor (16). In this manner, the power/speed split between the bypass fan and the compressor can be optimized under all conditions. In the example shown, the turbine drives a sun gear (74), which drives a planet carrier (78) and a ring gear (80) in a differential manner. One of the planet carrier and the ring gear is coupled to the bypass fan, while the other is coupled to the compressor.

Abstract: The cross sectional flow area of a fan discharge nozzle on one side of a central plane of an associated gas turbine engine power plant is greater than the corresponding flow area of the fan discharge nozzle on an opposite side of the central plane to compensate for the blockage of fan airflow by a pylon.

Abstract: A control device for controlling a variable section nozzle of an aircraft power plant, the variable section nozzle including one or several movable parts capable of modifying the nozzle section and connected by a mechanical transmission chain to an actuator. The control device includes a system for regulating the power plant connected to a control member configured to control the actuator. The control device includes a single control member, an immobilization unit configured to immobilize all the movable parts which are deactivated only when the regulation system controls the positional change of the movable part or parts and a determination unit configured to determine the actual position of the movable part or parts.

Abstract: A nacelle assembly for a bypass gas turbine engine includes a variable area fan nozzle having a first fan nacelle section and a second fan nacelle section. The variable area fan nozzle is in communication with a fan bypass flow path, the first fan nacelle section defines an intermittent trailing edge which defines a multiple of ports and the second fan nacelle section defines a multiple of doors, each of the multiple of doors match each of the multiple of ports such that a fan nacelle trailing edge is continuous when the second fan nacelle section is selectively translated to a closed position relative to the first fan nacelle section.

Abstract: A gas turbine engine has a fan at an axially outer location. The fan rotates about an axis of rotation. The fan delivers air into an outer bypass duct, and across a booster fan positioned radially inwardly of the outer bypass duct. The booster fan delivers air into a radially middle duct, and across a cold turbine into a radially inner core duct being directed into a compressor. From the compressor, air flows axially in a direction back toward the fan through a combustor section, and across an exhaust of the turbine section as directed into the middle duct. A gear reduction drives the fan from a fan drive turbine section. A method is also disclosed.

Abstract: A heat exchange system for use in fluid operated equipment to provide air and working fluid heat exchanges to cool the working fluid in airstreams on a stream side of a wall. An actuator is mounted to be substantially located on a side of the wall opposite the stream side thereof having a positionable motion effector. A heat exchanger core having a plurality of passageway structures therein to enable providing the working fluid to, and removal therefrom. The heat exchanger core is mounted on the motion effector so as to be extendable and retractable thereby through the opening for selected distances into that region to be occupied by the airstreams.

Abstract: A dual clutch arrangement is disclosed herein. The dual clutch arrangement includes a first input rotatable member. The dual clutch arrangement also includes an output rotatable member. The dual clutch arrangement also includes a first clutch coupling the first input rotatable member and the output rotatable member such that the first input rotatable member drives the output rotatable member in rotation. The first clutch is operable to be overrun. The dual clutch arrangement also includes a second input rotatable member. The dual clutch arrangement also includes a second clutch operable to selectively couple the second input rotatable member and the output rotatable member such that the second input rotatable member drives the output rotatable member in rotation. The first clutch is overrun when the second clutch is engaged.

Abstract: A turbine engine for aircraft includes a gas generator as well as to a receiver driven by a power turbine, and to an electric power generator, the rotor of which centered on a rotor axis distinct from the longitudinal axis, is driven into rotation by the power turbine, via a mechanical gear transmission, including a toothed drive wheel centered on the axis. Further the receiver includes a first propeller driven by a free power turbine and the toothed drive wheel is attached on an outer case of this power turbine.

Abstract: The turbojet engine inlet and exhaust covers are permanently installed at the inlet and exhaust outlet or nozzle of an aircraft turbojet engine installation. The covers preclude the entrance of small animals, windblown debris, snow, and/or other contaminants into the engine inlet or exhaust while the airplane is inoperative. The covers have an iris-type shutter configuration, each cover having a plurality of leaves pivotally attached to a circumferential stationary housing that is attached to the engine or nacelle. A rotary ring is rotated through a small arc to pivot the shutter leaves inward and outward to close and open the engine opening. The rotary ring may be driven by an electric motor, or by hydraulic or pneumatic power. The system may be automated to close according to engine temperature after flight, and/or to open upon initiation of the engine start sequence for the aircraft.

Abstract: A method and a system for detecting the ingestion of an object by an aircraft turbine engine during a mission is disclosed. The method includes: acquiring during the mission digital images of the operating fan of the turbine engine, these images being acquired at an acquisition frequency proportional to the speed of rotation of the fan and to the number of blades of the fan; identifying the different phases of the mission of the aircraft for each phase of the mission; comparing the images of the fan with at least one reference image corresponding to sound operation of the fan; and, if necessary, identifying each abnormal image of the fan which differs from the corresponding reference image, the identification of an abnormal image of the fan corresponding to the crossing of an alert level for detecting the ingestion of an object by the turbine engine.

Abstract: This disclosure relates to a gas turbine engine including a core engine that drives a fan. A fan nacelle surrounds the fan and provides an inner surface facing the fan. The fan nacelle includes an inlet lip adjoining the inner surface. A ring is arranged at the inner surface and is axially upstream from the fan. The ring includes a deployed position in which the ring is spaced radially inwardly from the inner surface to provide a gap therebetween through which air flows. The ring also includes a stowed position in which the ring is arranged radially inwardly from the deployed position and proximate to the inner surface. The ring is commanded to the deployed position in response to detecting a first condition, which corresponds to a flow separation condition at the fan nacelle inlet. The ring may be stowed during cruise conditions, for example.

Abstract: A propulsion system includes an engine with a fan and a nacelle that circumscribes the engine. The engine defines an axis. A distance between a forwardmost point at the top of the nacelle and at the front of the nacelle to a forwardmost point of the fan is greater than a distance between a forwardmost point at the bottom of the nacelle and at the front of the nacelle.

Abstract: A system and method are provided for controlling the temperature of engine bleed air from a turbofan gas turbine engine. The system includes a fan air valve and a fan air valve controller. The fan air valve is adapted to receive a flow of fan air from a turbofan gas turbine engine intake fan. The fan air valve is coupled to receive valve position commands and is configured, in response to the valve position commands, to move to a valve position to thereby control the engine bleed air temperature. The fan air valve controller is configured to implement a linear quadratic regulator (LQR) control. The fan air valve controller is adapted to receive a plurality of sensor signals, each sensor signal representative of one or more system parameters, and is configured, in response to the sensor signals, to supply the valve position commands to the fan air valve.

Abstract: A gas turbine engine includes a shaft defining an axis of rotation. An outer turbine rotor directly drives the shaft and includes an outer set of blades. An inner turbine rotor has an inner set of blades interspersed with the outer set of blades. The inner turbine rotor is configured to rotate in an opposite direction about the axis of rotation from the outer turbine rotor. A gear system couples the inner turbine rotor to the shaft and is configured to rotate the inner set of blades at a faster speed than the outer set of blades. The gear system is mounted to a mid-turbine frame.

Abstract: A turbofan gas turbine engine comprises a nacelle cowl and a core engine. A bypass duct is between an outer surface of a casing of the core engine, and an inner surface of the nacelle cowl. An air channel is in the nacelle cowl, an inlet and an outlet of the air channel being in an outer surface of the nacelle cowl. An oil cooler has at least one oil passage for oil circulation, the air cooler having a first heat exchange surface in the air channel exposed to air circulating in the air channel, the air channel having a second heat exchange surface in the bypass duct exposed to air circulating in the bypass duct. A method for cooling oil in a turbofan gas turbine engine is also provided.

Abstract: A compressor section includes a counter rotating low pressure compressor that includes outer and inner compressor blades interspersed with one another and are configured to rotate in an opposite direction than one another about an axis of rotation. A transmission couples at least one of the outer and inner compressor blades to a shaft. A turbine section includes a counter rotating low pressure turbine having an outer rotor that includes an outer set of turbine blades. An inner rotor has an inner set of turbine blades interspersed with the outer set of turbine blades. The outer rotor is configured to rotate in an opposite direction about the axis of rotation from the inner rotor. A gear system couples at least one of the outer and inner rotors to the shaft.

Abstract: A bleed valve bleeds hot gases from a compressor into a bypass duct of a gas turbine engine. The valve comprises a diffuser having opposing walls and a divider located between the walls and which define at least two passages through which a bleed fluid flows and into a fluid flow through the bypass duct. The passages are angled towards each other to force the two gas flows together to form a high aspect ratio plume. This plume has a relatively large surface area that that improves mixing with the bypass flow to cool it and prevented otherwise hot bleed flows from impinging on heat sensitive components adjacent the bypass duct.

Abstract: A gas ejection cone for an aircraft turbojet, the cone including a hollow main body defining, on the outside, a radially inner skin of an annular primary flow channel, and a device generating turbulence in the primary flow limiting jet noise, mounted so as to move on the main body so as to be able to be displaced from an extracted position in which the device projects toward downstream in relation to a downstream end of the hollow main body, and a retracted position in which the device is retracted into the hollow main body, and vice versa. Further, the device includes a cylindrical support body having an axis parallel to an axis of the ejection cone, and at least one fin supported by the body.

Abstract: A method of assembling a gas turbine engine is provided. The method includes providing at least one heat exchanger assembly including a heat exchanger and a mixer extending downstream from the heat exchanger. The mixer includes a plurality of windows formed therein. The method also includes coupling the at least one heat exchanger assembly within a bypass duct of the engine such that the at least one heat exchanger assembly is coupled to at least one of an outer engine casing and an inner engine casing of the turbine engine.

Abstract: An epicyclic gear train includes planetary gears rotatably mounted on spindles supported by an annular carrier and including axially spaced apart forward and aft sets of output teeth extending radially outwardly from a planetary gear hub and axially spaced apart forward and aft roller bearings disposed between planetary gears and spindles. The forward and aft roller bearings are axially aligned with or adjacent to spaced apart forward set of output teeth and input gear respectively. A ring gear meshes with forward set of output teeth and an external gear meshes with aft set of output teeth. An input gear fixedly attached to hub aft of aft set of output teeth and engaged with a sun gear. The output teeth, input gear, ring gear, external gear, and sun gear may all be helical. A turbofan gas turbine engine may include counter-rotatable first and second fan stages driven by a low pressure turbine through the gear train.

Abstract: A gas turbine engine includes a low spool along an engine axis with a forward fan section and a low pressure turbine section. A high spool along the engine axis with a high pressure turbine section and an aft core driven fan section (CDFS), the aft core driven fan section (CDFS) axially aft of the high pressure turbine section along the engine axis.

Abstract: A gas turbine engine includes a core housing that has an inlet case and an intermediate case that respectively provide an inlet case flow path and an intermediate case flow path. The shaft supports a compressor section that is arranged axially between the inlet case flow path and the intermediate case flow path. The shaft includes a main shaft and a flex shaft having bellows. The flex shaft is secured to the main shaft at a first end and includes a second end opposite the first end. A geared architecture is coupled to the shaft, and a fan coupled to and rotationally driven by the geared architecture. The geared architecture includes a sun gear supported on the second end. A first bearing supports the shaft relative to the intermediate case and a second bearing supports the shaft relative to the inlet case. The second bearing is arranged radially outward from the flex shaft.

Abstract: A gas turbine engine includes a propulsion unit mounted to rotate about a first axis, and a core engine mounted to rotate about a second axis, and wherein the first and second axes are non-parallel. A gas turbine engine includes a propulsion unit driven by a free turbine which is adjacent to the propulsion unit and an associated fan, and having a gas generator core engine including a compressor, combustor and turbine section. A method is also disclosed.