Abstract: A gas turbine engine according to an exemplary aspect of the present disclosure includes a windmill pump driven by a spool. A first pump driven by said spool with an air-oil cooler is located downstream of the first pump. A second pump is also driven the spool with an air-air precooler located downstream of the second pump. A method of operating a gas turbine engine during a “windmilling” condition includes driving a windmill pump with a spool during a “windmilling ” condition. A lubricant is communicated to a geared architecture with the windmill pump. A first pump is driven by the spool and an air-oil cooler is located downstream of the first pump.

Abstract: A gas turbine engine has a rear mounting structure incorporating a mounting apparatus attached to a bypass duct wall with a link device of six rods for transferring core portion related inertia-induced loads, from an inner case of the core portion in a short circuit across an annular bypass air passage to the bypass duct wall.

Abstract: A gas turbine engine includes a first and second pump driven by a spool. An Air-Oil Cooler downstream of the first pump. An air-air precooler is downstream of the second pump, the air-air precooler downstream of the Air-Oil Cooler.

Abstract: An oil management system includes an engine housing assembly which defines a first rotor volume and a second rotor volume. An oil cooler assembly arranged between the first rotor volume and the second rotor volume.

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

Abstract: A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a fan section including a fan, a geared architecture, a spool which drives the fan through the geared architecture, and at least one component geared to the spool and being operable to control a speed of the spool during a “windmilling” condition. A method of gas turbine engine operations during a “windmilling” condition is also disclosed.

Abstract: A gas turbine engine includes a constant speed transmission driven by a spool. The constant speed transmission drives a first pump and a second pump. In a further non-limiting example, the first pump is driven by a first drive shaft which is driven by the constant speed transmission and the second pump is driven by a second drive shaft which is driven by the first shaft through a gearbox to provide a constant speed ratio between the first shaft and the second shaft.

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

Abstract: Disclosed is a turbofan engine for an aircraft. An inner fan cowl of the turbofan engine bounds an annular cold-stream duct proximate to the engine's central hot-stream generator, and an outer fan cowl bounds the annular cold-stream duct proximate to an outer nacelle cowl. An annular boss is provided along the inner fan cowl, with the annular boss having a rounded cross section that projects with respect to a smooth shape configuration of the inner fan cowl. The annular boss is configured to locally change speed of a cold stream flow in the annular cold-stream duct from a subsonic range to a supersonic range and to originate a first shockwave characteristic in a succession of shockwave characteristics, with the first shockwave characteristic being inclined from its origin toward the rear portion of the turbofan engine.

Abstract: A gas turbine engine flow duct comprising a flow duct disposed along an engine centerline of the gas turbine engine and defining a stream flow passage, and first and second rows of heat exchangers disposed along the engine centerline of the gas turbine engine and integrated in the flow duct in fluid communication with the stream flow passage of the flow duct.

Abstract: A separate propulsion unit incorporating a free turbine and a fan receives gases from a plurality of core engines. The core engines each include a compressor, a turbine and a combustion section. The core engines in combination pass gases across the free turbine. A method is also disclosed.

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

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

Abstract: An engine comprises a gas generator having an exhaust plenum and a propulsor comprising a propulsion fan coaxially mounted within an annular turbine. The annular turbine comprises a turbine duct and a plurality of turbine rotor blades rotationally coupled to the propulsion fan. A plurality of hollow struts extend axially and radially from the gas generator to the annular turbine. The hollow struts comprise flow ducts connecting the exhaust plenum to the turbine duct.

Abstract: A gas turbine engine comprises a high spool, a low spool and an intermediate spool. The high spool comprises a high pressure turbine coupled to a high pressure compressor. The intermediate spool comprises an intermediate pressure turbine coupled to a ducted fan. The low spool comprises a low pressure turbine coupled to an open-rotor propeller. A variable area turbine section positioned between the intermediate pressure turbine and the low pressure turbine variable turbine section is configured to vary an expansion ratio across the intermediate pressure turbine to control rotational speeds of the low spool and the intermediate spool.

Abstract: The invention relates to a rear section (11) of an aircraft nacelle, that comprises two halves (13a, 13b) defining: a central portion (C) for receiving a turbojet engine (7), a cool-air annular passage (31) provided around said central portion (C), and at least one six-hour cavity (15) provided under said central portion (C). The rear section is characterized in that it comprises at least one duct (29a, 29b) for the fluidic communication between said annular passage (31) and said six-hour cavity (15) for maintaining the temperature inside the six-hour cavity (15) within a relatively low range.

Abstract: A gas-turbine engine with at least one compressor and at least one bleed-air tapping device, which includes an annular duct in a radially outer wall of a flow duct, and with an annular closing element, which is arranged in the region of the annular duct and can be moved in a substantially axial direction from a closed position to an open position, with the closing element having an annular flow divider projection which in the open position projects in the flow duct.

Abstract: An example gas turbine engine includes, among other things, a geared architecture rotatably coupled to the fan drive shaft, and a high pressure compressor. The gas turbine engine is configured so that a core temperature at an exit of the high-pressure compressor is approximately in a range of about 1150 to about 1350 degrees Fahrenheit at take-off. The gas turbine engine is configured so that an Exhaust Velocity Ratio, defined by a ratio of a fan stream exhaust velocity to a primary stream exhaust velocity, is approximately in a range of about 0.75 to about 0.90. A Bypass Ratio of the engine is greater than about 8.0.

Abstract: A turbine engine includes a fan shaft and at least one tapered bearing mounted on the fan shaft. The fan shaft includes at least one radially extending passage adjacent the at least one tapered bearing. A fan is mounted for rotation on the tapered bearing. An epicyclic gear train is coupled to drive the fan. The epicyclic gear train includes a carrier that supports star gears that mesh with a sun gear, and a ring gear that surrounds and meshes with the star gears. Each of the star gears is supported on a respective journal bearing. The epicyclic gear train defines a gear reduction ratio of greater than or equal to about 2.3.

Abstract: A gas turbine engine has a variable overall pressure rate (“OPR”). The engine includes a high pressure compressor having at least a primary stage having a set of primary rotors and a secondary stage having a set of secondary rotors. A clutch is provided to selectively engage the secondary rotors with the primary rotors. Engagement of the clutch may be controlled based on the vehicle travel mode, such as disengaging during a takeoff mode to reduce turbine entry temperature and engaging during a loiter mode to increase OPR.

Abstract: A gas turbine engine typically includes a fan section, a compressor section, a combustor section and a turbine section. A speed reduction device such as an epicyclical gear assembly may be utilized to drive the fan section such that the fan section may rotate at a speed different than the turbine section so as to increase the overall propulsive efficiency of the engine. In such engine architectures, a shaft driven by one of the turbine sections provides an input to the epicyclical gear assembly that drives the fan section at a speed different than the turbine section such that both the turbine section and the fan section can rotate at closer to optimal speeds providing increased performance attributes and performance by desirable combinations of the disclosed features of the various components of the described and disclosed gas turbine engine.

Abstract: A gas turbine engine includes an intercooling turbine section to at least partially drive one of a low spool and a high spool. An intercooling turbine section bypass to selectively bypass at least a portion of a core flow through an intercooling turbine section bypass path around the intercooling turbine section.

Abstract: A turbofan engine includes a core engine, having a high-pressure compressor, a combustion chamber and a high-pressure turbine which are coupled to one another via a high-pressure shaft, at least one fan from which gas is supplied into both a primary flow duct and a secondary flow duct of the turbofan engine, at least one low-pressure turbine arranged behind the core engine, and at least one low-pressure shaft, with each low-pressure shaft coupling a fan to a low-pressure turbine. It has been provided that no low-pressure shaft of the turbofan engine passes through the core engine.

Abstract: A turbine engine includes at least a compressor section and a turbine section, each having at least a first and second portion. A ratio of turbine section second portion stages to compressor section second portion stages is less than or equal to 1.

Abstract: A gas turbine engine having a rotational axis, an intake and a compressed gas source; the intake includes a lining having a facing which defines an inlet surface and an array of holes; the array of holes includes at least a first set of holes and a second set of holes, the holes of the first set of holes are angled a relative to a radial line and the holes of the second set of holes are angled ? relative to the radial line; the angles ? and ? are different. An active flow control arrangement including a compressed gas supply pipe, a valve arrangement, a controller, a compressed gas distribution pipe may be provided. Compressed gas may be provided to prevent the formation of separation of a main gas flow through the intake and prevent or remove ice accretion.

Abstract: The invention relates to a guiding device (100) for part of a jet engine nacelle that can move in translation, comprising: (i) a rail (101) fixedly mounted on part of the nacelle, and (ii) a slide (102) which is connected to the mobile part and which can slide along the rail. The invention is characterized in that the guiding device is provided with a built-in retractable system for preventing the translational movement of the slide, for which purpose the device includes: (i) at least one blocking means (104) mounted such that it can move on the rail and occupy alternatively an engaged position in which it forms a translational abutment for the slide and a released position in which it is moved away from the slide and enables the relative translational movement of the slide beyond the locking system, and (ii) at least one means (130) for controlling the blocking means.

Abstract: A double-body gas turbine engine, including a low-pressure body LP and a high-pressure body HP, rotatably mounted about a single shaft in a stationary casing, the low-pressure body LP including a compressor and a turbine connected by a low-pressure shaft LP. The low-pressure shaft LP is supported by an upstream LP bearing, a first downstream LP bearing, and an additional downstream LP bearing by the stationary casing, the high-pressure body being supported by an upstream HP bearing and a downstream HP bearing which is an inter-shaft bearing including an inner track rigidly connected to the HP turbine rotor and an outer track rigidly connected to the LP shaft.

Abstract: A supersonic nacelle design employing a bypass flow path internal to the nacelle and around the engine is disclosed herein. A set of aerodynamic vanes may be used to shape a supersonic airflow within a bypass around an engine. The vanes may be capable of compressing the supersonic airflow into a subsonic airflow, direct the subsonic airflow around the engine, and then expand the subsonic airflow into a supersonic exhaust. The vanes may shape the airflow by reducing sonic boom strength, cowl drag, and airframe interference drag.

Abstract: A propulsion system and an aircraft that includes the propulsion system are provided. The propulsion system includes an engine and a cowling that surrounds at least a portion of the engine. The propulsion system also includes an inlet. The inlet includes a compression surface that is upstream and external to the cowling. The compression surface can slow air flow from a supersonic speed to a subsonic speed before the air flow enters the cowling.

Abstract: A noise reducing device in a jet engine that has a cylindrical casing, a cylindrical partition wall that is inserted in the casing while protruding partially from a trailing edge of the casing, and a compressor that compresses air that is taken into the cylindrical partition wall, with the inside of the cylindrical partition wall serving as a duct in which a core stream of high-speed air flows, and the space between the cylindrical partition wall and the casing serving as a duct in which a bypass stream of low-speed air flows, and being connected with a wing of an airplane by a pylon that has a projection portion that extends beyond the casing to the downstream of the core stream and the bypass stream, the noise reducing device includes a nozzle, disposed at the pylon on the downstream of the core stream, that injects a fluid toward a noise generation source that is produced from the mutual approach of an ambient air stream that is produced outside of the bypass stream and the core stream.

Abstract: A turbojet engine nacelle, including an air intake structure for directing an airflow towards a fan of the turbojet engine, a middle structure attached to the air intake structure for surrounding the fan of the turbojet engine, the air intake structure including an inner panel attached to the middle structure and defining a fixed structure with the middle structure, and an outer panel including an air intake lip at the end thereof opposite the middle structure, the outer panel being longitudinally translatably movable relative to the inner panel. The nacelle further includes at least one rod having one end attached onto a flange of the air intake lip and a mechanism rigidly connected to the middle structure for securing and locking the other end of each rod by applying a traction force thereon.

Abstract: A running-in coating for a compressor housing is disclosed. The running-in coating has at least two layers where a first layer is dimensionally stable and at least one additional layer has run-in capability.

Abstract: A gas turbine engine typically includes a fan section, a compressor section, a combustor section and a turbine section. A speed reduction device such as an epicyclical gear assembly may be utilized to drive the fan section such that the fan section may rotate at a speed different than the turbine section so as to increase the overall propulsive efficiency of the engine. In such engine architectures, a shaft driven by one of the turbine sections provides an input to the epicyclical gear assembly that drives the fan section at a speed different than the turbine section such that both the turbine section and the fan section can rotate at closer to optimal speeds providing increased performance attributes and performance by desirable combinations of the disclosed features of the various components of the described and disclosed gas turbine engine.

Abstract: An adaptive cycle gas turbine engine is disclosed having a number of features. A fan arrangement is provided having counter-rotating fan stages, one fan stage is operable to be clutched and decoupled from the other stage. A high pressure compressor bypass is also provided. A clutch is provided to at least partially drive an intermediate pressure compressor with a high pressure turbine when the high pressure compressor is bypassed. A partial bypass of the high pressure turbine may be provided.

Abstract: A ram air fan inlet housing for containing a ram air fan rotor. The inlet housing includes a flanged surface and an interior surface. The flanged surface is perpendicular to an axis of the inlet housing and defines a flange plane at an axial end of the inlet housing. The interior surface is symmetric about the axis of the inlet housing and includes a flange section, a transition section, an outlet section, a rotor section, and an inlet section.

Abstract: A blade for use in a gas turbine engine has an airfoil including a leading edge and a trailing edge. A sheath is positioned at the leading edge and secured to the airfoil by a first adhesive formed of a first material. The sheath is formed of a second material that is distinct from said first material. The first material is less electrically conductive than the second material. A grounding element is in contact with the sheath. The grounding element is in contact with a portion of the airfoil formed of a third material that is more electrically conductive than the first material. The grounding element and the portion of the airfoil together form a ground path from the sheath into the airfoil.

Abstract: A three-spool turbofan engine (20) has a variable fan nozzle (35). The fan blades have a peak tip radius RT and an inboard leading edge radius RH at an inboard boundary of the flowpath. A ratio of RH to RT is less than about 0.40.

Abstract: A gas turbine engine includes a core housing that includes an inlet case and an intermediate case that respectively provide an inlet case flow path and an intermediate case flow path. A geared architecture is arranged within the inlet case. A shaft provides a rotational axis. A hub is operatively supported by the shaft. A rotor is connected to the hub and supports a compressor section. The compressor section is arranged axially between the inlet case flow path and the intermediate case flow path. A bearing is mounted to the hub and supports the shaft relative to one of the intermediate case and the inlet case.

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

Abstract: An intermediate casing for a gas-turbine engine has an outer ring (3), an inner ring (1) and an intermediate ring (2) mutually supported by mutually aligned outer and inner supporting struts (4, 5). A radial drive shaft (6), positioned through inner and outer supporting struts (4.1, 5.1) aligned at a 6-o? clock position, connects to an auxiliary gear drive (7). Four additional outer supporting struts (5.2, 5.3, 5.4, 5.5) are spaced apart 72°. A mounting bracket connects to the outer ring at a side facing an aircraft fuselage in areas of two of the additional outer supporting struts. The arrangement of supporting struts and type of mounting bracket attachment, saves weight, improves flow in the bypass duct and increase maintenance and installation access at the bottom side of the engine to nearly 144°.

Abstract: A disclosed control system for a gas turbine engine includes a controller configured to set a position of the variable area fan nozzle according to a predetermined schedule of variable area fan nozzle positions corresponding to a flight operating condition. The schedule is determined in view of a relationship between a position of the variable area nozzle and a performance level of the engine at current flight conditions.

Abstract: A turbofan engine is disclosed and includes a fan and a compressor in communication with the fan section, a combustor, a turbine and a speed reduction mechanism coupled to the fan and rotatable by the turbine. The turbine includes a first turbine section that includes three or more stages and a second turbine section that includes at least two stages. A ratio of airfoils in the first turbine section to a bypass area is less than about 170.

Abstract: A turbine engine includes at least a compressor section and a turbine section, each having at least a first and second portion. A ratio of turbine section second portion stages to compressor section second portion stages is less than or equal to 1.

Abstract: An aircraft wing turbine engine is designed to increase the ground clearance for larger modernized engines mounted under the wings of an airfoil. To this end, the aircraft wing turbine engine includes a hot flow generator with a turbine engine rotating about a first central axis, a cold flow blower rotating about a second central axis, and a pod surrounding the hot flow generator and cold flow blower. The first and second central axes are offset and non-collinear with each other such that the cold flow blower can be moved farther away from the ground to increase the ground clearance of the turbine engine.

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

Abstract: Disclosed is a 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. 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. A moveable shroud can be used to open and close a passage of an ejector which can be used to assist in withdrawing a boundary layer or entrain an ambient air. 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 turbofan engine has an engine case and a gaspath through the engine case. A fan has a circumferential array of fan blades. The engine further has a compressor, a combustor, a gas generating turbine, and a low pressure turbine section. A speed reduction mechanism couples the low pressure turbine section to the fan. A bypass area ratio is greater than about 6.0. The low pressure turbine section airfoil count to bypass area ratio is below about 170.

Abstract: A turbofan engine includes a gas generator section for generating a gas stream flow with higher energy per unit mass flow than that contained in the ambient air and a power turbine that converts the gas stream flow into shaft power. The turbofan engine further includes a propulsor section including a fan driven by the power turbine through a geared architecture at a second speed lower than the first speed for generating propulsive thrust as a mass flow rate of air through a bypass flow path. An Engine Unit Thrust Parameter defined as net engine thrust divided by a product of the mass flow rate of air through the bypass flow path, a tip diameter of the fan and the first rotational speed of the power turbine is less than about 0.15 at a take-off condition.

Abstract: A propulsion system includes a fan, a gear, a turbine configured to drive the gear to, in turn, drive the fan. The turbine has an exit point, and a diameter (Dt) is defined at the exit point. A nacelle surrounds a core engine housing. The fan is configured to deliver air into a bypass duct defined between the nacelle and the core engine housing. A core engine exhaust nozzle is provided downstream of the exit point. A downstream most point of the core engine exhaust nozzle is defined at a distance from the exit point. A ratio of the distance to the diameter is greater than or equal to about 0.90.

Abstract: A thermal management system includes at least one heat exchanger in communication with a bypass flow of a gas turbine engine. The placement of the heat exchanger(s) minimizes weight and aerodynamic losses and contributes to overall performance increase over traditional ducted heat exchanger placement schemes.