Snecma

Safran Aircraft Engines

General Electric Company

Francois Gallet

Gilles Alain Charier

Wouter Balk

Nicolas Jerome Jean Tantot

Jean-Pierre Serey

Mathieu Ange Poisson

Alexandre Tan-Kim

Ivan Guy Rouesne

Boris Briantais

Michel Gilbert Roland Brault

Stephane Rousselin

Daniel Edward Mollmann

Craig William Higgins

Guillaume Glemarec

Jonathan Edward Coleman

Mark Edward Linz

Francois Robert Bellabal

Patrick Joyez

Valerio Capasso

Jean-Baptiste Manuel Nicolas Vignes

Audrey Grede

Nicolas Maurice Hervé Aussedat

Daniel Thomas Gerard Riviere

Michael Cline

Antoine Hellegouarch

Daniel E Mollman

Methods, apparatus, systems and articles of manufacture for mounting a gas turbine engine are described. An apparatus for mounting a gas turbine engine to a pylon, the gas turbine including an upstream section and a downstream section, the gas turbine defining a roll axis, a yaw axis, and a pitch axis, the apparatus including: a first mount to couple the upstream section of the gas turbine engine to the pylon; a second mount to couple the upstream section of the gas turbine engine to the pylon, the second mount downstream of the first mount; a thrust linkage to couple the upstream section to the pylon, wherein the downstream section is decouplable from the upstream section without decoupling the first mount, the second mount, and the thrust linkage.

Methods and apparatus for mounting a gas turbine engine

Counter-rotating turbine of a turbomachine extending about an axis of rotation and comprising an inner rotor rotating about the axis of rotation, and comprising at least one inner movable blade rotatably supported by a first shaft, an outer rotor rotating about the axis of rotation in a direction opposite to the inner rotor, and comprising at least one outer movable blade rotatably supported by a second shaft coaxial with the first shaft, the first and second shafts extending axially from upstream to downstream of the turbine, wherein the first shaft is guided in rotation by a first bearing disposed between the first shaft and an upstream casing of the turbine, and the second shaft is guided in rotation by a second bearing disposed between the second shaft and said upstream casing of the turbine.

Architecture of a turbomachine with counter-rotating turbine

The invention relates to an aircraft propulsion system () intended for being built into the rear of an aircraft fuselage, the propulsion system comprising at least two gas generators () supplying a power turbine () having two counter-rotating turbine rotors () for driving two fans (), and separate air inlets () for supplying each gas generator, characterised in that it comprises an electrical drive device () configured to rotate at least one of the turbine rotors, at least one electrical generator () configured to transform part of the energy of the flow from the gas generators into electrical power and an electric motor () supplied by said electrical generator and capable of rotating at least one of the turbine rotors, said electrical generator being installed on one of said gas generators, and in that said turbine rotor is capable of being rotated simultaneously by a flow from said gas generators and by the electrical drive device.

Aircraft propulsion system and aircraft powered by such a propulsion system built into the rear of an aircraft fuselage

A turbomachine including a stator, a rotor mounted on a shaft, a first bearing and a second bearing downstream of the first bearing in a direction of air circulation at the inlet of the turbomachine, the first bearing forming an axial abutment linked to the stator, the second bearing including an inner ring secured to the shaft and an outer ring mounted on an arm secured to the stator, the second bearing including rolling elements in axial support on the inner ring and on the outer ring. The outer ring includes a radially extending ring portion. The turbomachine also includes an axial force application device mounted between the arm and the ring portion.

Turbomachine with axial force adjustment at a bearing

The invention relates to an aircraft comprising a fuselage (), which is propelled by a turbine engine with two coaxial fans, namely an upstream fan () and a downstream fan (), driven by two contra-rotating rotors () of a power turbine (). The two fans () and the turbine () are integrated into a nacelle () which projects downstream from the fuselage () and through which air flows. According to the invention, at least one of the fans () of the aircraft and, in particular, the downstream fan () comprises variable-spacing blades, and at least one stator-forming variable-spacing blade ring () in the aircraft is placed upstream of the upstream fan (). The variable-spacing stator blades () and the variable-spacing blades of the downstream fan () are mutually configured to direct the air flow in a first mode in which the air flows through the nacelle () from upstream to downstream and in a second mode in which the air is pushed back upstream through the nacelle ().

Aircraft comprising two contra-rotating fans to the rear of the fuselage, with spacing of the blades of the downstream fan

A turbine blade including an airfoil () extending above a platform () and a root () extending below the platform, the root of the blade having the form of a radially extending arm (), being produced from composite material and including at least one cylindrical face (). The direction of the generatrices which are oriented in the direction of the platform to allow a clamping piece () to slide towards the latter on the cylindrical face under the action of centrifugal force. The arm includes, at its lower end, a protuberance () forming a radial bearing point for engagement with an element for retaining the blade on a turbine disc (). The root () is located entirely on the side of the protuberance with respect to the face.

Self-clamping fastener for CMC turbine blade

The invention concerns an aircraft propelled by a turbine engine having contrarotating fans (), the turbine engine being incorporated at the rear of a fuselage () of the aircraft, in the extension of same and comprising at least two gas generators () that supply, via a shared central stream (), a power turbine (), the turbine () comprising two contrarotating rotors () for driving two fans () disposed downstream from the gas generators (), said aircraft comprising means () arranged for separating the gas flow in the power turbine () into at least two concentric streams () and a device comprising first means for distributing the gas flow (-) between said streams () from the central stream (), the first distribution means being configured to be able to open or close the supply of at least one so-called sealable stream () of the streams () of the power turbine ().

Aircraft comprising a turbine engine incorporated into the rear fuselage with variable supply

The invention relates to a bypass turbojet comprising an exhaust housing having a central hub () and connecting means () that can transmit the forces generated by the turbojet to the structure of the aircraft that it propels, said connection means being two arms extending radially from the central hub in order to cross the cold flow of said turbojet and being characterized in that they are secured to said central hub and positioned in a diametrically opposed manner in relation to each other. An additional connection means () extends between the hub () and the area () for securing the exhaust housing to the structure () of the aircraft in order to transmit the exceptional dimensioning loads, said connection means being on standby during normal use, without any transmission of force between said hub and said area.

Isostatic suspension of a turbojet by rear double support

A bypass turbojet engine nacelle that forms a fan casing includes a first, upstream, streamlining element and a second streamlining element that forms a jet pipe. The second element can move translationally between a position ensuring aerodynamic continuity of the nacelle and a downstream position uncovering flow reversal openings. A thrust reverser device is housed in the nacelle and includes flow reverser flaps and cascades of vanes providing radial guidance for the flow. The cascades of vanes providing radial guidance for the flow can move in translation along the axis of the nacelle between a retracted position into the first steamlining element and an active flow-guidance position, the second streamlining element being secured to the cascades of vanes, the reverser flaps being mounted to rotate about axes that are transverse with respect to the axis of the nacelle and secured to the cascades of vanes.

Thrust reverser device for compact jet pipe

A variable geometry suspension structure for suspending a turboprop from a structural element of an aircraft, the structure including a rear cradle for fastening to a structural element of an aircraft and a front cradle for supporting a front portion of the turboprop, the front cradle being connected to the rear cradle, firstly via a pivot connection configured to enable the front cradle to tilt vertically relative to the rear cradle, and secondly by a return spring configured to prevent the front cradle from tilting while thrust from the turboprop is below a predetermined threshold.

Suspension structure with variable geometry of a turboprop engine on a structural element of an aircraft

A compressor rotor for a turbomachine including a shaft and at least two disks mounted on the shaft to hold a single set of vanes that are rotatably mobile about the rotational axis of the shaft, at least one first disk being mounted in a mobile manner on the shaft to be able to generate an angular gap in relation to the second disk and to trigger a rotation of the vanes about the radial axis of same, at least one of the two disks being shaped to receive at least one mechanism for attaching the vanes, the connection between the disk and the attaching mechanism enabling a rotation of the attaching mechanism about a radial axis, and connection is a ball-jointed connection that is tangentially and axially mobile.

Fan having a variable setting by means of differential rotation of the fan disks

A system for pylon attachment for mounting an engine on a structure of an aircraft, or to wings of the aircraft, is provided. The system includes a first attachment mechanism that can be secured to the pylon and a second attachment mechanism that can be secured to the structure of the aircraft. A connection mechanism connects the first and second attachment mechanism. The connection mechanism allows movement of the first attachment mechanism in relation to the second attachment mechanism between first and second positions.

Pylon for mounting an engine on the structure of an aircraft

A device for feeding fluid to a hydraulic actuator for controlling a pitch of fan blades of a two-propeller turboprop, the device including: a fluid manifold secured to a rotor of the turboprop and including a cylindrical portion including at least two fluid-circulation grooves, each presenting an outlet orifice; a fluid-admission support secured to a stationary portion of the turboprop and including a cylindrical portion including at least two distinct fluid passages opening to fluid-admission tubes and each leading radially into one of the fluid-circulation grooves of the manifold; and a tub constrained to rotate with the fluid manifold and including a cylindrical portion including at least two fluid-flow channels, each fed with fluid by a respective one of the outlet orifices of the fluid dispenser, each fluid-flow channel leading to a respective chamber of a control actuator.

Device for feeding fluid to a hydraulic actuator for controlling the pitch of the fan blades of a two-propeller turboprop

A turbine blade for a turbojet engine, made of a ceramic matrix composite material and including a vane that extends above a platform and a root that extends below the platform is provided. The pressure face and the suction face of the vane is formed by sheets of a composite material, the sheets meeting at the leading edge and trailing edge of the vane. The pressure face and suction face are formed by the same sheet bent back on itself, in a radial direction, at the leading edge or trailing edge.

High-pressure turbine blades made of ceramic matrix composites

A sealing device for passing a connecting rod of a system for controlling a pitch of fan blades of a turboprop through a partition. The device includes a tube for fastening to the partition that is to be sealed, and a frustoconical sheath through which the connecting rod is to pass, the sheath configured to slide axially inside the tube and including, at its wider end, a sealing mechanism co-operating with the tube, and, at its narrower end, a leaktight fastener fastening to a corresponding end of the connecting rod.

Sealing device having a sleeve for the passage of a connecting rod of a system for controlling the orientation of the blower blades of a turboprop engine through a partition

Growing community of inventors

Paris, France

Francois Gallet